4 Final Report - Emits - ESA
4 Final Report - Emits - ESA
4 Final Report - Emits - ESA
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Geo-Oculus: A Mission for Real-Time Monitoring through<br />
High-Resolution Imaging from Geostationary Orbit<br />
All the space you need
<strong>ESA</strong> CONTRACT No.<br />
21096/07/NL/HE<br />
* <strong>ESA</strong> CR( )No<br />
<strong>ESA</strong> STUDY CONTRACT REPORT<br />
SUBJECT<br />
Geo-Oculus: A Mission for Real-Time<br />
Monitoring through High Resolution<br />
Imaging from Geostationary Orbit<br />
*STAR CODE<br />
No of volumes: 1<br />
This is volume no: 1<br />
CONTRACTOR<br />
Astrium GmbH<br />
GOC-ASG-RP-003<br />
Issue 1-0<br />
ABSTRACT:<br />
This <strong>Final</strong> <strong>Report</strong> summarises the results of the study “Geo-Oculus: A Mission for Real-Time<br />
Monitoring through High Resolution Imaging from Geostationary Orbit” performed from October 2007<br />
to April 2009 and lead by Astrium GmbH in Friedrichshafen.<br />
In the frame of the study a comprehensive survey of the potential user needs has been performed with<br />
the result that a demand has been defined within the existing political and institutional framework for a<br />
high resolution and high revisit mission from geostationary orbit. Out of the identified applications four<br />
have been selected as primary mission objectives used to size a system and confirm principle<br />
feasibility on a Phase-0 level.<br />
Those primary objectives have been: Disaster monitoring, fire monitoring, algal bloom detection and<br />
monitoring and water quality monitoring.<br />
Taking the user requirements for these applications a set of mission and system requirements has<br />
been derived and a first iteration of the payload, spacecraft and ground segment design has been<br />
elaborated.<br />
On the payload side the design lead to a telescope with an aperture of 1,5 m diameter and five focal<br />
planes. The feasibility of implementing the envisaged GSD of around 10 m (at the equator) has been<br />
confirmed. The feasibility of all selected applications has also been confirmed.<br />
To tackle the stringent LoS requirements various techniques from disturbance suppression over image<br />
processing and active LoS control have been studied. Also the application of image post processing<br />
on ground with landmark detection (INR) has been considered.<br />
On the spacecraft design emphasis has been placed on the AOCS. It has been confirmed that the<br />
required agility of the system can be realized. It has been demonstrated that the allocated manoeuvre<br />
time including tranquilisation is feasible which leads to an imaging capability of around 42 images per<br />
hour for Geo-Oculus.<br />
<strong>Final</strong>ly a first iteration of the ground segment architecture has been elaborated investigating the main<br />
challenges fast data dissemination and flexible mission planning taking into account on-demand<br />
imaging (emergency missions) but also cloud dynamics.<br />
The work described in this report was done under <strong>ESA</strong> Contract. Responsibility for the<br />
contents resides in the author or organisation that prepared it.<br />
Names of authors:<br />
Astrium study team, lead by Astrium Study Manager Ulrich Schull<br />
** NAME OF <strong>ESA</strong> STUDY<br />
MANAGER:<br />
Jean-Loup Bézy (EOP-PIO)<br />
Earth Observation Programmes Directorate<br />
* Sections to be completed by <strong>ESA</strong><br />
** Information to be provided by <strong>ESA</strong> Study Manager<br />
** <strong>ESA</strong> BUDGET HEADING:<br />
OUTPUT: 60 GSP<br />
SUB-HEADING: 510 Special Studies
DL <strong>Final</strong><br />
Distribution List<br />
<strong>Report</strong><br />
Quantity Type * Name Company / Department<br />
1 PDF J.-L. Bezy <strong>ESA</strong><br />
1 PDF M. Aguirre <strong>ESA</strong><br />
1 PDF F. Gascon <strong>ESA</strong><br />
1 Word & PDF U. Schull Astrium GmbH<br />
1 Word & PDF U. Schäfer Astrium GmbH<br />
1 Word & PDF T. Knigge Astrium GmbH<br />
1 Word & PDF X. Sembely Astrium SAS<br />
1 Word & PDF L. Vaillon Astrium SAS<br />
1 Word & PDF N. Leveque Astrium Ltd<br />
* Type: Paper Copy or Electronic Copy (e.g. PDF or WORD file etc.)<br />
Doc. No: GOC-ASG-RP-002 Page i<br />
Issue: 2<br />
Date: 13.05.2009 Astrium GmbH
CR <strong>Final</strong><br />
Change Record<br />
<strong>Report</strong><br />
Issue Revision Date Sheet Description of Change<br />
1 28.01.2009 All First version<br />
2 13.05.2009 Consideration of <strong>ESA</strong> comments:<br />
2-1, 5-87 Reference to GMES<br />
2-2 Reference to S-2 and S-3<br />
3-20 SSD units changed from km to m<br />
3-23 Description of cloud cover change<br />
4-27 Figure 4.1-1 repaired<br />
4-29 Explanation of versions a and b of MWIR/TIR channels<br />
4-32 Explanation of columns 6 and 7 of Figure 4.3-3<br />
4-53 Additional information on data rate<br />
4-53 Information on TRL of downlink system<br />
4-56 Impact of small mobile stations on downlink system<br />
4-58 Comment on active damping of flexible modes<br />
4-60 to 62 Probability level of pointing performance<br />
4-63, 4-65 Units for manoeuvre times added in tables<br />
4-80 Information on alternative decentralised PDGS<br />
5-87 New section added<br />
Doc. No: GOC-ASG-RP-002 Page iii<br />
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TOC <strong>Final</strong><br />
Table of Content<br />
<strong>Report</strong><br />
Distribution List..........................................................................................................i<br />
Change Record.........................................................................................................iii<br />
Table of Content ........................................................................................................v<br />
1 Introduction ...................................................................................................1-1<br />
1.1 Scope of the document...........................................................................................................1-1<br />
1.2 References ............................................................................................................................... 1-2<br />
1.2.1 Applicable Documents ........................................................................................................... 1-2<br />
1.2.2 Reference Documents ........................................................................................................... 1-2<br />
2 Product and Mission Baseline Description.................................................2-1<br />
2.1 Overview................................................................................................................................... 2-1<br />
2.2 Survey for Mission Objectives ............................................................................................... 2-1<br />
2.3 Mission Objectives .................................................................................................................. 2-3<br />
3 System Requirements and Mission Scenarios .........................................3-12<br />
3.1 System Requirements........................................................................................................... 3-12<br />
3.2 Major System Trade-Offs...................................................................................................... 3-17<br />
3.3 Mission Scenarios ................................................................................................................. 3-18<br />
3.3.1 Mission Scenario Baseline................................................................................................... 3-19<br />
3.4 Cloud Coverage Analysis ..................................................................................................... 3-22<br />
4 Mission and System Level Analyses .........................................................4-27<br />
4.1 Mission Architecture............................................................................................................. 4-27<br />
4.2 Mission Analysis ................................................................................................................... 4-27<br />
4.3 Payload................................................................................................................................... 4-29<br />
4.3.1 Imaging capability ................................................................................................................ 4-29<br />
4.3.2 Radiometric & image quality performances......................................................................... 4-30<br />
4.3.3 Instrument design ................................................................................................................ 4-33<br />
4.3.4 PLM budgets........................................................................................................................ 4-42<br />
4.4 Line of Sight (LoS) Stabilisation Concepts......................................................................... 4-43<br />
4.4.1 LoS stabilisation main issues: microvibrations and post-integration ................................... 4-43<br />
4.4.2 Microvibrations..................................................................................................................... 4-44<br />
4.4.3 Post-integration.................................................................................................................... 4-46<br />
4.5 Satellite................................................................................................................................... 4-48<br />
4.5.1 Configuration........................................................................................................................ 4-48<br />
4.5.2 Electrical Architecture .......................................................................................................... 4-51<br />
4.5.3 Power Subsystem ................................................................................................................ 4-52<br />
4.5.4 Payload Data Handling and Transmission........................................................................... 4-53<br />
4.5.5 Telemetry and Telecommand .............................................................................................. 4-56<br />
4.5.6 Attitude and Orbit Control .................................................................................................... 4-58<br />
4.5.7 Propulsion System ...............................................................................................................4-66<br />
4.5.8 Structure and Thermal Concept........................................................................................... 4-72<br />
4.5.9 Satellite Budgets .................................................................................................................. 4-77<br />
4.6 Ground Segment ................................................................................................................... 4-77<br />
4.6.1 Ground Segment Architecture ............................................................................................. 4-77<br />
4.6.2 Geo-Oculus dedicated Ground Segment issues ................................................................. 4-80<br />
Doc. No: GOC-ASG-RP-002 Page v<br />
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TOC <strong>Final</strong><br />
<strong>Report</strong><br />
5 Recommendations on further Analysis .................................................... 5-87<br />
5.1 System Analysis ....................................................................................................................5-87<br />
5.2 Mission Objectives and Data Processing ...........................................................................5-87<br />
6 Conclusion .................................................................................................... 6-1<br />
Annex A Abbreviations .................................................................................. A-1<br />
vi Page Doc. No: GOC-ASG-RP-002<br />
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1 Introduction<br />
1.1 Scope of the document<br />
This document provides the main findings of the study "Geo-Oculus - A Mission for Real-Time<br />
Monitoring through High Resolution Imaging from Geostationary Orbit".<br />
The study teaming is as follows:<br />
• Astrium GmbH Study Prime<br />
• Astrium SAS Instrumentation and LoS<br />
• Astrium Limited Mission Analysis<br />
• DLR Institute for Optical Systems Focal Plane Analysis<br />
The following consultants have contributed to the definition of the suitable applications and the<br />
collection of the product requirements:<br />
• ACRI-ST Traffic & Security at Sea, Earth Science Applications<br />
• Brockmann Consult Marine Applications<br />
• Infoterra GmbH Land Applications<br />
The first task of the study has been an open minded survey for applications for a mission that<br />
combines fast-response, high-revisit, near-real-time and high-resolution capabilities to introduce a new<br />
class of Earth observation missions. Figure 1.1-1 illustrates the uniqueness of Geo-Oculus to provide<br />
high resolution images in the scale of Sentinel 2 at the revisit time and the timeliness of MTG. Based<br />
on the survey for applications a set of mission objectives has been selected in consultation with <strong>ESA</strong><br />
for sizing of the system during this study.<br />
Chapter 2 gives a short overview on the starting point for the survey and describes the approach for<br />
the selection of the mission objectives. The mission objectives are briefly described in 2.2.<br />
The second task of the study has been to derive and analyse the system requirements and to<br />
establish the preliminary candidate mission concepts.<br />
In chapter 3 the driving system requirements are summed up. The major system trade-offs "Field of<br />
View vs. resolution", " Magnetic Bearing Wheels vs. Electric Propulsion for manoeuvres", "Manoeuvre<br />
time vs. image post-integration effort" and "Image post integration and Inter-channel co-registration"<br />
are described in 3.2<br />
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Revisit Capability [min]<br />
100000<br />
10000<br />
1000<br />
100<br />
10<br />
<strong>Report</strong><br />
Revisit vs. Resolution<br />
Bubblesize indicates Timeliness (small=fast)<br />
1<br />
0,1 1 10 100 1000 10000<br />
Resolution @ SSP [m]<br />
Geo-Oculus<br />
Disaster<br />
Geo-Oculus<br />
Fire<br />
Geo-Oculus<br />
Marine<br />
MTG (HRFI)<br />
Page 1-2 Doc. No: GOC-ASG-RP-002<br />
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MSG<br />
Metop +<br />
NOAA<br />
Envisat<br />
EOS<br />
Terra & Aqua<br />
Landsat 7<br />
SPOT 5<br />
Ikonos<br />
GeoEye-1<br />
Pleiades<br />
RapidEye<br />
Sentinel 2<br />
(2Sats)<br />
Sentinel 3<br />
Figure 1.1-1 Geo-Oculus compared to other EO-missions in terms of revisit against resolution<br />
1.2 References<br />
1.2.1 Applicable Documents<br />
[AD 1] ESTEC Contract No. 21096/07/NL/HE<br />
[AD 2] Statement of Work 'Geo-Oculus: A Mission for Real-Time Monitoring through High<br />
Resolution Imaging from Geostationary Orbit'; TEC-EEP/2006.93/FG Iss.: 01, Rev.: 01<br />
Date: 03.04.2007<br />
1.2.2 Reference Documents<br />
[RD 1] Geo-Oculus: A Mission for Real-Time Monitoring through High Resolution Imaging<br />
from Geostationary Orbit; EADS Astrium GmbH Proposal No., A.2007-4200-0-1:<br />
Author: Dr. Ralf Münzenmayer; Friedrichshafen; July 2007<br />
[RD 2] Rapport de l’étude « Contraintes induites par une instrumentation d’observation HR<br />
sur une plateforme Géostationaire », PFGEO.ASTR.TN.001.06, Edition 2.1,<br />
08.06.2007<br />
[RD 3] System Requirements <strong>Report</strong>, GOC-ASG-TN-002, Iss.: 01, Rev.: 00,<br />
Date: 02.06.2008<br />
[RD 4] Product Requirements <strong>Report</strong>, GOC-ASG-TN-001, Iss.: 01, Rev.: 02,<br />
Date: 30.05.2008<br />
[RD 5] LoS Stabilisation Concepts, GOC-ASF-IN-002, Iss.: 01, Rev.: 00,<br />
Date: 30.05.2008<br />
[RD 6] Candidate Instrument Concepts <strong>Report</strong>, GOC-ASF-IN-001, Iss.: 01, Rev.: 00,<br />
Date: 30.05.2008
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[RD 7] Candidate Mission Concepts <strong>Report</strong>, GOC-ASG-TN-003, Iss.: 02, Rev.: 00,<br />
Date: 19.11.2008<br />
[RD 8] Instrument Analysis <strong>Report</strong>, GOC-ASF-IN-003, Iss.: 02, Rev.: 00,<br />
Date: 16.01.2009<br />
[RD 9] LoS Stabilisation Analysis <strong>Report</strong>, GOC-ASF-IN-004, Iss.: 02, Rev.: 00,<br />
Date: 16.01.2009<br />
[RD 10] Preliminary Mission Analysis <strong>Report</strong>, GOC-ASG-TN-004, Iss.: 01, Rev.: 00,<br />
Date: 28.01.2009<br />
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2 Product and Mission Baseline Description<br />
2.1 Overview<br />
The first task of the study has been the identification and selection of applications for Geo-Oculus in<br />
order to define preliminary mission objectives including the technical requirements.<br />
Geo-Oculus is set as an independent mission with the objective to enable observations of the Earth so<br />
far not feasible with current or planned systems or missions. Among the wealth of EO-activities on<br />
European and international level, <strong>ESA</strong> has identified the lack of the capability for a combination of fastresponse,<br />
high-revisit, near-real-time and high-resolution observations. Therefore, the starting point for<br />
the study is a geo-synchronous satellite mission with high resolution optical imaging instrumentation,<br />
real-time control and agile platform.<br />
A survey for mission objectives covering a very wide field of Earth observation applications has been<br />
performed to identify applications that require or profit from the basic characteristics of a mission like<br />
Geo-Oculus. This survey is briefly described in 2.2 and in more detail in [RD 4]. Although, the survey<br />
identified a number of applications that benefit substantially from Geo-Oculus, one major finding is that<br />
this mission lays the foundation for new kinds of Earth observation applications which yet have to be<br />
recognized.<br />
For the selection process a ranking scheme has been chosen that rates all applications, that where<br />
identified in the survey, in terms of "Political Importance", "Institutional / Non-Profit Importance"<br />
"Commercial Importance" and "Suitability of Geo-Oculus". Based on this ranking, a final selection of<br />
the preliminary mission objectives has been conducted in consultation with <strong>ESA</strong>.<br />
2.2 Survey for Mission Objectives<br />
For the survey for mission objectives, a comprehensive review for user requirements, potential<br />
applications and the related product requirements has been conducted. The scope of this survey<br />
covers the political framework in terms of ongoing or future European initiatives, especially the GMES<br />
initiative, as well as international treaties and European and national directives, policies and protocols.<br />
Synergies with European and international Earth observation systems and missions, like the<br />
Sentinels, GEOSS and EPS were identified and considered for the identification of suitable<br />
applications for Geo-Oculus.<br />
Mission of Choice<br />
The analysis of user requirements and potential applications is conducted with an open mind for user<br />
demands that will especially benefit from the mission characteristics in fields of e.g.:<br />
• Ecological, economical and humanitarian incidents<br />
• Rapidly evolving events<br />
• Local to regional monitoring<br />
• Instantaneous situation awareness<br />
• Regions regularly covered with clouds<br />
The survey points out that Geo-Oculus is the 'mission of choice' for the above mentioned fields of<br />
applications. Yet, another finding is that only few applications already exist that require the specific<br />
features of Geo-Oculus to become possible. This is not due to missing interest but due to missing<br />
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capability of existing missions. Nevertheless many existing applications are identified that can profit<br />
from Geo-Oculus, some can significantly profit or first become possible in an operational manner.<br />
The range of applications covers a large variety of remote sensing applications of the Earth's surface.<br />
For the structuring of the documents and to account for the expertise of and task of the consultant it<br />
has been chosen to categorize the applications into four fields of services, the:<br />
• Land Applications, covering all services related to the Earth's solid surface, including the<br />
land part of the coastal zones and disaster monitoring;<br />
• Marine Applications, covering the services related to the marine ecosystem, hydrology,<br />
oceanography, sustainable exploitation of marine resources and anthropogenic forcing and<br />
threat to the environment;<br />
• Traffic and Security at Sea Applications, covering marine operations, natural threat to the<br />
citizen and law enforcement related to the oceans;<br />
• Earth Science Applications, covering climate research (esp. role of the ocean), the Earth's<br />
radiation budget, monitoring of rapid events and data assimilation.<br />
For all applications the necessary and optional products have been identified and defined to a sound<br />
level of detail to provide the required technical parameters for the definition of the mission and the<br />
instruments. A threshold, a breakthrough and a goal value have been given for most of the<br />
parameters, if available through generally accepted literature.<br />
Important synergies<br />
Geo-Oculus provides strong assets for synergies with current and planned European EO-missions.<br />
The optimisation for cloud cover, which is considered as a central benefit of Geo-Oculus, is only<br />
possible with support data from Meteosat and EPS. On the other hand, Geo-Oculus can support other<br />
missions to improve quality of service. Some synergies, receiving and supportive, are listed below:<br />
Receiving synergies:<br />
• Real-time cloud cover information from Meteosat and EPS<br />
• Fire presence by any means<br />
• Highest resolution support data e.g. for disaster monitoring from SPOT, Pleiades, Ikonos etc.<br />
Supportive synergies:<br />
• Oil slick verification<br />
• Gap filling due to cloud cover for Sentinel 2 & 3<br />
• Fine scale and real-time spotlight support to meteorology, e.g. for severe weather events<br />
These synergies are considered as a prerequisite for the selection of mission objectives.<br />
Selection process<br />
In a preliminary selection process based on the criteria "Political Importance", "Institutional / Non-Profit<br />
Importance", "Commercial Importance" and "Suitability of Geo-Oculus" a set of applications was<br />
proposed to <strong>ESA</strong> and in consultation with <strong>ESA</strong> the preliminary mission objectives were set. These<br />
were used for sizing of the system. For that reason, only sufficiently elaborated applications with<br />
available technical requirements could be taken into consideration. Nevertheless, these mission<br />
objectives represent a realistic case, demanding a challenging system without overtightening the<br />
requirements.<br />
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2.3 Mission Objectives<br />
The mission objectives for Geo-Oculus that have been selected in consultation with <strong>ESA</strong> based on the<br />
survey described in 2.2 are:<br />
Primary Mission Objectives:<br />
• Disaster Monitoring<br />
• Fire Monitoring<br />
• Algal Bloom Detection / Monitoring<br />
• Water Quality Monitoring with respect to European Regulation<br />
Secondary Mission Objectives:<br />
• Oil Slick Environmental Information<br />
• Erosion / Sediment Transport on the European Shoreline Monitoring<br />
An overview of each of the mission objectives is given in the following:<br />
Disaster Monitoring Service - Primary Objective<br />
The disaster monitoring service is aimed at providing overview information in case of natural hazards<br />
with significant geographic extend. Based on the findings of PREVIEW (2006) on the priorities<br />
adopted for Civil Protection activities of the Member States concerning the risk management and the<br />
suitability of high-resolution imaging, the following hazards are considered for the Geo-Oculus disaster<br />
monitoring service:<br />
• Large landslides<br />
• Floods<br />
• Windstorms<br />
It is the goal of the disaster monitoring service to deliver geospatial information with short acquisition<br />
delay and timeliness of less than an hour on demand of civil protection organisations. This service<br />
shall be tailored for early warning, crisis, and post crisis support to the users.<br />
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Figure 2.3-1: Flood Waters surrounding Yangon City, Myanmar, Cyclone Nargis<br />
Source: MODIS on Terra and Aqua, 28.5m/pixel resolution.<br />
Acquired: 05/05/2008 and 18/03/2008<br />
Satellite-detected flood waters over Yangon, as of 5 May 2008. Red areas shown in<br />
the map represent standing flood waters identified from Landsat 7 satellite imagery<br />
acquired on 5 May 2008 at a spatial resolution of 28.5m.<br />
Blue areas represent pre-flood waters identified from Landsat 7 acquired on 18 March<br />
2008. Preliminary analysis not yet verified in the field. Credit: Credit NASA/USGS<br />
2008<br />
Image processing, map created 05/05/2008 by UNOSAT.<br />
Disaster monitoring services will benefit substantially of the significant advantages of the GeoOculus<br />
mission in terms of acquisition delay, observation cycle and timeliness, all of which in the range of an<br />
hour or less.<br />
Fire Monitoring Service - Primary Objective<br />
The fire monitoring service is an on occasion service that becomes active once a fire in the service<br />
region is present, which has been detected and reported by other means. Therefore it is the objective<br />
of the fire monitoring service is to provide timely fire observations on demand of fire fighting and<br />
mitigation organisations. The data products shall provide accurate information on fire location, extend,<br />
temperature and development over time to allow for optimized planning of mitigation efforts.<br />
This requires a very agile and responsive overall system to assure acquisition delays and observation<br />
cycles (revisit time) shorter than about 10 min and data delivery after acquisition in less than about<br />
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15 min. These requirements are not possible to be fulfilled by LEO missions and underline the<br />
strengths of the GeoOculus mission.<br />
The primary region of service provision is the European mainland, nevertheless can it be considered<br />
to provide this service for the whole visible regions (e.g. Africa) with reduced priority to avoid<br />
interruption of other services.<br />
Figure 2.3-2: Forest fires near Los Angeles in October 2003 imaged by the DLR BIRD satellite in<br />
the course of 24 hours. The GSD is 185 m squared. The scale is as follows: yellow =<br />
0,1 MW / Pixel; orange = 1 MW / Pixel; red = 10 MW / Pixel. © DLR<br />
Additional observations of manifold high temperature events, like volcanic activity, tropical peat land<br />
fires and coal seams, mainly for scientific purposes, shall also be covered by this service; hence not<br />
drive the system requirements.<br />
Algal Bloom Detection Service - Primary Objective<br />
An algal bloom refers to a quick and local increase in the abundance of a phytoplankton species. The<br />
so-called Harmful Algal Blooms (HAB) are special cases of the former, with deleterious effects on<br />
human health or marine resources (natural or cultured). Therefore the algal bloom detection service is<br />
aimed to detect and locate an algal bloom in European waters. This information shall be incorporated<br />
into the respective GSE MARCOAST service line.<br />
Users include national environment agencies and fisheries industries (aquaculture, shellfish). All<br />
coastal and offshore European waters are concerned (region enclosed between latitudes 35°N and<br />
70° N and longitudes 12°W and 30°E).<br />
Although blooms are relatively well detected by remote sensing technique, the identification of their<br />
possible toxicity from space is far from being achieved. Satellites play a crucial role for forecasting<br />
(combination of EO data with models) and visualising the extent.<br />
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Figure 2.3-3: Algal bloom at Cape Rodney, NZ; Photo by Miriam Godfrey.<br />
Attention shall be paid to the definition of an Algal bloom. For some cases it is an anarchic increase of<br />
biological material compared to a climatology (seasonal evolution), for other application it may just be<br />
the sudden increase of chlorophyll concentration (thus covering the seasonal trends) as it is presently<br />
done in the Algal Bloom service line of GSE-Marcoast.<br />
Algal Bloom Monitoring Service - Primary Objective<br />
Algal blooms in European waters, either detected by Geo-Oculus or reported by local authorities, shall<br />
be monitored within the algal bloom monitoring service. The service is aimed to provide detailed<br />
information on position, extend and persistence to the users (cp. chapter 0). The data can be derived<br />
either by dedicated observations or within the required routine scanning of the algal bloom detection<br />
service, if applicable.<br />
Figure 2.3-4: Algal bloom east of Scotland, May 7th, 2008, Envisat-MERIS image, unusually strong<br />
algal bloom degrades bathing water quality and threatens the local ecosystem.<br />
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Figure 2.3-5: Algal bloom in the North Sea and west of France, © Y. Park,<br />
MUMM (MarCoast)<br />
Water Quality Monitoring Service with respect to European Regulation - Primary Objective<br />
The water quality monitoring service with respect to European regulation addresses mainly the user<br />
needs of the WFD and conventions like the Bathing Water Directive. For the later it is required to<br />
examine the bathing-water quality with concern to public health criteria. The objective of this service is<br />
to provide timely status reports of the water quality of the European waters.<br />
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Figure 2.3-6: Chlorophyll concentration on 31.03.2007 (Envisat MERIS). Processed for Marcoast.<br />
© Brockmann Consult / LANU (MarCoast)<br />
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Figure 2.3-7: Turbidity of the lake Lohjanjärvi on 20.5.2002 (Landsat 7 ETM+).<br />
© Finish Environment Institute (SYKE)<br />
Oil Slick Environmental Monitoring Service - Secondary Objective<br />
The detection of oil spills is coordinated at European level through the CleanSeaNet initiative. All<br />
information about oil detection is directed towards EMSA who has the mandate to track and survey<br />
illegal discharges at sea in order to intercept polluters. The system of oil observation is run in parallel<br />
with Automatic Identification Systems (AIS) and Vessel Monitoring System (VMS) that allows<br />
identification of polluters. The system is operational since 2007.<br />
Figure 2.3-8: Left: Fresh oil slick spread widely into a thin film.<br />
Right: Partly dispersed oil slick as seen by an airplane.<br />
© Cedre [RD T8]<br />
The oil spill detection today relies especially on SAR images but should be complemented with<br />
ancillary information (such as SST and Ocean colour) in order to improve the level of confidence of the<br />
detection and corresponding reporting – this statement is especially valid in the Baltic where biogenic<br />
spills due to biological material may lead to misdetection of oil spills (see EMSA documentation and<br />
GSE-Marcoast phase 2 recommendation). This is the objective of the oil slick environmental<br />
monitoring service.<br />
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Erosion / Sediment Transport on the European Shoreline Monitoring Service - Secondary<br />
Objective<br />
The erosion / sediment transport on the European shoreline monitoring service is dedicated to assess<br />
the impacts of natural events like floods and storms, as well as increased river discharge on the<br />
European shoreline. Even though that the formulated user needs (see TMAP [RD M4]) demand for a<br />
spatial resolution in the range of 1-5 m, which is not achievable by Geo-Oculus, the short delay of<br />
image acquisition and delivery combined with a still high spatial resolution, denote Geo-Oculus as an<br />
indispensable mission for this service; hence the focus of the service is to timely provide data for<br />
damage assessment on demand of the authorities.<br />
Figure 2.3-9: Sediment classification of the Wadden Sea, © K. Stelzer, Brockmann Consult<br />
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Figure 2.3-10: Wadden Sea largest intertidal area worldwide is suspect to erosion due to storms and<br />
flooding<br />
Table 2.3-1 gives an overview of the mission objectives including the basic requirements for each<br />
application/mission.<br />
Table 2.3-1 Mission Objectives for Geo-Oculus<br />
Application<br />
Mission 1:<br />
Disaster<br />
Monitoring<br />
Geo-Oculus Mission Objectives<br />
Primary Mission Objectives Secondary Mission Objectives<br />
Mission 2: Fire<br />
Monitoring<br />
Mission 3: Algal<br />
Bloom Detection /<br />
Monitoring<br />
Mission 4: Water<br />
Quality Monitoring<br />
wrt. European<br />
Regulation<br />
Mission 5: Oil<br />
Slick<br />
Environmental<br />
Information<br />
Mission 6:<br />
Erosion /<br />
Sediment<br />
Transport on the<br />
European<br />
Shoreline<br />
Monitoring<br />
Type of service on demand on demand<br />
all European fire<br />
routine / on<br />
demand routine on demand on demand<br />
endangered areas<br />
European coastal<br />
Service regions Europe<br />
up to 45° N all European waters all European waters all European waters waters<br />
Product Image Size 150 x 150 km² 100 x 100 km² 100 x 100 km² 100 x 100 km² 100 x 100 km² 100 x 100 km²<br />
Service period all year summer-early fall all year all year all year all year<br />
Daily service period (solar<br />
24 hours (sun<br />
zenith angle, time span)
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3 System Requirements and Mission Scenarios<br />
3.1 System Requirements<br />
The system requirements for Geo-Oculus have been derived from user and product requirements,<br />
iterated during the first part of this study and documented within the System Requirements <strong>Report</strong>,<br />
RD [3]. These requirements have been further evolved after MTR. Within this chapter, a summary of<br />
the driving requirements is given.<br />
Major Challenges for the Geo-Oculus Mission<br />
The unprecedented high resolution combined with large areas to be covered within a short period of<br />
time drives the system concept of Geo-Oculus. The high resolution requires a large telescope<br />
diameter and high pointing stability, whereas the coverage drives the detector Field of View and short<br />
repeat cycles ask for short manoeuvre times. A large numbers of required channels drives the<br />
instrument optics and focal plane assembly (number of detectors and filter wheel), whereas the MTF<br />
and SNR requirements ask for image post-integration techniques.<br />
Definition of Missions<br />
Six mission objectives are defined as follows:<br />
Primary objectives:<br />
• Mission objective 1: Disaster Monitoring;<br />
• Mission objective 2: Fire Monitoring;<br />
• Mission objective 3: Algal Bloom Detection / Monitoring;<br />
• Mission objective 4: Water Quality Monitoring with respect to European Regulation.<br />
Secondary objectives:<br />
• Mission objective 5: Oil Slick Environmental Information;<br />
• Mission objective 6: Erosion / Sediment Transport on the European Shoreline Monitoring.<br />
Both primary and secondary mission objectives are considered for the system requirements definition.<br />
In order to derive system observation requirements, these six defined missions objectives are<br />
compared in terms of system driving requirements (observation cycle, coverage requirements, etc.),<br />
see following table.<br />
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Table 3.1-1: Geo-Oculus mission objectives<br />
Mission<br />
objective<br />
1 - Disaster<br />
Monitoring<br />
2 - Fire<br />
Monitoring<br />
3 - Algal<br />
Bloom<br />
Detection /<br />
Monitoring<br />
4 - Water<br />
Quality<br />
Monitoring<br />
Observation cycle Coverage Observation time / period<br />
Goal Threshold Goal Threshold<br />
1 hour 2 days Land areas (on<br />
demand)<br />
10 min 1 hour Land areas (on<br />
demand)<br />
1 day 3 days Full coverage of<br />
European coastlines<br />
1 day 3 days Full coverage of<br />
European coastlines<br />
5 - Oil Slick 1 hour 6 hour Specific areas of<br />
European coastlines<br />
(on demand)<br />
6 - Erosion /<br />
Sediment<br />
Transport<br />
1 hour 6 hour Specific areas of<br />
European coastlines<br />
(on demand)<br />
SZA
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Furthermore, the mission shall be capable of providing images over the whole Earth disk, with relaxed<br />
geometric, revisit, sensing, etc. requirements.<br />
Following figure shows the area to be covered by the marine application mission. The whole area can<br />
be covered by 65-70 images with a FoV of 285km x 285km:<br />
Figure 3.1-1: Coverage for marine application mission with 65-70 285km x 285km FoV images<br />
The Sun Zenith Angle (SZA) shall be < 80 deg for images acquired for the disaster monitoring, oil<br />
slick, erosion and marine application mission, whereas for the fire monitoring mission, no SZA has to<br />
be specified. The definition of the SZA determines the time window, during which the area can be<br />
observed. The operational concept shall consider an optimisation of the SZA for the marine<br />
applications. The average SZA of all acquired images within the extended observation area of one<br />
observation cycle shall be minimised. The radiometric requirements (definition of minimum radiance)<br />
shall assume a SZA < 75 deg for the disaster monitoring and < 60 deg for oil slick, erosion and marine<br />
application mission.<br />
The observation times and periods for all missions, except the fire monitoring mission, depend on<br />
the specification of the maximum allowed sun zenith angle and on the season. For summer solstice,<br />
the observation periods are longest. For this case, a mean observation time for the marine application<br />
mission of 9 hours has been considered for the mission scenarios presented within the next chapter.<br />
Due to cloud coverage, the effective coverage (=cloudfree coverage) will differ from the nominal<br />
coverage. From system side, the impact of cloud coverage can only be minimised by optimising the<br />
observation strategy of the marine applications mission. As it is assumed that emergency missions<br />
(fire and disaster monitoring) are conducted in parallel to the marine applications mission, the 2°<br />
manoeuvres which have already to be considered for the emergency missions, allow the optimisation<br />
of the image acquisition cycle for the marine application mission.<br />
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marine<br />
fire<br />
disaster<br />
Figure 3.1-2: Optimisation of marine application is possible, since parallel fire / disaster monitoring<br />
missions assumed. As for these missions 2°-manoeuvres are considered, an<br />
optimisation without any additional manoeuvres is possible.<br />
The choice of the Field of View of one image is driven by the trade-off between resolution and the<br />
size of image FoV (constrained by detector technology).<br />
The choice of image FoV size directly influences observation scenarios by the number of required<br />
manoeuvres and image takes to cover the observation area (for marine application). Actually, the<br />
observation scenario (i.e. number of parallel fire / disaster monitoring missions per time, see below) is<br />
not a fixed user requirement. In general, the marine application mission asks for a large FoV and<br />
medium resolution, whereas high resolution is first priority for the disaster monitoring mission.<br />
For fire, disaster monitoring and oil slick / erosion missions following product FoVs are specified:<br />
Table 3.1-2: Product FoV requirements at SSP<br />
[in km x km at SSP] Threshold Goal<br />
Disaster monitoring 150 x 100 300 x 200<br />
Fire monitoring 50 x 33 100 x 66<br />
Oil slick 100 x 66 500 x 333<br />
Erosion 100 x 66 500 x 333<br />
An effective FoV of 285² km² has been implemented for all missions. Only for the panchro channel of<br />
the disaster mission, mosaic imaging with smaller single FoV sizes is considered.<br />
These FoV sizes consider pointing errors, which reduce the effective FoV compared to the<br />
implemented detector FoV.<br />
The spatial sampling distance (SSD) in N/S direction defined in the mission requirements refers to a<br />
certain latitude on Earth and do not consider the degradation of the N/S-SSD from nadir to higher<br />
latitudes. Depending on the maximum latitude in which the SSD requirement shall be fulfilled, N/S-<br />
SSD at sub-satellite point (SSP) can be derived. For the product requirements, a max. latitude of 52.5<br />
deg, leading to a degradation factor of two has been considered.<br />
The most challenging SSD requirement is for the disaster panchro channel with an SSD in N/S<br />
direction of 5 m (goal) to 50 m (threshold) at SSP. The SSD requirements for the other channels are<br />
more relaxed (between 50 m goal to 500 m threshold).<br />
The absolute geolocation knowledge of each sample of an image observed at one instance shall be<br />
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better than 0.5 SSD at SSP (threshold) / 0.25 SSD at SSP (goal), with a confidence level of 99.73%<br />
over each image. This requirement shall be fulfilled for any images, where land areas are included<br />
(presence of GCP). For images which contain only water areas (no GCP), the requirement can be<br />
relaxed. For these cases, additional AOCS requirements have to be established, assuring a certain<br />
knowledge drift stability between two marine images (see pointing requirements table below).<br />
The absolute geolocation knowledge requirement assures also the image-to-image registration<br />
(knowledge) performance, which is twice (worst case) the absolute geolocation knowledge (1 SSD at<br />
SSP (threshold) and 0.5 SSD at SSP (goal) for a sample of two consecutive images).<br />
The inter-channel co-registration requirement (knowledge accuracy) is 0.3 px (99.73%) between<br />
each two channels (referring to pixel size of channel with worse resolution).<br />
Pointing Requirements<br />
Following pointing requirements have been derived and established for Geo-Oculus:<br />
• Pointing coverage – European area shall be covered nominally, with the potential to cover<br />
the whole Earth disc;<br />
• APE – to acquire a coverage without gaps for the marine applications;<br />
• RPE over integration time – to assure high resolution for the panchro channel;<br />
• PDE over integration time – to limit image post-integration effort;<br />
• PDE for mosaic imaging – to have products without gaps;<br />
• PDE knowledge (reference to images with landmarks) – for marine images without coastline.<br />
Timing Requirements<br />
The timing requirements can be found in detail in RD [3]. Following requirements have been defined:<br />
• Acquisition Delay;<br />
• Timeliness;<br />
• Product Acquisition Time;<br />
• Temporal Co-registration.<br />
Sensing and Instrument Requirements<br />
A set of sensing and instrument requirements have been established for Geo-Oculus, documented<br />
within RD [3]. They are also discussed within the Instrument section of this document. Most of these<br />
requirements are purely instrument related, therefore they are not discussed in detail within this<br />
system chapter. What should be mentioned, is that several system pointing requirements can be<br />
derived from these instrument requirements, which has then be traded on system level (see next<br />
chapter). Following requirements have been established:<br />
• Radiometric Requirements;<br />
• Spectral Accuracy;<br />
• Modulation Transfer Function (MTF);<br />
• Polarisation.<br />
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Ground Segment Requirements<br />
The ground segment requirements can be found in detail in RD [3]. It is based on a centralised Flight<br />
Operations Segment (FOS). Requirements have been specified for:<br />
• Functionalities of FOS;<br />
• S-Band TM/TC Ground Station and X-Band Ground Station for PDT;<br />
• Centralised Payload Data Ground Segment (PDGS);<br />
• Standardised User Portal;<br />
• Centralised processing facilities;<br />
• High-speed communication connections to the PDT receiving stations;<br />
• Interfaces to meteorological service providers for the provision of nowcasting and very short<br />
range forecasting information of cloud coverage.<br />
3.2 Major System Trade-Offs<br />
In this chapter the major system trade-offs performed after the MTR, leading to the proposed Geo-<br />
Oculus baseline, are summarised. They are discussed in detail within the dedicated chapters /<br />
documents.<br />
• Field of View vs. resolution<br />
The combination of instrument FoV and resolution is limited by the detector technology<br />
(number of pixels). Additionally, both the maximum FoV size and the resolution are limited by<br />
the telescope size. Due to the coverage requirements for the marine applications, the choice<br />
is to go for a maximum possible FoV size (300km x 300 km with the proposed telescope<br />
concept), with medium resolution. For disaster monitoring, the best resolution possible with<br />
the proposed telescope concept has been chosen. This leads to a smaller FoV size, which<br />
requires mosaic imaging for disaster monitoring.<br />
• Magnetic Bearing Wheels vs. Electric Propulsion for manoeuvres<br />
MBWs allow high torques and therefore short manoeuvre times. The drawback are the<br />
microvibrations, which are much lower than with ball bearing wheels, but still impact the<br />
image quality. EPS would create no microvibrations, but increase the manoeuvre time due to<br />
the relative small torque, which leads then to a small number of missions. Furthermore, the<br />
propellant demand is significant, especially for EPS with high thrust. As a consequence, it<br />
has been decided to go for the MBW solution.<br />
• Manoeuvre time vs. image post-integration effort<br />
The pointing instability impacts the image quality. With a good pointing stability, no postintegration<br />
(including image motion compensation) is needed, as long as the MTF<br />
requirements are met. The major contributors to pointing instability are microvibrations<br />
(which are not time varying at a short timescale) and solar array oscillations. The solar array<br />
oscillations are a direct function of the waiting time after a manoeuvre. The trade-off is<br />
therefore between a long manoeuvre time, needing no post-integration and shorter<br />
manoeuvre times, requiring post-integration.<br />
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• Image post integration and Inter-channel co-registration<br />
If motion compensation on-board is required, several options are feasible:<br />
− 1. Post-integration with respect to the first frame of each channel on-board and interchannel<br />
co-registration based on landmarks on ground. Major drawback are successive<br />
processing steps involving resampling.<br />
− 2. Post-integration with respect to the first frame of the all channels on-board.<br />
Advantageous is the avoidance of multiple resampling, however it is necessary to<br />
perform image matching between different spectral channels which might lead to a<br />
somewhat reduced matching performance. No inter-channel co-registration processing by<br />
landmarks has to be performed on-ground.<br />
− 3. Read-out of the panchro channel simultaneous to the first frame of each channel.<br />
Post-integration will be performed as for option 1. The panchro channels will provide due<br />
to their high resolution high matching accuracy regarding image motion compensation.<br />
− 4. Read-out of the panchro channel simultaneous to each frame of each channel. Postintegration<br />
and inter-channel co-registration are performed on-board with respect to the<br />
first panchro image acquired. Resampling is therefore applied only once. This would lead<br />
to the best performance for both image motion compensation and inter-channel coregistration.<br />
The choice for one of these options depend highly on the processing capabilities available<br />
on-board. The currently proposed baseline is a motion compensation on-board using only<br />
integer pixel shifts (due to lower processing power), which would not meet the inter-channel<br />
co-registration requirements. Therefore, the inter-channel co-registration processing is based<br />
on landmark processing on-ground.<br />
3.3 Mission Scenarios<br />
The main focus for the analysis of the Geo-Oculus mission scenarios is the combination of a back<br />
ground marine application mission with high coverage needs (European coastlines) and fast revisit<br />
emergency missions. Both kind of missions have to some extend contrary mission requirements (e.g.<br />
need of large FoV for marine and high resolution, combined with high revisit for the emergency<br />
missions.<br />
A balancing between effective (cloudfree) coverage for the marine application mission and number of<br />
emergency missions has to be performed. This depends on:<br />
• cloud coverage statistics;<br />
• importance of emergency missions.<br />
The following diagram shows the correlation between effective coverage and number of emergency<br />
missions schematically. For precise numbers, a detailed cloud coverage analysis would be needed.<br />
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Effective Coverage<br />
100 %<br />
TBD%<br />
TBD%<br />
TBD%<br />
Eff. cov. after 65 images<br />
Optimised pattern<br />
Eff. cov. LEO mission<br />
non-optimised pattern<br />
Number of<br />
Emergency Missions<br />
Theroetical maximum coverage<br />
during one observation cycle<br />
(cloudfree min. once per observation cycle)<br />
65 130 195<br />
10<br />
Improvement of effective<br />
coverage by Geo-Oculus<br />
Number of marine<br />
Images<br />
Figure 3.3-1: Correlation between effective coverage for marine application mission and number of<br />
emergency missions<br />
For the system baseline, a mission scenario with 2.5 times coverage of the European coastlines<br />
(= 165 marine images) has been chosen.<br />
3.3.1 Mission Scenario Baseline<br />
The key parameters for sizing the proposed mission scenario baseline are:<br />
• Manoeuvre time (based on the proposed magnetic bearing reaction wheel baseline);<br />
• Image acquisition time;<br />
• Product FoV for marine applications;<br />
• Number of images for marine applications.<br />
The number of marine missions and parallel emergency missions have to be traded and balanced<br />
against each other. The minimum requirements for Geo-Oculus are:<br />
• Full coverage of European coastlines (about 65 images within 9 hours);<br />
• At least one fire monitoring mission in parallel (10 min revisit time);<br />
• At least one disaster and one oil slick mission in parallel (60 min revisit time).<br />
The time, which is still left can be used for either increase the effective (cloudless) coverage for marine<br />
applications or increase the number of emergency missions. The following table gives an overview on<br />
the used baseline parameters:<br />
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Table 3.3-1: Baseline parameters for Geo-Oculus mission scenarios<br />
Missions: Disaster Fire<br />
Oil Slick /<br />
Erosion<br />
Marine<br />
Observation times Daytime 24 h Daytime Daytime<br />
Observation cycle (min) 60 10 60 540<br />
Product FoV (km E/W x km N/S, at SSP) 300 x 141 285 x 285 285 x 285 285 x 285<br />
Image FoV (km E/W x km N/S, at SSP)<br />
157 x 157<br />
0.25° x 0.25°<br />
300 x 300<br />
0.48° x 0.48°<br />
300 x 300<br />
0.48° x 0.48°<br />
300 x 300<br />
0.48° x 0.48°<br />
APE (orbit + attitude)<br />
+/- 7.5 km<br />
PDE (mosaic imaging) 700m - - -<br />
Number of images per product 3 1 1 1<br />
FoV at nadir<br />
Required manoeuvres<br />
0.25 deg (70 sec) / 0.4 deg (70 sec)* 2 - - -<br />
2 deg (70 sec) 1 1 1 1<br />
Total manoeuvre time (sec) 210 70 70 70<br />
Single image acquisition time (sec)<br />
panchro: 0.4<br />
others: 7.7<br />
1.2 24.6 24.6<br />
Total image acquisition time (sec) 8.1 1.2 24.6 24.6<br />
Total time per one mission (sec) 218 71 95 95<br />
Number of channels<br />
SSD (m E/W x m N/S, at SSP)<br />
13 5 21 21<br />
Panchro 21 x 10.5<br />
-<br />
UV-VNIR 40 x 20 40 x 20 80 x 40 80 x 40<br />
MWIR, SWIR -<br />
150 x 150<br />
TIR -<br />
375 x 375<br />
Product data amount (Mbits) 3.86E+04 3.95E+03 3.30E+04 3.30E+04<br />
* 1.2 deg with Korsch configuration<br />
Based on the mission scenario, the number of manoeuvres (and images) per day have been<br />
determined and a schematic mission schedule is shown in Figure 3.3-2. For this missions schedule,<br />
the 10 min repeat cycle for the fire monitoring is the sizing parameter.<br />
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Fire<br />
2° Manoeuvre<br />
Fire<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Disaster panchro<br />
0.25° Manoeuvre<br />
Disaster panchro<br />
0.4° Manoeuvre<br />
Disaster other channels<br />
Figure 3.3-2: Mission schedule<br />
2° Manoeuvre<br />
Margin<br />
Margin<br />
Fire<br />
2° Manoeuvre<br />
Fire<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Oil slick<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Marine<br />
2° Manoeuvre<br />
Margin<br />
Mission Schedule:<br />
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Manoeuvre Times and Image Acquisition Times<br />
The manoeuvre times can be separated into two contributors:<br />
• Actual manoeuvre time;<br />
• S/A tranquillisation time.<br />
The actual manoeuvre time depends on the size of manoeuvre, chosen technology (wheel size, etc.)<br />
and manoeuvre strategy.<br />
The S/A tranquillisation time depends mostly on the pointing drift required for the image acquisition<br />
(the more challenging the requirement, the longer the tranquillisation time). An allocation of 70<br />
seconds per manoeuvre has been considered.<br />
3.4 Cloud Coverage Analysis<br />
Two of the key features of Geo-Oculus, the possibility for real-time commanding and the capability for<br />
short revisit cycles have been found to give an essential asset in order to maximise the mission<br />
performance - the optimisation of mission planning for cloud cover. The intention of this analysis has<br />
been to identify the potential of Geo-Oculus that can be gained, to validate the system requirements,<br />
to identify a possible optimisation strategy and to assess the performance compared to reference<br />
missions.<br />
Due to its geostationary orbit, Geo-Oculus has the capability to access every spot within its footprint at<br />
the time the spot becomes cloud free. Considering the applied FOV and the possible agility of the<br />
system, this capability confined. In result only a certain image acquisition frequency is achieved;<br />
hence a selection of the images is required. This leads to the point that the system will have to apply a<br />
permanently updated optimisation of the mission planning, to gain maximum possible ground<br />
coverage. This optimisation should take into account the current cloud cover situation, possibly<br />
supplied by MTG and Metop, the changing illumination conditions throughout the entire day, nowcasting<br />
and short range forecasting information on the expected cloud cover situation and the<br />
constraints placed by the on-demand missions.<br />
In the analysis described in here, a simplified optimisation strategy and mission planning have been<br />
used, considered to represent a realistic approach. This strategy accounts for the illumination<br />
conditions and maximises the achieved ground coverage.<br />
The entire cloud coverage analysis is based on cloud mask data from MSG with a revisit time of 15<br />
min. The time span, considered in this analysis range from 01/2004 to 05/2007. In a preliminary low<br />
level analysis representative days for a detailed evaluation of the cloud coverage are filtered out of the<br />
complete dataset. To gain representative results from the analysis, representative days are indicated<br />
by analysing every day concerning:<br />
• Cloud amount<br />
• Cloud coverage changes<br />
• Time of sufficient illumination conditions<br />
By comparing the values of each day with the mean value of the whole data set, several days for<br />
detailed analyses have been indicated.<br />
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Referring to these days the detailed cloud coverage analysis is conducted. It comprises four stages:<br />
• Analysis of evolution of geometrical conditions through the day (Illumination situation)<br />
• Analysis of cloud coverage amount and dynamics<br />
• Evaluation of the performance of Geo-Oculus for different system set ups<br />
• Comparison of the performance provided by Geo-Oculus with other planned EO Systems<br />
The analysis of geometrical conditions regards especially the system requirements on the solar zenith<br />
angle and the view zenith angle. For Geo-Oculus the view zenith angle of every region is constant all<br />
times. By contrast, the solar zenith angle, hence the illumination condition changes through the day<br />
and is depended to the season. To consider this in the analysis, the illumination conditions are<br />
calculated for each cloud mask file by computing VZA and SZA for each pixel. These information are<br />
one necessary input for the simplified mission planning, applied in the performance evaluation of Geo-<br />
Oculus.<br />
The second necessary input information are evaluated in the analysis of cloud coverage amount and<br />
dynamics. Herein the cloud mask data is evaluated concerning cloud amount and cloud coverage<br />
changes through one day.<br />
• Cloud amount is defined as how long a pixel was clouded in the time between 06.00 UTC<br />
and 18.00 UTC. It is provided in percent. Analysing the cloud amount allows to point out<br />
areas, where observation is possible, in general.<br />
• Cloud cover changes is defined as number of changes of a pixel from clouded to unclouded<br />
or vice-a-versa within the considered time-frame (06.00 UTC to 18.00 UTC) in the 15 min<br />
time interval of the MSG data. Since 49 cloud mask files are available in this time-frame, a<br />
maximum of 48 cloud cover changes can occur.<br />
With the evaluation of the cloud coverage changes, the dynamics of the cloud situation are indicated.<br />
With the help of this, it is possible to point out areas where (nearly) cloud free products can be<br />
generated, by acquiring the same area several times, as it is possible with Geo-Oculus. Some results<br />
of this analysis are to be seen in figure 3.4-1.<br />
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Cloud Amount [%]<br />
Number of Cloud Coverage Changes<br />
Figure 3.4-1: Cloud amount and Cloud coverage changes during 6.00UTC and 18.00 UTC at<br />
30.09.2005<br />
The plots in figure 3.4-1 shows cloud amount and cloud coverage changes other Europe. It is to be<br />
seen, that nearly complete Europe and its coastlines are clouded at least 50% of the day (1 st plot). But<br />
there are also a lot of areas within Europe or its coastlines, where the cloud situation changes during<br />
the day (2 nd plot). One can assume, that observations allowing only one acquisition per day for a<br />
certain area, as it is provided by a LEO system, will lead to a rather small ground coverage. The<br />
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ground coverage can be increased by acquiring the same spot several times, as already mentioned.<br />
This is useful especially in areas, where high cloud coverage dynamics occur, hence clouds are<br />
moving or dissolving a lot. In these areas, indicated by a high amount of cloud coverage changes,<br />
Geo-Oculus can increase the ground coverage through multiple acquisitions of the same areas. In<br />
result Geo-Oculus provides a higher performance in means of ground coverage than LEO systems.<br />
These results have been correlated with the impacts of changing illumination situation, to indicate the<br />
areas where the ground coverage can be increased most by observations with Geo-Oculus for the<br />
handled day.<br />
The performance evaluation of Geo-Oculus has been conducted for the Marine Applications mission,<br />
which is accomplished as background mission. The results are considered to be representative for<br />
these missions and also illustrate the capacity of the system for on Demand missions in the sea areas,<br />
like Oil Slick Monitoring. For the Marine Applications, an image pattern has been implemented. The<br />
simplified mission planning considers that the acquisition sequence is updated immediately when an<br />
update on the cloud coverage information becomes available to the system; hence with every cloud<br />
mask file (one new cloud mask file every 15 min) the mission plan is optimised and updated.<br />
According to the agility of the system a certain number of acquisitions is possible within 15 min and a<br />
selection of the images observed within the next 15 min has to be accomplished. The number of<br />
acquisitions within 15 min is also depended to the number of parallel on-demand missions, which have<br />
to be accomplished. The selection is based on the results of the analysis of geometrical conditions<br />
and of cloud amount and cloud coverage changes. The current baseline foresees 4 images per 15<br />
min. The final product is achieved by combining all the acquired images. In the analysis the<br />
combination of the images leads to the total observed area which identifies the performance of Geo-<br />
Oculus.<br />
<strong>Final</strong>ly the cloud coverage analysis compares the performance of Geo-Oculus with LEO Systems like<br />
Sentinel 2 and Sentinel 3, regarding the total ground coverage which can be achieved at the handled<br />
day. For this, different LEO swaths are implemented and superposed with the same cloud mask data,<br />
as used for the performance evaluation for Geo-Oculus. Image 3.4-2 shows the performances of Geo-<br />
Oculus and LEO systems by highlighting the ground coverage for one day:<br />
Figure 3.4-2: Ground coverage within one day for Geo-Oculus (left) and LEO (Sentinel 3, right) are<br />
highlighted blue<br />
It can be seen that Geo-Oculus provides considerably more ground coverage (~83,3% of the<br />
maximum possible coverage, ~55% effective) than a LEO system (Sentinel 3 ~35% of the maximum<br />
possible coverage,~23% effective). This is due to the fact that the area for observation is accessible to<br />
Geo-Oculus the whole day, whereas a sun-synchronous LEO mission provides commonly ~3 passes<br />
over Europe during one day. This provides Geo-Oculus the advantage to benefit already from dynamic<br />
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cloud scenes where the cloud coverage, although the spots might feature high cloud amount. The only<br />
restrictions for observations on Geo-Oculus, are areas clouded the whole day with no cloud<br />
movement.<br />
Conclusion<br />
Geo-Oculus has been found to provide the best possible ground coverage at high resolution with a<br />
significant improvement compared to LEO-missions. The achievable ground coverage with Geo-<br />
Oculus at ~40 m GSD over Europe is ~2.5 times more than Sentinel 3 at 300 m GSD. This advantage<br />
results from the swath width, the orbit geometry of LEO-missions which results in three swaths per day<br />
over Europe at fixed local times and on the other side the capability of Geo-Oculus to access whole<br />
Europe and to pick the cloud free points in time. The unique feature of multiple acquisitions and near<br />
real time mission plan updating is bund to geo-synchronous missions and can not be provided by<br />
LEO-missions.<br />
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4 Mission and System Level Analyses<br />
4.1 Mission Architecture<br />
A visualisation of all elements contributing to the Geo-Oculus mission architecture is shown in Figure<br />
4.1-1.<br />
Figure 4.1-1: Mission Architecture<br />
4.2 Mission Analysis<br />
Mission analyses issues have already been traded in [RD 7] for the following topics:<br />
• type of orbit,<br />
• orbit inclination,<br />
• orbit determination performance,<br />
• orbit transfer and launcher.<br />
The preferred mission parameters which were assumed for the subsequent analyses of technical<br />
solutions for the spacecraft are summarised hereafter.<br />
Type and inclination of orbit<br />
A geostationary orbit with 0 degree inclination is the suggested baseline. It requires only one<br />
spacecraft and provides constant observation conditions (view Zenith angle) but it is linked to about<br />
300 kg of fuel consumption to be assigned for North South station keeping.<br />
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Alternative solutions which have been investigated are:<br />
• geostationary orbit with limited station keeping,<br />
• inclined geosynchronous orbit,<br />
• Molniya orbit (highly elliptical inclined orbit).<br />
The appealing features of these alternatives are on one hand potential propellant savings for station<br />
keeping and on the other hand improved viewing conditions over Europe.<br />
The maximum propellant saving may be achieved for a geostationary orbit with 7.5 deg inclination and<br />
0 deg initial right angle of ascending node. This orbit is rather stable such that no orbit corrections are<br />
needed. However, as it is the case for all inclined geosynchronous orbits, the orbit trace projected onto<br />
the Earth surface is a figure of eight which the satellite passes through once per orbit. This yields that<br />
the satellite is half an orbit over Northern latitudes (with better viewing conditions over Europe) and the<br />
other half orbit it is over Southern latitudes. If no corrections of the orbital plane are performed, the<br />
local time of the equator crossing will change over the year. This means that the satellite is at the most<br />
Northern position e.g. at 12:00 at a certain day of the year but half a year later this position is achieved<br />
at midnight. Hence, the good viewing conditions at daytime are achieved for a certain part of the year<br />
only and become even worse for the other part of the year. Since this feature is a significant constraint<br />
for the mission flexibility, all alternatives with inclined orbits have currently been dropped.<br />
Orbit determination performance<br />
The achievable precision for the position of the spacecraft is important since it contributes to the<br />
overall pointing budget which is rather stringent. The preferred solution with the best performance is a<br />
ranging technique based on 2 or 3 ground stations and using a spread spectrum method for the<br />
ranging signal. The orbit determination accuracy is 100 m to 150 m (3σ) along track and 10 m to 20 m<br />
(3σ) across track. The interesting feature is that this technology shows the same performance right<br />
after a manoeuvre when using 3 ground stations. Moreover, this technique is routinely used by SES<br />
Astra which gives strong evidence that it can be successfully applied to Geo-Oculus.<br />
The following alternative technologies have also been investigated:<br />
• single ground station ranging + line of sight measurements,<br />
• dual ranging,<br />
• DARTS,<br />
• short baseline interferometry,<br />
• long baseline interferometry,<br />
• optical telescope,<br />
• use of landmarks,<br />
• GPS.<br />
The GPS option is currently being investigated for geostationary orbits. It may provide a similar<br />
nominal performance as the ground based spread spectrum method but a significant degradation is<br />
expected for a couple of hours after a manoeuvre. All performance data for above listed options are<br />
found in [RD 7].<br />
Orbit transfer and launcher<br />
A geostationary transfer orbit strategy is suggested as baseline. The initial elliptical orbit provided<br />
by the launcher is changed to the final circular orbit with a liquid apogee engine installed in the<br />
spacecraft. The achievable maximum mass of the spacecraft in the final orbit is slightly higher<br />
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compared to a strategy where the launcher provides a direct injection. Moreover, this strategy is a<br />
common standard with a high level experience whereas the direct injection is offered by few launchers<br />
only.<br />
Depending on the final launch mass of the spacecraft, the preferred launchers are Soyuz from<br />
Kourou (up to ~3 tons) and Ariane 5 (more than 3 tons). For alternatives of non-European launch<br />
service providers see next table.<br />
Table 4.2-1: Launcher Survey: Standard Launch into GTO including Performance<br />
Injection into GTO Launch European Perf. into ΔV to GSO Remark<br />
Service LSP GTO [m/s]<br />
Launcher Name Provider<br />
[kg]<br />
Ariane 5 ECA Arianespace Y 9000 1500 flight qualified<br />
Soyuz Fregat / Kourou Arianespace Y 3000 1480 under development<br />
Soyuz Fregat / Baikonur STARSEM Y 1840 1500 flight qualified<br />
Atlas 5 ILS N 8670 1804 flight qualified<br />
Delta 2 Boeing N 2120 1840 flight qualified<br />
Delta 4M Boeing N 6470 1800 not commercially available<br />
Delta 4H Boeing N 10819 1800 not commercially available<br />
Proton * ILS N 5530 1500 flight qualified<br />
Sea Launch Sea Launch N 5850 1500 flight qualified<br />
Land Launch * Sea Launch N 3600 1500 flight qualified for direct GEO<br />
GSLV Antrix N 2400 1650 flight qualified up to 2t<br />
H-2A MHI N 6000 1840 flight qualified up to 5t<br />
Long March 3B CGWIC N 5000 1840 flight qualified<br />
Falcon 9 Space X N 5070 TBD under development<br />
Angara 3 * ILS N 2400 1500 under development<br />
Angara 5 * ILS N 5400 1500 under development<br />
* performance for S/C + adapter; all other launchers are S/C separated masses<br />
4.3 Payload<br />
4.3.1 Imaging capability<br />
In order to support Fire Monitoring & Marine applications, the instrument provides simultaneous<br />
imaging of Earth scenes on four multi-spectral focal planes (UV-blue, Red-NIR, MWIR and TIR) with a<br />
ground FoV of 300x300 km (0.48x0.48 deg). The spectral channels are defined in the following figure<br />
together with the achieved ground resolution over Europe (worst case given at 52.5 °N corresponding<br />
to a viewing zenith angle of 60 deg). For some channels (e.g. for IR ones), subscript "a" refers to Fire<br />
Monitoring mission while "b" refers to Marine application and corresponds to different radiometric<br />
requirements (e.g. SNR & typical radiance). The VNIR resolution for marine applications is 80 m, twice<br />
that of other missions because pixel binning is necessary to meet the challenging SNR requirements<br />
of these applications (see next section).<br />
In addition, the Disaster Monitoring applications requires a VIS panchromatic (PAN) focal plane with<br />
higher resolution (10.5 m nadir, 21 m over Europe) and reduced FoV (157x157 km, i.e. 0.25x0.25 deg)<br />
imposed by the use of the same detector array as the UV-blue & Red-NIR channels. The PAN channel<br />
is separated in the field from the other channels.<br />
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The achieved ground resolution & coverage are generally between threshold (T) and goal (G)<br />
requirements, with the following exceptions:<br />
• VNIR channels have a resolution (40 m & 80 m for marine) better than the G requirement<br />
• T requirement is not met in TIR for Fire Detection due to sensor size limitation. This is deemed<br />
acceptable because the MWIR band is actually used for fire area monitoring, whereas TIR<br />
bands are used to monitor fire temperature, for which resolution is not critical.<br />
Channel ID Center<br />
wavelength<br />
Bandwidth Focal planes<br />
(nm) (nm)<br />
UV1 318 10<br />
UV2 350 10<br />
VNIR1 412 10<br />
VNIR2 443 10 UV-blue<br />
VNIR3 490 10<br />
VNIR4 510 10<br />
VNIR5 555 10<br />
VNIR7 655 155 PAN<br />
VNIR6 620 10<br />
VNIR8a 665 10<br />
VNIR8b 665 10<br />
VNIR9 681 8<br />
VNIR10 709 10<br />
VNIR11<br />
VNIR12<br />
753<br />
779<br />
8<br />
15<br />
Red-NIR<br />
VNIR13a 865 20<br />
VNIR13b 865 20<br />
VNIR14 885 10<br />
VNIR15 900 10<br />
VNIR16 1040 40<br />
SWIR 1375 50<br />
MWIRa 3700 390 SWIR MWIR<br />
MWIRb 3700 390<br />
TIR1a 10850 900<br />
TIR1b<br />
TIR2a<br />
10850<br />
12000<br />
900<br />
1000<br />
TIR<br />
TIR2b 12000 1000<br />
Mission<br />
Disaster<br />
Monitoring<br />
Fire<br />
Monitoring<br />
Marine<br />
Applications<br />
Figure 4.3-1: Spectral channels (optional channels are in blue) & imaging capability summary<br />
4.3.2 Radiometric & image quality performances<br />
Ground Pixel Size at<br />
52°N [m]<br />
Image ground<br />
coverage [square, km]<br />
T G T G<br />
Disaster monitoring 100 10 100 200<br />
Fire monitoring 250 100 100 200<br />
Marine applications 1000 100 100 500<br />
Ground resolution & coverage requirements<br />
Channels<br />
Number of<br />
channels<br />
GSD (m) at<br />
52.5°N<br />
FOV (km)<br />
PAN 1 21.0 157x157<br />
UV-blue<br />
Red-NIR<br />
4<br />
8<br />
40<br />
40<br />
300x300<br />
Red-NIR 2 40<br />
SW/MW IR 2 300 300x300<br />
TIR 2 750<br />
UV-blue 7 80<br />
Red-NIR<br />
SW/MW IR<br />
10<br />
2<br />
80<br />
300<br />
300x300<br />
TIR 2 750<br />
4.3.2.1 Disaster Monitoring<br />
The high resolution PAN channel drives telescope diameter (set to 1.5 m) & pointing stability<br />
requirements. The acquisition is performed in 4 successive images, all downloaded for on-ground<br />
processing for SNR & MTF recovery. Indeed, the Nyquist MTF requirement for the raw images is<br />
relaxed from 10% to 5% to allow best resolution. The required SNR is increased in the same ratio to<br />
keep constant the SNRxMTF figure of merit to account for SNR degradation in MTF recovery by<br />
ground processing.<br />
Thanks to that, a Ground Sampling Distance (GSD) as good as 10.5 m (nadir), i.e. 21 m at 52.5 N is<br />
achieved, close to Goal requirement (10 m) for Disaster Monitoring. This resolution is achieved with<br />
LoS pointing stability requirements of 5 µrad/s and 0.15 µrad p-p, well within the capability of the<br />
selected AOCS design based on magnetic reaction wheels. As shown in §4.4, selecting conventional<br />
ball bearing wheels mounted on isolators (jitter increased 0.25 µrad p-p) has however a modest<br />
impact, with nadir GSD degraded to 11.5 m (23 m at 52.5 N).<br />
The same CMOS detectors are used for all UV-VNIR focal planes to reduce development cost & risks.<br />
The GSD of the UV-blue & Red-NIR channels is then obtained from the PAN GSD in the ratio of the<br />
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FoV (300/157 = 1.9), i.e. 20 m (nadir). Since no SNR requirements were derived from user needs, a<br />
value of 150 was selected for all VNIR channels, consistently with PAN channel and Sentinel 2<br />
requirements (around 150 and up to 170 for some channels).<br />
The instrument parameters for each channel are summarised in Figure 4.3-2. The GSD at the<br />
maximum latitude where full performance is required is recalled in column 3.The MTF at Nyquist<br />
frequency in is given in column 4, showing that the 10% requirement is met for all multispectral<br />
channels. For the PAN channel, despite relaxation to 5%, the requirement is met without much<br />
margin. This clearly demonstrates that the maximum resolution achievable with a 1.5 m telescope is<br />
reached for this channel (10.5 m nadir).<br />
1 2 3 4 5 6 7 8 9<br />
Mission Channel<br />
GSD at 52°N<br />
(m)<br />
MTF at<br />
Nyquist<br />
<strong>Final</strong><br />
SNR<br />
Nb<br />
of images<br />
Post<br />
Integration<br />
Integ. Time<br />
of image (s)<br />
Channel acq.<br />
time (s)<br />
Disaster VNIR2 40 0.117 150 1 No 0.009 0.088<br />
Disaster VNIR3 40 0.116 150 1 No 0.011 0.088<br />
Disaster VNIR4 40 0.115 150 1 No 0.014 0.088<br />
Disaster VNIR5 40 0.111 150 1 No 0.017 0.088<br />
Disaster VNIR6 40 0.110 150 1 No 0.021 0.088<br />
Disaster VNIR7 21 0.051 300 4 No 0.013 0.352<br />
Disaster VNIR10 40 0.111 150 1 No 0.029 0.088<br />
Disaster VNIR11 40 0.115 150 1 No 0.005 0.088<br />
Disaster VNIR12 40 0.112 150 1 No 0.034 0.088<br />
Disaster VNIR13b 40 0.113 150 1 No 0.034 0.088<br />
Disaster VNIR14 40 0.098 150 1 No 0.073 0.088<br />
Disaster VNIR15 40 0.117 150 1 No 0.011 0.088<br />
Disaster VNIR16 40 0.110 150 1 No 0.027 0.088<br />
Figure 4.3-2: Performances for Disaster Monitoring<br />
The SNR obtained after accumulation of the number of successive images indicated in column 6 is<br />
given in column 5. All multi-spectral channels can be acquired in a single image, i.e. without postintegration,<br />
while keeping good image quality (MTF at Nyquist > 10%, see column 4). The integration<br />
time of individual raw images is given in column 8. The time to acquire the channel (last column) is<br />
obtained by multiplying the number of images by the largest value between the integration time and<br />
the array readout time.<br />
4.3.2.2 Marine applications<br />
For marine applications with challenging SNR requirements, 2x2 pixel binning is used to increase the<br />
collected signal, so the final ground resolution is 40 m nadir and 80 m at 52.5°N, well within the goal<br />
requirement of 100 m. Thanks to this lower resolution, the requirements on the pointing stability during<br />
imaging periods can be relaxed to 10 µrad/s, allowing to largely reduce the tranquilisation time<br />
following a slew manoeuvre. For channels with high SNR requiring long exposure time, the image<br />
acquisition is split in several successive images to avoid pixel saturation & image smear due to<br />
pointing drift. These images are summed on-board, with for the most critical ones (UV-blue channels)<br />
compensation of the image motion (so-called "post-integration").<br />
The instrument parameters for each channel are summarised in Figure 4.3-3. The MTF at Nyquist<br />
requirement (10%) is met with good margins for all channels. Post-integration with LoS motion<br />
compensation is only required for the bands of the UV-blue focal plane (UV1 to VNIR5). Other bands<br />
require several images to avoid saturation at typical flux, these images are simply added in the on-<br />
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board image processing. Column 7 gives the number of images requiring post-integration with LoS<br />
motion compensation. For instance, channel VNIR1 requires an accumulation of 18 images (column 6)<br />
with image motion compensation between 6 packets (column 7) of 3 simply accumulated images.<br />
Channel UV1 requires post-integration with motion compensation of 30 successive images (same<br />
value in columns 6 & 7).<br />
1 2 3 4 5 6 7 8 9<br />
Mission Channel<br />
GSD at 52°N<br />
(m)<br />
MTF at<br />
Nyquist<br />
<strong>Final</strong><br />
SNR<br />
Nb<br />
of images<br />
Post<br />
Integration<br />
Integ. Time<br />
of image (s)<br />
Channel acq.<br />
time (s)<br />
Marine UV1 80 0.156 1000 30 30 0.082 2.637<br />
Marine UV2 80 0.162 1000 10 10 0.083 0.879<br />
Marine VNIR1 80 0.146 1500 18 6 0.035 1.582<br />
Marine VNIR2 80 0.161 1300 15 5 0.031 1.318<br />
Marine VNIR3 80 0.154 943 6 3 0.050 0.527<br />
Marine VNIR4 80 0.175 748 6 3 0.041 0.527<br />
Marine VNIR5 80 0.171 557 2 2 0.085 0.176<br />
Marine VNIR6 80 0.144 418 2 No 0.059 0.176<br />
Marine VNIR8b 80 0.159 376 1 No 0.105 0.105<br />
Marine VNIR9 80 0.140 339 1 No 0.118 0.118<br />
Marine VNIR10 80 0.165 323 1 No 0.098 0.098<br />
Marine VNIR11 80 0.216 478 2 No 0.019 0.176<br />
Marine VNIR12 80 0.186 258 1 No 0.073 0.088<br />
Marine VNIR13b 80 0.193 213 1 No 0.051 0.088<br />
Marine VNIR14 80 0.137 213 1 No 0.108 0.108<br />
Marine VNIR15 80 0.201 259 1 No 0.023 0.088<br />
Marine VNIR16 80 0.164 250 1 No 0.055 0.088<br />
Marine SWIR 300 0.271 250 1 No 4.19E-04 0.025<br />
Marine MWIRb 300 0.095 378 1 No 0.017 0.025<br />
Marine TIR1b 750 0.097 1818 1 No 5.64E-04 0.016<br />
Marine TIR2b 750 0.102 2001 1 No 5.46E-04 0.016<br />
Figure 4.3-3: Performances for Marine applications<br />
4.3.2.3 Fire Monitoring<br />
Fire monitoring is based on three IR channels, a MWIR channel with 300 m resolution to monitor the<br />
fire area & location and two TIR channels with 750 m resolution to measure fire temperature.<br />
Two VNIR channels with moderate SNR are also required, for which 40m resolution is possible without<br />
post-integration if VNIR13a SNR requirement is relaxed from 213 (user requirement) to 150. This<br />
value is assumed to simplify the image acquisition scheme.<br />
1 2 3 4 5 6 7 8 9<br />
Mission Channel<br />
GSD at 52°N<br />
(m)<br />
MTF at<br />
Nyquist<br />
<strong>Final</strong><br />
SNR<br />
Nb<br />
of images<br />
Post<br />
Integration<br />
Integ. Time<br />
of image (s)<br />
Channel acq.<br />
time (s)<br />
Fire VNIR8a 40 0.112 80 1 No 0.017 0.088<br />
Fire VNIR13a 40 0.103 150 1 No 0.062 0.088<br />
Fire MWIRa 300 0.095 233 1 No 2.62E-05 0.025<br />
Fire TIR1a 750 0.097 670 1 No 4.72E-05 0.016<br />
Fire TIR2a 750 0.102 736 1 No 5.05E-05 0.016<br />
Figure 4.3-4: Performances for Fire Monitoring<br />
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4 <strong>Final</strong><br />
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4.3.3 Instrument design<br />
4.3.3.1 Optical design<br />
The instrument is based on a 1.5 m diameter all-SiC monolithic telescope, i.e. the same size as<br />
Aeolus/ALADIN, formed by M1 & M2 mirrors in Cassegrain configuration. The PAN channel that<br />
requires a long focal length is imaged by a Korsch telescope formed with a third converging mirror<br />
following a flat folding mirror placed out of the M1-M2 axis. The other channels are separately imaged<br />
by the Cassegrain telescope formed by M1 & M2 mirrors (as shown in Figure 4.3-5), or alternately by<br />
the symmetric Korsch configuration. A first dichroïc plate is used to separate the UV-blue & Red-NIR<br />
channels from the IR channels, and within each group a second dichroïc plate provides separation<br />
between the focal planes. Four filter wheels (one for each focal plane) are used to select the narrow<br />
channels in each band. Cold stops are required in front of the IR focal planes which need to controlled<br />
at low temperatures (130 K for MWIR and 50 K for TIR).<br />
M2<br />
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M1<br />
M3<br />
PAN high resolution<br />
detector<br />
SW/MW IR<br />
detector &<br />
filter wheel<br />
Figure 4.3-5: Multi-spectral imaging telescope optical architecture<br />
Field correction<br />
& focal length<br />
adjustment<br />
TIR detector<br />
& filter wheel<br />
UV-Blue detector<br />
& filter wheel<br />
Red-NIR detector<br />
& filter wheel<br />
The retained optical combination is based on three Korsch combinations with dioptric correctors (PAN,<br />
UV-VNIR, IR) for focal length adjustment & aberration correction in the large FoV. Several folding<br />
mirrors are required to accommodate the five channels in the volume below the M1 mirror.<br />
From M2<br />
TIR<br />
MWIR<br />
PANCHRO<br />
RED-NIR<br />
UV-BLUE<br />
Figure 4.3-6: Optical configuration (after M2)<br />
View from SVM
4 <strong>Final</strong><br />
<strong>Report</strong><br />
4.3.3.2 UV-VNIR focal planes<br />
For UV and VNIR spectral bands (from 315 to 1040 nm), silicon semi-conductor is the only material<br />
considered, thanks to its large maturity, its lower cost, its ability to build large format arrays and its low<br />
dark current at ambient temperature. Monolithic CMOS array is preferred to CCD for its good maturity<br />
to build large arrays, its better immunity to GEO harsh radiation environment (as shown by<br />
GOCI/COMS CMOS detector qualification) and because smearing during transfer rules out large CCD<br />
arrays for Earth observation.<br />
Monolithic CMOS imagers manufactured with processes optimised for imaging applications (so-called<br />
“CIS”) look by far as the most promising technology for Geo-Oculus. Indeed, thanks to improved<br />
photodiode processes, CIS arrays are featuring excellent electro-optics performances even for small<br />
pixel pitches. Thinned backside CMOS technology is considered to improve the fill factor and therefore<br />
the detection efficiency. Back-illuminated CMOS are already available in the USA and are being<br />
investigated in Europe, so availability in a 5-year frame is very likely.<br />
Considering only mandatory channels, the spectral range to be covered by the CMOS detector (0.4<br />
0.9 µm) is compatible with a conventional “broadband” detector. When accounting for optional<br />
channels, it is impossible to have a good detection efficiency in the large spectral range to be covered<br />
by the CMOS detector (0.315 to 1.04 µm), so two detectors are used, one optimised for UV & short<br />
visible wavelengths (“UV-blue” detector) and the other for red & NIR wavelength (“Red-NIR” detector).<br />
The largest space CMOS arrays currently available in Europe are in the range of 1.5k x 1.5k pixels<br />
(e.g. COBRA2M 2 Mpixels detector developed by Astrium & ISAE/CIMI for the COMS/GOCI<br />
instrument). The next step will be 3k x 3k arrays, expected for 2010-2011. There is therefore a major<br />
step to be performed in order to reach the typical 100 Mpixel array size required for Geo-Oculus. Even<br />
though there is no strict limitation in CMOS array size, such large arrays set many technical<br />
challenges. In particular, the manufacturing process shall be well mastered to guaranty a sufficient<br />
yield, i.e. a reasonable cost & development schedule. Moreover, to build a very large format array<br />
without dead zone with good electro-optics performances and not to constraint too much the planarity,<br />
stitching (i.e. gap-less array assembly during manufacturing process) will be required (see Figure<br />
4.3-7) since the total array will not fit within the stepper field (currently limited to 22x22 mm²).<br />
– Sub-blocks are exposed one after<br />
another<br />
– Some blocks are used multiple<br />
times<br />
– Ultimate limit is given by wafer size<br />
22mm<br />
V<br />
1<br />
Stepper field<br />
horiscan2<br />
horiscan1<br />
V V<br />
2 3<br />
array<br />
Stitched CMOS Sensor<br />
horiscan1 horiscan2<br />
array array array<br />
array array array<br />
array array array<br />
Figure 4.3-7: Wafer-level stitching is used to build arrays with size larger than the stepper field<br />
As for the 2 Mpixels detector of the GOCI instrument, it is proposed that the array is divided 4 subblocks<br />
independently operated, offering a redundancy level in case of failure (see Figure 4.3-8). Each<br />
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V<br />
1<br />
V<br />
2<br />
V<br />
3
4 <strong>Final</strong><br />
<strong>Report</strong><br />
sub-block has16 video outputs with 20 Mpixels/s data rate, allowing reading the total array in less than<br />
100 msec.<br />
Column decoder<br />
16 outputs 16 outputs<br />
Column readout circuit<br />
4 stitched 25 Mpix arrays<br />
Column readout circuit<br />
16 outputs 16 outputs<br />
Column decoder<br />
Figure 4.3-8: Architecture for the CMOS detector for PAN , UV-Blue & Red-NIR focal planes<br />
Each detector is interfaced with a Proximity Electronics Module (PEM) housing all the functions<br />
requiring to be located close to the detector, i.e. detector sequencing (e.g. clock generation), bias<br />
voltage supply and video signal pre-amplification.<br />
4.3.3.3 MWIR focal plane<br />
The selected technological approach is that photo-detectors arrays are manufactured by using the<br />
adequate detection material and hybridised on top of a CMOS Read Out Integrated Circuit (ROIC).<br />
The ROIC is in charge of providing the reference bias voltage to each photo-detector, injecting the<br />
signal at the output of the photo-detector within the corresponding integration capacitance and<br />
multiplexing the analogue signals from all the pixels through a reduced number of outputs.<br />
AlGaAs/GaAs or InGaAs QWIP technology was initially preferred to HgCdTe for its better yield,<br />
operability and uniformity. QWIP main drawback is its lower sensitivity. However, as MWIR integration<br />
time is low with respect to read out time, the sensitivity should not be the driver of the choice, whereas<br />
cost, stability, cosmetics and uniformity are important Nevertheless, this choice had to be<br />
reconsidered when an additional SWIR channel had to implemented. Indeed, QWIP technology does<br />
not allow wide-band detectors with good detection efficiency from 1.3 to 3.7 µm, so a dedicated SWIR<br />
focal plane would be required. The right choice is then HgCdTe technology which allows such a<br />
combined SWIR/MWIR detector with good detection performances, but also with the yield drawbacks<br />
pointed out above.<br />
Driven by the minimum pixel pitch that can be achieved for European indium bump hybridized<br />
detectors (i.e. 15 µm), the 30x30 mm 2 area of the 2k x 2k photo-detector array assumed for Geo-<br />
Oculus is larger than the today European state of the art (20 mm diagonal for QWIP and 25 mm for<br />
HgCdTe, but seems reachable within a few years provided the necessary pre-developments are<br />
performed. The CMOS ROIC associated to the photo-detector array is another challenge. Stitching will<br />
be required, 30x30 mm 2 being larger than the stepper field.<br />
The MWIR detector architecture is similar to the UV-VNIR one (see Figure 4.3-9).<br />
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Column decoder<br />
4 outputs 4 outputs<br />
Column readout circuit<br />
4 stitched 1k x 1k arrays<br />
15 µm pixel pitch<br />
Column readout circuit<br />
4 outputs 4 outputs<br />
Column decoder<br />
Figure 4.3-9: Architecture for the MWIR CMOS hybrid detector<br />
The 2k x 2k array with 15 µm pitch consists in 4 independent 1000x1000 pixels sub-arrays. Each subblock<br />
is read out via 4 video outputs with a 10 Mpixels/s output rate, so the read out period is 25 ms,<br />
much larger than the max. integration time (2.5 msec for Marine application).<br />
The operating temperature of the detector is dictated by the level of dark current, which can reduce<br />
the useful dynamic range and significantly increase the detection noise. A temperature of 130K is a<br />
typical choice for a 3.7 µm MWIR band.<br />
4.3.3.4 TIR focal plane<br />
As for MWIR, the two candidate technologies are TIR HgCdTe and AlGaAs/GaAs QWIP. From a<br />
qualitative point of view, the trade-off between both materials is identical: better sensitivity for HgCdTe<br />
and better yield / uniformity (spatial and spectral) / temporal stability / cosmetics for QWIP. The<br />
situation is however worse for TIR HgCdTe as its metallurgy complexity is strongly increasing with cutoff<br />
wavelength. A 25 µm pixel pitch is considered as the smallest achievable pixel pitch. The targeted<br />
format of 0.8k x 0.8k pixels has a 28 mm diagonal, larger than the HgCdTe state of the art. The<br />
development of the GIFTS array made by BAe in the US has shown that, despite important<br />
technological and financial efforts, it looks hard to produce with acceptable operability a 20 mm<br />
diagonal HgCdTe 2D array with very long cut-off wavelength. On the other hand, 640x480 pixels<br />
QWIP arrays with 25 µm pitch are currently produced by few manufacturers. The HgCdTe problem is<br />
well known by <strong>ESA</strong> detection experts, particularly in the framework of MTG studies, justifying the two<br />
ways approach proposed by the Agency:<br />
• Improve the weaknesses of HgCdTe, via technological development. This is the object of the<br />
running contract "initial design of thermal infrared detector array for MTG"<br />
• Improve performances of alternative ways. This is the object of the contract "Enhanced<br />
QWIP/Sb-superlattice Array Detector".<br />
About ROIC, the 800x800 pixels format with 25 µm pitch avoids the need for stitching.<br />
An alternative to quantum detectors that could be figured for Fire Monitoring applications (which have<br />
much relaxed noise requirements) is the emerging microbolometer technology.<br />
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Microbolometers measure changes of electrical resistances with the help of pulsed currents. The<br />
major advantages are the possibility to operate at room temperature and the monolithic silicon<br />
structure which allows cheap production. Microbolometers are sensitive between 7 and 14 µm, but the<br />
responsivity is much lower than for quantum detectors. Microbolometers are not retained in the<br />
baseline for Fire Monitoring because the performances and the maturity level for GEO applications<br />
needs to be consolidated.<br />
The TIR detector architecture is similar to the MWIR one, but simpler thanks to the reduced number of<br />
pixels: two 400x800 pixels sub-arrays read out via 2 10 Mpixels/s video outputs, so the read out period<br />
is 16 ms, much larger than the max. integration time (
4 <strong>Final</strong><br />
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4.3.3.5 Mechanical architecture<br />
The mechanical configuration is driven by the Cassegrain telescope with 1.5 m diameter primary<br />
mirror (M1), as shown in Figure 4.3-11).<br />
2931 mm<br />
GEO-<br />
OCUL<br />
US-1<br />
2342 mm<br />
Figure 4.3-11: Overall PLM configuration<br />
646 mm<br />
2320 mm<br />
GEO-OCULUS-3<br />
The M1 mirror is mounted on the top side of the MIP (Main Interface Plate), whereas the bottom face<br />
carries all the focal planes and associated optics. The secondary mirror (M2) is supported by a spider<br />
attached to an hexapod structure which also carries the 2.5m long baffle. This configuration minimises<br />
the obscuration and provides a high dimensional stability between the M1 & M2, which drives the<br />
telescope optical quality.<br />
The instrument is interfaced with the SVM through an hexapod allowing high mechanical and thermal<br />
decoupling with respect to the platform.<br />
4.3.3.6 Thermal control<br />
The thermal control of the instrument is rather simple because Sun illumination of the interior of the<br />
telescope, and in particular the M1 mirror, is avoided by a Sun avoidance manoeuvre when the Sunto-LoS<br />
angle reaches 30 deg, i.e. +/-2h around midnight at equinoxes. This interrupts the imaging<br />
sequence (anyway limited to IR bands during night time). The Geo-Oculus telescope thermal<br />
architecture is therefore very classical and based on proven concepts & technology, with a<br />
combination of passive and active thermal control. The telescope is protected against Sun and cold<br />
space by the baffle and the focal plane & external structures are isolated thanks to MLI. Thermal<br />
washers are used to decouple the various assemblies and the whole PLM from the SVM. Active<br />
thermal control with heaters & thermistors is used to control telescope temperature, though radiative<br />
screens on the back of M1 & M2 mirrors and by direct conductive coupling for other elements.<br />
The temperature of the mirror will be very stable during daylight (6h to 18h), when the Sun aspect<br />
angle is larger than 90°, i.e. does not illuminate the inner baffle. This corresponds to the phase of UV-<br />
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VNIR imaging, where the best accuracy in pointing, defocus and WFE is required.<br />
During night time, the illuminated part of the baffle generates a disturbing flux on the mirror, the<br />
resulting thermo-elastic deformations which could generate defocus and wave front error (WFE) are<br />
minimised thanks to the high conductivity of SiC material. Moreover, the response of the mirror to this<br />
smoothly-varying flux is quick thanks to the combined effect of the high SiC conductivity and the low<br />
mass-to-area ratio of the mirror. Thermo-elastic distortions experienced during the night time are<br />
therefore not affecting high resolution daytime imaging performances.<br />
The required radiometric performance implies a temperature stabilised environment for each of the<br />
detectors, with the following operational temperatures: 50 K for TIR sensor, 130 K for MWIR and 20°C<br />
for UV & VNIR detectors. While the obvious solutions are passive cooling for UV & VNIR, and active<br />
cooling for TIR, the MWIR sensor could be in principle controlled with one or the other technique,<br />
provided that a sufficient radiating area with full view to cold space can be implemented. This is<br />
however not possible for the selected dual wing spacecraft configuration, since solar arrays are in<br />
view of possible radiating areas on the north & south walls.<br />
The three CMOS detectors and their proximity electronics are cooled by coupling with a small radiating<br />
area (0.06 m²) through conventional heat pipes. IR focal planes are housed in cryostats (single stage<br />
for MWIR, two stage with intermediate enclosure at 150 K for TIR) and cooled by mechanical<br />
cryocoolers. Coolers can be selected among several European products (see Figure 4.3-12), with two<br />
candidate technologies, Stirling-cycle coolers or pulse tube coolers. Astrium UK Stirling coolers are<br />
proven devices flown on numerous missions. Pulse tube coolers are completing space qualification<br />
and should be fully mature for Geo-Oculus. This technology is selected to minimise the number of<br />
units (Stirling coolers have to be operated in back-to-back pairs to avoid excessive vibrations) and<br />
therefore the mass and complexity. Three redunded coolers are necessary, two Miniature Pulse Tube<br />
(one for MWIR and the other for TIR outer enclosure) and a Large Pulse Tube for 50 K TIR enclosure.<br />
Manufacturer ASTRIUM-UK ASTRIUM-UK AIR LIQUIDE AIR LIQUIDE<br />
Model 50-80 K<br />
Miniature<br />
Pulse-Tube<br />
Large Pulse Tube<br />
Cooler (LPTC)<br />
Miniature Pulse Tube<br />
Cooler (MPTC)<br />
Type Stirling cooler Pulse Tube Pulse Tube Pulse Tube<br />
Performance<br />
1850 mW à 80 K<br />
>> 3 W à 130 K<br />
1400 mW à 80 K<br />
>> 2,5 W à 130 K<br />
6 W à 80 K 1300 mW à 80 K<br />
Mass 7,3 kg / cooler 6,4 kg / cooler < 8 kg / cooler < 6 kg / cooler<br />
Figure 4.3-12: Air Liquide Miniature Pulse Tube Cooler (left) and Astrium UK 50-80 K Stirling cooler<br />
(right)<br />
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4 <strong>Final</strong><br />
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4.3.3.7 Electrical architecture<br />
The electrical architecture is the essentially the same for all five focal planes:<br />
• The focal plane comprising the detector array (for and ) and the Proximity Electronics Module<br />
(PEM) housing all the functions requiring to be located close to the detector<br />
• The Remote Electronics Module (REM) hosting the other functions specific to each focal plane<br />
and providing the interface with the spacecraft data handling system. In order to minimise the<br />
power dissipation on the PLM, the REM units are implemented on the SVM.<br />
All detection chains are connected to a data bus interfacing with the spacecraft central processing unit<br />
and the data downloading function. This modular architecture with an independent detection chain for<br />
each focal plane allows flexibility in the PLM design and ensures robustness of the mission to a failure.<br />
The control electronics (for thermal control and activation of calibration devices and filter wheels) can<br />
be hosted in one of the REM as depicted in Figure 4.3-13 or in a dedicated electronics unit.<br />
Thermal control<br />
Calibration<br />
Telescope<br />
PAN optics UV-VNIR optics SWIR/MWIR optics TIR optics<br />
CMOS array<br />
PEM<br />
PAN FPA<br />
Video chain<br />
Memory<br />
Command/control &<br />
data processing<br />
Power supply<br />
PAN REM<br />
CMOS array<br />
PEM<br />
UV-Blue FPA<br />
Video chain<br />
Memory<br />
Command/control &<br />
data processing<br />
Power supply<br />
UV-blue REM<br />
CMOS array<br />
PEM<br />
Red-NIR FPA<br />
Video chain<br />
Memory<br />
Command/control &<br />
data processing<br />
Power supply<br />
Red-NIR REM<br />
TM/TC & data bus interface with SVM<br />
Figure 4.3-13: Electrical architecture of the instrument<br />
Hybrid array<br />
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PEM<br />
MWIR FPA<br />
Video chain<br />
Memory<br />
Command/control &<br />
data processing<br />
Power supply<br />
MWIR REM<br />
Hybrid array<br />
PEM<br />
TIR FPA<br />
Video chain<br />
Memory<br />
Command/control &<br />
data processing<br />
Power supply<br />
TIR REM<br />
Cryocooler Cryocooler
4 <strong>Final</strong><br />
<strong>Report</strong><br />
4.3.3.8 Calibration<br />
Challenging absolute (Goal: 1%, Threshold: 2%) and relative (0,2% inter-band) radiometric accuracy<br />
requirements impose careful calibration of absolute and inter-band offsets and gains. The calibration<br />
process shall enable to recover the scene reflectance values from instrument measurements. This<br />
implies a complete calibration of radiances/irradiances according to following calibration process:<br />
• Corrections of detectors systematic errors (Offset, Dark signal, Dark Signal Non Uniformity<br />
(DSNU), Pixel response non uniformity (PRNU), Dead/bad pixels)<br />
• Correction of absolute value and variation of the overall detection gain (optical transmission,<br />
detector efficiency, electronics gain).<br />
• Possibly, stray light correction based on ground characterisation<br />
To reach the above accuracy, in-orbit calibration is required, using a calibration device with properties<br />
well characterised on the ground and stable over the mission lifetime.<br />
For PAN & UV-VNIR channels, a retractable Sun diffuser will be used. The preferred candidate<br />
diffuser technology are QVD (Quasi Volumic Diffuser) and perforated plates for their low sensitivity to<br />
GEO environment. Since sighting the Sun with the instrument is not possible for thermal reasons and<br />
full pupil diffuser implementation at telescope entrance is not feasible because of large aperture, two<br />
complementary Sun diffusers are proposed:<br />
• A full pupil diffuser implemented near the intermediate focus, sighting the Sun through a<br />
window in the baffle. This diffuser does not monitor M1 & M2 possible degradation.<br />
• A small pupil diffuser implemented near the M2 spider to calibrate M1 & M2 transmission<br />
(local degradation of the mirrors outside the small pupil area covered by the diffuser is not<br />
monitored).<br />
In both cases, calibration is performed during the 4h interruption of measurements around midnight,<br />
thus do not interfere with the mission imaging capability. The Sun avoidance manoeuvre performed to<br />
keep the Sun-to-LoS angle larger than 30 deg is used to sequentially orient each Sun diffuser towards<br />
the Sun. The Sun diffuser LoS (defined by the window in the Sun shield) the has an offset from the<br />
instrument LoS equal to the magnitude of the Sun avoidance manoeuvre, and the spacecraft is rotated<br />
about the instrument LoS, as illustrated in the following figure.<br />
GEO Sun diffuser<br />
θman − βsun<br />
North<br />
Figure 4.3-14: Geometry of Sun diffuser sighting during Sun avoidance manoeuvre<br />
For IR channels, calibration can be made against two stable radiometric references, a small black<br />
body mounted on a flip-flop mechanism or periodic sighting to the cold space<br />
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βsun<br />
Equator
4 <strong>Final</strong><br />
<strong>Report</strong><br />
4.3.4 PLM budgets<br />
The PLM budgets are computed for the baseline configuration with the following assumptions:<br />
• A 1.5 m diameter instrument with 5 focal planes.<br />
• Dual wing solar arrays, which is less favourable for thermal control efficiency because<br />
radiators on the NS walls have a reduced viewing factor.<br />
• Conventional 2.5 m baffle and Sun avoidance manoeuvres around midnight to keep the LoSto-Sun<br />
angle larger than 30 deg, i.e. preventing that Sun enters the telescope.<br />
• Remote Electronics Modules used for digital processing of the images are implemented in the<br />
SVM and accounted for in the SVM budgets.<br />
The PLM budgets are provided in the following figure:<br />
2930 mm<br />
2340 mm<br />
Mass Power<br />
Best estimate 505 kg 423 W<br />
Margins: 20% 101 kg 85 W<br />
TOTAL with margins 606 kg 508 W<br />
Figure 4.3-15: Geo-Oculus instrument interface budgets<br />
650 mm<br />
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4.4 Line of Sight (LoS) Stabilisation Concepts<br />
4.4.1 LoS stabilisation main issues: microvibrations and post-integration<br />
The LoS pointing stability requirements are very stringent for the aspects of the mission involving high<br />
resolution imaging. This is particularly the case during PAN imaging for disaster monitoring, since the<br />
20 m resolution over Europe corresponds to 0.28 µrad nadir resolution, and to a lower extent when<br />
acquiring full resolution VNIR images (40 m resolution over Europe, 0.56 µrad nadir resolution) for<br />
disaster or fire monitoring. Image quality is then highly sensitive to LoS motion over the short<br />
integration time (up to ~100 msec) required to meet the moderate SNR requirements.<br />
On the contrary, the Marine applications, with a resolution relaxed to 80 m over Europe (i.e. 1.1 µrad<br />
nadir), are less sensitive to pointing stability over the image integration time. Nevertheless, since<br />
several images need to be post-integrated to reach high SNR requirements, image quality is more<br />
sensitive to pointing drifts over the total image acquisition time (several sec).<br />
LoS Stabilisation requirements derived from instrument design are summarised in the following table:<br />
RPE: Relative Pointing Error<br />
(stability over the integration time)<br />
RME: Relative Measurement Error<br />
(over image acquisition time)<br />
PDE: Pointing Drift Error<br />
(drift over integration time)<br />
0.15-0.2 µrad peak-to-peak for high frequency jitter (>10 Hz)<br />
0.1 µrad over 5 s max. acquisition time<br />
Marine applications: 10 µrad/s<br />
Fire/Disaster monitoring: 5 µrad/s<br />
Such specifications can not be met without proper management of the LoS pointing stability issue.<br />
Depending on the type of disturbances that challenge the LoS stability requirements, and in particular<br />
depending on the frequency band affected, different solutions might be proposed for Geo-Oculus :<br />
• High frequency perturbations, with period lower than typical integration time (100 msec), i.e.<br />
frequency > 10 Hz, require disturbance reduction techniques. Such disturbances are mainly due to<br />
microvibrations generated by moving parts (e.g. reaction wheels and cryocoolers).<br />
• Medium frequency disturbances (with period in the range of a few sec, corresponding to the time<br />
to acquire an image based on accumulation of several shots) require image processing<br />
techniques to enable post-integration. Such disturbances are mainly due to solar array flexible<br />
mode excitation after slew manoeuvres.<br />
• Low frequency disturbances are handled by the AOCS for those observed by the attitude sensors<br />
and by ground-based processing (so-called INR, Image Navigation & Registration) for LoS<br />
pointing errors due to orbit errors and thermo-elastic distortions between LoS and AOCS<br />
reference.<br />
Therefore, the LoS stabilisation issues to be addressed are:<br />
• Microvibrations: level reduction through careful selection of actuators and identification of<br />
potential other disturbances.<br />
• Post-integration: Evaluation of the technique to be used to estimate the shift between one<br />
image and another one before adding them.<br />
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4.4.2 Microvibrations<br />
4.4.2.1 Candidate mitigation actions<br />
The following strategies are defined to limit the impact of microvibration on LoS stability:<br />
• Stopping cryocoolers during PAN imaging. This solution is successfully used in orbit for high<br />
accuracy LEO Earth observation. Since switching on and off would not be acceptable in terms of<br />
number of electronic cycling, the cryocoolers are in fact kept on, but the amplitude of the engine is<br />
simply turned to 0 during imaging and then turned back to full power, without any ageing effect.<br />
The drawback of this technique is that thermal control of cold detector is effectively turned-off, and<br />
its temperature raises by less than 1 K/s. In the case of GEO-Oculus, since the PAN image<br />
acquisition is very short (0.4 s), the temperature raise shall be much less than 1K. These very<br />
small temperature cycles are deemed to be acceptable for the detector.<br />
• Elastomeric suspension to isolate the spacecraft from microvibrations generated by cryocoolers<br />
or conventional ball-bearing reaction wheels (BBW). Elastomeric mounts sustaining launch efforts<br />
without clamping have been developed and qualified for reaction wheel isolation and will be flight<br />
proven with Pleiades in 2009. With suspension frequency around 15 Hz, elastomeric mounts allow<br />
efficient attenuation of disturbances above ~50 Hz, where major BBW harmonic disturbances are<br />
reported. They are also efficient for high order harmonics which dominate cryocoolers<br />
disturbances when the main disturbance at cooler rate (40 to 50 Hz) is cancelled by design (backto-back<br />
Stirling coolers or pulse tube technology). This is therefore the most mature solution to<br />
drastically reduce the high-frequency components of BBW & cryocoolers disturbances.<br />
• Magnetic Bearing reaction Wheels (MBW) is a reaction wheel where no mechanical contact<br />
between moving parts is established during normal operation. This is achieved by magnetic<br />
levitation and position control of the rotor. The direction of the rotation axis can be actively<br />
controlled within certain limits by adjustment of the magnetic fields. This feature allows creating<br />
relatively high torques perpendicular to the wheel rotation axis. Hence, a MBW can be used for<br />
limited agile slewing manoeuvres. MBW are known to generate much less perturbations than their<br />
ball-bearing equivalent. With the availability of such equipments, the resulting high-frequency<br />
micro-vibration at instrument level should be reduced.<br />
In the following sections, the two reaction wheel options (MBW and BBW + elastomeric isolator) are<br />
compared in terms of microvibration disturbance levels and technology maturity. Cryocooler<br />
microvibrations are assumed to be mastered by the combination of elastomeric suspension and<br />
cryocooler stop during the most jitter-sensitive phase, PAN imaging.<br />
4.4.2.2 Magnetic bearing wheel (MBW) evaluation for Geo-Oculus<br />
The magnetic bearing wheels envisaged for Geo-Oculus corresponds to the new design of Rockwell<br />
Collin’s Teldix (RCT) wheels. The currently existing MBW has the status of a technology<br />
demonstration, with drive and control electronics located outside of the wheel, only sensor electronics<br />
placed inside. In the flight design, the complete electronics equipment will be put inside the wheel.<br />
Table 4-3 compares the technical data of Teldix's 15 Nms BBW to the data of the prototype MBW and<br />
the foreseen data of two future flight MBW: MWI 30-400/37 is the basic model adequate for Geo-<br />
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Oculus, whereas MWI 100-100/100 has a bigger rotor and different motor design.<br />
Prototype MBW<br />
Figure 4.4-1: MBW characteristics and photography of the prototype MBW (courtesy of RCT)<br />
The microvibration levels generated by the prototype MBW wheels have been characterised in 2007<br />
by EADS Astrium GmbH in the frame of the DLR study “High Precision Attitude Control of Earth<br />
Observation Satellites”. The results of this study show that the MBW disturbance levels are lower by a<br />
factor of 10 to 240 (depending on frequency and wheel rotation rate) than typical BBW levels.<br />
However, it shall be noted that the comparison is supposing hard-mounted wheels, whereas a BBW<br />
mounted on an elastomeric suspension would be the actual competitor for a high accuracy pointing<br />
mission. Moreover, microvibrations is analysed at the source, whereas its impact on LoS at PLM level,<br />
largely dependent on structure transmission is the relevant parameter. It is undoubted that the<br />
microvibrations will be lower, but the above ratios shall not be taken as granted.<br />
The MBW appear in all cases as a good candidate for the Geo-Oculus mission, due to its inherent low<br />
microvibration content. The low maturity level (TRL ~4) and the associated development and technical<br />
risks shall also be accounted for. The pre-development needs to be actively pursued to reach TRL 5 at<br />
the beginning of phase C/D.<br />
4.4.2.3 Ball Bearing Wheel (BBW) option<br />
A second option is to use standard ball bearing wheels mounted on elastomeric mounts developed for<br />
LEO observation missions. This option has been analysed in 2007 in the frame of the CNES study<br />
“Constraints for High resolution observation on GEO”. A microvibration analysis was performed for<br />
three different structural transmissions, without elastomeric suspension, with a 15 Hz suspension, and<br />
with a 30 Hz one. Microvibration levels measured on 8 flight models of Pleiades BBW (18 Nms Teldix<br />
RSI) are used as input to the structural model, typical of a GEO spacecraft equipped with an optical<br />
payload for Earth observation. The results are post-processed so that to show the peak-to-peak<br />
variation of the LoS over an integration time of 70 ms. The results of this study are therefore directly<br />
relevant for Geo-Oculus.<br />
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Figure 4.4-2: Structure FEM used for the BBW microvibration analysis and computed stability over 70<br />
msec for a typical 2-day wheel rate profile<br />
The following conclusions are drawn from this analysis. First, elastomeric suspension is mandatory<br />
since it allows a reduction by a factor 100 of the high-rank harmonics perturbations. Second, the<br />
elastomeric suspension frequency shall be set to 15 Hz and the wheels velocity shall be maintained<br />
below 45 Hz (2700 rpm) thanks to adequate wheel off-loading process at AOCS level. Then,<br />
conservatively considering a linear summation of the harmonics and the worst wheel FM, the worst<br />
case performance over a typical wheel velocity profile is 0.24 µrad peak-to-peak over 70 ms. This is<br />
above the Geo-Oculus requirement (0.15-0.2 µrad peak-to-peak), but the impact on the achievable<br />
resolution would be limited, with a degradation of the nadir ground resolution from 10.5 m to 11.5 m.<br />
Of course, the level of this preliminary analysis cannot give commitment on these figures, but the order<br />
of magnitude is believed to be valid.<br />
Therefore, the use of BBW for Geo-Oculus shall not been ruled out by microvibration aspects.<br />
Furthermore, the technology is fully qualified, flying on previous programmes.<br />
4.4.3 Post-integration<br />
Stabilité sur 70 ms (µrad)<br />
0.25<br />
0.2<br />
0.15<br />
0.1<br />
0.05<br />
Performance en fonction du temps<br />
FM09<br />
FM06<br />
Pire cas<br />
0<br />
0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2<br />
Temps (jour)<br />
4.4.3.1 Principles<br />
Post-integration consists in taking several successive images of the scene with short integration time<br />
and to add them together to obtain long-integration images, i.e. with high SNR. Keeping the integration<br />
time small is necessary for avoiding pixel saturation and for relaxing LoS stability when drift is the<br />
dominant error. Post-integration is of no use for micro-vibrations mitigation and applies only to drift<br />
mitigation, by reducing the integration time. One option is to perform this post-integration on ground,<br />
but this dramatically increases the downlink data rate since up to several tens of images are required<br />
on the most critical channels (e.g. for marine applications with high SNR requirements and low<br />
reflectance). On-board post-integration is therefore the baseline for Geo-Oculus, based on the of the<br />
experience gained on COMS satellite developed by Astrium for Korea.<br />
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If the LoS motion during the summation of the successive images is large (typically > 1 pixel), the<br />
motion must be compensated for by shifting the pixels. The simplest correction is the so-called<br />
"nearest pixel motion compensation", where the shift is limited to an integer number of pixels, simply<br />
achieved by a shift in memory and an accumulation. The average MTF loss at the Nyquist frequency is<br />
0.64 on the accumulated image, i.e. equivalent to a shift of one pixel over the whole accumulation. To<br />
reduce this significant degradation of the MTF, refined offset correction methods based on pixel<br />
interpolation are possible, but not retained for Geo-Oculus for their low maturity and the required large<br />
on-board computation and storage capabilities.<br />
The first step is however to measure the LoS motion between two integration phases, with an<br />
accuracy significantly better than half a pixel (0.2 pixel i.e. ~0.1 µrad).<br />
4.4.3.2 LoS drift measurement<br />
The LoS motion information can be extracted from gyroscope measurements, provided they are<br />
mounted close to the focal plane. The following figure shows the LoS drift estimation error for two high<br />
accuracy gyros (Pleiades Astrix 200 FOG and SIRU HRG): Over the maximum image acquisition time<br />
(5 sec), the error is 0.3-0.5 µrad, well above the 0.1 µrad requirement. Gyros are therefore not<br />
adequate for LoS drift estimation.<br />
Gyro drift (µrad)<br />
0,5<br />
0,4<br />
0,3<br />
0,2<br />
0,1<br />
0,0<br />
0 1 2 3 4 5<br />
Time (secs)<br />
Figure 4.4-3: LoS drift estimation accuracy using gyroscopes<br />
ASTRIX200<br />
In the case of GEO-observation with a staring instrument, the motion information can also be<br />
extracted from the image itself, which removes the need for additional motion sensor. The principle is<br />
to correlate in real-time on board the spacecraft the incoming image with the accumulated image, so<br />
as to determine the relative image to be corrected. Either the full image or vignettes of interest are<br />
used. In the first case, the processing load is high, but the algorithm is simple (correlation over a small<br />
moving window) and repetitive, which is well adapted to FPGA or ASIC implementation. In the latter<br />
case, the system shall identify vignettes of interest within the first image using an algorithm detecting<br />
areas with contrasted variations. Then the correlation is performed between the selected vignettes<br />
extracted from the accumulated & current image.<br />
Such techniques are actively investigated at Astrium, primarily for on-ground processing to improve<br />
image quality without relying on pre-defined landmarks. The resulting accuracy of image correlation is<br />
in the order of 10% to 20% of a pixel, that is to say below 0.03 to 0.06 µrad, well with in the 0.1 µrad<br />
requirement. Image correlation is therefore the selected approach for LoS drift measurement.<br />
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4 <strong>Final</strong><br />
4.5 Satellite<br />
<strong>Report</strong><br />
4.5.1 Configuration<br />
The basic features of the suggested baseline configuration as derived by the trade-offs documented in<br />
[RD 7] are as follows:<br />
• dual wing steerable solar array,<br />
• payload on top panel (launch) / Nadir panel (operation),<br />
• no yaw flip manoeuvre.<br />
The selected spacecraft configuration based on above inputs has been influenced by the choice of the<br />
favourable design of the chemical propulsion system as well as by the heritage available from similar<br />
projects. The resulting external configuration of the spacecraft in stowed condition is shown in Figure<br />
4.5-1.<br />
The payload is connected to the platform via 3 bi-pods providing iso-static mounting conditions. The<br />
configuration of these bi-pods has still to be iterated and the provision of suitable hard-points in the<br />
platform, accordingly.<br />
The solar array wings are stowed on the side panels which correspond to the North and South panel<br />
during operational mode.<br />
The PDT antenna system comprises a deployable boom and is also folded to a side panel. The<br />
deployable boom is needed in order to provide visibility between antenna and ground station which is<br />
challenged by the big payload and, moreover, by the manoeuvres which point the complete spacecraft<br />
to the scene of interest.<br />
A similar problem concerns the S-Bd antenna which is roughly Nadir oriented (the complementary<br />
Zenith oriented antenna is hidden behind the spacecraft). Since this S-Bd antenna needs a<br />
hemispherical field of view, a mounting on the payload close to the entrance of the big baffle has been<br />
selected.<br />
This accommodation of platform equipment on the payload does not seem to be necessary for the<br />
infrared Earth sensors (IRES). Their field of view requirement is about 20 degree half cone angle and<br />
may be provided by putting these sensors on a pedestal.<br />
The situation is even a little more comfortable for the star trackers. They shall be aligned as close as<br />
possible with the payload line of sight direction in order to achieve the maximum attitude knowledge<br />
accuracy. However, they must also consider a certain Earth exclusion angle which yields an off-Nadir<br />
viewing direction. This together with the even narrower field of view (~15 degree half cone angle)<br />
allows to accommodate the star trackers directly on the top panel. This position may be revised in a<br />
later phase since the neighbouring payload radiator may evolve and the contribution of thermo-elastic<br />
deformations to the pointing knowledge budget may be optimised by mounting the star trackers<br />
directly on the payload.<br />
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S-Bd<br />
Antenna<br />
Instrument<br />
X 3500 mm<br />
Y 2300 mm<br />
Z 2300 mm<br />
S/C Body<br />
X 2450 mm<br />
Y 2600 mm<br />
Z 2200 mm<br />
PDT<br />
Antenna<br />
Space<br />
Environment<br />
Sensor<br />
Figure 4.5-1: Stowed Configuration<br />
S-Bd<br />
Antenna<br />
Star<br />
Tracker<br />
IRES<br />
Sensors<br />
Figure 4.5-2 depicts the deployed configuration of the suggested baseline concept for Geo-Oculus.<br />
The upper solar array wing is oriented towards the North direction and the lower wing towards the<br />
South direction. The wings are steerable around the North – South axis thereby always providing an<br />
optimised sun inclination angle.<br />
The PDT antenna is also deployed to achieve a sufficient clearance between the antenna field of view<br />
and the payload. A 2-axis antenna pointing mechanism is installed directly under the antenna dish<br />
which is foreseen to compensate the manoeuvres for acquiring a new scene.<br />
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4 <strong>Final</strong><br />
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Star<br />
Tracker<br />
Figure 4.5-2: Deployed Configuration<br />
Solar Array<br />
IRES<br />
Sensors<br />
PDT<br />
Antenna<br />
<strong>Final</strong>ly, in Figure 4.5-3 the side panels of the spacecraft body are folded away such that the internal<br />
arrangement of equipment becomes visible.<br />
All equipment is spread over the North panel (on the left) and the South panel (on the right). These are<br />
the preferred locations since the conditions for heat rejection are optimum there which is beneficial for<br />
the sizing of the thermal control system. Since the available mounting surface of these panels is<br />
comfortable also in view of a later optimisation of the spacecraft balancing, no equipment needs to be<br />
mounted on the East and West panels. Hence, the East and West panels are designed as light weight<br />
closure panels.<br />
As a conclusion it can be stated that for the overall spacecraft configuration no major criticality has<br />
been identified. All design concepts applied are based on heritage thus providing a solution with low<br />
risk.<br />
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Instrument<br />
Prop. Tanks<br />
PDHT<br />
Electronics<br />
Instrument<br />
Electronics<br />
Star<br />
Tracker<br />
IMU<br />
IMU<br />
Electronics<br />
SADM<br />
<strong>Report</strong><br />
Figure 4.5-3: Internal Configuration<br />
Battery<br />
PSR<br />
SPU<br />
He-Tank<br />
(in central tube)<br />
Reaction<br />
Wheels<br />
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SCU<br />
ADE5<br />
S-Bd<br />
Transponder<br />
PLIU<br />
SADM<br />
Coarse IMU<br />
4.5.2 Electrical Architecture<br />
The electrical architecture satisfies the need for high reliability and availability. A low risk approach is<br />
followed which means that heritage from previous projects is used whenever possible. Geostationary<br />
heritage can be derived from the Eurostar platform series and specifically from the COMS satellite<br />
which already implements meteorological and communication services on a 3-axis stabilized<br />
geostationary satellite, similar to the GEO-Oculus mission.<br />
Minimum complexity is achieved by separation of the electrical architecture into functional modules<br />
consisting of elements independent of the mission needs and customized elements which are tailored<br />
for mission specifics especially in the payload section. As such, the architecture references already<br />
the hardware breakdown as used on Eurostar for the bus elements while still giving flexibility for<br />
mission specific adaptations on platform side.<br />
High autonomy and reliability is satisfied by the redundancy concept within the overall electrical<br />
architecture and the on-board computer. It provides redundant modules and bus systems thus<br />
minimizing the amount of single point failures. Autonomy is also a key requirement on<br />
telecommunication satellites and therefore inherently available.<br />
Growth potential is achieved by a scalable electrical architecture. This is given by a scalable solar<br />
array in terms of amount of panels, regulator stages, battery size and regulator capability as well as<br />
adaptability of the amount of command and data handling interfaces from and to the on-board<br />
computer.<br />
For the payload side, a mission specific architecture is necessary due to the data rates of the<br />
instrument. Therefore, MIL-1553 bus has been selected for instrument TM/TC between SCU
4 <strong>Final</strong><br />
<strong>Report</strong><br />
(Spacecraft Computer Unit) and instrument and high-speed interface for the instrument data to be<br />
transferred to the PDH. Cross-coupling is achieved by redundant Milbuses and cross-coupled<br />
interface to the PDH. Since G-Link is now obsolete, newer technology such as Aeroflex UT54 series<br />
may be more appropriate.<br />
The proposed electrical architecture is given in the following figure. The payload subsystem contains<br />
an Instrument Control Unit (ICU) for self-standing instrument mode control and to facilitate instrument<br />
testability.<br />
Figure 4.5-4 Geo-Oculus Electrical Architecture (Overview)<br />
4.5.3 Power Subsystem<br />
The Electrical Power System (EPS) shall serve the satellite in sun and eclipse with the required power<br />
as derived from the power budget. The following essential EPS sizing requirements apply:<br />
• Life time: 10 years<br />
• Orbit: Geostationary<br />
• Maximum eclipse duration: 72min<br />
• Unit margin: between 5% to 30% pending maturity<br />
• System margin: 10% at EOL plus 10% time margin on energy recharge<br />
The main functions of the electrical power systems are<br />
• generation of electrical power with a photo-voltaic solar array<br />
• storage of excess energy during sun phases into the battery<br />
• safe distribution of the electrical power to the on-board users<br />
• protection of the battery against overcharging and deep discharging<br />
• fully autonomous operation.<br />
The EPS consists of solar array, battery, Power Shunt Regulator (PSR), battery charge and discharge<br />
regulators as well as bus distribution and protection. Two one axis drive mechanism provide optimum<br />
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sun orientation for each of the two wings of the solar array.<br />
Power Budget<br />
Solar array and battery sizing assumptions:<br />
• Solar array temperature: 57°C<br />
• Summer solstice EOL<br />
• Solar array degradation for required life time<br />
• Unit power figures include maturity margins and an overall system margin of 10%.<br />
Power sizing<br />
The instrument power value is based on a mean figure without sun avoidance and a dual wing solar<br />
array configuration. For the other instrument/satellite alternatives (dual wing solar array with sun<br />
avoidance or single wing with/without sun avoidance) the solar array and the battery will be slightly<br />
smaller.<br />
This gives the following results:<br />
• Average load power: 1800W<br />
• Required solar array area: Minimum 10.1 m 2<br />
• Battery size: 135Ah (11 string 3 parallel configuration)<br />
• Battery mass: 48kg<br />
W<br />
2500<br />
2000<br />
1500<br />
1000<br />
500<br />
Figure 4.5-5: Power Profile<br />
0<br />
Power Profile GEO-Oculus<br />
396 796 1196<br />
min<br />
100,00<br />
90,00<br />
80,00<br />
70,00<br />
60,00<br />
50,00<br />
40,00<br />
30,00<br />
20,00<br />
10,00<br />
0,00<br />
Power SA W Power profile Load W Battery SoC<br />
The solar array area includes 3.5% margin for string failures.<br />
The battery is based on G5 technology. Cell losses are covered by two additional modules (parallel<br />
cells). The maximum battery DoD is approx. 62%. The battery sizing is based on nominal in-orbit<br />
operations.<br />
4.5.4 Payload Data Handling and Transmission<br />
Data rate assessment<br />
For the sizing of the PDHT, an average data rate of 250Mbps coming from the instrument is assumed.<br />
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This is based on a worst case data rate scenario with the following assumptions:<br />
• Post-integration with (simple) image motion compensation is done on-board (necessary for the<br />
UV-blue channels of the marine applications with up to 30 images to be summed up).<br />
• Due to the required image summation, 18 bits per pixel are assumed to provide an adequate<br />
quantisation. For the sake of a simple and reliable algorithm, the 18 bits per pixel are assumed<br />
for all channels even if no image summation is required.<br />
• For the panchromatic channel VNIR7 required for the disaster monitoring mission, it is<br />
assumed that 4 images are down-linked which may then be subjected to motion compensation<br />
and post-integration with a more accurate and sophisticated algorithm.<br />
• 2 x 2 pixel binning of the UV-blue and red-NIR channels of the marine applications has not<br />
been considered for the data rate.<br />
With these partly conservative assumptions, an evolution of the data rate throughout the next study<br />
phases may be compensated such that no change of the selected data transmission technology<br />
becomes necessary. For details about the choice of the location and technology of post-integration<br />
and image summation, see Ref. [RD 8].<br />
The instrument data are routed via cross-coupled high rate serial interface to the PDH. In the PDH the<br />
data are buffered and formed to a continuous data stream with formatting (CADU generation), RS<br />
encoding and scrambling. The buffer size has to be determined in the coming study phase since it<br />
strongly depends on the ratio of average to peak instrument data. The PDH has a fully redundant<br />
structure with the input modules, the buffer and the TMFE output modules interfacing with the cold<br />
redundant transmission chains. The TMFE outputs provide full cross-coupling to the modulators.<br />
The PDT is based on cold redundant transmit chains with each consisting of modulator and SSPA,<br />
followed by a non-redundant chain selection switch and an output filter. In order to reduce power<br />
consumption, a high gain satellite transmit antenna of 0.8m has been selected together with a ground<br />
station receive antenna of 13m diameter. The satellite transmit antenna has a beamwidth of approx. 3<br />
degree (3dB double sided beamwidth). The peak gain of the antenna is considered in the link budget<br />
requiring antenna pointing via a 2 axes pointing mechanism in case of satellite re-orientation.<br />
Payload data handling, modulator, amplifier and antenna are based on existing components or require<br />
minor modifications to be suitable for Geo-Oculus. The only exception is the antenna pointing<br />
mechanism due to the large amount of operational cycles. It is expected that upgrading of existing<br />
designs or delta qualification will be sufficient.<br />
Carrier frequency selection (ITU constraints)<br />
The choice of carrier frequency is dependent on a number of technical and regulatory constraints,<br />
including the ease of frequency coordination and location of ground station. For the less than 300 MHz<br />
required bandwidth proposed, use of X-Band has been chosen as the baseline (maximum bandwidth<br />
available in the 8 GHz downlink band is 375 MHz). Use of the currently unused EES spectrum at 26<br />
GHz (Ka Band) could also be feasible, if a ground station is capable of overcoming propagation<br />
issues.<br />
Antenna design baseline<br />
Due to the high downlink data rate (250 Mbit/s), a High Gain Antenna has been selected as a solution<br />
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for the PDHT system. The need for a body-mounted instrument affects the contact to the ground<br />
station. Due to the necessity to transmit while the satellite is repointing, a steering mechanism for the<br />
HGA will be required. This will probably involve a boom mounting scheme on the spacecraft. The<br />
impact on the spacecraft in terms of pointing disturbance has not yet been assessed. Control of the<br />
HGA motion will be within the PLIU.<br />
Other options such as an electrically steerable (phased array) antenna or a Low Gain Antenna (LGA)<br />
do not provide sufficient gain for the high data rate and have now been removed.<br />
Modulation and coding scheme<br />
As OQPSK has been selected as modulation scheme, standard RS encoding 255/223 is proposed.<br />
This results with an instrument output rate of 250 Mbps in an occupied bandwidth of < 300 MHz in Xband.<br />
The Reed-Solomon code to be used is in agreement with ECSS standard ECSS-E-50-01A<br />
(Telemetry Sync and coding), Chapter 6. A pseudorandomiser could be used to provide sufficient<br />
channel symbol transitions and hence improve received symbol lock.<br />
Link Budget<br />
Considering a bit error rate of 10 -9 and a transmit power of 8W, a satisfactory link margin of 3.9dB is<br />
achieved with a ground station of 13m diameter.<br />
Ground station interface<br />
Selection of ground station locations would be dependent on the geostationary longitude of the<br />
spacecraft, as described in the System requirements document. A location in central Europe such as<br />
Italy or Germany would be capable of viewing a spacecraft regardless of whether it is situated at a<br />
longitude more than 45 degrees East or West. This would provide some flexibility in choosing a<br />
suitable position for the spacecraft. In all cases the spacecraft should be continually more than 10<br />
degrees above the horizon as seen from the ground station in order to provide sufficient link margin.<br />
Additional considerations to the ground station would include the interface to the user segment,<br />
including archiving availability, offsite data links, processing centre capability and user access and<br />
security. These factors are especially important where mobile terminals could be deployed as part of a<br />
network.<br />
The block diagram of the PDT is shown in Figure 4.5-6.<br />
Nominal<br />
Chain<br />
Data<br />
Clock<br />
Redundant<br />
Chain Data<br />
Clock<br />
OQPSK-<br />
Modulator<br />
OQPSK-<br />
Modulator<br />
Figure 4.5-6: PDT Block Diagram<br />
SSPA<br />
X-Bd<br />
SSPA<br />
X-Bd<br />
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Reception of the instrument data by mobile terminals is also possible, but affects the necessary onboard<br />
transmit power. For example, assuming a 3m diameter mobile reception antenna, the on-board<br />
RF power will increase to approx. 50W requiring the use of TWTAs for the amplification instead of<br />
SSPAs.<br />
4.5.5 Telemetry and Telecommand<br />
General<br />
The S-Band Subsystem provides all classical Tracking, Telemetry and Command (TT&C) services.<br />
Ranging and range rate services are implemented according to the <strong>ESA</strong> ranging standard. Power Flux<br />
density requirements are respected during all normal operational phases (except launch).<br />
TM<br />
TC - Data<br />
& Clock<br />
TC - Data<br />
& Clock<br />
TM<br />
Transmitter 1<br />
(nominal)<br />
Ranging<br />
&<br />
Coherency<br />
Receiver 1<br />
(nominal)<br />
Transponder 1<br />
Transponder 2<br />
Receiver 2<br />
(hot redundant)<br />
Ranging<br />
&<br />
Coherency<br />
Transmitter 2<br />
(cold<br />
redundant)<br />
Figure 4.5-7: S-Band Subsystem<br />
Diplexer<br />
Diplexer<br />
3 dB -<br />
Combiner<br />
Nadir<br />
Antenna<br />
Zenith<br />
Antenna<br />
The S-band communications subsystem consists of two transmitters, two receivers, a 3dB Combiner,<br />
two antennas, and RF harnessing. The nominal RF transfer from and to ground will be achieved using<br />
a combined receive/transmit quadrifilar helix (QFH) antenna mounted on the nadir side. An identical<br />
QFH antenna on the zenith side is used to establish ground contact for off-nominal attitude conditions.<br />
While both receivers are running in permanent hot redundancy, one of the two transmitters will be<br />
switched on via the Spacecraft Computer Unit (SCU) when required to perform ranging or to downlink<br />
the housekeeping telemetry data to the ground station and will, usually, also be switched on prior to<br />
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launch, remaining on until completion of LEOP.<br />
RF signals from the ground stations at Usingen in Germany, or Maspalomas in Gran Canaria, will be<br />
received from both receive antennas, superposed in the combiner and routed to both receivers. The<br />
first receiver achieving a subcarrier lock will be selected by the SCU telecommand decoder as the<br />
‘active’ receiver.<br />
The nominal transmitter will send the generated RF signal to both nadir and zenith antennas for<br />
transmission to ground. In the case of a failure, the nominal transmitter will be deactivated and the<br />
redundant transmitter will take over operation.<br />
Telemetry, Tracking and Command<br />
The uplink frequency will be within the 2025 – 2110 MHz band whereas the downlink frequency will be<br />
within the 2200 – 2290 MHz band. Coherence will be implemented, applying the turnaround ratio of<br />
221/240, and can be enabled or disabled using appropriate commands.<br />
During TT&C operation, the uplink Telecommand data and the downlink Housekeeping Telemetry<br />
data are modulated on to subcarriers with typical bitrates of 2000 bps and 8192 bps, respectively. The<br />
subcarriers are phase modulated onto the respective uplink and downlink carriers, together with the<br />
ranging signal.<br />
The helix antennas provide hemispherical coverage with a worst case antenna gain of about -3 dBi.<br />
The Link Budget has been developed to achieve required minimum link margins, e.g. >3dB nominal<br />
TM recovery margin, while fully respecting the maximum Power Flux Density (PFD) requirements.<br />
Link Performances<br />
A sample link budget for S-Band TT&C and Ranging link is shown in Table 4.5-1 below. Basic 223/255<br />
Reed-Solomon encoding has been assumed for the telemetry downlink. The chosen ground station is<br />
Maspalomas on Gran Canaria, which has a 15m diameter dish and a reception G/T of 29.2 dB/K.<br />
Table 4.5-1: Assumptions for S-Band TTC Links<br />
Programme: Geo-Oculus Orbit: GEO<br />
Ground Station: Maspalomas-1 (S-Band)<br />
Height: 35800.00 km<br />
S/C Antenna: AEOLUS LGA (S-Band)<br />
Elevation: 10 degrees<br />
Type: Ranging<br />
Amplifier RF Output: 5.0 W Ranging Possible: TRUE<br />
Downlink Data Rate: 8,192 s/s Uplink Data Rate: 2,000 s/s<br />
Information Rate: 7,142 bps<br />
Coding Scheme: 223/255 Reed-Solomon Coding<br />
Modulation Scheme: PCM(NRZ-L)/PSK/PM<br />
Subcarrier Type: Sine Wave<br />
The analysis yields very healthy recovery margins for Telecommand (26.45 dB nom.) and<br />
Telemetry (18.65 dB nom.). Power Flux density limits are not violated.<br />
It should be noted that two S-band stations are required for the envisaged spread spectrum ranging<br />
method - for an optimal performance, an additional third station should be implemented.<br />
Consequently, it makes sense to consider not only Maspalomas but also a second station, such as an<br />
S-Band station in Redu, for instance. Table 4.5-2 shows the differences in EIRP and G/T for these two<br />
stations. Redu will have a marginally better link than Maspalomas.<br />
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Table 4.5-2: Characteristics of Maspalomas and Redu S-Band stations<br />
MAS-1 RED-1<br />
EIRP 72.1dBW 72.5dBW<br />
G/T 29.2dB/K 29.6dB/K<br />
4.5.6 Attitude and Orbit Control<br />
4.5.6.1 Introduction<br />
The AOCS plays a significant role within the functional process chain to fulfil the demanding pointing<br />
requirements of a very high performance mission, both in terms of absolute and relative pointing /<br />
pointing knowledge. This overall process chain consists of the payload, the platform (including the<br />
AOCS) and the on-ground post-processing (INR).<br />
Based on system pointing requirements and an associated pointing budget for the whole process<br />
chain, preliminary requirements for the AOCS have been derived. Out of the various pointing<br />
requirements, the following ones are driving the AOCS concept and will be checked in this chapter:<br />
• the absolute pointing error (APE),<br />
• the absolute measurement error (AME),<br />
• the pointing drift error (PDE) over 100 msec.<br />
A feature special to the Geo-Oculus mission is that the whole spacecraft is turned in order to move to<br />
the next image which shall be acquired in a step and stare mode. Based on the findings related to the<br />
mission scenario trade-offs, a medium agility is requested from the AOCS concept in order to support<br />
a reasonable number of mission products within the dedicated revisit cycles. The allocated budget<br />
assigned to this medium agility is 70 sec for the total manoeuvre time between 2 images. The agility<br />
itself is driven by the AOCS actuator selection and the overall platform design (moments of inertia,<br />
flexible modes). Especially the flexible modes come into play when high torque actuators are used.<br />
This is due to a high initial deflection and the related long tranquilisation time in order to reach again<br />
the required pointing budgets. Sun avoidance manoeuvres in regular intervals also represent an agility<br />
aspect but are not driving the design because the slew times can be reasonably long. In order to<br />
reduce the negative impact of flexible modes on the manoeuvre time, an active damping strategy for<br />
the solar array modes could be assessed. This is currently kept in mind as a back-up solution but, so<br />
far, only the performance of actuators without such a damping technique has been evaluated for this<br />
study.<br />
For the assessments performed in this chapter it is assumed that the AOCS architecture to support the<br />
various operational modes (transfer, acquisition, nominal operation, orbit maintenance, safe mode)<br />
can be established on the basis of existing E/O or telecomm platforms (e.g. Eurostar) in order to<br />
benefit from long-time heritage, risk mitigation and cost minimisation. The focus for this study is to<br />
select a suitable set of sensors and actuators which fit with the dedicated pointing and agility<br />
requirements of the Geo-Oculus mission.<br />
4.5.6.2 Pointing budgets for AOCS<br />
The requirements in Table 4.5-3 represent requirements assigned for the AOCS performance (thermoelastic<br />
distortions are covered by a separate budget):<br />
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Table 4.5-3: AOCS 100% pointing requirements<br />
Pointing<br />
Index<br />
Preliminary Values (100%) Remark<br />
APE ±100 μrad Derived to minimise overlap of neighbouring images<br />
PDE 0.5 μrad over 100 ms<br />
1.0 μrad over 100 ms<br />
For VNIR7 channel (panchro)<br />
For other channels<br />
AME ±100 μrad Currently the same value assumed as for APE<br />
4.5.6.3 AOCS sensors and actuators<br />
Given the high performance requirements for Geo-Oculus, only high performance sensors are<br />
considered . The main sensor will be a star tracker (STR) which will be operated together with an<br />
inertial measurement unit (IMU) in a gyro-stellar estimator set-up (GSE). In the GSE the STR data is<br />
combined with the IMU data to benefit from the advantages of both sensors. The STR provides noisy<br />
but stable attitude information and the IMU provides low-noise data which drifts over time. The IMU<br />
cancels the noise from the STR and the STR cancels the drift in the IMU to a large extent. Several<br />
options for STR are available on the European market:<br />
• Sodern Hydra<br />
• Jena Optronik Astro APS<br />
• Galileo AA<br />
Several options also exist for the IMU selection:<br />
• EADS Astrium Astrix 120 HR<br />
• EADS Astrium Astrix 200 GEO<br />
• Northrop Grumman Scalable SIRU<br />
The baseline STR is the Astro APS and the baseline IMU is the Astrix 200 GEO, as these are the<br />
baseline sensors for a reference mission which is similar in many respects. The performance of the<br />
three listed STR are similar and the baseline can easily be changed if necessary. For the IMUs there<br />
is a clear performance difference between the Astrix 120 on one hand and Astrix 200 and SIRU on the<br />
other hand. The performance level of the Astrix 200 and SIRU is necessary to meet the relative<br />
pointing requirements. The SIRU is produced in the US and is subject to ITAR restrictions. It can<br />
therefore not be selected as baseline.<br />
The attitude control and manoeuvre actuator selection has gone through several iterations, and the<br />
current choice stands between using Magnetic Bearing Wheels (MBW) and an Electric Propulsion<br />
System (EPS).<br />
• Rockwell Collins MBW<br />
• EPS system<br />
The Rockwell Collins (RCD) MBW is the only MBW option available in the European market. It is<br />
currently not flight proven but RCD indicates that they will have flight proven models available in 2013.<br />
Thales in Ulm, Germany, is currently developing a HEMPT based EPS system called HEMPT 3050.<br />
This system is much to powerful for fine attitude control and a theoretical, scaled down microHEMPT<br />
thruster has been developed for comparison.<br />
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Table 4.5-4: Comparison of HEMPT 3050 and microHEMPT<br />
HEMPT 3050 microHEMPT<br />
Force 30-50 mN 100-500 µN<br />
Mass flow 1.2 mg/s 12 µg/s<br />
For the final suggestion of preferred actuators for the Geo-Oculus mission, the properties of above<br />
options in terms of pointing performance, agility and fuel consumption is checked for attitude control<br />
tasks and manoeuvres in the following.<br />
4.5.6.4 Attitude control performance<br />
The attitude control performance of Geo-Oculus has been derived from simulations using a high<br />
performance attitude control simulator as already applied in a similar project. The baseline sensor<br />
suite has been used for simulations for both the MBW and EPS option. Below follows a summary of<br />
the attitude performance figures for steady state attitude control and manoeuvres.<br />
MBW system and performance<br />
The configuration of a MBW based actuator system is equal to that of a ball bearing reaction wheel<br />
system. Four or five MBW can be selected, depending on the impact zero-crossings has on the<br />
pointing accuracy. Five wheels are needed if zero-crossings are judged to be of importance, as a five<br />
wheel system will not experience zero-crossings even if one of the wheels should fail. A four wheel<br />
configuration is the minimum for redundancy, but will have wheel zero-crossings if one wheel should<br />
fail.<br />
The worst values of the analysed performances are<br />
• APE 14.1 μrad, AME 11.3 μrad, PDE 0.32 μrad/100ms (all values for 100% probability).<br />
All performance values are better than the requirements of Table 4.5-3.<br />
EPS system and performance:<br />
The EPS analysis is based on the recently finished HOPAS-3 study, investigating the use of EPS as<br />
the sole actuator on spacecraft in GEO. The study has been done for DLR by EADS Astrium GmbH.<br />
The baseline EPS configuration is derived from this study.<br />
The EPS based attitude control system will have a 12 thruster configuration based on the<br />
microHEMPT. A HEMPT 3050 based system has been analysed and rejected based on its high fuel<br />
consumption (~100 kg over 10 years) and problems with meeting the given minimum lever arm and<br />
torque level requirements. Also, the high power consumption of the HEMPT 3050 (1.2 kW per active<br />
thruster) does not allow more than 4 thrusters for attitude control to be active at the same time. This<br />
impacts the attitude control performance, especially after the completion of manoeuvres, when only<br />
attitude control thrusters are used for settling. In comparison, the microHEMPT can operate six<br />
thrusters at the same time, with a total power consumption of 0.2 kW.<br />
The attitude control thrusters are configured with four thrusters around each axis, as can be seen in<br />
Figure 4.5-8. The special configuration has been developed especially to meet the lever arm and<br />
thruster plume direction requirements, and at the same time provide decent torque levels and fuel<br />
consumption.<br />
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Figure 4.5-8: EPS thruster configuration<br />
The analysed performance values are<br />
• APE 32.7 μrad, AME 11.3 μrad, PDE 0.21 μrad/100ms (all values for 100% probability).<br />
The APE and AME steady-state performances are well below the requirements with the APE being<br />
slightly worse than for the MBW solution. The EPS attitude performance is largely a function of the<br />
thruster controller dead zone. The dead zone determines the level the torque command must reach<br />
before the thruster will start firing. It is there to avoid excessive thruster firing due to noise and must be<br />
selected large enough to prevent continuous firing and counter firings. A small dead zone gives higher<br />
pointing accuracy whereas a large dead zone reduces the fuel consumption. The PDE steady-state<br />
performance for a EPS system is better than that for the MBW system, due to the lower torque<br />
exercised on the system from the EPS thrusters.<br />
The power and fuel consumption is acceptable but it may be further decreased by a larger dead zone.<br />
This is possible since there is still a good margin to the absolute attitude requirements. Doubling the<br />
dead zone from 1/6 of the available torque to 1/3 reduces the fuel consumption by more than 80%.<br />
The attitude performance is reduced but still within the requirements.<br />
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Table 4.5-5: EPS fuel and power consumption and system mass for attitude control over 10 years<br />
Deadzone size Fuel consumption [kg] Power consumption [kW] System mass [kg]<br />
microHEMPT DZ 1/3 0.7 0.2 61<br />
DZ 1/6 3.9 0.2 64<br />
HEMPT 3050 120 2.3 307<br />
Thrusters active [-]<br />
Thrusters active [-]<br />
2<br />
1.5<br />
1<br />
0.5<br />
Thruster activation timeline (DZ: 1/6)<br />
0<br />
1000 1500 2000 2500 3000<br />
Time [s]<br />
3500 4000 4500 5000<br />
Thruster activation timeline (DZ: 1/3)<br />
2<br />
1.5<br />
1<br />
0.5<br />
0<br />
1000 1500 2000 2500 3000<br />
Time [s]<br />
3500 4000 4500 5000<br />
Figure 4.5-9: Thruster activation timelines for dead zone sizes of 1/6 (top) and 1/3 (bottom) of<br />
maximum available torque<br />
The analysed performance values are<br />
• APE 61.1 μrad, AME 11.2 μrad, PDE 0.24 μrad/100ms (all values for 100% probability).<br />
A major issue with the EPS based attitude control system is that it is based upon currently nonexistent<br />
technology which is believed to be available on the European market within the next five to<br />
seven years. No major technological showstoppers are identified for the development of a<br />
microHEMPT system but should such a system prove itself to be infeasible for use on Geo-Oculus,<br />
other technologies such as microHET and FEEPT thrusters can be considered.<br />
4.5.6.5 Manoeuvre performance<br />
One of the key issues for Geo-Oculus is the ability to image multiple locations throughout Europe<br />
several times per day. The limiting factor for manoeuvrability is the available torque to perform the<br />
manoeuvre in the shortest time possible, and the settling time needed after the manoeuvre to reach<br />
the required attitude performance again. Manoeuvrability is only required around the x- and y-axis, to<br />
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scan North/South and East/West, respectively. A summary of the manoeuvres required over 24 h is<br />
presented in Table 4.5-6.<br />
Table 4.5-6: Manoeuvre summary<br />
Manoeuvre [deg] Daytime manoeuvres<br />
over 9 h [-]<br />
Nighttime manoeuvres<br />
over 15 h [-]<br />
Total manoeuvres<br />
over 24 h [-]<br />
Manoeuvre time<br />
allocation [s]<br />
0.25 27 0 27 70<br />
0.40 27 0 27 70<br />
2.00 324 720 1044 70<br />
All manoeuvre performance data presented in this section are based on one-axis manoeuvres.<br />
There are several options for manoeuvre actuators. The MBW option is clear, but also two EPS<br />
options, using either the HEMPT 3050 or microHEMPT, have been considered. The MBW option uses<br />
the same wheels for attitude control and manoeuvres, as the torque output from one wheel is<br />
scaleable from 0 to 400 mNm. The wheel drive electronic quantisation is 55 µNm. The EPS can not<br />
throttle its output to the same degree and an additional EPS system is needed for manoeuvres. The<br />
first option is to use the HEMPT 3050 thrusters for manoeuvres which can produce a torque of ±85<br />
mNm around each axis.<br />
In the EPS analysis another option has been introduced as well. MicroHEMPT for attitude control has<br />
a few advantages such as lower fuel and power consumption and system mass, and one obvious<br />
drawback: the long resulting manoeuvre time. An overview of the theoretical, time optimal manoeuvre<br />
time for the various options are listed in Table 4.5-7. Note that these numbers do not allocate time for<br />
a settling period after the completion of the manoeuvre.<br />
Table 4.5-7: Theoretical, time optimal manoeuvre times for HEMPT 3050 and microHEMPT based<br />
configurations, and duty cycle over 24 h<br />
Manoeuvre [deg] HEMPT 3050 (30 mN) MicroHEMPT (2 mN) MicroHEMPT (3 mN) MBW (400 mN)<br />
0.25 24.2 s 132.6 s 108.3 s 9.4 s<br />
0.40 30.6 s 167.8 s 137.0 s 11.9 s<br />
2.00 68.5 s 375.2 s 306.3 s 26.5 s<br />
Duty cycle over 24 h 84% 463% 378% 33%<br />
Est. fuel consumption 635 kg<br />
over 10 years<br />
18 kg 27 kg 0 kg<br />
In the following, the complete manoeuvres are assessed which includes the time where the actuators<br />
are operated (corresponds to the manoeuvre times of above table) plus the settling time (mainly driven<br />
by the solar array) needed to reach the pointing requirements. Only the APE and PDE are evaluated<br />
according to the requirements of Table 4.5-3. The absolute measurement error AME shows the same<br />
performance before and after a manoeuvre such that this parameter does not need to be checked.<br />
From above figures, the EPS options could already be eliminated since the HEMPT solution needs too<br />
much fuel (635 kg of noble gas) and the micro HEMPT system is not suitable because of a duty cycle<br />
significantly higher than 100% which means that all the required manoeuvres can not be performed in<br />
the required time frame. Nevertheless, all of above options are evaluated to give an impression what<br />
is feasible with each of the options.<br />
MBW performance<br />
The MBW option has the most available torque for both the manoeuvres and the stabilizing after the<br />
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manoeuvre. It also has a close to linear torque response to the commanded torque within its torque<br />
limits, something that favours the MBW option over the EPS solution.<br />
In the following table the total times needed for the selected 3 typical manoeuvres are listed. The<br />
criterion for the end of the manoeuvre is the fulfilment of the pointing requirements according to Table<br />
4.5-3 where for the PDE the less stringent value has been considered (1μrad / 0.1sec).<br />
Table 4.5-8: Total MBW manoeuvre time, including settling time for APE and PDE over 0.1 sec<br />
Manoeuvre [deg] APE [s] PDE 0.1 s [s]<br />
0.25 8 58<br />
0.40 14 129<br />
2.00 63 126<br />
It can be seen that the APE settling time for all manoeuvres is within the allocated manoeuvre time of<br />
70 seconds. However, the PDE settling is far longer due to vibrations of the solar array induced by the<br />
manoeuvre. The length of the PDE settling can be possibly reduced by increasing the stiffness of the<br />
solar arrays and by lowering the applied torque when performing a manoeuvre, thus increasing the<br />
manoeuvre duration. Another option is to increase the simulated damping factor of the solar array,<br />
thus reducing the PDE settling time. A conservative value of 0.3% is used as default, but increasing<br />
the damping factor to 0.5% reduces the PDE settling from 126 to 74 seconds for a 2 deg manoeuvre.<br />
It is assumed that if the manoeuvre is optimized further, it will be possible to reduce all settling times to<br />
below 70 seconds.<br />
EPS performance:<br />
An EPS based manoeuvre system requires higher torques than the EPS based attitude control<br />
thrusters can produce in nominal operations. Therefore an additional set of manoeuvre thrusters are<br />
needed.<br />
A manoeuvre system based on the HEPMT 3050 thrusters can produce a torque of ±85 mNm around<br />
the x- and y-axis, giving the theoretical time optimal manoeuvre times and estimated fuel consumption<br />
listed in Table 4.5-7.<br />
It is also possible to use an additional set of microHEMPT thrusters, in combination with the attitude<br />
control thrusters, for manoeuvres. This will cause the manoeuvre times to increase dramatically, as<br />
the available thrust force only will be in the range of a few mN. This requires the microHEMPT thruster<br />
to be able to operate at both 100 µN and 500 µN. The 100 µN operational mode is used for attitude<br />
control and 500 µN for manoeuvres. If the two thruster pairs are operated at maximum force<br />
simultaneously, a total of 2 mN will be available. To reduce total manoeuvre time additional sets of<br />
microHEMPT thrusters can be added.<br />
When using the HEMPT 3050 configuration the following performance can be achieved.<br />
Table 4.5-9: Total HEMPT 3050 manoeuvre times, incl. settling time for APE and PDE over 0.1 sec<br />
Manoeuvre [deg] APE [s] PDE 0.1 s [s]<br />
0.25 18 20<br />
0.40 153 57<br />
2.00 300 160<br />
It can be seen from Table 4.5-9 that the APE settling is the largest problem. Even though the time<br />
optimal manoeuvre time for the HEMPT 3050 configuration is below the 70 seconds allocated to<br />
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manoeuvres, the APE settling time is so long that the duty cycle becomes as high as 368%. This only<br />
allows Geo-Oculus to perform 27.2% of the required manoeuvres over 24 hours, which is around the<br />
same performance as a configuration using microHEMPT can achieve. The 2 mN option can perform<br />
21.7% of the required manoeuvres and the 3 mN option can perform 26.5% of the required<br />
manoeuvres, assuming that the low force manoeuvres are so slow that no settling time is needed after<br />
the completion of the manoeuvre. In this context it can be seen that some fuel mass can be saved<br />
using microHEMPT for manoeuvres.<br />
Table 4.5-10: Total manoeuvre times for HEMPT 3050, microHEMPT and MBW based<br />
configurations, duty cycle over 24 h<br />
Manoeuvre [deg] HEMPT 3050 (30 mN) MicroHEMPT (2 mN)** MicroHEMPT (3 mN)** MBW (400 mN)<br />
0.25 20 s 132.6 s 108.3 s 58 s<br />
0.40 153 s 167.8 s 137.0 s 70 s *<br />
2.00 300 s 375.2 s 306.3 s 70 s *<br />
Duty cycle over 24 h 368% 463% 378% 89%*<br />
*Assumption, performance currently not achieved in simulations<br />
** Theoretical performance, no settling time included<br />
Overall performance<br />
A MBW configuration provides very good attitude control performance and is assumed to be able to<br />
meet the manoeuvre requirements with some additional tuning of certain parameters. The EPS<br />
attitude control performance is also very good, but the manoeuvre performance is far worse than what<br />
is required. By using large EPS thrusters, the actual manoeuvre time is within the allocated time, but<br />
the settling time after a manoeuvre is much to long for the 0.4° and 2° manoeuvres due to the low<br />
available torque from the attitude control thrusters. This leads to the result that not all manoeuvres can<br />
be performed which reduces the mission value. Also, using the HEMPT 3050 causes a high fuel and<br />
power consumption. The microHEMPT option for manoeuvres has not been investigated in detail, but<br />
it is clear that the fuel consumption will be lower, and that such a configuration will be able to perform<br />
as many manoeuvres as a HEMPT 3050 system if the assumption that no settling time is needed after<br />
the manoeuvre completion holds.<br />
The MBW option is the clear favourite of the two, and is selected as a baseline nominal mode<br />
actuator. The only drawback is that the MBW development is in an early phase, and might not be<br />
available as expected in 2013. If indications of significant delays in the MBW development, or<br />
shortcomings in the performance surfaces, the EPS option can be considered again. Figure 4.5-10<br />
shows the baseline AOCS configuration and Table 4.5-11 summarizes in which operational modes the<br />
various AOCS equipment is used.<br />
The hybrid option discussed in [RD 7] has not been considered further, as it leads to unnecessary<br />
high costs and system complexity. The CPS system can perform wheel offloading and East/West<br />
station keeping every three weeks, with a total outage of less than 10 min each time. The North/South<br />
station keeping is performed twice every year. If the total mass of the spacecraft should exceed the<br />
capabilities of the desired launcher, the EPS system can again be considered as it has the potential to<br />
lower overall system mass.<br />
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Figure 4.5-10: Baseline AOCS configuration<br />
Table 4.5-11: AOCS equipment and modes<br />
Sensors Actuators<br />
Mission phase Coarse GYP Earth IRES Star STR Sun BASS Fine LiASS IMU MBW CPS<br />
Transfer & Acquisition � � � � �<br />
On station: Normal mode � � �<br />
On station: Station Keeping � � � �<br />
Safe mode Transfer � � �<br />
Safe mode On-Station (�) � �<br />
4.5.7 Propulsion System<br />
The Propulsion System of the mission is composed by:<br />
• Chemical Propulsion System (CPS): it is intended for GTO-GEO Transfer and, in the option<br />
without EPS, also for Station Keeping, wheel off-loading and de-orbiting.<br />
• Electric Propulsion System (EPS): it is optional and, if present, it is intended for reaction<br />
wheel off-loading or for active pointing (for an AOCS without reaction wheels).<br />
The following table summarises the possible options and the tasks allocated to CPS and EPS.<br />
Table 4.5-12: Propulsion System Options for Geo-Oculus<br />
Option 1 2 3<br />
No EPS EPS + MBWs EPS only<br />
GTO-GEO CPS CPS CPS<br />
NSSK + EWSK CPS EPS EPS<br />
Deorbiting CPS CPS EPS<br />
Wheel off-loading CPS EPS /<br />
Pointing Manoeuvres / / EPS<br />
The main propulsion requirements, as deriving from AOCS analysis, are the following:<br />
• Transfer ΔV: 1500 m/s + 50 m/s margin,<br />
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• De-orbit ΔV: 10 m/s<br />
• NSSK ΔV 400 m/s,<br />
• EWSK ΔV 10 m/s<br />
• Wheel off-loading + Rate Damping + Safe Mode = 15 kg (applicable to CPS only)<br />
The main trade-off regarding the propulsion system is about the use of EPS:<br />
• no EPS onboard (propulsion tasks entirely performed by CPS, all attitude control tasks<br />
performed by reaction wheels)<br />
• EPS only for reaction wheels downloading (GTO-GEO transfer and de-orbiting performed by<br />
CPS, fine pointing manoeuvres for image acquisition performed by reaction wheels)<br />
• EPS for attitude control (GTO-GEO transfer and de-orbiting performed by CPS, no reaction<br />
wheels)<br />
In this scenario, the results of the various propulsion system options are then to be analysed in a<br />
trade-off analysis at system level, i.e. involving also AOCS and main satellite level design choices.<br />
Therefore, the purpose of this section is just to prepare the input for such trade-off analysis.<br />
In following sections, the mass budgets (main trade-off criteria) for the options in Table 4.5-12 are<br />
derived from requirements and briefly analysed.<br />
4.5.7.1 CPS<br />
Geo-Oculus is a geostationary mission. Astrium has a long heritage of supplying geostationary<br />
spacecraft, dating back to the 1970s. In addition to the telecom fleet there is a successful fleet of<br />
scientific and earth observation missions including Mars Express which has achieved two years in<br />
Mars Orbit, Venus Express currently in Venus orbit, and Rosetta and Cluster missions which are now<br />
flying with bipropellant NTO / MMH propulsion systems. The combined experience of this wealth of<br />
heritage shall enable Astrium to complete the study and return conclusions for the optimal CPS to<br />
meet the Geo-Oculus mission requirements.<br />
Astrium currently has 3 generic platforms for geostationary missions which can be considered for Geo-<br />
Oculus. These are:<br />
Eurostar 2000+<br />
MON-3/MMH bipropellant propulsion system, an evolution of the Eurostar 2000 platform, featuring 4<br />
propellant tanks, on a central cylinder supported structure<br />
Eurostar 3000<br />
MON-3/MMH bipropellant propulsion system, a larger version of the E2000+ platform. The design has<br />
been expanded to include larger tanks to increase the mission capabilities of the design, featuring 4<br />
propellant tanks on a central cylinder supported structure, in 4 sizes<br />
Eurostar 3000C<br />
The E3000C is an evolution of the E3000 design, based upon the successful Mars Express and Venus<br />
Express spacecraft. It remains a MON-3/MMH bipropellant propulsion system, with heritage from<br />
E3000 and Mars/Venus Express, but the platform is smaller than both E3000 and E2000+ to suit a<br />
smaller payload requirement. It features 2 propellant tanks are (supported by a “single H” type<br />
structure, demonstrated by the Mars/Venus Express spacecraft). As with Eurostar 3000, the tank size<br />
is interchangeable.<br />
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The CPS considered for Geo-Oculus is the same of E3000 platform. The variable is given by the<br />
propellant tanks capacity.<br />
The CPS main features are:<br />
• He pressurized bi-propellant system (MMH + MON-3)<br />
• Four cylindrical tanks / one central pressurant tank<br />
• Common propellant storage and feed system<br />
• One 450 N LAE<br />
• Seven pairs of 10N RCT’s<br />
Figure 4.5-11: Geo-Oculus CPS Schematic<br />
The basic CPS sizing option is performed for option 1 of Table 4.5-12 which is the solution without<br />
EPS. The result is presented in the following table.<br />
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Table 4.5-13: CPS Mass Budgets for considered options<br />
Budget Element Option 1<br />
Total CPS Dry Mass 162.9<br />
Total residual 26.82<br />
Total CPS EOL Mass 189.7<br />
Useful Propellant load 1793.9<br />
Total CPS BOL Mass 1983.6<br />
4.5.7.2 EPS<br />
The Electric Propulsion System (EPS) architecture is deriving from:<br />
• Option 2: existing and available commercial solutions<br />
• Option 3: iterations with AOCS definition and is the one used for most simulations of section<br />
4.5.6 (Attitude and Orbit Control)<br />
The necessary iterations with AOCS definitions are needed to optimise on board resources request<br />
(power, mass, volume) while fulfilling mission requirements.<br />
The architecture is summarised in the following table:<br />
Table 4.5-14: EPS Architecture summary for the two options, Option 2 and Option 3<br />
Option 2: EPS + MBW Option 3: EPS only<br />
Main Manoeuvre Thruster (MMT) 2 main + 2 redundant<br />
+ 2 thruster pointing mechanisms<br />
8 main + 8 redundant<br />
MMT Thrust 80 mN 30 mN<br />
MMT Duty Cycle 100% during NSSK<br />
85% during<br />
manoeuvres (twice a day)<br />
re-pointing manoeuvres<br />
MMT PCU 1 main+1 redundant, each driving 2 2 PCU, each driving 4 main + 4<br />
thrusters<br />
redundant MMT<br />
Fine Pointing Thruster (FPT) / 12 main + 12 redundant<br />
FPT Thrust /<br />
FPT Duty Cycle / 1% during fine pointing<br />
FPT PCU / 2 PCU, each driving 4 main + 4<br />
redundant MMT<br />
EPS for Option 2<br />
The EPS for Option 2 can be based on existing EPS for Eurostar 3000, with an architecture as per<br />
Option 2 of Table 4.5-14, using SPT-100 as thrusters (HET type, 80 mN of thrust, 1510 s of Isp).<br />
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PPU B*<br />
FU<br />
XST<br />
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FU<br />
FDV<br />
FDV<br />
XFC XFC XFC XFC<br />
Orientation mechanism (PY) Orientation mechanism (MY)<br />
HET2PY<br />
Thruster<br />
FU<br />
XRFS<br />
VBA<br />
HET1PY<br />
Thruster<br />
PV<br />
XEF<br />
HPT (2)<br />
Isolation and regulator<br />
solenoid valves<br />
Plenum<br />
FDV<br />
LPT (4)<br />
HET1MY<br />
Thruster<br />
PPU A*<br />
FU<br />
HET2MY<br />
Thruster<br />
PY TMA MY TMA<br />
TSU TSU<br />
Figure 4.5-12: GeoOculus EPS Schematic, Option 2, Eurostar 3000 type
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As secondary trade-off, the same architecture can be analysed by using HEMPT thrusters with 80 mN<br />
nominal thrust and an Isp of 3000 s.<br />
Following table summarises the results for mass and power<br />
Table 4.5-15: EPS Option 2 Mass Budgets for considered thruster options (figures in kg)<br />
SPT-100 80 mN HEMPT<br />
TOTAL EPS DRY MASS 92.8 103.2<br />
Total Propellant Load 83.0 41.1<br />
TOTAL EPS BOL MASS 175.8 144.3<br />
Table 4.5-16: EPS Option 2 Power Budgets for considered thruster options (figures in W)<br />
SPT-100 80 mN HEMPT<br />
Main Thruster(s) Assembly 2404 5553<br />
PCU & Ancillary 227 464<br />
Total 2632 6017<br />
Total (Including Margins) 2895 6619<br />
Although the system based on HEMPT technology is about 20% lighter, it needs twice the power to be<br />
run. Decreasing power to match the system based on SPT-100 technology could be achieved by:<br />
• either having a 30 mN HEMPT, but thrust times would increase around the nodes,<br />
decreasing the efficiency of firing and thus increasing the quantity of propellant to be used<br />
• or setting the HEMPT at a lower Isp, but being a more massive thruster of SPT for the same<br />
combination of Thrust and Isp, this would not be an option to be considered<br />
The final trade-off solution will depend on spacecraft level trade-off analysis.<br />
EPS for Option 3<br />
The EPS for Option 3 is based on the general architecture envisaged for carrying on AOCS<br />
simulations:<br />
• 8 Main (Attitude) Manoeuvre Thrusters (MMTs) of 30 mN each, Isp of 3000 s;<br />
• 12 Fine Pointing Thruster (FPT) providing down to 0.1 mN of thrust each<br />
The rest of the system has been completed as per Option 3 of Table 4.5-14.<br />
Two MMT options have been considered, namely HEMPT and GIT, while smaller GIT (Isp of 3000 s)<br />
and FEEP (Isp of 6000 s) have been considered as FPT, making four possible combinations for<br />
Option 3. There is little difference in propellant mass (between 574 and 578 kg); the overall EP mass<br />
and power budget are shown in Table 4.5-17and Table 4.5-18, respectively..<br />
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Table 4.5-17. EP mass budget for Option 3 [kg]<br />
Fine Pointing<br />
Thruster<br />
Table 4.5-18. EP power budget for Option 3 [W]<br />
Fine Pointing<br />
Thruster<br />
Main Manoeuvre Thruster<br />
HEMPT GIT<br />
GIT 860 902<br />
FEEP 834 876<br />
Main Manoeuvre Thruster<br />
HEMPT GIT<br />
GIT 2571 2138<br />
FEEP 2600 2167<br />
4.5.7.3 Propulsion System Summary<br />
The mass of the propulsion system for the three options is summarised in Table 4.5-19.<br />
Option 3 is not practical since too massive, complex and expensive. Option 1 is the traditional<br />
configuration and is feasible. Option 2 allows to save between 250 and 340 kg (depending on the<br />
selected technology) on Option 1 by using EP for station-keeping. However, this mass reduction<br />
should be somewhat reduced as it does not take into consideration the additional mass due to an<br />
increase in solar array as well as batteries, and potentially PCDU too. An in-depth analysis would be<br />
required at a later stage to determine the better of Options 1 and 2.<br />
Table 4.5-19. Geo-Oculus propulsion options summary<br />
S/C dry mass (no PS)<br />
EPS dry mass<br />
CPS dry mass<br />
TOTAL PS DRY MASS<br />
EPS Total Prop. load<br />
CPS Total Prop. load<br />
TOTAL PROP. LOAD<br />
TOTAL PS MASS AT LAUNCH<br />
Power requirements [W]<br />
Option 1<br />
1668.3<br />
/<br />
189.7<br />
189.7<br />
/<br />
1793.9<br />
1793.9<br />
1983.6<br />
/<br />
Option 2<br />
HET<br />
1668.3<br />
92.8<br />
157.2<br />
250.0<br />
83<br />
1404.2<br />
1487.2<br />
1737.2<br />
2895<br />
4.5.8 Structure and Thermal Concept<br />
4.5.8.1 Structure<br />
Option 2<br />
HEMPT<br />
1668.3<br />
103.2<br />
133.8<br />
237.0<br />
41.1<br />
1364.4<br />
1405.5<br />
1642.5<br />
6619<br />
Option 3<br />
HEMPT+FEEP<br />
1668.3<br />
287.4<br />
189.7<br />
477.1<br />
573.9<br />
1838.3<br />
2412.2<br />
2889.3<br />
Option 3<br />
GIT+GIT<br />
1668.3<br />
354.4<br />
194.3<br />
548.7<br />
578.2<br />
1888.0<br />
2466.2<br />
3014.9<br />
2138<br />
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2600<br />
Option 3<br />
HEMPT+GIT<br />
1668.3<br />
311.7<br />
189.7<br />
501.4<br />
576.3<br />
1857.7<br />
2434<br />
2935.4<br />
2571<br />
Option 3<br />
GIT+FEEP<br />
1668.3<br />
330.1<br />
189.7<br />
519.8<br />
575.6<br />
1869.5<br />
2445.1<br />
2964.9<br />
Structure design<br />
Figure 4.5-13 depicts the structure of the satellite with 4 propellant tanks, based on Astrium’s Eurostar<br />
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The overall S/C structure has a classical "box" shape with a central cylinder (800 mm) as the main<br />
structural load path to the launcher. The design will fulfil the satellite strength and stiffness<br />
requirements. The instrument is mounted at the top of the platform.<br />
Currently, the instrument is connected to the platform by means of three isostatic mounts. It is<br />
recommended for future work to revisit the instrument mechanical configuration so that it is supported<br />
by four sets of isostatic mounts, turned upside down. This would allow the “head” of an isostatic mount<br />
to be fixed to the top of the shear walls, and thus provide a better load path than what is currently<br />
depicted. Clearly, this topic has to be iterated with the mechanical design of the payload considering<br />
the mechanical load path and thermo-elastic distortions.<br />
It should also be noted that due to the width of the instrument being much larger than the central<br />
cylinder, it is not possible to fix the isostatic mounts in the current configuration directly to the central<br />
cylinder.<br />
Figure 4.5-13. Satellite Configuration showing the main structure<br />
Primary Structure<br />
The satellite primary structure consists of,<br />
• A launcher interface ring<br />
• A central cone/cylinder structure<br />
• 4 shear walls<br />
• ±X Upper and lower floors<br />
• Upper and lower tank floors<br />
• Tank support struts<br />
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YS/C<br />
XS/C<br />
ZS/C
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The primary structure design is founded on the E3000 heritage design.<br />
The central structure consists of a lower conical section and two upper cylinder segments. With the<br />
exception of the Aluminium Alloy upper ring segment all other parts are made from filament wound<br />
CFRP sandwich panel section. Bonded rings are located at each end of the CFRP cone cylinder and<br />
at the intersection between the cone and cylinder.<br />
Considering each of these rings the lower ring provides the launch vehicle adapter (LVA) interface,<br />
through this ring interfacing with an opposite ring on the launch vehicle, combined with the use of a<br />
clamp band, the launch vehicle interface is made. The at the cone cylinder junction supports the tank<br />
floors, to this an important part of the central structure, the tank support struts, linking this floor at the<br />
tank interfaces to the LVA ring, these provide axial support to the tank. The upper tank floor, which<br />
provides lateral support only allowing tank expansion, mounts to the ring at the top of the cylinder.<br />
Focusing on E3000 heritage all shear wall and tank floors are made from Aluminium alloy sandwich<br />
panels. If future, more detailed, distortion analyses would show the need for a CFRP panels, the<br />
material of the shear walls could be switched from aluminium skin to CFRP skin but retaining the<br />
Aluminium alloy honeycomb core.<br />
Secondary Structure<br />
The satellite secondary structure consists of,<br />
• ±Z equipment panels and ±Y closure panels<br />
• Local support brackets/panels as e.g.<br />
− Connector brackets<br />
− Thruster supports<br />
− EMC covers<br />
− Liquid apogee engine and pressurant tank supports<br />
− etc.<br />
Focusing on E3000 heritage the equipment panels are made from Aluminium alloy sandwich panels.<br />
Currently proposed is the use of CFRP skinned panels, however if future, more detailed, distortion<br />
analyses would show that Aluminium Alloy panels could be accommodated, the material of the<br />
equipment panel walls could be switched from CFRP to aluminium skin but retaining the Aluminium<br />
alloy honeycomb core.<br />
The panel thickness will be typically 35-40mm. The design and the materials of local support<br />
structures will be defined in a later project phase.<br />
Solar Array<br />
The solar array substrate will be a lightweight CFRP sandwich panel with typically 20 mm thickness.<br />
Structure Load Paths<br />
The circular central structure will collect the individual loads over its height and will ensure a<br />
homogeneous load distribution over the launcher interface circumference. Hence the launcher I/F<br />
overflux requirement will be fulfilled.<br />
The axial (in-plane) equipment panel loads will be transferred to the central cylinder via the shear<br />
webs.<br />
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The lateral in-plane equipment panel loads will be transferred to the central cylinder via the floors in<br />
shear. The lateral out-of-plane equipment panel loads will be transferred to the central cylinder via the<br />
shear webs and floors in tension/compression.<br />
The tanks are grouped closely around the central structure. Thus the tank load path is very short and<br />
mass-effective, the tanks will laterally be supported by the tank floors and axially by struts from the<br />
lower tank boss to the LVA ring side.<br />
CFRP Outgassing<br />
Water and carbohydrate evaporation in the vicinity of the instruments needs to be limited to avoid e.g.<br />
ice on cold optical surfaces. Any critical CFRP surface shall be sealed by an aluminium barrier foil of<br />
typically 20 microns. In the past, this was done successfully e.g. on all inner surfaces of the 6.8 m long<br />
XMM telescope tube structure.<br />
Distortion Assessment<br />
Pointing stability depends on thermal and moisture release distortions and therefore on configuration,<br />
material selection, method of construction, changes in average temperature and temperature<br />
gradients.<br />
A similar stringent requirement also exists for MTG. It is therefore assumed that, through the high<br />
degree of similarity between the two satellites, that the Geo-Oculus environment would yield<br />
distortions of similar magnitude. Therefore, it is confidently believed that distortions should not be an<br />
issue for Geo-Oculus.<br />
Launch Vehicle Vibrations<br />
The Frequency requirements for the S/C hard mounted at the I/F to the launch adapter are taken from<br />
Soyuz since it is the worse case compared to Ariane 5:<br />
• 15 Hz in lateral + 15% margin<br />
• 35 Hz in longitudinal + 15% margin<br />
Considering stiffness the main driver for the fist axial frequency is tank mass and the stiffness of the<br />
underlying support. To optimise stiffness performance the heavy and light tanks are diagonally<br />
opposite as to maintain a central centre of gravity position, also the supporting struts are categorised<br />
into heavy or light struts each providing the best support stiffness for the respective tank mass.<br />
The first lateral mode is largely driven by the X axis centre of gravity position, this is heavily influenced<br />
by the mass of any +X top floor mounted equipments and instruments.<br />
To estimate the first lateral frequencies of the Geo-Oculus spacecraft, the performance of the reported<br />
performance of MTG spacecraft has been examined and scaled as appropriate. Scaling has taken<br />
account of the comparative propellant mass and instrument mass associated with the Geo-Oculus<br />
spacecraft.<br />
The MTG spacecraft is viewed as a suitable foundation given the spacecraft architecture is similar to<br />
Geo-Oculus and likewise is largely based on E3000 heritage, the most significant differences is an<br />
extra 150kg approx located on the spacecraft +X floor and an extra propellant mass of 178Kg.<br />
The first lateral and longitudinal frequencies are as follows,<br />
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Direction MTG Scaled MTG Requirement Margin<br />
Lateral 19.32Hz 18.86Hz 17.25Hz 1.61Hz<br />
Longitudinal 40.65Hz 40.37Hz 40.25Hz 0.12Hz<br />
4.5.8.2 Thermal Control<br />
Thermal Environment<br />
The Geo-Oculus spacecraft will circle the Earth in a geostationary orbit. It is positioned directly over<br />
the equator and follows it’s path in the equatorial plane at a speed matching the Earth’s rotation. Thus,<br />
the spacecraft completes one rotation around its North / South axis per day.<br />
The sun will traverse through an angle of ±23.5° perpendicular to the orbit plane during the year. The<br />
extremes will occur at the Winter Solstice and the Summer Solstice. Eclipses will occur during the<br />
Equinox seasons, with maximum eclipse duration of 72 minutes. The North face of the spacecraft will<br />
receive direct solar illumination for 6 months centred on the Summer Solstice, while the South face will<br />
receive direct solar illumination for 6 months centred on the Winter Solstice. The other faces of the<br />
spacecraft will receive varying solar illumination during each day.<br />
The Earth varies it’s distance from the Sun over a period of 1 year. This means that the solar constant<br />
at Earth’s location changes over the year from 1420 W/m 2 at Winter Solstice to 1327 W/m 2 at Summer<br />
Solstice.<br />
Thermal Control Concept<br />
The Geo-Oculus spacecraft body thermal control will rely primarily on passive means supported by<br />
electrical heaters. The North and South faces of the spacecraft are used as the main heat rejection<br />
paths. Externally they will be covered by Optical Solar Reflectors (OSR). The exact area of OSRs<br />
exposed to space will be regulated by the use of Multi-Layer Insulation (MLI).<br />
The inside of the panels will have a black finish. Aluminium doublers and heat pipes, as appropriate,<br />
will be used to spread the heat within the panel. All electronic units are mounted inside the spacecraft<br />
primary structure. Heat transfer from the dissipating units to the radiators relies mainly on conduction..<br />
Thermal Performance<br />
The total dissipation of the equipments on the spacecraft is 1402 watts. 423 watts is the dissipation of<br />
the externally mounted units, leaving 979 watts dissipated within the spacecraft body. The payload<br />
electronics, 300 watts, is mounted on the North radiator. In addition, some of the bus equipment will<br />
also be mounted on the North panel such that the total amount of heat rejection capability adds up to<br />
479 watts. The rest of the spacecraft bus electronics, 500 watts, is mounted on the South radiator.<br />
This allows the calculation of the radiator sizes and the required heater power. The analysis results<br />
are shown in Table 4.5-20 and Table 4.5-21.<br />
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Table 4.5-20: Thermal Results for North Radiator<br />
North Radiator<br />
Dissipation 479 watts<br />
Margin 153 watts<br />
Total dissipated 632 watts<br />
Required radiator area 3.24 m 2<br />
Heater power – Equinox sunlight 50 watts<br />
Heater power – Equinox eclipse 105 watts<br />
Table 4.5-21: Thermal Results for South Radiator<br />
South Radiator<br />
Dissipation 500 watts<br />
Margin 160 watts<br />
Total dissipated 660 watts<br />
Required radiator area 3.54 m 2<br />
Heater power – Equinox sunlight 88 watts<br />
Heater power – Equinox eclipse 139 watts<br />
The North and South panel provide up to about 5 m 2 of radiator surface each which leaves sufficient<br />
margin for further evolution.<br />
4.5.9 Satellite Budgets<br />
Geo-Oculus Budgets<br />
S/C Mass<br />
Power<br />
4.6 Ground Segment<br />
Propulsion<br />
Dry Mass 1858 kg<br />
Launch Mass 3652 kg<br />
Power Demand 1800 W<br />
S/A Size (installed) 11 m 2<br />
Communication<br />
Battery 135 Ah<br />
Tanks 4x406 ltr<br />
PDT 250 Mbit/sec<br />
4.6.1 Ground Segment Architecture<br />
The architecture of the Ground Segment for the Geo-Oculus system takes into consideration the<br />
heritage of the Agency in operating EO satellites and within this context the consistency of the<br />
functionalities and of the implementation solutions with other EO systems operated by the Agency. In<br />
particular, the interoperability of Geo-Oculus with other EO systems serving the GMES needs is of<br />
paramount importance.<br />
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4.6.1.1 High Level Functional Architecture<br />
At the first level of breakdown, the functional architecture identifies two main components:<br />
• The Flight Operations Segment (FOS),<br />
• The Payload Data Ground Segment (PDGS).<br />
This breakdown is illustrated and further refined in Figure 4.6-1.<br />
Mission Control<br />
System<br />
• Spacecraft Operational<br />
Database<br />
• Housekeeping Telemetry<br />
Processing<br />
• Time Management<br />
• Telecommand<br />
• Mission Planning<br />
• On-Board Software<br />
Maintenance<br />
• Ground Station Network<br />
Interface<br />
• Off-line analysis<br />
• Authentication and<br />
Encryption<br />
Facilities Ground Segment<br />
• Acquisition<br />
• Ingestion<br />
• Processing / reprocessing<br />
• Archiving and Inventory<br />
• Production Requests Handling<br />
• Dissemination<br />
• Circulation<br />
• Monitoring and Control<br />
FOS<br />
Flight Dynamics<br />
• Orbit Determination,<br />
Prediction and Control<br />
• AOCS Monitoring<br />
• AOCS Command<br />
Generation<br />
• Test and Validation<br />
Spacecraft Simulator<br />
• Platform Model<br />
• Payload Model<br />
• Ground Segment Model<br />
Ground Stations and<br />
Networks<br />
• TMTC Ground Stations<br />
• NDIU<br />
• PSS<br />
• Ground Communication<br />
Network<br />
User Services & Mission<br />
Planning<br />
• General Web<br />
• Catalogue<br />
• On-line Ordering and Order Handling<br />
• User Management<br />
• Mission Planning<br />
• Help and Documentation Desk<br />
• Statistics & <strong>Report</strong>ing<br />
PDGS<br />
Figure 4.6-1: Geo-Oculus -Ground Segment breakdown into Domains and Functions<br />
Sensor Performance, Products<br />
& Algorithms<br />
• Routine Quality Control<br />
• Product Quality Control<br />
• Product Calibration<br />
• Product Validation<br />
• Instrument Calibration<br />
• Processing and Instrument Data<br />
Files Generation<br />
• Instrument Performance Monitoring<br />
• Algorithms and Instrument<br />
Processing Facility Development<br />
• Instrument Processing Facility<br />
Maintenance & Evolution<br />
• User Support<br />
• Precise Orbit Determination<br />
4.6.1.2 Proposed Architecture<br />
In the following, the proposed overall architecture of the Geo-Oculus Ground Segment is presented.<br />
The single elements of the architecture will be considered in the subsequent subsections.<br />
The Geo-Oculus Ground Segment Architecture features:<br />
For Mission Monitoring and Control:<br />
• A Flight Operations Control Centre,<br />
• a nominal TM/TC station,<br />
• a back-up TM/TC station and<br />
• a network of additional TM/TC stations used only for LEOP.<br />
For Payload Data Reception, Exploitation and Processing:<br />
• A “core” Payload Data Ground Segment and<br />
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• one or several deployable Payload Data Reception and Processing facilities.<br />
The Figure below depicts the Ground Segment Architecture. It also identifies the main internal and<br />
external interfaces. The external interfaces comprise the interfaces to the users and additional data<br />
sources, which provide auxiliary data for processing and meteorological forecast data supporting the<br />
scheduling of observations.<br />
EGSE<br />
Instrument Raw Data<br />
LEOP<br />
Network X-Band<br />
TMTC<br />
TMTC<br />
station<br />
S-Band<br />
S-Band<br />
Flight Operations<br />
Control Centre<br />
FOS PDGS<br />
Receiving<br />
Station(s)<br />
Backup<br />
Payload Data<br />
Reception<br />
Decryption, Processing,<br />
Archiving &<br />
Dissemination<br />
Users Services<br />
Coordination & Control<br />
Payload Mission<br />
Planning<br />
User<br />
Requests<br />
External<br />
Auxiliary Data<br />
Sensor<br />
Performance,<br />
Products and<br />
Algorithms<br />
Meteo Forecast Data<br />
+ MTG real time data<br />
Users<br />
Figure 4.6-2: Preliminary Architecture of the Geo-Oculus Ground Segment<br />
External<br />
External<br />
Data<br />
Data<br />
Sources<br />
Sources<br />
Basic Products<br />
User Reception<br />
and Processing<br />
Terminal<br />
Service<br />
Segment<br />
Customised<br />
Services<br />
Payload Data<br />
Reception<br />
Decryption,<br />
Processing,<br />
Archiving &<br />
Dissemination<br />
Basic Products<br />
The concept for the GEO-Oculus Ground Segment takes into account the particular principles of<br />
operations and technical constraints resulting from a spacecraft on a geostationary orbit. Moreover,<br />
the ground segment architecture is adapted to the needs of its customers for the different targeted<br />
applications, in terms of revisit time, flexibility in satellite observations programming and latency from<br />
observation to end of product delivery. In addition, the Ground Segment concept for the Sentinel<br />
missions, which will be operated by <strong>ESA</strong> in the GMES era, has been considered as a "loose" design<br />
guideline.<br />
The mission for an optical high spatial resolution satellite operating from a GEO orbit must be<br />
regarded as the conjunction of routine monitoring missions (sometimes termed “background” mission)<br />
and of one or several emergency monitoring missions, which are by nature less schedulable than the<br />
routine ones. An example of routine monitoring mission is the monitoring of coastal areas for which the<br />
revisit times and the response times are comparatively long. Emergency monitoring missions are e.g.<br />
fire or disaster monitoring. In this case, the revisit time as well as the response time need to be much<br />
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shorter. As a consequence, the Ground Segment baseline architecture takes into account both types<br />
of missions in order to be equally suited for routine type applications and emergency type applications.<br />
In the frame of a study for CNES, it has been envisaged that routine type applications could be served<br />
via a “Central” PDGS, due to the fact that (1) they are less time critical than emergency applications<br />
and (2) they require some value added processing (via service segment entities) in order to actually<br />
satisfy the needs of end users who are not experts in the manipulation and understanding of satellite<br />
images. This is compliant with the generic PDGS architecture model with, however, the PDGS<br />
providing products up to Level 1b only and further processing up to the delivery of a value added<br />
service being performed by specialised value added providers within the service segment.<br />
For emergency applications such as disaster monitoring (fires, floods), the study for CNES privileged<br />
an architecture model with specific decentralised user’s facilities for data reception, processing and<br />
dissemination. The crisis headquarters would be equipped with a small reception terminal and with the<br />
means to at least perform basic processing and dissemination towards the crisis actors on the<br />
operations theatre. The rationale for such a decentralised model is that it will provide quicker response<br />
times by delivering the data directly on the operations field.<br />
The decentralised model supposes that (1) the crisis headquarters or even the mobile command posts<br />
are equipped with all means to process data and (2) the end users can cope with a relatively basic<br />
level of processing.<br />
The projects currently on-going in the field of natural risks management do not foresee a direct<br />
delivery of basic products (e.g. Level 1b, which could be provided by processing on-board the S/C) to<br />
the end users but rather still foresee the involvement of specialised service providers in the value<br />
adding chain. To be easily understandable by non experts, the basic products must be geo-referenced<br />
and combined with geographical or socio-economic information, and finally be integrated into a<br />
Geographic Information System (GIS).<br />
All of these steps might be performed in the spirit of a Service Oriented Architecture (SOA) of the<br />
overall GEO Oculus PDGS even in the field. However, they are basically relying on the know-how of<br />
the service providers for executing the value adding such as Web Map Services (WMS), Web Feature<br />
Services (WCS) or Web Processing Service (WPS). Note that these services are Web-based by<br />
principle, such that a high-speed WAN connection is mandatory for this kind of applications.<br />
In addition, the processing cannot be executed without input of auxiliary data such as calibration data,<br />
orbit data and attitude data, which in turn requires specific interfaces between the “Core” Payload Data<br />
Ground Segment and the decentralised part of the Payload Data Ground Segment.<br />
So, the main advantage of the decentralised architecture model, which is to optimise the response<br />
time is counterbalanced by the need for a high-speed WAN connection even in the field (which might<br />
as well provide the products generated by the "Central" PDGS) or even the shortcomings of providing<br />
only products of low value, possibly usable only by experts.<br />
For more details, see [RD 10].<br />
4.6.2 Geo-Oculus dedicated Ground Segment issues<br />
This section addresses specific issues linked to the characteristics of the Geo-Oculus mission, i.e. the<br />
fact that (1) the satellite orbit allows permanent contact with the ground both for TM/TC operations and<br />
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for the reception of payload telemetry and (2) the mission is focused on frequent revisit for systematic<br />
observations and real time instantaneous access emergency management.<br />
4.6.2.1 Command and Control<br />
Geo-Oculus will allow permanent contact with the Mission Control System. One main advantage is<br />
that commanding the satellite is possible at any time, without having to wait for a ground station<br />
contact. On the other hand, the Mission Control System will receive a permanent flow of<br />
Housekeeping telemetry, during day and night. Hence, a reasonable schedule for operating the<br />
spacecraft should be established in order to minimise the operations cost.<br />
4.6.2.2 Flight Dynamics<br />
As a baseline for the Geo-Oculus orbit determination the spread spectrum ranging method using a S-<br />
Band repeaters and S-Band Ground Stations has been selected. The stations will have to include the<br />
necessary ranging equipment, and the Flight Dynamics System will have to process the ranging<br />
information from the stations in order to perform high accurate orbit restitution.<br />
4.6.2.3 Mission Planning<br />
Efficiency of mission planning operations is essential to gain the full benefit of the satellite agility. As<br />
far as the routine (or background) mission is concerned, the baseline observation schedule can be<br />
established well in advance and loaded on-board the satellite at given times. However, actual<br />
meteorological conditions must also be taken into account in order to minimise the likelihood of cloudy<br />
scenes and thus useless observations.<br />
For the emergency monitoring mission(s), reactivity is at stake. This means that, as soon as a catastrophic<br />
event requiring fast scheduling of an observation occurs, it shall be possible to superimpose<br />
an emergency observation to the nominal plan. As for the routine missions, the emergency observations<br />
shall also take into account the meteorological conditions during the mission planning stage.<br />
From a mission planning point of view, simple conflicts management rules can be implemented, which<br />
allow for giving priority to emergency observations over routine ones in an automatic manner.<br />
Additional conflict management strategies are needed to dissolve possible resource conflicts between<br />
different emergency monitoring missions, which may coexist within the same period of time. For<br />
handling these situations it will be useful to dynamically assign priority levels to the different<br />
emergency events.<br />
Strategy baselines on the implementation of emergency observation requests into the schedule have<br />
already been given in [RD 3], including scenarios for combining routine and emergency observations.<br />
The straight forward approach to avoid congestion is to allocate only a given percentage of<br />
observation capabilities to routine observations, so that in total, enough resources will be available to<br />
include both routine and emergency observations.<br />
As can be seen from the above considerations, one major issue in the context of the on-demand<br />
scheduling of emergency observations is staffing, unless the whole process could be automated. If not<br />
the case, personnel will need to be available both on PDGS and FOS side to take into account and<br />
schedule unforeseen requests. For solving this issue, an intermediate approach like for Sentinel-3 Fire<br />
Monitoring mission can be adopted, i.e. working during normal working hours only (8/24 5/7) outside<br />
periods of natural disasters, and working round the clock during periods when such events are the<br />
most likely to occur (e.g. April to October for the fire season in Southern Europe).<br />
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4.6.2.4 Cloud Cover Nowcasting and Very Short Range Forecasting<br />
Currently, mission planning for optical satellites uses cloud forecast issued by e.g. Meteo France at 6<br />
hour intervals. Such cloud forecast belongs to the category of “Nowcasting" (NWC) and “Very Short<br />
Range Forecasting" (VSRF), which is defined in a very broad sense as “user-driven services using<br />
appropriate meteorological and related science to provide information on expected conditions up to 12<br />
hours ahead”, covering inter alia air pollution, ocean, and hydrology at these timescales.<br />
The Satellite Application Facility (SAF) for Nowcasting provides operational services to ensure the<br />
optimum use of meteorological satellite data in Nowcasting and Very Short Range Forecasting.<br />
Currently the SAF NWC generates so called Cloud Mask Products at three hour intervals.<br />
In addition to EUMETSAT's SAF NWC, there exist numerous public web sites which provide prediction<br />
maps of weather, temperatures, and cloud cover at various resolutions and various time intervals. A<br />
cloud forecast map is published each hour from current time until 5 days later. However, the reliability<br />
of the information is difficult to be verified.<br />
Regarding the perspectives for the timeframe from 2015 onwards, one can obviously consider the new<br />
generation of meteorological satellites, i.e. Meteosat Third Generation (MTG). The MTG missions<br />
capitalise on the continuation and enhancement of the MSG capabilities.<br />
In addition to the afore-mentioned remote sensing missions, there are also scientific research works<br />
that have demonstrated the ability of so called “advanced advection methods” to provide robust shortterm<br />
top forecasts of cloud motion. The advection technique is based on a cross-correlation algorithm<br />
that computes local motion vectors by tracking identifiable cloud features across pairs of timesequential<br />
satellite images. Satellite data are first processed by cloud detection and cloud property<br />
retrieval algorithms to identify, classify, and stratify cloudy features by altitude. Cloud information is<br />
remapped to a standard map projection and the correlation algorithm applied. If available, NWP winds<br />
are used to reduce processing time and to eliminate obviously incorrect motion vectors.<br />
All these elements concur to the conclusion that significant progress for nowcasting of cloud cover is<br />
being made, which, together with the advent of a next generation of LEO and GEO meteorological<br />
satellites in the time frame 2015 – 2025 should contribute to improve very significantly the accuracy,<br />
timeliness and update frequency of cloud cover forecast products used to optimise the scheduling of<br />
operations for space based optical remote sensing.<br />
For the optimisation of the instrument's schedule, the NWC and VSRF information are needed to be<br />
provided to the payload mission planning. The interface between the mission planning and the meteo<br />
services should be realised such that the cloud cover information can be directly extracted. Based on<br />
this information the cloud cover ratio of the scenes to be acquired shall be computed automatically<br />
within the PDGS. After this the schedule of the instrument will be elaborated accordingly to be then<br />
provided to the FOS, which is in charge of incorporating it into the spacecraft's overall mission plan.<br />
4.6.2.5 User Access<br />
The realisation of the user interfaces for the access to the system shall be realised via a standardised<br />
central user portal. Via this user portal it shall be possible to place general user requests. These<br />
general user requests can be related to catalogue inquiries, ordering of available products from the<br />
archive up to the placement of new acquisition request for the Spacecraft.<br />
The user portal should be based on state-of-the-art Web technology. When realising the user portal, it<br />
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is mandatory to take into account not only existing but also evolving standards, which are commonly<br />
agreed not only in the remote sensing domain but also within a broad community of Geo Information<br />
Service providers and users. Utilising standards guarantees interoperability between different systems<br />
and a unified access to these from a user point of view.<br />
At present, the Open Geospatial Consortium (OGC) is the most relevant entity which defines and<br />
drives the relevant open standards. Within the OGC, the Sensor Web Enablement (SWE) focuses on<br />
sensors and sensor networks. The definition of the SWE standards aims to access and, where<br />
applicable, to control all types of sensors, instruments and imaging devices via the Web.<br />
For this purpose, the SWE comprises seven major elements, i.e.:<br />
(1) Observations & Measurements Schema (O&M) – Standard models and XML Schema for<br />
encoding observations and measurements from a sensor, both archived and real-time.<br />
Current standard: OGC 07-022r1 Observation and Measurements – Part 1 – Observation<br />
Schema<br />
(2) Sensor Model Language (SensorML) – Standard models and XML Schema for describing<br />
sensors systems and processes; provides information needed for discovery of sensors,<br />
location of sensor observations, processing of low-level sensor observations, and listing of<br />
taskable properties.<br />
Current standard: OGC 07-000 Sensor Model Language<br />
(3) Transducer Markup Language (TransducerML or TML) – The conceptual model and XML<br />
Schema for describing transducers and supporting real-time streaming of data to and from<br />
sensor systems.<br />
Current standard: OGC 06-010r6 Transducer Markup Language<br />
(4) Sensor Observations Service (SOS) - Standard web service interface for requesting, filtering,<br />
and retrieving observations and sensor system information. This is the intermediary between<br />
a client and an observation repository or near real-time sensor channel.<br />
Current standard: OGC 06-009r6 Sensor Observation Service<br />
(5) Sensor Planning Service (SPS) – Standard web service interface for requesting user-driven<br />
acquisitions and observations. This is the intermediary between a client and a sensor<br />
collection management environment.<br />
Current standard: OGC 07-018 Sensor Planning Service Application Profile for EO Sensors<br />
(6) Sensor Alert Service (SAS) – Standard web service interface for publishing and subscribing<br />
to alerts from sensors.<br />
Draft standard (not yet released): OGC 06-028r5 Sensor Alert Service<br />
(7) Web Notification Services (WNS) – Standard web service interface for asynchronous<br />
delivery of messages or alerts from SAS and SPS web services and other elements of<br />
service workflows.<br />
With relation to the "Inspire" directive of the EC, which aims at establishing a Geo-data infrastructure<br />
for Europe, the OGC specifications are the key drivers for defining the standardised exchange of Geoinformation<br />
within the European Community.<br />
As can be seen from the above considerations, the realisation of the user portal following the OGC<br />
standards allows to define generic user requests via the Web, where user requests can be among<br />
others:<br />
• Catalogue browsing (primarily the request/mission catalogue)<br />
• Image requests from the archive and re-processing requests of archived data<br />
• Performing new acquisitions (routine and emergency)<br />
• Monitoring of events and submittal of alerts and notifications<br />
Apart from the above-mentioned user services, the user portal is also in charge of handling all related<br />
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management functions, e.g. provisions of access rights, granting of user privileges, etc. By defining<br />
the access rights, a user may have the privilege only to browse the request/mission catalogue, to<br />
subscribe to request to currently scheduled acquisitions products, up to requesting new acquisitions.<br />
In this context, the User Portal is also responsible of providing prioritized access for users requesting<br />
new acquisitions in the frame of emergency applications. This is in particular necessary to guarantee<br />
the high reactivity of the overall system for emergency situations.<br />
The UACC (User Access, Coordination and Control) domain is currently responsible of the interface<br />
with the end-user and for all the services pertaining to the user interfaces including the ordering<br />
function, the mission planning, the master catalogue and the help desk (with the exception of the<br />
dissemination). The UACC domain is composed of a number of heterogeneous elements, some of<br />
them being natively multi-mission (e.g.: MUIS, MMMC, EOLI), some other elements, which were notnatively<br />
multi-mission, have been adapted. The UACC domain is located at ESRIN.<br />
For Geo-Oculus, the evolvement of the current UACC towards multi-mission planning can be<br />
considered as a preliminary baseline for defining the realisation of user accesses and all related<br />
functionalities.<br />
4.6.2.6 Payload Data Reception, Processing and Dissemination<br />
As already stated earlier, the reception of Payload data can make use of:<br />
• A nominal Payload Receiving Station collocated with the main Processing and Archiving<br />
Facility<br />
• A set of smaller user dedicated Payload Receiving Stations, with smaller size dishes.<br />
At time of ingestion, the Payload data will be decrypted first. After that, the Payload data have to be<br />
screened and the metadata attached to this Payload data are extracted in order to feed the catalogue.<br />
A Moving Window Display function may be included to provide the capability to display the raw data<br />
before it is processed.<br />
Within the Payload Data Ground Segment, the processing should proceed up to Level 1B or even<br />
further, depending on the availability of auxiliary and ancillary data required for processing. These<br />
include auxiliary data resulting from calibration (generated inside the PDGS), predicted or restituted<br />
orbit data (generated by the FOS) and other auxiliary / ancillary data for instance for more precise<br />
geo-localisation or ortho-rectification.<br />
Dissemination should occur mainly towards the service segment, the latter being in charge of further<br />
processing and delivery of value added services in compliance with the GMES Service provision<br />
model.<br />
The above-mentioned processing chain for the image product has to be fully automatic in order to<br />
minimise the processing delays in the context of emergency observation missions. The assignment of<br />
different priority levels for each self-contained image product may be suitable to additionally accelerate<br />
the processing and transmission of urgent data.<br />
Regarding the required processing performance, it is suitable to analyse a worst-case scenario. For<br />
oil-slick detection, an image size of approx. 6E+09 Bits can be assumed. A product consists of 15<br />
image parts, giving in total 9E10 Bits to be processed. As a reference for a first approximation we can<br />
assume that the required floating point operations (FLOPS) per bit are comparable to the processing<br />
of a Spot 5 image. For a Spot 5 image, 2000 FLOPS per bit are required, which results in a total of<br />
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1.8E14 FLOPS per product.<br />
The currently available hardware capabilities can be estimated using the example of a SUN SPARC<br />
Enterprise M 9000 Server, which provides a computing power of 1.032 TFLOPS. The M 9000 is one of<br />
the most powerful servers available today. As a result, the M 9000 would be able to compute the<br />
above-mentioned product within 180 seconds (1.8E14 TFLOPS / 1E12 TFLOPS).<br />
The automatic dissemination of the products to the dedicated Service Segment for the generation of<br />
higher level products / value adding can be facilitated by assigning Request-IDs, by which each user<br />
request, the subsequent tasking within the Spacecraft's schedule and the finally delivered product can<br />
be identified and automatically routed to the end-user.<br />
4.6.2.7 Encryption Concept<br />
From a generic point of view, the encryption of the payload data shall ensure the confidentiality of its<br />
content. The reason why this data shall be kept confidential is basically motivated by the following two<br />
headlines:<br />
• Public safety<br />
• Commercial aspects<br />
The consideration of public safety aims at safeguarding the EO data from misuse by illegal groupings<br />
such as criminal associations or even terrorists. This aspect has already been considered by different<br />
legalisation authorities worldwide.<br />
The commercial aspects cover all those issues which are related with the property rights of the image<br />
data. In this respect, the satellite data should be safeguarded from eavesdropping or theft in order to<br />
be able to retail the image products to commercial users.<br />
As a result of the previous considerations, the encryption concept should be simple. A commercial<br />
level encryption concept should be adequate.<br />
One of the main issues will be to ensure that, in case of deployment of users’ terminals for emergency<br />
applications, only the user terminal which submitted the observation request should be able to receive<br />
the image(s) acquired (in addition to the core Payload Data Ground Segment, which would receive all<br />
images in parallel for the sake of their long term archiving).<br />
Among the possible concepts for encryption, the following alternatives can be considered:<br />
• One key per image: this is the most secure but also the most complex alternative<br />
• One key per time slot: in this concept, the keys would be changed at regular time intervals<br />
e.g. each day, each week, each month<br />
• One key per receiving station: each “user” receiving station would then be able to decrypt<br />
only the images which have been encrypted with this station’s key. The core PDGS would<br />
receive all images and then would also need to receive all keys.<br />
One specific issue with respect to users’ receiving terminals will be how to distribute the keys, which<br />
may be an issue in case of deployment of the stations where no permanent network is available. Note<br />
that this issue extends to the distribution to the users’ stations of any other kind of auxiliary data<br />
needed for data reception (pointing data if the satellite is not on a geostationary orbit, time slot for data<br />
reception) and for processing (auxiliary data, orbit data).<br />
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4.6.2.8 Ground Stations for Geo-Oculus<br />
For the preliminary architecture of the Geo-Oculus ground segment it is foreseen to establish two<br />
TT&C stations to provide redundancy for the TT&C functionality.<br />
TT&C Ground Stations<br />
The TT&C Stations are responsible for exchanging telecommands and telemetry with the satellite and<br />
to provide ranging functionality. For the envisaged orbit determination based on the spread spectrum<br />
ranging method, at least two S-Band stations are required; to achieve optimum performance three S-<br />
Band Ground stations should be foreseen. It is recommended to use dedicated S-Band GEO ground<br />
stations for the TT&C functionality providing permanent contact with the satellite.<br />
Using the S-Band for TM/TC and ranging makes the system compatible to the LEOP G/S network.<br />
During LEOP, commissioning and verification phase the LEOP G/S network can provide backup and<br />
failsafe capabilities for the operational system within the initial verification phase. Additionally using a<br />
lower frequency band like the S-Band in combination with a big G/S antenna size increases also the<br />
ranging accuracy of the station.<br />
The location of the primary S-Band TT&C ground station can principally be selected at free choice.<br />
The only constraint is that the S-Band G/S needs a direct communication link to the operations<br />
facilities (being ESOC in Darmstadt as a baseline), which allows a seamless exchange of the TM/TC<br />
data between ESOC and the G/S. At present stage, the Agency's ESTRACK facilities located at<br />
Maspalomas (Spain) is considered as primary ground station. The secondary ground station is<br />
assumed to be located in Redu, Belgium. The monitoring and control of the S-Band ground stations<br />
can be achieved remotely from ESOC by the staff already in place.<br />
TT&C Standards and Interfaces<br />
For compatibility reason to the ESOC/ GSOC G/S network, it is also recommended to follow the<br />
CCSDS standard for the TM/TC data packets and to provide SLE (Space Link Extension) interfaces.<br />
PDT Ground Stations<br />
As a baseline for the Payload Data Ground Segment the main data reception facility shall be equipped<br />
with a dedicated ground station to provide permanent contact with the satellite. As has already been<br />
mentioned before, it is recommended as far as feasible that the receiving station is collocated with the<br />
Payload Data Ground Segment in order to reduce the latency between data reception and processing.<br />
What concerns the usage of a dedicated frequency band for payload data transmission, the ITU allows<br />
for the data reception of GEO earth observation satellites to use the X-, DBS- or Ka-Band. At present<br />
stage, the X-Band has been selected as baseline for the Geo Oculus PDT. The X-band is in the earth<br />
observation domain the most commonly used frequency range for TM data transfer.<br />
However, a GEO based S/C the utilisation of the X-band for payload data transmission may produce<br />
interferences with the LEO systems. A LEO Ground Station located within this foot print can probably<br />
cross this permanent GEO TM link through the tracking process of its LEO spacecraft. As a result of<br />
the interference, it may be disturbed in its link and might loose its track. A small foot print of the GEO<br />
S/C can reduce this risk, on the other hand, this increases the size of the TM transmit antenna on<br />
board the S/C.<br />
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5 Recommendations on further Analysis<br />
5.1 System Analysis<br />
Detailed analysis of the manoeuvre times<br />
A driving aspect of the mission performance is the manoeuvre time of the system to point from on<br />
observation pattern to the next. Depending on the configuration of the AOCS, the limiting factor is the<br />
settling time to achieve the desired level of stability after the active part of the manoeuvre. It has been<br />
identified that especially the characteristics of the solar array are of relevance for the settling time.<br />
It is highly recommended to analyse the manoeuvre times in more detail, especially concerning the<br />
characteristics of large structures as the solar array, in order to consolidate this important aspect of<br />
mission performance at a high level of confidence.<br />
Detailed investigation of microvibration aspects<br />
At the required level of attitude stability, the microvibrations from momentum wheels, solar array drive<br />
and, if applicable, of an antenna pointing mechanism have to be minimized and/or compensated. A<br />
detailed analysis of the microvibrations and the means of reductions is highly recommended for the<br />
further study phases.<br />
5.2 Mission Objectives and Data Processing<br />
At least two major issues remain at this stage of the GEO study. First the need to strongly consolidate<br />
the user’s requirement, second the temporal coverage specificity of the GEO (compute the optimal<br />
revisit frequency to get one clear image per days, based on the Eumetsat cloud products archive and<br />
taking into account ocean colour geometrical limitations).<br />
Additional proposed tasks:<br />
• Justification of the GEO concept for OC. A less demanding requirement on the temporal<br />
coverage is to be able to detect the daily oceanic structures (like Chlorophyll gradient).<br />
Contrary to purely numerical techniques of “optimal interpolation” trying to fill the gap of LEO,<br />
the GEO concept could directly supply the physical data in a progressive way among the<br />
day. It is proposed to analyse the progressive detection of Chlorophyll structure (i.e.<br />
progressive improvement of the gradient computation with increasing clear zones along the<br />
day), as a function of the number of acquisitions, and to derive the minimal revisit<br />
requirement which suits current operational services in structure detection.<br />
• Air mass issue and atmospheric correction. There is a big need to have a radiative<br />
transfer modeling tool in spherical coordinates, in order to access the realistic air mass<br />
requirements. To our knowledge such a code is not available to the Ocean Colour<br />
community in Europe and could be developed.<br />
• Coverage analysis The user requirement on temporal coverage refers to two main aspects:<br />
− need to have several images per day in order to built one daily cloud-free synthesis (e.g.<br />
phytoplancton map)<br />
− need to have several clear images per day in order to follow rapid events (e.g. tides, NRT<br />
water quality monitoring).<br />
The analysis conducted so far on "availability coverage" used a very high revisit time (15<br />
min) and is thus not exactly scaled to the requirements and potential of Geo-Oculus (agility<br />
for pointing on cloud-free region). An acquisition scenario that would optimise the cloud-free<br />
region, taking into account the realistic duration of acquisition, pointing and stabilization<br />
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<strong>Report</strong><br />
based on EUMETSAT images could be elaborated. The analysis should also include other<br />
constraints: minimization of air mass fraction and glint avoidance.<br />
• Data processing. Ocean colour measurements from GeoOculus are characterised on one<br />
side by high frequent, spatial high resolution, which are superior to current LEO orbit data,<br />
on the other side by radiometrically less favourable conditions, including probably lower<br />
SNR, possibly spectral and spatial misregistration between bands and a very large viewing<br />
angle at higher latitudes.<br />
For detailed assessment and identification of required developments of data processing<br />
methods, the relevant processing chain should be analysed for sensibility to the<br />
characteristic of geostationary observation.<br />
Beside the traditional approach it should be studied to develop products which require<br />
simplified processing. The processing would either use directly TOA radiance without explicit<br />
atmospheric correction, or perform a simplified AC.<br />
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<strong>Report</strong><br />
6 Conclusion<br />
As a conclusion the principle feasibility of the proposed Geo-Oculus mission is confirmed. The results<br />
show that the challenges in terms of instrument design and LoS performance can be met.<br />
It is however clear that a lot of assumptions concerning the selected applications, the derived product<br />
requirements and the mission scenarios have to be re-iterated before the coming study phases. As a<br />
consequence the modified requirements will then lead to a re-iteration of the system design and<br />
performance.<br />
Nevertheless, the results of the study are very promising to allow a clear recommendation for a<br />
continuation of the activities also in view of the potential of Geo-Oculus to become an operational<br />
mission e.g. as a part of the GMES program.<br />
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Annex A Abbreviations<br />
AD Applicable Document<br />
ADCS Attitude Determination and<br />
Control System<br />
AEF Apogee Engine Firing<br />
AGA 8PSK Eight Phase Shift<br />
Keying<br />
ATSR Along Track Scanning<br />
Radiometer<br />
AATSR Advanced ATSR<br />
AIDCO EuropeAid Co-operation<br />
Office, EU<br />
AIT Assembly Integration & Test<br />
AIV Assembly Integration &<br />
Verification<br />
ALADIN Atmospheric Laser Doppler<br />
Instrument<br />
AMAP Arctic Monitoring and<br />
Assessment Programme<br />
AME Absolute Measurement Error<br />
AOCS Attitude and Orbit Control<br />
System<br />
APE Absolute Pointing Error<br />
APS Antenna Pointing System<br />
APSK Advanced Phase Shift Keying<br />
ASAR Advanced Synthetic Aperture<br />
Radar<br />
ASD Astrium D (Deutschland)<br />
ASF Astrium F (France)<br />
ASIC Application Specific Integrated<br />
Circuit<br />
ASU Astrium UK (United Kingdom)<br />
AVHRR Advanced Very High<br />
Resolution Radiometer<br />
BAe British Aerospace<br />
BCR Battery Charge Regulator<br />
BDR Battery Discharge Regulator<br />
BOL Begin of Life<br />
BRDF Bi-directional Reflectance<br />
BTDF<br />
Distribution Function<br />
Bi-directional Transmission<br />
Distribution Function<br />
CCD Charged Coupled Device<br />
CDH Command and Data Handling<br />
CFRP Carbon Fibre Reinforced<br />
Plastic<br />
CMG Control Momentum Giros<br />
CMOS Complementary Metal Oxide<br />
Semiconductor<br />
CNES Centre Nationale d´Etude<br />
Spatiale<br />
COMS Communications Operational<br />
Meteorological Satellite<br />
CORINE Coordinated Information on the<br />
European Environment<br />
COTS Commercial of the shelf<br />
CPS Chemical Propuslions System<br />
CSA Canadian Space Agency<br />
CSS Coarse Sun Sensor<br />
DC Direct Current<br />
DET Direct Energy Transfer<br />
DG European Union Directorate<br />
General<br />
DGA La délégation générale pur<br />
l’armement<br />
DLR Deutsches Zentrum für Luft<br />
und Raumfahrt<br />
DMC Disaster Monitoring<br />
Constellation<br />
DSNU Dark Signal Non-Uniformity<br />
DUE Data User Element<br />
DUP Data User Project<br />
EADS European Aeronautic Defence<br />
and Space Company<br />
EC European Community<br />
ECSS European Cooperation for<br />
Space Standardization<br />
EEA European Environmental<br />
Agency<br />
EIONET European Environment<br />
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ENS<br />
Information and Observation<br />
Network, EEA<br />
Earth Observation, Navigation<br />
and Science<br />
ENTR Enterprise<br />
ENV Environment<br />
EO Earth Observation<br />
EOGeo Earth Observation from GEO<br />
EOL End of Life<br />
EOPP Earth Observation Preparatory<br />
Programme<br />
EOS Earth Observation System<br />
EPS Electrical Propulsion System<br />
or Electrical Power System<br />
ERS Earth Remote Sensing<br />
Satellite<br />
<strong>ESA</strong> European Space Agency<br />
ESOC European Space Operation<br />
Centre<br />
ESRIN European Space Research<br />
Institute<br />
EU European Union<br />
EUROSTAT Statistical Office of the<br />
European Union<br />
EWSK East West Station Keeping<br />
fAPAR Fraction of Absorbed<br />
Photosynthetically Active<br />
Radiation<br />
fCover Fraction of Land Cover type<br />
FDIR Failure Detection Identification<br />
and Recovery<br />
FDS Flight Dynamics<br />
System/Software<br />
FM Flight Model<br />
FOG Fibre Optic Gyro<br />
FOS Flight Operations Segment<br />
FOV Filed of View<br />
FPGA Free Programmable Gate<br />
Array<br />
FTS Fourier Transform<br />
Spectrometer<br />
GAC GMES Advisory Counsel<br />
GCP Ground Control Points<br />
GEMS Global Earth System Modelling<br />
Using Space and in-situ data<br />
(FP6 IP)<br />
GEO Geostationary Earth Orbit<br />
GFRP Glass Fiber Reinforced<br />
Polymer<br />
GIFTS Geostationary Infrared Fourier<br />
Transform Spectrometer<br />
GIS Geo Information System<br />
GMES Global Monitoring Environment<br />
and Security<br />
GMFS Global Food Security Service<br />
GNC Guidance Navigation and<br />
Control<br />
GNSS Global Navigation Satellite<br />
System<br />
GOFC-GOLD Global Observation for Forest<br />
and Land Cover Dynamics<br />
GOME Global Ozone Monitoring<br />
Experiment<br />
GOSIS GMES Organisation and<br />
Systems Integration Scenarios<br />
GPS Global Positioning System<br />
GS Ground Sampling<br />
GSD Ground Sampling Distance<br />
GSE Ground Support Equipment<br />
GSE GMES Service Element (<strong>ESA</strong><br />
Projekts)<br />
GSO Geosynchronous Orbit<br />
GTO GEO Tranfer Orbit<br />
HEMP High Efficient Electromagnetic<br />
Plasma<br />
HEMPT High Efficient Electromagnetic<br />
Plasma Thruster<br />
HRS High Resolution Instrument<br />
HW Hard Ware<br />
ICEMON Sea ice monitoring in the polar<br />
regions (GSE project)<br />
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ICU Instrument Control Unit<br />
IEEE Institute of Electrical and<br />
Electronics Engineers<br />
IG Implementation Group<br />
IGSO Inclined Geosynchronous Orbit<br />
IMU Inertial Measurement Unit<br />
INR Image Navigation and<br />
Registration<br />
IP Integrated Project (European<br />
Commission)<br />
IR Infrared<br />
ITAR International Traffic in Arms<br />
Regulations<br />
ITD Infoterra Deutschland<br />
ITT Invitation to Tender<br />
ITU International<br />
Telecommunications Union<br />
LAI Leaf Area Index<br />
LC Land Cover<br />
LEO Low Earth Orbit<br />
LEOP Lauch and Early Operations<br />
LHCP Left Handed Circulary<br />
Polarized<br />
LMCS Land Monitoring Core Service<br />
LNA Low Noise Amplifier<br />
LOS Line of Sight<br />
LR Low resolution<br />
LST Local Satellite Time<br />
LUSI Land Use and Spatial<br />
Information (European Topic<br />
Centre)<br />
LWIR Long Wave Infrared<br />
MARS Monitoring of Agriculture by<br />
Remote Sensing<br />
MCT Mercury Cadmium Telluride<br />
MIR Medium Infrared<br />
MLI Multi Layer Insulation<br />
MMU Minimum mapping unit<br />
MOU Memorandum of<br />
MPPT<br />
Understanding<br />
Maximum Power Point Tracker<br />
MR Medium resolution<br />
MS Member State of the European<br />
Union<br />
MSG Meteosat Second Generation<br />
MTF Transfer Function<br />
MTG Meteosat Third Generation<br />
MTR Mid Term Review<br />
MWIR Medium Wave Infrared<br />
NDVI Normalised Differential<br />
Vegetation Index<br />
NEDL Noise Equivalent Delta<br />
NIR Near Infrared<br />
NRC National Reference Centre<br />
NRT Near Real Time<br />
NSSK North South Station Keeping<br />
OBC Onboard Computer<br />
OC Ocean Colour<br />
OCM Orbit Control Manoeuvre<br />
OD Orbit Determination<br />
OQPSK Offset Quaternary Phase Shift<br />
Keying<br />
OSPAR The 1992 OSPAR Convention;<br />
the current instrument guiding<br />
international cooperation on<br />
the protection of the marine<br />
environment of the North-East<br />
Atlantic<br />
PB Programme Board<br />
PCM Pulse Code Mode<br />
PDGS Payload Data Ground<br />
Segment<br />
PDH Payload Data Handling<br />
PDHT Payload Data Handling and<br />
Transmission<br />
PDT Payload Data Transmission<br />
PF Platform<br />
PFD Power Flux Density<br />
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PM Progress Meeting<br />
PPS Pulse per Second<br />
PRM Product Requirements Meeting<br />
PRNU Pixel Response Non<br />
Uniformity<br />
PSK Phase Shift Keying<br />
QE Quantum Efficiency<br />
RAMSAR Convention on Wetlands,<br />
signed in Ramsar, Iran, in<br />
1971<br />
RAAN Right Ascension of Ascending<br />
Node<br />
RD Reference Document<br />
REGIO Directorate General Regional<br />
Development (Europ. Com.)<br />
RF Radio Frequency<br />
RFI Request for Information<br />
RFP Request for Proposal<br />
RHCP Right Handed Circulary<br />
Polarized<br />
RME Relative Measurement Error<br />
ROM Rough Order of Magnitude<br />
RPE Relative Pointing Error<br />
RS Reed Solomon<br />
RW Reaction Wheel<br />
SA Solar Array<br />
SADM Solar Array Drive Mechanism<br />
SAGE Service for the Provision of<br />
Advanced Geo-Information on<br />
Environmental Pressure and<br />
State (<strong>ESA</strong> GSE project)<br />
SAR Synthetic Aperture Radar<br />
SCU Spacecraft Computer Unit<br />
SK Station Keeping<br />
SNR Signal to Noise Ration<br />
SOLAS Surface Ocean Lower<br />
Atmosphere Study<br />
SOW Statement of Work<br />
SSPA Solid State Power Amplifier<br />
SSH Sea Surface Height<br />
SSP Sub Satellite Point<br />
SST Sea Surface Temperature<br />
STR Star Tracker<br />
SWIR Short Wave Infrared<br />
TBC to be Confirmed<br />
TBD to be determined<br />
TC Telecomand<br />
TESI TerraSAR Exploitation and<br />
Service Infrastructure<br />
TIR Thermal Infrared<br />
TM Telemetry<br />
TMA Three Mirror Anastigmat<br />
TMTC Telemetry and Telecomand<br />
TN Technical Note<br />
TOA Top of Atmosphere<br />
TOC Table of Contents<br />
TRL Technology Readiness Level<br />
TV Thermal Vacuum<br />
TWTAs Travelling Wave Tube<br />
Amplifiers<br />
UK United Kingdom<br />
UN United Nations<br />
UNCLOS United Nations Convention on<br />
the Law of the Sea<br />
UNFCCC United Nations Framework<br />
Convention on Climate<br />
Change<br />
UV Ultra Violet<br />
VGT Vegetation<br />
VHR Very high resolution<br />
VIS Visible<br />
VLWIR Very long wavelength Infrared<br />
VNIR Visible Near Infrared<br />
WBS Work Breakdown Structure<br />
WFD Water Framework Directive<br />
WFE Wave Front Error<br />
WP Working Package<br />
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