Prepared by Antonio De Luca DG-SET
Reference ESA-DG-SET-2016-1378
Issue/Revision 2.0
Date of Issue 14/09/2017
estec
European Space Research and Technology Centre Keplerlaan 1 2201 AZ Noordwijk The Netherlands T +31 (0)71 565 6565 F +31 (0)71 565 6040 www.esa.int
INTERFACE REQUIREMENT / SPECIFICATION / INTERFACE CONTROL DOCUMENT / EID
ESEO Satellite-Launcher Interface Requirements Document
Table of contents:
1 INTRODUCTION ... 4
2 APPLICABLE AND REFERENCE DOCUMENTS ... 4
2.1 Applicable Documents ... 4
2.2 Reference Documents ... 4
3 DEFINITIONS AND ABBREVIATIONS ... 4
3.1 Definitions ... 4
3.2 Abbreviations ... 5
3.3 Document Conventions... 5
4 MISSION REQUIREMENTS ... 6
4.1 Overview ...6
4.2 Specific Design Characteristics of relevance to the launcher ... 7
4.2.1 De-Orbiting Mechanism... 7
4.2.2 Cold-gas µpropulsion ... 7
4.2.3 Optical Payload ... 7
4.2.4 Spacecraft Activation... 7
4.2.5 ESEO Reference Frame ... 8
4.3 Orbit Injection ... 8
4.4 Separation ...9
5 MECHANICAL INTERFACE REQUIREMENTS ... 9
5.1 Spacecraft Data ...9
5.2 Mass, Alignment, Inertia ... 13
5.3 Separation System... 14
6 ELECTRICAL INTERFACE REQUIREMENTS ... 15
6.1 Isolation, grounding and bonding ... 15
6.2 Monitoring signal ... 16
6.3 Separation Command ... 16
7 ENVIRONMENTAL REQUIREMENTS ... 16
7.1 Spacecraft Data ... 16
7.1.1 Low Frequency sine vibration ... 16
7.1.2 High Frequency Random Vibration ... 18
7.1.3 Shock ... 19
7.2 Mechanical Environment ... 20
7.3 Thermal Environment ... 20
7.4 EMC Environment ... 20
8 OPERATIONAL INTERFACE REQUIREMENTS ... 22
8.1 Spacecraft Data ... 22
8.2 Cleanliness ...23
8.3 EGSE ...23
8.4 FGSE ... 24
8.5 Stay-out zones ... 26
8.6 Hoisting Points, Handles ... 26
8.7 Remove Before Flight Items ... 26
8.8 Pre-Launch Activities ... 26
1 INTRODUCTION
The Satellite-to-Launcher Interface Requirements Document is aimed at the definition of the requirements to be satisfied by the interfaces between the satellite and the Launch Segment for the ESEO project. Requirements concerning different aspects are presented: mechanical and thermal properties, electrical and RF characteristics and environment. Main handling, storage, and pre- launch activities requirement are also defined.
2 APPLICABLE AND REFERENCE DOCUMENTS 2.1 Applicable Documents
Ref. Title
[AD. 1] ESA-DG-SET-2016-1377, Provision of Launch Services for ESEO
Table 2.1: Applicable Documents
2.2 Reference Documents
Ref. Title
[RD. 1] VEGA User’s Manual, Issue 3, Arianespace, 2006 [RD. 2] DNEPR User’s Guide, Issue 2, Kosmotras 2001
[RD. 3] EHB0003, ROCKOT User’s Guide, Issue 5, EUROCKOT Launch Services, 2011 [RD. 4] ISRO/ VSSC/PSLV/05, PSLV User’s Manual, Issue 5, ISRO
[RD. 5] 2000785F MkII MLB User Manual
[RD. 6] AS\12_0005|SYS\PLA\PS\AR\01, ESEO Power System DDF [RD. 7] European Directive Equipment under Pressure
[RD. 8] 2002204A-Separation-Switch-Data-Sheet2, MkII MLB Separation switches
Table 2.2: Reference Documents
3 DEFINITIONS AND ABBREVIATIONS 3.1 Definitions
The definitions and glossary of terms from ECSS-S-ST-00-01C apply to this document.
3.2 Abbreviations
AD Applicable Document
AIT Assembly Integration and Test BP Battery Pack
DL Downlink
ESA European Space Agency
ESEO European Student Earth Orbiter GS Ground Station
LEO Low Earth Orbit LMP Langmuir Probe
MAIT Manufacturing, Assembly, Integration and Test MCC Mission Control Center
MDD Mission Description Document MLI Multi-Layer Insulation
MPS Micro-propulsion System OBDH On Board Data Handling PA Product Assurance P/L Payload
PFM Proto Flight Model RD Reference Document SA Solar Array
Red Redundant S/C Spacecraft S/S Subsystem
SSO Sun Synchronous Orbit Tbl Time before launch TC Telecommand TM Telemetry
UL Uplink
UTC Universal Time Coordinates
3.3 Document Conventions
Within this document requirements are identified by an annotation in the left margin:
IRD-XXX-nn
Where IRD-XXX is a mnemonic code;
nn is a consecutive numerical identifier.
Paragraphs without this annotation are not requirements, and only provide information.
Requirements marked TBC (To Be Confirmed) have some uncertainty associated, usually with a numerical value. They will be confirmed, by the Agency, at a later stage. This confirmation will be made in agreement with the Contractor.
Where a value is identified as TBD (To Be Defined), the Agency does not have sufficient information to define it yet. The definition will be provided later, in agreement with the Contractor.
Comments or Info, where necessary, appear in an oblique typeface, indented between horizontal rules, as shown in this paragraph. They are normally used to provide a necessary interpretation for a requirement.
4 MISSION REQUIREMENTS 4.1 Overview
The European Student Earth Orbiter (ESEO) is a micro-satellite mission to Low Earth Orbit. It is being developed, integrated, and tested by European university students as an ESA Education Office project.
ESEO will orbit the Earth taking pictures, measuring radiation levels and testing technologies for future education satellite missions.
It is aimed at providing students with unparalleled hands-on experience to help prepare a well- qualified space-engineering workforce for Europe’s future.
In particular, ESEO satellite has the following mission objectives:
to take pictures of the Earth and/or other celestial bodies from Earth orbit for educational outreach purposes;
to provide dosimetry and space plasma measurement in Earth orbit and its effects on satellite components;
to test technologies for future education satellite missions.
The first one will be achieved by the use of a micro camera (uCAM) operating in the visible spectrum.
To fulfill the second objective two instruments will be operated on board:
plasma diagnostic probe (LMP)
tri-dimensional dosimeter instrument (TRITEL) In particular, the LMP shall measure:
electron density
electron temperature while the TRITEL shall measure
LET spectra
absorbed dose
dose equivalent
In order to provide high-speed datalink for payload data transmission a dedicated S-band transmitter (HSTX) will be provided as payload complement.
The realization of third objective will consist in the flight-testing of GPS receiver for orbit determination and a De-Orbit Mechanisms (DOM) to be activated at the end of the mission.
4.2 Specific Design Characteristics of relevance to the launcher 4.2.1 De-Orbiting Mechanism
The ESEO spacecraft incorporates, as a payload, an external De-Orbiting Mechanism (DOM), activated upon release of a dedicated command from the spacecraft bus. The DOM is deployed at the end of the mission lifetime to enlarge the effective satellite area in order to increase the orbital drag. The DOM deploys a sail, which is attached to elastic boom arms. The boom arms as well as the sail are rolled up around a central spool in the middle of the device and are held in position by three Kevlar chords. Thus, the device is compactly stored in a single unit of 100 x 100 x 60 mm before actuation. Upon deployment, the Kevlar chords are cut by thermal knives and the strain energy stored in the boom arms is used to deploy the sail.
Figure 4.1 – ESEO De-Orbiting Mechanism
4.2.2 Cold-gas µpropulsion
Another technology demonstrator on board is the cold-gas µpropulsion subsystem, which uses nitrogen as propellant. The nitrogen will be loaded by the S/C Contractor in the relevant tank (150 bar) at launch site premises. This subsystem is directly inherited by the one present on ALMASat-1 and launched in 2012 with the VEGA Maiden Flight.
4.2.3 Optical Payload
The camera payload is sensitive to particulate contamination and it is protected by a dedicated cover installed at the baffle and must be kept closed on ground. Such cover will be removed immediately before the mechanical integration of the satellite with the launcher. Class 100,000 (or ISO8) cleanroom conditions are required when the protective covers are installed and they will be applied for the whole duration of the launch campaign till closing of the fairings.
4.2.4 Spacecraft Activation
The ESEO activation will be initiated at the separation; neither the on-board computer nor the UHF receiver will be ON during launch and ascent.
4.2.5 Components and materials
All electronic components and materials used for the manufacturing of the ESEO Satellite are COTS (Commercial Off The Shelf); no items subject to any inport/export restriction are present on board.
4.2.6 ESEO Reference Frame
The following figures report the definition of the ESEO body reference frame; the origin is placed at the centre of the bottom plate in correspondence with the crossing point of the diagonals; the Z axis is negatively oriented in direction of Zenith once in orbit. The Y axis is normal to the radiator’s surface, the X axis completes the right-handed frame.
Figure 4.2 – ESEO Body Reference Frame
4.3 Orbit Injection
IRD-SSO-010 The ESEO Orbit Shall be Sun-Synchronous (SSO).
IRD-SSO-020 The orbit injection should be made according to the following parameters, and including launcher dispersion:
Apogee: 535.9977 km
Perigee: 517.3837 km
Inclination: 97.4788 deg
Y X
Y
X
Z Z
X
Y
Z
LTAN: baseline option: 10:30 am (backup option: 01:30 pm) Note: the reference design altitude of ESEO is 523 km.
Different orbit injection parameters, meeting in any case SSO requirement, may be proposed; but they shall be negotiated and agreed before contract signature.
4.4 Separation
IRD-SEP-010 The maximum allowed separation speed, with respect to the launcher, is 0.5 m/sec (tbc) along satellite Z axis.
IRD-SEP-020 The maximum allowed angular rate at separation is 5 deg/sec around each of the satellite axes.
Note: there is no specific ejection direction required by the satellite.
5 MECHANICAL INTERFACE REQUIREMENTS 5.1 Spacecraft Data
The current estimation of the spacecraft mass is 60.00 kg (including 10% margin). The CoG position is estimated in:
Table 5.1: CoG coordinates
X [mm]
Y [mm]
Z [mm]
0.894 -0.25 339.7 The inertia tensor is the following:
ESEO MoI [kg/m3]
Ixx = 1.811368 Ixy = -0.00431 Ixz = 0.023103 Iyx = -0.00431 Iyy = 1.807942 Iyz = -0.00114 Izx = 0.023103 Izy = -0.00114 Izz = 0.348481
The inertia tensor is known with ±10% precision.
The overall spacecraft launch envelope for the current configuration of ESEO is reported in the following figure.
Figure 5.1: ESEO spacecraft launch envelope
On the bottom side, in order to provide a suitable payloads layout the following constraints in terms of stay-out zones and FOV have been considered:
Langmuir probe (LMP) stay-out zone
High-speed S-band Transmitter (HSTX) beam width
Micro-camera (uCAM) FOV
Launch vehicle (LV) interface stay-out
The Langmuir probe detector is installed in nadir pointing configuration and a proper volume around the sensing element shall be guaranteed. The radius of the stay-out zone has been estimated in 91 mm. As represented in Figure 5.2.
Figure 5.2: LMP stay-out zone verification
The beamwidth of the antenna of the High-speed S-band transmitter (HSTX)determines the stay- out zone necessary to guarantee optimal performance. The envelope is reported in Figure 5.3.
Figure 5.3: HSTX antenna beamwidth and stay-out zone
The micro-camera FOV is represented in Figure 5.4.
Figure 5.4: uCAM FOV
The bottom side of ESEO is predisposed to accommodate the Mark II Lightband manufactured by Planetary Systems Corp. as separation system. This is the baseline solution considered for ESEO
Figure 5.5: Mark II Lightband and adapter
The selected size of the adapter ring, is 13 inches diameter.
Figure 5.6: Mark II Lightband sizing table
The drawings of Mark II Lightband adapter ring interface with the launch vehicle are reported in Figure 5.7.
Figure 5.7: Mark II lightband adapter
In case of the clamp-band separation system will not be the baseline one, its adapter ring can be attached to the mechanical interface at the bottom plate of the ESEO spacecraft by means of 20 ISO 7462 Stainless Steel screw, at least grade A2-70, size M6.
Figure 5.8: Adapter ring mechanical interface dimensions (aft flange detail)
A further back-up configuration is kept in case of unavailability of the baseline solution;
The concept behind this technical solution is represented and highlighted in green in Figure 5.. The bottom plate could be constrained to the adapter and separation system in two or more (up to four in the case of ESEO) points using a 45° pin. A series of centering cones allows the correct positioning of the spacecraft on the adapter.
Figure 5.9: Details of the spacecraft bottom plate acting as interface with the launch vehicle
5.2 Mass, Alignment, Inertia
IRD-MAI-010 Reference frames of Satellite and Launcher shall be reported in the ICD.
IRD-MAI-020 Misalignment of Spacecraft CoG shall be reported in the ICD.
5.3 Separation System
IRD-SEP-010 The baseline separation system is based on Lightband Mk II. Any other separation system shall be mechanically compatible with mechanical interfaces reported in figures 5.8 and 5.9.
IRD-SEP-020 After separation, the FOVs relevant to the payloads and reported in sect. 5.1 shall be respected by the adapter for the separation system
IRD-SEP-030 The adapter ring shall mount the 4 separation switches required to provide the redundant signal confirming the detachment from the launch adapter and switching on the ESEO satellite.
IRD-SEP-040 The mass of the adapter ring, remaining connected to the spacecraft after separation, shall be limited to 0.8 kg (tbc).
IRD-SEP-050 The contact surface between spacecraft mechanical interface and the launch vehicle adapter ring shall have the following characteristics:
Roughness: 1.6
Flatness: 0.05
Surface finishing: Alodine according to MIL-DTL-5541 Type I, Class 3
IRD-SEP-060 The separation system and adapter shall be able to guarantee that the specified volume inside the adapter as showed in figure 5.10 is completely free.
Figure 5.10: Required Stay out zone for separation system and adapter
6 ELECTRICAL INTERFACE REQUIREMENTS 6.1 Isolation, grounding and bonding
The present Section considers the ESEO S/C in the ready-to-launch configuration with the separation mechanism installed. Therefore it is assumed that, as standard practice and flight heritage, the whole separation system is capable to guarantee the required electrical continuity between the bottom of the ESEO S/C and the interface between the LV and the separation system.
IRD-EIF-010 Isolation of input power lead and return from chassis shall be equivalent to a parallel combination of a resistor of 1 MΩ minimum and a capacitor C < 50 nF, as per ECSS-E-ST-20-07C
IRD-EIF-020 The bonding resistance shall not exceed 0.1 ohm DC per bonding junction between adapter and separation system chassis and bond strap and between bond strap and launch vehicle structure.
The bonding stud shall have the following characteristics:
perpendicular to a flat surface
length of 15 mm +/- 2 mm
M4 threads
no obstructions within 30 mm from the bonding stud in all directions surface around the bonding stud (within 20 mm from the stud) treated to provide low contact resistance
IRD-EIF-030 Bonding connections shall be installed such that vibration, expansion, contraction or relative movement incident to normal service use will not break or loosen the connection to such an extent that the resistance will vary during the movement.
IRD-EIF-040 Bolts or screws shall not be relied on as a ground path, i.e. shall not carry any intentional current.
The bonding through metallic structures shall be preferred and the bonding resistance shall not exceed 0.1 ohm DC per bonding junction.
IRD-EIF-050 Non-structural mechanical equipment without electrical function, which are not expected to work as fault current path, do not contain electrical parts or equipment, and are prevented from inadvertent or accidental contact with electrical parts or equipment, shall have a bonding resistance to the structure not exceeding 1000 ohm DC.
IRD-EIF-060 All isolated conducting items having an area greater than 100 square centimetres which are subject to frictional charging or plasma-induced current flow or charging, shall have a mechanically secure conducting connection to adjacent conductive structure. The resistance of the connection shall not exceed 1 ohm DC.
IRD-EIF-070 All composite materials which are subject to frictional charging or plasma–
induced current flow or charging shall have a mechanically secure conductive connection to adjacent conductive structural. The resistance of the connection shall not exceed 1000 ohm DC.
6.2 Monitoring signal
IRD-EIF-080 Separation switches for detachment detection shall be installed on the separation system and acquired by the ESEO on board avionics, to flag the beginning of the in-orbit operation.
6.3 Separation Command
IRD-EIF-090 The separation command shall be provided by the LV to the Separation System.
7 ENVIRONMENTAL REQUIREMENTS 7.1 Spacecraft Data
ESEO has been designed and developed taking into consideration the environmental conditions imposed by different launchers; because of its reduced dimensions and mass, it has always been considered as secondary payload, and therefore it was decided to have it structurally compatible with different launchers.
Details on the environmental conditions have been collected from the most common known launch vehicle user-manuals. In order to guarantee compatibility with a range of different launch vehicles the Design Limit Loads to be considered for the sizing and verification of the ESEO spacecraft have been calculated considering the envelope of all the reported loads and the margin policy reported in ECSS-E-ST-32-10C, ECSS-E-ST-32C and ECSS-E-10-03A.
The following launch vehicles have been taken into account: VEGA, DNEPR, PSLV, ROCKOT.
7.1.1 Low Frequency sine vibration
In Error! Reference source not found.1 the summary of all the FoS applied in the E SEO mission for the low frequency
sine vibration levels analysis are reported.Table 7.1: Mechanical Factors Of Safety for low frequency sine vibration
The resulting load envelope is shown:
LONGITUDINAL Frequency
[Hz]
DLL [g]
4 – 10 13.3 mm (0 - peak)
10 5.0
100 5.0
LATERAL Frequency
[Hz]
DLL [g]
2 6.6
13 6.6
15 8.3
25 8.3
25 4.2
100 4.2
Figure 7.1: ESEO low frequency sine design limit levels (launchers envelope) – Design Limit Loads
7.1.2 High Frequency Random Vibration
In Error! Reference source not found.2 the summary of all the FoS applied in the ESEO m ission for the high frequency random vibration levels analysis are reported.
Table 7.2: Mechanical Factors Of Safety for high frequency random vibration
The resulting load envelope is shown in Error! Reference source not found.2.
LONGITUDINAL Frequency
[Hz]
DLL PSD [g2/Hz]
20 0.014
80 0.014
160 0.044
320 0.07
640 0.07
1280 0.034
2000 0.01
Overall grms 8.88
Figure 7.2: ESEO high frequency random design limit levels (launchers envelope) – Design Limit Loads
7.1.3 Shock
In Error! Reference source not found.3 the summary of all the FoS applied in the ESEO m ission for the high frequency random vibration levels analysis are reported
Table 7.3: Mechanical Factors Of Safety for shock
The resulting load envelope is shown in Error! Reference source not found.3.
LONGITUDINAL Frequency
[Hz]
DLL SRS [g]
30 5
100 60
1000 1500
1500 2000
4000 3000
5000 3000
10000 1500
Figure 7.3: ESEO shock design limit levels (launchers envelope) – Design Limit Loads
7.2 Mechanical Environment
IRD-MEC-010 Compatibility between launcher performance and Spacecraft Limit Loads due to Low Frequency Sine vibration shall be confirmed and demonstrated.
IRD-MEC-020 Compatibility between launcher performance and Spacecraft Limit Loads due to High Frequency Random vibration shall be confirmed and demonstrated.
IRD-MEC-030 Compatibility between launcher performance and Spacecraft Limit Loads due to Shock shall be confirmed and demonstrated.
7.3 Thermal Environment
IRD-THE-010 Non Operative temperature range for the Satellite shall always maintained between -10 and +40 deg. C for the whole launch campaign including lift off and flight till separation.
7.4 EMC Environment
In order to guarantee the EMC between the spacecraft and the launcher during operation, a list of operative uplink and downlink frequencies are reported, as well as the radiated emission and susceptibility masks of the S/C. Please note that the threshold values shown in this section take into account for a margin of 6dB, so they can be considered as final value for comparison with launcher masks. On the other hand, this limits can be tailored on the including the radiation pattern of each antenna, which is neglected at this level of description and assumed as 1 dB omnidirectional.
Finally, thresholds refer to 1 m of distance from the S/C.
Downlink frequencies
Unit involved Central frequency [MHz] Bandwidth [kHz]
ESEO TMTC subsystem 437,000 50
HSTX payload 2280,000 6000
AMSAT payload 145,930 25
Uplink frequencies
Unit involved Central frequency [MHz] Bandwidth [kHz]
ESEO TMTC subsystem 435,200 MHz 50
AMSAT payload 1264,000 MHz 1000
GPS 1575,420 1000
Table 7.4: summary of UL and DL frequencies in use on ESEO S/C
Figure 7.4 shows the radiated susceptibility mask of the ESEO S/C. Please note that the notches included in the mask (one for each RF receiver included on-board the S/C, refer to Table 7.-UL freq.) is to be considered only when the involved unit is switched on.
Figure 7.4: RS mask of ESEO S/C
Figure 7.5 shows the radiated emission mask of the ESEO S/C. Please note that the notches included in the mask (one for each RF receiver included on-board the S/C, refer to Table 7.-DL freq) is to be considered only when the involved unit is switched on.
The Spacecraft will be OFF for the whole launch phase, and switched ON after separation.
Figure 7.5: RE mask of ESEO S/C
IRD-EMC-010 Radiated emissions from the launcher shall not induce any damage or malfunction to the ESEO Spacecraft.
IRD-EMC-020 Radiated emissions from the launcher shall not provoke the auto switch-ON of the ESEO spacecraft
8 OPERATIONAL INTERFACE REQUIREMENTS 8.1 Spacecraft Data
In order to allow preparation of the spacecraft before launch and diagnostic activities during ground operations a specific set of skin connectors to EGSE has been implemented on the top plate (+Z) of the structure as represented in Figure 8.1 (tbc).
Figure 8.1: Position of the ESEO EGSE/Skin Connectors
A protective cap on each of the skin connectors will be added before launch.
Battery discharge time: in order to guarantee the correct operability of the ESEO satellite once in orbit, it is strictly required a maximum delay time between the last batteries recharging operation and the launch of less than 7 days (ESEO Maximum parking time).
This is due to the minimum current absorption due to the isolation switch installed to keep the satellite switched off until launch. The estimation of the ESEO maximum Time-before-launch (Tbl) is provided in the following table:
Parameter Value Units
Average cell capacity (@4.15V) 2200 mAh
number of packs 6
Total batteries capacity in mAh 13200 mAh
Total batteries capacity in Wh 293.04 Wh
Time before launch (Tbl) in days 7 dd Time before launch (Tbl) in hours 168 hh
Isolation Switch consumption 5 mA
Complete batteries depletion time 2640 h
Residual capacity after Tbl 12360 mAh
Residual capacity after Tbl 274.392 Wh
Batteries SOC after Tbl 93.6 %
Separation MODE consumption 20 W
maximum Sun acquisition time in
minutes 35 min
maximum Sun acquisition time in 0.58 hh
hours
Residual capacity after Separation 240.11 Wh Batteries SOC at Sun acquisition 81.9 %
Table 8.1: Maximum parking time before launch estimation
8.2 Cleanliness
IRD-CLN-010 Class 100,000 (or ISO8) cleanroom conditions shall be provided and they shall be applied for the whole duration of the launch campaign till closing of the fairings.
8.3 EGSE
ESEO EGSE and ESEO Spacecraft will be delivered in two separate transport containers. These containers have been already used in past missions, including Vega VV-01 flight and spacecraft integration activities at CSG launch site.
The ESEO EGSE will be connected to the ESEO satellite by means of a dedicated umbilical on the top of the structure, for integration reasons, and it will communicate with the other P/L and S/S via CAN protocol as an external device.
The operating modes of the EGSE are indicated in the related user manual, that will be delivered together with the EGSE at the end of the review of the internal instrumentation for the ESEO purposes.
In the test procedure the user manual and the test specific operation to be performed via the EGSE shall be reported.
Briefly, the ESEO EGSE is configured as in
Figure 8.
2.
Figure 8.2: ESEO EGSE Configuration
The ESEO EGSE container dimensions are: 0.6 m x 0.8 m x 1.2 m.
Figure 8.3: ESEO EGSE Transport Container
8.4 FGSE
The ESEO S/C will be provided with a dedicated FGSE, in order to perform the final filling of the onboard N2 tank (ESEO MPS MPV) before the final integration on the Launch Vehicle. The ESEO FGSE is composed of the following items:
Nitrogen Cylinder vessel, with handle valve (pressure > 200 bar, volume > 5 liters)
Pressure regulator, with relief valve, high pressure analogic manometer and low pressure analogic manometer
Needle valve, for gas flow control
Digital Manometer
Female “Quick release coupling”
The FGSE will be interfaced to the ESEO MPS Fill-valve, accessible from the radiator panel on the +Y direction, as shown in Figure 8.4.
Figure 8.4: ESEO MPS MPV Fill-Valve positioning (in red circle)
The ESEO Fill-Valve is provided with a male “Quick Release coupling” connector to complete the setup shown in Figure 8.5.
Figure 8.5: ESEO MPS MPV filling setup
8.5 Stay-out zones
IRD-OIF-010 In order to guarantee safety for operators and avoid damages to the spacecraft, during operations (including the activities to be performed after integration on the launch vehicle fairing) the stay-out zone represented in the following figure shall be guaranteed.
8.6 Hoisting Points, Handles
Tbd
8.7 Remove Before Flight Items
IRD-RBF-010 For safety during ground operations, the ESEO separation system shall be provided with a mechanical device that shall avoid the release of the spacecraft in case of inadvertent activation of the release device on ground;
the device shall be removed before encapsulation into the Launch Vehicle fairings. The device shall be red/orange-colour and be provided with a red/orange ‘Remove Before Flight’ flag.
8.8 Pre-Launch Activities
IRD-PLA-010 It shall be possible to recharge the ESEO on-board batteries by means of the dedicated EGSE, via the umbilical cable to be connected on top of the satellite.
IRD-PLA-020 It shall be possible to recharge the ESEO on-board batteries once the spacecraft is integrated on the launcher and just before closure of the fairings.
IRD-PLA-020 If the launch does not occur before the expiration of the Tbl, the on-board batteries shall be charged again in order to guarantee the feasibility of the mission.
IRD-PLA-030 Tank filling: it shall be possible to fill the ESEO on board propeller tank by means of the dedicated FGSE, via the related fluidic to be connected on the side of the satellite, in correspondence of the radiator panel, on the +Y side of the spacecraft, wrt the S/C reference frame. The FGSE is manually operated and provides the required N2 supply to the tank, up to the 150 bar operating pressure.
IRD-PLA-040 Female connector from N2 bottle side shall be compatible with ESEO male
“Quick Release coupling” connector
