preliminary!study!of!a!space!mission!...

C o n a e , C o r d o b a -‐ 1 7 O c t o b e r 2 0 1 3

Cordoba -‐ Master AERTE

Preliminary Study of a Space Mission using the Concurrent Engineering

Facility Claudia Facchinetti

PRELIMINARY STUDY OF A SPACE MISSION USING THE CONCURRENT ENGINEERING FACILITY

2

Summary

1. INTRODUCTION .......................................................................................................................................... 3 2. PHASE OF A SPACE MISSION ........................................................................................................... 3 3. CONCURRENT ENGINEERING ........................................................................................................... 5 3.1 PRINCIPAL COMPANIES AND ORGANIZATIONS THAT USE CE ...................................................................................... 6 3.2 THE APPROACH OF CONCURRENT ENGINEERING (CE) ............................................................................... 7

4. CEF ASI STUDY CASE ..................................................................................................................................... 14 4.1 MISSION STUDY ................................................................................................................................................................. 15 4.2 MISSION OBJECTIVES ....................................................................................................................................................... 16 4.3 REQUIREMENTS ................................................................................................................................................................ 16 4.3.1 Mission Requirement ............................................................................................................................................... 16 4.3.2 Mission Operative Requirements .................................................................................................................... 17 4.3.3 Mission Constraints .................................................................................................................................................. 17 4.3.4 System Requirements .............................................................................................................................................. 17 4.3.5 Space Segment Requirements ............................................................................................................................. 18

4.4 SUBSYSTEM STUDY ........................................................................................................................................................... 18 4.4.1 System ............................................................................................................................................................................ 18 4.4.2 Mission ........................................................................................................................................................................... 19 4.4.3 Payload .......................................................................................................................................................................... 20 4.4.4 Thermal ......................................................................................................................................................................... 23 4.4.5 Mechanical & Structure ......................................................................................................................................... 23 4.4.6 Configuration ............................................................................................................................................................. 24 4.4.7 AOCS ............................................................................................................................................................................... 25 4.4.8 GrounSegment ............................................................................................................................................................ 26 4.4.9 Thecnical Risk ............................................................................................................................................................ 26 4.4.10 Management ............................................................................................................................................................ 27 4.4.11 Cost ............................................................................................................................................................................... 28

4.5 CONCLUSION SEO ............................................................................................................................................................ 28 5. LESSON LEARNED AN RECOMMENDATIONS ......................................................................................... 29 6. CONCLUSION .................................................................................................................................................... 30 ACRONYM .............................................................................................................................................................. 30 BIBLIOGRAFY .................................................................................................................................................... 31 ANNEX A ..................................................................................................................................................................... 32

CONCURRENT ENGINEERING

3

1. Introduction This document describes the different steps of a preliminary study of an Earth’s Observation space mission. In particular, when the mass budget is allocated, the relevant parameters to make a choice on the orbit will be defined, a set of useful characteristics of the payload will be identified, and, pertinent considerations for this particular project (such as revisit time, resolution, etc…) will be eventually discussed. The aim of this study is to provide the Costumer with the instruments necessary to establish an assessment of the cost, a schedule of work (and risk) and a total time necessary to complete the project (from start-up to the launch of the satellite). This goal will be achieved using the Concurrent Engineering Facility (CEF) at the Italian Space Agency (ASI) premises. The methodology and procedures employedare hereafter described. In general, a team of 12 expert working in “concurrent engineering” mode can be able to elaborate a consistent solution for a study case, in line with the expectations and general constraints.

2. Phase of a Space Mission The life cycle of space projects (in ECSS standard [7]) is typically divided into 7 phases, as follows:

Phase 0 - Mission analysis/needs identification Phase A - Feasibility Phase B - Preliminary Definition Phase C - Detailed Definition Phase D - Qualification and Production Phase E - Utilisation Phase F - Disposal

Project phases are closely linked to activities on system and product level. Depending on the specific circumstances of a project and the acceptance of involved risk, activities can overlap project phases. ü Phases 0, A, and B are focused mainly on

§ the elaboration of system functional and technical requirements and identification of system concepts to comply with the mission statement, taking into account the technical and programmatic constraints identified by the project initiator and top level customer.

§ the identification of all activities and resources to be used to develop the space and ground segments of the project,

§ the initial assessments of technical and programmatic risk, § initiation of pre‐ development activities.

ü Phases C and D comprise all activities to be performed in order to develop and qualify the space and ground segments and their products.

ü Phase E comprises all activities to be performed in order to launch, commission, utilize, and maintain the orbital elements of the space segment and utilize and maintain the associated ground segment.

ü Phase F comprises all activities to be performed in order to safely dispose all products launched into space as well as ground segment.


4

This set of definitions may change within national Space Agencies (see Fig. 1). Most missions will require some new technology to be developed. This is natural, as new scientific goals are pursued and the means required to meet them must advance on previous technologies; but new requirements can pose tough challenges for engineers and spacecraft designers. Sometimes specific circumstances and opportunities allow this potentially lengthy process to be shortened.

Fig. 1 – Definition of life cycle of a projects [8]

The time required to progress from the initial concept to deorbiting or death of the space asset appears to be independent of the sponsor. Large complex mission typically require 10-15 years to develop and operate from 5 to 15 years, whereas small, simple missions require as few 12 to 18 month to develop and operate for 6 month to several years. A Concurrent study can help reduce time and risk. The scope of CE activities is usually limited to (pre-) Phase A design and to Phase B reviews of the space project lifecycle (see Fig. 2). In ESA, the majority of these studies are conducted internally at ESA’s Concurrent Design Facility (CDF). The ASI's approach is to conduct internally studies (using CEF) mainly of phases -0 and -A and, more recently, study of projects that require costs’ estimation.


5

Fig. 2 – The project life-cycle, ASI proposes only studies of phase 0 and A even if the CEF can

handle studies of phase B (as shown by ESA/CDF in red parts, [6][13] )

3. Concurrent Engineering The definition of Concurrent Engineering that we have adopted for the Concurrent Design Facility is: "Concurrent Engineering (CE) is a systematic approach to integrated product development that emphasises the response to customer expectations. It embodies team values of co-operation, trust and sharing in such a manner that decisionmaking is by consensus, involving all perspectives in parallel, from the beginning of the product life-cycle." Essentially, CE Concurrent Engineering is a systematic approach to integrated product development focused on the team values of cooperation, trust and sharing, that focuses on the response to customer expectations. The concurrent engineering approach is based on five key elements: • a process • a multidisciplinary team • an integrated design model • a facility • a software infrastructure

The spacecraft design is based on mathematical models, which make use of custom software and linked spreadsheets. By this means, a consistent set of design parameters can be defined and exchanged throughout the study, and any changes which may have an impact on other disciplines can immediately be identified and collectively assessed. In this way, a number of design iterations can be performed, and different design options can easily be analysed and compared. It is primarily used to assess the technical and financial feasibility of future space missions and new spacecraft concepts (e.g. internal pre-phase A or Level-0 assessment studies) providing:

• new mission concept assessment

ASI

CEF


6

• space system trade-offs and options evaluation • new technology validation at system/mission level

as well as: • payload instrument conceptual design • reviews of industrial phase A studies • scientific requirements definition and consolidation • anomaly investigation • education and training

The number of companies and organizations developing facilities oriented towards the application of the principles of Concurrent Engineering is steadily growing worldwide

To overcome the communication gaps between the “designer” (who produces design information) and the “user” (who utilizes the design information)

A methodology approach based on the concept of a “driving parameter” is implemented to produce one or more scenarios. Each scenario represents a trade-off between a mission, a system which achieve the mission and programmatic perspectives.

3.1 Principal Companies and organizations that use CE In the Italian Space Agency (ASI) the Concurrent Engineering Facility is present with the name CEF and it’s based on the Integrated Design Model (IDM) distributed by ESTEC. Similar tools are already operating in the mayor Space Agency, in the principal European Spatial Industries and in many institutes, universities and Academies all over Europe (see fig Fig 3). As an example, a list of some institute, industries, and Agency that use a concurrent Engineering methodology is the following: § CEF - ASI : operating since 2009, the main studies performed in ASI involved Phase

0 and A. § CIC – CNES(Center National d’études Espatiales): operating since 2005. § CEF - DLR(German Aerospace Center): operating since 2008, already 20 condensed

studies have been peformed at the DLR CEF [4]. § CDF – ESA(Europe Space Agency): At ESA establishment European Space Research

and Technology Centre (ESTEC), Noordwijk, the Netherlands, the application of concurrent engineering principles is undertaken at the site known as Concurrent Design Facility (CDF). The CDF was established at ESTEC in November 1998 within the framework of the General Studies Programme.and until today, 160 studied have been performed[3].

§ Team X - in the NASA JPL, operating since 1994. § The DARPA Initiative in Concurrent Engineering (DICE), is a program launched in

1988 by the US Dept. of Defense's Defense Advanced Research Projects Agency (DARPA) to encourage the practice of concurrent engineering in the US military and industrial base.

§ CDF (Australia –Australia) - Victorian Space Science Education Centre (VSSEC) § SDO (Satellite Design Office) - EADS Astrium - [2] operating from 1998, use the

MUSSAT tool for Satellite Design § CDP (Concurrent Design Platform), commercialised by JAQAR Concurrent Design

Services (J-CDS) - Astrium-ST, operating since 2011. [9]


7

§ ISDEC (Integrated System Design Center) - TAS-I (Roma), operating from 2003 [11].

§ VCDF(Virtual Concurrent Design Facility) –TAS-I & F (Torino, Cannes, Toulose) operating from 2003.

§ CEF - University Roma1 (rome –Italia).

Fig 3 - The Agencies, Universities, Industries that use a Concurrent Engineering (ESA 2013 -

[13])

3.2 THE APPROACH OF CONCURRENT ENGINEERING (CE) The Concurrent Engineering (CE) approach generally consists of the following 5 key elements [3] which are: § The CE process: conceptual an practical § The interdisciplinary team of experts § The Integrated Design Model (IDM) § The subsystem definitionThe software and hardware infrastructure § The CE facility

The CE Process: conceptual and practical The conceptual model of the design process is shown in Fig. 4:it is highlighted that a space system has many interdependencies between components. This implies that the definition and evolution of each component has an impact on other components and that any change will propagate through the system. Early assessment of the impact of changes is essential to ensure that the design process converges on an optimised solution (see Fig. 5). The CE approach is intended to improve the means of achieving this early review and verification, step by step[4].


8

Fig. 4 – Conceptual model of the mission and spacecraft design process (ESA [13]).

The process starts with a phase in which the Customer and some exponents of the engineering team meet to refine and formalise the mission requirements, to define the constraints, to identify design drivers, and to estimate the resources needed to achieve the study objectives. The whole process is conducted in a number of sessions (minimum 8 – with 2-3 session for week) in which all specialists must participate. This reduces the risk of incorrect or conflicting design assumptions, because each major decision is debated and agreed collectively [4]. A first step in the modelling process is to establish the model suited to the mission scenario before it can be parameterised to perform the iterative design process. The design process starts with the kick-off; then , the subsequent sessions are based on the iterative process that addresses all aspects of the system design in a quick and complete fashion. Each specialist presents the solutions for his domain, proposes options and discusses the implications for the other domains. After each iteration the processed data are stored in a shared database and consolidate. Repeating these sequences for several times will lead to final convergence towards a optimal solution, according to the Spiral Model in Fig. 5. In general, in every study it’s possible identify a default path, but at any stage it must be possible to take advantage of alternative path (different persuasive decision in the process).


9

Outside of the principal session the specialist can meet and discuss the approach in private session. The participation to the study of the customer, guide the process in the expected direction and can be carried out on the open issues.

Fig. 5 – The spiral model converging design in the CEF

The last step of any study is the final study report. The specialists (using Microsoft Word) prepare the report of their section as a sub-document that will be incorporated in the master document. The standard document template has been provide by ASI (each Agency has its own). The practical process follow a formal technical procedure. At the start of a every study case, the creation of some directories is required: § A Project model directory: contains all excel sheets used during the study case; § An Admin directory: contains the documents deemed relevant to all the experts

involved; § A Meeting Directory: for each works session an identified directory (with date) is

generated; this contains the documents generated during the specific session (pdf, doc, xls, presentations, team leader questions and system assigned work)

§ A Notes directory: it contains the subsystem works directory, where each expert insert data, consulted document (article, report, ppt, etc.), alternative sw adopted; in general, the material used/consulted during a study case.

§ A Presentation directory: contains the final presentation of the experts. § A Report directory: contains the report (per discipline) generated by each expert,

and the assembled final report. § A miscellaneous directory: contain other not relevant documents, but managed

during the study case. The ideal duration required for each study is 8 sessions, 2 per week, with a duration of at least 3-4 weeks to work.

The interdisciplinary Team The engineering team is composed by: one team leader, one system engineer and


10

selected specialists. The team leader moderates all the process: directs the work by facilitating and regulating the communication between specialists, also decides the strategy of the engineering sessions. The specialists must work together in one room, using sophisticated tools, and perform the design work in real-time. During this time the conflicts are amplified and discussed [4]. The team must to be multi-disciplinary and highly motivated. The essential elements that contribute to personal motivation are for example the collective approach, the co-operative environment, the intense and focused effort and a clear and short term goal. The difficulties increase because team members have to accept, e.g., the use of a new technology, perform design work and give a answer in real time. It’s required the availability of 2 persons per role during the build-up phase. In addition to the technical and scientific team, also the customer’sspecialists participate in the CE-process to discuss and correct, in real time, the evolution of the design, in case it’s not in line with their expectations. Furthermore, their participation in the study permits a direct interaction with the team and a continuous monitoring of the design evolution.

Fig. 6 – IDM model (ESA 2007 [6])


11

The Integrated Design Model The iteration process of the study development requires an engineering model which is able to connect and update the different subsystems. This is realized by an Integrated Design Model (IDM) containing an interdisciplinary data structure (see Fig. 6) and based on the application software MS Excel [4]. Each IDM consist of an input/output/calculation/results areas. The Input/Output areas are used to exchange parameters with the rest of the systems (i.e. internal/external tools and models). The Calculation area contains equations and specifications data for different technologies in order to perform the actual modeling process. The Results area contain a summary of the numeric results during the design process and as part of the report at the end of the study [5]. The experts insert the calculated values of their own discipline (e.g. mass, dimensions, power consumption, temperature, etc) into the IDM. The different output data are shared by a central data exchange workbook and then distributed to the corresponding input sheets of other workbooks. It is possible to link values for internal calculation and external distribution automatically or manually. Furthermore, users are free to change, delete or add parameters for adapting their IDM working environment relating to the study specific mission contents.

The Subsystem definition In each session it is possible to configure and use a different subsystem. In the ASI facility about 15 disciplines are foreseen; the list is proposed in Table 1. Within the facility, for each discipline a position is created and assigned to an expert in that particular technical domain. Each position is equipped with the necessary tools for design modeling, calculations and data exchange. The choice of disciplines involved depends on the level of detail required and on the specialisation of the available expertise.

Table 1 – Satellite Subsystem and technical discipline in the CEF

Postazioni WB's Main Function [9]

ASICEF-11 AOCS Measurement and control of the satellite orientation and orbit.

ASICEF-12 Communication Communication with ground stations/other satellites, tracking of the satellite.

ASICEF-14 Configuration Definition & control of the overall mechanical configuration of the satellite i.e. equipment accommodation consistent with mechanical/thermal environment.

ASICEF-5 Cost & Risk Definition of the life cycle cost of the mission and analysis of risk

ASICEF-2 System Plan and managment of the data exchange and in general of the project.

ASICEF-9 Command & Data_Handling

Processing, storage and formatting of mainly payload data. Distribution of commands and collection of housekeeping


12

data for payload and subsystems.

ASICEF-13 Planning & GSE (Operation)

Plan and Manage the ground station. Planning and ground support equipment (GSE), planning of the design & development & assembly, verification and integration (AIV) efforts as well as planning of the required ground support equipment.

ASICEF-2 Instruments/Payload

Collection and conveyance of the primary mission data

ASICEF-15 Mission & Simulation

Mission analysis, including orbit analysis, functional analysis, satellite operations and on-board software definition.

ASICEF-10 Power Generation of the required power for the satellite, including storage, regulation and distribution functions.

ASICEF-6 Propulsion & Pyrotechnics

Provide thrust for orbit maintenance and orbit control.

ASICEF-4 Programmatics

ASICEF-7 Structures & Mechanisms

Carriage, support and alignment of the satellite equipment. Interface structure with the launcher.

ASICEF-8 Thermal Temperature control

ASICEF-1 Team Leader

Overall responsibility for organising the study, for coordination with the customer, for organising the study in cooperation with the customer and for planning and conducting group activities

The specialists for the subsystems System, Cost & Risk, Programming and The Team Leader are normally chosen inside the agency.

The Software/Hardware Infrastructure

In general, each Space Agency implements the Facility with external software (SW) and/or Hardware (HW) or other infrastructures. In the ASI CEF all the subsystems have been provided with an integrated peculiar tool. Furthermore, there are some infrastructures common to all disciplines that permit: § to work remotely with other facilities (inside/outside ASI) § to exchange information easily between the normal office working environment

and the Facility environment. In general, the configuration adopted is based on the the products already available in the Agency’s infrastructure or within the technical domain of the participating engineers. This case is shown in Table 2 (left). In other cases, it is chosen to incorporate specific tools for modeling and/or complex calculation for which additional licences (where required) are purchased ( the major software products to be employed are Matlab, CATIA, STK, ESARAD, etc).


13

Table 2 – left: general tools chosen as basic infrastructure items used by all team members and right: domain-specific tools used by the domain experts

Function Tool used Function Tools Used ASI

Tools Used ESA

Documentation storage & archive

Note database Configuration & Structural design

CATIA & Matlab

CATIA

Electronic communication within the team

Agency’s mail Mission analysis & orbit control

STK IMAT, Matrix X

Storage area for all data files

ftp directory server

Mission Simulation & Visualisation

STK EUROSIM

System modeling

Excel spreadsheets

Cost Modeling PRICE + DBTE+ ECOS

ECOS cost /technical database & small satellite cost

Project Documentation

MS word Risk MATED

(?)

Remote audio/visual communication

Video conferencing & net meeting

3.3 Thermal ESATAN

ESATAN/ESARAD

The CE facility

The ideal case would be to have access to one main conferencing room and some smaller rooms for splinter meetings. Another room would be available for the costumer team to follow the study sessions. An adequate infrastructure room can facilitate the study. In the main conferencing room, every team member needs to view the central video screens and the team leader. In addition, a opportune distribution of the inner structure is required, so that every team member is able to do a face-to-face discussion with all other team members. In the following figures the typical CDF/ESA and ASI facility are showed.


14

Fig. 7 - CDF facility of ESA

Fig. 8 –Configuration of the CEF at previous ASI premises

4. CEF ASI Study Case This study has been conducted in the ESTEC CDF by an ASI Team in the June 2010. The team leader was Mr. M. Bandecchi, ESTEC head of CE section and the team was


15

supported by a group of ESTEC experts. The team held 8 sessions in the CDF, supported by 8 half-day off-line preparation work, splinter meetings and discussions. The main objectives of the CEF Team were the following: § Give a set of coherent requirements and mission constraints. § Generate a certain number of trade off to offer at ASI management different choice

at high level. In the CEF study, the following steps are fundamental: 1) Identify the Mission objective. 2) Identify the Requirements. These can change during the study. 3) Define the mass budget 1.

4.1 Mission study

Fig. 9 - SEO image of the interest area.

The subject of the study is a Space system for Earth Observation (SEO). The study was aimed to develop: § The Mission Concept, with a high level description of the SEO mission. § SEO project feasibility assessment and preliminary system and subsystems

1 S/C dry mass depends on payload mass and spacecraft bus dry mass. Early estimates for spacecraft bus dry mass and the allocation of this mass over the bus subsystems can again be made based on mass data for prior spacecraft designs and known value of payload mass. Using such data a relationship can be established between payload mass (PLM) versus vehicle dry mass (VDM) and in a next step the mass available for the spacecraft bus can be distributed amongst the various subsystems. VDM = 3.3 *PLM (range is from 2-6 times PLM) [8],

§ All Earth -orbiting spacecraft: VDM = 4.8 PLM (based on 48 data points) [20] § GEO Comsats: VDM = 3.6 PLM § Planetary spacecraft: VDM = 7.5 PLM

a reference table for the Dry mass distribution in %: small (<500 kg) satellites/probes [17], GEO telecommunications satellites [19] ,[18], subsystem mass allocation %of VDM [8].


16

configuration § mission programmatic elements (summary bar chart and AIV philosophy), as well

as application and technological aspects § risk evaluation § cost estimation, detailed per products and activities (Management, Development,

AIT, etc.).

4.2 Mission Objectives The SEO mission shall consent the continuous monitoring of the Amazon region for at least 5 years

PRIMARY OP1. SEO is a space borne SAR mission to provide data of the vegetated terrain of

the Amazon Region. Due to the scarcity of up to-date information, which is fundamental for planning and strategic decision-making about environment assessment and monitoring and management of natural resources in the Brazilian Amazon, the proposed mission should be strongly oriented to a quasi-operational (“application-oriented”) system. The application is dedicated to thematic mapping (vegetation and deforestation, hydrology such as flooding, oil spills etc).

OP2. The monitoring activities will be used also to support civil protection intervention in case of flooding near to any urban city of the region.

OP3. SEO will allow ASI and national industries the maximum utilisation of the already existing qualified technologies and know-how.

SECONDARY OS1. All residual SEO resources shall be exploited as much as possible to observe

other forestal zones.

4.3 Requirements The second step in the CE work is to define the requirements using ECSS standards [14]. A list of principal requirements identified in this study case (come after an example in SEO case – invented data) follows:

1. Mission requirements; 2. Mission Operative requirements; 3. Mission constrain; 4. System requirements; 5. Space segment requirements;

4.3.1 Mission Requirement MR-F-1 SEO shall be a system of one or more satellites carrying on board an L

band radar payload MR-F-2 The SEO payload shall be able to do Stripmap observations.

SAR Products shall have the following characteristics: • Stripmap

− Geometric Resolution: 10x10 m; − Incidence Angle: 20-50°; − ………….

Geolocalisation accuracy: 30 m; • ScanSAR is an option

MR-F-3 The SEO system shall provide 3-look Stripmap products.


17

MR-F-4 The SEO products are foreseen at different levels: • Level 0 Product (raw); • Level 1 Products:

- 1A: Single look, Slant range; - 1B: Multi-looked, Ground range; - 1C: Geocoded Ellipsoid Corrected (on the WGS84 ellipsoid);

• Higher Level Products: - Mosaicked (at level 1C); - Cropped (at level 1C).

MR-F-5 Each SEO satellite shall be able to acquire strips of at least 10 minutes in Stripmap mode.

MR-F-6 The satellites shall provide raw data in order to cover the Amazonian region every month (Stripmap mode).

MR-F-7 The SEO system shall provide 100 products per day up to level 1C. MR-F-8 The SEO system shall provide the following time performances:

• Revisit time: 1 day • Response time: 1 day • Information age: 5 hours

MR-F-9 The system shall be able to archive all the raw data files generated by the mission for at least 5 years after the satellites end of life.

4.3.2 Mission Operative Requirements MR-O-1 SEO shall have a full operational capability of at least 5 years (LEOP

and IOT time excluded). MR-O-2 Each SEO satellite shall be operative after a commissioning (LEOP and

IOT) period of 3 months. MR-O-3 SEO shall allow a direct access of the user to the system via a web

interface and MR-O-4 SEO mission operations shall be accomplished by means of a dedicated

ground segment MR-O-5 SEO mission ground segment operations shall include at least LEOP,

IOT, Reorbiting for repeating ground trace maintenance, collision avoidance tasks and at EOL, final satellite passivity and disposal.

4.3.3 Mission Constraints MC-1 The cost of the SEO Mission shall not exceed 500 M€ (launch costs

excluded) and including the ground segment and operations. MC-2 SEO should be based on international co-operation in terms of space

assets, ground segment and operations. MC-3 SEO shall be designed and realized in 5 years from the Kick-Off to be

held in early 2011 MC-4 SEO mission will be able to use GEO Data Relay Satellite for the

transmission to the ground segment

4.3.4 System Requirements SR1 The science data downlink rate shall be of 155 Mbit/s.


18

4.3.5 Space Segment Requirements

SS-‐R-‐1 The wet mass of each SEO satellite shall be less than 1500 kg.

SS-‐R-‐2 SEO overall space segment shall have an end-of-life full performance reliability not less than 60%.

SS-‐R-‐3

The satellite pointing : • accuracy will be 0.3 deg. • stability < 5 deg during maneuvers • ……..

SS-‐R-‐4 The SEO satellites shall have dimensions compliant with medium-‐class launchers.

SS-‐R-‐5

Satellite pointing requirement in deployment phase: • Reach sun-‐pointing in less than 3 orbits (start-‐up mode); • …………

SS-‐R-‐6

Satellite pointing requirement in Sun and Earth acquisition mode: • point the sun in less than 1000 s (sun acquisition mode); • point the earth in less than 1000 s;

SS-‐R-‐7

Observation mode pointing requirements: • pointing knowledge: 60 arcsec in 10 min of target

acquisition • pointing stability < 100 arcsec in 10 min of target acquisition • WGS84 surface height compensation • yaw steering compensation • geo-‐location accuracy < 40 m

4.4 Subsystem Study A detailed analysis would require a study of all subsystem, but in this example, only the “System”, “Mission” and “Payload” subsystems will be discussed in detailed. In the other Subsystem will be shown only the major results and principal considerations.

4.4.1 System The main design drivers for the system are:

§ Ensure a correct interpretation of all SEO requirements. § Set the proper Operational Mode and Mission Phases. § Minimization of Mass and Cost, promoting the use of off–the–shelf

equipments, existing qualified technologies and know-how.


19

The mission includes descriptions of the scientific goals, size, cost of the mission, together with programmatic and implementation details. The system monitors the entire process of data exchange, the realization of the main objectives and requirements and constrain. Keeps track of the historic. Monitors the solution of the baseline system. Opens other options (trade-off) if the study met interesting alternatives. Manages the mass budget by a margin of 20% of the total.

4.4.2 Mission The orbit has been optimised to fulfil the Primary Objective (see par. 4.2); for this reason the orbital inclination has been chosen in order to assure the maximum effectiveness in terms of Amazonian region coverage regardless for the global coverage of the Earth, which is not a mission objective. Therefore, the selected inclination was 18°. As SEO is a system for Earth observation, a repetitive ground track has been considered to add value, since it assures deterministically predictable time performances, and allows to produce a periodic request handbook (which is an interesting issue, provided that the mission’s primary objective is to have a periodic coverage of a geographic area). As the mission is based on the use of an active instrument (SAR), it is highly desirable to have a circular orbit not higher than 800 km and possibly below 650 km. Evolution of the Ground Track : The orbit has been designed to guarantee the full coverage of the target region (Amazonia) using only the ascending (or only the descending) passes. On one side this allows to guarantee the full coverage by design; on the other hand, half of the passes can be used for further observations of Amazonia or secondary target regions.

Fig. 10 - Real ground track shift with respect to the ideal one


20

The orbit cycle is 147 orbits or 9.82 days long. It has been assumed to keep the real ground track in a range of ± 1.5 km with respect to the ideal one. The real orbit considered is circular, a few hundreds meters higher than the ideal one, in order to compensate for the drag effect. To keep the real ground track in a range of ± 1.5 km with respect to the ideal one, from the Fig. 10 it is clear that about 1 station keeping manoeuvre (0.098 m/s) every 6 days is necessary. Orbit Performance: The assumed acquisition strategy foresees the coverage of the whole Amazonian region in two orbit “cycles” (a few less than 20 days): during the first cycle the odd passes are acquired, during the second cycle also the even passes are acquired. With this hypothesis, the system can accomplish the primary mission (by downloading all the data from the satellite) in 20 days. Of course, the actual strategy during operations can foresee to cover the Amazonia in 3 orbit cycles (less than 30 days, thus still fulfilling the requirement) resulting in further possibility to perform secondary targets imaging or emergency acquisitions on the Amazonia. End of Life : As the satellite re-enters the Earth atmosphere in less than 12 years, there is no need of an active de-orbiting or re-orbiting to comply with the disposal requirements of the Inter-Agency Space Debris Coordination Committee or to the European Code of Conduct on Space Debris (which prescribe to reduce the residual orbital lifetime below 25 years).

4.4.3 Payload The motivation for this mission proposal comes from the increasing science requirements for a continuous and global monitoring of climate and environmental variables with high resolution and on a reliable way [15]. At first, on the basis of a review of previous missions (Tandem L, BIOMASS ALOS, SAOCOM) the P and L Bands were considered. P-band SAR penetration capabilities are significant with regard to vegetation canopies, glacier or sea ice, and soil. Its vegetation canopy imaging capability is considered a key element in estimating vegetation biomass by means of remote sensing. The P Band was however excluded because the bandwidth allocation (ITU Regulation) is 6 MHz, limiting the resolution at about 60 m at 25° incidence. This limitation does not fit with the requested geometric resolution. The L-Band is the most appropriate frequency due to : • penetration capability in vegetated area, • low RF interference and ionospheric perturbations when compared to lower

frequency bands • high bandwidth of the available frequency allocation: 85 MHz on the basis of ITU

regulation. In this case the resolution can reach a value of few meters, fitting with the mission requirements. A Full polarimetric mode is suggested to satisfy wider customers needs. It improves the results with respect to vegetation classification and the analysis of motionless ground targets. A very long SAR antenna is needed to meet the SEO mission coverage/resolution requirements (the maximum expected antenna length is 20 m). The chosen radar central frequency for SEO-SAR is 1275 MHz, corresponding to a wavelength λ= 23.5 cm. Baseline [15]: The design of a space-borne synthetic aperture radar (SAR) with large swath coverage in full polarimetric mode is a challenging task in L band. The


21

requirement for full-polarimetry involves a doubling of the PRF, reducing the available swath and making range ambiguities a serious problem. An example is given by the case of the next Argentinean constellation, SAOCOM, made of two identical L band SAR. For the SEO mission, the SAR concept is based on a deployable planar array antenna, suited for conventional platform. A passive planar array antenna was selected in order to reduce mass, costs and complexity and to be compliant with smaller launchers. It means that the electronic part is not integrated directly in the antenna; the SSPAs are integrated in the electronic box located in the spacecraft body The antenna configuration consists of 9 elements with 12 sub-elements rows per panel. The central panel has sizes adapted to fit with the platform panel : 1.4 m (width) and 2.2 m (height). Sizes of the lateral panels are different. The total antenna length is 18,00 m. The antenna height was limited in the design phase at 2.2 m by the launcher requirement: as a matter of fact, in principle the idea was to use the VEGA launcher. Cross-section of the panel sandwich is about 4 cm thick. The antenna mass is 90 kg. The total SAR mass budget is about 400 kg including a 20% mass margin. The Peak Power Transmission is 2500 W.

Fig. 11 - SEO’s SAR antenna configuration

In the CEF model, the payload is composed by 4 units in order to simplify the configuration. Unit-1 and Unit-2 corresponds to the SAR electronics, whereas Unit-3 and Unit-4 were used to differentiate the lateral panels and the central panel of the SAR antenna (see ). Follown an instrument block diagram for SAR electronics:

Fig. 12 - SAR Equipment list

Interferometry mode: The estimation of dynamic processes on Earth surfaces requires systematic, long term and continuous observation strategies in order to detect short and long term changes with a sufficient accuracy.

ANTENNA

UNIT-3 Lateral panel of the SAR

antenna

UNIT-4 Central panel

of the SAR antenna

UNIT-1 Integrated

Central Electronic

UNIT-2 TRU (TX-RX Unit)

SAR ELECTRONIC

l


22

Depending on the environmental and/or anthropogenic process to be observed, there is a need for having different time intervals in the acquisition plan. The main goal in the analysis of this option is to establish if one satellite is sufficient for interferometric acquisitions related to the Earth’s dynamic processes or a constellation of at least two satellite is required. The Fig. 13 shows the required time intervals for sampling of different Earth’s dynamic processes. For deforestation, forest biomass changes and reforestation a time interval of weeks or months, one spacecraft is sufficient. The interferometry based on one spacecraft is acceptable to characterise the mentioned dynamic processes.

Fig. 13 - Dynamics of different Earth spheres and the requested interval for sampling the corresponding process in an unambiguous way [16].

SCANSAR acquisition mode: To improve the swath width, the Scansar acquisition mode is considered with electronic steering capability in elevation. Scansar acquisition mode requires a more complex software; in this case the acquired image is elaborated for each sub-swath. A summary of specifications for Scansar acquisition mode is given in the following table.

Table 3 – SCANSAR acquisition mode parameters

Item Specifications Average value subswath 40 km Total swath width 160 km Incidence angles 27°-‐30°-‐33°-‐36.2°

Single look Resolution 50 m x 50 m Band Width 680 Hz

Technology Readiness: SAR electronic is based on a consolidate technology and previous mission heritage. Reuse of existing X- and C-Band components was considered. The most critical part is the antenna which requires a pre-development


23

phase and bread-boarding activity. A review of the design is required. Cold redundancy was selected for the SAR electronics.

Table 4 - Payload technology requirements

4.4.4 Thermal The thermal analysis uses a simulation of Sun’s and Earth’s irradiance at the attitude selected in the study [15] and, given the internal power budgets (for bus and payload needs), provides as an output the design of a thermal control subsystem that allows all components inside and outside the S/C to remain inside strict temperature ranges.

Fig. 14 - S/C attitude as propagated along the orbit considered

4.4.5 Mechanical & Structure This analysis permits to define the structure of the Satellite and the main mechanisms [15].

Equipment and Text Reference

Technology Suppliers and TRL Level

Technology from Non-‐Space Sectors

UNIT 1 Consolidate heritage from previous missions 6 NO UNIT 2 Consolidate heritage from previous missions 6 NO

UNIT 3 and 4 Design based on heritage from previous missions

4 NO


24

Fig. 17 – SEO main elements

4.4.6 Configuration The Assumptions of the configuration subsystem in the preliminary study are: § The payload antenna must be aligned with flight direction and properly

oriented with respect to nadir; § The sensing devices (i.e. AOC sensors) are placed so that their fields of

view is unobstructed, in deployed configuration, by large appendages as solar arrays or antennas.

§ The selected low-inclination orbit imposes the orientation of solar arrays perpendicular to orbital plane and capable to rotate in order to maximize the sun incidence.

X BAND ANTENNA

PAYLOAD ANTENNA

SOLAR ARRAY

PL ANTENNA TILTING

MECHANISM

PAYLOAD

POWER

DATA HANDLING

AOCS

PROPULSION

COMMS

MECHANISM

STRUCTURES

COLOR CODE

PAYLOAD

POWER

DATA HANDLING

AOCS

PROPULSION

COMMS

MECHANISM

STRUCTURES

COLOR CODE

STAR TRACKERS

PAYLOAD ELECTRONICS

THRUSTERSTANK

RADIATOR

Fig. 15 - SEO overall view

Fig. 16 - SEO stowed configuration


25

The result in the configuration subsystem are:

Fig. 18 - Localization and pointing requirements for main components

4.4.7 AOCS It’s necessary define the Assumptions on this preliminary study: § Possibility to decouple yaw, pitch and roll dynamics § A PD controller is used to evaluate the attitude performances. Kalman filtering

techniques would strongly improve the attitude control performances § Star-trackers are placed as close as possible to the payload chassis so as to

minimize the thermo elastic distortions § The SAGA anomaly does not affect the AOCS equipment performances The result of this is the definition of the baseline design, see figure.

Fig. 19 - AOCS physical architecture


26

4.4.8 GrounSegment The SEO ground segment architecture can represented as in the following figure:

Fig. 20 - GS & Ops Architecture.

4.4.9 Thecnical Risk Risk management is an organized, systematic decision making process that efficiently identifies, analyzes, plans, tracks, controls, communicates, and documents risk to increase the likelihood of achieving the project goals. The risks themselves are characterized by the combination of the likelihood that a program or project will experience an undesired event and the impact or severity of the undesired event on mission success criteria. In the SEO case the assessment of the Risk is shown in the following table:

Table 5 - Risk Index Chart for SEO Mission, where 1A: low risk à 4D: very hight risk


27

4.4.10 Management Taking into account mainly the P/L needs in terms of pre-development/development//qualification activities and the assumed margin between the satellite delivery to the launch site date and the launch date itself, the launch date of January 2016 cannot be met. A more realistic schedule requests to set the launch date at the beginning of 2017.

Fig. 21 - SEO Bar Chart


28

4.4.11 Cost In order to minimize both technical risk and overall mission costs, a large reuse of existing and still flying equipments and technologies has been done. Notwithstanding this relevant amount of design reuse especially in many of the bus subsystems, some newness in the payload design and in the data handling subsystem have led to an increase of the satellite design complexity. The L-band frequency choice, coupled with the resolution and revisit time requirements do imply very large antenna overall dimensions. Therefore particular care in the mechanical, electrical and electromagnetic antenna design and tests must be devoted, bringing to an increase in development and integration costs. A Design Maturity Cost Margins (DMCM) of 15% has been considered. The following cost estimate is based upon the data available and collected from the various specialists after the last iteration

Fig. 22 - SEO Space Segment cost

4.5 Conclusion SEO The Primary Objective of the SEO satellite is to acquire SAR images of the amazonian region in stripmap mode, covering in one month the total area. The satellite/mission is required to meet revisit time, response time and information age requirements with an operative life of 5 years and to download the acquired data with a rate of at least 155 Mbit/sec. From a programmatic standpoint, the development is required to be completed in five years and to be compatible with a medium-class launcher. There is a constraint on the mission overall cost. The complete set of requirements has been collected in the “Space system for Earth Observation (SEO) - Mission Needs & User Objectives” document. During the study, two options were developed as for the TM/TC link: one employs a link with DLR + Ka-band and one another doesn’t. In the first case, two Ground stations are required, in the second case only 1. The best orbit was identified to be a low inclination in circular orbit. The VEGA launcher was discarded, due to the satellite dimensions exceeding the fairing envelope. The selected launcher was a PLSV indian launcher. The launch costs, should be shared with a copassenger. The system mass budgegt, in the worst case, amounts to 1312 kg including a 20% margin in the. The power budget (worst case) amounts to 1650 W. Structure, Thermal, Propulsion and AOCS S/S have been assessed to be feasible using well established/high TRL(see appendix) technologies.

STRUCTURE4,4%

MECHANISM0,4%

THERMAL CONTROL1,5%

POWER10,7%

TTC3,4%

DATA HANDLING8,4%

AOCS5,0%

PROPULSION2,1%SOFTWARE

1,9%

SEO SAR (ICE+TRU+Radar Antenna wings)

45,3%

AIT SYSTEM6,3%

PROGRAM LEVEL5,1%

GSE5,1%


29

The satellite configuration includes, as far as the communication s/s, 2 S-Band antennas , which typically provide a wide field of view and a link capability without steering capability and one fixed X-band antenna. The radar payload, in L-band, is a passive deployable planar array SAR antenna. The electronic is in a dedicated box located on the spacecraft body. The SAR antenna operates in stripmap mode, in full polarisation and the total mass is 80 kg. It has low TRL, thus requiring an extensive breadboarding/engineering modeling and qualification activities. The interferometry based on one spacecraft is acceptable as for deforestation, forest biomass changes and reforestation a time interval of weeks or months; one spacecraft is sufficient. From the schedule point of view, the study resulted in identifying that the system development cannot meet the 5 years requirement. One additional year is needed, mainly to accommodate adequate development activities for the most critical parts: the payload and its interfaces. The P/L is the design driver for the overall satellite design and planning. Riskwise, the schedule and, from a technical standpoint, the interfaces between P/L, COMM s/s, DH s/s and PW s/s are introducing the most relevant risks, to be mitigated by moving the delivery date and executing an extensive test campaign at avionic test bench level. The SEO SAR mission cost estimate has shown that the system may be designed, built and operated remaining well under the allocated budget. As expected the most costly element of the space segment is the SAR payload with its very wide radar antenna that deserves a special consideration for the assembly, integration and test phases. The use of a DRS system could be of interest in order to test existing technologies in the Ka band RF spectrum for scientific data transmission.

5. Lesson learned an Recommendations As often happens, the adaptation of a process to take full advantage of a new method is not straightforward. For a time the process goes on as before, taking partial advantage of the new method, but suffering from resistance to change. Adopting a new method often needs a change in the mentality of the people involved, and only when these actors are convinced can the method itself be fully exploited. The team leader is a very important figure, so its competence/skill is essentially to achieve the purpose of the study. The performance of the entire study depends from him. The task of this figure is the moderator and guide and this allows to highlight critical issues and problems under study. From him also depends the timing of project development. A figure inappropriately chosen could frustrate some important and essential points of the CEF. A cohesive and competent expert team also is essential to achieve effective interaction at appropriate work times. The iterative approach to the mission design allows the depth of the final product to be controlled. It is possible to study a mission at very high level in a very short time, or to go to detailed design over a longer period. The advantage of this procedure is that the project can be stopped to clarify some points, without loss of information. This activity of the CEF will have to be combined with the consolidation and evolution of the CEF infrastructure, in particular software, methodology, procedures and


30

standardization of document (format comun with other infrastructures). Other areas liable for improvements will be investigated (i.e. improvement of the database: for the collection of the data produced during the studies, or in support to cost and risk knowledge, etc.) Also the costumer contribution is relevant for the purpose of the study, because they can help adjust in real time the correct orientation of the project (avoiding loss of time).

6. Conclusion The assessment studies have shown that a mission design at the level of Phase-0 and Phase-A can be efficiently performed though the CEF in a much shorter time and with higher quality results than with traditional methods. From the bibliography is clear that the products of this instrument are considered more consistent than those produced via the classical approach. This approach in ASI is still at experimental stage, although its usefulness has already been widely demonstrated and verified with the studies performed up to date.

Acronym ASI Italian Space Agency AIV Assembly, Integration and Verification AOCS Attitude and Orbital Control System CDF Concurrent Design Facility CEF Concurrent Engineering Facility CIC Centre d'Ingénierie Concourante CNES Center National d’études Espatiales DH Data Handing DLR German Aerospace Center - Deuch DRS Data Relay Station EADS European Aeronautic Defence and Space Company ECSS European Cooperation for Space Standardization EGSE Electronic Ground Support Equipment EO End of Life ESA European Space Agency EU European Union FM Flight Model H/W Hardware I/F Interface I/O Input/Output IOT In Orbit Test LEOP Launch and Early Orbit/Operations Phase RF Radio frequency SAR System Requirement Review S/C Spacecraft SEO Space system for Earth Observation TM/TC Telemetry/Telecommand TRL Technology Readiness Level


31

TT&C Tracking, Telemetry and Command VGF Vegetation Gap Filler

Bibliografy

[1] Building Skills in Concurrent Design: A Partnership Between Education and Industry, SECESA 2010, 13-15 October 2010 N. Mathers (1), G. Graham (2) and M. Pakakis

[2] The Satellite Design Office at Astrium - A Success Story of an Industrial Design Center Application, Proceeding EusEC 2010, Rolf Mager, Ralf Hartmann System Engineering Division, Astrium GmbH 88039 Friedrichshafen, Germany

[3] CDF Presentation Info Pack 2008; http://www.esa.int/SPECIALS/CDF/SEM1OF1P4HD_0.html

[4] Status of the concurrent engineering facility at dlr bremen, Romberg, O., Braukhane, A., Schumann, H., DLR German Aerospace Center

[5] Web The ESA/ESTEC Concurrent Design Facility, Proceeding EusEC 2010, M. Bandecchi, B. Melton, B. Gardini, F. Ongaro, Noordwijk

[6] Concurrent Engineering at ESA: from the Concurrent Design Facility (CDF) to a Distributed Virtual Facility, The 14th ISPE International Conference on Concurrent Engineering, Sao Jose’ dos Campos, SP, Brazil, 16-20 July 2007, M.Bandecchi

[7] www.ecss.nl with ECSS-M-ST-10C_Rev.1(6March2009).pdf [8] Space mission Analisis and Design, Wiley J. Larson and James R. Wertz, 3ed. [9] An Advanced Methodology for the Design Process of a Satellite, Professor

Heinz Stoewer, Ralf Hartmann, L.A.J. Baron von Richter [10] Challenges for Concurrent Engineering on Launcher Design, EUCASS2013, D.

Haerens, G.Collange, D. Denier-Gegu, G.Hericher, A. Huet, A. Matthyssen [11] TAS-I Integrated System DEsign Center Activities for Remote Sensing

Satellites, SECESA 2010, 13-15 October 2010, M. Marcozzi, G. Campolo, L. Mazzini, R.Cialdini, G. Landella, R. Campanella internet

[12] Technology Readiness Levels Handbook For Space Applications, TEC-SHS/5551/MG/ap, iss.1 vers.6 , 2008

[13] The ESA Concurrent Design Facility: Concurrent Engineering applied to space mission assessments, CDF 2013, ESA/ESTEC Noordwijk – NL, presentation unclassified

[14] www.ecss.nl with ECSS-E-ST-10C(6March2009).pdf [15] SEO (System of Earth Observation), ASI CEF Study Report, team ASI/CEF,

July 2010 [16] An improved focusing method for geosynchronous, SA, Cheng Hu, Zhipeng

Liu, Teng Long, Volume 51, Issue 9, 1 May 2013, Pages 1773–1783, China, Advance in space research

[17] Chapter on spacecraft structures, Anarella C., taken from space-‐craft-‐structure.pdf [18] Lecture series ae2-‐S02, Wijker J. Delft University of Technology, 2002. [19] MediaGlobe study, SpaceTech 1989-‐1999, TopTech studies, TU-‐Delft [20] Elements of Spacecraft Design, AIAA, C. D. Brown , 2003.


32

Annex A TRLs are a set of management metrics that enable the assessment of the maturity of a particular technology and the consistent comparison of maturity between different types of technology, all in the context of a specific system, application and operational environment[11]. The following Table provides the complete set of basic definitions and explanations of the TRLs applicable to hardware. Analogue guidelines exist for the definition of software technology readiness levels.

Readiness Level Definition Explanation

TRL 1 Basic principles observed and reported

Lowest level of technology readiness. Scientific research begins to be translated into applied research and development. (See Paragraph 4.2)

TRL 2 Technology concept and/or application formulated

Once basic principles are observed, practical applications can be invented and R&D started. Applications are speculative and may be unproven. (See Paragraph 4.3).

TRL 3 Analytical and experimental critical function and/or characteristic proof-of-concept

Active research and development is initiated, including analytical / laboratory studies to validate predictions regarding the technology. (See Paragraph 4.4)

TRL 4 Component and/or breadboard validation in laboratory environment

Basic technological components are integrated to establish that they will work together. (See Paragraph 4.5)

TRL 5 Component and/or breadboard validation in relevant environment

The basic technological components are integrated with reasonably realistic supporting elements so it can be tested in a simulated environment. (See Paragraph 4.6)

TRL 6 System/subsystem model or prototype demonstration in a relevant environment (ground or space)

A representative model or prototype system is tested in a relevant environment. (See Paragraph 4.7)

TRL 7 System prototype demonstration in a space environment

A prototype system that is near, or at, the planned operational system. (See Paragraph 4.8)

TRL 8 Actual system completed and “flight qualified” through test and demonstration (ground or space)

In an actual system, the technology has been proven to work in its final form and under expected conditions. (See Paragraph 4.9)

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

The system incorporating the new technology in its final form has been used under actual mission conditions. (See Paragraph 4.2.10)

preliminary!study!of!a!space!mission!...

Documents