ayload efinition ocument - european space...

AYLOAD EFINITION OCUMENT

prepared by/préparé par L. d’Arcio, Darwin Team, Darwin Technical Officers reference/réference SCI-A/2005/301/DARWIN/DMS/LdA issue/édition 1 revision/révision 2 date of issue/date d’édition 31 August 2005 status/état issued Document type/type de document

Technical Note

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DARWIN Payload Definition Document

Ref: SCI-A/2005/301/Darwin/DMS/LdA Date: 31 August 2005

Issue: 1.2 Page: 2

A P P R O V A L

Title titre

1 revision revision

2

author auteur

date date

approved by approuvé by

date date

C H A N G E L O G

date issue revision pages reason for change 23 jan 05 1 0 all First draft 2 jun 05 1 1 all Inputs A. Peacock, A. Karlsson 21 aug 05 1 2 Sections 11.1.2,

11.3 Updated milli and microthruster information

D I S T R I B U T I O N

name organisation A. Peacock Sci-A A. Karlsson Sci-A M. Fridlund Sci-SA M. Bavdaz Sci-A Ch. Erd Sci-A R. d. Hartog Sci-A N. Rando Sci-A L. d’Arcio Sci-A



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R E F E R E N C E D O C U M E N T S

RD 1 DARWIN, The infrared Space Interferometer, Concept and

feasibility study report, ESA-SCI(2000)12 Jul 2000

RD 2 DARWIN System Requirements Document, ASPI-00-OSM/IF.29 1.0 Oct 2000

RD 3 DARWIN Science Requirement Document SCI-SA Darwin 2005 :1 1.1 May 2005

RD 4 DARWIN Mission Requirement Document SCI-A/2005/287/Darwin/DMS 1.0 Apr 2005

RD 5 Darwin Prime Target Stars – planar TTN configuration SCI-A/2005/289/DARWIN/DMS/AS 1.0 Apr 2005

RD 6 Darwin Prime Target Stars- Emma configuration SCI-A/2005/290/DARWIN/DMS/AS 1.0 Apr 2005

RD 7 DARWIN TTN+ array assessment report SCI-A/2004/187/Darwin/DMS 1.3 Oct 2004

RD 8 DARWIN: Feasibility of H2O spectroscopy at 18-20 micron SCI-A/2004/348/Darwin/LdA 1.0 Dec 2004

RD 9 Darwin Variability Noise Sensitivity Analysis, SCI-A/2005/048/Darwin/LdA 1.0 Mar 2005

RD 10 M. Kilter, Micropropulsion Technology Assessment for Darwin, SCI-A/2003/296/DARWIN/MK1 5.0 Mar 2004

RD 11 Darwin Mission Analysis: Operational Phase, Transfer and Rendez-Vous, MAO Working Paper 2.2 Mar 2005

RD 12 Emma assessment, CR-DCS-TN10 1.1 May 2005



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T A B L E O F C O N T E N T S

1. INTRODUCTION 13

2. MISSION OVERVIEW 14 2.1. Scientific objectives................................................................................................... 14 2.2. Mission implementation ............................................................................................ 14

2.2.1 Three Telescope Nuller ......................................................................................... 14 2.2.2 Array configurations.............................................................................................. 15 2.2.3 Planar TTN ............................................................................................................ 19 2.2.4 Emma TTN............................................................................................................ 19 2.2.5 Trade-off between planar and “Emma” TTN ........................................................ 20

2.3. Mission phases........................................................................................................... 21

3. PAYLOAD DEFINITION AND OVERVIEW 23 3.1. Collector Spacecraft................................................................................................... 23

3.1.1 Functionality.......................................................................................................... 23 3.1.2 Description ............................................................................................................ 24

3.2. Beam combiner spacecraft......................................................................................... 26 3.2.1 Functionality.......................................................................................................... 26 3.2.2 Description ............................................................................................................ 26 3.2.3 Fine steering mirror ............................................................................................... 30 3.2.4 Optical Delay line.................................................................................................. 30 3.2.5 Differential Delay line........................................................................................... 30 3.2.6 Fringe Tracker ....................................................................................................... 31 3.2.7 Photometric sensor ................................................................................................ 31 3.2.8 Modulator .............................................................................................................. 32 3.2.9 Dichroic device...................................................................................................... 32 3.2.10 Beam splitters.................................................................................................... 32 3.2.11 Achromatic phase shifters ................................................................................. 32 3.2.12 Beam shaper ..................................................................................................... 33 3.2.13 SMW coupling unit ........................................................................................... 33 3.2.14 Single Mode Waveguide ................................................................................... 33 3.2.15 Spectrograph ..................................................................................................... 33 3.2.16 Detector ............................................................................................................. 34 3.2.17 Focal Plane Assembly and Cooler .................................................................... 34 3.2.18 BCA implementation ........................................................................................ 34

4. TELESCOPE SPACECRAFT PAYLOAD 37 4.1. Main telescope ........................................................................................................... 37

4.1.1 Functionality.......................................................................................................... 37 4.1.2 Description ............................................................................................................ 37 4.1.3 Performance........................................................................................................... 41 4.1.4 Interface and Physical Resource Requirements..................................................... 43 4.1.5 Open Points and Critical Issues ............................................................................. 44

4.2. Send transfer optics.................................................................................................... 44



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4.3. Wide Field Camera .................................................................................................... 47 4.3.1 Functionality.......................................................................................................... 47 4.3.2 Description ............................................................................................................ 48 4.3.3 Performance........................................................................................................... 49 4.3.4 Interface and Physical Resource Requirements..................................................... 50 4.3.5 Open Points and Critical Issues ............................................................................. 50

5. BEAM COMBINER SPACECRAFT PAYLOAD 51 5.1. Receive Transfer Optics............................................................................................. 51


5.2. Fine steering mirror ................................................................................................... 55 5.2.1 Functionality.......................................................................................................... 55 5.2.2 Description ............................................................................................................ 55 5.2.3 Performance........................................................................................................... 56 5.2.4 Interface and Physical Resource Requirements..................................................... 57

5.3. Optical Delay Lines ................................................................................................... 58 5.3.1 Functionality.......................................................................................................... 58 5.3.2 Description ............................................................................................................ 58 5.3.3 Performance........................................................................................................... 59 5.3.4 Interface and Physical Resource Requirements..................................................... 61

5.4. Differential Delay Line Unit...................................................................................... 62 5.4.1 Functionality.......................................................................................................... 62 5.4.2 Description ............................................................................................................ 62 5.4.3 Open Points and Critical Issues ............................................................................. 62

5.5. Fringe sensor.............................................................................................................. 63 5.5.1 Functionality.......................................................................................................... 63 5.5.2 Description ............................................................................................................ 64 5.5.3 Performance........................................................................................................... 65 5.5.4 Interface and Physical Resource Requirements..................................................... 66 5.5.5 Open points and critical issues .............................................................................. 69

5.6. Modulator .................................................................................................................. 69 5.6.1 Functionality.......................................................................................................... 69 5.6.2 Description ............................................................................................................ 69 5.6.3 Performance........................................................................................................... 70 5.6.4 Interface and Physical Resource Requirements..................................................... 72 5.6.5 Open issues and critical items ............................................................................... 72

5.7. Beam Splitters and dichroic devices .......................................................................... 73 5.8. Achromatic Phase Shifter .......................................................................................... 73

5.8.1 Functionality.......................................................................................................... 73 5.8.2 Description ............................................................................................................ 73



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5.8.3 Performance........................................................................................................... 75 5.8.4 Interfaces and Physical Resource Requirements ................................................... 77

5.9. Beam shaper............................................................................................................... 78 5.9.1 Functionality.......................................................................................................... 78 5.9.2 Description ............................................................................................................ 78 5.9.3 Performance........................................................................................................... 80 5.9.4 Interface and Physical Resource Requirements..................................................... 80

5.10. Single Mode Waveguide Coupling Unit.................................................................... 81 5.11. Single Mode Waveguide............................................................................................ 81

5.11.1 Subsystem description....................................................................................... 81 5.11.2 Functionality ..................................................................................................... 81 5.11.3 Description ........................................................................................................ 82 5.11.4 Performance ...................................................................................................... 83 5.11.5 Interface and Physical Resource Requirements ................................................ 84

5.12. Spectrograph .............................................................................................................. 84 5.12.1 Functionality ..................................................................................................... 84 5.12.2 Description ........................................................................................................ 85 5.12.3 Performance ...................................................................................................... 88 5.12.4 Interface and Physical Resource Requirements ................................................ 88 5.12.5 Open Points and Critical Issues......................................................................... 89

5.13. FIR Linear Detector Array......................................................................................... 89 5.13.1 Description ........................................................................................................ 89 5.13.2 Performance ...................................................................................................... 90 5.13.3 Interface and Physical Resource Requirements ................................................ 90

5.14. Sorption Cooler.......................................................................................................... 91 5.14.1 Functionality ..................................................................................................... 91 5.14.2 Description ........................................................................................................ 91 5.14.3 performance ...................................................................................................... 92 5.14.4 Interface and Physical Resource Requirements ................................................ 94

6. ON-BOARD PROCESSING AND AUTONOMY 96

7. CALIBRATIONS AND CHECKOUT 98 7.1. Ground Calibration .................................................................................................... 98 7.2. In-flight Calibration ................................................................................................... 99

8. OPERATIONS 99 8.1. Baseline Operations ................................................................................................... 99

8.1.1 Detection phase ................................................................................................... 100 8.1.2 Spectroscopy phase ............................................................................................. 100

8.2. Telemetry Estimate .................................................................................................. 101

9. APPENDIX A: TABLES 102

10. APPENDIX B: TECHNOLOGY DEVELOPMENT STATUS 104 10.1. Development status: overview................................................................................. 104 10.2. Optical Delay Line................................................................................................... 105



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10.2.1 Development status ......................................................................................... 105 10.2.2 Performance .................................................................................................... 109 10.2.3 Interface and Physical Resource Requirements .............................................. 110 10.2.4 Open Points and Critical Issues....................................................................... 112

10.3. Fringe sensor............................................................................................................ 114 10.3.1 Development status ......................................................................................... 114 10.3.2 Functionality ................................................................................................... 114 10.3.3 Description ...................................................................................................... 114 10.3.4 Performance .................................................................................................... 118 10.3.5 Open Points and Critical Issues....................................................................... 120

10.4. Beam Splitters and dichroic devices ........................................................................ 122 10.4.1 Development status ......................................................................................... 122

10.5. Achromatic phase shifters........................................................................................ 122 10.5.1 Development status ......................................................................................... 122 10.5.2 Functionality ................................................................................................... 122 10.5.3 Description ...................................................................................................... 124 10.5.4 Performance .................................................................................................... 124 10.5.5 Interface and Physical Resource Requirements .............................................. 126

10.6. Single Mode Waveguide Coupling Unit.................................................................. 127 10.6.1 Development status ......................................................................................... 127

10.7. Single Mode Waveguide......................................................................................... 127 10.7.1 Development status ......................................................................................... 127 10.7.2 Single Mode Fibers for Darwin (TPD/TNO) .................................................. 128 10.7.3 Single Mode Fibers for Darwin (Astrium)...................................................... 134

10.8. Integrated Optics for Darwin ................................................................................... 140 10.8.1 Development status ......................................................................................... 140 10.8.2 Description ...................................................................................................... 140 10.8.3 Performance .................................................................................................... 143 10.8.4 Interface and Physical Resource Requirements .............................................. 143 10.8.5 Open Points and Critical Issues....................................................................... 144

10.9. FIR Linear Detector Array....................................................................................... 145 10.9.1 Development status ......................................................................................... 145 10.9.2 Description ...................................................................................................... 145 10.9.3 Functionality ................................................................................................... 146 10.9.4 Performance .................................................................................................... 148 10.9.5 Interface and Physical Resource Requirements .............................................. 149 10.9.6 Open Points and Critical Issues....................................................................... 149

10.10. Sorption cooler .................................................................................................... 150 10.10.1 Development status ......................................................................................... 150 10.10.2 Activity overview............................................................................................ 150 10.10.3 Performance .................................................................................................... 156 10.10.4 Interface and Physical Resource Requirements .............................................. 158 10.10.5 Open Points and Critical Issues....................................................................... 160

11. APPENDIX C: FORMATION FLYING COMPLEMENT 162 11.1. FF complement overview ........................................................................................ 162

11.1.1 Metrology........................................................................................................ 162 11.1.2 FF propulsion .................................................................................................. 165 11.1.3 Control aspects ................................................................................................ 172



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11.2. FF complement: overview of development status ................................................... 175 11.3. Formation Flying RF Metrology Technology Development ................................... 176

11.3.1 Development status ......................................................................................... 176 11.3.2 Functionality ................................................................................................... 176 11.3.3 Description ...................................................................................................... 177 11.3.4 Performance .................................................................................................... 179 11.3.5 Interface and Physical Resource Requirements .............................................. 184 11.3.6 Open Points and Critical Issues....................................................................... 185

11.4. High-Precision Optical Metrology (HPOM) Technology Development ................. 185 11.4.1 Development status ......................................................................................... 185 11.4.2 Functionality ................................................................................................... 186 11.4.3 Description ...................................................................................................... 187 11.4.4 Performance .................................................................................................... 193 11.4.5 Interface and Physical Resource Requirements .............................................. 193 11.4.6 Open Points and Critical Issues....................................................................... 194

11.5. RITA-10 mN ion thrusters Technology Development ............................................ 194 11.5.1 Activity overview............................................................................................ 194 11.5.2 Functionality ................................................................................................... 195 11.5.3 Description ...................................................................................................... 196 11.5.4 Performance .................................................................................................... 198 11.5.5 Interface and Physical Resource Requirements .............................................. 199 11.5.6 Open Points and Critical Issues....................................................................... 199

11.6. T-5 mN ion thrusters Technology Development ..................................................... 200 11.6.1 Activity overview............................................................................................ 200 11.6.2 Functionality ................................................................................................... 201 11.6.3 Description ...................................................................................................... 203 11.6.4 Performance .................................................................................................... 205 11.6.5 Open Points and Critical Issues....................................................................... 206

11.7. RMT mN ion thrusters Technology Development .................................................. 207 11.7.1 Activity overview............................................................................................ 207 11.7.2 Functionality ................................................................................................... 207 11.7.3 Description ...................................................................................................... 208 11.7.4 Performance .................................................................................................... 210 11.7.5 Interface and Physical Resource Requirements .............................................. 211 11.7.6 Open Points and Critical Issues....................................................................... 211

11.8. ALTA mHall Thruster ............................................................................................. 212 11.8.1 Executive Summary ........................................................................................ 212 11.8.2 Functionality ................................................................................................... 212 11.8.3 Description ...................................................................................................... 214 11.8.4 Performance .................................................................................................... 216 11.8.5 Interface and Physical Resource Requirements .............................................. 217 11.8.6 Open Points and Critical Issues....................................................................... 218

11.9. RIT µN ion thrusters Technology Development ..................................................... 218 11.9.1 Activity overview............................................................................................ 218 11.9.2 Functionality ................................................................................................... 219 11.9.3 Description ...................................................................................................... 220 11.9.4 Performance .................................................................................................... 222 11.9.5 Interface and Physical Resource Requirements .............................................. 223 11.9.6 Open Points and Critical Issues....................................................................... 224



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11.10. In µN FEEP thrusters Technology Development ................................................ 224

11.10.1 Functionality ................................................................................................... 224 11.10.2 Description ...................................................................................................... 224 11.10.3 Performance .................................................................................................... 227 11.10.4 Interface and Physical Resource Requirements .............................................. 227 11.10.5 Open Points and Critical Issues....................................................................... 227

11.11. ALTA µN FEEP thrusters Technology Development........................................ 228 11.11.1 Activity overview............................................................................................ 228 11.11.2 Functionality ................................................................................................... 228 11.11.3 Description ...................................................................................................... 229 11.11.4 Performance .................................................................................................... 233 11.11.5 Interface and Physical Resource Requirements .............................................. 234 11.11.6 Open Points and Critical Issues....................................................................... 234

11.12. Guidance, Navigation and Control (ICC) Technology Development ................. 235 11.12.1 Activity Overview........................................................................................... 235 11.12.2 Functionality ................................................................................................... 235 11.12.3 Description ...................................................................................................... 236 11.12.4 Performance .................................................................................................... 237 11.12.5 Interface and Physical Resource Requirements .............................................. 238 11.12.6 Open Points and Critical Issues....................................................................... 238

11.13. Guidance, Navigation and Control II: Interferometer Deployment (ICD) Technology Development ..................................................................................................... 239

11.13.1 Activity overview............................................................................................ 239 11.13.2 Functionality ................................................................................................... 240 11.13.3 Description ...................................................................................................... 240 11.13.4 Interface and Physical Resource Requirements .............................................. 241 11.13.5 Open Points and Critical Issues....................................................................... 242



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A C R O N Y M N S

APCS Attitude and Position Control System APS Achromatic Phase Shifter BCA Beam Combiner Assembly AR AntiReflection BCM Baseline Control Mode BCS Beam Combiner Spacecraft BSC Beam shaper & SMW coupling CGMT Cold Gas Micro-Thruster CM Cruising Mode CS Collector Spacecraft DM Deployment mode DSx Detector Spacecraft x EGSE Electrical Ground Support Equipment EM Engineering Model FEEP Field Emission Electrical Propulsion FAM Fringe Acquisition Mode FEEP Field Emission Electrical Propulsion FFS Fine Fringe Sensor FM Flight Model FoV Field of View FS Fringe Sensor FTS Fine tilt sensor HEO Highly Elliptical Orbit HZ Habitable Zone ISD Inter Satellite Distance L2 The second Earth-Sun libration point LGA Low Gain Antenna LoS Line of Sight LZ Local Zodiacal light mas Milliarcsecond (4.848 × 10-9 rad) MMZ Modified Mach Zender MS Metrology Sensor NOM Normal Observation Mode OCM Orbit Correction Mode ODL Optical Delay Line



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OGSE Optical Ground Support Equipment OPD Optical Path Difference OPL Optical Path Length pc Parsec (astronomical unit equivalent to ~ 3.10 13 km or ~ 3.3 light-

years) PCU Power Control Unit PLM Payload Module PM Propulsion Module RF Radio Frequency RM Reconfiguration Mode RMT Radio-frequency with Magnetic-field ion Thruster Rms Root, Mean ,Square RR Retro-reflectors ROIC Readout Integrated Circuit RSx Relay Spacecraft X SMW Single Mode Waveguide SNR Signal to Noise Ratio STM Structural Thermal Model STR Star Tracker SVM Service Module TBC To Be Confirmed TBD To Be Determined TE-SAT Terrestrial Exo-planets Science Advisory Team TM Transmission Map TRP Technology Research Programme TTN+ Three Telescope Nuller (+ indicate dual output modulation) WFC Wide Field Camera wrt With respect to



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Definitions Continuously Habitable Zone, CHZ: The region where conditions have been benevolent to life as we know it for all of the Sun’s evolutionary history: Currently believed to be 0.9 to 1.1 AU-equivalent distance scaled by the square root of the stellar luminosity (Sun=1).

Habitable Zone, HZ is defined as the area where liquid water could exist on a planetary surface of terrestrial type orbiting a solar type star. The Habitable zone around a star is a uniform zone with radii 0.7 to 1.5 (1.9 TBD) AU-equivalent [RD9]. Eccentricity (TBD). ηEarth : The percentage of solar-type stars that have Earth-like planets. Giant planet: A planet with a mass larger than 15 times the Earth. InterSatellite Distance, ISD: The distance between the geometric centres of two satellites, see [RD 3]. The geometric centre of a telescope flyer shall be the intersection of the telescope line of sight and the transfer optics line of sight, while the geometric centre of the hub shall be the intersection of the transfer optics line of sight. ISDx,y shall refer to the intersatellite distance between spacecraft x and y, where x ∈ {H, A, B, C) and y ∈ {A, B, C). Modulation Map, MM: is the arithmetic difference between two transmission maps. Nulling: Destructive interference by application of achromatic phase shifts in on or more arms of an interferometer. Planetary system: Defined as a system of multiple planets orbiting a star. Prevalence of terrestrial type planets is the number of rocky planets per solar type star). In the solar system this number is > 3 (Venus, Earth and Mars, with Mercury optional). Rejection Ratio: Ratio between the flux transmitted through an interferometer and the flux transmitted by the same interferometer, when adjusted to null the light. (RR > 1) Solar type star: is F5 – K9 (same order of evolutionary time scales, no bound rotation in HZ, equivalent to 0.43 < B-V< 1.31 Terrestrial planet: A minimum of 0.5 Earth radii and a maximum of 2 Earth radii. Earth albedo, Earth emissivity, Effective temperature between 260K and 373K. A density > 3g cm-3 and smaller than 7g cm-3 (TBC). Transmission map, TM: The response of an nulling interferometer as a function of wavelength and source coordinates in reference to the telescope plane.



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1. INTRODUCTION

The Darwin Payload Definition Document (PDD) defines the reference payload design for the ESA Darwin mission, and the current status of associated technology development. This PDD applies to the current Darwin baseline configuration, the planar Three Telescope Nuller (TTN), which was introduced mid 2004 [RD 7]. TTN is an array of three 3.15 m telescope s/c and a beam combiner hub, to be launched by two Soyuz-Fregat rockets. The spacecraft can be arranged in two configurations, triangular and linear, that are optimized for planet spectroscopy and detection, respectively. TTN features a multi-beam injection scheme where beams are combined in the image-plane, and simultaneously injected into the single mode fibre [RD 7]. In this scheme, the fiber acts simultaneously as beam combiner, and modal filter. The content of this PDD is based on:

• internal ESA work on TTN, with special reference to the TTN+ Mission Design Assessment [RD 7] describing the planar TTN array.

• A 3D TTN architecture (the so-called Emma TTN) is also being evaluated [RD 12].

The trade-off between the planar and 3D architectures is TBD. However, most technology development activities presented here are relevant for Emma TTN as well.

• Darwin requirement documents, especially the Science Requirement Document (SRD,

[RD 3]) and Mission Requirement Document (MRD, [RD 4]) • the result of the Alcatel Darwin system study 2000 [RD 2]. This study defined the

original baseline Darwin design, the hexagonal array, based on six 1.5 m free-flying telescope spacecraft and one beam-combining spacecraft, possibility augmented by a spacecraft dedicated to ground communication and metrology. The beam combiner and the telescope spacecraft fly in one plane with the telescope spacecraft at equal distances from the beam combiner. The Alcatel study is included for completeness, since many of the technology studies described here were initiated when the hexagonal array was the baseline design.

• the results of completed and ongoing Darwin technology development activities • the results of the Darwin mission analysis by ESOC [RD 11]

This document is compiled in order to provide input to the Darwin Mission Assessment study planned to begin mid 2005.



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2. MISSION OVERVIEW

2.1. Scientific objectives The main scientific objective of the Darwin mission is to detect signs of earth-like life outside the solar system. This leads to the following top-level science requirements [RD 3]

1. To survey nearby stars for terrestrial planets (i.e. planets similar to Venus, Earth and Mars in our Solar System), orbiting within the habitable zone.

2. To study in detail any found planets as what concerns composition, orbital mechanics,

geophysical conditions (atmospheric composition, evolutionary status), and to determine its evolutionary status.

3. To search any planetary atmosphere detected for signs of what are commonly called bio-

markers (i.e. signs of biological activity as we know it), or their pre-cursors. During the mission a large number of exo-solar systems, of different ages and initial parameters shall be studied. This shall provide a sufficient sample of planetary systems to allow a comparison, with each other but also with our own Solar system. It is expected that this shall increase the understanding and the scientific models concerning the causes/conditions of planetary system formation, as well as evolutionary pathways. There is a second mission objective augmenting the science capabilities of the high-resolution astrophysical imaging. This option has not been considered in the PDD at this stage, as its implementation is not supposed to drive the mission design. The design of the high resolution imaging capability should be based on the system design as mandated by the nulling interferometer.

2.2. Mission implementation

2.2.1 THREE TELESCOPE NULLER

The current baseline implementation of Darwin is the planar Three Telescope Nuller (TTN) [RD 7]. This is a nulling interferometer consisting of 4 spacecraft: three Collector Spacecraft, CS, each carrying an identical telescope assembly, collecting the light from the science target and relaying this light to a Beam Combiner spacecraft, BCS. In the BCS, the three incoming beams are matched in intensity and phase and combined under suitable achromatic phase shifts, resulting in destructive interference on-axis (at the star location) and at the same time, constructive interference at possible planet locations. In the TTN design, the light is combined using multi-axial beam combination with a single mode waveguide, a new nulling technique described in [RD 7]. This technique allows deep nulling and efficient internal modulation with three beams. It has been demonstrated that a 3 spacecraft mission (where the beam combination optics is lumped at one of the CS) could be



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implemented using this nulling technique. However, this would a) lead to a complex beam relay scheme and b) not allow efficient use of two launch vehicles. It is to be assumed for the purpose of this assessment study that the beam combination is to be implemented on a separate beam combiner spacecraft (BCS) resulting in a 4 s/c configuration. In order to separate the signal of the planet from the background, it is necessary to modulate the signal from the planet and that from the zodiacal dust at different frequencies. This is done either by switching between different sets of achromatic phase shifts or by moving the individual telescopes, see Section 3.2.8 for further details. Accurate matching of the optical properties of the interfering beams is critical for a nulling interferometer. The rejection ratio for the stellar photons will be limited by (among other effects) polarization and amplitude mismatching and phase defects. Using current optical technology, the required rejection rates of 105 – 106, cannot be reached unless the interfering wave fronts are filtered. Wavefront filtering will be implemented by propagation in a single mode wave-guide (SMW). Further to the beam combiner assembly, the BCS carries the detector assembly and the associated cooler. Additional equipment as the X-band transponder and the medium gain antenna, used for nominal communication with ground, are also accommodated in the hub, as well as major components of the optical metrology subsystems. The ability to “fly” several spacecraft at precise distance and attitude is mandatory for Darwin. Challenges are to be found in the field of precise and independent measurements, control system architecture and actuation system. For the metrology system both radio frequency ranging and high precision laser technology is foreseen. For the “nulling mission” there is always a central star towards, which the interferometer and its individual telescopes are pointing. Nevertheless, the spacecrafts’ positions (and rotation around the line of sight) relative to the BCS have to be monitored. The control system will have to collect several measurements on different spacecraft and control the individual spacecraft of the interferometer. For the actuation highly precise and proportional micro-propulsion units will be used, for example based on field emission electric propulsion (FEEP) technology. When several spacecraft are flown in close formation there is an apparent risk for spacecraft collisions, which, in the worst case, could put the mission to a premature end. It is therefore of utmost importance that spacecraft collisions are avoided and that the spacecraft formation can be formed from an almost arbitrary distribution in space. The latter is important when the formation is detached from the launch vehicle and deployed, and after contingencies when the control of the formation has been interrupted.

2.2.2 ARRAY CONFIGURATIONS

The response of the interferometer in the vicinity of the target star is determined primarily by the relative position of the collector s/c as seen from the target, and by the set of achromatic delays imparted to the various beams. Two configurations have been studied in the detail so far (see Figure 1):

- the linear TTN (LinTTN), where the three collector s/c are equally spaced along a line

- the equilateral TTN, where the three collector s/c are at the vertexes of an equilateral triangle

The corresponding modulation maps are shown in Figure 2.



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A

B

CISDA,B

A

B

C

ISDA,B ISDA,hub

Figure 1 Linear TTN (left) and equilateral TTN (right). The two arrays are sized to resolve the same planet at a certain wavelength and angular distance from the star. The three collector s/c are denoted as A,B,C. Note that the distance ISDAB between contiguous collector s/c is twice as large for the equilateral TTN than for Lin TTN. The linear TTN (right) has the three TS on a straight line and the BCS in the same plane orthogonal to the pointing direction, at a right angle with respect to the central TS. Note that the central TS is closer to the BCS than the outer s/c. The equilateral TTN has the BCS in the centre and the three TS at equal distance from the BCS, separated by 120º. The telescopes are pointed to the same target and relay the collected light to the BCS, as indicated by red arrows.

Figure 2 Modulation map generated by switching between the two transmission maps for the triangular and linear configuration of the TTN configuration (FoV 2.5AU at10 µm). Note that green shows the zero modulation level



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In the detection phase, and possibly also in the spectroscopy phase of the mission, the interferometer will be rotated around the line of sight (LoS) during observations. This rotation will modulate the planet signal allowing its detection against background signals, and discrimination of multiple planets. The LinTTN modulation map exhibits a single primary maximum. The peak modulation efficiency is 93.3 %, while the mean value and standard deviation of the modulation efficiency within the field of view are 44.6% and 33.9%, respectively. The planet signal is essentially modulated at the array rotation frequency. Unambiguous detection with LinTTN requires a rotation by 180º. The modulation map of the equilateral TTN has three well-separated primary peaks, again with a modulation efficiency of 93.3 % and standard deviation of 33.9%. The mean value of the modulation efficiency within the field of view is 28%, which is significantly lower than for LinTTN. The planet signal is modulated essentially at the third harmonic of the array rotation frequency. The equilateral triangular configuration requires a 60º rotation only. The LinTTN clearly outperforms the equilateral TTN for planet detection since

- it has a higher modulation efficiency - the modulation efficiency is rather uniform through the whole Habitable Zone. This is

an asset in order to detect a planet, whose location inside the habitable zone is a priori unknown, with as little reconfiguration as possible.

The two configurations have a comparable performance for spectroscopy. However, the equilateral configuration has superior imaging capabilities, i.e., a more compact (about 2 times narrower) PSF, see Figure 3. This is advantageous for planet localization and for disentangling multiple planets. Therefore, the configuration of choice for spectroscopy would be the equilateral TTN. The Lin and Equilateral TTN are two extreme cases of the generic triangular array, whose properties have not been explored systematically. A limited study has been performed of the so-called 90 degree array, where two collector s/c are arranged at three of the corners of a square, and whose properties are intermediate between those of the Lin en equilateral TTN In general, when the collectors form a non-equilateral triangle, the modulation map will be asymmetric as shown in Figure 4. The absence of rotational symmetry makes it easier to locate the planet.



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Figure 3 Correlation or PSF maps for equilateral TTN (left) and linear TTN (right). The outer limit corresponds with the FoV, and the ratio of baseline to wavelength was scaled such that the planet appears in the same position inside the FoV for both configurations.

Figure 4 The modulation map of non-equilateral triangular configurations show asymmetries in the modulation map, allowing for better determination of the planet’s location. Note that green shows the zero modulation level.



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2.2.3 PLANAR TTN

Two different TTN architectures are being considered, which differ significantly in term of achievable sky coverage, intersatellite distance requirements, collector satellite and relay optics implementation. The planar TTN, where the spacecraft fly in a common plane, is shown in Figure 5.

Figure 5: Planar versions of the equilateral TTN (left) and linear TTN (right)

The planar configuration has the main advantage of minimizing radiative transfer between s/c, since all hot thermal shields are nominally in the same plane. The instantaneous sky coverage of the planar TTN is the ± 45 degrees cone centered in the antisun direction. The science signal is relayed from the collector s/c to the hub as collimated beams. The distance between s/c A and B (or B and C) needed for observations ranges between 15 m and 85 m for LinTTN, and between 15 m and 170 m for the Equilateral TTN. The shortest s/c distance is set by collision avoidance considerations, and corresponds to an edge-to-edge distance between sunshields of 5 m, when assuming 10 m sunshields.

2.2.4 EMMA TTN

The Emma TTN architecture is three-dimensional. The three collector s/c are located in a parabolic 3D surface, and the BCS is at the focus of the paraboloid, see Figure 6. Preliminary trade-off show indicates that the optimum focal length of the paraboloid will be in the order of 1 km. The collector optics consists simply of a spherical mirror, whose focal length matches that of the underlying paraboloid, and that also serves as beam transfer optics to the hub. The instantaneous sky coverage of Emma is a hollow cone covering approximately angles between 45 and 71.8 degrees (TBD) from the antisun direction. However, almost the whole sky can be accessed during one year.



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α

Sun

Science target

BCS

CS 1

CS 2

CS 3

αα

Sun

Science target

BCS

CS 1

CS 2

CS 3

Figure 6 Emma configuration. The three (or four) TS are located in a parabolic 3D surface, and the BCS is at the focus of the paraboloid. The TS are located along a circle of radius variable between 15 m – 100 m, forming a triangle or a trapezoid. The range between TS-BCS is about 1 km.

The distance between s/c A and B (or B and C) needed for observations ranges between 7 m (TBC) and 85 m for LinTTN, and between 7 m (TBC) and 170 m for the Equilateral TTN. The shortest s/c distance is set by collision avoidance considerations, and corresponds to an edge-to-edge distance between sunshields of 5 m, when assuming 5 m sunshields. Note that arrays with 4 or more collectors can be also implemented according to the Emma principle.

2.2.5 TRADE-OFF BETWEEN PLANAR AND “EMMA” TTN

The main potential attractiveness of the Emma design as compared to the planar design are:



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- simple collector s/c implementation: a spherical mirror replaces the telescope and the relay optics of the planar architecture, improving mass and volume budgets, and optomechanical stability

- the hub serves as an out-of-plane reference for metrology, which simplifies the control of the array geometry

- better sky access, targets are available for up to six months per year

- the minimum distance between collector s/c can be reduced to about 7 m (TBD) since sunshields are much smaller than in the planar case. As compared to the planar array, the array can be operated in more compact configurations, which reduces geometrical leakage and improves the snr at short wavelengths.

Conversely, Emma requires a more complex beam receive optics at the beam combiner level. Moreover, beams will be effected by rather large wavefront errors when the array is operated with long baselines (i.e., for distant and/or cold targets). A Mission Design Assessment is available for the planar array [RD 7]. A study of TTN/Emma, with a more limited scope, is currently being performed [RD 12]. A detailed trade-off between the two architecture options is not available yet.

2.3. Mission phases The science operations time of the mission is 5 years, this excludes in-orbit commissioning and the transfer to the operational orbit around the second Lagrange point. The 5 years are tentatively1 subdivided into 2 years planet detection phase and 3 years of planet spectroscopy phase. The launch marks the beginning of the Launch and Early Operations phase (LEOP). The LEOP phase, as described in [RD 7], is estimated to be completed in 2 weeks. By the end of LEOP the spacecrafts are separated from the upper stage, the formation deployed, the spacecrafts are sun pointing and on a trajectory for free injection into a Lissajous orbit around the second Lagrange point. The sunshields are deployed and the cooling down of the payload module has begun (assisted by formation and sunshield deployment). The cruise phase to L2 will last for approximately 100 days. During this time the payload module should reach its operational temperature of 40 K (TBC). During the cruise phase the spacecraft subsystems will be checked out, including the metrology sub-systems and the formation flying capability. The last metrology system to be checked for performance will be the fringe sensor. When this has been completed, and the payload modules have reached the operational temperature, the scientific commissioning phase can commence. The duration of the commissioning phase shall be less than 3 months.

1 The duration of the spectroscopy phase may be reduced, depending on the number of detected possible earth-like planets during the detection phase.



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When the instrument’s scientific performance has been established the science operations phase shall begin, which shall have a nominal duration of 5 years. There are basically two tasks to be completed during the science operations phase: planet detection and spectroscopy of detected planets. The amount of time that can be dedicated to the latter obviously depends on the results from the former, i.e. if few planets are detected during the first phase only little time needs to be allocated for spectroscopy of these few planets, and vice-versa. In this study it is assumed that 2 years of the science operations phase is required for planet detection and the remaining 3 years for spectroscopy. It should be noted that formation reconfigurations, e.g. slews and resizings (change of baselines), and fringe acquisition will require a significant fraction of the time during the science operations phase. It is assumed that science duty cycle (fraction of observation time as compared to elapsed time) is more than 70% during the science operations phase of the mission. The mission should not be restricted to the nominal lifetime by exhaustion of consumables. This has influenced the choice of micro-propulsion technology (FEEP) and for detector cooling (sorption cooler). It is therefore possible to extended the mission once the nominal mission lifetime has been reached.



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3. PAYLOAD DEFINITION AND OVERVIEW

3.1. Collector Spacecraft

3.1.1 FUNCTIONALITY

The function of the Collector Spacecraft (CS) optical payload is to collect flux from an observed star and its (expected) planets, and to transmit this flux to the central hub, where nulling interferometry is performed by combining the incoming beams under prescribed phase and intensity prescriptions.

M1

M2

M3

- one field intermediateimage)

Field stop

Folding mirror

Laser beamcoming from the hub

Folding mirror

Field stop

Transmitter:

-two off-axis parabolic mirrors- one field stop at intermediate image

Science Telescope:

- Korsch type - parabolic primary f-no ~ 1- Distance M1-M2 ~ 3.15 m- Diameter M1: 3150 mm- Diameter M2: 630 mm- Diameter M3: 84 mm

Wide Field Camera- relay optics - 2D detector- fov ~ 2.7 arcmin

M1

M2

M3

- one field intermediateimage)

Field stop

Folding mirror

Laser beamcoming from the hub

Folding mirror

Field stop

Transmitter:

-two off-axis parabolic mirrors- one field stop at intermediate image

Science Telescope:

- Korsch type - parabolic primary f-no ~ 1- Distance M1-M2 ~ 3.15 m- Diameter M1: 3150 mm- Diameter M2: 630 mm- Diameter M3: 84 mm

Wide Field Camera- relay optics - 2D detector- fov ~ 2.7 arcmin

Figure 7 Schematics of the optical payload of the Collector s/c. Red rays indicate on-axis light, being sent to the hub. Blue and green rays indicate light from off-axis sources, being sent to the Wide Field Camera.



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Figure 8 Close up of the field selector device. Blue and green rays indicate light from off-axis sources, being sent to the Wide Field Camera.

3.1.2 DESCRIPTION

As shown in Figure 7, the optical payload of the CS is composed of the following units:

• The main telescope including mirrors M1…M3 and a folding mirror/field stop unit M4 separating the science field of view from the surrounding fov

• A Wide Field Camera (WFC) serving as high-accuracy star tracker for spacecraft stabilization purposes

• A transmit telescope for relaying the science beam towards the hub The complete optical payload is operated at 40 K in order to reduce the thermal noise. The WFC must also be cooled to this temperature, since the WFC optics is seen by the science detector though the central hole in the field mirror. The various units are described in the detail in Chapter 4.



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Table 1 Collector s/c payload complement: main characteristics.

Payload Acro

nym Functionality Mass

[kg](*) Physical size [cm^3]

Ops Power[W]

Remarks

Main Telescope

TEL Receives and compresses the beam from the science target and surrounding FOV

307 (**)

340x340x380 (TBC)

TBC Power consumption determined by active tilt/focus control needs

Wide Field Camera

WFC high-accuracy star tracker for CS stabilization

6 TBD 5

Send transfer optics

STO Sends the science beam to the hub s/c

3 200x300x400 (TBC)

0

(*) mass figures do NOT include margins (assumed to be 20%, see Table 50) (**) Telescope mass excludes baseplate, which is considered to be part of the s/c structure

DARWIN Payload Definition


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3.2. Beam combiner spacecraft

3.2.1 FUNCTIONALITY

The beam combiner optical payload is located in the hub, and consists of a large number of optical modules that are responsible for the conditioning of the interferometer beams, their combination and detection (see Figure 10). In this section, the overall rationale, organization and design of the beam combiner payload are discussed. Its various components will be discussed in the detail in Chapter 5.

3.2.2 DESCRIPTION

The optical scheme of the TTN beam combiner is shown in Figure 9. The design is characterized by two levels of beam splitter/combiners. In principle, beam combination and modulation could also be achieved with a single beam-splitter level. The modulation efficiency in this case would be limited to 50%. The introduction of a second beam combination level allows to obtain a modulation efficiency of 93%. The first BS level splits each received beam into two components with nominally equal intensity, producing 6 output beams. Moderate deviations from the nominal 50:50 beam splitting ratio do not degrade nulling performance, as long as the beam splitters are identical. The 3 reflected beams are then phase shifted by 90° and pair-wise combined with the complementary 3 beams. The left outputs of the cross combiners are combined in a Single Mode Waveguide (SMW) after having been phase shifted by -120°, 0° and +120° respectively. The right outputs of the cross combiner are processed in the same way, but with the phase shifts +120°, 0° and -120° respectively. Moderate deviations from the nominal 50:50 beam splitting ratio are acceptable for the cross combiners, as long as the three optical components are identical. The optical design of the BC must ensure proper matching of the following optical properties of the beams at combination:

- pathlength (sensor: fringe sensor FS, actuators: delay lines ODL) - beam intensity (sensor: TBD, actuator: fine steering mirror FSM) - polarization. All interferometer beams must see identical numbers and types of

reflections, occurring in the same sequence. As discussed in more detail in Section 3.2.18, this requirement poses important constraints on the BCA layout.

The overall organization of the BCA is shown in Figure 10.. The BCA assembly receives the beams collected by the receive optics. Each beam is processed by a number of modules, whose function is briefly discussed below.



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Table 2 Beam combiner s/c payload: main characteristics.

Payload Acro Functionalitynym

Mass Physical [kg](*) size [cm^3]

Ops Power[W]

Remarks

Receive optics

RTO Receives and compresses the science beam from CS

18 Six times 250x350x400 (TBC)

N.A. Two separate sets of 3 RTO for linear & equilat. TTN

Select optics

SEL Connects optically the light from one of the two RTO sets to the BC

TBD TBD TBD

Fringe Tracker

FT Measures path differences, tilt and higher aberrations between interferometer beams

19 300x300x900 (optics) 300x300x400 (FPA/electr)

20 (TBC)

Fine steering mirror

FSM Fine/fast adjustment of beam pointing wrt the optical axis of the SMW

1 (TBC) 50x50x50 (actuator, TBC) 300x300x100 (electr., TBC)

2.5

Optical Delay Line

ODL

Provide fine/fast adjustment of the optical path of the beams

10.5 100x100x300 2.5

Differential Delay Line

DDL

Compensates shorter pathlength of central beam in Linear TTN

50 (TBC) TBD 2.5 Only needed for the linear array

Dichroic mirror

SUB Splits the spectral band into subbands 1 40x40x40 (TBC) N.A. Assuming two sub-bands SW = 6-11 µm and LW = 11-20 µm (TBC)

Achromatic Phase Shifter

APS Insert ±120 degrees achromatic phase shift 1.5 190x480x80 TBD

Fiber injection device

FID Reimages the beams onto the fiber tip. May include a beam shaper to increase fiber coupling efficiency

TBD TBD 0 (TBC)

Modal filter

SMW Performs modal filtering of the combined beams

Few grams 300x7x7 0

Spectrograph SP Performs spectral dispersion of the beams TBD 100x100x50 (each SW/LW) 0 Detectors DET Linear array, detecting the planet radiation TBD TBD < 10 mW Detector cooler

Provides vibration-free cooling of the detector to 6-8 K

20 TBD 200 Sorption cooler

(*) mass figures do NOT include margins (assumed to be 20%, see Table 51) These components are referred collectively to as the Beam Combiner Assembly (BCA). Beam combination physically occurs in the SMW.



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+90º

-120º

+90º+90º

+120º -120º+120º

A CB

Fine Steering MirrorOptical Delay Lines

Beam Splitters

Cross Combiners

Single Mode Waveguides

SMW Coupling Units

+90º

-120º

+90º+90º

+120º -120º+120º

A CB

Fine Steering MirrorOptical Delay Lines

Beam Splitters

Cross Combiners

Single Mode Waveguides

SMW Coupling Units

Figure 9: Optical beam combination scheme. Note that the two outputs are complementary, i.e. the signal registered on one detector corresponds to sources in the FoV which do not create a signal on the other detector. This allows chopping between the two channels, which enables background noise subtraction.



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Spectro-graph,

Detector&

Cooler

APS

SMWs

APS

4 5 6 7 8 9

APS

Beamshaper

& SMWcouple

Spectrograph,

Detector&

Cooler

from previous arm

to next arm

APS

SMWs

APS

4 5 6 7 8 9

APS

Beamshaper

& SMWcouple

to next arm

FS

Modul-

ator

1 2 3

SW sub-band

ODL

FSM PS

LW sub-bandfrom previous arm

Figure 10 The modules of the beam combiner assembly. 1) Optical Delay Lines, 2) Fringe Sensor and Photometric Sensor, 3) Modulator, 4) Beam splitter and 90º Achromatic Phase Shifter, 5) Cross combiner 6) 120º and -120º Achromatic Phase Shifters, 7) Beam shaper and SMW coupling unit, 8) Wavefront filters (SMWs) and 9) The detector assembly and the associated cooler



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3.2.3 FINE STEERING MIRROR

The Fine Steering Mirror (FSM) stabilizes the pointing of the interferometer beams within the BCS, in order to maximize the coupling efficiency of the science beams into the SMW. Moreover, the FSM is also used to control the intensity of the beam coupled into the SMW, by introducing small offsets with respect to the nominal pointing reference defined by the optical axis of the SMW.

3.2.4 OPTICAL DELAY LINE

The Optical Delay Line (ODL) is the actuator of the Fringe Tracking control loop and provides fine/fast pathlength corrections over a 2 cm range, with 1 nm accuracy at 10 Hz (TBD). The operating bandwidth is dictated by the spectrum of the OPD disturbance, including FF motion and microvibrations. Three ODL are needed in order to maintain complete symmetry of the optical trains. However, only two of them need to be actuated. The preferred location of the ODLs is in the beam combiner satellite, since the Fringe Tracker is located here. In order to minimize the torques induced by the moving parts, especially during the fringe acquisition phase, the delay lines shall be located as close as possible to the centre of mass of the satellite and their design shall minimize exported microvibrations.

3.2.5 DIFFERENTIAL DELAY LINE

A long-stroke differential delay line (DDL) is to be implemented in the BCS to compensate differential paths between the collector s/c and the hub occurring for the linear TTN. As shown in Figure 11 the collector s/c of a linear TTN are not equidistant from the hub. The s/c B is closer to the hub by a length δl that is a function of ISDAB and varies between about δl ~ 2.4 m (for the most compact array) and δl ~ 12.8 m (the most diluted array) for the geometry shown in the figure. In the scenario described above, a stroke of 10.4 m is required.



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s/c A

s/c C

Sphere centered at hub

δl

Hub

Compact array

Extended array

R ISDAB

s/c B

Figure 11 Possible array resizing strategy for the linear TTN. The distance ISDA,B between adjacent collector s/c will range between 15 m (most compact array) and 80 m (most diluted array). Assuming a fixed pointing direction of the hub receive telescopes, the distance of the collectors to the hub will ranges from R = 47 m at ISDA,B = 15 m to R = 250 m for ISDA,B = 80 m. However, in order to have a linear array, telescope B must be closer to the hub than the other two s/c by a length δ l~0.5 ISDA,B 2/R. Note that δ l~ 2.4 m for the most compact array, and δ l~ 12.8 m for the most diluted configuration.

3.2.6 FRINGE TRACKER

The Fringe Tracker (FT, or equivalently Fringe Sensor, FS) measures the relative phase of the three science beams to sub-nm accuracy at a 10 Hz rate. The FS rate is TBC, ODL development activities indicate that the FS rate might be in the order of 100Hz. This information is used to drive the ODLs. The Fringe Sensor also operates as a fine tilt sensor, measuring relative pointing of the beams at recombination with respect to an internal reference. This information drives the FSM mirror implemented before the ODL. The FT also measures higher-order aberrations up to and including spherical aberration, although at a much slower rate (TBD) than OPD and tilt.. The FT should be as close as possible to the beam combination optics in order to minimize differential effects between the FT readings, and the actual phases of the science beams at combination.

3.2.7 PHOTOMETRIC SENSOR

The photometric sensor measures the relative intensities of the science beams to 10-3 relative accuracy at a 10 Hz rate (TBC). This information is used to match the intensities of the beam at recombination using the FSM. The photometric sensor should be as close as possible to the beam combination optics in order to minimize differential effects between its readings, and the actual intensities of the science beams at combination.



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3.2.8 MODULATOR

The modulator induces a periodic change in the optical transfer function of the interferometer such that the detector sees two different transmission maps alternating in time. Subtraction of the readings in the two modulation states removes uniform background contributions (for example, thermal emission) from the signal, while it modulates the signal from a planet transiting close to the transmission peak. The highest modulation efficiency is achieved for the positions in the sky that lie on a peak in one transmission map and in a trough in the other transmission map.

3.2.9 DICHROIC DEVICE

After the modulator, each interferometer beam is split into a number of spectral sub-bandwidths. This is necessary since the Darwin APS device, which consists of 2 or 3 dispersive wedges in each interferometer arm, can be designed to operate properly limitedly to a spectral bandwidth of about one octave. In the following, we shall provisionally assume that the Darwin science bandwidth is split into two sub-bandwidths, namely short wavelengths (SW band, covering 6-11 micron) and long wavelengths (LW, 11-20 micron) (TBD). The spectral separation is by a dichroic mirror assembly in each arm. The three dichroic mirrors shall have closely matched spectral response in order to avoid chromatic mismatches between the beams at combination. Requirements and design are TBD, pending a trade-off about the number and extent of the sub-bandwidths needed for combination. Note that also single mode waveguides (SMW) have a limited spectral bandwidth of about one octave over which only the fundamental mode is propagated, which is determined by the refractive indices of core and cladding and the geometry of the waveguide. Nominally, each sub-bandwidth is fed to one single waveguide. In the case where this result into unacceptable insertion losses, the recombined beams will have to be split into more spectral subbands, each feeding a separate fiber. A trade-off concerning the optimal number of spectral sub-bands is required.

3.2.10 BEAM SPLITTERS

The beam splitters and cross combiners should be as close as possible to the ideal component, which reflects 50% of the light and transmits the remaining 50%, over the complete spectral bandwidth, at 45º incidence angle. However, moderate deviations from this nominal splitting ratio would not result in a loss of function but just result in performance degradation, provided that identical components for beam splitting and identical components for cross combination are used. A limited analysis indicates that the performance degradation would restrict itself to a loss in modulation efficiency.

3.2.11 ACHROMATIC PHASE SHIFTERS

The achromatic phase shifts required for TTN are +90°, +120° and –120°, see Figure 9. The baseline device for generating such phase shifts consists of stacks of two or three dispersive wedges, inserted in each interferometer arm, such that the difference in OPL increases linearly



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with wavelength, i.e. achromatically. Note that this design allows in-flight compensation of small chromatic errors introduced elsewhere in the optical train, which cannot be done with reflective APS devices.

3.2.12 BEAM SHAPER

The theoretical maximum optical power that can be coupled into a single mode waveguide is approximately 80% of the input power, assuming a flat intensity distribution over the pupil plane. This coupling loss is due to the mismatch between the Airy pattern, as created in the focal plane when the beam is imaged at the SMW tip, and the Gaussian shape of the fundamental mode in the SMW. In the three-beams case, the coupling efficiency into the SMW could be improved by shaping the input pupils to form a single circular one, with amplitude distribution as close as possible to the Gaussian pattern of the fundamental mode of the SMW. Depending on the selected design, the beam shaper device could be integrated in the SMW coupling unit.

3.2.13 SMW COUPLING UNIT

The SMW coupling unit performs beam injection into the SMW, i.e. arranging the beam(s) such that the optical power will propagate inside the core of the SMW. This is done by focusing the beam on the center of the SMW core, matching the numerical aperture of the SMW. In its basic form, the SMW coupling unit is a focusing telescope with the same f/number as the fiber.

3.2.14 SINGLE MODE WAVEGUIDE

In the case of the TTN architecture the single mode waveguide has a dual functionality;

a) beam combination, i.e. arrange the beam such that they propagate along the same optical axis and interfere, and

b) modal filtering, removing residual wavefront errors, leaving only the fundamental mode propagating in the waveguide and thus contributing to the output signal.

The spectral bandwidth over which SMWs can propagate the fundamental mode with high efficiency is approximately one octave. Since the Darwin spectral range is more than 1 octave (but less than two) at least two SMW are needed, covering for example the spectral bands between 6 – 11 and 11 – 20 micron respectively.

3.2.15 SPECTROGRAPH

Each of the SMW units associated to the Darwin sub-bands is followed by a spectrograph, which disperses the planet signal onto the detector array associated to that sub-band. A modest spectral resolution of about R =36 through the full science bandwidth is sufficient, except for the 9.2-10.1 micron range (TBC) containing the ozone line, where a higher resolution of about R = 120 is needed. The number and characteristics of the spectrograph units will depend on how the Darwin science bandwidth is split into sub-bands.



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3.2.16 DETECTOR

The baseline detector for both the the SW and LW spectral sub-bandwidths is a Si:As linear (1D) detector array. The same detector serves both the detection and the spectroscopy phases, since binning is possible without significant sensitivity degradation. The detailed geometrical design of the spectrograph will define the orientation of these detector channels. A total of four independent detectors is needed to cover the two science sub-bands (SW and LW), and two independent outputs of the TTN beam combines, see Figure 9. Different pixel sizes may be required for the two sub-bandwidths.

3.2.17 FOCAL PLANE ASSEMBLY AND COOLER

The focal plane assembly will incorporate line array detectors in a detector housing. The required opening for the incident light is not considered to be critical, provided that the closest optical component is sufficiently far away such that thermal leakage is minimized. Focussing and uniformity of the dispersed wavelength scale require more attention. The focal plane assembly requires good thermal insulation in order to reduce the necessary cooling power. Dark current generation limits the operation temperature effectively to <15 – 17 °K. Note that these figures are device specific. This temperature range is at the limit of a single-stage cooler, and therefore more detailed investigation on both the design of the focal plane, as well as on developments and measurements of candidate device technologies should be performed. The read-out circuit could be implemented in so-called Cryo-CMOS, which meets the noise performance requirements. The readout circuit is flip-chip-bonded onto the array. It may be necessary to custom manufacture a line readout assembly. The thermal load will be dominated by the ROIC power consumption and by parasitic loads mainly from connector leads, providing the necessary bias and clocking from the warm FEE to the cold FEE. The thermal load from detector bias and leakage, and from the light intensity through the aperture is marginal, even in the case when the constructive output of the interferometer is sent to the focal plane. Early estimates of the power consumption indicate that the ROIC could operate at <10 mW (for the whole detector assembly). Another possibility to reduce the required thermal lift-off that should be looked into is the staging of the operating temperature of the focal plane assembly. At least two separate chips will be necessary to fill the required wavelength band. Then, it may be possible to provide different operating temperatures to different segments, where the detectors covering the shorter wavelength band would be able to operate at a slightly elevated temperature. This would allow operating the detectors warmer at areas where the sensitivity to created dark current is less.

3.2.18 BCA IMPLEMENTATION

The beam combiner assembly (BCA) is implemented on the beam combiner spacecraft. It has an overall shape of an extended triangular tube, with the input interfaces at one end and the detectors at the opposite end. This arrangement provides identical numbers and types of



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reflections, occurring in the same sequence, for each of the three input beams. This minimizes polarization differentials that would reduce the achievable null depth. The beam combiner assembly has three arms, rotated by 120 degrees relative to each other around the longitudinal axis of the BCA. A number of stages can be defined for the BCA. The first stage accommodates the three optical delay lines (ODLs). In the second stage a dichroic mirror splits the metrology spectrum from the science band. In order to minimize the path length not measured by the FS, the FS should be located in the optical train as close as possible to the detector. However, it must be located before the first beam recombination. The third stage accommodates the modulator, which enables modulation between two different transmission maps.

APS

APS

APS

Mirror

Beam Splitter

Beam input

Beam output Stage 4 Stage 5

Figure 12 Stage 4 and 5 of the BCA, as seen along the propagation axis of the light.

Th fourth stage holds the beam splitter and the 90º achromatic phase shifter. The two beams in each arm of the interferometer are passed to the fifth level, which holds the cross combiners. These two stages are illustrated in Figure 12, showing how the beams from the different arms are routed. Note that polarization is maintained up to beam recombination, by an odd number of reflections in both arms. The reflections are in parallel planes. The sixth stage holds the +120º and the -120º phase shifters, passing the beams to the seventh stage where the beams are reshaped and injected into the modal filter, stage 8. The Beam shaper & SMW coupling (BSC) evens the intensity distribution over the pupil to better match the first mode of the SMW, thereby reducing the coupling losses. The resulting pupil is then focused on the SMW input face. The number of SMW is twice the number of sub-bandwidths. The ninth and final stage is the spectrograph, detector assembly and its cooler. An alternative design has been recently proposed as part of the Emma assessment study [RD 12], and is shown in Figure 13. Each of the three beams first propagates through a delay line so that the optical paths can be equalized. Each beam is then split into two beams and propagates through an APS. Note that the beam combiner scheme of Figure 9 includes a first stage of APS introducing a 90 degrees phase shift. In the considered implementation, this delay is provided by



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the BS reflection so that no dedicated APS device is required. One of the two beams propagates through an intensity matcher (IM) so to make sure that the intensities are of the correct level. The six created beams are recombined at three beam crosscombiners so to create six beams, each beam containing the signal from two different telescopes. Finally, the 120o phase shifts are applied. These phase shifters are represented by the white and black boxes in the drawing. The three beams that propagate through the “white” (“black”) APS's will be focused and combined at the same fiber. To this purpose, these beams will be directed towards the SMW coupling unit (not depicted in the figure). Notice that the optical axis is orientated perpendicular to the plane of this optical scheme.

Figure 13 Alternative BCA arrangement from [RD 12]. IM = intensity matcher. APS = Achromatic Phase Shifter.



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4. TELESCOPE SPACECRAFT PAYLOAD

4.1. Main telescope

4.1.1 FUNCTIONALITY

The main telescope collects the incoming radiation from the science target. It has a diameter of 3.15 m and a diffraction-limited field of view of about 2.7 x 2.7 arcmin. An integrated field separator extracts the inner 1” field of view, containing the science target. This beam is collimated to a beam of 20 mm diameter and sent to the hub via the beam transfer optics. The outer part of the field of view is directed towards the Wide Field Camera for s/c attitude stabilization. The main technical requirements for the science telescopes are listed in Table 3. Table 3 Technical requirements for the science telescopes

TECHNICAL REQUIREMENTS

Parameter Requirement Goal

Aperture diameter 3.15 m Largest diameter compatible with launch in a Soyuz fregat

Optical Bandwidth 0.3-20 micron 0.3-1 µm used for star tracking (WFC) 1-4 micron used for fringe/tilt tracking 6-20 micron used for science

Diffraction-limited fov 2.7 × 2.7 arcmin Central 1” (TBC) used for science outer field used for s/c stabilization

4.1.2 DESCRIPTION

The current baseline telescope design was originally proposed during the Alcatel system study [RD 1] and is based on a Korsch three-mirror afocal design, with the addition of a fold mirror M4. The optical design of the telescope is shown in Table 4. The design uses a relatively fast primary mirror, in order to stack two telescope spacecraft within one Soyuz launcher, see Figure 15. Close to the primary focus of the telescope there is a combined field stop / pick-off mirror M4. This is a flat mirror with a central hole that transmits the science field of view (about 1” around at the science target, TBC) to be collimated by M3 and reach the transfer optics for relay to the hub. The outer field is directed towards an internal high performance star tracker, the Wide Field Camera (WFC), see Section 4.3. The same diameter of 20 mm of the science beam after M3 has been retained as in the original Alcatel design, for compatibility with the following optics. Accordingly, the beam compression factor of the Darwin telescope is M = 3150/20 = 157.5.



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Table 4 Optical design of the DARWIN telescopes

Curvature

radius [mm] Thickness [mm]

Aperture radius [mm]

Type

M1 -6300.0 -2505.3 1575.0 parabola M2 -1620.8934 3572.373 336.0 hyperbola,

k = -2.29275(*) M3 -400.0 -200.0 25.0 Parabola M4 infinite 2100 50 Flat mirror * This value controls the cancellation of the spherical aberration in the system and should be checked since the radius of M4 has changed The optical quality of the nominal Darwin optical design is excellent within the 1” instrumental field of view used for science. The wavefront error is only 0.7 nm rms. This figure does not increase significantly for any combination of rigid translations and tilts of the satellites, within the limits imposed by the GNC (1 cm translation, ~10” attitude). Clearly, the final optics error budget will be dominated by manufacturing and mirror alignment errors.

4.1.2.1 Optics The concept and the manufacturing of the primary mirror M1 present several complexities. This mirror has large dimensions (3150 mm diameter) and is optically fast. Extensive lightweighting is needed for mass reduction. Still, the mirror must have a good optical quality and dimensional stability including during cool down to 40 K. The baseline material for M1 is SiC, which allows for mirror surface densities as low as of 20 kg/m2 for a 3 m class mirror. Therefore, the estimated mass of M1 is 156 kg. We assume that all other telescope mirrors are also made of SiC, which ensures a high degree of thermo-elastic uniformity for the entire telescope. The mirror surface quality requirements are relatively moderate, since wavefront errors are filtered out by the Single Mode Waveguide. A preliminary WFE budget based on [RD 2] would require rms surface roughness of better than 40 nm for M1, 25 nm for M2, 20 nm rms for M3 and 15 nm rms for M4. The surface roughness of the mirrors should be less than 0.5 nm rms per mirror, in order to minimize the scattered starlight flux. For all mirrors, the baseline coating material is bare gold because of its high transmission throughout the required spectral range and benign polarization properties.

4.1.2.2 Structure The mechanical architecture of the telescope is sketched in Figure 14, and again is inherited from the Alcatel Darwin system study [RD 1].



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M2Spider

Cylindre

M1

Baseplate

WFC

M3

Beam trackerTransmitter

Figure 14 Mechanical architecture of the Free-flyer telescopes

This design has been selected with the objective of ensuring high stability between the optical components comprising the payload. The mechanical concept adopts the use of highly stable structural materials, taking into account the manufacturing feasibility of these elements. The mechanical paths between optical elements, between which a high degree of stability is required, should be minimised as well as the number of mechanical links. In order to ensure a high degree of alignment stability, a thick, rigid, and stable base-plate is used to provide support for all of the optical elements. The primary mirror M1 is attached to the baseplate by means of a specific Mirror Fixation Device (MFD) which provides an isostatic mount, such that eventual distortions of the base-plate induce only very low wavefront perturbations to the optical surface of the mirror. All other optical components of the Collector s/c are mounted to the bottom of the baseplate. The main telescope structure is basically comprised of a structural cylinder and a secondary mirror spider to support the secondary mirror. Deployable telescope secondaries have not been envisaged here, firstly in order to limit the number of complex technologies required to ensure the mission's success, and secondly because the current design already comes close to the upper limit of the launch mass capability.



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Table 5 Telescope mass budget

Component Diameter (mm)

Mass (kg)

Remarks

Primary mirror M1 3150 155.9 SiC technology assumed (20 kg/m2) Secondary mirror M2 540 4.6 “ M2 support 5.4 Fold mirror M3 and support

85 1

Telescope tube 141 Total mass 306.9 no mass margins included

The cylinder is composed of two concentric layers and naturally fulfils the required optical and thermal baffling functions. It is directly interfaced with the base-plate. The internal one must provide excellent performance in terms of mechanical stability, since this will have a direct influence on the alignment accuracy and stability of the (M1, M2) telescope mirrors. Preliminary analysis shows that a 3.2 mm thick, cylindrical skin of C/C could satisfy the mechanical stability requirements of the telescope optics. The external tube must have a low emissivity with respect to neighbouring CSs, for which purpose an aluminium coating is required on its external surface, and a black paint coating on its internal surface. The external tube must also support locking devices for the sunshield arms during launch, and must therefore exhibit good resistance and stiffness under launch loads. For these reasons, a classical aluminium honeycomb sandwich using aluminium skins is foreseen. It should be noted that this architecture needs to be modified due to specific constraints of the TTN design. The Soyuz ST fairing provides a user volume with a basic diameter of 3.8 m and a height of 9.5 m. Figure 15 shows that a structural cylinder supporting M2 is not compatible with the need to stack two telescope s/c within the Soyuz fair. For the upper telescope, the cylinder must be shortened, and M2 must be supported by a spider structure, presumably directly attached to M1. The total mass of this modified telescope assembly including mirrors, spiders and autofocus mechanism is estimated to be 307 kg including mirrors, spiders and structural cylinders. The mass of the supporting baseplate is about 100 kg.



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Umbilical Umbilical

Umbilical

Propulsion Module

Star Tracker

Figure 15 Accommodation of two telescope spacecraft under the Soyuz-ST fairing, for a telescope mirror diameter of 3.15 m.

4.1.2.3 Alignment mechanisms The WFE requirements in Table 3 lead to severe stability requirements on the M1-M2 mirror pair, which can be summarised as follows: Defocus M1/ M2 ± ~ 1 µm bias, 30 nm stability during observation Tilt of M1 with respect to M2 ± ~ 1 arcsecond Decentering between M1/ M2 ± ~ 1 µm These requirements apply to the Alcatel design (1.5 m apertures) and should be revisited for the TTN case. However, the numbers are not expected to change substantially. The telescope is required to maintain this stability performances over periods of at least the integration time required for detection, but is not expected to maintain this level of stability with respect to launch misalignments. Due to this high level of mechanical stability required, the tolerable thermal fluctuations (asymmetric gradients or differences between CSs) are expected to be of the order of ~ 0.1 K. The 1 µm defocus requirement defines the starting condition for the operation of an active focus control, which will reduce further defocus to the ultimate accuracy requirement of ~ 30 nm needed during observation. The need for active control of the position and tilt of M2 is TBD.

4.1.3 PERFORMANCE

No measured/expected performance can be listed, since no specific activity concerning the telescope development is running as yet. Note that the optical properties of the three telescopes



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(in particular, polarization and spectral response) must be matched to a high degree in order to preserve deep null capability. Table 6 Comparison of required and measured (or expected) performance for the main telescope

Performance

Required

Remarks

Rms WFE per mirror

M1: 50 nm M2: 23 nm M3: 20 nm M4: 15 nm

Manufacturing errors only

Mirror roughness < 0.5 nm rms Ensures low scattered starlight

Spectral response uniformity

< 5×10-4

Between any two telescope assemblies

M1-M2 misalignment

Defocus: ± ~ 1 µm Rel. tilt: ± ~ 1” Rel. decenter: ±~1 µm

Static misalignments wrt nominal positions

M1-M2 alignment stability during observ.

Defocus: 30 nm Tilt: ~ 1” Decenter: TBD

Requires active focus ctrl Need for active ctrl TBD Need for active ctrl TBD



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4.1.4 INTERFACE AND PHYSICAL RESOURCE REQUIREMENTS

Table 7 Interface requirements for the main telescope

Interface requirement

Value

Remarks

Volume (mm×mm×mm) 3400x3400x 3850(TBC)

Mass (kg) 307 Whole telescope, excluding baseplate Accommodation requirements Telescope is fixed to baseplate with three isostatic mounts.

All CS optical payloads are fixed to the opposite side of the baseplate

Power

Average consumption (W) TBD M2 defocus actuator Tilt/decenter actuators TBD

Peak consumption (W) TBD

Thermal interfaces

Operational temperature (K) 40 K Complete Collector s/c payload Operational temp. range (K) 0.1 K Allowable gradient within a telescope

and/or between two telescopes

Optical interfaces

Aperture diameter 3.15 m Largest diameter compatible with launch in a Soyuz fregat

Exit beam diameter 20 mm (TBD) Angular magnification of 157.5

Optical Bandwidth 0.3-20 µm 0.3-1 µm used for star tracking (WFC) 1-4 micron used for fringe/tilt tracking 6-20 micron used for science

Alignment telescope axis /WFC axis

TBD

Alignment telescope axis /send optics axis

TBD

Alignment telescope axis/metrology beams from hub

TBD

Other

Exported vibration level (µN) TBD



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4.1.5 OPEN POINTS AND CRITICAL ISSUES

4.1.5.1 Telescope mechanical design The mechanical design of the telescope is inherited from the Alcatel study. However, this design is not compliant with a number of constraints of the TTN implementation. In particular, the use of a structural cylinder supporting the secondary mirror M2 does not allow stacking two telescopes s/c within the Soyuz fairing. A spider structure is required to this purpose, and the cylinder must be shortened or made deployable. This will also modify the thermal characteristics of the telescope assembly.

4.1.5.2 Alignment accuracy and stability The telescope has severe stability requirements on the M1 and M2 mirrors, especially in terms of decentering. The system must be capable of preserving correct alignment and WFE performance at 40 K, while most of the telescope alignment operations will be performed at room temperature.

4.1.5.3 Primary mirror manufacturing and polishing Several 3-m class mirrors are being built using SiC or related technology, including the Herschel and SPICA primaries. However, the Darwin M1 wavefront error requirement is ten times more stringent that for the other mission. Moreover, the cryogenic performance at 40 K has never been tested before and thus must be confirmed. The roughness requirement of 0.5 nm rms at the limit of what is achievable at present, and for much smaller mirrors. However, it is expected that polishing technology will evolve sufficiently to ensure that this requirement can be met in useful time for the Darwin mission.

4.2. Send transfer optics

4.2.1 FUNCTIONALITY

At each telescope s/c, the send transfer optics expands the collimated science beam received from the science telescope to the optimum size for transfer, and launches it towards the hub. At the hub, the receive transfer optics receives the beam, and compresses it back to the size required for beam combination. The receive transfer optics is described in Section 5.1. The transfer optics also re-images the telescope entrance pupil at the hub location. The actual position of the reimaged pupil depends on the distance between the spacecraft.

4.2.2 DESCRIPTION

The send transfer optics consists of an afocal telescope that expands the output beam from the science telescope (whose diameter is about 20 mm) into a collimated beam with a diameter of 150 mm. The size of the transferred beam is the result of a trade-off between the contrasting needs of low beam divergence for the transferred beam (or high coupling efficiency at the receiver, which favours large optics) and low mass/complexity/cost of the relay system (which calls for small apertures).



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Figure 16 Photometric beam transfer efficiency for assumed diameters of the transmit/receive optics of 300 mm (solid line), 250 mm (dashed line), 200 mm (dotted line) and 150 mm (dash-dot line).

As shown in Figure 16, 200 mm diameter apertures result in a photometric transfer efficiency of 95% at all wavelengths. In reality, this size is only needed at the receive optics, where the effective beam footprint is increased by diffraction and s/c jitter. For the send optics, a diameter of 150 mm has been shown to be sufficient. Table 8 Technical requirements for the send transfer optics



Input beam diameter 20 mm Compression ratio 1:7.5

Output beam diameter 150 mm

Transmit fov TBD Typically 1-10 Airy disks at 20 µm Optical Bandwidth 1-20 micron 1-4 micron used for fringe/tilt tracking

6-20 micron used for science The send transfer optics is implemented as a two-mirror Gregorian telescope. The off-axis design has the advantage of having no obstructions that may generate straylight, see Figure 17.



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Figure 17 Transmitter telescope with 2 off-axis mirrors and a dedicated assembly structure

The Gregorian design provides an intermediate focus, where a field stop could be placed in order to reject straylight generated within the telescope (from the spiders of the main telescope, etc.). The need for such field stop for the send optics is TBD. The send transfer optics is supported by a dedicated stable structure, in order to ensure that its manufacturing, alignment and ground tests can be performed independently from other optical components of the CS payload. Once integrated and tested, the complete transmitter assembly is then mounted and aligned on the lower side of the telescope base-plate.

4.2.3 PERFORMANCE

Table 9 Comparison of required and measured (or expected) performance for the send transfer optics

Performance

Required

Measured/ expected

Remarks

Overall WFE 28 nm TBD Including manufacturing and alignment Surface roughness 0.5 nm TBD



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Table 10 Interface requirements for the send transfer optics

Interface requirement SEND TRANSFER OPTICS

Value

Remarks

Volume (mm×mm×mm) 200x300x400 (TBD) Baffling excluded Mass (kg) 6 Baffling excluded Accommodation requirements Transfer optics may be mounted high above the plane of

the sunshield to minimize view factor to them

Power

Average consumption (W) N.A. Peak consumption (W) N.A.

Thermal interfaces

Operational temperature (K) 40 Same as for main telescope Operational temp. range (K) 0.1 Same as for main telescope

Optical interfaces

Input beam diameter (mm) 20 TBD Output beam diameter (mm) 150 TBD

Other

Exported vibration level (µN) N.A.


The send transfer optics does not present critical issues, although operation at 40 K implies a careful choice of optical and mechanical components. The opportunity of introducing a field stop at the intermediate focus of the send optics is TBD and should be addressed in the course of a end-to-end straylight analysis.

4.3. Wide Field Camera

4.3.1 FUNCTIONALITY

The Wide Field Camera (WFC) acts as a fine pointing sensor, allowing the telescope flyer to achieve very accurate and stable attitude control, using the flyer's fine AOCS system as a corrective actuator. The WFE operates as a conventional star tracker, with the difference that the WFC detector is fed by the light collected by the 3 m class science telescope, rather than by a



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dedicated (and smaller) collector. This allows measurements of the s/c attitude to sub-arcsecond level with excellent sensitivity.

4.3.2 DESCRIPTION

Each telescope spacecraft has a high performance star tracker (also referred to as the Wide-Field Camera, WFC). The WFC is used for target acquisition, and as primary pointing sensor in all operational modes except NOM. The WFC instrumental requirements for TTN are illustrated in Table 11.

Figure 18 Interface of the WFC interfaces to the intermediate focus of the science telescope. The field selector mirror M4 diverts the outer portion of the telescope fov towards the WFC, which consists of a reimaging optics and a 2D detector array.

In the TTN design, the beam coupling efficiency into the SMW is optimized by fine tip/tilt mirrors installed in the hub. Spacecraft attitude control requirements are driven by the need to provide efficient beam transfer from the CS to the hub, resulting into a WFC measurement accuracy of 0.1” (1 σ) at a sampling rate of 10Hz. This accuracy holds for pitch and yaw, while the attitude around the telescope LOS will be measured to a lower accuracy, typically 30 arcseconds @ 10Hz. As shown in Figure 18, the WFC interfaces to the intermediate focus of the science telescope. The field selector mirror M4 diverts the outer portion of the telescope fov towards the WFC, which consists of a reimaging optics and a 2D detector array. The wavelength range of operation of the WFC is provisionally specified as 0.3-1 micron, although, depend on the selected detector type, the final range could be extended in order to increase sensitivity or reduce fov requirements. The re-imaging optics projects a 2.7 x 2.7 arcmin2 field on the sky at the WFC sensor. The field is dimensioned to ensure a 95% probability to find at least one sufficiently bright star for nominal WFC performance anywhere in the sky. A trade-off should assess whether the field size could be reduced by a factor of two, while maintaining a > 95% sky coverage.



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Table 11 Technical requirements for the Wide Field Camera for TTN. Angles are referred to the sky.

PARAMETER REQUIREMENT

Operating bandwidth 0.3-1 micron

WFC fov 2.7 × 2.7 arcmin

WFC sensitivity 0.1 arcsec @ 10 Hz (pitch/jaw) 30” (roll)

Star limiting magnitude mV = TBD (> 19)

Detector dimensions 2k × 2k pixels

Readout rate 1 Hz

This would allow for considerably more compact WFC optics, and would reduce the total number of detector pixels. The actual design of the WFC optics depends on the design of the telescope optics and on the detector pixel size, and has not been addressed yet. However, most likely it will be a refractive system. The detector is a CCD-like detector of 2048 x 2048 pixels. Note that the science sensor will see the WFC optics through the pinhole in fold mirror M3. Accordingly, the WFC unit must be cooled at 40 K. For standard Si CCDs, cooling below 77 K will result in the so-called "carrier freeze-out" process, i.e., the mobility of electrons in the silicon will drop. HgCdTe arrays can be used as an alternatice. 2kx2k HgCdTe arrays have been developed by Rockwell for the JWST shortwave channel (0.6-2.3 micron) which can be operated well below 40 K. WFC data processing equipment is TBD. In this context it should be noted that the star catalogue would need to be larger than for a conventional star tracker, since the FoV of the WFC is smaller, but the limiting magnitude is much fainter.

4.3.3 PERFORMANCE

Table 12 Comparison of required and expected performance for the Wide Field Camera

Performance

Required value

Goal Remarks

Operating bandwidth 0.3-1 micron WFC fov 2.7 × 2.7 arcmin Possibly can be reduced to

80 × 80 arcsec WFC sensitivity 0.2 arcsec @ 10 Hz

(pitch/jaw) 30” (roll)

Ensure efficient beam transfer from CS to hub



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Star limiting magnitude mV = TBD (> 19) Increased telescope size, relaxed

pointing requirement Detector dimensions 2k × 2k pixels Possibly can be reduced to

1k × 1k Readout rate 10 Hz


Table 13 WFC interface requirements


Value

Remarks

Volume (mm×mm×mm) 500x100x100 Preliminary estimate Mass (kg) 5 Accommodation requirements WFC is accommodated on

the back side of the telescope baseplate

Power

Average consumption (W) TBD Peak consumption (W) TBD

Thermal interfaces

Operational temperature (K)

40 The WFC detector sees the science detector and must be cooled to 40 K. Standard CCD cannot be used.

Operational temperature range (K)

2 (TBD)

Optical interfaces

Alignment WFC origin/ LOS science telescope TBD


The WFC does not present critical issues. A trade-off should assess whether the field size could be reduced, while maintaining a > 95% sky coverage. This would allow for considerably more compact WFC optics, and would reduce the total number of detector pixels. The WFC detector must be cooled to 40 K. The on-board star catalogue would be considerably larger than for a conventional star tracker, since the WFC limiting magnitude is mV = 19 or fainter (TBD).



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5. BEAM COMBINER SPACECRAFT PAYLOAD

5.1. Receive Transfer Optics

5.1.1 FUNCTIONALITY

The receive transfer optics receives the science beam as sent from the send transfer optics on the CS, and compresses it to the diameter of 20 mm required for beam combination.

5.1.2 DESCRIPTION

Table 14 Technical requirements for the receive transfer optics


Parameter Requirement Remarks

Input beam diameter 200 mm Receive optics is oversized wrt send optics to accommodate for diffraction and beam wander due to CS attitude jitter

Output beam diameter 20 mm Compression ratio 7.5:1

Receive fov TBD Receive fov is oversized to allow for s/c attitude jitter. However, it shall also minimize collected straylight

Optical Bandwidth 0.5-20 micron 0.5-4 micron used for fringe/tilt tracking 6-20 micron used for science

The receive optics is implemented as a two-mirror Gregorian telescope (see Figure 17), similar to the send optics (see Section 4.2). However, the primary mirror of the receive telescope is oversized to a clear aperture of 200 mm to accommodate the effective widening of the received beam footprint due to s/c attitude jitter, and diffraction. No steering parts are involved in the receive telescope. The receive transfer optics is supported by a dedicated stable structure, in order to ensure that its manufacturing, alignment and ground tests can be performed independently from other optical components of the payload. Once integrated and tested, the complete transmitter assembly is then mounted and aligned on the hub base-plate. With reference to Figure 20, the field of view of the receive optics should be dimensioned in order not to collect heat radiated from the edge of the CS sunshield, at the longest intersatellite required distance.



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From CS

Figure 19 Receive telescope with 2 off-axis mirrors and a dedicated assembly structure

hubHeight of Transfer optics

Transfer LOSCS

SunshieldSun light

Receiveoptics

Sendoptics

hubHeight of Transfer optics

Transfer LOSCS

SunshieldSun light

Receiveoptics

Sendoptics

Figure 20 Beam transfer and straylight from CS sunshade. TheCS sunshield must remain outside the field of view of the receive optics at the longest ISD

The sunshade view angle to the receive optics can be reduced with a combination of several countermeasures: - Placing a suitable field stop at the intermediate focus of the receive telescope - increasing the angle between the star light and the CS sunshade at the BC, by positioning the relay optics high above the plane of the sunshield



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- baffling the receive optics with long, deployable baffles extruding from the main s/c body. Since the baffle must remain inside shade cone of the s/c, the allowable baffle length decreases as the height of the baffle over the sunshield plane is increased. Therefore, a trade-off between these two parameters (baffle length and height above sunshield) must be performed.

- Angling sunshade downwards

Straylight mitigation is a still largely unaddressed issue that may have a strong impact on the design of the relay optics and its interface to the s/c. The straylight issue is a critical one and must be addressed in depth. There will be two separate sets of receive telescopes on the beam combiner spacecraft, corresponding to the linear and the triangular array configuration, see Figure 21. The same beam combiner assembly will be used for both aperture configurations. For this purpose, a beam select mechanism will allow to re-direct the beams from either set of receive telescopes to the combiner.

Beam combiner spacecraft

Telescope spacecraft C

Telescope spacecraft B

B

Telescope

A Cspacecraft A spacecraft C

Telescope

Beam combiner spacecraft

Telescope Telescopespacecraft B spacecraft A

A CB

Figure 21 Beam relay scheme for the equilateral and linear TTN



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5.1.3 PERFORMANCE

Table 15 Comparison of required and measured (or expected) performance for the receive transfer optics

Performance

Required

Measured/ expected

Remarks

Overall WFE (send+receive)

28 nm TBD Including manufacturing and alignment

Surface roughness 0.5 nm TBD


Table 16 Interface requirements for the receive transfer optics RECEIVE TRANSFER OPTICS

Value

Remarks

Volume (mm×mm×mm) 200x300x400 (TBD) Per unit, six units needed, baffling excluded

Mass (kg) 6 Per unit, six units needed, baffling excluded

Accommodation requirements Transfer optics may be mounted high above the plane of the sunshield to minimize view factor to them

Power


Thermal interfaces

Operational temperature (K) 40 Same as for main telescope Operational temp. range (K) 0.1 (TBC) Same as for main telescope

Optical interfaces

Input beam diameter (mm) 200 Nominal input beam diameter of 150mm, extra margin accommodates for diffraction and s/c attitude jitter

Output beam diameter (mm) 20 TBD

Other




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The receive transfer optics does not present critical issues, although operation at 40 K implies a careful choice of optical and mechanical components. Straylight mitigation is a still largely unaddressed issue that may have a strong impact on the design of the relay optics and its interface to the s/c. The straylight issue is likely to be a critical one and must be addressed in depth.

5.2. Fine steering mirror

5.2.1 FUNCTIONALITY

The Fine Steering Mirror (FSM) stabilizes the pointing of the interferometer beams within the BCS, in order to maximize the coupling efficiency of the science beams into the SMW. Moreover, the FSM is also used to control the intensity of the beam coupled into the SMW, by introducing small offsets with respect to the nominal pointing reference defined by the optical axis of the SMW. Table 17 Technical requirements for the Fine Steering Mirror



Input beam diameter 20 mm

Tip/tilt range TBD Receive fov is oversized to allow for s/c attitude jitter. However, it shall also minimize collected straylight

Tip/Tilt accuracy TBD Equivalent to TBD on the sky

Tip/Tilt accuracy TBD Equivalent to on TBD the sky

Optical Bandwidth 1-20 micron 1-4 micron used for fringe/tilt tracking 6-20 micron used for science

5.2.2 DESCRIPTION

While no baseline concept for the FSM has been selected yet, a suitable concept could envisage a steering mirror suspended on a flexure mount, and driven by voice-coil or piezo actuators. Differential capacitive sensors are included in the design, in order to achieve high positional accuracy and resolution. A specific requirement on the Darwin FSM is that it shall not induce significant amounts of OPD. This requires gymbal-like operation, where the position of the center of the mirror is kept constant during operation.



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5.2.3 PERFORMANCE

Table 18 Performance requirements for the Fine Steering Mirror

Performance

Required value

Remarks

FSM optical requirements Wavelength range 0.45 – 20 µm

Free optical diameter > 25 mm Induced OPD

(over the full actuation range) TBD Requires gymbal-like actuation

Wavefront distortion < λ/20 RMS (λ=633 nm)

Optical transmission losses TBD Not critical, single reflection

Coating manufacturing reproducibility • Relative spectral response • Chromatic phase differences • Relative polarization

o Rotation o Ellipticity

< 10-4

OPD contribution < 0.1 nm RMS < 0.1 ° < 0.1 °

It depends on corner cube manufacturing and angular mounting tolerance

Polarization variation (over the full actuation range)

• Rotation • Ellipticity

< 0.1 ° < 0.1 °

FSM functional requirements Tip/tilt stroke TBD

Tip/tilt resolution TBD

Tip/tilt stability TBD

Assuming low measurement noise and low microvibration environment (100nm RMS, 1Hz cut-off, -40dB/dec slope). Under typical ground microvibration conditions the FSM stability is expected to be <TBD RMS

absolute tilt accuracy TBD FSM dynamic response

requirements

Max. tip/tilt rate TBD Settling time (in closed loop) < 20 ms

Overshoot none



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Table 19 Interface requirements for the FSM unit


Required value

Remarks

Volume (mm×mm×mm) 50x50x50 (TBC) Mass (kg) 1 (TBC) Accommodation requirements

Power

FSM average power consumption TBD

average value of the power consumption dissipated in a dynamic FSM (including harness) operating at 40K

FSM peak power consumption

peak value of the power consumption dissipated in a dynamic FSM (including harness) operating at 40K

Total average power consumption (W) < 2.5 W (TBC)

average value of the total power consumption of an FSM (operating at 40K) including flight version of the associated electronics box operating at room temperature

Total peak power consumption (W)

peak value of the total power consumption of an FSM (operating at 40K) including flight version of the associated electronics box operating at room temperature

Thermal interfaces Thermal environment FSM: 40 ±2 K

Electronics: 290 ± 20 K

Optical interfaces

Input Beam divergence Collimated beam Output Beam divergence Collimated beam

Other

Exported vibration level TBD



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5.3. Optical Delay Lines

5.3.1 FUNCTIONALITY

An ODL is an opto-mechanical system that is able to introduce well-defined optical path variations without introducing significant wavefront errors, beam tilt and beam lateral deviation, in the full actuation range. In the context of high precision optical metrology (HPOM) systems and guidance, navigation and collision avoidance (GNC) control systems for formation flying, ODLs constitute the actuators of the “fine” OPD control loop. The ODL will be operated in three different modes:

• Fringe acquisition (mm range): to scan over the whole ODL range in order to find the coherence length range where fringes appear (interferences).

• Zero path difference locking (µm range): to scan within the fringe range (coherence

length of the light source) in order to find the zero path difference, where a constructive interference occurs (destructive in case an achromatic π-phase shifter has been introduced in one of the arms of the interferometer) for all wavelengths.

• Path stabilization and tracking (nm range): to compensate the OPD variations due to

optical path disturbances during the whole observation time (several hours), maintaining highly stable the required deep nulling value.

Table 20 Main technical requirements for the optical delay line





Optical stroke > ± 10 mm

Resolution < 0.5 nm Max. ODL rate 250 µm/s


5.3.2 DESCRIPTION

The ODL mechanism comprises an optical subsystem, an actuator and a mechanical subsystem costraining the motion of the optics along a line. Two technological solutions are being considered, as depicted in Figure 22 (see Section 10.2 for details):



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- a cat-eye delay line, guided by a magnetic suspension - a corner cube optics guided by a flexure mechanism

In both cases, the required stroke and with the required resolution can be obtained with a single-stage voice coil actuator. A semi-active launch lock mechanism compatible with cryogenic operation is also part of the ODL. Three ODL are needed in order to maintain complete symmetry of the optical trains. However, only two of them need to be actuated. All the ODLs must be placed in the beam combiner satellite, where the fringe tracker subsystem is located. The ODL moving mass is about 0.5 kg for both concepts. In order to minimize the torques induced by the moving parts, especially during the fringe acquisition phase, the delay lines shall be located as close as possible to the centre of mass of the satellite. Note that the operation of the ODL (as part of the “fine” OPD control loop) is in cooperation with other actuators and sensors (e.g. thrusters, fringe sensor unit, RF metrology, laser metrology, etc). Due to the limited stroke of the ODL, it is mandatory that the global GNC controller desaturates continuously the ODL in order to avoid the ODL reaching the edges. In other words, the ODL should operate in its central position under normal circumstances.

Figure 22 DARWIN delay line concepts by TPD/TNO (left) and CONTRAVES (right)

5.3.3 PERFORMANCE

The following table summarizes the required performances of all functionalities of the equipment. In particular, note that the ODL is itself a potential source of microvibrations, and therefore the design has minimize the exported forces/torques from the ODL to the optical bench.



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Table 21 Required performance for the Optical Delay Line

Performance

Required value

Remarks

ODL optical requirements Wavelength range 0.45 – 20 µm

Free optical diameter > 25 mm Output beam tilt

(over the full actuation range) < 0.05 arcsec

Output beam lateral deviation (decenter)

< 100 µm over the full actuation

range


Optical transmission losses < 15% from 0.65 to 4µm < 6% from 4 to 20µm

Coating manufacturing reproducibility • Relative spectral response • Chromatic phase differences • Relative polarization


< 10-4




< 0.1 ° < 0.1 °

ODL functional requirements

ODL stroke > ± 10 mm Optical stroke ODL resolution < 0.5 nm

ODL stability < 1 nm RMS

Assuming low measurement noise and low microvibration environment (100nm RMS, 1Hz cut-off, -40dB/dec slope).

ODL absolute position accuracy < 100 µm RMS ODL dynamic response

requirements

Max. ODL rate 250 µm/s Settling time (in closed loop) < 20 ms

Overshoot none



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Table 22 Interface requirements for the Optical Delay Line


Required value

Remarks

Volume (mm×mm×mm) 100 x 100 x 300

Mass (kg) < 10 kg (goal 6 kg)

Power

ODL average power consumption < 25 mW

average value of the power consumption dissipated in a dynamic ODL (including harness) operating at 40K

ODL peak power consumption

peak value of the power consumption dissipated in a dynamic ODL (including harness) operating at 40K

Total average power consumption (W) < 2.5 W

average value of the total power consumption of an ODL (operating at 40K) including flight version of the associated electronics box operating at room temperature

Total peak power consumption (W) total power consumption of an ODL

(operating at 40K) including electronics box Thermal interfaces Thermal environment ODL: 40 ±2 K

Electronics: 290 ± 20 K

Vacuum operation ODL: 10-6 mbar Elect.: ambient

pressure

Optical interfaces


Other

Exported vibration level

<50 microN RMS TBC

Residual exported force/torque during Fringe Acquisition Mode (ramp response). During nominal Tracking Mode the exported vibrations should be smaller.

Exported vibration frequency range High frequency noise.



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5.4. Differential Delay Line Unit

5.4.1 FUNCTIONALITY

The differential delay line (DDL) has the function of compensating differentials paths between the collector s/c and the hub occurring for the linear TTN. Resolution and operating bandwidth depend on the array reconfiguration requirements. The requirements listed in Table 20 refer to the baseline observation scenario, where the array size is constant during observation, or observation is stopped during resizing.

Table 23 Main technical requirements for the differential delay line





Optical stroke 2.4 m to 12.8 m Required delay depends on array size

Resolution 1 mm TBD


5.4.2 DESCRIPTION

No detailed concept has been developed yet. A typical DDL realization consists of two parallel mirrors, one of which can be translated in order to adjust the overall delay. The distance between mirrors and the incidence angle of the beam on the mirrors are dimensioned such that the beam undergoes multiple reflections before leaving the system. This makes it possible to introduce large pathlength differentials with a system with relatively moderate geometrical dimensions.


The DDL device itself will be complex and large. In practice, the distance between the DDL mirrors cannot be much shorter than about 1 m. Also, a long-stroke cryogenic actuator will be required in order to adjust the overall delay provided by the DLL. Note also that the device breaks the symmetry between the various optical trains, since the beam from the central CS undergoes a number of extra reflections as compared to the beam from the outer s/c. This can potentially result in a degradation of the nulling performance, whose magnitude is still TBD.



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The DDL could be avoided by placing the central SC further away from the hub, such that all three SC have the same distance from the hub. The array is no more linear, but resembles a flattened triangle with vertex angle ranging from about 178 degrees for compact arrays, to about 165 degrees for the most diluted ones. Although detailed analysis still needs to be performed, these arrays are anticipated to have lower modulation efficiency, as compared to the linTTN.

5.5. Fringe sensor

5.5.1 FUNCTIONALITY

The fringe sensor is a core component in the DARWIN system. It measures in real time the differential piston/tip/tilt between the telescope beams due to s/c displacements, thermal effects and vibrations. It also provides an estimate of the amplitudes of a selected number of higher order aberrations for offline investigation, at a slower rate (TBD) rate. The main technical requirements for the Fringe Sensor Unit are summarized in Table 24. Measurement rates have to be confirmed after a more accurate characterization of the expected input disturbances, see also Section 10.3.

Table 24 Technical requirements for the fringe sensor unit



Spectral range 600 nm - 1,5 µm Three input beams OPD measurement Range20µm

Repeatability 0,75nm RMS Rate 10Hz TBC

Tip/tilt measurement Range 18µm Repeatability 1,21nm RMS (equivalent to 0,05 arcsec)

Rate 10Hz TBC

Defocus (Zernike 4) measurement

Range 173µm Repeatability 9,23nm RMS

Rate 10Hz TBC

Astigmatism (Zernike 5,6)\ measurement

Range 400µm Repeatability 10nm RMS

Rate: TBD

Coma (Zernike 7,8) measurement

Range 400µm Repeatability 10nm RMS;

Rate: TBD

TrifoldComa (Zernike 9,10) measurement

Range 400µm Repeatability 10nm RMS;

Rate: TBD

Spherical Aberration (Zernike 11)

Range 90µm Repeatability 7.45nm RMS;

Rate: TBD

Limiting magnitude 11.1 (G-type stars) 12 (K type stars) 12 (M type stars)



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5.5.2 DESCRIPTION

The Fringe sensor consists of a Schmidt-Cassegrain telescope focussing three incoming beams, a defocus generator and a detector to acquire the fringes generated by these three input beams in the focal plane and in a out of focus plane. The schematic optical layout is shown in Figure 23.

Figure 23 Schematic drawing of the fringe sensor. Two beams only are shown. This FS is fed by the star light in the 0.4 – 1.5 micron range, which is extracted from the interferometer beams by means of dichroic mirrors, see Figure 10. The focused and defocused images contain interference patterns between the three beams. Detectors with about 500x500 pixel, with small pixel size (≈13µm), are required for adequate spatial sampling. The detected fringes are processed by phase diversity algorithms to derive the corresponding values for the primary aberrations for realtime (10Hz) correction of the DARWIN optical train (Piston, Tip, Tilt, Defocus (tbc)).



Issue: 1.2 Page: 65

5.5.3 PERFORMANCE

Table 25 Performance requirement for the Fringe Sensor Unit

Aberration

Required Remarks

Piston

(Zernike 1)

Range 20µm Repeatability 0,75nm RMS; Rate 10Hz

Note: Accuracy: Reference is defined by nuller

Tip/Tilt

(Zernike 2,3)

Range 18µm Repeatability 1,21nm RMS; (0,05 arcsec),

Rate 10Hz

The tip/tilt range tested with the BB is just limited by the aperture of the defocus generator used in the BB. For the FM a different focal plane setup will be selected which removes these restrictions.

Defocus (Zernike 4)

Range 173µm Repeatability 9,23nm RMS; Rate 10Hz

Currently, Defocus can only be estimated with iterative phase diversity or siclope algorithms. Thus it is not possible to detect it at 10 Hz.

Astigmatism

(Zernike 5,6)

Range 400µm Repeatability 10nm RMS; Rate: best effort

Up to now results only by simulation Experimental results tbc.

Coma (Zernike 7,8)



TrifoldComa

(Zernike 9,10)



Spherical Aberration

(Zernike 11)

Range 90µm Repeatability 7.45nm RMS; Rate: best effort




Document

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Table 26 Interface requirements for the fringe sensor unit

Interface

requirement

Value FM (expected) Remarks

Volume (mm×mm×mm)

300x300x900 FSTelescope300x300x300 FS FPA 300x300x100 Electronics (all tbc.)

Rough estimate assuming a stable passively compensated structural design and broadband detectors without need for isolating dewars. The FPA could be mounted outside the cryogenic beam combiner spacecraft payload.

Mass (kg) Telescope: 10 kg Focal Plane Assembly: 5 kg Electronics box: 4 kg

Estimations for telescope and focal plane assembly are just order of magnitude values: exact numbers have to be derived by structural analysis and after detector selection.

Accommodation requirements Cryogenic payload

Ideally placed on rotational axis of the Beam Combiner Spacecraft (rotational invariance); FPA assembly has to be thermally isolated

See Figure 24 for accommodation concepts with FPA mounted in or attached to the BCS cryogenic compartment.

Accommodation requirements “warm” electronics payload

The electronics box can be accommodated anywhere in the warm payload compartment of the BCS.

Average consumption (W)

20W tbc. (two detectors and all electronics) assuming detector operation at 40K. tbc.

Strongly dependant on detector choice and thermal concept (detector operational temperature /dewar). (in case operation at 40K is not feasible and dewars have to be applied to run the detectors at 150 K significant more power will be needed & dissipated)

Peak consumption (W)

- Not expected to be significant different from average consumption.

Operational temperature (K)

40 K The telescope as passive all aluminium structure will be integrated in the cryogenic compartment of the BCS. The FPA containing the Detectors might have to be isolated properly or placed outside the cryogenic compartment reducing any heat dissipation problems into the cryogenic compartment.

Operational temp. range (K)

± tbd /10 K Stability of the optical performance of the FS telescope and of the FPA geometry should not be critical for variation of a few K. The heat dissipation of the FS detectors inside the FPA has to be stable and carefully controlled) see accommodation.

Number of beams 3 3 beams tested during the BB study and assumed as minimum baseline for the final DWARF configuration.



Document

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Beam geometry See Figure 26 Derived from the BB study as target values

Wavelength Band 600 nm – 1500nm Resulting from the BB Study the extension in the NIR is required to be able to detect the M type stars down to the required magnitude and also to improve fringe acquisition performance (if fringe tracking is lost in the VIS range it can still be performed in the NIR range.

Input polarisation Stable and same orientation on all three input beams.

If the state of polarisation is different on the individual input beams (rotated due to the triangular DARWIN telescope configuration) this difference has to be compensated for at least two of the three input beams before entering the FS.

Data interfaces Schematic diagram See Figure 25

Data interface to ODL is unidirectional (tbc.) Data rate for transmission of the PA for piston tip tilt and defocus) to the ODL: ≤ 115,2 kBit/s (compatible with serial RS232) at 10Hz ≈ 10% load (tbc.)

Exported vibration level (µN)

tbc. if detector cooler is required it has to be microvibration compensated

In the ideal FS FM configuration with detectors operating at 40K no vibration generating component is required. If the detector development does not reach these requirements, complex cooler/heater/dewar designs are required. (tbd. after detector selection )

Radiation tolerance

10 krad (tbc) LET of 13 MeV-1cm-2 from protons (tbc)

DWARF Fringe Sensor

Beamcombiner, cryo optical bench

DWARF Fringe Sensorelectronics

DWARF Focal Plane AssemblyDetector Assembly @ 40 K

DWARF Fringe Sensor

Beamcombiner, cryo optical bench

DWARF Fringe Sensorelectronics

DWARF Focal Plane AssemblyDetector Assembly @ 40 K

Figure 24 Schematic drawing of FS accommodation with Focal Plane Assembly (which will dissipate tbd. W due to its detectors) inside the cryogenic part of the beam combiner spacecraft (left) and with the FPA at the interface of the cryogenic compartment. The front end electronics are integrated in the FPA, all other electronics DPU, supply etc. are accommodated in a warm electronics compartment.



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ODL 1Filtering Optics

DWARF Telescope

DWARF Control Unit(computer)

ODL (OPD)Controller(computer)

Input beams from telescopes

Cryogenic satellite environment (40K)

Output to cryogenicscience detector

DWARF Fringe SensorODL

ODL 2

ODL 3

Focal Plane Assemblyredundant detector assembly

40K

40 K

Satellite electronics compartment (

ODL 1Filtering Optics

DWARF Telescope

DWARF Control Unit(computer)

ODL (OPD)Controller(computer)

Input beams from telescopes

Cryogenic satellite environment (40K)

Output to cryogenicscience detector

DWARF Fringe SensorODL

ODL 2

ODL 3

Focal Plane Assemblyredundant detector assembly

40K

40 K

Satellite electronics compartment (

Command I/F

Optical I/F

Data I/F

Power

Thermal /Mechanical I/F

Sync 1MHz

Sync 1MHz ?

BUS OBDH system

270K tbc.)

Command I/F

Optical I/F

Data I/F

Power

Thermal /Mechanical I/F

Sync 1MHz

Sync 1MHz ?

BUS OBDH system

270K tbc.)

Figure 25 Schematic diagram of DWARF FS - System Interfaces

20,0

30,0

20,0

30,0

20,0

30,0

Figure 26 Optical interface beam geometry tbc.



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The presently assumed 10Hz fringe sensor rate may not be sufficient to reduce the DARWIN OPD environment as currently assumed (100nm RMS, 1Hz cut-off, -40dB/dec slope) down to 1nm RMS. The input OPD disturbance needs to be characterized in the detail by means of end-to-end simulations.

The internal OPD / tilt references of the fringe sensor need to be referenced to the ones in the science channel. This will require accurate calibration procedures and high performance thermal control of the associated part of the optical bench.

5.6. Modulator

5.6.1 FUNCTIONALITY

The modulator induces a periodic change in the optical transfer function of the interferometer such that the detector sees two different transmission maps alternating in time. Subtraction of the readings in the two modulation states removes uniform background contributions (for example, thermal emission) from the signal, while it modulates the signal from a planet transiting close to the peak in one transmission map and in a trough in the other transmission map.

Table 27 Main technical requirements for the optical delay line





Operating frequency 1-10 Hz Modulation frequency is not frozen yet


5.6.2 DESCRIPTION

The modulation finction has not been addressed in the detail yet. Only rough performance requirements and instrumental concept are available. For the TTN the simplest way to implement such modulation is by swapping the positions of two input beams, at a frequency of several Hz (TBD). Referring to Figure 9 this means, for example, that the beam collected by telescope B is exchanged for the beam collected by telescope C. Two possible design options have been considered for Darwin, that are shown in Figure 27.



Issue: 1.2 Page: 70

Figure 27 Possible modulator designs for TTN, using separate mirrors for the various beams (top design) and common mirrors for all beams (bottom).

The upper scheme makes use of a pair of moving mirrors per beam, that are steered simultaneously in order to displace the beams laterally. Beam C does not pass through the modulator. This may modulate the OPD between beam C and beams A,B at the modulation rate, and introduce a polarization unbalance between beams C and A,B.

The second design consists of three mirrors in a triangular configuration. One of the mirrors is switched in and out of the beams so that in that case the beam reflects only once (left), while in the other case, the beam reflects twice (right).

This second design is the preferred baseline since it is inherently symmetric, i.e., all three beams see the same type and number of reflections, in each of the modulation states.

5.6.3 PERFORMANCE

Key performance requirements for this device are expected to be the following:

• High reproducibility of the two mirror stand from one modulation cycle to the following one



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• Low settling time

• Low systematic differential distortions between the two transmission maps, that could lead to false detection

• Low induced microvibration level

Further formulation of these requirements is TBD.

Table 28 Required performance for the Modulator Unit

Performance

Required value

Remarks

optical requirements Wavelength range 0.45 – 20 µm

Free optical diameter > 25 mm

Differential beam tilt < 0.05 arcsec (TBC) Between the two modulation states

Differential output beam lateral deviation (decenter) < 100 µm (TBC) Between the two

modulation states


Optical transmission losses TBD Coating manufacturing reproducibility

• Relative spectral response • Chromatic phase differences • Relative polarization


< 10-4


It depends on corner cube manufacturing and angular mounting tolerance



< 0.1 ° < 0.1 °

functional requirements

OPD reproducibility TBD Between the two modulation states

tilt reproducibility TBD Between the two modulation states

dynamic response requirements

Operating frequency 1-10 Hz Modulation frequency is not frozen yet

Settling time (in closed loop) < 20 ms (TBC) Overshoot none



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Table 29 Interface requirements for the Modulator Unit


Required value

Remarks

Volume (mm×mm×mm) TBD Mass (kg) 3 (TBC)

Power

average power consumption TBD

average value of the power consumption dissipated in a dynamic device (including harness) operating at 40K

peak power consumption TBD

peak value of the power consumption dissipated in a dynamic device (including harness) operating at 40K

Total average power consumption (W) TBD

average value of the total power consumption of a device (operating at 40K) including flight version of the associated electronics box operating at room temperature

Total peak power consumption (W) total power consumption of a device

(operating at 40K) including electronics box Thermal interfaces Thermal environment 40 ±2 K

Optical interfaces


Other


< TBC microN RMS

Exported vibration frequency range High frequency noise.

5.6.5 OPEN ISSUES AND CRITICAL ITEMS

Modulation is a critical function. Beam swapping must occur with an extremely high repeatability in pathlength and tilt, short setting time and with an extremely low level of induced microvibrations.



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Depending on the design, modulation may result in the interruption of the optical path to the FT and other optical sensors at a rate of several Hz (TBD). These sensors should be able to operate correctly during modulation.

5.7. Beam Splitters and dichroic devices TBD

5.8. Achromatic Phase Shifter

5.8.1 FUNCTIONALITY

In the TTN design, starlight suppression relies on the introduction of a 120º phase shift among the interferometer beams. The main technical requirements for the phase shifter are summarized in Table 30. In particular, note that very stringent requirements are imposed on the achromaticity of the phase shift throughout the operating bandwidth. This achromaticity requirement cannot be met through the whole science bandwidth. Therefore, the science bandwidth will be split into two sub-bands, each one having its own dedicated APS device. Table 30 Technical requirements for the APS devices.



Spectral range 6 - 18 µm 4 - 20 µm

Nominal phase shift 120 degrees For TTN

Absolute accuracy of Phase shift

< 0.5 10-3 rad Through the nominal spectral range

Transmission > 95% Depends on A/R coatings

Rejection rate 106 (through 6 - 18 µm) Achievable on two sub-bandwidths (6-11 µm and 11-18 µm)

WFE <λ /20 rms @ λ=0.6µm Goal

Polarisation disturbance

≤ 1 · 10-2 rad

5.8.2 DESCRIPTION

While all-reflective APS concepts exist, they can only provide 180º phase shift and therefore they cannot be used for TTN. As shown in Figure 28, the selected APS method consists in introducing a given number of glass or dielectric plates in each arm of the interferometer. This method is inspired by the techniques used by optical designers to minimise lenses chromatic aberrations. This arrangement can be designed to produce any desired achromatic phase shift



Issue: 1.2 Page: 74

between the interferometer beams by optimizing the choice of the material and thickness of the plates.

1

2

3

4

5

6

Arm 1

Arm 2

translation

Figure 28 Scheme of the Dielectric Plates APS principle.

The total phase difference between the two arms of the interferometer results from the differential thickness of each couple of plates (see Figure 28 couples (1,2) (3,4) or (5,6) for example). Thus, in order to adjust the total phase difference, plates are wedged ones that can move perpendicularly to the beam axis. Thanks to that functionality, it is possible to compensate for deviations from temperature, from nominal plate’s dimensions The subsystem development study led to a design of the DPAPS based on the use of three plates of different materials. It has been demonstrated that this number of 3 plates is a minimum to achieve the requested performance to achieve the requested performance while the Darwin spectral band is split in two sub-bands only. Among the IR materials, three of them, namely Ge, KRS5 and ZnSe, have been chosen regarding their properties and in particular their theoretical refractive index dispersion. Beside, it has been demonstrated that the whole spectral band (6µm to 18µm) has to be split in two sub-bands (6µm to 11µm and 11µm to 18µm) to reach the target performance. In that case, geometrical parameters of every dielectric plates have been optimised. These ones are shown in the Table 31 and Table 32 below.

Table 31 Geometrical parameters for the dielectric plates for 6-11µm subband.

6-11µm Plate 1 Plate 2 Plate 3 Plate 4 Plate 5 Plate 6

Material Ge (35 K) KRS-5 (35 K) ZnSe (35 K) Thick. (mm) 6.000 6.159 6.000 4.621 6.000 6.359 Height (mm) 30 30 30 Length (mm) 60 60 60 Wedge (rad) -8.80 10-3 2.41 10-2 -6.50 10-3




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Material Ge (35K) KRS-5 (35K) ZnSe (35K) Thick. (mm) 6.000 5.772 6.000 5.653 6.000 6.057 Height (mm) 30 30 30 Length (mm) 60 60 60 Wedge (rad) -8.80 10-3 2.41 10-2 -6.50 10-3

5.8.3 PERFORMANCE

Figure 29 and Figure 30 describe the expected final nulling performance obtained from the design parameters of Table 31 and Table 32. The Figures demonstrate that the nulling remains below 10-6 with a strong margin throughout the whole 6-18µm spectral band. Figure 31, illustrates the performance of a variant of the same design, that is optimized for operation over an extended bandwidth of 11-20 micron. This design is still compliant with the 10-6 nulling ratio specification, even if the margins are modest as compared to the previous case.

Phase-shifter Nulling Ratio

1.00E-10

1.00E-09

1.00E-08

1.00E-07

1.00E-06

1.00E-05

6 6.5 7 7.5 8 8.5 9 9.5 10 10.5 1

Wavelength (microns)

Nul

ling

Rat

io

1

Figure 29 Nulling Ratio performance of the 3 dielectric plates on the 6-11µm sub band.



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1.00E-10

1.00E-09

1.00E-08

1.00E-07

1.00E-06

1.00E-05

11 12 13 14 15 16 17 18


Nul

ling

Rat

io


Phase-shifter Nulling Ratio Ge/KRS5/ZnSe

1.E-10

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

11 13 15 17 19 Wavelength [micron]

Null Ratio

Figure 31 Nulling ratio vs. wavelength for Ge/KRS5/ZnSe 3 plates APS, optimized for the 11µm to 20µm. Courtesy of Alcatel Space.



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Table 33 shows a comparison between requested and measured (or expected) performances on design parameters of DPAPS to match targeted specification.

Table 33 Comparison of required and measured (or expected) performance for the wedge APS

Performance

Required

Measured/ expected

Remarks

Wedge accuracy 7.10-4 rad 1.10-4 Expected value from supplier comment

thickness accuracy 50µm 50µm Plate Translation

range +/-20mm +/-20mm Resolution 5µm 5µm

5.8.4 INTERFACES AND PHYSICAL RESOURCE REQUIREMENTS

Three dispersive plates have to be moved linearly over a 40 mm range, in one arm of the interferometer, while the 3 others remain static. In the test bench, stepper motors actuate a linear table through ball bearing screws. When operating point is found, motors are stopped.

Table 34 Interface requirements


Value

Remarks

Volume (mm×mm×mm) 190 x 480 x 80 Motorized APS Mass (kg) 1.5 Motorized APS Volume (mm×mm×mm) 120 x 380 x 80 Static APS Mass (kg) 0.5 Static APS Accommodation requirements

Power

Average consumption (W) 1 Power is 5 W to find operating point, then 0 W. Cyclic rate is estimated to 0.2

Peak consumption (W) 5

Thermal interfaces

Operational temperature (K) 100 100 K in test breadboard Operational temp. range (K) 80 - 300 in test breadboard

Optical interfaces

Wave front quality plane wavefront λ/10 pv

Other

Exported vibration level (µN) TBD, zero during Darwin data acquisition



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5.9. Beam shaper

5.9.1 FUNCTIONALITY

The theoretical maximum optical power that can be coupled into a single mode waveguide is approximately 80% of the input power, assuming a flat intensity distribution over the pupil plane. This coupling loss is due to the mismatch between the Airy pattern, as created in the focal plane when the beam is imaged at the SMW tip, and the Gaussian shape of the fundamental mode in the SMW. The beam shaper has the task to maximize the efficiency with which beams are coupled into the Single Mode fiber, by re-shaping the input beams to match the Gaussian profile of the SMW main mode.

Table 35 Main technical requirements for the beam shaper



Optical Bandwidth 6-20 micron Darwin science bandwidth


Coupling efficiency > 90% (TBC) best achievable coupling efficiency without beam shaper is limited to 78%

Overall transmission > 90% (TBC) Implies a lossless design

5.9.2 DESCRIPTION

In the three-beams case, the coupling efficiency into the SMW is improved by adjustment of the input pupils to form a single circular one, with amplitude distribution as close as possible to the Gaussian pattern of the fundamental mode of the SMW. This can be done either by creating three input pupils, each having a 120° segment of each input beam, or by distorting the originally circular input pupils to approximate 120° circular segment, see Figure 32

A B C B B A CA

A B

C C B AC

Figure 32 Three different ways to combine three circular beams in a circular aperture, left: no beam shaper (leading to a coupling efficiency of 67%). Centre: splitting each beam in three segments and creating three combined apertures and right: re-shaping the beams to better fit the aperture. The latter two arrangements can yield a higher coupling efficiency.



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0.001 0.01 0.1 0.50

10 20 30 40 50 60 70 80 90

100

relative beam spacing ∆ d/d

inse

rtion

loss

[%]

3 segments full pupil

∆d

d

2 3 -3

-0.5

Total Coup-

ling Effici-ency [dB]

Beam Shaper

No Beam Shaper

4 5 6

Figure 33 Insertion losses for different pupil configurations. Left panel: insertion loss as a function of beam spacing for the case when three 120º segments are combined. Right panel: insertion loss as a function of number of input beams for the case of three circular beams (no beam shaper) and three distorted beams (beam shaper).

A different option consists in coupling only the central interference fringe to the fiber (see Figure 34). This system has a twice as small field of view as compared to coupling options considered above. The relatively low coupling efficiency of this technique (0.67, or 1.71 dB) is compensated by a four times more effective rejection of background radiation achieved by the reduced field of view. One that reducing the fov is an advantage only at wavelengths where background dominates, i.e., above 10 micron.

1 0.5 0 0.5 10

0.5

1

Normalized Airy disk radius

Nor

mal

ized

inte

nsity

0

0.5

1

Normalized Airy disk radius1 0.5 0 0.5 1

Planet at transmission max. Star

couple Into fiber

Figure 34 Red lines: interference patterns at the fiber tip for two-beams injection scheme for a planet at a transmission maximum (left) and for the star (right). Blue dashed lines indicate the Airy disk envelope. Green lines indicate the fundamental fiber mode.



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The beam shaper is the selected baseline pupil configuration, since it offers relatively high coupling efficiency and reducing the number of used fibres, i.e. relatively low hardware complexity. Optical scheme is TBD.

5.9.3 PERFORMANCE

Table 36 Required performance for the Beam Shaper Unit

Performance

Required value

Goal Remarks

Operating bandwidth 6-20 micron Science bandwidth fov 3 arcmin (TBC) Coupling efficiency > 90% (TBC)

Overall transmission > 90% (TBC) Implies a lossless design


Table 37 Interface requirements for the Beam Shaper Unit

RECEIVE TRANSFER OPTICS

Value

Remarks

Volume (mm×mm×mm) TBD Mass (kg) TBD Accommodation requirements

Power


Thermal interfaces

Operational temperature (K) 40 Same as for main telescope Operational temp. range (K) 0.1 (TBC) Same as for main telescope

Optical interfaces

Input beam diameter (mm) 20 Output beam diameter (mm) N.A.

Other




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5.10. Single Mode Waveguide Coupling Unit TBD

5.11. Single Mode Waveguide

5.11.1 SUBSYSTEM DESCRIPTION

The DARWIN planet finding mission requires that wavefront errors are reduced to a very high degree in order to achieve the required nulling quality. The required high wavefront quality can only be achieved with adequate wavefront filtering, as the one provided by single mode fibers. The single mode optical fiber makes use of the fact that in the operational spectral bandwidth of the fiber only the fundamental mode (of nearly Gaussian mode profile) propagates, whereas any higher order modes are strongly attenuated. Light can be coupled into an optical fiber with a high efficiency only over an optical bandwidth of about one octave. Therefore, at least two fibers will be needed in order to cover the wavelength region relevant for Darwin, i.e., 4-20 micron.

5.11.2 FUNCTIONALITY

A modal filter is an optical element that provides at its output a field with unique amplitude and phase distribution, irrespective of the input field. Single-mode waveguides show this behavior. The input field is projected onto the waveguide’s orthogonal eigenmodes Fi, i.e. the field is represented as Σi aiFi, where the ai are the modal amplitudes. If, by proper choice of wavelength, refractive indices, and waveguide geometry, only one mode can propagate (ai=0 for i ≠0), the waveguide acts as a modal filter.

Figure 35 Fundamental layout and principle of operation of a modal filter realized by a single-mode step-index fibre



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The basic setup of a modal filter is shown in the upper part of Figure 35. The input field Ein is focused into the waveguide’s input facet by a lens of focal length f and free radius a, resulting in the field Efoc. In general, an infinite number of modes are excited. However, all modes but the waveguide’s fundamental mode are radiated off, thus leading to a spatial steady state after a certain distance where only the fundamental mode is present in the vicinity of the fibre axis. The filter’s output field is characterized by a plane wavefront and an amplitude profile, which is solely determined by the waveguide’s physical properties. For an ideal waveguide, the output field is proportional to the waveguide’s fundamental mode F0, i.e. Eout= ζ.F0, where F0 is normalized to an overall power of unity. The peak amplitude and phase shift are determined by the complex field coupling efficiency ζ between the waveguide’s input field Efoc and the fundamental mode F0, i.e. by ζ = ∫∫AEfoc F0

* dA. With F0* the complex conjugate of eigenmode

F0, A is the coupling area. For an ideal waveguide, which provides only its fundamental mode at the output, the coupling efficiency ζ is the only quantity determining the filter action and the insertion loss of a modal filter. As indicated in the lower part of Figure 35, the output field always equals the waveguide’s fundamental mode, irrespective of the input field. Of course, input wavefront and amplitude profile perturbations cause a reduction of the overall output power (determined by the squared modulus of ζ) and a slight additional phase shift (given by the phase of ζ). The main technical requirements for the SMW are listed in Table 30. Table 38 Technical requirements for the single mode fiber



Spectral range 4 - 20 µm two or more SMW will be needed to cover the whole Darwin range

High-order mode suppression ratio

106

Total insertion loss < 1.5 dB (71% throughput)

5.11.3 DESCRIPTION

Several kind of fibers operating in the Darwin range are being studied, based on chalcogenide glasses (Te-As-Se (TAS) and GaAsSeTe (GAST)) for the 4-11 micron range, silver halides (for the 10-20 micron range) and Tellurium glasses, which could in principle cover the whole Darwin spectral range. Typical core dimensions of a few micron. Mid-IR fibers have a high refractive index, which causes strong back-reflections from the cladding/air interface. Therefore, and absorbing coating is mandatory to obtain single-mode operation within a reasonable fiber length (approx. 20-50 cm). The minimum fiber length needed in each sub-band will have to be determined during future development activities.



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The intrinsic absorption of the fibers is in general low. However it has been concluded that future fibers shall need AR coatings to avoid high Fresnel losses in its facets. In order to achieve broadband operation, special attention shall be given to the coupling optics and the need to split the overall spectral range into 2-4 bands. For each band the fiber parameters need be optimized to maximize the coupling efficiency.

5.11.4 PERFORMANCE

Table 39 Comparison between required and achieved performance of Single Mode Fibers

Parameter Required

Value Achieved Results Remarks

Operational wavelength range

4 – 20 µm 4 - > 10,6 µm (TAS fibers) > 5.8 µm (silver halide) ~ 3-11 µm (GAST fiber)

Tests performed at 10,6 µm

Higher order mode suppression ratio

10-6

Gaussian FFI distribution demonstrated at 10,6 µm

For all fiber types; best value ~ few hundreds, limited by measurement setup

Total insertion loss goal

< 1.5 dB (about 30%)

< 0,18 dB/cm (TAS fibers) < 0,23 dB/cm (silver halide) < 0,73 dB/cm (GAST fibers)

Intrinsic absorption is low Figures are with absorbtion coating Fresnel losses not included!

Operational temperature

40 Kelvin No mechanical or transmission degradation on bulk materials down to 77 K ( silver halide down to 4 K)

No teste performed on fibers yet

Radiation susceptibility

Not susceptible to radiation levels <104 Rads

No transmission degradation after exposure to < 10 kRads (TAS) Silver halide material is very susceptible to radiation but probably this is only a superficial effect

Fiber manufacturing process reproducibility

Should enable 5% or better parameter realisation

Difficult to control refractive index contrast between core and cladding (TAS fiber) No major problems with reproducibility (silver halides, GAST fibers)



Issue: 1.2 Page: 84


Table 40 Interface requirements for the SMW units


Value

Remarks

Volume (mm×mm×mm) 300x7x7 total length of fiber needs to be determined for each spectral interval. Bending is not desirable.

Mass (kg) a few grams The mass driver might be the need of proper shielding

Accommodation requirements

no bending Chalcogenide glass is very brittle so it needs careful handling and fiber should no be bended.

Power

Average consumption (W) none Peak consumption (W) none

Thermal interfaces

Operational temperature (K) any Temperature gradients during operation are undesirable

Operational temp. range (K) Between room temperature and 4 K

Optical interfaces

Numerical aperture of coupling lenses

match fiber numerical aperture to optimize coupling

optimized for each spectral interval for which the fiber is being used

Other


n.a.

Radiation shielding TBD

5.12. Spectrograph


Each of the modal filter units associated to the Darwin sub-bands is followed by a spectrograph unit, which disperses the planet signal onto the detector array associated to that sub-band. The number and characteristics of the spectrograph units will depend on how the Darwin science bandwidth is split into sub-bands. In the following, two bands, SW (short wavelengths, covering the 6-11 micron range) and LW (long wavelengths, covering the 11-20 micron range) are assumed.



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5.12.2 DESCRIPTION

The needed spectral resolution is driven by the scientific requirement of measuring the Equivalent Width2 of the H2O, CO2, and O3 absorption lines to an accuracy of 20% [RD 3]. In the following we use the absorption spectrum of the earth as a reference, see Figure 36.

Figure 36 Schematics of the infrared Earth spectrum showing a number of absorption lines due to key tracers for life, including the two water lines centered at 6.5 and 28.6 micron respectively

Early indications have shown that the width of the O3-line is independent of abundance, whereas the shape and width of the CO2-line changes significantly as a function of atmospheric composition. Compared to various models of exo-planetary atmospheres, the shape of the absorption line of the earth's atmosphere is somewhat average in width. To accurately determine the width, amplitude, and equivalent width of an absorption line, it follows from statistical considerations, that a sampling of 7 channels across the FWHM is required (corresponding to σ/3 for a Gaussian profile). This roughly yields the following requirements on resolution: Table 41 Assumed properties of absorption lines for the calculation of required spectral resolutions. Line widths are estimated from a typical earth type spectrum, and are only used for indication of resolution here. The mid-IR H2O line is centered at 28.6 micron. However, it will be measured in the 18-20 micron range since the instrumental transmission drops at longer wavelengths.

Absorption line

Center of line in µm

Assumed FWHM in µm

Required width of channel in µm

Resolution

H2O 6.5 1.39 0.20 32.5 O3 9.65 0.58 0.083 116

CO2 15 2.92 0.42 35.7 H2O 28.6 12.2 1.75 16.3

2 The Equivalent width of an absorption line is defined as the width of an ideal absorption line running all the way down from the continuum to zero, with the same surface area as the real feature,



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The table shows that a spectral resolution of about R =36 through the full science bandwidth would be sufficient, except for the 9.2-10.1 micron range (TBD) containing the ozone line, where a higher resolution of about R = 120 is needed. As explained in Section 5.8 and Section 5.11, a single Achromatic Phase Shifters or Single Mode Fiber cannot be designed to perform properly over the whole Darwin science bandwidth (6-20 micron). At least two sub-bands will be needed, each equipped with dedicated APS, fiber and spectrograph units. In the following we shall assume two spectral sub-bands, namely LW (short wavelengths, covering the 6-11 micron range) and LW (long wavelengths, covering the 6-11 micron range). The split at about 11 µm is driven by cross-calibration issues and the need to determine the blackbody continuum emission for a Earth-like spectrum. Then, the following technical requirements can be formulated:

Table 42 Technical requirements for the spectrometers. With TTN a minimum of two devices are needed, operating at short wavelengths (SW = 6-11 micron), and at long wavelengths (LW = 11-20 micron)

Parameter Requirement


Spectral bandwidth and resolution

SW: R = 36 between 6-11 micron, but

R = 120 between 9.3-10 micron (ozone range, TBD)

LW: R = 36 between 11-20 micron

Transmission > 95% (goal)

Operating Temperature 40 K

Fiber output

20 mm input beam

collimator

Focusingoptics

dispersingprism

FPA

Fiber output

20 mm input beam

collimator

Focusingoptics

dispersingprism

FPA

Figure 37 General scheme of a spectrograph unit



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The general scheme of both spectrographs is depicted in Figure 37 and is composed by a collimating mirror that collimates the light exiting the modal filter into a 20 mm diameter beam, a dispersive element, and a focusing element that matches the dispersed beam to the sub-band FPA. The dispersion will be achieved by a prism or grism. The distance between the dispersing element and the focal plane detector, is determined by the geometrical parameters of the prism/grism and by the smallest available pixel size of the detector. Considering that the dispersion requirements are moderate, the spectrograph design is driven by throughput and accommodation. There is therefore a slight preference for a prism. No detailed optical design is available yet for the spectrograph units. The LW spectrometer has a modest resolution of 36. Since the dispersion by a prism (or grism) is linear in wavelength, we have to sample the spectrum per sub-band at the smallest channels, resulting in 30 spectral bins assuming equal-sized detector pixels. The distance between the dispersing element and the focal plane detector, is determined by the geometrical parameters of the prism/grism and by the smallest available pixel size of the detector. Considering that the dispersion requirements are moderate, the spectrograph design is driven by throughput and accommodation. There is therefore a slight preference for a prism. The material of the dispersing element will be one of KRS-5, Ge, or ZnSe, to be selected mainly based on their transmission properties. Due to the smallest gradient of thermal expansion, ZnSe is preferred. The SW spectrometer is also based on a prism. The dual-resolution capability can be implemented in one of the following ways: a) Dimensioning the spectrometer for a resolution of R = 36, corresponding to 30 detector pixels. Higher resolution for the O3 line is achieved by fitting the corresponding portion of the detector plane with 120/36 ~ 3.5 times narrower detectors as compared to the ones used for the rest of the science bandwidth. About 7 to 10 detectors will be needed to cover the O3 line, depending on the assumed line boundaries, resulting in a total of about 35 detector elements. As for the LW spectrometer, the required spectral resolution can be readily achieved with a prism.

b) Implement a spectrometer with a resolution of 120 throughout the LW band, corresponding to a total of 99 detector pixels. For this amount of pixels, dark current and readout noise are still expected to be insignificant with respect to thermal and zodiacal background. Outside the wavelength range of the ozone line, the spectroscopic resolution can be reduced electronically at no sensitivity loss by re-binning adjacent spectral channels. This implementation requires about three times more detector pixels as the previous option. The first option is considered as baseline, since it is simpler and has a lower parasitic detector thermal load in virtue of the fewer pixels used. For both the SW and LW during the detection phase the spectral channels will be re-binned to yield a spectral resolution of a few (TBD) over the whole science bandwidth, in order to optimize the detection snr. Re-binning can be performed on-ground, as the impact on telemetry will be marginal, as the telemetry channel has to be sized to provide data at full spectral resolution. The level of spectral detail could thus be decided during data reduction. The spectral resolution is driven by the required spectroscopic resolution for planet characterization.



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5.12.3 PERFORMANCE

Table 43 Performance requirements for the Spectrograph Units

Performance

Required value

Goal Remarks

SW SPECTROGRAPH Operating bandwidth 6-11 micron (TBC) fov 3 arcmin (TBC) Transmission 80% (TBC)

LW SPECTROGRAPH Operating bandwidth 11-20 micron (TBC)

fov 5.5 arcmin (TBC)

Transmission 80% (TBC)


Table 44 Interface requirements for the spectrograph units


Value

Remarks

Volume (mm×mm×mm) SW: 100x100x50 mm LW: 100x100x50 mm

TBD

Mass (kg) NA Depends on instrument design Accommodation requirements NA Depends on instrument design

Power

Average consumption (W) 0 Peak consumption (W) 0

Thermal interfaces



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Operational temperature (K) 40 K Operational temp. range (K) NA

Optical interfaces

Entrance beam diameter 20 mm TBD

Other

Exported vibration level (µN) NA


The number and final design of the spectrographs depends on the chosen repartition of the Darwin science bandwidth into sub-bands, which is not frozen yet. The SW spectrograph must provide high resolution at the ozone line. A trade-off between two possible implementations must be performed.

5.13. FIR Linear Detector Array

5.13.1 DESCRIPTION

The performance requirements for Darwin are listed in Table 45. These are very similar to that of MIRI detector (5 – 28 µm) on JWST, which are also listed in the table. These requirements are met by existing detectors, namely Blocked Impurity Band (BIB) / Impurity Band Conductor (IBC) Si:As extrinsic semiconductor detectors. A Si:As array, produced by Raytheon, which has been selected for MIRI. The same technology is baselined for Darwin. However, Si:As detectors must operate at 8 K. As described in Section 10.9, alternative detector technologies are also investigated which could deliver similar performance with higher operating at temperatures.

Quantum Efficiency 80% over whole 50% @ 5 µm

Table 45 Comparison of Darwin and MIRI detector performance requirements

Parameter Darwin Requirement Darwin Goal

MIRI Requirement

Wavelength coverage 6 – 18 µm 4 – 22 µm 5 – 28 µm

Format Linear, ~35 pix (SW) Linear, ~ 30 pix (LW) 1024×1024

(3 chips of)

Pixel pitch 30 µm (~ 9 µm for O3 line) 18 – 30 µm

Operating temperature > 13-15 K > 40 K ~7 °K

Frame time ~3 s Integration time 1-10 s 1000 s Total noise (incl. Fower sampling) <19 e–

(goal 2.5 e–) Read noise for single read 10 e– 3 e–

Dark current < 25 e- sec-1 pixel-1 < 10 e- sec-1 pixel-1



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50% over whole spectral

range

spectral range

Well capacity 105 e– 105 e– (goal 2×105 e–)

Pixel crosstalk < 10-6 Power dissipation per chip

< 10 mW (at cold part of detector < 5 mW < 1 mW

In order to cover the entire Darwin spectral range, several detector modules optimised for smaller spectral intervals must be developed, possibly using different technologies per module. Integration times are determined by the frequency at which internal modulation is performed, see Section 3.2.8.

5.13.2 PERFORMANCE

See Table 45.


Table 46 Interface requirements for the FIR linear detector array


Value

Remarks

Volume (mm×mm×mm) NA Depends on instrument design Mass (kg) NA Depends on instrument design Accommodation requirements NA Depends on instrument design

Power

Average consumption (W) 10 mW Depends on instrument design Peak consumption (W) NA

Thermal interfaces

Operational temperature (K) 8 K Operational temp. range (K) NA

Optical interfaces

Depends on instrument design

Other




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5.14. Sorption Cooler


The detectors used in Darwin require a cooling power of 10mW at 5K and are very sensitive to micro-vibrations and temperature fluctuations. The lowest temperature that can be reached at L2 by passive radiation was estimated to be 40K if only few mW are dissipated. To cover the temperature range 50K-40K down to 5K, two independent Sorption Coolers will be implemented: H2 Sorption Cooler (provides 15K stage) and a He Sorption Cooler providing 10mW at 5K. This is a new technology under development, which is described in Section 10.10. The advantages of Sorption coolers are:

• They can be stored and launched in warm conditions.

• They are installed on cold environment (50K radiators) and have no connections with the warm part of S/C – reducing the parasitic heat and facilitate the integration

• Sorption coolers are almost vibration-free; the gas running through the piping being

the only source of noise; which is expected to be lower than 1µN/√ Hz.

5.14.2 DESCRIPTION

The cooling principle is based on isenthalpic expansion (change in pressure, at a constant enthalpy), when the gas is passing through a restriction area. The expansion can be on helium, to reach 5K or on hydrogen to reach 15K. In order to insure temperature stability the best working point is where the liquid/vapour coexist; the isotherms being horizontal lines in the P-H system. Small temperature fluctuations or changes in heat load will affect only the liquid quantity produced, not the temperature on the cold end, as far as the system is able to keep a constant pumping pressure. By changing the pumping pressure, the working point is moving from one isotherm to another isotherm producing temperature fluctuations.

Figure 38: a. Single stage Sorption Cooler schematic and the 4Two stage sorption cooler schematic.

52 K, 2.92 WA = 9.7 m2

A B C D

phases of a single compressor cell .b) shield

pL

pI

49 K, 1.65 WA = 6.5 m2

4.5 K,10 mW

to H2-stage



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The system insures a constant injection pressure or a constant pumping pressure by temperature change of an activated carbon. The activated carbon has the property of absorbing large quantity of gas, when its temperature decreases. Once the carbon is saturated in gas, by increasing the temperature, the gas will be release, creating high pressure in the system. The absorption is a physical process, which doesn’t degrade the carbon properties over time.

To insure a continuous process, the system needs minimum 4 independent cells, filled with activated carbon. By expanding the working gas, a high percentage of the nominal flow rate will be liquefied. The evaporation heat of the liquid is providing cooling at the interface with the instrument. To avoid having non-controlled liquid in the system, the excess liquid, not used by the Instrument will be evaporated in so-called evaporator. All the cells are connected to the cold end via check-valves. Each cell will go successively through 4 distinct phase: see Figure 38: A: warming up, B: desorption,C: passive cooling, D: active cooling.

The cooling of carbon is reached by heat sinking the cells on the passive radiator at 50K. When the carbon needs heated, the cells are insulated from radiator and heated-up. The sinking and the insulation of carbon to the radiator is made via a Gas Gap Heat Switch ( GGHS) which insure this function.

The helium expansion, to produce liquefaction, is made very close to the detectors interface; the connections are flexible and easy to implement. Part of the liquid produced will be hold in place at the detector interface plate, and part of liquid and the vapours will leave the interface to insure a continuous running. This will insure also a increased thermal stability at the interface with the detectors.

The actual design of 5K cold end contains a thermal shield at 15K provided by the H2 SC pre-cooling. It is important to be able to implement the 15K thermal shield in Darwin configuration thermal design.

5.14.3 PERFORMANCE

In table below are presented the performances for He and partial H2 Sorption Cooler.

Cooler mass for He 10Kg for H2

Table 47 Comparison of required and measured (or expected) performance for the He and partial H2 Sorption Cooler.

Performance of He Sorption Cooler

Required

Goal

Remarks

Cold plate temperature 5K Cooling capacity 10 mW Power consumption 10W Temperature fluctuation

< 1 mK for 1 h < 10 mK for 2 weeks < 100mK for 5 years

Requires active control PID. Temperature fluctuation of H2 stage is less important if still less than 0.1K.

Distance between radiators and cold tip

2m This distance is flexible. A CHEX is planed for 2m length.



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SC and He SC Supply voltage 28V Mechanical resonance frequency

Higher than 100Hz

lifetime 6 years during 5 years is: 275 000 Depressurisation profile

Withstand profile Ariane5

Mechanical strength -qualification 1.5 x operational load - yield strength 1.8 x operation load - ultimate strength 2.0 x operational load

Cold stage orientation System works independent of orientation

System works independent of orientation

Heat dissipation at 49K- stage 1

1.7W at 49K Average 1.7W The cooler performance improve if passive radiator can reach lower temperature than 49K during absorbing phase. -Stage 1 is heated 49K-80K


3W at 50K In average 3W at 50K

The maximum heat load is reached in the beginning phase of active cooling and depends on GGHS performance. -Stage 2 is heated 50K-120K

Precooling stage needed for He SC

-15K precooling T - 25mW cooling power


H2 SC dissipation on 80K radiator

3W 3W Stage 1 is heated 80K-120K Stage 2 is heated 80K-240K

Working pressure for He SC

Low P/ inter.P/ High P: 1.3b/ 4b/ 17b


Working pressure for H2 SC



GasGapHeatSwitch- for He SC

Resistance : -ON 2K/W -OFF 1000K/W


Degradation in time will be verified by test H2 is the working gas for GGHS

Check-valves for He SC;

*nominal flow: 0.8mg/s He *leakage flow closed direction

*nominal flow: 0.8mg/s He *leakage flow closed direction

Check-valves are identical for all stages. Operating Temperature 50K



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2µg/s *max dP in closed direction 20b *cracking pressure 1-10mb *pressure drop<15mb

2µg/s *max dP in closed direction 20b *cracking pressure 1-10mb *pressure drop<15mb


Table 48 Sorption Coolers interface requirements to the S/C


Value

Remarks

Compressor Cells Volume of single cell D=16mm L=150mm Attachment on radiator: 16x150mm

-copper interface Nb of cells: minimum 8 Mass (kg)/ cell 0.5 Accommodation requirements

extra volume is required on top of the cell to accommodating GGHS; -space allocated : D=50 L=200mm -minimum distance between cells shall be planed such that T fluctuations during active cooling are not propagate to the cells in adsorption phase.

The T fluctuation of outershell shall not perturb the detection

Distance between cells

Minimum distance between 2 cells can go from 5cm to 5m, depending on radiator accommodation to provide at any time temperature fluctuation on adjacent cells less than 0.5K

Effort shall be put in designing: • low mass radiators, • easy integration of the cells • if possible have late access

Check-valves unit 2 check-valves / cell 1 unit has 2 check-valves Volume of unit 40x20x10mm Is attached to the radiator Total number of check-valves 16 Mass 1 unit (2 check-valves) 0.07Kg/unit



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Accommodation requirements Require thermal coupling and attachment to the radiator

Heat exchanger CFHEX Volume Total number Mass Accommodation requirements Buffers Volume 2 litres In aluminium Total number 3 Mass 0.625Kg each Accommodation requirements Attached on radiator at

50K One buffer at each pressure level: 1b, 4b, 17b

Precooling +thermal shield at 15K

Volume D=100mm; L=20mm Requires a flange to install 15K thermal shield

Total number 1 Mass 0.5 Kg minim Accommodation requirements Requires mechanical

supports and thermal insulation from radiator

15K thermal shield has to cover the 5K stage to reduce the parasitic heat load

Cold End Volume To be considered

D=60mm; L=40mm Actual volume is small but has to be interconnected with the detectors

Total number 1 Mass 0.5Kg Accommodation requirements The actual supports are

in kevlar; it may require redesign.

Power

Average consumption (W) for He Sorption Cooler

5W

Peak consumption (W) for He Sorption Cooler

6W Some margins are required and consider 10W peak in power

Average consumption (W) for H2 Sorption Cooler

3W

Peak consumption (W) for H2 Sorption Cooler


Electrical harness 16 wires Goes along HEX

Thermal interfaces

Operational temperature (K) He Sorption Cooler Innershell

Innershell : 49K-80K stage1 50K-120K stage2

Operational temperature (K) Innershell :



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H2 Sorption Cooler Innershell 80K-120K stage1

80K-240K stage2 Operational temp. radiator (K) for He / H2 SC

-50K & 5W dissipation -80K & 3W dissipation

Beam Combiner Passively 40K Pre-cooling temperature 15K &25mW cooling

Other

Exported vibration level (µN) 1µg/ √Hz By analysis; measurements laser interferometers

Electromagnetic emission none measured

10Hz fringe sensor rate requirement may not be sufficient to reduce the DARWIN

6. ON-BOARD PROCESSING AND AUTONOMY

Due to the relatively small number of spectral channels, and long integration times, no substantial on board science data processing is foreseen, which would lead to a standard, or even low performance, on board data handling system. The GNC will interface to the following entities:

- Gyros (needed during deployment, detumbling and safety) and star trackers (coarse

s/c attitude sensors) - Wide field camera (fine s/c attitude sensors) - RF Navigation system (coarse s/c position sensor) - Laser metrology system (divergent, lateral, range: fine s/c position sensor) - Fringe sensor + delay line (OPD control loop) - TM/TC system - Launch structure propulsion and on board system - mN and µN thrusters (s/c position/attitude actuators)

The research and development activities, up to today, have led to a system sampling time of 1 second, having typically no control loop bandwidths faster than 0.1 Hz. GNC processing power requirement is TBD but is not considered to be critical. However, a number of sensors/control systems require fast local processing: RF metrology, see Section 11.3. OPD/fine tilt tracking: Both OPD and tilt will be measured by the fringe sensor FS, see Section 5.5. The FS is equipped with a 1k x 1k detector, which is read out at 10 Hz. Real-time data reduction for OPD and tilt has been demonstrated at this rate, using a conventional PC. Faster processing is needed to retrieve defocus, since an iterative algorithm is needed in this case. The data rate to the ODLs and fine pointing mirrors will be < 115 kbps. The data rate to the autofocus mechanism will be < 115 kbps. Note that the ODL technology development studies (see Section 10.2) indicate that the present



Issue: 1.2 Page: 97

microvibration environment down to 1nm RMS. If the need for higher fringe sensor rates will be confirmed, this will have a serious impact on the FS processing capabilities. Note that the modulation process, as described in Section 3.2.8, will interrupt the optical path to the fringe/tilt sensors at a rate of several Hz (TBD). This will interfere with the operation of the FS. Therefore, the FS detector must be occluded (with a diaphragm) during the switch/set phases of the modulator mirror,

- before modulation mirror is moved, the processing of the FS detector readouts is disabled; ODL/fine tilt mirrors are kept in their actual stand

- the modulation mirror is moved to its new position, setting time is allowed - normal operation of the FS is restored. As several sampling cycles have been

elapsed, some time will be needed in order to converge to actual OPD/tilt values. The AO optics of the VLT telescopes operates in similar conditions, since during chopping the VLT secondary mirror is tilted periodically, leaving the AO sensor without reference signal. However, the feasibility of such a procedure for Darwin must be confirmed by detailed studies. Wide Field Camera: The WFC is equipped with a 2k x 2k detector, which is read out at 1 Hz (TBC). WFC data processing equipment is TBD, but will be comparable to those of a high-performance star tracker. In this context it should be noted that the star catalogue would need to be larger than for a conventional star tracker, since the FoV of the WFC is smaller, but the limiting magnitude is much fainter.



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7. CALIBRATIONS AND CHECKOUT

7.1. Ground Calibration The satellite production, assembly and test will be staggered, reusing manufacturing and testing resources. This will ensure maximal use of gained experience. Staggering shall be applied for both telescopes, service modules and as far as possible for other payload elements. Environmental testing shall be performed up to spacecraft level, i.e. the spacecraft shall not be mounted together in the launch configuration for mechanical tests. Compatibility with the launch environment shall be demonstrated by means of analysis (TBC). The spacecraft-to-spacecraft interfaces will be tested functionally, using OGSE and EGSE. This means that the OGSE will have simulate telescope spacecraft as seen from the beam combiner spacecraft, and vice versa. Some of these tests should be performed in a cryogenic environment (e.g. fringe acquisition and lock). The OGSE will as a minimum consist of the following items:

• Narrowband source simulator, strong narrowband source used for photometric performance testing.

• Wideband source simulator, used for verification of interferometer performance over metrology and science wavelengths.

• Pupil generator used in combination with above source simulators to generate the input beams to the BCA.

• Engineering models of optical metrology systems that are distributed over several spacecraft, including:

• Retro-reflectors for tests of the laser metrology system • Sensors / sources mounted on other spacecraft • Telescope test equipment • Large diameter source • Sensor, measuring beam tilt, wavefront errors.

The EGSE will support standard spacecraft testing including thermal, mechanical and performance tests. The EGSE shall also support testing of the radio frequency equipment, i.e. RF metrology and ground communication. A special facility will be built in order to test formation flying. The tests shall, as far as possible, make use of “hardware-in-the-loop” (engineering models) and run in real time. The following systems shall be incorporated: RF and optical metrology, control system (APCS) and representative actuation system. It may prove necessary to build two separate test set-ups: one for the purpose of coarse maneouvres and one for fine control. The MGSE will consist of various handling, lifting and transport equipment as well as deployment test support equipment (zero-gravity simulators). Additionally there will be MGSE required to support the mechanical test programme (launch configuration not tested).



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7.2. In-flight Calibration The cruise phase to L2 will last for approximately 100 days. During this time the payload module should reach its operational temperature of 40 K (TBC). During the cruise phase the spacecraft subsystems will be checked out, including the metrology sub-systems and the formation flying. The last metrology system to be checked for performance will be the fringe sensor. When this has been completed, and the payload modules have reached the operational temperature, the scientific commissioning phase can commence. The duration of the commissioning phase shall be less than 3 months. The purpose of the in-orbit commissioning is to calibrate and verify the performance of the instrument before scientific operations. The optical as well as the formation flying performance of the instrument have to be characterised such that their influence on the output signal can be well described. A strong target star together with a relatively bright and well-characterised companion would allow for accurate calibration of the instrument. A binary system with equally bright stars is not suitable for calibration since the Fringe Tracker would equalize the optical paths for an object with a photo-centre in between the two stars, in which case deep nulling is not achievable. Alternatively, a star-Brown dwarf system could be used for instrument characterisation purposes. However, it would requires more integration time as compared to a binary system, but could also characterise the spectrograph’s performance and the arrays performance in nulling mode, as the spectra of the brown dwarf could be well known and a Brown dwarf is not as bright as a stellar companion. The flux on the detector should be considerably less than what could be expected for a binary system.

8. OPERATIONS

8.1. Baseline Operations When the instrument’s scientific performance has been established the science operations phase shall begin, which shall have a nominal duration of 5 years. There are basically two tasks to be completed during the science operations phase: planet detection and spectroscopy of detected planets. The amount of time that can be dedicated to the latter obviously depends on the results from the former, i.e. if few planets are detected during the first phase only little time needs to be allocated for spectroscopy of these few planets, and vice-versa. In this study it is assumed that 2 years of the science operations phase is required for planet detection and the remaining 3 years for spectroscopy. It should be noted that formation reconfigurations, e.g. slews and resizings (change of baselines), and fringe acquisition will require a non-insignificant fraction of the time during the science operations phase. It is assumed that science duty cycle (fraction of observation time as compared to elapsed time) is more than 70% during the science operations phase of the mission. This number needs to be consolidated in follow-on assessments.



Issue: 1.2 Page: 100

8.1.1 DETECTION PHASE

The targets for the Darwin observations will be selected from the target list, considering the current pointing direction and the remaining duration that the target can be observed. The latter is a consequence of the fact that each star is accessible only during a 3 month period every year. Obviously the targets will be observed in an order that minimizes the slew times. This objective may be in conflict with the strategy to observe the “easy targets”, i.e. the ones with the shortest integration time, first. The detailed planning and sequencing of the target stars will require a considerable work. We assume here that a reasonable sequencing of the target stars can be achieved. Once the target star has been selected, the interferometer is slewed to point in this direction. Note that this involves both attitude changes of the spacecraft as well as relative translations in order to maintain the formation configuration. The new target star will most likely require a different baseline. The baseline change could be performed simultaneously with the slew. It is expected that only one baseline per target star would be required during the detection phase. The optimal baseline will vary between 11m and 34m for 165 prime target stars in the Darwin target star catalogue [RD 5, RD 6]. However, “safe formation flying” imposes a minimum ISD of 15m. The minimum ISD will be defined by the sunshade radius and the distance needed for “safe formation flying”. Once the formation has been slewed to the new target star and the optimal baseline achieved the interferometer should be arranged to obtain fringe lock, i.e. the relative spacecraft positions and attitudes and the OPD control loop adjusted such that the stellar light from the target system interferes destructively. This is achieved by using a chain of metrology systems, each refining the measurements of the previous system. The coarse system, i.e. the RF metrology system, will always be operated as a fall-back system and allow to reconfigure the interferometer to a safe operational state. The fine metrology system, i.e. the fringe sensor, will finally bring the interferometer into fringe lock, which means that the optical path lengths, as seen by the stellar light when collected by the different telescopes, is identical to within a fraction of a wavelength (20 nm). The typical duration of a single detection run will be in the order of one day. However, each target has to be revisited a number of times, either to properly establish the orbital parameters of any planet detected, or to ensure that no planets were missed during earlier unsuccessful screening. In this study we have adopted an average of 3 visits, but again this number is not based on a detection strategy studied in detail.

8.1.2 SPECTROSCOPY PHASE

The observations during the spectroscopy phase will be either staring spectroscopy or rotating the formation, similar to the observations in the detection phase. The typical duration of a spectroscopy run is driven by water measurement, and will span between several days and a few hundreds days. The mission lifetime will not be restricted by consumables (FEEP for micro-propulsion and sorption cooler for detector cooling) (TBC). It is therefore possible to extended the mission once the nominal mission lifetime has been reached.




8.2. Telemetry Estimate The DARWIN requirements on telemetry bit rate are modest in that the payload will generate low-resolution spectra at a low frequency, due to the long integration times. Most of the telemetry will therefore be housekeeping data and is not expected to exceed 20 kbps. A telecommand rate of 4 kbps is deemed sufficient. Nominal communication between ground and the satellites occurs via the beam combiner spacecraft using a 0.5 m medium gain antenna, pointed to a 15 m antenna on ground, e.g. Kourou. The RF communication sub-system is based on X-band communication for the space / ground links. During the 8 hours of ground coverage repointing of the MGA may require that the science observations to be suspended until stable pointing has been reacquired. The inter-spacecraft communication is realized by means of a number of data-channels modulated on the RF metrology signal. In order to allow communication in non-nominal situations all satellites are equipped with 2 low-gain X-band antennas for ground communication. The telecommand signal is split by a 3dB coupler such that both receivers work in hot redundancy. The on-board system of the beam combiner automatically reconfigures from MGA to LGA in case of loss of telecommand signal.




9. APPENDIX A: TABLES

Table 49 Mass and power budget for the Beam Combiner Spacecraft [RD 6]

BC SPACECRAFT

Equipment No Mass[kg]

Totalmass [kg]

EclipsePower

[W]

OpsPower

[W]Structure Structure 1 N 137.7 137.7 0 0Thermal Sun shield (180 deg.) 1 N 76.1 76.1 0 0

Thermal control equipm. 1 Y 5.0 5.0 0 100Sorption cooler 2 Y 10.0 20.0 0 2003 m2 radiator 1 N 10.5 10.5 0 0

ACPS Coarse sun sensors 8 Y 0.0 0.2 1 1WFC incl. Pick-off & relay 1 Internal 6.0 6.0 5 5RF metrology incl. 7 antennas 2 Y 8.9 17.8 29 29Coarse lateral 4 Y 3.0 12.0 0 8Fine lateral 4 Y 3.0 12.0 0 0Absolute longitudinal (3 branches) 3 Partly 13.0 39.0 0 17Fringe sensor, (incl. elec. redundan 1 Internal 9.0 9.0 0 8

Propulsion In-FEEP Multiemitter (uN) 16 Internal 2.5 40.0 0 48FEEP power electronics 16 Internal 0.0 0.0 0 7.2Ion engine (mN) 16 Internal 1.8 28.8 15 720Ion engine fuel tank 1 N 1.0 1.0 0 0Ion engine power electronics 2 Internal 10.5 21.0 1.5 72

Power Solar Array 1 N 60.0 60.0 0 0Battery 2 Y 5.0 10.0 0 0

Avionics Low gain antenna 2 N 0.2 0.4 0 0Low gain transponder 2 Y 5.0 10.0 22 22X-band LGA antenna 2 N 0.5 1.0 10 0X-band MGA antenna 1 N 5.0 5.0 2 0X-band transponder 2 Y 7.0 14.0 0 25Computer 1 Internal 15.0 15.0 0 10Harness 1 N 30.0 30.0 0 0

Payload Relay. Telescope 6 N 3.0 18.0 0 0Receive telescope selector 1 N 10.0 10.0 2Long delay line (fixed) 1 N 50.0 50.0 0Optical Delay Line 3 N 3.5 10.5 2.5Beam combiner assemb. 1 N 250.0 250 0subtotal 701 85.5 1276.720 % margin 140DRY MASS 841




Table 50 Mass and power budget for the Collector Spacecraft [RD 6]

TELESCOPE SPACECRAFT

Equipment No Redun-dancy

Mass[kg]

Totalmass [kg]

EclipsePower

[W]

OpsPower

[W]Structure Structure 1 N 113.3 113.3 0 0

Mechanism connection SV / PLM 1 N 20.0 20.0 0 0Thermal Sun shield (180 deg.) 1 N 76.1 76.1 0 0

Thermal control equipm. 1 Y 5.0 5.0 0 100ACPS Coarse sun sensors 8 Y 0.0 0.2 1 1

WFC incl. Pick-off & relay 1 Internal 6.0 6.0 5 5RF metrology incl. 7 antennas 2 Y 8.9 17.8 29 29Coarse lateral (retroreflector) 2 Y 0.5 1.0 0 0Fine lateral 2 Y 0.5 1.0 0 8Absolute longitudinal 2 Y 0.5 1.0 0 0

Propulsion In-FEEP Multiemitter (uN) 16 Internal 2.5 40.0 0 48FEEP power electronics 16 Internal 0.0 0.0 0 7.2Ion engine (mN) 16 Internal 1.8 28.8 15 720Ion engine fuel tank 1 N 1.0 1.0 0 0Ion engine power electronics 2 Internal 10.5 21.0 1.5 72

Power Solar Array 1 N 50.0 50.0 0 0Battery 2 Y 5.0 10.0 0 0

Avionics Low gain antenna 2 N 0.2 0.4 0 0Low gain transponder 2 Y 5.0 10.0 22 22CDMS 1 Internal 15.0 15.0 10 10Harness 1 N 30.0 30.0 2 0

Payload Primary mirror 1 N 155.9 155.9 0 0Telescope tube 1 N 141.0 141.0 0 0Secondary + spider 1 N 10.0 10.0 0 0Relay. Telescope 1 N 3.0 3.0 0 0subtotal 758 85.5 1022.220 % margin 152DRY MASS 909




* These items will require complete redesign for a mission based on the 3D (Emma) TTN

10. APPENDIX B: TECHNOLOGY DEVELOPMENT STATUS

10.1. Development status: overview Table 51 summarizes the current development status for the Darwin payload. Development activities have been performed for a substantial number of components. The associated results and critical issues are discussed in more detail in the following sections. Table 51 Technology development status for the Darwin payload components

COLLECTOR SPACECRAFT Main telescope* Preliminary design inherited from the hexagonal array, needs to be

actualized to TTN+ Send transfer optics* Concept Wide Field camera* Concept BEAM COMBINER SPACECRAFT Receive transfer optics* Concept Optical Delay Line Ongoing manufacturing/testing of two breadboards representative

in form/fit/function of an engineering model. See Section 10.2 Differential Delay Line Concept Fine steering mirror Concept Fringe sensor Fine tilt sensor

Ongoing development of a functional demonstrator BB for a combined OPD/tilt sensor, see Dwarf activity in Section 10.3

Modulator Concept Beam splitters and dichroic devices

Concept. These components will be addressed during the planned activity “Passive components for interferometry”.

Achromatic phase shifters

Ongoing development of a cryogenic demonstrator BB, see Nulltimate activity in Section 10.5

SMW coupling unit A functional demonstrator BB for this component will be built and tested during the activity “Advanced Fiber Coupling Technology”

Single Mode Waveguide Various IR fibers (chalcogenide, silver halide, TAS) manufactured and demonstrated single-mode @ 10 µm. See Section 10.7.2

Integrated Optics for Darwin

First prototypes of hollow waveguides available, see Section 10.8 Development of IO based on Telluride glasses started

Spectrograph Concept FIR detector array Development of Quantum Well Infrared Photoconductors ongoing,

see Section 10.9. Activity on IBC Si:As BiB array to be started Sorption cooler Testing and performance characterization of a demonstrator

breadboard ongoing, see Section 10.10 Optical bench Dedicated activity “High stability Optical Bench” to be started




10.2. Optical Delay Line

10.2.1 DEVELOPMENT STATUS

Two parallel activities have been run with the objective of designing, manufacturing and testing a breadboard of an ODL that is representative in form/fit/function of an engineering model (EM). The activities, whose prime contractors were Contraves Space AG and TPD/TNO, and are now in the test/performance characterization phase. The ODL mechanism contains an optical subsystem and a mechanical subsystem. For the Contraves design the optical subsystem comprises a cornercube and the associated mounting interface. The mechanical subsystem consists of a voice coil actuator that moves the cornercube along the stroke by means of a guidance mechanism based on flexure technology. The ODL has a single stage actuation and it is decoupled from the baseplate by an intermediate isolation suspension stage. A semi-active launch lock mechanism compatible with cryogenic operation is also part of the ODL. For the TPD/TNO design, the optical subsystem consists of a 2-mirror cat ’s eye mounted on the moving part of the ODL. If needed, a cat ’s eye could accommodate an optional pupil imaging (not implemented in the breadboard). An isostatic mounting interfaces with the moving part of the actuator. The mechanical subsystem consists of a voice coil actuator that displaces the moving structure containing the 2-mirror cat’s eye (both mirrors simultaneously) along the stroke. The guidance mechanism is based on magnetic bearings technology with no friction and no wear. The ODL has a single stage actuation. A 1-g compensation device is included in the dynamic ODL. Operation in horizontal position is mandatory. The breadboard includes two separate ODLs, a fixed one and a dynamic one. The design of the ODL and the selection of space-qualified materials/components are driven by the DARWIN science/metrology spectral range (0.65 –20 µm), the radiation (20 krad) and the thermal environment in the hub (vacuum, 40 K). Emphasis has been put on minimizing heat generation at the ODL, and imported micro-vibrations. The breadboard shall be subjected to a functional performance verification campaign for the on ground validation of the developed concept. Full mechanical testing under the expected launch environment is not required. However, the Ariane 5 launch environment is considered in the design of the ODL. The locking mechanism, required to survive the launch, has been investigated at conceptual level and its characteristics have been included in the ODL performances (e.g. mass, size, power consumption). The static delay line have exactly the same optical layout but the mechanical support is simpler. The opto-mechanical units (i.e. static and dynamic ODLs) are compatible with vacuum environment (from 10-6 mbar to ambient pressure), cryogenic conditions (from 40K to 25°C) and 1-g operation. The rest of units (i.e. electronics, controller and user I/F and the optical test equipment) all operate at ambient pressure and at room temperature.




Figure 39 ODL engineering overview drawing (Contraves)




Figure ODL engineering overview drawing (TPD) 40




Figure 41 ODL engineering overview drawing (TPD, continued)




requirements

10.2.2 PERFORMANCE

The following table summarizes the required and expected performances of all functionalities of the equipment. At the present stage of the activity no tests results are available; compliance is expected for all items, for both BB. Table 52 Comparison of required and expected performance

Performance

Required value

Expected

value

Remarks

ODL optical requirements

Wavelength range 0.45 – 20 µm C Free optical diameter > 25 mm C

Output beam tilt (over the full actuation

range) < 0.05 arcsec C

Output beam lateral deviation (decentre)

(over the full actuation range)

< 100 µm C

Wavefront distortion < λ/20 RMS (λ=633 nm) C TBC by test results Optical transmission

losses < 15% from 0.65 to 4µm < 6% from 4 to 20µm

C C

In the case of no protective coating

Coating manufacturing reproducibility

• Relative spectral response

• Chromatic phase differences (OPD contribution)

• Relative polarization

o Rotation o Ellipticit

y

< 10-4

< 0.1 nm RMS (for λ<1µm)

<λ/10000 (for λ≥1µm )

< 0.1 ° < 0.1 °

C

C

C C

Polarization variation (over the full actuation

range) • Rotation • Ellipticity

< 0.1 ° < 0.1 °

C C

ODL functional




version” electronics

ODL stroke > ± 10 mm C Optical stroke (physical stroke > ± 5 mm) TBC by test results

ODL resolution < 0.5 nm C

ODL stability < 1 nm RMS C

The sampling frequency of the fringe sensor available in DARWIN will determine the OPD that can be compensated for. In the breadboard a high frequency interferometer system will be used for OPD control in an earth environment.

ODL absolute position accuracy < 100 µm RMS C

ODL dynamic response requirements

Max. ODL rate 250 µm/s C Presently amply margin Settling time (in closed

loop) < 20 ms C TBC by test results

Overshoot Damping factor > 0.9 C C (compliant) TBC (to be confirmed)


The following table summarizes all relevant interface items for the ODL Table 53 Interface requirements


Required

value

Expected

value Contraves

Expected value

TPD/TNO

Remarks

Volume (mm×mm×mm)

100 x 100 x 300

170 x 130 x 250

114 x 116 x 210 (w/o launch lock

mechanism) 114 x 116 x 280

(with launch lock mechanism)

The non-compliance is accepted

Mass (kg) < 10 kg (goal 6 kg) <6 kg <6 kg

Total mass of the breadboard, i.e. fixed and dynamic ODLs (including mass of the associated “flight




module).


Roughness TBCµm

Power

ODL average power consumption

< 25 mW

< 10 mW < 20 mW

average value of the power consumption dissipated in a dynamic ODL (including harness) operating at 40K and earth environment

ODL peak power consumption

< 50 mW

< 10 W (at 300 K)

< 1 W (at 40 K)

peak value of the power consumption dissipated in a dynamic ODL (including harness) operating at 40K and earth environment

Total average power consumption (W)

< 2.5 W < 2.5 W < 2.5 W

average value of the total power consumption of an ODL (operating at 40K) including flight version of the associated electronics box operating at room temperature

Total peak power consumption (W)

< 8 W

< 13 W (ODL at 300 K)

< 4 W (ODL at 40 K)

peak value of the total power consumption of an ODL (operating at 40K) including flight version of the associated electronics box operating at room temperature

Thermal interfaces

Thermal environment

ODL: 40 ±2 K Electronics: 290 ± 20 K

C C

Vacuum operation

ODL: 10-6 mbar Elect.:

ambient pressure

C C




compliant with the necessary control bandwidth is used. As a conclusion, a better

Optical interfaces

Input Beam position

Input Beam divergence Collimated

beam Collimated

beam

Output Beam position

Output Beam divergence Collimated

beam Collimated

beam

Other


<50 µN RMS <50 µN RMS

<50 µN RMS <1 mN RMS

Residual exported force during nominal Tracking Mode due to the actuator (space and ground environment respectively). The residual exported forces due to the guidance mechanism are expected ~1mN in any environment.

Exported vibration frequency range

50-500 Hz ~0.1-100 Hz due to the actuator.

C (compliant) TBC (to be confirmed)


The main critical points associated to the ODL are summarized in the following list:

• ODL stability testing (requirement 1nm RMS) under ground laboratory conditions. The ground microvibrations levels are much larger than the (possibly representative) DARWIN microvibration environment specified in the SOW (100nm RMS, 1Hz cut-off, -40dB/dec slope). Therefore the 1nm RMS ODL stability is found to be challenging given typical ground microvibrations levels.

• Usable control bandwidth. The ODL is not a standalone system. It operates in conjunction with other subsystems (e.g. fringe sensor unit, OPD controller) and its ultimate performances depend on the performances of the other subsystems. The present 10Hz fringe sensor rate is not sufficient to reduce the (possibly representative) DARWIN microvibration environment specified in the SOW (100nm RMS, 1Hz cut-off, -40dB/dec slope) down to 1nm RMS. For the ground testing, a fast interferometer with a rate




estimation of a representative DARWIN microvibration environment (inducing OPD fluctuations in the science beams) is mandatory. Based on this microvibration environment, better estimations of the needed fringe sensor rate can be made and the impact to the global linkbudget.

• Overall testing at 40K. This activity has to prove operation of the ODL in cryogenic temperatures (40K). This has an impact on the overall design and type of materials to be used. In addition, it does complicate the testing of the ODL (cooling system, good thermal conductivity to avoid thermal gradients, minimization of heat generation, more complex test set up, etc). In particular, in the TPD design the moving part of the ODL is physically decoupled (i.e. contactless). Then, the time required to cool the complete ODL down 40K increases significantly and it has to be considered for ground testing (not an issue for space).

• In order to test the ODL in 1-g environment, a 1-g compensation device is included in the breadboard. An extra mechanism to remove the 1-g compensation device once in space is necessary (not included in the breadboard). Testing under 1-g conditions impose the ODL to be oriented in certain directions. In this respect, it restricts the ODL orientations in which the ODL performances are compliant with specifications. This could have an impact at system level where the full hub spacecraft (with several ODLs) has to be tested on ground. ODL designs that are not orientation dependent would therefore be preferable.

• Optical coatings. There are materials with optical properties compatible with the DARWIN wavelength range requirement. Presently, such materials are extremely fragile and are normally protected by means of protective coatings. However such protective coatings might introduce some asymmetry between the different arms of the interferometer, and the accumulated effect may ruin the nulling measurement. The development of robust and reliable un-protected coatings is mandatory (not only for the ODLs).

• Minimization of exported forces/torques from the ODLs to the spacecraft. Likewise the ODL has to be well decoupled of any source of vibration, the ODL is itself a potential source of microvibrations, and therefore the design has minimize the exported forces/torques from the ODL to the optical bench. The residual exported forces due to the guidance mechanism are expected to be ~1mN both in earth and ground environment.

• Fringe acquisition mode (FAM). The objective of the FAM is to search for the fringe envelope (coherence length) minimizing both the Fringe Acquisition Time (FAT) and the exported forces of the ODL to the spacecraft. A trade-off between different approaches to drive the ODL during FAM needs to be performed at system level.




10.3. Fringe sensor


The purpose of the fringe sensor activity is to identify and validate a high-precision fringe detection and tracking sensor called DWARF (DARWIN Astronomical Fringe Sensor) by setting up, characterizing and testing a respective sensor breadboard. The main study goals were: The identification and validation of the best high precision fringe detection and tracking

sensor concept to measure all relevant perturbations and to provide the necessary information to cophase the DARWIN free-flying telescopes

The definition, preliminary and detailed design of a demonstration breadboard Manufacturing, characterisation and performance testing of a representative breadboard The consideration of space technological issues in order to allow a direct derivation of a

space compatible design Development planning for a future DWARF flight model

The activity was carried out by Kayser-Threde as prime contractor, supported by the subcontractors ONERA and Alcatel Space.


The DWARF BB functionality is shortly summarized as follows:

Measurement of differential wave front perturbations of at least three parallel input beams

Calculation of piston, tip, tilt from BB data sets at a 10 Hz rate, all others, i.e. Zernike coefficients 4 through 11, at a slower rate

Breadboard design compatible to the ONERA OTE test bench (used for BB performance testing)

BB compatible to operation under normal laboratory conditions (RT, no vacuum)

10.3.3 DESCRIPTION

The DWARF Fringe sensor as designed for the BB study consists of a Schmidt-Cassegrain telescope focussing three incoming beams, a defocus generator and a detector to acquire the fringes generated by these three input beams in the focal plane and in a out of focus plane. The schematic optical layout is shown in Figure 42. In order to be able to test the performance in the BB study with just one detector the defocus generator was realised as shown in Figure 43.




50/50Beam-splitter

DWARF FS telescope

“Defocus” Detector

∆f

50/50Beam-splitter

DWARF FS telescope

“Defocus” Detector

∆f

Figure 42 Schematic drawing of the fringe sensor concept realised in the BB study

Figure 43 Defocus generator as realised for the DWARF BB Study. This setup allows to measure both focussed and defocussed fringes on one detector; DFG P1 is a 50/50% beamsplitter optimised for 633 nm)

The Optical Test Equipment (OTE) at ONERA consists of an optical bench that generates 3 separate collimated beams from one point source and delivers these three beams to the DWARF FS. Actuated mirrors inside the OTE optical train allow to introduce defined primary aberrations. A 90° folding mirror between OTE and DWARF telescope allows to introduce known amounts of higher order aberrations (see Figure 42). The detected fringes are processed by analytical or numerical algorithms to derive the corresponding values for the primary aberrations for realtime (10Hz) correction of the DARWIN optical train (Piston, Tip, Tilt, Defocus (tbc)). The tests during the BB study where performed in the visible (VIS) spectral range with various source bandwidths and various point source intensities, simulation the relevant star brightnesses expected in the DARWIN mission.




Figure 44 Schematic diagram showing Optical Test Equipment (OTE) and the DWARF FS System

Figure 45 CAD 3D view of the DWARF FS BB with interface bellows, DFG folding mirror housing, DWARF telescope and Defocus generator mounted in front of the CCD camera. All components are mounted on the BB optical bench.




Figure 46 DWARF FS BB (left side: baseplate with camera, telescope and DTE folding mirror) mounted on the OTE bench at ONERA.




10.3.4 PERFORMANCE

Table 54 Conditions / assumptions under which the performance has been demonstrated (BB) respectively is expected (FM):

Operational requirements / assumptions

Assumption for BB study Assumption for DARWIN FM design concept

Operating wavelength BB operational in Vis -band 600 nm - 1,5 µm

Separation between input beams Defined by OTE pupil mask; about 30mm

Telescope diameter Darwin: 1,5 m Darwin: 1,5 m Telescope surface Darwin: 1,77 m² Darwin: 1,77 m² Optical throughput Darwin: 0,25 Darwin: 0,25

Light collection power of optical input

Nbr. of telescopes 6 6

or equivalent light collecting power with less than 6, but larger telescopes

Operating temperature RT Darwin: 40 K Operating atmosphere air vacuum

Radiation tolerance N/A for BB Darwin: 10 krad (tbc) LET of 13 MeV-1cm-2 from protons (tbc)

Lifetime N/A for BB Darwin: 5 (ground) + 7 (orbit) years Number of input beams 3 (optical power corresponding to

6 DARWIN telescopes) Available actuators upstream OTE, DTE Darwin: ODL and TTC Other available resources N/A for BB Darwin: Nuller spectrometer output

Budget for the nuller OPD N/A for BB Darwin: 20 nm Budget for the OPD control system

NA for BB Darwin: 5 nm

Free aperture for each DWARF beam

20 mm 20 mm

Darwin: output from BB study

See

Figure 26

Power dissipation N/A for BB Darwin: to be minimized

Darwin: < 6




Table 55 Comparison of required and measured performance for the DWARF FS BB activities

Performance Required Measured (BB study) Remark

(Zernike 1)

PA Linearity: ok Repeatability: ok Max rate 10 Hz: ok Acquisition range: fringes detectable over 10 µm

Note: Accuracy: Reference is defined by nuller

Tip/Tilt

(Zernike 2,3)

PA Linearity: ok Repeatability: ok Max rate 10 Hz: ok Acquisition range: Notes: Full Tip/tilt range is not testableAccuracy: Reference is defined by fibre head

Range 18µm Repeatability 1,21nm RMS; (0,05 arcsec),

Rate 10Hz

The tip/tilt range tested with the BB is just limited by the aperture of the defocus generator used in the BB. For the FM a different focal plane setup will be selected which removes these restrictions.

PA Range 173µm Repeatability 9,23nm RMS; Rate 10Hz

Currently, Defocus can only be estimated with iterative phase diversity or siclope algorithms. Thus it is not possible to detect it at 10 Hz.

(Zernike 5,6)

HA

Coma (Zernike 7,8)

HA Range 400µm Repeatability 10nm RMS; Rate: best effort

Up to now results only by simulationExperimental results tbc.

Accuracy: tbc. Repeatability: achieved

TrifoldComa

(Zernike 9,10)

HA Range 400µm Repeatability 10nm RMS; Rate: best effort



Spherical Aberration

(Zernike 11)

HA Range 90µm Repeatability 7.45nm RMS; Rate: best effort



G type stars to be used between magnitude

3.5 – 11.1 Experimental data exceeds the limiting magnitude 11.1

Assumption of 6 DARWIN telescopes with 1,77 m² collecting surface each

K type stars to be used between magnitude-

2 - 12 Experimental data exceeds the limiting magnitude 12


M type stars to be used between magnitude

5.8 - 12 Experimental data exceeds the limiting magnitude 12

Piston Range 20µm Repeatability 0,75nm RMS; Rate 10Hz

Defocus (Zernike 4)

Accuracy: tbc Repeatability: tbc Max Rate 10 Hz not feasible !

Astigmatism Range 400µm Repeatability 10nm RMS; Rate: best effort








The following table summarises open points and critical issues, which have to addressed in order to be able to design an optimised DWARF FS FM. Especially the detector and beamsplitter issue need to be addressed to allow an elegant and Darwin system-compatible design. Table 56 DWARF FS open points and critical issues for system design and FM development

Open points / critical issues

Description Remark

Detector issue Development on the detector manufacturer side in the direction of 1Mega pixel broadband detector (600nm -1500nm) with small pixel size (≈13µm). Optimization of detector materials, packaging and front-end electronics for cryogenic operation (40K).

With currently available detector technology the DWARF FM detector assembly will be quite complex, heavy and will require dedicated dewars. Broadband detectors that are able to operate at cryogenic temperatures seem to be within reach of midterm detector development.

FS onboard calibration and Stability

Detailed analysis of structural design/ stability for an all-aluminium FS telescope and a cryogenic FPA is required to evaluate the necessity of active calibration during flight. In case of an active alignable FPA (/Focus/Defocus) experimental HW demonstration should be performed (possibly with the adapted DWARF BB).)

Focus and defocus location relative to the FS detectors have to be known to a high accuracy during flight. Depending on the selected design concept (very stiff and stable passively compensated cryogenic design versus active alignable detector assembly) a calibration during flight might be required (WFE , Focus/Defocus) and needs further investigation

Realtime detection of Defocus

A real time, analytical version of these phase retrieval /siclope algorithms has to be derived to allow detection of Defocus at 10Hz or faster.

Currently, defocus can only be estimated with iterative phase diversity or siclope algorithms. Thus it is not yet possible to detect it at 10 Hz.

Siclope Algorithm

Mathematical reshaping and optimisation of the algorithm shall allow fast analytical operation at 10 Hz or faster.

The Siclope algorithm, as tested during the BB activities is a numerical algorithm, which has to run through a high number of iterations to calculate the required values. It is therefore not capable to provide the required results at 10Hz or faster.

40 K Beamsplitter

In order to be able to integrate the FPA of the FS into the cryogenic

Dichroic beamsplitters will be required throughout the whole




environment of the beam combiner spacecraft the required dichroic beamsplitters have to be mounted and have to perform without degradation at 40K. (thermal stress, stability). Respective investigations have to be performed.

DARWIN system, e.g. in front of the fringe sensor, where the short wavelength FS wavelength band is split from the long-wavelength science band. This issue could be addressed for system and FS simultaneously.




10.4. Beam Splitters and dichroic devices


These components will be addressed during the planned technology development activity “Passive components for interferometry”. Starting date TBD.

10.5. Achromatic phase shifters


The purpose of the Nulltimate activity (ESA contract n°17 005/02/NL/JA) is to demonstrate deep nulling interferometry at cryogenic temperatures and in the thermal spectral band domain (6µm to 18µm). The activity’s goal is to demonstrate three different Achromatic Phase Shifting functions. Two reflective APS designs are considered, which provide a 180 degrees phase shift. The third design is based on a dispersive plates arrangement and can produce any desired phase shift. This solution is baselined for TTN, where 120º phase shifts are required. A separate technology development study has been recently initiated addressing a different APS concept, based on Total Internal Reflection Zeroth Order Grating (TIRG APS). Zeroth Order Grating (ZOG) consists of sub-wavelength structures for which only zeroth-order transmitted and reflected orders are allowed to propagate. The use of ZOG structures as antireflection layer for Darwin refractive optics (beam splitters, cross combiners, wedge APS) is also being investigated


The proposed solution is based on a dispersive plates arrangement. This activity is divided in two distinct phases. During Phase 1, a model of the proposed solution has been developed that allowed for a first design of the Dispersive Plates Achromatic Phase Shifter (DPAPS). Both performances and parameters sensitivity have been studied (see document NULLTIMATE CDR 2004, October 15th Technical Notes). During the second part of the activity, so called Phase 2, the design parameters of the DPAPS will be adjusted following measurements of material characteristics (in particular the refractive index as a function of wavelength). The main technical requirements for the phase shifter are summarized in Table 57. Note that very stringent requirements are imposed on the achromaticity of the phase shift throughout the operating bandwidth. This achromaticity requirement cannot be met through the whole science bandwidth. Therefore, the science bandwidth will be split into two sub-bands, each one having its own dedicated APS device.




Table 57 Technical requirements for the APS devices. The phase shift requirement applies to the TTN configuration.



Spectral range 6 - 18 µm 4 - 20 µm

Nominal phase shift 120 degrees For TTN

Absolute accuracy of Phase shift

< 0.5 10-3 rad Through the nominal spectral range

Transmission > 95% Depends on A/R coatings

Rejection rate 106 (through 6 - 18 µm) Achievable on two sub-bandwidths (6-11 µm and 11-18 µm)

WFE <λ /20 rms @ λ=0.6µm Goal

Polarisation disturbance

≤ 1 · 10-2 rad

In the TTN design, starlight suppression relies on the introduction of a 120º phase shift among the interferometer beams. All-reflective APS concepts can only provide 180º phase shift. As shown in Figure 9, the selected APS method consists in introducing along each arm of the interferometer a given number of glass or dielectric plates. This method is inspired by the techniques used by optical designers to minimise lenses chromatic aberrations. This arrangement can be designed to produce any desired achromatic phase shift between the interferometer beams by optimizing the choice of the material and thickness of the plates.

1

2

3

4

5

6

Arm 1

Arm 2

translation

Figure 47 Scheme of the Dielectric Plates APS principle.




The total phase difference between the two arms of the interferometer results from the differential thickness of each couple of plates (see Figure 9 couples (1,2) (3,4) or (5,6) for example). Thus, in order to adjust the total phase difference, plates are wedged ones that can move perpendicularly to the beam axis. Thanks to that functionality, it is possible to compensate for deviations from temperature, from nominal plate’s dimensions

10.5.3 DESCRIPTION

The subsystem development study led to a design of the DPAPS based on the use of three plates of different materials. It has been demonstrated that this number of 3 plates is a minimum to achieve the requested performances. Among the IR materials, three of them, namely Ge, KRS5 and ZnSe, have been chosen regarding their properties and in particular their theoretical refractive index dispersion. Beside, from the model of DPAPS developed during the Phase 1 of the activity, it has been demonstrated that the whole spectral band (6µm to 18µm) has to be split in two sub-bands (6µm to 11µm and 11µm to 18µm) to reach the target performance. In that case, geometrical parameters of every dielectric plates have been optimised. These ones are shown in the Table 9 below.



Material Ge (35 K) KRS-5 (35 K) ZnSe (35 K) Thick. (mm) 6.000 6.159 6.000 4.621 6.000 6.359 Height (mm) 30 30 30 Length (mm) 60 60 60 Wedge (rad) -8.80 10-3 2.41 10-2 -6.50 10-3



Material Ge (35K) KRS-5 (35K) ZnSe (35K) Thick. (mm) 6.000 5.772 6.000 5.653 6.000 6.057 Height (mm) 30 30 30 Length (mm) 60 60 60 Wedge (rad) -8.80 10-3 2.41 10-2 -6.50 10-3

10.5.4 PERFORMANCE

Figure igure 29 and F 30 describe the expected final nulling performance obtained from the design parameters of Table 31 and Table 32. The Figures demonstrate that the nulling remains below 10-6 with a strong margin throughout the whole 6-18µm spectral band.




Table 60 shows a comparison between requested and measured (or expected) performances on design parameters of DPAPS to match targeted specification. A complete sensitivity analysis appears in document NULLTIMATE CDR 2004, October 15th Technical Notes.


1.00E-10

1.00E-09

1.00E-08

1.00E-07

1.00E-06

1.00E-05

6 6.5 7 7.5 8 8.5 9 9.5 10 10.5 1


Nul

ling

Rat

io

1



1.00E-10

1.00E-09

1.00E-08

1.00E-07

1.00E-06

1.00E-05

11 12 13 14 15 16 17 18


Nul

ling

Rat

io





Table 60 Comparison of required and measured (or expected) performance

Performance

Required

Measured/ expected

Remarks

Wedge accuracy 7.10-4 rad 1.10-4 Expected value from supplier comment

Plate thickness accuracy

50µm 50µm

Plate Translation range +/-20mm +/-20mm

Resolution 5µm 5µm


Three dispersive plates have to be moved linearly over a 40 mm range, in one arm of our interferometer, while the 3 others remain static. In the test bench, movements are operated in our bench by stepper motors actuating a linear table through ball bearing screws. When operating point is found, motors are stopped.

Table 61 Interface requirements for the APS


Value

Remarks

Volume (mm×mm×mm) 190 x 480 x 80 Motorized APS Mass (kg) 1.5 Motorized APS Volume (mm×mm×mm) 120 x 380 x 80 Static APS Mass (kg) 0.5 Static APS Accommodation requirements

Power

Average consumption (W) 1 Power is 5 W to find operating point, then 0 W. Cyclic rate is estimated to 0.2

Peak consumption (W) 5

Thermal interfaces

Operational temperature (K) 100 100 K in test breadboard Operational temp. range (K) 80 - 300 in test breadboard

Optical interfaces

plane wavefront λ/10 pv

Other

TBD, zero during Darwin data acquisition

Wave front quality





10.6. Single Mode Waveguide Coupling Unit


Two parallel technology development activities have been recently completed, led by Astrium Gmbh (D) and TPD/TNO (NL). These activities aimed to develop single-mode fibers to be used as a wavefront filters for the wavelength region relevant for Darwin, i.e., 4-20 micron. A major challenge associated with this activity was to find a material that can be used for the wavefront filtering device. This difficulty arises from the fact that silica is no longer transparent for wavelengths beyond 4 microns. The activities were structured as follows:

2) Assessment of suitable candidate materials

5) Experimental assessment of the manufactured samples

A functional demonstrator BB for this component will be built and tested in the framework of the recently started activity “Advanced Fiber Coupling Technology”

10.7. Single Mode Waveguide


1) Study, research, and trade-off of possible fiber design solutions

3) Development of suitable manufacturing techniques 4) Production of test samples

Chalcogenide glasses are good candidate materials for the wavelength range up to about 10 µm, comprising such elements as Se, S or Te, associated with Ge, Si, As or Sb.

The industrial team led by TPD/TNO addressed the development of fibers based on chalcogenide glass Te-As-Se (TAS), manufactured by a novel preform technique, developed at the University of Rennes. Currently (May 2005) the activity is almost closed. Single mode behavior of various TAS fiber samples have been demonstrated at 10 µm and characterisation methods, both at 4.8 µm and 10.6 µm, have been verified. As the TAS fibers can only be used up to 10 – 12 µm a separate ESA activity has been started to study glasses based on Tellurium having a transmission up to above 20 µm.

The industrial team lead by Astrium Gmbh has investigated in detail chalcogenide GAST fibers (GeAsSeTe) for the near IR and poly-crystalline silver-halide materials for the mid-IR. Several fibers, made of chalcogenide glass or silver halide crystal, have been successfully manufactured and the tests performed demonstrated their single mode characteristics.




10.7.2 SINGLE MODE FIBERS FOR DARWIN (TPD/TNO)

10.7.2.1 Description

2

From the results of the initial part of the activity, the study was finally focussed on the lower wavelength range TAS-fibers, manufactured by the university of Rennes applying their patented preform drawing technique. Table 62 summarizes the targeted design parameters for these single mode TAS fibers.

Table 6 Targeted parameters for the single mode fiber based on TAS glass.

A fiber designed for single-mode operation at 4 microns is necessarily single-mode at all longer wavelengths. Therefore, in theory the entire wavelength for DARWIN from 4 to 20 microns range could be covered by a single fiber. However, this is not feasible in practice because the coupling efficiency drops dramatically if a single fixed collimating optics is used and also the guiding of the fundamental mode becomes weaker and weaker for increasing wavelength. A broadband single-mode fiber therefore needs some special coupling optics and a very large cladding.

It turned out that an absorbing coating is mandatory to obtain single-mode operation within a reasonable length (approx. 23 cm). Such an absorbing coating is of particular importance for mid-infrared fibers as because the high refractive index’s cause strong back-reflections from the cladding/air interface.

Two types of preform manufacturing techniques have been applied. The first one consists in the insertion of the core material inside the cladding structure using a special ampoule and a complex procedure that was patented by the University of Rennes. This method has the advantage of ensuring a low level of contamination in the core/cladding interface. However, this method proved to be very complex and difficult to perform due to large demands on processing sequence and thermal control. It also involved operations that were time consuming. No reproducible fibers samples could be accomplished within the limited timeframe of this activity. Moreover, the only fiber attained using this method resulted in an elliptical intensity distribution.




Hence a second method was developed in the course of the study, starting with a one-step fiber-in-tube technique and finally resulting in a three-step rod-in-tube method. This latter technique consists on drawing a large cladding tube against a slightly smaller rod of core material. This operation is performed under vacuum to prevent inclusions of air or other contaminants at the core-clad interface. By manufacturing a fiber with a high refractive index contrast between core and cladding it is demonstrated that the rod-in-tube method provides samples with good symmetrical core geometry. Several fiber samples have been manufactured and characterised.

Figure 50 Some of the manufactured fiber samples

Visual inspection with conventional microscope and SEM is performed to demonstrate the absence of cracks or (air) inclusions at the cross section of the TAS fiber samples. It has been demonstrated that both the existing preform method and the newly developed rod-in-tube technique give a defect free interface between core and cladding. The geometry of the core could not be visualized by microscopic inspection nor by SEM-EDS due to the small difference in refractive index between core and cladding material. Therefore, a fiber with a considerably higher contrast between core and cladding has been produced to demonstrate that the new rod-in-tube method results in a well-defined core geometry.

Characterization of the various fiber samples manufactured within this activity has been performed at 2 different wavelengths:

10.7.2.2 Performance

It has been demonstrated on bulk TAS material (not on fiber samples) that the TAS-glass does not deteriorate (mechanical damage nor change in optical transmission) after cooling down to liquid nitrogen temperature (77 K).

Also the susceptibility to radiation has been examined on bulk TAS material, demonstrating that TAS-glass is not susceptible to radiation levels <10 kRads.

• 4.8 µm tuneable lead-salt laser (Far Field Intensity Distribution, Intermodal Interference Measurement)

• 10.6 µm CO2-laser (Far Field Intensity Distribution, Attenuation)




Tests performed on 2 different TAS samples to estimate the fiber attenuation by means of the cut-back method showed that the intrinsic attenuation of an absorption coated TAS fiber at a wavelength of 10.6 µm varies between 0,1 dB/cm and 0,18 dB/cm. For the designed operational wavelength range from 4 – 9 µm the attenuation is expected to be even lower. This value does not take into account the coupling and Fresnel losses. Characterization results showed that the far field radiation distribution of some of the absorption coated TAS fiber samples have Gaussian beam profiles. However, to achieve this, the length of the fiber samples tested ranges between 20-36 cm, which is considerably larger then the theoretically predicted value of 5 – 10 cm.

1

A number of additional tests are carried out to gain more understanding of the performance and characteristics of the TAS fiber. The effect of misalignment is demonstrated using the far-field intensity set-up at 10,6 µm. It is demonstrated that misalignment of the fiber with respect to the optical beam couples light into higher order modes and cladding modes.

Figure 5 Far field radiation distribution showing the near Gaussian profile of the manufactured and absorption coated single mode TAS fiber (at 10 µm).

It is also demonstrated that bending of the fiber causes non-uniform mechanical stress in the fiber resulting in a change of the mode profile and the far-field intensity distribution.




Table 63 Comparison between required and achieved performance of TAS single mode fiber

Parameter Required Value

(SOW) Achieved Results

4 – 20 µm


10-6 Gaussian FFI distribution demonstrated at 10,6 µm; filter level unknown

< 1.5 dB (about 30%) < 0,18 dB/cm (with absorbtion coating)

Operational temperature

40 Kelvin No mechanical or transmission degradation down to 77 K

Not susceptible to radiation levels <104 Rads

No transmission degradation after exposure to < 10 kRads

Fiber manufacturing process reproducibility

Should enable 5% or better parameter realisation


4 - > 10,6 µm

Total insertion loss goal

Fresnel losses not included!

Radiation susceptibility

Difficult to control refractive index contrast between core and cladding




10.7.2.3 Interface and Physical Resource Requirements Table 64 Interface requirements for the TAS single mode fiber

Interface requirement Value

Remarks

Volume (mm×mm×mm) 300x7x7 total length of fiber needs to be determined for each spectral interval. Bending is not desirable.


no bending Chalcogenide glass is very brittle so it needs to be handle carefully and fiber should not be bended. Bending has also an influence on the geometry of the Far Field Intensity distribution

Power

none Peak consumption (W) none

Thermal interfaces

Operational temperature (K) any no gradients desired during operation







Optical interfaces

It is difficult to realize a broadband wavefront filter covering the entire 4-20 micron wavelength range required for DARWIN. A broadband wavefront filter must be always designed for the shortest wavelength in question. The mode field diameter varies with wavelength and so guiding becomes weaker and weaker for longer wavelengths. In addition the coupling efficiency has to be optimized for each wavelength by using separate optics otherwise the coupling efficiency drops dramatically. Hence, it is recommended that the final Darwin range must be split into several (2-4) sub-bands, and for each sub-band an optimized fiber can be used.

10.7.2.4 Open Points and Critical Issues




The following critical points have been identified during this study:

3- An immersive coating on the cladding surface is required to effectively suppress propagation of higher order modes in the fiber. Current method for applying this coating is useful for laboratory measurement but doesn’t ensure a homogeneous coating. It is preferable to investigate a more liable, reproducible and stable method. Current coating material applied provides significant cladding mode suppression. However, a coating material with stronger absorption is desired to reduce the required length of the fiber.

4- AR coatings need to be developed to avoid Fresnel losses. 5- The Darwin spectral range needs to be divided in several bands (2-4) depending on the

allowable transmission for the fiber. Coupling optics must be also optimized for each sub-band.

6- Fibers need to be tested at more wavelengths to demonstrate their complete performance. With the current equipment available at TNO, characterization can only be performed at 4.8 µm and 10.6 µm.

9- One of the possible causes of the elliptical far field intensity distribution is the mechanical stress in the fiber. Annealing is a well-known method to remove mechanical stress in optical components. For using annealing to remove residual mechanical stress in the TAS fiber, the optimal annealing temperature and time has to be investigated.

1- TAS fibers can be used up to 10 to 12 µm (depending of fiber length). To cover the upper spectral band TNO and the University of Rennes are studying and developing another glass based on Tellurium that has high transmission up to above 20 µm. This is being performed in an ongoing ESA activity.

2- The fiber length must be larger than the theoretical predictions of a few cm. Instead the fiber shall require 20-30 cm in length to guarantee the removal of the high order modes.

7- The cut-off wavelength could not be measured during the study using the set-up developed by TNO due to the size of the TGS detector and the divergence and the size of the optical beam in the FTIR spectrophotometer. As far as known, the proposed basic design provides the only possibility to determine the cut-off wavelength of an IR fiber. To ensure the ultimate IR fiber is single mode for the design operational wavelength range, it is therefore recommended to develop the proposed test set-up to an operational system. For this to happen, it is necessary to find a suitable light source and detector.

8- The intermodal interference measurement method applied has proved to be a highly sensitive method to reveal the presence of the first higher order mode LP11 in the output of the fiber. However, presence of other modes and cladding modes can disturb the measurement results because the phase modulation can be caused by interference between any modes. So, this method is most suitable for the testing of a fiber of which the higher order modes are suppressed significantly and is regarded to be single mode by other test methods as far field intensity distribution




10.7.3 SINGLE MODE FIBERS FOR DARWIN (ASTRIUM)

10.7.3.1 Description

Table 65 targeted design parameters for parameters and resulting single-mode cutoff wavelength λc for mid-infrared fibers.

2) capillary drop and capillary suction method, and

A fiber designed for single-mode at 4 microns is necessarily single-mode at all longer wavelengths. Therefore, in theory the entire wavelength for DARWIN from 4 to 20 microns range could be covered by a single fiber. However, this is not feasible in practice because the coupling efficiency drops dramatically if a single fixed collimating optics is used and also the guiding of the fundamental mode becomes weaker and weaker for increasing wavelength. During the project it turned out that a lower limit is given by the composition difference between core and cladding of a silver-halide fiber. The lower limit is as high as about 2% and it is obtained for the eutectic composition with 25% AgCl and 75% AgBr. It is a result of the physical structure of the silver halide crystal. A further reduction of the composition difference would cause merging of the core/cladding boundaries and no step index could be observed. Fibers with a large variety of composition differences and core diameters have been manufactured to experiment with the two key parameters and to find an optimum solution. Finally, composition differences in the order of 15-20% turned out to be optimum. Large composition differences therefore were traded against small core diameters (down to 10 µm) to obtain high quality fibers. Further it turned out that an absorbing coating is mandatory to obtain single-mode operation within a reasonable length (between 20-50 cm). Such an absorbing coating is of particular importance for mid-infrared fibers as the high refractive indices cause strong back-reflections from the cladding/air interface.

To manufacture step-index fibers made of poly-crystalline material, first preforms are produced. The preforms are extruded in multiple steps to reduce the core diameter. With a final extrusion step the single-mode fiber with the required small core diameter is realized. During this project, the preforms were produced by three different methods:

1) mechanical combining of core and cladding preform,




3) preform growth from the melt.

More than 30 fiber samples have been manufactured from all three methods. Method 1, based on mechanical deformation, was found to produce many scattering defects. Nevertheless, some of the experimental samples have been made according to the required design with small core diameter and low refractive index difference between core and cladding. However, the produced single-mode fibers were of poor quality. The samples produced by method 2, based on capillary suction, had by far the best quality. Several samples had the targeted geometry and numerical aperture. After applying an immersion layer to the fiber cladding a single spatial mode could be observed at the fiber output. The method of preform growth from the melt was found too complex and the process controlling too difficult to produce fiber samples with correct geometry and composition.

Several glassy fibers (GeAsSeTe) have been manufactured to demonstrate single-mode behavior at least in a restricted wavelength range up to about 12 micron. The big advantage of glassy fibers is that they can be drawn, thus yielding small core diameters and perfectly shaped cores. Several samples have been manufactured and characterized. Measurements with and without primary coating have been conducted. The fibers with core diameters of 17, 23, and 25 micron yielded single-mode operation after applying a proper immersion coating onto the cladding surface. This cladding mode stripper is mandatory for single mode operation.

2

Figure 5 Some of the manufactured fiber samples




10.7.3.2 Performance

Table 66 Comparison of required and measured (or expected) performance for the Astrium SMW


This activity demonstrated that its possible to developed single mode fibers based on crystal materials that have, at least in bulk material, high transmission properties up to the highest wavelength value of the DARWIN spectral range (20 µm). In addition chalcogenide glass fibers designed to be single mode and with high transmission up to 12 µm, were also successfully manufactured and demonstrated. The test performed in this activity, were only performed at 10 µm (with the exception of some far field radiation profiles tests at 5 µm). Although the bulky crystal material is transparent up to more than 20 µm, the absorption coefficient of the fibers above 10 µm is unknown. But in theory the scattering centres, that




fibers made of crystal material have and that are associated with the low quality of single mode behaviour and higher intrinsic absorptions, are more critical at lower wavelengths regions.

Total insertion loss

Operational temperature The silver halide fibers were placed at 4 K and neither damage nor performance degradation were reported. No tests were performed to the glass fibers because it was not foreseen for this activity. Future tests shall be required to validate and qualify the fibers for different temperatures and wavelengths. In theory only the glass fibers might have some problems due to their coefficient of expansion dependence with temperature.

Radiation susceptibility Silver halide material is very susceptible to radiation but probably this is only a superficial effect. At the time of the radiation tests the impact of the immersion coating on the single mode behaviour of the fibers was not know, so it was not possible to determine the exact reason for the decrease of transmission after radiation. It is possible that this is just absorption of higher order modes. The tested fiber was multi-mode, since that at the time of the test no single mode fiber was available. In future activities it shall be necessary to make the radiation test to single mode fibers to determine the exact origin of the absorption (or not). In addition annealing effects need to be studied. Another point that needs to be better determined is the level of radiation that the fibers will be subjected at instrument level. Special jackets can also be used to protect the fibers.


To demonstrate the single mode behaviour, two major tests were conducted. The first test is to observe the far field radiation distribution and the second is to implement the fiber in a nulling interferometer to see the filter action of the device. The first test demonstrated to be very useful to detect the presence of higher order modes and clearly distinguish between good (Gaussian profile) from bad (no Gaussian profile) samples. The second allowed to demonstrate the higher order suppression action of the fibers, but it was found out that the maximum suppression measurement was limited by the quality of the entire set-up and not the fiber itself. The test was only performed at 10 µm.

The intrinsic absorption of the fibers was found to be low. However it has been concluded that future fibers shall need AR coatings to avoid high Fresnel losses in its facets. The broadband operation require that special attention shall be given to the coupling optics and the need to divide the overall spectral range into 2-4 bands. For each band the fiber parameters need to be optimized to maximize the coupling efficiency. Again these tests were only performed at 10 µm. Another important point that can influence the total insertion loss is the fiber length. From theory it was assumed that the fiber only required a few centimetres to reject the higher order modes. During the tests at different fiber lengths it was demonstrated that this was not correct. At least at 10 µm, the fiber needs several cm (approx. 20) to reject efficiently the high order modes. Further fiber developments shall require the determination of the minimum fiber length to the entire wavelength region.




Manufacturing process reproducibility No major problem with this requirement since many samples were manufactured.

10.7.3.3 Interface and Physical Resource Requirements Table 67 Interface requirements for the Astrium SMW development


Value

Remarks

Volume (mm×mm×mm)

none

300x7x7 total length of fiber needs to be determined for each spectral interval. Bending is not desirable.



Dependent of material Silver halide material can not get in contact with metals so other holders have to be considered.

Power

Average consumption (W) Peak consumption (W) none

Thermal interfaces

Operational temperature (K) any no gradients desired during operation


Optical interfaces




Other


10.7.3.4 Open Points and Critical Issues It is difficult to realize a broadband wavefront filter covering the entire 4-20 micron wavelength range required for DARWIN. A broadband wavefront filter must be always designed for the shortest wavelength in question. The mode field diameter varies with wavelength and so guiding becomes weaker and weaker for longer wavelengths. In addition the coupling efficiency has to be optimized for each wavelength by using separate optics otherwise the




coupling efficiency drops dramatically. Hence, it is recommended that the final Darwin range must be split into several sub-bands(2-4), and for each sub-band an optimized fiber can be used. The lower end of the required wavelength range can be covered by glassy fibers transparent up to around 10-12 micron. No glass is available yet for the longer wavelength and therefore poly-crystalline silver-halide fiber is at the moment the only choice. Glassy fibers can be easily drawn to very thin core diameters and the core shape is reasonably circular. Immersive coatings exist to strip off all cladding modes and compatible materials have been found. The upper wavelength range from 10 to 20 micron can be covered by poly-crystalline silver-halide fibers. One material chemically compatible with the halide and providing high damping of the cladding modes has been found and successfully applied. As a result of this study, the following improvements on crystal quality, manufacturing technology, and coating technology have been identified:

1. Immersion coatings are mandatory to absorb cladding modes. The materials are very limited and their attachment to the fiber is critical as well. Further investigations on immersion coatings will be required.

2. Improvement of the material quality concerning graining and impurities. The use of more sensitive instrumentation would help to identify the nature and the type of certain scattering sources.

3. Improvement of core/cladding boundary. Disruptions and micro-cracks are formed by the multiple extrusions. The number of preform extrusions must be reduced and the extrusion speed must be controlled more precisely to reduce these cracks.

4. Application of radiation shielding to better resist to hard radiation. The influence of a possible annealing over time has to be investigated.

Further improvements to the test interferometer are required. Furthermore, some measurement equipment would be required to speed up the manufacturing process and to get reasonable feed-back in an early state of fiber production:

1. The established test interferometer must be extended to several other wavelengths by replacing the key components and the laser sources.

2. Sensitive FTIR device to identify type and nature of scatterers and to measure material attenuation over wavelength.

3. Use of a 2D-camera to investigate the obtained mode-structure in an early state and to judge mode suppressing measures like bending, different coatings, etc. in real time.

In order to finally achieve a high-quality modal wavefront filter for the DARWIN mission, the following activities should be performed:

1. Full characterization of manufactured fibers in terms of spectral attenuation, single mode behavior over wavelength, and radiation hardness.

2. Improve the test interferometer to precisely measure high-order mode suppression of the manufactured fibers.

4. Research and improvement of immersive absorbing coatings for silver-halide fibers.

3. Improvement of boundary and material quality of silver-halide fibers, focusing on the promising preform manufacturing methods of mechanical combining and capillary suction.




5. Development of AR coatings. 6. Photonic crystal fibers made of poly-crystalline silver-halides are seen as potential

alternative technology to realize broad-band single-mode waveguides. Good single-mode guiding over an extended wavelength range is expected.

10.8. Integrated Optics for Darwin


Integrated optics (IO) beam combiners and modal filters have been successfully operated on ground-based stellar interferometers in recent years. As compared to a realisation of the same functionality by bulk optics, an IO device provides large benefits by its low mass and its low sensitivity to temperature and other environmental factors. While IO components are not currently baselined for Darwin, the potential of this technique is being explored for functions such as wavefront filtering, beam recombination and beam splitting. The activity is led by IMEP (F) and is split in two technical phases. The first phase deals with the identification and development of suitable transmissive materials as well as the technology to actually manufacture single mode waveguides. The second phase is dedicated to the implementation of different functions such as beam splitting or recombination and testing of the manufactured devices. Phase 1 has been completed. After technological developments to manufacture waveguides the contractor identified hollow metallic waveguides as the most mature technology. Note that the technology development of integrated optics is time consuming and probably not all goals of the activity can be achieved within the current contract.

10.8.2 DESCRIPTION

The components to be developed will provide wavefront (modal) filtering. The final component will also contain splitters allowing for beam recombination (beam interference) or splitting (e.g. a control signal). The requested integrated optics components will provide two inputs for two π-phaseshifted signals and three outputs: 1 interferometric (recombined signals), 2 control signals split from each incoming signal (see sketch below).

Figure 53 The requested topology for the integrated optics components for DARWIN




The components will cover the total DARWIN wavelength range, but each component will only cover about 1/4 of the whole range in order to limit losses. Since the detailed design of the components has not yet started the actual layout may vary from the one shown above. In the first phase a number of substrate/superstrate combinations are investigated: ZnS/ZnSe, which are characterized by well-matched indices of refraction. The development was successful and single mode planar waveguides have been manufactured. However, they will not be taken into consideration for phase two of the development since their transmission cannot be extended above 10µm due to the fixed material combination. This represents only a fraction of the full range to be covered and does not justify further development effort. Hollow metallic waveguides based on the same principle as microwave hollow guides were successfully manufactured. Out of two, one technology based on photo-lithographic techniques was implemented and able to realise hollow waveguides with good quality. A minor problem still represents the dicing which is the process that cuts the individual integrated optics chips out of the silicon wafer. This process partially damages the coupling facets and effectively results in a low manufacturing gain. Still, single mode operation of channel waveguides was successfully demonstrated. The components can provide transmission over the full DARWIN range since the guiding material is air/vacuum, but transmission is physically limited. Wavefront filtering is very effective and requires only short propagation lengths. Consequently, small components can keep the transmission losses low. Chalcogenide (As2Se3/As2S3) waveguides generated by photo-exposition has not proven successful. Energy transfer to the guiding material by light illumination at an absorption band may introduce permanent changes of the index of refraction. This effect can be used to generate waveguides. However, the underlying principles are not well understood. Attempts to generate index changes with various sets of parameters failed and the method was discarded. Chalcogenide (As2Se3/As2S3) ridge waveguides generated by CVD deposition of a guiding layer and subsequent etching was successful. So far only multimode waveguides (by design) have been demonstrated. Since the material parameters have been identified the realisation of single mode waveguides is not considered to be a major technological step. They could provide a transmission up to 16µm. In an attempt to have a second single technology available that covers the full Darwin spectral range Tellurium based chalcogenide glasses (Te2As3Se5/As2S3 and TeAs4Se5/As2S3) form the basis of waveguide developments that are still ongoing. They appear to be promising and would transmit up to 20µm on a suitable substrate. Since they are dielectric waveguides they would benefit from the design heritage of conventional telecom waveguides operating in the NIR. The present selection for further developments in phase two of the contract are hollow metallic waveguides. However, since the developments of the chalcogenide glass waveguides are

promising the contractor is currently asked to propose an updated workplan that covers the development of hollow metallic waveguides and chalcogenide glass based waveguides in phase 2.




Figure 54: Wafer with Hollow Metallic Waveguide Chips

Figure 55: Individual HMWG chip (5x1x1mm3)




10.8.3 PERFORMANCE

Table 68 Performance requirements for the hollow metallic waveguides

Performance (hollow metallic waveguides)

Requirement

Measured/Expected

Remarks

spectral range 4...20µm na / by design modal behaviour single mode proven / by design transmission 25% ~2...5% / <25% 1mm straight waveguide

at 10.6 µm, including coupling losses

The spectral range that each individual component will cover is only a part of the Darwin spectral range (e.g. 4...6µm, 6...10µm, 10...18µm, 18...30µm). The modal behaviour is truly single mode, i.e. only one polarization will be transmitted by the waveguide. In order to avoid losses a polarization components need to be separated and individually brought to interference. Transmission is currently much lower than the theoretical limit since coupling losses are included. Since hollow metallic waveguides are based on “strong” guiding (E-field is zero at the boundaries) a large NA of the coupling optics is required for efficient coupling. .


The interface requirements specified in the table below refer to the hollow metallic waveguide based components. Currently, only straight waveguides have been tested. The actual design of the components might change and, hence, the interfaces are not yet well defined: Table 69 Interface requirements for the hollow waveguide Integrated Optics


Value

Remarks

Volume (mm×mm×mm) 5x1x1 size may vary depending on the number of functions implemented

Mass (kg) <10g Accommodation requirements tbd coupling will demand high

positional stability (<1µm)




Power

Average consumption (W) 0 Peak consumption (W) 0

Thermal interfaces

Operational temperature (K) tbd thermal testing not foreseen in Operational temp. range (K) tbd this activity;

Optical interfaces

Numerical aperture tbd detailed design of device not yet done

Other

Exported vibration level (µN) 0


The basic technology development for waveguides in the thermal infrared turned out to be more difficult than anticipated. There is only little inheritance of existing technology. The activity is delayed due to sample characterization problems related to the specific spectral range. Although the manufacturing technology of hollow metallic waveguides is demonstrated, the use of this technology might be limited since:

• The theoretical attenuation of hollow metallic waveguides is relatively high, resulting in tiny components (waveguides of 1mm length or shorter).

• Hollow metallic waveguides are truly single mode, i.e. only one polarization component is transmitted. This requires polarization splitters and leads to increased complexity of the optical layout.

• Hollow metallic waveguides are based on strong guiding that imposes a high numerical aperture for coupling and consequently a very fast optics (though this could possibly be improved by tapers).

These problems could be avoided with dielectric waveguides based e.g. on chalcogenides (weak guiding). They would allow more flexibility for coupling, have the potential of a higher transmission and their design could be based on the heritage from telecom waveguides in the NIR. But,

• it requires a capable and experienced glass manufacturer (usable chalcogenides are not commercially available),

• it requires development effort for basic technologies involved (layer deposition, structuring),

It is, therefore, envisaged to redirect phase 2 of the activity towards development of both MHWG and dielectric chalcogenide waveguides with less emphasis on the implementation of specific functions such as splitters but instead concentrating on the realisation of single mode straight waveguides with wavefront filtering capability.




10.9. FIR Linear Detector Array


The current baseline technology for the Darwin science detectors is the Blocked Impurity Band (BIB) / Impurity Band Conductor (IBC) Si:As extrinsic semiconductor detector technology. However, these devices have an operational temperature of 8 K. In view of the fact that DARWIN will be passively cooled down to 40 K, any further cooling will have to rely on additional cryo-cooling equipment, thereby increasing system design complexity caused by the cryo-cooler vibrations and weight. The activity “FIR Linear Detector Array” aims to develop a novel detector technology to realize the linear focal plane array (FPA) for the DARWIN mission spectrometer which can meet the performance requirements ad a higher operating temperature. A consortium consisting of ACREO (S), for the detectors, and IMEC (B), for the readout, was selected to perform this task. The consortium proposed a single technology to cover the whole band: Quantum Well Infrared Photoconductors (QWIP). These detectors have the advantage that their energy gap can be tailored and their operating temperature for the required performance has been estimated to be above 16 K. The activity began in November 2002 and is scheduled to end in January 2006. It includes two phases. Firstly, the feasibility phase, which would determine whether the selected technology could achieve the required performance requirements at higher operating temperatures. Secondly, the realization phase, which would proceed with the design and fabrication of the demonstrator. The activity is currently in its second phase. In addition, the 2-year technology development activity "BIB Detector and Readout Electronics" was kicked off in November 2004 and is expected to end in October 2006. The objectives of this activity are twofold:

1. To achieve the QE and dark current performance of the BIB Si:As detector performance required for DARWIN over the 4-18 micron band. This objective has been reached.

2. Once the required sensitivity has been achieved, the technology will be optimised so as to maximise the operating temperature. Ideally, the operating temperature should be increased to at least 13 K in order to reduce the complexity of the cooling system.

10.9.2 DESCRIPTION

The linear array must cover the wavelength band going at least from 6 to 18 µm with state-of-the-art sensitivity and take into account the mission thermal budget. the sought novel technology should achieve the same performance at higher operating temperatures. As a matter fact, the requirements may not be satisfied by a single technology for the whole spectral band. Thus, the Statement of Work (SOW) allowed for separate technologies to be developed, provided the resulting detector modules could be butted on two sides in order to realize a linear array without spectral gap. Therefore, the nature of this activity is to cover basic technology




development resulting in a demonstrator covering relevant portions of the required spectral band.


The required functionality of the FIR Linear Detector Array is completely described by the technical requirements provided in the SOW, which were derived from the DARWIN system study. The technical requirements are reproduced in Table 70. Table 70 Technical requirements for the FIR technical array



Detector Dimensions Array of 1 × N pixels NA

Pixel Dimensions 30µm × 30µm (TBC) NA

Number of Pixels N is dictated by the spectral resolution (TBD) NA

Quantum Efficiency > 50% over entire spectral range 80%

Read Noise < 10 e- 3 e-

Dark Current < 25 e- sec-1 pixel-1 10 e- sec-1 pixel-1

Spectral Response 6 – 18 µm 4 – 22 µm

Integration Time 1 – 10 sec NA

Operating Temperature > 13 – 15 K 40 K

Pixel Cross-Talk < 10-6 NA

Power Consumption 10 mW (at cold part of detector) 5 mW

Note that in order to cover the entire spectral range of 6 – 20 µm, several detector modules optimised for smaller spectral intervals must be developed, possibly using different technologies per module. Therefore, in order to demonstrate the required performance over the entire spectral range, a minimum number of modules covering relevant portions of the spectrum must be developed and tested. From its inception, this technology development activity focused on two crucial parameters: high QE and low dark current. The latter is particularly important for DARWIN and was considered critical, even though simulations predicted sufficiently low dark current values at 17

treated in detail here.

K. Other performance parameters such as, for instance, array dimensions, power consumption and radiation tolerance as well as the design and development of the readout electronics were not thought to be of major concern, as was indeed subsequently confirmed, and will not be




During the Preliminary Design Review (PDR), marking the end of the first phase, the preliminary dark current results were disappointing: the measured value was six orders of magnitude larger than predicted (see Figure 56). Ideally, a research package was in order before proceeding to phase II aimed at determining the causes of such a discrepancy and working out suitable counter measures. However, because of funding problems within the Basic Technology Research Programme (TRP) this was not feasible at the time. Instead, it was decided to go ahead with the project as initially defined and fund the needed research package separately, through the GSTP programme. This additional activity is being initiated at the time of writing.

14 16 18 20 22 24 26 28 30 32 34 36 38 40 42 44 46 48 5010-19

10-18

10-17

10-16

10-15

10-14

10-13

10-12

10-11

10-10

10-9

10-8

10-7

10-6

7⋅10-12 A

I d [A]

T [K]

Vb = -0.5 VVb = -0.25 VTheoretical fit

25 e-/s

Figure 56 Experimental dark current as a function of temperature for long wave (17 – 18 µm) sample at two bias voltage values.

The other parameter of interest for the DARWIN mission is Quantum Efficiency. The results obtained for this parameter were better than predicted. They were obtained by measuring the photoresponse of a fully processed sample and multiplying the result by the expected photoconductive gain of the whole readout chain. The Quantum Efficiency, η, is estimated to lie lie in the range 50 – 80%. It should be emphasised here that state-of-the-art results in the literature state similar RQE values for Si:As arrays for practical bias voltage values. Finally, it must be noted that the best dark current and QE results were not obtained for the same samples. Therefore, further work is required to combine those two performance parameters in a single device.




While the activity is now proceeding towards the realization of a demonstrator satisfying these preliminary results, the hope is to optimize the results obtained so far within the additional GSTP activity. The work to be performed here includes both quantum well subband energy level fine-tuning, to maximize QE for a given sample, and dark current reduction by 4 to 5 orders of magnitude for the long wavelength samples (17 – 18 µm). If these results are achieved, it remains to be seen whether a couple of hundreds e-/sec represent a low enough dark current to work with within the DARWIN mission framework.

10.9.4 PERFORMANCE

The performance parameters obtained so far are summarised in Table 71. Table 71 Comparison of required and measured (or expected) performance for the FIR linear array

Performance

Required

Measured/ expected

Remarks

Pixel Dimensions 30µm ×30µm

30µm × 30µm By design

Quantum Efficiency > 50% 50 – 80% RQE = 25 – 41%. Read Noise < 10 e- < 10 e-

feasible Provided DC contribution is reduced by 4 – 5 orders of magnitude

Dark Current 4·10-17 A 10-12 – 10-11 A Reduction to ~ 10-16 A should be possible

Spectral Response 6 – 18 µm 6 – 18 µm 4 – 22 µm extension3 Integration Time NA 590 s Depends on readout electronics

configuration Operating Temperature

> 13 – 15 K 17 K Provided DC is reduced by 4 – 5 orders of magnitude

Pixel Cross-Talk < 10-6 NA Not measured but should not be a problem in view of the MESA geometry

Power Consumption 10 mW Feasible Depending on number of pixels

3 Extension above 18 µm is possible with the current structure but a 22µm structure would result in some 5 K lower operation temperature. In addition, the characterisation of samples with peak wavelength above 18 µm would face severe experimental difficulties again leading to additional cost. Extension of the wavelength range below 6 µm would require the use of indium-enhanced structure and associated separate optimisation.





It is still premature to discuss interfaces for the linear array subsystem since the final design for the linear array is not available. Therefore, the following table is still incomplete. Table 72 Interface requirements for the FIR linear detector array


Value

Remarks


Power


Thermal interfaces

Operational temperature (K) 16 K Detectors operating at 18-22 µm would require some 5 K lower operation temperature.

Operational temp. range (K) NA

Optical interfaces


Other



The critical issue for the QWIP technology is the dark current. Unless the scheduled research package achieves the needed improvements this technology cannot be used for the current DARWIN performance requirements. An alternative technology, Si:As BIB/IBC arrays, already mature in the US, exists but operates below 8 K.




10.10. Sorption cooler


In the frame of TEC-MCT TRP program a 5K Sorption Cooler is in development for Darwin with the University of Twente (NL). This activity passed successfully the PDR, CDR and shall be ready for a TRR in March 2005. The tests results and performance characterization will be available in July 2005.

10.10.2 ACTIVITY OVERVIEW

The cooling principle is based on isenthalpic expansion (change in pressure, at a constant enthalpy), when the gas is passing through a restriction area. The expansion is on helium, allowing to reach 5K. Considering that the pre-cooling temperature available on S/C is 50K, the cryogenic chain was completed using an additional 15 K Sorption cooler whose design is similar with the 5K cooler with the exception of the working gas that has to be hydrogen and the check-valves will be different. In order to insure temperature stability the best working point is where the liquid/vapour coexist; the isotherms being horizontal lines in the P-H system. Small temperature fluctuations or changes in heat load will affect only the liquid quantity produced, not the temperature on the cold end, as far as the system is able to keep a constant pumping pressure (see isotherm curves presented in P-H system). By changing the pumping pressure, the working point is moving from one isotherm to another isotherm producing temperature fluctuations. In Figure 57, Isotherms are indicated in blue colour, while the Cooler cycle is in red colour.

Liquid+ vapour

P(atm)

Liquid+ vapour

P(atm)

Figure 57 Isotherms curves for He, presented in Pressure-Enthalpy system.




The system insures a constant injection pressure or a constant pumping pressure by temperature change of an activated carbon. The activated carbon has the property of absorbing large quantity of gas, when its temperature decreases. Once the carbon is saturated in gas, by increasing the temperature, the gas will be release, creating high pressure in the system. The absorption is a physical process, which doesn’t degrade the carbon properties over time. Special attention will be taken to avoid particle formation by friction between the carbon and inner shell of the compressor cells. Applying a heat load, evaporating only part of liquid, the working point will stay, on the same isotherm, moving towards right. To insure a continuous process, the system needs minimum 4 independent cells, filled with activated carbon. By expanding the working gas, a high percentage of the nominal flow rate will be liquefied. The evaporation heat of the liquid is providing cooling at the interface with the instrument. To avoid having non-controlled liquid in the system, the excess liquid, not used by the Instrument will be evaporated in so-called evaporator. All the cells are connected to the cold end via check-valves as presented in the. Each cell will go successively through 4 distinct phase: see Figure 58-a.as: A: warming up, B: desorption,C: passive cooling, D: active cooling.

Figure 58: a. Single stage Sorption Cooler schematic and the 4 Two-stage sorption cooler schematic.

The compression phase works from 1b to 17b. This can be implemented, see Figure 58 b): where the pressure values these cells is permuted after TBD minutes (about 200secondsTo improve the efficiency of the cooler, few sets of heat exthe compressor cells and the cold end. The complete crysorption coolers) is presented in

The cooling of carbon is reached by heat sinking50K. When the carbon needs heated, the cells are insulated

52 K, 2.92 W2

A B C D

Figure 62.

phases of a single compressor cell .b)

improved if 2 stage compression is are: 1b /4b /17b. The function of ). changers are implemented between o-cooling chain schematics (both

the cells on the passive radiator at from radiator and heated-up. The

shield

pL

pI

49 K, 1.65 WA = 6.5 m2

4.5 K,10 mW

A = 9.7 m

to H2-stage




sinking and the insulation of carbon to the radiator is made via a Gas Gap Heat Switch (GGHS) which insure this function. The helium expansion, to produce liquefaction, is made very close to the detectors interface; the connections are flexible and easy to implement. Part of the liquid produced will be hold in place at the detector interface plate, and part of liquid and the vapours will leave the interface to insure a continuous running. This will insure also an increased thermal stability at the interface with the detectors. A brief description of each part of He Sorption Cooler is presented below. The H2 Sorption Cooler is similar and is not described separately.

10.10.2.1 Compressor Cells

Compressor cells:

Figure 59 Schematics of the compre

Each cell has following characteristi

Figure 59): Outer diameter 16mm; Length 150mm; Mass 350g; Radiator connection with flat surface 16x150mm. Inner/Outer Shell are made in SS 316L, outer-shell has 0.2mEach Sorption Cooler has minimum 8 cells (1 redundancy p Inner part is cycled in temperature: from 50K to 80K- stag80K-120K stage1 and 80K-240K stage 2 for H2.The inner small gap, which can be under vacuum or filled with gas ca

Inner Shell, filled ith carbon; installed

w inside outershell

Gtbs

Outer shell, bottom part fixed on the radiator

GHS- making hermal connection etween Inner/Outer hells

ssor cells

cs (see

m Cu layer on outside er compression stage)

e1 and 52K-120K stage2 for He and part is separated from outer part via a lled Gas Gap Heat Switch.




The GGHS resistance for ON/ OFF position is expected to be: 2K/W and 1000K/W, for He cooler. The gap–with =0.4mm; Contact area 720cm2, and filling gas is H2. The average heat dissipation on radiator at 50K by the He SC is about 4.6W. For the H2 SC is about 1.6W. As result large radiator area shall be implemented in S/C.

10.10.2.2 Installation on Radiator The only constraint imposed by the SC on the radiator is a minimum distance between the compressor cells. The distance shall be such that during thermal cycling of one cell the adjacent cells shall not see any increase in temperature higher than 0.5 K.

Figure 60 He SC cells implementation on 50K radiator, and H2 SC cells implemented on 80K radiator. This must be considered as an implementation example; which can be modified by the prime.

The outer-shell will have a temperature fluctuation over the cycling time, and these fluctuations shall not be source of perturbation for the detectors (which are located in the Beam Combiner) A preliminary accommodation of the 50K Radiator, for 1st and 2nd stage of He and H2 Sorption Cooler was proposed as a fix radiator install between sun shield and the Beam Combiner (see Figure 60).

10.10.2.3 Counter Flow Heat exchangers

A discrete gas heat exchanger will be present at the exit of each compressor cell, such that the warm gas will be cooled at same temperature as the radiator; average He flow:0.8mg/sec. The size is TBD but not larger than 20x10 mm. The pressure drop in low pressure line is very important and is about 4mb.

10.10.2.4 Cold End The cold end is composed from 3 independent parts: the Joule Thomson expansion, the detector interface and the evaporator block (see Figure 61).

Between the compressor cells and the cold end, it is foreseen to implement a CFHX. The tubing will be stainless steel and the external diameter about 3.3mm (inner tube SS 1/16”outer tube SS 1/8”). The actual length foreseen is 2m. There are no specific restrictions in routing of HEX (no attachments on parts warmer than 50K).




Figure 61 Cold End parts: 15K-5K

The JT is a sintered SS porous metal mounted inside a SS tube ( D tube~3mm). The detector interface is a copper plate; in this BB the interface surface is small but can be adapted to the detector needs. The evaporator is thermally decoupled from the detector plate and its function is to evaporate the He excess liquid. This BB contains only a limited number of wires, necessary to the cooler function. If a large number of electrical wires are foreseen for the detectors than integration of the harness with the HEX and the evaporator shall be study and implemented in order to reduce the parasitic heat loads. The actual design of 5K cold end contains a thermal shield at 15K provided by the H2 SC pre-cooling. It is important to be able to implement the 15K thermal shield in Darwin configuration thermal design.

JT restriction

evaporatorDetector interface

Load

heater power heaterpower

4.5 K > 4.5 K

6 - 8 K

heatexchanger

Detector interface

evaporator

CFHEX15K plate




The lowest temperature that can be reached at L2 by passive radiation was estimated to be 40K if only few mW are dissipated. To cover the temperature range 50K-40K down to 5K, two independent Sorption Coolers are need it: H2 Sorption Cooler (provides 15K stage) and a He

re

Figure 62 H2 Sorption Cooler, He Sorption Cooler and radiators interface

In principle all materials used in 5K Sorption Cooler are compatible with hydrogen but the porosity of stainless steel to hydrogen is higher than for helium, resulting in a slight degradation of cooler performances at long run. A 15K Sorption Cooler breadboard has to be built in near future to prove the feasibility of the complete cryogenic chain. The 6 years lifetime performance will be demonstrated based on accelerated lifetime tests of critical components. The micro-vibration level induced by the cooler will be measured in a non-representative configuration, designed in purpose to increase the vibration level such that the system can be characterised and the real micro-vibration performances will be found by analysis.

Sorption Cooler providing 10mW at 5K. A schematic of the two Coolers is presented in Figu62.

Heat exchange on radiators: He Sorption Cooler: 52K; average heat load 2.92W 49K; average heat load 1 65W

14.5 K

pL

pI

pH

49 K

52 KpL

pI

pH

80 K

4.5 K

shield

radiator

H2 Sorption

He Sorption

Compressor cells

Stage

Stage

Detector interfac

Pre-cooling of He

1 3b

4

17b0 1b

50b




10.10.3 PERFORMANCE

In table below are presented the performances for He and partial H2 Sorption Cooler. Table 73 Comparison of required and measured (or expected) performance for the Sorption Cooler

Performance of He Sorption Cooler

Required

Measured/ expected

Remarks

Cold plate temperature 5K 4.5K * 4.5K without temperature control; * 5K with PID temperature control to insure fluctuations lower than 1mK

Cooling capacity 10 mW 9.5 mW 10mW without margins; it is expected 9.5mW

Power consumption 10W Average: 5W; Max: 6W

This power excludes the driving electronics of the cooler 6W input is expected for He cooler and 4W for H2 Cooler

Temperature fluctuation



Reached by active control PID. Temperature fluctuation of H2 stage is less important if still less than 0.1K.

Distance between radiators and cold tip

2m 2m This distance is flexible. A CHEX is planed for 2m length.

Cooler mass for He SC

10Kg for H2 and He SC

8Kg for He SC 3Kg for H2 SC

70% of mass will be install on radiator, cold plate will be less than 0.5Kg and the difference (piping, supports etc). in between.

Supply voltage 28V For BB will be 220V

Mechanical resonance frequency

Higher than 100Hz

Higher than 100Hz

lifetime 6 years 6 years Accelerated lifetime tests for carbon, GGHS, check valves; all other parts presents no

degradation




Nb. of temp-pressure cycles during 5 years is: 275 000

Depressurisation profile



Mechanical strength -qualification 1.5 x operational load - yield strength 1.8 x operation load - ultimate strength 2.0 x operational load

Cold stage orientation System works independent of orientation

System works independent of orientation


1.7W at 49K Average 1.7W The cooler performance improve if passive radiator can reach lower temperature than 49K during absorbing phase. -Stage 1 is heated 49K-80K


3W at 50K In average 3W at 50K

The maximum heat load is reached in the beginning phase of active cooling and depends on GGHS performance. -Stage 2 is heated 50K-120K

Precooling stage needed for He SC



H2 SC dissipation on 80K radiator

3W 3W Stage 1 is heated 80K-120K Stage 2 is heated 80K-240K

Working pressure for He SC



Working pressure for H2 SC



GasGapHeatSwitch- for He SC



Degradation in time will be verified by test H2 is the working gas for GGHS

Check-valves for He SC;

*nominal flow: 0.8mg/s He *leakage flow closed direction 2µg/s *max dP in closed direction

*nominal flow: 0.8mg/s He *leakage flow closed direction 2µg/s *max dP in closed direction

Check-valves are identical for all stages. Operating Temperature 50K




20b *cracking pressure 1-10mb *pressure drop<15mb

20b *cracking pressure 1-10mb *pressure drop<15mb


Table 74 Interface requirements for the sorption cooler


Value

Remarks

Compressor Cells Volume of single cell D=16mm L=150mm Attachment on radiator: 16x150mm

-copper interface Nb of cells: minimum 8 Mass (kg)/ cell 0.5 Accommodation requirements extra volume is required

on top of the cell to accommodating GGHS; -space allocated : D=50 L=200mm -minimum distance between cells shall be planed such that T fluctuations during active cooling are not propagate to the cells in adsorption phase.

The T fluctuation of outershell shall not perturb the detection

Distance between cells Minimum distance between 2 cells can go from 5cm to 5m, depending on radiator accommodation to provide at any time temperature fluctuation on adjacent cells less than 0.5K

Effort shall be put in designing: • low mass radiators, • easy integration of the cells • if possible have late access

Check-valves unit 2 check-valves / cell 1 unit has 2 check-valves Volume of unit 40x20x10mm Is attached to the radiator Total number of check-valves 16 Mass 1 unit (2 check-valves) 0.07Kg/unit Accommodation requirements Require thermal

coupling and attachment to the radiator




Heat exchanger CFHEX Volume Total number Mass Accommodation requirements Buffers Volume 2 litres In aluminium Total number 3 Mass 0.625 Kg each Accommodation requirements Attached on radiator at

50K One buffer at each pressure level: 1b, 4b, 17b

Precooling +thermal shield at 15K

Volume D=100mm; L=20mm Requires a flange to install 15K thermal shield

Total number 1 Mass 0.5 Kg minim Accommodation requirements Requires mechanical

supports and thermal insulation from radiator

15K thermal shield has to cover the 5K stage to reduce the parasitic heat load

Cold End Volume To be considered

D=60mm; L=40mm Actual volume is small but has to be interconnected with the detectors

Total number 1 Mass 0.5Kg Accommodation requirements The actual supports are

in kevlar; it may require redesign.

Power

Average consumption (W) for He Sorption Cooler

5W

Peak consumption (W) for He Sorption Cooler


Average consumption (W) for H2 Sorption Cooler

3W

Peak consumption (W) for H2 Sorption Cooler


Electrical harness 16 wires Goes along HEX

Thermal interfaces

Operational temperature (K) He Sorption Cooler Innershell


Operational temperature (K) H2 Sorption Cooler Innershell


Operational temp. radiator -50K & 5W dissipation




(K) for He / H2 SC -80K & 3W dissipation Beam Combiner Passively 40K Pre-cooling temperature 15K &25mW cooling

Other

Exported vibration level (µN) 1µg/ √Hz By analysis; measurements laser interferometers

Electromagnetic emission none measured


The driving factor of the Sorption Cooler is the temperature; the performance is strongly dependent on radiator temperature and temperature fluctuation. The required radiator area, to meet the cooler performances, is quite large and margins to account for degradation over time shall be taken. This makes overall S/C design heavy and impose constraints for the test facility in terms of size. In meantime the heat loads on detectors are mainly due to harness, supports, electronics etc. The heat dissipation on detectors due to the radiation is negligible (in PACS, 400 pixels, the heat dissipation due to the radiation is in microW region and main heat dissipation is coming from pre-amplifier). In this specific case the cooler integration with the detectors may provide 5K cooling for detectors only and a 15K pre-cooling for the electronics, supports etc, so the cooling power at 5K will be reduced. If the needs in cooling power can be split in 2mW at 5K plus 5mW at 10K, than another cryogenic design may become interesting, which scheme is presented in Figure 63. The cryogenic chain is composed by a Stirling reaching 20K-17K, a Stirling compressor, running He through a JT expansion to reach 4.5K with 15mW cooling, and a He 4 Sorption Pump providing 5K for 2weeks, for 2mW cooling power

SVMSun-Shield

4KV-Grooves40K

Detector interfaceHe Liq.Bath

Carbon pump-40K

JT CompressorSVMSun-Shield

4KV-Grooves40K

Detector interfaceHe Liq.Bath

Carbon pump-40K

He Liq.Bath

Carbon pump-40K

JT CompressorJT Compressor

Stirling Compressor

20K

Stirling Compressor

20K




Figure 63 Alternative cryogenic design providing 2mW at 5K plus 5mW at 10K

The Sorption Pump will be recycled every 2 weeks (TBC). Since no vibration is allowed during observation, recycling will be performed during slew manoeuvres. The advantages of this design are:

- The 4K Cooler already qualified in Planck Mission

- 20K Stirling is already developed; Compressors for both coolers are in the process of redesign to improve the cooling power and performances

- Sorption Pump already qualified for SPIRE- Herschel Mission for lower temperature

0.3K; adjustments and BB need it for 4K. The system is passive during observation and heat rejection required is about 0.1W. No large radiators needed, helps reaching lower temperature and provide an extra cooling stage by using the cold He vapour to cool harness and electronics. This design can also be considered the radiators size and radiator performances for the baseline design will prove problematic at Spacecraft level.




11. APPENDIX C: FORMATION FLYING COMPLEMENT

11.1. FF complement overview The ability to fly and manouver several spacecraft in close formation is essential for Darwin and a number of future science missions, including XEUS. Challenges are to be found in the field of precise and independent measurements, control system architecture and actuation system. For the metrology system both radio frequency ranging and high precision laser metrology are foreseen. For the nulling mission there is always a central star towards, which the interferometer and its individual telescopes are pointing. Nevertheless, the spacecraft positions (and rotation around the line of sight) relative to the combiner spacecraft have to be monitored. The control system will have to collect several measurements on different spacecraft and control the individual spacecraft, forming the interferometer. For the actuation highly precise and proportional micro-propulsion units will be used, most likely field emission electric propulsion (FEEP) technology and ion thrusters. When several spacecraft are flown in close formation there is an apparent risk for spacecraft collisions, which, in the worst case, could put the mission to a premature end. It is therefore of utmost importance to avoid that the spacecraft collide and that the spacecraft formation can be formed from an almost arbitrary distribution in space. The latter is important when the formation is detached from the launch vehicle and deployed, and after contingencies when the control of the formation has been interrupted.

11.1.1 METROLOGY

A number of metrology systems will measure attitudes and relative positions of the telescopes, as needed by the control system to deploy and control the formation. A chain of metrology systems allows the measurement accuracy to be refined both in terms of spacecraft pointing and relative positions. A number of coarse sensors, including coarse sun sensors, star trackers and RF metrology, are utilized in the initial stages ensuring that the attitudes and positions are good enough to hand over to the subsequent laser metrology systems. These will bring the relative attitudes and positions to sufficiently accurate level to start the Fringe Acquisition Mode (FAM). In FAM the differential optical path lengths are controlled to nanometer accuracy and the Optical Delay Lines (ODLs) are actuated such that stellar fringes can be detected on the Fringe Sensor. The distribution of the metrology systems over the formation is illustrated in Table 75. Each satellite will be equipped with coarse and fine attitude sensors. All laser metrology beams will be launched from the hub towards each of the three telescope satellites. The laser source will, as far as possible, be shared between the different metrology systems. The telescope spacecraft host metrology retroreflectors (RR) or metrology sensors (MS), depending on the considered metrology device. Note that the inter-satellite metrology path is separate from the science

optical path, i.e. metrology beams are not overlayed with the science beam. Beam tilt and OPD




metrology are implemented in the Fringe Sensor. This is accommodated in the hub, as well as the associated actuators, i.e. Optical Delay Line and Fine Steering Mirror.

Table 75 Distribution of the metrology systems between the various s/c

Metrology

hub

CS 1

CS 2

CS 3

Pointing metrology Coarse sun sensor X X X X Set of gyros X X X X Standard star tracker X X X X WFC X X X s/c co-alignment laser MS (WFC) MS (WFC) MS (WFC) Fine tilt sensors (3x) X Position metrology RF metrology X X X X Coarse lateral Laser+MS RR RR R Fine lateral laser MS MS MS Absolute longitudinal Laser+MS RR RR RR Fringe Sensor X

11.1.1.1 Pointing metrology

Coarse sun sensor

Coarse sun sensors will be implemented on each spacecraft, allowing the spacecraft to acquire initial sun-pointing after separation from the upper stage. Furthermore the coarse sun-sensors are used in the recovery procedure following a loss of attitude event. The main purpose is to ensure that the solar panels are oriented towards the sun.

CS star tracker

Each telescope spacecraft has, a high performance star tracker (also referred to as the Wide Field Camera, WFC). The WFC is used for target acquisition, and as primary pointing sensor in all operational modes except NOM. This subsystem is further discussed in Section 4.3.

Hub star tracker

The hub will host an autonomous star tracker, equipped with its own telescope. ing measurement requirement is approximately 1 arcseconds (1 σ) for both pitch and yaw, while the attitude around the telescope LOS will be controlled within 100 arcseconds, at a rate of 10Hz. The star tracker is used for target acquisition, and for monitoring variations of the S/C pointing with respect to inertial space.

Spacecraft co-alignment

The fine co-alignment of the LOS of the transmitters and hub is monitored by imaging a laser beacon from the hub onto the WFC at the telescope spacecrafts. The WFC measurement accuracy for these laser stars is 7.5 times better than for natural ones, since the net angular

The point




magnification of the laser beam from the hub to the WFE entrance pupil is unity. Changes in the pointing of the hub will be detected as correlated motions of the laser spots at the three WFCs. It may prove possible to remove the spacecraft co-alignment metrology, as describe d above. In this case, two star trackers are needed on each s/c, one pointing in the LOS direction, and the second one pointing orthogonally to the LOS. This second device measures the s/c attitude around its LOS.

Figure 64 Spacecraft co-alignment is achieved by launching a laser beam (in green in the figure) from the hub towards each CS. This beam appears as an artificial star at the wide field camera.

Fringe sensor

The Fringe Sensor measures the relative phase of the three beams as close as possible to beam recombination. The relative phases of the interferometer beams determine the fine orientation of the interferometer LOS. This device is further discussed in Section 5.5.

11.1.1.2 Position metrology The optical metrology equipment will be located on the BCS and measure the position in three dimensions of the three telescope spacecraft.

RF metrology

The RF metrology system measures the distance between any pair of antennas. By using several separated antennas three-dimensional relative positions can be measured. The RF system shall have an operating range of 8 km with coarse position determination capability. For equilateral and linear configurations, the accuracy in elevation and azimuth is better than 0.4 degrees and 0.7 cm in range at 1 Hz. During the deployment phase or non-nominal situations, relative navigation is computed between pair of spacecrafts, and the accuracy in elevation and azimuth is better than 1.1 degree within a conus of +/- 60 deg (two sides), and 0.7 cm in range, with significant degradation outside of the conus. The performance of the spacecraft position control,




based on RF metrology, is compatible with the acquisition range of the coarse lateral and longitudinal optical metrology. The coverage of the RF metrology subsystem shall be omni-directional in order to enable measurements to be performed immediately after formation deployment, when the spacecrafts are randomly distributed in position and attitude. See Section 11.3 for details on this subsystem.

Coarse lateral metrology

The coarse lateral metrology provides absolute measurements of the position of a S/C in the plane perpendicular to the beacon from the reference satellite, with an accuracy of 1 mm @ 10 Hz. See Section 11.4 for more details.

Fine lateral metrology

The fine lateral metrology monitors absolute displacements of a S/C in the plane perpendicular to the beacon from the reference satellite, to an accuracy of 32 µm rms @ 10 Hz, within a range of ±1 mm. This measurement, together with the fine longitudinal metrology, allows to damp OPD drifts to a level compatible with fringe acquisition. See Section 11.4.2.1 for details on this subsystem.

Fine longitudinal metrology

The distance between two S/C will be monitored by an absolute laser metrology system to an accuracy of 32 µm rms @ 10 Hz. The bulk of this metrology system (two frequency-locked Nd-Yag lasers, send/receive optics, electronics) is hosted on the beam combiner satellite. The metrology beam is sent to the second spacecraft, where it is retro-reflected. The metrology path is independent from the science one, which greatly simplifies beam routing. Analysis has shown that the non-common paths between the science and metrology beams can be kept well within the required measurement accuracy. This subsystem is described in detail in Section 11.4.3.2 and 11.4.3.3.

Technology development activities have been performed addressing the RF metrology (see Section 11.3) and laser metrology (HPOM activity, see Section 11.4).

11.1.2 FF PROPULSION

Spacecraft propulsion is required for two functions: 1) orbit correction, in order to correct for launcher dispersion, and possibly for injection in

small-amplitude orbit at L2. This could be achieved with chemical propulsion system, or with milliNewton thrusters (see next paragraph)

2) formation flying (FF), including both coarse manoeuvres, e.g. slew, and precision formation control

For FF coarse formation manoeuvres, a thrust capability of a few milliNewton and a resolution of ~ 0.1 mN (TBD) would be required. Precision formation flying will make use of µN thrusters with a maximum thrust of ~ 0.1 mN and µN-level resolution, which will allow performing fine attitude/position corrections directly during an observation, without perturbing fringe tracking. Possible technologies for mN and µN




propulsion have been examined and traded-off during a recent ESA internal trade-off [RD 10], as described in the following two subsections.

Milli-Newton propulsion trade-off For the DARWIN coarse manoeuvres, which include reconfiguration and rotations, Cold Gas Microthrusters (CGMT) and Electric Propulsion System were found to be the most suitable propulsion technologies. Following the change of the Darwin launch vehicle from Ariane-5 to Soyuz-Fregat, however, the allowable mass budget has been considerably reduced. Because CGMT thrusters require a considerable amount of fuel due to their low specific impulse, the Electric Propulsion System technology is today considered the baseline for this application. Several electric propulsion systems exist nowadays and have already been operated in space. Two main technologies can be assessed as propulsion system for Darwin coarse manoeuvres: Hall Effect thruster (HET) and Ion Thruster (IT). The difference between the two technologies is in the way they generate and accelerate the ion of the Xenon plasma. Usually ITs have higher specific impulse than HET, which may translate into significant propellant savings, especially in the case of high ∆V requirements. However, due to the high voltage and power required, the weight of the Power Conditioning and Supply Unit needs to be taken into account. Also, IT provides higher controllability of the thrust then HET. HET have lower power to thrust ratio then ITs, which means a reduced demand in term of Power for the same thrust required. They also have a more compact design and less complex PCU. In general the choice between those two technologies is likely dependent on the mission specifics, such as ∆V required, performances, power and mass constrain. Following the most recent DARWIN requirements [RD 4], the following options could be considerate even if the thrusters do not belong in the same class:

• “RIT-10” developed by EADS Space Transportation, Flight Proven in ARTEMIS (see Section 11.5),

• “T5” developed by QinetiQ, Flight Proven in ARTEMIS (see Section 11.6),

• “Radio-frequency with Magnetic-field ion Thruster” (RMT) developed by Alenia

Spazio, Laben Proel, Engineering Model (see Section 11.7),

• “Mini-HET” developed by ALTA, Engineering Model (see Section 11.8). The properties of all the systems are compared in Table 76.




Table 76 IT and HET properties

RIT-10 T5 RMT mini-HALL

Flight Flight EM EM

205x185x165 190x190x242 150x150x130 100x100x100 (1)

1.7 2.95 1.6 <0.6(2)

10% of the Xe mass required




15.1(3) 33.75(4) 9.7(5) 3.5 (6)

<30.6 (7) 30.8 25-40 25

15 18 TBC 4

0.3-40 (8) 1-20 2-12 0.48-91 0.2@1mN level 1 TBC

2500-3700 500-3500 2200-3600 >960

<11.5° <25? at 1mN; <12? at 20mN 10-15? <45

>30000 38000(10) >10000 >3500 (14)

20000 (9) 4500(10) 500 100

1.07*106 (11) 1.5*106 (12) >1*106 (12) >52500(12)

>6714 5500 (13) TBC >3000(12)

Predicted Lifetime [hours]

Total impulse [Ns]

Tested hours

(13) cathode cycles in 15000h; up to 10000 cycles at components level

(1) Thruster envelope with neutralizer (2) inclusive of neutralizer and after optimization; Today 0.71kg

(8) for RIT 10 EVO

(7) 25-27 W/mN demonstrated for the RIT 10 EVO

(6) inclusive of neutralizer, harness and PSCU. The PSCU drives up to 4 Thrusters.

Demonstrated Thrust range [mN]

Thrust resolution [♦N]

Interface requirements and Performance

Thruster Dimensions [mm]

Mass of the Thruster [kg]

Propellant Tank mass estimation [kg]

Power-to-thrust ratio @ nominal thrust [W/mN]

(14) According to the present requirements

(4) inclusive of IPCU and PXFA

Cycles

(3) inclusive of Power Supply and Control Unit (PSCU), RF Generator, Flow Control Unit FCU, Harness, Tubing and Support

Demonstrated Isp range [s]

Total system mass[kg]

Nominal Thrust [mN]

Status

Beam divergence (half-cone)

(12) predicted

(5) inclusive of GFCU, RFGM, PSCU, Harness and piping.

(9) demonstrated during ground test at ESTEC in the EPL, >30000h predicted(10) 4500 demonstrated for GOCE, 38000h predicted for GOCE(11) demonstrated for 20000 hours @ 14.7mN

To be highlighted: 1. In [RD 10] is written that “for coarse manoeuvres DARWIN mission will require a

maximum thrust levels of 1mN or more, in order to be able to reconfigure the spacecraft position within reasonable time frames. The noise and accuracy requirements for coarse manoeuvring are not as stringent as for the precision pointing, since their main purpose is




to reorient and reconfigure the spacecraft as fast as possible, to maximise science observation time” (TBC).

2. Both the RIT10 and the T5 have already reached a high level of maturity and are flights proven being used in ARTEMIS.

3. Both the RIT10 and the T5 have been characterized or are under characterization for different application in which thrust level less then 1mN are not required. Only in the case of RIT10-EVO the thrust range demonstrated covered a range of 0.3-41mN during 2400 hours of test performed. If DARWIN will demand to have lower thrust level a test campaign will be necessary to assess the minimum thrust achievable by each Thruster and the resolution.

4. The mini-Hall belongs to a different class of Thruster compare to the T5 and RIT10 one. It has been characterize for a thrust range between 0.48mN and 9mN. The mini Hall is driven by a Power Control Unit able to drive 4 thrusters at the same time. Depending on DARWIN micro-Newton manoeuvres the m-Hall could also be used as single Multipurpose Propulsion System. Lower level of thrust could be reached.

5. Both ITs and HETs use Xenon as propellant, resulting in benign thrust plumes that minimize contamination issues.

6. If DARWIN coarse manoeuvres will require a maximum thrust of around 1mN, a choice to develop one Multipurpose (µN and mN) Propulsion System for both coarse and micro-Newton manoeuvres could be taken into account for mass and size saving. Milli-Newton FEEP as well as micro-Newton ion Thruster (RIT4 or RMT) could be possible candidate.

11.1.2.1 µN propulsion trade-off For the DARWIN micro-Newton manoeuvres, such as fine pointing and slow OPD control during science operation, two different technologies, Cold Gas Microthrusters (CGMT) and Electric Propulsion Systems were taken into consideration as propulsion system. CGMT option is currently not suitable due to mass constrains. According to RD16 the requirement for DARWIN micro-Newton manoeuvres are:

Thrust range between 1µN to 100µN (up to 150µN); Noise level less than 1.65µN/Hz1/2 for thrust level below 100µN; Resolution less than 3µN; Accuracy less than 0.5µN.

The following options are currently under investigation:

• “FEEP-8” developed by ALTA, Engineering Model (see Section 11.11), • “Indium FEEP Multi-emitter” developed by ARC, Engineering Model (see Section

11.10),

• “RIT-4 micro-Newton ion thruster” developed by GIESSEN University, Prototype (see Section 11.9),

“Radio-frequency with Magnetic-field ion Thruster” (RMT) developed by Alenia

• Spazio, Engineering Model (see Section 11.7).




A summary of characteristics of those different systems is given in Table 77.

Table 77 Micro-Newton Propulsion System properties

FEEP-8 In-FEEP Multiemitter RIT-4 RMT

EM EM Prototype EM

169xφ10 TBD 70xφ83 150x150x130

1.25 TBD 0.6 1.6

9.1(1) TBD TBD 9.7(7)

5400-5900-6400s @150µN with - 4,-3,-2kV;

6900-7400-7850s @30µN with - 4 ,-3,-2 kV

1600-8000 (TBC) 3850@230µN 2200-3600

0.1-250 0.1-40 (6) 20-190 2000-12000

0.1 0.1 TBC 100057-61-64W/mN @150µN

with - 4,-3,-2kV; 53-57-60 W/mN @ 30µN

with - 4,-3, -2kV

120 TBC 25-40

in-plane_15˚; out-of-plane_40˚ 25-60˚ TBC 10-15˚

9000hours @75µN (2)

9000hours @100µN (3)

18000hours @150µN (4)

9000hours @75µN (2)

9000hours @100µN (3)

18000hours @150µN (4)30000 >10000

1600 (5) 2400 (6) TBC 500

2460 (2), 3100 (3), 10000 (4) 2460 (2), 3100 (3), 10000 (4) TBC >1*106

600 240 (6) TBC TBC300 milion, 100 milion (3), 600

milion (4)300 milion, 100 milion (3),

600 milion (4) TBC TBC

(7) inclusive of GFCU, RFGM, PSCU, Harness and piping.

Beam divergence (half-cone)

Demonstrated Thrust range [µN]

(5) Lifetime verification test for Microscope foreseen in 2006

Demonstrated Thrust resolution [µN]

(3) Microscope requirements

Predicted Lifetime [hours]

Predicted Total impulse [Ns]

(2) LISA PF requirements

Demonstarted Total Impulse [Ns]

Total Mass of the system (g)

Number of actuations

Mass of Thruster (kg)

(4) GAIA requirements

(1) According with Microscope, including 3 TA, 1 PPU, 2 NA

Tested hours

Demonstrated Isp range [s]

Power-to-thrust ratio (W/mN)

(6) Test with 4 emitters each of them delivering a max thrust of 10µN

System properties

Thruster dimensions (mm)

Status

To be highlighted:

• FEEP-8 Subsystem is nowadays qualified to satisfy the future ESA mission

requirements. It has been selected for MICROSCOPE and is candidate for LISA PF, GAIA and LISA. When MICROSCOPE will be launched (namely 2008) the system will be also FLIGHT PROVEN.

• ALTA FEEP system could also offer a larger range of thrust that can cover up to 1mN. If DARWIN coarse manoeuvres will require a maximum thrust around 1mN per




thruster, this system could allow having one multipurpose (µN and mN) propulsion system meaning mass and size reduction.

• Indium FEEP Multiemitter consists of ‘cluster’ of emitters. Each of them delivers up to 10µN. To date a study on using one PPU able to drive up to four emitters is being performed. In any case, if higher thrust level is required an increase of mass (more emitters and more PCU) as well as an increase of risk should be taken into account.

• The RIT-4 as well as the RMT µN Ion Thruster could guarantee a much large range of Thrust. If DARWIN coarse manoeuvres will require a maximum thrust of more than 1mN, both thrusters could allow having one multipurpose (µN and mN) propulsion system.

• To date the minimum thrust level reachable by the RIT-4 Thruster is under investigation in GIESSEN University. The qualification of the Thruster as Engineering Model is in progress.

• The RIT-4 as well as the RMT �N Ion Thruster use Xenon as propellant, resulting in benign thrust plumes that minimize contamination issues.




11.1.2.2 Arrangement of milliN and µN thrusters To be capable of performing rotations and translations in all directions, a minimum of 12 thrusters are required. This is true for both the mN-thrusters and µN-thrusters, resulting in a total of 24 thrusters. It is however desirable to have some redundancy, and for this reason 4 extra thrusters of each type are baselined, resulting in a total of 16 mN-thrusters and 16 µN-thrusters.The foreseen thruster configuration is to have 4 pairs of thrusters of each type on both the service module and the payload module, see Figure 65. The thruster pairs or pods, share some hardware, for example control and power electronics and for the ion engines the propellant supply lines. The thruster pods mounted on the cryogenic side of the spacecraft will have to be thermally isolated in order to ensure an operational temperature close to room temperature, note that this is also true for the tubing, transporting the Xenon gas to the ion engines (an alternative gas is Argon).

Figure 65 Temptative distribution of the FF thrusters on the s/c.

Top view

Side view

Top view

Side view




11.1.3 CONTROL ASPECTS

This section describes the Guidance, Navigation and Control (GNC) system based upon the main modes as anticipated in the system at this point in time. These modes are illustrated in Figure 66. In several of these modes there will be a number of sub-modes, which are not described in detail here. Different sets of sensors and actuators are used by the control system depending on the active control mode. In this way it is possible to increase the formation flying precision in relative position and attitude as more accurate sensors and actuators are utilized. The research and development activities, up to today, have led to a system sampling time of 1 second having typically no control loop bandwidths faster than 0.1 Hz.

Figure 66 State transition diagram for the top level modes in the GNC system. The dotted line separates the 2 safe modes, used by the respective modes on each side.

Separation Mode: This mode is activated at separation of the spacecraft from the launcher upper stage. The launch vehicle delivers the each spacecraft in a transfer orbit to L2. The spacecraft are initiated and automatically despun and acquire coarse sun-pointing, ensuring that the solar arrays receive sun illumination. The spacecraft are controlled by their own on-board system and fly in a loose formation. The sun shield and solar array are deployed as soon as




possible in this mode. The GNC system provides coarse pointing and control as required for slew maneuvers, using non-mechanical gyros, coarse sun sensors and star tracker(s) as sensors. Formation Deployment Mode: The spacecraft are separated by the control system to a safe distance. All spacecraft are pointed anti-sun with coarse precision while maintaining relative position control. Based on the RF metrology, which is also used for the continuously active collision avoidance, the four spacecraft are brought into the same line by the GNC, i.e. coorbital. At some point during the transfer the trailing pair of spacecraft will arrive at the two leading ones. The four spacecraft will be created within range of the RF system. Minimum Delta-V will lead to a formation with all 4 spacecraft on the same trajectory one behind the other. Cruise Mode: This mode is used after separation and up to arrival at the Lissajous orbit about L2. It includes coarse pointing of the spacecraft, providing power and protecting the payload from direct sun illumination during the transfer, thus allowing the payload to slowly reach the cryogenic operational temperature. Same type of avionics equipment as in the Formation Deployment Mode is used. Orbit Correction Mode: This mode is entered to perform orbit correction maneuvers during the parabolic transfer to L2. The GNC system points the spacecraft to an accuracy better than 0.5 deg, in order to ensure that the Delta-V is being delivered in the correct direction. Same type of avionics equipment as in the formation deployment Mode. The maneuvers shall ensure that the spacecraft are correctly injected in the intended orbit at L2 and to ensure that the two trailing spacecraft “catch up” with the leading two spacecraft during the approximately 100 day transfer. Coarse Formation Mode: This mode is similar to the formation deployment mode, though the target formation is different. This is also the mode that the formation enters, from safe mode, which is entered following detection of an control anomaly. The telescope spacecraft form a circle, with a radius of about 50 m and angular separation of about 120 degree, around the hub spacecraft. Relative attitude and position determination is based on RF metrology. Spacecraft attitudes and relative positions are controlled by actuating milli-Newton thrusters. The relative positions accuracy in this mode is 5-8 cm along each axis. Baseline Control Mode: The accuracy of the relative positions are improved. In order to control the lateral relative positions to better than 1 cm a divergent laser metrology system is used. This sensor will bridge the gap between the RF metrology and the 1 cm acquisition range of the fine lateral sensor. Lateral position determination, as provided by the divergent laser, and range information by the RF system are the main sensor in this mode. The actuation is by a micro-Newton propulsion system, which is used in subsequent modes. The architecture of the GNC system is centralized, typically located on the hub. Measurements are as far as possible performed on the hub and commands for actuation is communicated back to the telescope spacecraft. Fringe Acquisition Mode: The fine laser metrology system is used for both attitude and position control loops. The optical path difference is scanned, using the full stroke (2 cm) of the Optical Delay Line, until fringe lock is dedtected on the fringe sensor. The control loop for the ODL is faster that the outer position loops. The ODL locks on the central (white) fringe, even if the outer loop moves with in its 1 cm allowed range. The design of the inner and outer control

loop is performed considering couplings and interactions between them. When the ODL




approaches one of the end positions, the actuator is slowly moved towards its center position by moving the telescope spacecraft on the outer loop guidance. Normal Observation Mode: In this mode the relative positions and attitudes of the formation are kept fixed while fringes are observed. The optical path difference is controlled to an accuracy better than 5 nm using the ODL actuator, while spacecraft relative positions are controller to better than 1 cm. The relative attitude between spacecraft is critical with respect to the wave front tilt at combination. This leads to either very stringent pointing requirement for the telescopes or a fast tip-tilt mirror. This control system would be similar to the two stage control of relative positions. In this mode high precision sensors for position and attitude and micro-Newton actuators are used. It has been demonstrated that the resolution of micro-Newton thrusters have to be around 0.1 micro Newton to achieve the required pointing and position control.

During normal observations it should be possible to rotate and resize the formation, maintaining fringe lock and high pointing accuracy. It shall be possible to rotate or resize the formation with relative velocities up to 3.7 mm/s. Reconfiguration Mode: When an observation is completed, the formation is slewed to the next target. In most cases the formation will have to be resized as well, in order to adjust for the angular separation between the star and the habitable zone. During the reconfiguration the attitude and position requirements are relaxed. The RF metrology or a divergent laser could be used as the main sensor. Note that the hub spacecraft is in charge of formation control, i.e. centralized formation control. In this mode the formation reconfiguration, i.e. switching between equilateral to linear array, take place. Safe Mode: Safe mode can be triggered either as a result of an individual spacecraft loosing position and / or attitude w.r.t. the formation, or more severely, more spacecraft loosing position and / or attitudes. The control system architecture remains central with a backup decentralized system. The development of guidance, navigation and control algorithms for precision formation flying missions is addressed in the on-going TRP studies “Interferometer Constellation Control” and “Interferometer Constellation Deployment” that are further discussed in Sections 11.12 and 11.13, respectively.




11.2. FF complement: overview of development status Table 51 summarizes the current development status for the Darwin formation flying complement, including metrology, propulsors and Guidance and Navigation System. Table 78 Technology development status for the Darwin formation flying complement

FF METROLOGY RF metrology BB demonstrator available, including the main features of the RF

Metrology S/S concept, see Section 11.3. Laser metrology BB demonstrators available for most laser metrology subsystems,

see HPOM activity in Section 11.4. FF propulsion mN thrusters No Darwin-specific activities, relying on heritage from LISA

Pathfinder µN thrusters No Darwin-specific activities, relying on heritage from LISA

Pathfinder Guidance and Navigation Control Constellation Control GNC simulator available for the 3 fundamental modes of the

Darwin interferometer, namely the Baseline Control Mode (BCM), the Fringe Acquisition Mode (FAM) and the Nominal Mode (NOM). See activity ``Interferometer Constellation Control'' (ICC) in Section 11.12

Deployment Control GNC simulator available for the deployment modes of the Darwin interferometer, namely the jettisoning from the launch dispenser, the transfer from Earth to the L2 orbit and the creation of the baseline control mode of the formation. In the deployment phase See activity ``Interferometer Deployment Control'' in Section 11.13.




11.3. Formation Flying RF Metrology Technology Development


The RF FF metrology S/S is the first element in the metrology system chain (together with sun sensor and star tracker). These sensors will ensure initial good attitude and position accuracy for the subsequent metrology systems (coarse lateral metrology, fine lateral metrology, fine longitudinal metrology). The FF RF metrology S/S provides autonomous restitution of coarse position and attitude (option), throughout all mission phases. This note constitutes the summary of the activity “Formation Flying RF Sub-System” under ESA contract n°15511/02/NL/EC. The consortium selected for this study was led by Alcatel Space and comprised GMV as sub-contractor. Alcatel Espacio has participated in phase 2 of the study in the design of the RF transmitter. The activity was divided in 2 phases. Phase 1 was focused on design of the Darwin hexagonal array and Smart-2 (later for Smart-3) RF Metrology S/S. Phase 2 consisted in the development of a breadboard (BB) intended to demonstrate the main features of the RF Metrology S/S concept, and the performances based on laboratory experimentation and analysis of the results. Future activities will cover the development of an engineering model, overcoming the limitations of the current BB.


The FF RF metrology S/S comprises of one RF terminal on each spacecraft. The RF terminal includes a set of transmit and receive antennas, and one payload equipment unit, consisting of transmitter, receiver and navigation processing units. The FF RF metrology S/S computes the relative navigation between spacecrafts for the GNC S/S. The main functionalities of the RF FF metrology S/S are: • Computation of relative navigation: position, attitude (option), velocity, attitude rate

(option) and local time, for all spacecrafts in the formation, in all operational phases (Formation Deployment, Cruise Mode, Coarse Formation, Baseline and Reconfiguration Control Modes); with two basic functions, collision avoidance and coarse position. During the deployment, it is used for anti-collision and setting up of the nominal formation. During the nominal formation, it is used for formation reconfigurations.

• Navigation signal and message generation, transmission and reception to/from all other FF RF terminals, in all operational phases. The signal includes also the provision of a local inter-spacecraft data link (TM/TC) of 9 kbps for the BCS and 3 kbps for the TS.

• Provision to each S/C of a physical synchronized clock reference (Local System Time) with the rest of S/Cs.

• Autonomy and robustness. No hybridization with other sensors is required, except for the case of deployment of the two spacecrafts configuration. In this case, hybridization is required with other attitude sensors (for instance, star trackers) to provide full relative attitude and attitude rate restitution with respect to the body related frame (and, eventually,




with respect to an inertial frame). In centralised mode operation, tolerability to a single central-node spacecraft (BCS) failure, and tolerability to multiple telescope s/c failure, with navigation accuracy performance degradation.

11.3.3 DESCRIPTION

The phase 1 of the study was devoted to the definition and to the architectural design of the Formation Flying RF S/S for Darwin and Smart-2. The logic of the study was to analyse first the concept for the Darwin mission, and then in a second step to derive requirements and architecture for the Smart-2/ Smart-3 missions. The objective of phase 2 was to demonstrate the formation flying concept through a laboratory breadboard. The BB was used for testing and validation of the RF S/S for 3-satellite and 2-satellites configurations. The validation of RF S/S for Darwin configuration was performed using a SW simulator developed in phase 1. The FF RF S/S equipment for each spacecraft consists of one RF terminal, including a set of antennas, a Navigation Processing Unit (NPU) and a Receiver/Transmitter unit (Rx/Tx). The Rx/Tx unit broadcasts FF RF signals (for use by other Rx/Tx units), receive FF RF signals (from others Rx/Tx units) and provide measurements to the NPU. The NPU computes ranging and angular measurements (elevation and azimuth) within the formation, and passes this information to the GNC algorithms. The RF link is in S-band, based on GPS receiver heritage. Two frequencies (S1: 2100.00 MHz, S2: 2210 MHz) are used to allow robust sub-centimeter accuracy and short initialisation times. The first frequency, S1, is modulated in QPSK, with only ranging signal in I and only data in Q. The second frequency, S2, is modulated in BPSK with only ranging signal. The access is TDMA/CDMA. Measurements consist of pseudoranges at two frequencies (P1 and P2) between transmitter and receiver of different spacecrafts, and single difference between carrier phases (S1 and S2) for the several antennas located on the same spacecraft. These measurements allow the NPU to compute the relative navigation within the formation. The NPU provides in this manner the navigation output data to the GNC S/S. The navigation output data is formed by the ranging and angular measurements (elevation and azimuth) between space-crafts @1 Hz , as well as the TM/TC data. The relative position and velocity is computed and provided in

The FF RF BB developed within the phase 2 was used to demonstrate feasibility and performance of the concept proposed for the Darwin mission and for the Smart-2/Smart-3

Cartesian and polar coordinates, for all the spacecrafts in formation. The relative attitude and attitude rate (when applicable) of the spacecrafts are expressed in Euler angles (321 sequence). FF RF metrology S/S operation, and consequently the NPU processing, can be either in centralized or in distributed mode or in both simultaneously. For the deployment phase, the distributed mode is used, whereas for the nominal formation, the centralized mode is used. Integer Ambiguity Resolution (IAR) is used for both ranging and angular positioning data, based on carrier phase and dual frequency, to meet sub-centimeter accuracy for the inter-spacecraft range and sub-degree accuracy for the angular measurements. The Rx and Tx sub-units of a same vehicle share a common clock, in order to have a unique clock bias. The basis for the design is a spaceborne GPS receiver. The Rx/Tx unit is composed of receiving RF hardware, digital baseband chip, microprocessor and associated memory, transmitting up-converters and RF power amplifiers, OCXO and DC/DC converter. The equipment is equipped with 7 antennas, comprising 3 Rx /Tx antennas and 4 Rx-only antennae. The baseline for the antenna type is S-band quadri-filar helix. The design of the Rx/Tx unit is modular, for accommodating different antenna configurations, from a basic and small 3-antenna configuration, up to a complex 7-antenna configuration.




demonstrators. This breadboard comprises of 3 RF terminal units and 1 PC hosting the real-time navigation processing software. Each RF Terminal transmits and receives a TDMA/CDMA L-band GPS-like signal, to and from all other RF Terminals. The breadboard is set up in a laboratory and RF Terminal units are connected through cables. Specific devices are used to represent the effect of dynamics, different formations configurations and multipath.

Figure 67: FF RF Metrology BB. View of the test bench. The navigation processing is running in a PC instead of embedded on the Tx/Rx unit. Two Rx/Tx units are located on the racks (S/C 02 and S/C 03), and the third one is based on an elegant BB, located on the table (S/C 01).

Major differences of the BB with the proposed design were the following. The signal transmission occurs in L-band and single frequency, instead of S-band (as required by the SCFG group) and dual frequency. Note that Integer Ambiguity Resolution (IAR) is not possible with a single frequency approach. The data transmission is performed at low data rate (50 bps) in one terminal only, for the initial TDMA synchronisation, instead of several Kbps in quadrature to the ranging signal. The navigation processing is running in a PC instead of embedded on the Tx/Rx unit. The breadboard is illustrated in Figure 67. Two Rx/Tx units are located on the racks, and the third one is based on an elegant BB, located on the table. The BB developed during the activity was not fully representative of the final design for Darwin, as already mentioned above. It is the intention of ESA to develop an Engineering Model (EM) of the FF RF Metrology S/S within the current GSTP-4 programme, to overcome the limitations of the BB and to meet the complete Darwin design as defined within the activity. This EM will also be used to validate the feasibility of the critical issues identified during the activity (see Section 11.3.6). The architecture and functionality of the EM will be fully representative of the final Flight Model (FM), with the same electrical design, but using commercial components in some parts of the Tx and Rx sub elements, that will be replaced by space qualified components for the FM.




The complete SW will also be embedded on the equipment, including the signal and navigation processing.

11.3.4 PERFORMANCE

11.3.4.1 General performance assumptions The following assumptions have been made for performance assessment: Deployment ranges:

• The distance between spacecrafts during the deployment phase are: min 5 m (the wingtip-to-wingtip for 2 satellites is 10 m ) and max 8 km (with data demodulation, and 15 km with only ranging function)

• The relative angular rate during deployment is < 5 deg/s

Reference frames:

• During nominal configuration: The state vector is computed in a reference frame attached to the satellite body. In nominal configuration, the state vector is computed in the Rotating Body Frame (RBF), defined by :

Origin Centre of Mass of the Hub. +X axis Pointing towards CS number 1 along the science beam. +Y axis Y = Z x X +Z axis Perpendicular to the plane of science-beams.

• During deployment, the positioning is referenced to the Geometrical Fixed Frame

(GFF) of each S/C, defined by : Origin Centre of Mass of the spacecraft. +X axis Pointing to the direction defined by the science beam. +Y axis Y = Z x X . +Z axis Along the bore-sight of the telescope.

S-band antenna assumptions:

• Same antenna is used for both carrier frequencies and for Tx and Rx • Phase centre accuracy < 1 mm (rms) for nominal, and < 5 mm (rms) for deployment • Antenna gain (omni-directional) is higher than –3 dB over off-axis angle [-90°;+90°] • 4 Rx and 3 Tx/Rx antennas. Full sphere (4 π steradian) visibility with a minimum –3 dB

gain Navigation assumptions for the deployment mode:

• The FF RF S/S navigation processing is decentralised during deployment, where each S/Cs computes the relative position to the rest of S/Cs one by one (equivalent to a case of two S/Cs)




• At least 3 RF antennas are located in a plane perpendicular to the +X axis (3 antennas in

each side of the S/C), with the sun-shield not yet deployed (or RF transparent). RF measurements error budget:

• The receiver pseudorange and carrier phase error were tested and validated with the FF RF S/S BB. However, the multipath error and biases are theoretical assumptions (maximum bounds). These two elements are key parameters that need demonstration. The antenna baseline for the angular measurements was assumed to be 1 m.

Table 79 Measured performance with the FF RF metrology BB

Measurement accuracy (1-σ) Pseudorange (P1/P2) 1.2 m Carrier phase (S1/S2) 2.5 deg SD Carrier Phase (S1/S2)

5 deg

Mpath error (in P1 or P2) < 3m Mpath error (in S1 or S2) < 6 deg chan

n frq antenna

Biases (rms) 1mm 1 mm 1 mm Time tag accuracy 0.1 ms

The navigation performances presented in Table 80 and Table 81 are based on the FF RF metrology BB results for 2 and 3 spacecrafts formations (based on the Smart-2 and Smart-3 configurations), and extrapolated to Darwin hexagonal formation with a FF RF metrology SW simulator. The reference array configurations originally assumed for the FF RF metrology development activity are:

• Darwin hexagonal array. This is a 3D configuration consisting of six free-flyers (TS) located at the vertices of a hexagon, one hub satellite (BCS) at the centre of the hexagon, and one master satellite out of the plane formed by the TS/BCS. The hex Darwin array has been de-scoped in favour of TTN.

• Smart-2 nominal configuration: it consists of two satellites (TS and BCS) with the Centre of Mass placed in a straight line.

• Smart-3 nominal configuration: it consists of three satellites (two TS and BCS) located in a plane, forming an equilateral triangle.

• Deployment configuration: random position and orientation of the satellites within a sphere of radius r.

11.3.4.2 Measured performance: two s/c test case Table 80 summarizes the measured performances for the FF RF Metrology S/S for the two s/c case (Smart-2). Performance degrades for increasing azimuth and elevation angles. This is the configuration assumed during deployment. Table 81 summaries the measured performances for the case of three S/Cs (Smart-3). The addition of a third s/c allows maintaining planarity of the




formation, so that relative azimuth and elevation angles always remain within acceptable levels. In Table 81, the results are extrapolated to the case of the Darwin hex array. Table 80 Navigation accuracy (99.9%) measured with the FF RF metrology BB, for the case of two S/Cs (deployment phase). Note that these results also apply to current Darwin system during the deployment phase

Performance

Measured

2 S/Cs (Deployment)

Remarks

Position Range 0.7 cm Elevation 0.7° (El<30°)

1.1° (El<60°) 2.5° (El<80°)

For El>80° the performance degrades significantly with a singularity at 90°

Azimuth 1° (Az<60) 4.1° (Az<75)

For 105°Az>75° the performance degrades significantly with a singularity at 90°

Velocity Range rate 1.4 mm/s Elevation rate 0.05°/s (El<30°)

0.08°/s (El<60°) 0.2°/s (El<80°)

For El>80° the performance degrades significantly with a singularity at 90°

Azimuth rate 0.09°/s (Az<60) 0.32°/s (Az<75)

For 105°>Az>75° the performance degrades significantly with a singularity at 90°

Table 81 Navigation accuracy (99.9%) measured with the FF RF metrology BB, for the case of three S/Cs, and extrapolated for the Darwin Hexagonal array configuration

Performance

Measured

3 S/Cs

Darwin Hexagonal

(FF RF metrology SW simulator)

Remarks

Position

Range 0.7 cm 0.2 cm Elevation 0.7° (El<30°) 0.05 ° Azimuth 0.33 ° 0.04 ° Velocity Range rate 1.4 mm/s 0.4 mm/s Elevation rate 0.05°/s (El<30°) 0.03 °/s Azimuth rate 0.09°/s 0. 005 °/s Attitude Attitude (at d=25 m TS-TS)

- 0.28°/0.04°/0.22° Euler angles accuracy (99.9%) in 321 sequence

Attitude (at d=250 m)

- 0.15°/0.11°/0.23°

Attitude rate - 30 mdeg/s




11.3.4.3 Expected performance for planar and EMMA TTN Table 82 summarizes the expected performances for the FF RF Metrology S/S for the current Darwin configuration, the planar TTN (see Section 2.2.3) and for the 3D Emma TTN architecture (see Section 2.2.4). For the planar array, performance figures are extrapolated from the BB results for 2 and 3 spacecraft formations (Smart-2 and Smart-3 configurations), accordingly to the new number of satellites and geometries. The shown performance refers to three S/Cs forming an equilateral triangle around the BCS. For the case of four or five coplanar S/Cs the accuracy performances are expected to improve due to a better geometry and higher number of measurements. The range and azimuth performance are based on the results obtained with the FF RF S/S SW simulator and validated with the BB for the case of 3 S/Cs. The values presented for elevation are based on the results obtained with the FF RF S/S SW simulator and validated with the BB for the case of 2 S/Cs during deployment forEl<30°, and improved by a factor √3 due to the higher number of signals (3 instead of 1). At least 3 RF antennas are located in a plane perpendicular to the +X axis (3 antennas in each side of the S/C).

Table 82 Expected navigation accuracy (99.9%) during nominal formation for the planar configurations (three/four/five S/Cs located in the same plane) and EMMA configuration (four/five S/Cs, TS’s located in a parabolic surface, and the BCS out of the surface).

Performance

Planar arrays (triangular &

linear)

EMMA arrays

Remarks

Position Range 0.7 cm 0.3 cm Elevation 0.4° 0.21° Azimuth 0.33° 0.13° Velocity

instead of 6 TS’s in hexagonal formation with 120 deg vertices. This results into Dilution of

Range rate 1.4 mm/s 1.4 mm/s Elevation rate 0.05°/s 0.04°/s Azimuth rate 0.09°/s 0.08°/s Attitude Attitude (at d=25 m TS-TS)

- 0.28°/0.04°/0.22° Euler angles accuracy (99.9%) in 321 sequence

Attitude (at d=250 m)

- 0.15°/0.11°/0.23°

Attitude rate - 30 mdeg/s

For the Emma configuration, the presented accuracy performance are based on the results obtained with the FF RF S/S SW simulator and validated with the BB measurements for the case of 8 S/Cs based on the hexagonal Darwin array, and extrapolated to the EMMA equilateral triangular configuration. A degradation factor of √2 has been applied since the number of measurements is halved (4 S/Cs instead of 8 S/Cs). Also, a degradation factor of around 3 due to the different formation geometry (3 TS’s in triangular formation with 60 degrees vertices




Precision factors (DOP) in the three directions by XDOP=2/3 and ZDOP=5 for both configurations, however the YDOP passes from 3/2 to 2 from the hexagonal to the triangular configuration). For the case of the BCS it is expected similar performance than for the TS when the height of the BCS is similar to the distance between TS, however progressive degradation is expected for the accuracy performance of the BCS position (X and Y axis) for higher heights. Attitude accuracy figures refer to the hexagonal Darwin configuration with 8 S/C. These values have not been extrapolated to the EMMA configuration, but a degradation factor around 4 is expected. Finally, timing performance is addressed in Table 83.

Table 83 Measured (or expected) timing performance. These figures applies to any array configuration.

Time synchronization accuracy between S/Cs (99.9%) Time accuracy 0.1 ms Time To Fix Formation Navigation Worst case: quasi-static conditions TTFF: first fix (95%) TTFA: first fix with nominal performances (95%) Deployment Initialisation time < 60 s TTFF warm start <51 s TTFF cold start <300 s TTFA warm start < 151 s TTFA cold start < 400 s Nominal Formations

Initialisation time < 60 s TTFF warm start <281 s TTFF cold start <530 s TTFA warm start < 1921 s TTFA cold start < 2170 s TM/TC Data Data provision latency

0.5 s

Frame error rate 10-3 /s TDMA synchronisation

<0.1 ms





4

Remarks

Table 8 Interface requirements for the RF metrology


Value

Volume (mm×mm×mm) 290 x 180 x 130 RF front-end, Rx/Tx digital section and navigation processing unit implemented in one single equipment

5.4 Antennas not included. The mass of each antenna is in the order of 0.5 Kg

Interfaces 9 SMA connectors for connection with antennas 1 DM15 connector for RS422 links with OBC (RF

metrology data and local TM/TC) 1 DM9 connector

1 DM9 connector for ON/OFF commands and relay status output

Antenna accommodation requirements

Distance between antennas > 1m

TBD, not critical Antennas deployment maybe needed, depending on the S/C shape, in order to have always 4pi str visibility with at least 3 antennas.

Equipment accommodation TBD

Power

20 W During nominal phase Peak consumption (W) 30 W During deployment phase

Thermal interfaces

Operational temperature (C) [-25°C, +60°C] TBC Operational temp. range (C) TBD Heat dissipation TBD

Other

Exported vibration level (µN) [0 atm, 1 atm]

Radiation TBD Out-of-band spurious < -25 dBc inside [1.5 GHz,

3.0 GHz] < -TBD dBc outside [1.5 GHz, 3.0 GHz]

TBC

In-band spurious < -25 dBc

Mass (kg)

for primary power supply


TBD Operational pressure

TBC





The feasibility of antenna accommodation is to be confirmed with spacecraft design analysis, especially when it comes to the implantation of antennas on the deployable sun-shield. It is important also to note that the RF behaviour of the sun-shield has a high impact on multipath error (i.e. RF-transparent sun-shield is more favourable than RF-reflective). A precise calibration and stability of position of antenna phase centre is needed (1 mm level). This is a matter to be demonstrated during future EM development, through measurements in anechoic chamber with spacecraft model.

Final performance will be highly sensitive to multipath mitigation performance. It is obviously expected that spacecraft design will be constrained by other drivers than multipath mitigation and therefore will not be optimal to this respect (e.g. number of edges, reflection coefficient of surface). It has been shown that about 70% mitigation was needed at software level, for carrier phase measurements, to get reasonable performance. The only method deemed realistic for Darwin case is based on repeatability of multipath. It leads to calibrate multipath error from ground measurements performed in anechoic chamber. It has thus a major impact on spacecraft design and testing: this should be the object of careful analysis in the Darwin program.

11.4. High-Precision Optical Metrology (HPOM) Technology Development

Several critical issues have been identified during the activity.

A precise calibration and stability of Rx/Tx hardware biases is needed (1 mm level). Feasibility was demonstrated during phase 2, but it will have to be further validated during EM development.


The high-precision optical metrology (HPOM) contract started with an assessment of the positional accuracy requirements for the DARWIN satellites. Taking the Robin Laurance constellation of 6 telescope satellites (without an out of plane communication satellite) as a baseline, it was found that the longitudinal and the lateral positional accuracy requirements are identical. This finding relaxed the requirements on the longitudinal metrology, which could (in principle, by using interferometry) be made much more accurate compared to lateral metrology. Based on the assumption that a constantly operating radio frequency metrology system (for constellation deployment and satellite collision avoidance) would provide absolute and three dimensional positional accuracies within a cube of 10x10x10 cm3 and from the requirement that initial interferometric white-light fringes should be acquired within 400 seconds, the accuracy requirement for the optical metrology was found to be 10 µm/√Hz, which corresponds to ±32 µm @ 10 Hz. Subsequently, a fine lateral sensor (FLS) and two longitudinal sensors were breadboarded. The fine lateral sensor, as developed by TNO-TPD (The Netherlands) is based on a collimated beam whose position is measured by a position sensitive device (PSD). The two longitudinal sensors are based on a frequency scanning interferometer (FSI), developed by INETI (Portugal), and a dual-wavelength interferometer (DWI) developed by SIOS (Germany). It turned out that the FSI




required that the distance between satellites does not change during the measurement, which is not the case with the DWI and the DWI was able to take advantage of the parallel operation of the radio frequency metrology system. As the out of plane performance of the radio frequency metrology system turned out to be inferior to 10 cm the need for a coarse lateral sensor (CLS) was identified, which would be based on a divergent beam whose back-reflected light (from a corner-cube) is imaged onto a CCD camera. As the CLS technology is assumed to be state-of-the art no technology breadboarding was performed.


11.4.2.3 Dual-wavelength interferometer

Three critical sub-systems of the DARWIN optical metrology have been breadboarded, namely a fine lateral sensor (FLS) and two coarse longitudinal sensors, a frequency scanning interferometer (FSI) and a dual-wavelength interferometer (DWI). One of the lasers of the DWI been additionally fitted with a Pound Drever stabilisation scheme to create a fringe tracking interferometer for absolute distance measurements with nanometric accuracy over distances up to 250 meters.

11.4.2.1 Fine lateral sensor The fine lateral sensor transmits a collimated beam (25 mm) from one spacecraft to another, where a large area detector (CCD, CMOS or PSD) measures the lateral offset. The lateral offset is determined by fitting a Gauss function over the received irradiance (in the case of a multi-pixel CCD or a CMOS sensor) to calculate the “centre of gravity” of the received light, or the “centre of gravity” is obtained electrically in the case of a position sensitive device (PSD). The lateral measurement range specified was ±1 mm and the accuracy 10 µm/√Hz, which corresponds to ±32 µm @ 10 Hz.

11.4.2.2 Frequency scanning interferometer In order to determine the longitudinal distance, the laser wavelength is continuously changed in a frequency scanning interferometer (FSI) by a well-known amount. Scanning the wavelength causes the fringes in an unequal-arm interferometer to move and by counting the fringes the distance to a target can be unambiguously determined. The drawback is that there must be not distance change during the measurement.

In order to determine the longitudinal distance, an interferometer is fed from two frequency-locked lasers. The phase-difference of the interference fringes from the two laser frequencies determines the distance to a target. Such a dual-wavelength interferometer (DWI) acts like a single wavelength interferometer with a new (synthetic) wavelength, which is given by the product of the two laser frequencies divided by their difference. The synthetic wavelength is chosen such that the interferometer’s ambiguity range covers the inaccuracy range of the radio frequency metrology.

11.4.2.4 Fringe tracking interferometer One of the two lasers of the dual-wavelength interferometer (DWI) is stabilised to 10-12 by locking it to an iodine absorption line. In this way nanometric measurement accuracy over distance of up to 250 meters can be achieved. This is not required for the DARWIN metrology system because the DARWIN stellar interferometer will provide this information. I may




however become necessary for a DARWIN precursor demonstration mission to demonstrate that formation flying with interferometric accuracy is feasible.

11.4.3 DESCRIPTION

11.4.3.1 Fine lateral sensor The principle layout of the fine lateral sensor (FLS) is depicted in Figure 68. A red laser beam is fed via a single mode fiber, collimated to a diameter of 25 mm and transmitted from satellite 2 to satellite 1. A large area position sensitive device (PSD) measures the centre of gravity of the received beam and thus determines the lateral offset between the two satellites. The fine lateral sensor (FLS) collimator (manufactured in aluminum) is shown on the left in Figure 69 and the complete setup mounted together with the dual-wavelength interferometer (DWI) is shown on the right.

Figure 6 Principle layout of the fine lateral sensor (FLS) 8

igure 69 Fine lateral sensor (FLS) emitter (left) and full FLS system (right) shown together

dual-wa

Fwith the velength interferometer (DWI) on an optical bench.




11.4.3.2 Frequency scanning interferometer The principle of a frequency scanning interferometer (FSI) is shown Figure 70. The optical path difference (or distance L) is determined in an FSI by changing the frequency of the laser source and counting the number of interference fringes with a photo detector (D). Obviously, the amount of frequency change needs to be known very precisely and a convenient method to measure this is to use a highly stable Fabry-Perot interferometer.

Figure 7 Principle of a frequency scanning interferometer (FSI) 0

A tunable laser source (depicted as laser 1) together with a highly stable resonant cavity (Fabry-Perot interferometer) is shown in Figure 71

Figure 71: Frequency scanning interferometer laser locked to Fabry-Perot cavity by servo loop




During a distance measurement, the laser frequency is swept for a predefined integer number of Fabry-Perot resonances. For the distance measurement to be precise it is vital that the frequency sweep range is accurately known and that’s why Fabry-Perot resonances are well suited as they are very narrow. During the frequency sweep distance interferometer fringes are also counted whose number is directly proportional to the optical path difference and inversely proportional to the used wavelength. If higher distance resolution is required (this was not implemented), also the interferometer phase at the beginning and at the end of the frequency scan can be measured as depicted in the last line in Figure 72.

Figure 72 Operations to be performed during FSI distance measurement

The principle layout of the frequency scanning interferometer (FSI) as built by INETI (Portugal) is shown on the left in Figure and a picture of the set-up can be seen on the right. The problem with an FSI is that the distance to be measured must not change during the measurement.

73




Figure 73 Principle layout and picture of the FSI breadboard

11.4.3.3 Dual wavelength interferometer The interference phase in any interferometer is proportional to the optical path difference (OPD) in the both interferometer arms and inversely proportional to the laser wavelength. Thus, if the OPD is increased from zero until it corresponds to the wavelength the interference phase increases from 0 to 360 degrees, after which the interference phase and thus the target distance become ambiguous. As laser wavelengths are very short, so is the ambiguity range. To overcome the problem two wavelengths are used. The principle layout of a dual-wavelength interferometer (DWI) is shown in Figure 74

4Figure 7 Principle layout of a dual-wavelength interferometer (DWI)




In a dual-wavelength interferometer (DWI) the beat (difference) frequency of two ultra-stable lasers is measured by a high-speed photodiode, compared to an stable oscillator and one of the two lasers is servo controlled to maintain the frequency difference. An interferometer is fed with both laser wavelengths, which renders the phase difference between interferences of both wavelengths. This phase difference is still proportional to the OPD, but inversely proportional to a synthetic wavelength. The synthetic wavelength is given by the product of the two individual laser wavelengths and divided by their difference. Thus, if the laser wavelengths are close the synthetic wavelength can be made large. In the case of the DWI build for HPOM, the wavelength difference is so small that it is better expressed in a frequency difference, namely is 3 GHz. This in turn corresponds to a synthetic wavelength of 100 mm and this is the un-ambiguity range of the DWI. It is interesting to note that the stability of the individual laser frequency is irrelevant in a DWI; it is only the (electrically controlled) difference frequency that determines stability of the interferometric measurements. In order to perform accurate DWI phase measurements heterodyning was implemented. Four cousto-optical modulators (two per wavelength) create heterodyne beat frequencies of 100 kHz (for laser 1) and 175 kHz (for laser 2). The heterodyne frequency generation scheme and a picture of it are shown in Figure 75.

Figure 75 Dual wavelength interferometer heterodyne frequency generation scheme and picture (please ignore the optical components in the upper left of the picture - on the separate small optical bench -, they do not belong to the frequency generation equipment)




Figure 76 Dual wavelength interferometer head, principle layout and picture. The delay line shown in the layout is not shown in the picture. Please note that the corner-cube at the end of the distance to be measured is very close to the interferometer head for testing.

Two polarization maintaining single-mode fibers carry the heterodyne frequency shifted laser beams to the optical head of the DWI. Both detectors in the optical head are subjected to the same laser and heterodyne frequencies. Their only difference is the interferometer phase, which is introduced by the longitudinal distance to be measured.

11.4.3.4 Fringe tracking interferometer One of the two lasers of the dual-wavelength interferometer (DWI) has been chosen to give a frequency-doubled output (a wavelength of 532 nm), which is used to stabilise the laser to an iodine absorption line. This Doppler-free Pound-Drewer wavelength stabilization technique enables stabilities in the order of 10-12, which is necessary to achieve a distance resolution of 5 nm over distances of 500 meters. Figure 77 shows the schematics and a picture of the Pound-Drewer stabilization method.

Page: 193



Issue: 1.2

Figure 77 Doppler-free Pound-Drewer laser stabilization, principle layout and picture.

11.4.4 PERFORMANCE


Metrology instrument Required performance Measured/Expected RemarksCoarse lateral sensor 1“ (rms) @ ±1 degree, 10 Hz 1“ (rms) @ ±1 degree, 10 Hz * Fine lateral sensor 10 µm/√Hz @ ±5 mm, 10 Hz 10 µm/√Hz @ ±5 mm, 10 Hz Longitudinal sensor (FSI) 10 µm/√Hz @ 250 m, 10 Hz 10 µm/√Hz @ 250 m, 10 Hz ** Longitudinal sensor (DWI) 10 µm/√Hz @ 250 m, 10 Hz 10 µm/√Hz @ 250 m, 10 Hz *** OPD metrology sensor 2 nm/√Hz @ 250 m, 10 Hz 2 nm/√Hz @ 250 m, 10 Hz **** Remarks: * Straight-forward technology, has not been bread-boarded ** Measured distance must not change during measurement *** Makes additional use of RF metrology system to obtain absolute distance **** Iodine line stabilisation implemented in DWI



Metrology instrument Volume Mass Power Temp. Range Opt. Interface Coarse lateral sensor 200x100x100 3 kg 8 Watts Fine lateral sensor 50 x 50 x 50 0.5 kg 5 Watts Longitudinal sensor (FSI) 200x200x100 16 kg 20 Watts Longitudinal sensor (DWI) 400x200x100 20 kg 27 Watts Fringe tracking sensor 300x200x100 20 kg 27 Watts





The Coarse Lateral Sensor (CLS) is considered to require only straightforward engineering and does not pose development risks. The Fine Lateral Sensor (FLS) has been breadboarded and his performance is well to specifications. The Frequency Scanning Interferometer (FSI) is a relatively simple system (as compared to the DWI), which can measure longitudinal distance unambiguously. However, it suffers the problem of sensitivity to distance changes during the measurement. Its effect could be measured and compensated using a second interferometer. Due to its simplicity it would be worthwhile to continue to study FSI using two lasers to cancel its drift sensitivity. The dual-wavelength interferometer is a more complicated system, which needs the information from the radio frequency system to be unambiguous, but it is insensitive to distance changes during the measurement. While the optical head has already been miniaturised in the breadboard, dramatic simplifications in alignment and stability can be achieved with the used of fibre optics (fibre coupled Nd:YAG lasers and acousto-optic modulators) as developed to space-qualified standard for the LISA pathfinder mission. Problems with noise in the interferometer, caused by offsets and drifts in the interferometer electronics have so far delayed the final performance tests, but the causes have been identified and the tests will be performed in April 2005. The fringe tracking sensor, which requires the stabilisation of one of the DWI lasers to a iodine absorption line to achieve stabilities in the order of 10-12, is only required in a DARWIN precursor mission where no stellar interferometer is implemented.

11.5. RITA-10 mN ion thrusters Technology Development


For the Darwin coarse thruster (~ mN) segment, an internal ESA trade-off has evidenced ion propulsion as potentially promising technique due to: • High specific impulse (~3,000 s) • High maturity, lifetime and reliability • High thrust capacity (up to 20 mN) • Benign thrust plumes • Reasonable power consumption (~ 30 W/mN) The University of Giessen (D), initially conducted European research into radio frequency Ion propulsion in the 1960’s. In 1970 industrial involvement was realised when MBB (now EADS) joined the development team. For more than 30 years this team continued with the research, development and refinement of Radio-frequency Ion Thruster (RIT) technologies, associated propulsion systems, analytical tools and techniques, processes and materials technologies. The first Radio-frequency Ion Thruster Assembly (RITA) was successfully demonstrated in space aboard ESA's European Retrievable Carrier EURECA, launched by the Space Shuttle Atlantis in 1992. At that time, the RIT-10 system aboard EURECA provided a nominal specific impulse of 3,058 seconds. More recently, RITA-10 was used to retrieve the Artemis satellite from total loss to a full recovery, after thrusting for 6,430 hours.




RITA-10 is space qualified and has demonstrated thousands of hours of flawless operation in space and tens of thousands of hours on ground.


For ESA’s ARTEMIS mission both the Electron Ion-Bombardment Thruster (EIT) and Radio Frequency Ion Thruster (RIT) technologies were embarked. A common Propellant Management and Distribution Assembly (PSDA) was provided to store Xenon at high pressure and regulate the supply of propellant to each thruster assembly. Furthermore, a first generation Ion Thruster Alignment Mechanism (ITAM), developed by Austrian Aerospace was used to accommodate pairs of thrusters for the purposes of North South Station Keeping.

Figure 78 ARTEMIS Ion Propulsion Package General Block Diagram

Each Radio Frequency Ion Thruster Assembly (RITA) consists of one RIT thruster together with the units necessary for its operation. The RF-Ion Thruster ionises the propellant Xenon by collision of electrons, which are accelerated by an RF Field, with Xenon atoms. The RF-field is coupled to the Xenon plasma inside the electrically non-conductive discharge vessel by a RF-coil located outside the discharge vessel. The extraction system consists of the anode, the plasma holder, the acceleration electrode and the end electrode. The anode is mounted inside the discharge vessel in the middle of the flat bottom of the ioniser covering the outlet of the isolator tube. It is electrically connected through the discharge vessel to the positive high voltage of the PSU (part of the PSCU) .The holder is a flat disc, known as the screen grid with 253 holes or apertures. The graphite acceleration electrode is mounted downstream of the ion beam behind the plasma holder also with 253 apertures alighted with those in the screen grid. It is electrically connected to the negative high voltage, and can be switched over using a relay in the PSU to the igniting voltage. The end electrode (deceleration electrode) is mounted downstream of the ion beam behind the acceleration electrode. It is electrically connected to the thruster ground using hold-down rings and 6 fastening bolts. The breakdown voltage between the accelerator electrode and the end electrode is above 3.5kV at operational conditions. Ions from the plasma inside the discharge vessel are accelerated by an electrostatic field, which is generated by the different voltages applied to a grid system.




Electrons are generated by a hollow-cathode neutraliser and used for the ignition of the discharge at the beginning of the thruster operation. To avoid the satellite becoming electrically charged by the expulsion of only positively charged particles, electrons are fed into the beam by the neutraliser. Thrust control is achieved through varying the plasma density through a combination of RF power and propellant control. Each RITA consists of the following units: - 1 Flow Control Unit (FCU) - 1 RIT 10 thruster and neutraliser - 1 Radio Frequency Generator (RFG) - 1 Power Supply and Control Unit (PSCU)

NE

UT

RA

LIS

ER

(Cat

hod

e)

RF -G e ne ra to r

P la s m a H olde r( An ode )

A cc e le ra tor

Ig nit in gVo lta ge

R IT

Flo w Co ntro l Un it

X en on F e edS ys t e m

X e non Ta nk

Io ns Xe+ Elec tron s e -

+

+ -

-

- +

+

+

--

R F -Co il

Isola tingG as Feeder

D ischargeVessel

C a th ode H ea t e r

N eu tra lis erKe e pe r

Figure 79 Block diagram explaining the operation principle of a RIT thruster.

11.5.3 DESCRIPTION

The RF-ion thruster RIT10 has an inner ionisation chamber diameter of 100mm and a beam diameter of 87mm. Its development was started for North-South Station Keeping (NSSK) of geostationary satellites in the 1970’s with a 10mN thrust level. Since then, development of the thruster has been steadily continued and the thrust level for the ARTEMIS application was increased to 15mN. Nominally, the ARTEMIS mission qualification required a lifetime test on the thruster only. EADS-ST decided to demonstrate the performance of a complete RITA system in a full lifetime test at the ESTEC test facility. The RITA has successfully completed this test over the required 15,000 hours and subsequently was extended to support the mission during the orbit raising, ultimately achieving more than 20,000 hours. Based on these results, the predicted RIT-10 life is higher than 30,000 hours. Although the RIT-10 is the state-of-the-art engine with respect to lifetime, development of the thruster continued obtaining further improvements via an advanced grid system would improve




the thruster performance significantly and maintain lifetime capabilities at the high RIT-10 level. Based on RIT-10 for ARTEMIS the development of a high performance grid system was initiated. In order to keep the development linked as close as possible to a space-qualified concept, spare parts of an ARTEMIS RIT-10 were used to build this new thruster, called RIT-10-EVO. Beside in-house experiences, the results of ESA’s GSTP “Grid improvement” programme had decisive impact on the development and understanding of Ion Optics capabilities. In addition, the knowledge gained by the development of ESA-XX was introduced in the design. The RIT10-EVO was manufactured in 1997. Initial testing was performed in January/ February 1998 at the ESTEC test facility in Noordwijk. As predicted, the new EVO grid realised a significantly increased performance and even reduced grid erosion. • Thrust regulation bandwidth demonstrated: 0,1 to 41mN • Increased specific impulse +300s at const. Beam voltage • Reduced acceleration grid loss current from 1,5% to approx. 0,7% (indicating a reduced grid erosion and thus an increased lifetime) • Reduced specific power from 35 W/mN of the RIT10 thruster to 26 W/mN for the RIT10-EVO The significant higher thrust of RIT10-EVO qualifies the thruster for a broad field of applications. In addition, it is capable to operate at various thrust levels (1mN up to 40mN), with throttling achievable in a extremely short time scale. This unique behaviour was tested and verified during an ESA funded test during the GOCE phase A at the University of Giessen from February to May 2000.

Figure 80: T-5 & RIT-10 on ARTEMIS (left) and RIT-10 Thruster Unit (right)




11.5.4 PERFORMANCE

Table 87 Performance of the RIT-10 ion thurster system

Performance

Nominal

Demonstrated

Remarks

Thrust 15mN

1-40mN

>7,500 hrs operation achieved during ARTEMIS mission.

Specific Impulse 3300s 2500-3700s

Flow Rate 0.45mg/s

Total Impulse 0.81E6 N.s (Artemis Mission) 1.07E6 N.s

Qualification life test for Artemis mission performed within Electric Propulsion Facility at ESTEC. The RITA test activity at ESTEC was voluntarily terminated in November 2002.

On Time 15,000 Hrs (Artemis Mission) 20,083 Hrs

Predicted life exceeding 30,000hrs (Extrapolating data from end of life test)

Cycles 5,000 (Artemis Mission) 6,741





Table 88 Interfaces of the RIT-10 thruster system


Value

Remarks

Dimensions (mm) RIT-10 Thruster 205 x 185 x 165

PSCU 150 x 135 x 80

FCU 270 x 185 x 115

Dimensions of Artemis Flight Equipment

Mass Budget (kg) RIT-10 Thruster 1.7

1.3 RF Generator 8.0

FCU RITA total 15.1kg

Power (W) 575W At 15mN

Thruster Power 495W At 15mN Thruster Power 660W At 25mN (RIT-10 EVO)

Operational temperature (K) Delta T: 140K 223°K - 573°K

Other No vibration expected TBC

286 x 250 x 183 RF Generator

PSCU

2.0 Harness 2.1

Mass of Artemis Flight Equipment

PSCU Power

Thermal interfaces

Operational temp. range (K)



Currently there are no ongoing development programs related to the RIT-10. Final testing of the EVO version demonstrated that the predicted performances coming from the numerical simulation and modelling tools developed previously were achieved. As the original ARTEMIS design was used to evaluate the grid set, future activities or programs would benefit from the systems flight-qualified status. Current telecommunication satellite developments are utilising electric propulsion for NSSK, and a larger version of the RIT-10, the RIT-22 (4.5kW/150mN) has recently been tested. At the other extreme, the university of Giessen have initiated the development of a µN-RIT system by scaling down the RIT-10 design. Since the RIT-10 development for ARTEMIS, there has been significant progress in the area of PSCU and propellant management technologies. Any future exploitation of the RIT-10 thruster would benefit from these developments, reducing the mass while increasing controllability and autonomy of the system. Without a current market opportunity, the second-generation developments have yet to be demonstrated with the RIT-10, but have however been applied to the larger RIT-22 thruster.




Although very capable in terms of meeting and exceeding ARTEMIS requirements, the thruster was tested at a fixed operating point. For any future activities requiring a wide throttling range the RIT-10 design will require evaluation to ensure that lifetime requirements will be achieved. This would require further testing of the system and implementation of fine control propellant flow devices.

11.6. T-5 mN ion thrusters Technology Development

The mechanical design of the T-5 thruster, culminating in the T-5 MkV, is the product of three decades of research and development. During this time the basic plasma and ion beam features have remained virtually unchanged. Consequently, the majority of the test data from this long period became applicable to scientific missions such as GOCE. An engineering model (EM) T5 has been successfully tested for more than 5000 hours in support of the GOCE programme.


The T-5 thruster was developed from the UK-10 and UK-25 thrusters using well-established and proven scaling laws. A larger T-6 thruster is currently underdevelopment for future Telecommunication and interplanetary missions. All of these thrusters have exhibited good performance, resulting principally from refined ion optics and efficient discharge chamber designs.

In contrast to Telecommunication applications (e.g. ARTEMIS / ALPHABUS) the GOCE mission will demonstrate the capability of achieving wide throttle range and stringent thrust control requirements with Ion Engine Technologies.





For ESA’s ARTEMIS mission both the Electron Ion-Bombardment Thruster (EIT) and Radio Frequency Ion Thruster (RIT) technologies were embarked. A common Propellant Management and Distribution Assembly (PSDA) was provided to store Xenon at high pressure and regulate the supply of propellant to each thruster assembly. Furthermore, a first generation Ion Thruster Alignment Mechanism (ITAM), developed by Austrian Aerospace was used to accommodate pairs of thrusters for the purposes of North South Station Keeping.

Figure 8 ARTEMIS Ion Propulsion Package General Block Diagram 1

Thruster operation is based on the Kaufmann concept, which employs a Direct Current discharge into a divergent magnetic field to ionise the propellant gas. There are three feed pipes delivering Xenon propellant to the thruster, one each for the cathode, main flow assembly and neutraliser assembly. Electrons are emitted by the cathode, located at the centre of the discharge chamber back plate, and form a plasma with the Xenon gas flowing through it.

A cylindrical anode is located on the inside of the discharge chamber outer circumference, in order to accelerate electrons by means of the potential difference between it and the cathode. Electrons emitted from the cathode, and those knocked off atoms, are eventually drawn off at the anode. To increase the probability of collisions taking place and hence of ionisation, the paths of the electrons are increased by a magnetic field superimposed on the discharge chamber by an electromagnet.

The magnetic circuit consists of an inner pole located around the cathode outlet, the back plate and the discharge chamber, solenoids located around the outside of the chamber and an outer, front pole. The electrons are thus caused to spiral through the chamber by the magnetic field linking the inner and outer poles. The positively charged xenon ions drift through the discharge chamber towards the screen grid, which forms the front of the discharge chamber. A baffle disc in front of the cathode outlet prevents high-speed particles from passing directly out of the cathode and impinging upon the grid.




Located outside of the screen grid is an accelerating grid, which is maintained at a much lower potential than the rest of the spacecraft. The screen grid prevents the ions from being affected by the accelerating grid until they are very close to the screen grid, by which time the electrons have been attracted to the anode. Once the ions "see" the high negative potential of the outer grid they are accelerated through the holes in the grid across the narrow gap between the two and leave the thruster at a high exhaust velocity. To prevent charging the spacecraft due to the expulsion of the positively charged beam, a neutraliser is located with its outlet adjacent to the grids. The neutraliser operates in the same way as the cathode to produce a stream of electrons, which neutralise the positive beam. The discharge chamber and solenoids provide their own support structure upon which the outer earth screen is mounted.

Figure 82 T-5 Thruster Schematic.




11.6.3 DESCRIPTION

QinetiQ’s experience in the field of Electric Propulsion has been gained through more that 30 years of research and development. Extensive research has been carried out into understanding the operation of the thrusters and their respective interactions with propellant control and power systems. In the early 1990’s, the UK-10 ion thruster was selected for the ESA ARTEMIS mission and QinetiQ supported this program through formal qualification of the discharge and neutraliser cathodes. The UK-10 was based on the T-5 design and QinetiQ provided comprehensive technical support to the industrial team lead by EADS-Astrium. In 2000, the T-5 engine was selected for the ESA GOCE mission.

11.6.3.1 Ion Propulsion Control Unit (IPCU)

The GOCE requirements have created challenges in terms of the grid design. In particular the wide thrust range, coupled with the long lifetime requirement have resulted in considerable grid optimisation from the original UK-10 design. This was achieved through detailed modelling of the ion extraction and erosion processes using an ion optics-modelling tool. This work has resulted in a twin grid design (triple grid used for UK-10). Through a change in the accelerator grid material to graphite, a considerable increase in lifetime capability has been achieved whilst reducing mass. For the T-5 GOCE program, Astrium-CRISA is responsible for the Ion Propulsion Control Unit (IPCU) and Bradford Engineering for the Proportional Xenon Feed Assembly (PXFA). The challenging thrust control requirements for GOCE have resulted in considerable development work for power and propellant feed systems.

The ITA is controlled using a set of algorithms, specifically designed to meet the GOCE requirements. The control algorithm architecture is shown below. The ITA uses three of its input parameters to control the output thrust. The flow rate and anode current are adjusted relatively slowly, in an open-loop mode, to provide a coarse control of the thrust, while the solenoid magnet current is adjusted quickly, in a closed-loop mode, to provide fine control for high accuracy and quick response. The IPCU has been designed specifically to meet the GOCE requirements, utilizing EADS Astrium-Crisa’s extensive experience in space power equipments.

Figure 83 Control Algorithm Schematic




11.6.3.2 Proportional Xenon Feed Assembly (PXFA) The Proportional Xenon Feed Assembly provides accurately metered Xenon propellant flow to three separate flow components within the ITA. The PXFA is split into pressure control and flow control sections. The pressure control section receives Xenon from the tank at high pressure (between 5 – 125 bar) and regulates it down to a pressure of exactly 2.50 ± 0.05 bar. This section also includes particle filters, pressure transducers for monitoring the tank and supply pressures with isolation valves to allow switching between primary and redundant units. The flow control section receives the regulated pressure Xenon and provides accurately metered flow to the ITA main flow, cathode and neutraliser. The cathode and neutralizer flows are metered to fixed levels using advanced viscosity/thermal controlled passive flow restrictors, but the main flow must be varied over a wide range to allow the thruster to be throttled efficiently. This variable flow function is provided by a combined flow control valve and flow sensor system, which allows the flow sensor output to be used for closed loop control of the flow control valve.

Figure 8 PXFA for GOCE (left) and T-5 Thruster Unit (right) 4




11.6.4 PERFORMANCE

9

Remarks

Table 8 Performance table for the T5 thruster Unit

Performance

Nominal

Demonstrated

Thrust 18mN

1-20mN

Nominal-Artemis Mission Demonstrated – GOCE Mission

Predicted minimum level for T-5 thruster, optimised for maximum capability of 6mN is 200µN

5% of thrust set point. 50µN @ 1mN

Specific Impulse 3300s 500-3500s Across GOCE thrust range

Total Impulse >1.5 E6 N.s (GOCE Mission)

Under GOCE continuous throttling conditions. Predictions indicate capability up to 2.0 E6 N.s

Cycles >8500 Cycles Generic T-5 Capability

<±0.1º

<12º

Thrust Resolution GOCE requirement

Thrust Vector Stability GOCE requirement

Beam Divergence <25º 2 σ half cone angle at 1mN 2 σ half cone angle at 20mN




11.6.4.1 Interface and Physical Resource Requirements Table 90 Interface requirements for the T5 thruster Unit

Value Remarks


190 x 190 x 242

PXFA 150 x 200 x 250 GOCE T-5 System

Mass Budget (kg) EIT 2.95

IPCU 16.7 PXFA 7.5

GOCE T-5 System

Dimensions (mm) EIT

IPCU 300 x 250 x 200

Power (W) Thruster 55-585 Across GOCE thrust range

Thermal interfaces Conducted Dissipation <10W

Other Exported vibration level (µN) <1.1 E-6 m/s2 / √Hz PXFA Levels

The Critical Design Review of the GOCE Ion Thruster Assembly will be held during the second quarter of 2005. The challenges of stringent mission requirements in terms of thrust range and control have been achieved, whilst maintaining a strong heritage from previous designs and the implementation of lessons from the ARTEMIS experience.

Some of the more difficult aspects of program development have been related to demonstrating the thruster performances during ground testing, particularly with respect to thrust measurement. Fortunately, Ion thrusters benefit from an inherent ability to calculate actual thrust levels using measured electrical parameters during operation, the ion trajectories and their evolution over life are provided via modelling techniques and have been validated for a fixed thrust level on ARTEMIS. With demonstrated scalability, it is expected that development of Ion Thrusters for very low thrust applications is achievable. The biggest challenges may result from validating the “on-ground” performances of absolute thrust and thrust noise. It is anticipated that the GOCE mission will allow for further refinement through validation of both calculated and measured thruster performances this important for any future scaling of the design.


With most of the GOCE mission requirements demonstrated during ground test testing there is strong confidence of the generic Kaufman thruster being able to achieve even lower thrust levels, finer controllability and thrust resolution. However the impact of long life requirements and potential for retaining homogeneity of the discharge plasma requires specific investigation.




11.7. RMT mN ion thrusters Technology Development


The Radiofrequency Ion Thruster with applied Magnetic field (RMT) is an Ion thruster fine throttable in the milliNewton range (0-12 mN) with resolution less than 0.1 mN. This thruster design features a high efficiency (overall power consumption less than 500 W at max thrust) and high specific impulse (2200-3600 s). The addressed applications of the RMT system on platforms of the 300-1000 kg class are:

-Drag Compensation/ "Drag free" control of satellites orbiting in the altitude range 300-600 km,

- Orbit transfer (e.g. in the 300-600 km altitude range); - End-of-life orbit disposal; - Station keeping of GEO satellites. - General AOCS Tasks

The RMT has been developed in the frame of a R&D Contract awarded by the Italian Space Agency for application of Drag Free or Station Keeping tasks for Satellite platforms in the class of 300-1000 kg or for Attitude Control tasks of Scientific Satellites.


The Propulsion system is composed by the following sub-assemblies:

- Thruster Module (TM), - Gas feed line & Flow Control Unit (GFCU) including the High pressure regulator

control valve and low pressure thermo throttle flow chain - Radiofrequency Generator & Matching network (RFGM) - Power Supply & Control Unit (PSCU), - Xe Tank

The current development state of the various RMT sub-assemblies is presented in Table 91 The RMT thrust can be adjusted within the 0-12 mN range, while maintaining the maximum efficiency, by varying the coil power, the RF power, the xenon mass flow rate. Either the coil or the RF power, alone, can be varied to achieve fast (> 1 Hz) thrust variations around the selected thrust level. The specific impulse can be adjusted, according to the mission optimization criteria, by controlling the beam voltage and the propellant utilization efficiency. The RMT is well suited to be an actuator of an integrated system for the autonomous navigation in which GPS receivers are used as position and velocity sensors.




Table 91 Current development state of the various RMT sub-assemblies

RMT Sub-assembly State of development Comments TM (Thruster module) Engineering Model neutralizer already qualified and flown on

ARTEMIS mission RFGM Engineering Model GFCU Engineering Model The new GFCU include the Proportional

Control Valve PSCU (power supplies) elegant BB model PSCU (Control Unit) laboratory model

11.7.3 DESCRIPTION

This technology uses a RadioFrequency discharge, excited in the VHF range, in conjunction with a low level (∼ 100 Gauss) static magnetic field to establish a cold Xenon (or Ar or Kr) plasma. The excitation of resonance phenomena in the plasma is exploited to enhance the ionization process in the low gas flow rates regime. The design guidelines are: Use of external RF excitation to avoid the presence of critical components (e.g. the cathode

in the Kaufman thruster) in the interior of the discharge chamber; Avoid loss of efficiency when working at thrust level in the mN range Long operational lifetime (> 20,000 hrs in orbit) Operation in "continuous" and/or in "cycled" mode,

Figure 85 shows the RMT in operation inside vacuum chamber.

Figure 85 RMT operating in vacuum chamber

The RMT cross-section is shown in the left part of Figure 86. The baseline architecture of the RMT system, named RMTA (RMT Assembly) and including the thruster plus the auxiliary equipment, is shown in the right side of Figure 86.




PROPELLANTELECTRICALCONNECTOR PROPELLANT

COIL

PERMMAGAS

DISTRIBUTOR

NEUTRALIZER

GFCURegul. Xe flow

Insulator

Anode/Gas distrib.

Discharge ChamberPermanent Magnet Coil Screen Grid

Accel. GridDecel. Grid

Neutralizer

Keeper

Insulator

PA

+ -PC

+-PK

+ -PH PG

PSCU

S/C power busS/C Data link

command / telemetry line

powerline

RFGM

RF Antennacouplingsystem

Power I/FSC

Xe highpressure

inlet

PRF

TM

Sketch of the RMT cross-section RMTA functional block diagram

Figure 86 RMT cross section (left) and functional block diagram (right)

The RMTA includes the following sub-assemblies: a) Thruster Module (TM), composed by the following parts:

Discharge chamber with externally mounted RF (radiofrequency) coupling electrodes Gridded ion optics Magnetic circuit Anode/gas distributor Thruster case Neutralizer

b) Gas feedline & Flow Control Unit (GFCU), composed by the following parts:

Pressure reduction/Gas flow regulator towards the discharge chamber and the neutralizer

Command/Control Electronics c) Radiofrequency Generator & Matching network (RFGM), composed by the following parts:

Radiofrequency Generator in the VHF range (∼ 150 MHz) Matching Network to optimize the RF coupling to the plasma

d) Power Supply & Control Unit (PSCU), composed by the following parts:

Anode Power Supply (PA)

Coil Power Supply (PC) Grid Power Supply (PG)




Neutralizer Keeper Power Supply (PK) Neutralizer Heater Power Supply (PH) RFGM main Power Supply (PRF) Interface to the S/C Power Bus (ref. standard 28 V)

e) System Controller (SC), including the TLM/TLC interfaces to/from the S/C data handling

11.7.4 PERFORMANCE

Table 92 Measured performance of the RMT

Performance

Required

Measured/ Expected

Remarks

Thrust 0-12 mN beam voltage 1.5 kV 500 hours lifetime accumulated

Thrust Noise TBC

Specific Impulse 2200 @ 2 mN – 3600

@ 12 mN

a new very high specific impulse upgrade model is on going process. Beam Voltage 5000 v – Isp > 7000sec. 500 hours testing planned end of 2005

Flow Rate 0.1-0.4 mg/s of Xe varying according to thrust range

Total Efficiency 0.3-0.5





Table 93 Interfaces and resource requirements of the RMT

RMT sub-assemblies Dimensions (mm) Mass (kg) TM (Thruster module) ∅ 150 x 130 1.6 (including the neutralizer) GFCU 150 x 100 x 100 (TBC) < 1 (including the high

pressure stage based on proportional control valve )

RFGM 100 x 150 x 100 1.3 PSCU 250 x 250 x 185 (TBC) 4.8 (TBD) Harness & Piping 1 Total (tank & fill/drain valve excluded)

TBD

Power Average consumption (W) 50-500 Peak consumption (W) TBC

Thermal interfaces Operational temperature (K) Delta T: 140K Operational temp. range (K) 223°K - 573°K


The development of the RMT was terminated at EM level and no flight or QM process is on-going. An improved model with 7000 sec Isp is on going development and testing by the end of 2005 under ESA TRP Contract.




11.8. ALTA mHall Thruster

11.8.1 EXECUTIVE SUMMARY

During 2003 ALTA S.p.A. started the development of a laboratory model of a small Hall Effect Thruster with a nominal power of 100 W to cover the needs for low power, low cost propulsion system for the microsatellite market. In late 2004, following a successful preliminary test campaign, the development activity was aimed to the design of an engineering model of the unit under an ESA SME contract (No. 18731/04/NL/DC “Low Power, Small Hall Effect Propulsion System”). In this phase the development of a dedicated PSCU was also initiated involving in the same contract another italian company SkyTech Electronic Technolgies in charge of the design of a breadboard model. In 2005 efforts have been spent also in a preliminary design of the XFCU architecture. The SME contract includes mechanical tests (mechanical properties, sine and random vibration) and as a minimum an endurance test of 60 hrs at the end of a characterisation phase for a total operating time of 100 hrs. All the characterisation and endurance tests will be carried out with coupling with the PSCU model. An option has recently been approved to the programme to run two thruster units in parallel to assess the effects of cluster configurations. The output of the programme, expected to be completed in the first quarter of 2006 will provide information for the Qualification Programme of the subsytem thruster + PSCU.


The EPS subsystem is currently designed to operate in parallel from 1 up to 4 Hall thrusters with a nominal power of 100 W, thrust of 4 mN and Isp greater than 900 s and a throttling range between 50-200 W with thrust levels between 2 and 8 mN. The system is designed in order to have great flexibility and high modularity to be easily scaled and optimised for different mission scenarios and requirements.

Figure 87 XHT-100 Hall Thruster Laboratory Model during assembly (left) and testing (right)




In its base configuration the EPS comprises:

- up to 4 Hall Thrusters (Anode Unit + Neutralizer)

- No. 1 Power Supply Control System (PSCU)

- No. 1 Xenon Flow Control System (XFCU)

- No. 1 Xenon Storage Tank (XST)

In addition a Thruster Orientation Mechanism (TOM) and a Plasma Diagnostic Package (PDP) can be included in the system.

Logic Unit DC/DCConverter(s)

PM1

A PSU

TCL PSU

HIK1 PSU

HIK2 PSU

PM2

A PSU

TCL PSU

HIK1 PSU

HIK2 PSU

PM3

A PSU

TCL PSU

HIK1 PSU

HIK2 PSU

PM4

A PSU

TCL PSU

HIK1 PSU

HIK2 PSU

EPS Control(optional)

S/CComputer

S/C BusPower

Thruster 1 Thruster 4Thruster 3Thruster 2

Control Lines

TC Lines

PSCU

Figure 88 System (Thrusters+PSCU) Layout

The EPS is designed to operate with 28 V unregulated power from the satellite bus and no low- level control logic is implemented within the subsystem (i.e. all the settings for the EPS operating sequences are directly provided by the OBC via serial RS422 link) in order to reduce the system complexity. However for specific mission requirements the PSCU is designed to accomodate with minimum changes a Propulsion System Managing Unit (PSMU) to directly control the EPS operations requiring only high-level enable/disable commands.

Power Lines




11.8.3 DESCRIPTION

System design is currently focused on three main activities:

- optimisation of the Hall Thruster performance and configuration - development of the PSCU - development of the XFCU

Designs of the XST and of the PSMU are currently in stand-by waiting for specific requirements from the mission. The Hall Thrusters are based on the XHT-100 laboratory demonstrator designed and tested by Alta in 2004 in order to optimise its performance, remove some design criticalities (especially relaed to high voltage insulation) and to preliminarly qualify some of the technical solutions adopted. In particular a thermal endurance test has started in mid July 2005 to characterise the long term behaviour of the permanent magnets which are the key components for the thruster functionality. A parallel activity has also started for the development of a dedicated thrust stand (10 mN f.s., 10% accuracy) to directly and accurately characterise the thruster performance (thrust, Isp, efficiency). The thruster will have a single hollow cathode neutralizer to minimise the system complexity and mass budget, but the PSCU will have the possibility to include the power board for a redundant neutralizer unit if higher levels of reliability are required. The design of the PSCU is currently carried out by SkyTech Electronic Technologies and it’s aimed to the development of an EM breadboard in the version able to run two thrusters in parallel for an integration test to be carried out running on a cluster configuration. A single unit for a multiple thruster configuration was selected to reduce the system overall mass budget: the power modules will be modular and only the external case will be optimised for each flight mission. The XFCU design is still at a very preliminary stage and only the functional requirements have been identified. The work is now focused on the identification of the best system configuration since two different technologies are available for pressure reduction leading to different designs. The first option is based on bang-bang pressure regulators whereas the latter is based upon multifunction valves operating as pressure reducers and mass flow controllers. The baseline configuration, envisaged for microsatellites, includes a single XFCU feeding up to 4 thruster, but different solutions with distributed units (one for thruster) are still open. In any case the design will include a single MFC for anode and neutralizer with the two mass flow rates regulated at a fixed ratio by calibrated orifices to reduce system complexity and demand for telemetry/control lines.




Figure 89 Layout of XFCU

Figure 90 Thruster Performance




11.8.4 PERFORMANCE

Table 94 Main expected performance of the mHall thruster

Performance Nominal Demonstrated Remarks

Thrust 4 mN Calculated by eletrical parameters

Specific Impulse > 950 s Calculated by eletrical parameters

Total Isp: including anode and cathode mass flow rates

Thrust Efficiency > 20% Calculated by eletrical parameters Including total power consumption

Flow Rate 0.48 mg/s Total: including anode and cathode supply

Total Impulse >3.5.104 N s Estimated On Time > 2000 hrs Target Cycles >1000 Target





Table 95 Interfaces of the mHall thruster unit

Interface requirement Value Remarks Mechanical

Thruster Diameter (mm) 70

Thruster length (mm) 53 (80)

- (Including supporting structure)

Thruster Envelope (mm) 100 x 100 x 100 Including neutralizer Weight (g) 710 Including neutralizer

Weight after optimization (g) <600

Accommodation requirements TBD

The thruster will include a proper mounting interface for thermal

insulation with bolt pattern (TBD) for installation

Power


Inlet pressure (mbar)

100 @ thruster in nominal power mode

Peak consumption (W) 250 @ thruster in enhanced power mode

PSCU Efficiency >70% Propellant

Propellant Xe purity > grade4.8 100-1500

Thermal interfaces Operational temperature (K) 473 @ thruster TRP Operational temp. range (K) From 223 to 573 @ thruster TRP

Flux to/from S/C (W) < 10 W





Hall Thruster is rather robust technology and no major issues have to be pointed out, in any case the most important aspects of the thruster development can be summarised in:

- detailed thermo-mechanical analysis to determine the level of mechanical stresses into the ceramic discharge chamber to avoid cracking and formation of chips

- detailed thermo-magnetic analysis and test to ensure the durability of the permanent magnets at high temperature during lifetime at the proper magnetic induction level

- test of different ceramic materials (pure Boron-Nitride, BN-SiO2 compounds, BN-Si3N4 compounds etc.) to assess the erosion properties of different material under real operating conditions

- prediction of the thruster lifetime through appropriate models and experimental data

- verification of the electrical insulation durability along the operating life without appreciable degradation of the performance

- optimisation of the neutraliser position to reduce the voltage losses due unefficient coupling between electron source and magnetic field.

11.9. RIT µN ion thrusters Technology Development


Following the request of several scientific missions, asking for a thrust range in the order of µN level, a study on the process of scaling down the size of the Radiofrequency Ion-Thruster (RIT) has been started in Giessen University, Germany. Two Laboratory Prototype Thrusters, the µN-RIT4 and the µN-RIT2 have been already manufactured and are actually under test in Giessen University. The advantages of the mini Radiofrequency Ion-thruster, besides their low thrust range, will be their controllability and their save of mass in long-term missions due to their reduced size and the higher specific impulse. Moreover the using of inert gas, like Xenon, will avoid possible problems of Spacecraft contamination. Furthermore within the class of gas-discharge electric thruster, the RF-type seems to be the most suitable for scaling down, because it works without any discharge electrodes, magnetic pole inside the ionizer. The absence of discharge electrodes will also allow long lifetime. Another advantage of the RF-discharge is its regulation by varying gas flow, RF-power and extraction voltages independently.

The actual lowest thrust limit achieved by the µN-RIT4 Laboratory Prototype, according with the document UNI-GI-DOC-200503, is 70µN, limited by the capability of the test equipment. The Giessen University prevision is to achieve lower thrust and higher controllability, using an Automated Beam Current Control Unit with an integrated high precision power supply. A test in the ESA Propulsion Laboratory is in plan to verify the excepted thrust range (1-100µN).





The RIT Thruster generates thrust by accelerating Xenon ions out of plasma induced by RF-energy. A Flow Control Unit (FCU) delivers the propellant, Xenon, to the thruster and to the neutralizer. In order to prevent electric charging of the equipment and the spacecraft, the neutralizer injects electrons into the beam of ions.

Figure 91 µN-RIT4 Laboratory Prototype Drawing

Xenon for the thruster flows through the feed line interface via the isolator and the extraction anode, which also functions as gas distributor, into the discharge vessel of the thruster. The discharge vessel, made from quartz, is surrounded by an induction coil, which is connected to an RF-generator. The latter generates a high frequency (MHz) electrical field in the discharge vessel. Free electrons, within the xenon gas, collect energy from the induced electric field and ionize the neutral propellant atoms by inelastic collision. The neutralizer starts first to initiate the main discharge. The entire neutralizer includes an internal cathode and the neutralizer itself. The cathode, heated, produces electrons by thermo-ionic emission. These electrons are emitted into the neutralizer, which is filled with low-pressure xenon. Plasma is generated in the neutralizer tip by a low voltage arc discharge between cathode and neutralizer. Electrons from the neutralizer are drawn into the discharge chamber thanks to an electrostatic field generated by the voltage applied between the acceleration electrode and to the anode. After the ignition, thrust is generated by acceleration of ion in the electrostatic field applied to a system, which include an extraction anode, an isolating plasma holder, an acceleration electrode and a deceleration electrode. The thruster itself needs for its operation a Positive High Voltage, a Negative High Voltage and an RF-Power. Thrust level variation can be achieved by variation of the RF power at constant exhaust velocity.




11.9.3 DESCRIPTION

The development of the µN-RIT System started using a scaling down process from the well-known RIT10 geometry and performances. In the process of scaling down the size of the thruster and, consequentially, the thrust range, all design parameters and materials of the larger thruster were adopted. This concerning the ionizer/xenon injection system, the neutralizer, and the grid system, which is unchanged with respect to the materials (carbon composite), the thickness, the interspaces, the beam-let holes size and the grid voltage. The scaling down process takes also into account the value of the discharge parameters like the discharge vessel radius (R), the discharge pressure (P), and the radiofrequency (ν). Following the relation P~ ν~1/R the thrusters are manufactured. Moreover the development of the electronic subsystems is based on ARTEMIS qualification and flight model. Scaling down of electronic components has very low risk, as only low currents are needed. The Xenon flow control system is based on ALFA-BUS pre-developments (good potential to achieve low flow rates). The µN-RIT4 Laboratory Prototype has been manufactured by Giessen University and tested in the Big Mac Facility. It is equipped with a 19-hole grid system. The discharge vessel is made from quartz. The RF-coil has 6 turns.

Figure 92 µN-RIT4 Laboratory Prototype




During the test of the µN-RIT4 Laboratory Prototype, the second scaling down step has been started. The geometry of each beam-let electrode is identical with the µN-RIT4 Laboratory Prototype concept. The RF-coil of the µN-RIT2 has only 4 turns due to the envisaged higher frequency.

Figure 93 µN-RIT2 Laboratory Prototype




11.9.4 PERFORMANCE

Table 96:Comparison of required and measured (or expected) performance of the µNRIT-4 Laboratory Prototype

Performance

Required

Measured/ Expected

Remarks

Thrust 1-100µΝ

70µΝ @ constant beam voltage of 1400V measured

/1-100µΝ expected

A Test in the ESA Propulsion Laboratory is planned to verify the range required.

Thrust Noise <1µΝ/Ηz1/2 TBC

The Test Power Supply used by Giessen University was built to operate the ARTEMIS RIT10 (higher thrust level). A Test in the ESA Propulsion Laboratory will use equipments conform to the thrust level request.

Specific Impulse 3850s@230µΝ and

beam voltage of 1400V

Higher specific impulse could be reached with a beam voltage of 1700V. TBC

Flow Rate 0.006mg/s During the test in Giessen University the flow rate was kept constant.

Mass Efficiency Up to 84%

And down to 26% @lowest thrust level

The 26% is due to the impossibility to turn down the gas flow under 0.006mg/s with the used Giessen equipment according to UNI-GI-DOC-200503.





Following the interface requirements for the µNRIT-4 Laboratory Prototype and the µNRIT-2 Laboratory Prototype. Table 97: µNRIT-4 Laboratory Prototype interface requirements


Value Remarks

Thruster Diameter (mm)

83 Thruster length (mm) 70

Weight (g) 600 Weight after optimization (g) 400

Accommodation requirements TBC

Power Average consumption (W) 15-20

Peak consumption (W) TBC Thermal interfaces

Operational temperature (K) Delta T: 140K Operational temp. range (K) 223°K - 573°K

Optical interfaces

Other Exported vibration level (µN) Not vibration expected TBC




Table 98 µNRIT-2 Laboratory Prototype interface requirements


The development of the two µNRIT Thrusters is still in process. The two thrusters are prototype models. More details about the status of this technology will be available after the test in the ESA Propulsion Laboratory. During this test the Thrust range will be confirmed/refused.


Value

Remarks

Thruster Diameter (mm) 60 Thruster length (mm) 70

Weight (g) 350 Weight after optimization (g) 300

Accommodation requirements TBC

Power Average consumption (W) 10-15

Peak consumption (W) TBC Thermal interfaces

Operational temperature (K) Delta T: 140K Operational temp. Range (K) 223°K - 573°K

Optical interfaces

Other Exported vibration level (µN) Not vibration expected TBC

indium propellant is heated above its melting point of 156°C, and flows along a needle or capillary towards a sharp tip. By applying a high electric field between the tip and an extractor

11.10. In µN FEEP thrusters Technology Development


The Indium Field Electric Emission Propulsors (In FEEP) thruster can provide µN level thrusts with extremely low thrust noise and stability. The thruster is scalable to the required maximum thrust requirements for Darwin as it consists of cluster elements which can each deliver up to 10 µN of thrust, that can be simply connected together.

11.10.2 DESCRIPTION

The thruster is based on the space-proven ARCS Indium Liquid-Metal-Ion-Source technology. This technology is the core of active spacecraft potential control devices of mass spectrometer instruments flying on a number of satellites including CLUSTER-II, ROSETTA or DOUBLESTAR (accumulated > 6.600 h operation in space). With reference to Figure 91,




electrode, ions are directly field ionised and accelerated in the same field. Each emitter can safely produce up to 100 µA (with peaks up to several hundred µA) that translates into thrusts up to 10 µN. Depending on the acceleration voltage, the specific impulse ranges roughly between 4000 and 8000 s. Several elements can be clustered to obtain higher maximum thrust levels. The thruster has undergone a lifetime test of 3800 h, which is a record for FEEP thrusters. This allowed verifying available lifetime models as well to predict a total lifetime greater than 25.000 h. In order to counterbalance spacecraft charging, a low-power neutraliser is necessary to emit an electron current. Several options have emerged in this field ranging from microfabricated field emission cathode arrays (FEA), carbon nanotube or thermionic cathodes. The thermionic cathodes are flight-proven and tested for a long lifetime of more than 100.000 h. However, field emission/carbon nanotube neutralisers promise both power and volume/mass reduction. Lifetimes in excess of 10.000 h have been demonstrated already for carbon nanotube neutralisers.

Figure From Left to Right: Single Emitter (10 µN) breadboard including thermionic neutraliser and miniaturized HV electronics, 3x3 Cluster, 4x4 Cluster (demonstration for LISA PF)

The FEEP thruster itself has been subject to miniaturization and is presently being investigated/developed for the mN thrust range. A recent prototype from ARCS is shown in Figure 94 featuring a 20x20 array on an area of 5x5 mm2. The development aims at providing maximum thrust levels of several mN with the same thrust properties (low thrust noise, resolution, etc.) as in the µN thrust range.

94




Figure 95 Microstructured FEEP Thruster Prototype

From the present status, the indium FEEP thruster is in an advanced breadboard/early qualification model stage for thrust levels up to 150 µN. There is an established industrial consortium around ARCS/MAGNA/AAe/ASTRIUM to supply the full FEEP subsystem within this thrust range. The µFEEP thruster would offer a thrust extension into the mN thrust range together with a significant mass and volume reduction. mN thruster prototypes are presently under development (EO TRP) and first prototypes should be available in late 2007. Qualification and lifetime testing for this technology could be demonstrated before 2010.




11.10.3 PERFORMANCE

Required Demonstrated Expected Comments Minimum Thrust

0.1 µN Using laboratory power supplies without optimization.

Maximum Thrust

65 µN / emitter 5 mN per

emitter

Maximum thrust involves clustering of individual elements. µFEEP thrust range of 5 mN extrapolated from demonstrated array density.

Thrust Resolution

50 nN Using laboratory power supplies without optimization.

Specific Impulse

4000-8000 s Variable µFEEP will offer variable specific impulse up to 12.000 s (higher if necessary).

Beam Divergence

< 25° Using focusing electrodes, otherwise < 55°.

Total Impulse 490 Ns / emitter Derived from present reservoir size of 15 g (extrapolated using verified lifetime models). 205 Ns demonstrated in lifetime test. Higher values using clustering. µFEEP reservoir size might be even larger.

Thrust Noise < 0.1 µN/Hz0.5 Using present flight HV electronics.


TBD


µFEEP technology has to be demonstrated in running contract.




11.11. ALTA µN FEEP thrusters Technology Development


FEEP microthusters provide the following features:

- sub-µN thrust range, near instantaneous switch on/switch off capability, - fast dynamic response which enables accurate thrust modulation in both continuous

and pulsed modes with high resolution (0.1 µN). - high specific impulse (from 5000 up to 7500 s) e.g. very low consumption of the

propellant Presently baselined for the scientific missions satellites, this propulsion system has been also proposed for attitude control and orbit maintenance on commercial small satellites and constellations. The FEEP has been developed in the frame of ESA R&D Contracts and was recently selected for the µSCOPE mission and proposed for LISA PATHFINDER and GAIA Four Flight Units (EPSA) will be manufactured and flight qualified for the AOCS tasks of the µSCOPE mission. Each EPSA, is made up of 3 types of units: the FEEP Thruster Assemblies (TA), the Power Processing and Control Unit (PPCU), the Neutralizers (NA), and the Supporting Structure. The Thruster Assembly (TA) is a single-emitter module, formed by the emitter unit, the propellant reservoir and related insulating parts, heaters and sealed container, and electrical interface connectors. The Power Processing and Control Unit is made up by the power processing and interface assembly, data processing and interface assembly, thrusters and neutralizer control, HV

• self-protection from contamination during ground handling, storage and transportation;

assemblies, interface connectors to the thrusters, neutralizers and spacecraft, harness and support structure. The Neutralizer assembly (NA) is constituted by an electron-emitting device and related insulating and supporting parts. The current development state of EPSA and relevant items is at Engineering Level. Functional and Endurance test at ESTEC have been planned starting from mid of 2005. The Propulsion System CDR is planned by the end of 2005


The following functions are provided by the Thruster Assembly (TA): • thrust • propellant storage and feeding; • electrical insulation of the thruster electrodes; • grounding connection;




• heaters for active thermal control during initialization, cold start and normal operation; • mounting interfaces; • carrying the loads deriving from the mechanical environment Each TA is a stand-alone device, and can be commanded and controlled separately from other TAs.

11.11.3 DESCRIPTION

Field Emission Electric Propulsion (FEEP) is an electrostatic propulsion concept based on field ionization of a liquid metal and subsequent acceleration of the ions by a strong electric field. This type of thruster can accelerate a large number of different liquid metals or alloys. The best performance (in terms of thrust efficiency and power-to-thrust ratio) can be obtained using high atomic weight alkali metals, such as Cesium and Rubidium (133 amu for Cs, 85.5 amu for Rb). These propellants have a low ionization potential (3.87 eV for Cs and 4.16 eV for Rb), low melting point (28.7 o C for Cs and 38.9 °C for Rb) and very good wetting capabilities. These features lead to low power losses due to ionization and heating and the capability to use capillary forces for feeding purposes (i.e. no pressurised tanks nor valves are required). Moreover, alkali metals have the lowest attitude to form ionized droplets or multiply-charged ions, thus leading to the best attainable mass efficiency.

Figure 96 FEEP thruster concept




Figure 97 FEEP thruster under operation

An accelerating electrode (accelerator) is placed directly in front of the emitter. The emitter consists of two identical halves made from stainless steel, and clamped or screwed together. A nickel layer, sputter deposited onto one of the emitter halves, outlines the desired channel contour and determines channel height (a.k.a. slit height, typically 1 - 2 mm) and channel width (a.k.a. slit length, ranging from 1 mm up to about 7 cm). When thrust is required, a strong electric field is generated by the application of a high voltage difference between the emitter and the accelerator. Under this condition, the free surface of the liquid metal enters a regime of local instability, due to the combined effects of the electrostatic force and the surface tension. A series of protruding cusps, or “Taylor cones” are thus created. When the electric field reaches a value in the order of 10 9

V/m, the atoms at the tip of the cusps spontaneously ionize and an ion jet is extracted by the electric field, while the electrons are rejected in the bulk of the liquid. An external source of electrons (neutralizer) provides negative charges to maintain global electrical neutrality of the thruster assembly. The essential elements of the FEEP thruster are shown in Figure 98.

Figure 98 Schematic diagram of FEEP thruster

A stream of ions is liberated from the liquid metal meniscus in the emitter slit, by the field emission principle described in the previous section. The beam is then accelerated by the same




electric field established between the positive emitter and the negative accelerating electrode (accelerator). The beam exits the accelerator with a velocity determined by the total potential drop between the emitter and the accelerator, and by the charge-to-mass ratio of the ions employed (see Figure 99). After exiting the accelerator, the beam is decelerated by a potential rise, being the outer shield of the thruster usually grounded to the S/C potential. Subsequently, a stream of electrons joins the ion stream, producing a beam of zero net charge. The neutralizer also provides overall charge compensation to the S/C, keeping the S/C potential close to equilibrium with ambient plasma. The potential rise beyond the accelerator has two main effects: • it slows the exiting ions, therefore reducing Isp and power-to-thrust ratio (at first instance, power-to-thrust ratio is proportional to exit velocity); • it prevents free electrons from entering the accelerator-emitter gap, where they would be accelerated towards the emitter, with several negative drawbacks (emitter tip would be heated by electron bombardment, with ensuing increase in propellant evaporation).

Figure 99 Potential and velocity profiles for the FEEP thrusters




As it can be seen by the above figures, the final exit velocity (ve), and therefore Isp, is ultimately determined by the potential at the emitter electrode (emitter voltage, Ve). On the other hand, for the extraction mechanism the mass flow is determined by the electric field at the emitter tip, which is proportional to the total voltage difference, i.e. Ve-Va. (Va accelerator Voltage). As the thrust is proportional to both exit velocity and mass flow, operational parameters of FEEP (e.g. power-to-thrust ratio, Isp vs. thrust curve) can be modified to some extent by properly adjusting the voltage values. The Thruster Assembly (TA) designed for the µSCOPE program is depicted in Figure 100.

Figure 100 Thruster Assembly designed for the µSCOPE program




11.11.4 PERFORMANCE

Table 99 Performance of the ALTA µN FEEP thrusters

Performance

Units

Remarks

Thrust

0.1-150 µN

1500 hours lifetime accumulated

Thrust Direction In-plane divergence _ 15° Out-of-plane divergence _ 40°

The divergence semi-angles are defined such that the beam portion comprised therein includes at least 95% of the emitted current

Thrust Noise

For 10-4 [Hz] < f < 10-1 [Hz] Log(PSD) = 0.01 ⋅ (0.1 / f) [µN2/Hz] For 10-1 [Hz] < f < 1 [Hz] Log(PSD) = 0.01 [µN2/Hz]

Thrust resolution 0.1 µN for T<100 µN 0.3 µN for T>100 µN

Instantaneous Thrust direction

0.5 degrees (3-sigma value, uncertainty cone half angle).

Total Impulse 3100 Ns 4650 Ns with qualif margin

Specific Impulse

5400-5900-6400s @ 150 µN with - 4,-3,-2kV

6900-7400-7850s @ 30 µN with - 4 ,-3,-2kV

Specific Power

57-61-64 W/mN @ 150 µN with - 4,-3, -2kV

53-57-60 W/mN @ 30 µN with - 4,-3, -2kV

Total response time

Thrust step size

Maximum response time

1 µN <100 ms 10 µN <160 ms 30 µN <180 ms

150 µN <340 ms

Voltages Min 7 kV Max 15 kV





Table 100 Performance of the ALTA µN FEEP thrusters

FEEP System Dimensions (mm) Mass (g) EPSA module 410 (Xepsa) x 265 (Yepsa) x 280

(Zepsa) mm 9100 (3 TA, PPU, 2 NA)

EPSA baseplate 340 (Xepsa) x 225(Yepsa) TA only 94 (Xepsa) x 195 (Yepsa) x 135

(Zepsa) mm 1250

Power


53,4 W

25,3 W

3 thrusters ON , 150 µN each +1 neutralizer

3 thrusters ON , 30 µN each +1 neutralizer

Thermal interfaces

Mounting I/F range (C) - 10C + 50C S/C Operating - 35C + 60C S/C Non

Operating


The development of the RMT was terminated at EM level and no flight or QM process is on-going. An improved model with 7000 sec Isp is on going development and testing by the end of 2005 under ESA TRP Contract.




11.12. Guidance, Navigation and Control (ICC) Technology Development


This note constitutes the summary of the activity ``Interferometer Constellation Control'' (ICC), performed under ESA contract 15522/01/NL/FM. The consortium selected for the study was Astrium SAS with GMV (modelling and simulator), Astrium UK (open loop guidance function) and SciSys (autonomy, FDIR and Operations) as sub contractors. The activity deals with the design of the Guidance, Navigation and Control (GNC) system for the 3 fundamental modes of the Darwin interferometer, namely the Baseline Control Mode (BCM), the Fringe Acquisition Mode (FAM) and the Nominal Mode (NOM). The GNC system makes use of the micro Newton propulsion system and the Optical Delay Line (ODL) actuator as actuators. The sensors involved are the RF system for the coarse modes, and the high precision optical metrology (HPOM) systems for the other modes. The GNC must control all relative position and attitudes between spacecraft, as well as the delay line stroke for equalizing the optical paths of the beams at combination to within the coherence length. A GNC design tool as well as a simulator has been developed including a Monte Carlo capability for statistical analysis. The Darwin formation used in the study is the bow-tie layout of the 6 telescope spacecraft, with the beam combiner in the center, and a master spacecraft out of the plane of the interferometer itself.


The overall functionality of the GNC is to control the relative positions and relative attitudes of the s/c by use of the micro Newton propulsion system, on the basis of the measurements provided by the metrology system. In addition, the GNC design integrates directly the control of the Optical Delay Line by means of a faster inner control loop. All the couplings between the systems are taken into account directly in the overall combined design. The GNC system features a state estimator of all the 84 states in the system. There is a guidance function providing the relevant reference signals for station keeping as well as position and velocity reference signals for dynamic array manouvers. Additionally, the guidance provides the feed forward forces and torques required to perform the maneuvers. The state controller performs the function of reducing the control error to zero and to minimize the noise transmission to the platform. On top of the GNC, a state machine is designed and implemented performing mode management, and handling the criteria upon which the system leaves one mode and enters the next. Besides the 3 modes described earlier, Failure Detection Identification and Recovery (FDIR) system is also included, handling various submodes related to failures of various redundant avionics equipment as well as collision avoidance maneuvers.




11.12.3 DESCRIPTION

The development is kept largely generic for the on board system being expandable to a generic number of spacecraft. The performance requirements as discussed below are such to meet the Darwin requirements, as well as a 3 spacecraft demonstration mission with 2 telescopes and a beam combiner. The ICC activity makes use of a specific set of coordinate frames, rotation sequences and notation, which has been developed to harmonize with all relevant activities in the technology developments for Darwin. The environment at L2 has been described in terms of perturbation models, as well as by carefully modeling the spacecraft kinematics and dynamics properties. Flexible elements of the structural modes of the sun shield and solar arrays has been included. While no specific propulsion architecture has been implemented, engine characteristics are modeled as well as the typical cross couplings for such spacecraft. Interfaces to the other avionics subsystems have been defined, and their performance characteristics have been used as input for the GNC design wrt the performance requirements. The architecture for such distributed system is then analyzed for a central and a decentralized architecture. A trade-off between these two options has led to the selection of a central structure for the GNC system. For the deployment modes there might be other drivers leading to other conclusions. The strategy for the GNC synthesis has been to consider and model the complete formation, 7 spacecraft and 6 ODL as one large interconnected system. With such approach the model and the synthesis can easily capture the system interconnections. As the implementation is as a computer controlled system it was natural to perform a direct discrete design of the complete system.The guidance strategy has been designed to cover the reference signals needed and the feed forward forces and torques for the maneuvers. This covers rotations, resizing, formation slew to re-target and station keeping. Various strategies have been investigated concerning autonomy of the formation. A leader-follower architecture has been selected, where one s/c is the master and all other ones follow that. FDIR schemes and collision avoidance strategies has been investigated together with operational schemes for such type of formations. All the modes and their transition criteria have been designed.

Finally there are two test campaigns; for the demonstration mission using 3 spacecraft case, and for the Darwin modes starting from BCM. All results are evaluated and compared to the requirement and iterations performed where needed.

The functionality of the simulator has been defined taking the operational environment into account. Models have been designed and verified for spacecraft, space environment perturbations, orbital dynamics at L2 and all the equipment and GNC system. The simulator can run in separate or all modes from one end to the other. Post processing has been design for plotting the raw data, covariance analysis and all statistics following use of the Monte Carlo capability.




11.12.4 PERFORMANCE


Required performance

Remarks

sided PSD of 3E-8 m/sqrt{Hz} up to a cut off frequency of 10 Hz

R0001 All baselines must be equal within 1 cm (rms) over the observation duration

Fulfilled using the ODL off loading

R0002 All incoming beams must be cophased with an accuracy of 5 nm (rms) over the observation duration

Fulfilled

R0003 All incoming beams must be superimposed with a directionalaccuracy of 8.5 mas (rms per axis) over the observation duration

Fulfilled marginally

R0004 The tolerance on amplitude mismatch of incoming beams is 0.1 percent (rms) over the observation duration

Fulfilled (TBC by end-to-endSimulations)

R0110 The transverse offset of each telescope flyer wrt. its nominal position shall remain below 1 cm (rms)

Fulfilled using the lateral metrology measurements.

R0120 The radial offset of each telescope flyer wrt. its nominal position shall remain below 1 cm (rms)

Fulfilled. Larger ODL stroke recommended due to bias accumulation for the 1st

fringe acquisition. See Section 11.12.6

R0310 The OPD stability shall be better than 2x5=10 nm (peak to peak) over one integration period of the fringe tracker

Fulfilled, but change of req. proposed. See Section 11.12.6

R0410 The observation axis shall be within +/- 45 deg of the ecliptic plane in the anti solar direction

Fulfilled

R1110 The initial OPD uncertainty before the scanning process is below 2 cm

Fulfilled, but marginally due to accumulated biases. See Section 11.12.6

R1210 The total duration of OPD scan(s) for fringe acquisition(s) on one target must be less than 6*400 = 2400 s

Violated. See Section 11.12.6

R1211 The total duration of OPD scan(s) for first fringe acquisition on one target must be less than 5000 s

Violated.See Section 11.12.6

R1310 During the scanning process, the OPD stability shall be better than a fraction of the scanning step over fringe tracker integration time

Evaluated OPD stability during scanning process is ~ 0.1E-6 m/s. Acquisition steps are 0.3E-6 m at 1 Hz.

R1510 During the fine attitude acquisition process, attitude stability shall be better than a fraction of WFC pixel over WFC integration time

Achieved

R5250 Slew maneuvers should be made optimal Achieved using mN thrusters.D0140 The sunshield of the hub has an area that guarantees equal, within

5 percent, surface/mass ratios flyers and hub Violated, needs to be lower. See 11.12.6

D0150 The cantilever frequency of the flexible modes of the sunshields is in the range [0.05;0.5] Hz. The damping is in the range [0.1;1.0] percent.For the purpose of GNC design, a value of 0.1 Hz has been used successfully.

D3570 The distance measurements from the laser metrology have a one-

Achieved





The GNC system needs interfaces to the following entities: - Propulsion system - RF Navigation system - Laser metrology system (divergent, lateral, range) - Wide field camera - Fringe sensor - Delay line actuator - Gyros - Star trackers - TM/TC system - Launch structure propulsion and on board system The estimated mass for the on board computer system which will host GNC system including the boot part is estimated to be below 15 kg. The power consumption will be below 30 W. These figures exclude the power and bus harness as well as redundant units.


Several critical issues have been identified during this activity. • The thrust resolution (quantization) is on the lower side. Though workable a higher

resolution would benefit the smoothness of the performance (less OPD noise, smoother control).

• Currently, the thrust capabilities in the +Z axis directions are low. This lag of controllability is not a serious issue in the stationary NOM, but it becomes critical during resizing, rotation and slew.

• The OPD stability shall be better than the fringe tracker measurement range (peak to peak) over one integration period of the fringe tracker. The OPD stability will also come from vibrations due to ODL and therefore should also be specified in terms of PSD.

• In the case where fringe tracking is lost in normal observation mode, it is proposed to implement a fast reacquisition procedure by scanning the delay lines around its last position. If successful, this procedure would avoid time-consuming array reconfigurations.

• The pathlength scan velocity has been set to 0.3E-6 m/s for a 1 Hz output fringe tracker having an acquisition range of 1E-6m to avoid any fringes lost when acquisition is performed. After acquisition, the commands of the delay line are brought back to zero (scan order cancellation) slowly to avoid any fringes lost. Then ODLs are off-loaded. The overall duration of this process for +/-2 cm uncertainties on OPD is of the order of 200000 s (2.3 days). For subsequent acquisition (after calibration), the uncertainty can be as low as 2 mm, resulting in a de-saturation time of about 10 hours. It is proposed to alter this requirement from 400 s to about 10 h. It is also proposed to change the requirement for the total duration of the first fringe scan on a new target from 5000 s to about 2 days.

• When accounting for all biases in the system, the initial OPD uncertainty before the first scanning process and calibration is +/- 2 cm. The subsequent ones are smaller. This




leads to a possible redefinition of the ODL actuator to have a 4 cm stroke contrary to present one of 2 cm.

• The 5 percent tolerance for the surface/mass ratio between the various spacecraft is an issue. It leads to a constant load disturbance on the control loops, which is undesirable. This problem could be obviated by a tighter requirement (<1 %), or having the GNC system to design an estimation of this constant load, which is not trivial. A sharing of the requirement between GNC and mechanical is proposed.

• The divergent laser system used to handover from RF metrology to laser metrology has presently a half cone angle of 0.5 deg. This is very marginal and a half cone angle of 1 deg is proposed.

• For the fine lateral metrology the emitter beam is better located on the telescope spacecraft as better pointed and the metrology is sensitive to mispointing.

• Propulsion step response time at 95 percent shall be shorter than 1 s. • The fine distance measurement from the RF shall be better than 1 cm (99.9 percent)

including bias and noise. This measurement influences the ODL stroke requirement. • The position knowledge of the RF sensor shall be neglegible wrt the sensor accuracy,

i.e. in the order of 1 mm (99.9 percent) • The coarse lateral and the longitudinal metrology measurement rates does not need to

be higher than 1 Hz. (with the current assumptions on the noise PSD) • The fringe sensor shall acquire OPD measurements between a pair of flyers with an

accuracy of 2 nm/sqrt{Hz} if the OPD is less than 1 micro meter. • The micro propulsion shall have a range of 0-150 µN with a minimum thrust of 0.3 µN

and a quantization of 0.3 µN up to 50 µN and 3 µN above.

11.13. Guidance, Navigation and Control II: Interferometer Deployment (ICD) Technology Development


This note constitutes the summary of the activity ``Interferometer Deployment Control'' under ESA contract 16328/02/NL/JA. The consortium selected for the study was Astrium UK with GMV (modelling and simulator), Astrium SAS and SciSys as sub contractors. The activity deals with the Guidance, Navigation and Control (GNC) design for the Darwin mission. It covers the deployment modes of the Darwin interferometer, namely the jettisoning from the launch dispenser, the transfer from Earth to the L2 orbit and the creation of the baseline control mode of the formation. In the deployment phase, the GNC system makes use of the milli Newton propulsion actuators. The sensors it uses are the RF system for relative distance measurements between s/c, gyros for spacecraft despin, star trackers and coarse sun sensors. The GNC design control all relative position and attitudes between spacecraft. A GNC design tool as well as a simulator has been developed including a Monte Carlo capacity for the statistical properties. The Darwin formation assumed for this study is the bow-tie, consisting of 6 telescope spacecraft with the beam combiner in the center and a master spacecraft out of the plane of the interferometer itself. The consortium selected for the study was Astrium UK with GMV

(modelling and simulator), Astrium SAS and SciSys as sub contractors.





The overall functionality is to get measurement from all the sensor systems and control the relative positions and relative attitudes by use of the milli Newton propulsion system. This means that the spacecraft shall not collide, and shall stay within a sphere of diameter 8 km. The guidance function is build as an artificial potential function, which is a very convenient way to implement a 3D collision avoidance scheme. In addition to this the guidance provides the feed forward forces and torques required to perform the maneuvers. The state controller performs the function of reducing the control error to zero and to minimize the noise transmission to the platform. The control design is kept as a Single Input Single Output system and couplings are handled as disturbances to the system. Couplings are less critical than in the actual formation when operational. On top of the GNC is designed and implemented a state machine with the functionality of mode management and handling the criteria of which the system leaves one mode and enters the next. Beyond the modes described earlier is also a Failure Detection Identification and Recovery (FDIR) system. It handles various submodes related to failures of various redundant avionics equipment as well as collision maneuvers.

11.13.3 DESCRIPTION

The development is kept largely generic for the on board system. The performance requirements are such as to meet the Darwin requirements for the TTN configuration, and for other relevant configurations as for example a 3 spacecraft demonstration mission with two telescopes and a beam combiner. This activity has got defined coordinate frames, rotation sequences and notation, which has been used to harmonize with all relevant activities in the technology developments for Darwin. The environment at L2 has been described well in terms of perturbation models as well as a careful modeling of the spacecraft kinematics and dynamics properties. Flexible elements of the structural modes of the sun shield and solar arrays has been included. A specific propulsion architecture has not been implemented, but engine characteristics are modeled as well as the typical cross couplings for such size of spacecraft. The environment of a transfer orbit from Earth to L2 is also described in detail.

The strategy for the GNC synthesis has been to consider and model the single axes of each spacecraft. As the implementation is as a computer controlled system it was natural to perform a direct discrete design of the complete system but has not been a performed in practice by the prime contractor. Therefore the design is in continuous time and then discretized in an approximate manner.

Interfaces to the other avionics subsystems have been defined and their performance characteristic is the base for the GNC design wrt the performance requirements. The architecture for such distributed a system is then analyzed for a central and a decentralized architecture. The advantages and disadvantages of the two have been assessed, resulting in a decentralized structure for the GNC system in question here.




The guidance strategy has been designed to cover the reference signals needed and the feed forward forces and torques for the maneuvers. This covers jettison, despin, braking, collission avoidance, loose formation creation, transfer orbital maneuvers and finally the formation of the baseline control mode. Various strategies have been investigated concerning autonomy of the formation leading to a leader-follower architecture. FDIR schemes and collision avoidance strategies has been investigated together with operational schemes for such type of formations. All the modes and their transition criteria have been designed. The functionality of the simulator has been defined taking the operational environment into account. Models have been designed and verified for spacecraft, space environment perturbations, orbital dynamics at L2 and all the equipment and GNC system. The simulator can run in separate or all modes from one end to the other. Post processing software has been designed for plotting the raw data, covariance analysis and all statistics following use of the Monte Carlo capability. There has been developed a coherent working interface between the simulators of ICC and ICD studies.


The GNC system needs interfaces to the following entities. - Propulsion system - RF Navigation system - Gyros - Star trackers - TM/TC system - Launch structure propulsion and on board system The following values are estimated for the on board computer system which will host GNC system including the boot part. This is excluding the power and bus harness as well as redundant units. The mass and power is shared with the GNC system done in ICC study.






Value

Remarks


Power


Thermal interfaces

Operational temperature (K) NA Operational temp. range (K) NA

Optical interfaces


Other



As the study activity is severely delayed, there are no identification of critical issues really, but one. The RF system must be able to perform relatively well over angles of 4π steradians. The GNC systems designed in the study activities ICC and here ICD will clearly be one system in reality. It poses the difficulty that the implementation will have to be flexible to accomodate both a decentralized and centralized architecture for different modes. The estimated mass for the on board computer system which will host GNC system including the boot part is estimated to be below 15 kg. The power consumption will be below 30 W. These figures exclude the power and bus harness as well as redundant units.

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