Changes between Version 39 and Version 40 of TBR/UserApp/Space/Proba_2


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Timestamp:
Nov 13, 2018, 7:08:41 PM (7 months ago)
Author:
Sal
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  • TBR/UserApp/Space/Proba_2

    v39 v40  
    33[[TOC(TBR/UserApp/Space/Proba_2, depth=2)]]
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    5 PROBA stands for PRoject for OnBoard Autonomy. The PROBA satellites are among the smallest spacecraft ever to be flown by ESA, but they are making a big impact in the field of space technology. PROBA-2 is the second of the series, building on nearly eight years of successful PROBA-1 experience.
     5PROBA stands for PRoject for !OnBoard Autonomy. The PROBA satellites are among the smallest spacecraft ever to be flown by ESA, but they are making a big impact in the field of space technology. PROBA-2 is the second of the series, building on nearly eight years of successful PROBA-1 experience.
    66
    77[[Image(https://devel.rtems.org/old_images/Proba-2-in-orbit-rear-view.jpg, 400px)]]
     
    149149Figure 4: Phoenix GPS architecture for PROBA-2
    150150
    151 It is a miniature receiver specifically designed for high dynamics space applications. It is based on SigTech’s
     151It is a miniature receiver specifically designed for high dynamics space applications. It is based on !SigTech’s
    152152commercial-off-the-shelf MG5001 receiver board but operates a proprietary firmware developed by DLR.
    153153Though originally designed for automotive applications, the receiver board has been qualified for space use in a
     
    359359For further info [http://ilrs.gsfc.nasa.gov/docs/ESA4S_06_11d.pdf PROBA-2 MISSION AND NEW TECHNOLOGIES OVERVIEW]
    360360
    361 = RetroReflector Array (RRA) Characteristics
     361= !RetroReflector Array (RRA) Characteristics
    362362
    363363
    364364PROBA-2 will use the same retroreflector package as was installed on Cryosat-1.
    365 More information about the retroreflector characteristics can be found in the document "CryoSat-LRR-01 Laser Retro Reflector Technical Description" (V. Shargorodsky/2002).
     365More information about the retroreflector characteristics can be found in the document "!CryoSat-LRR-01 Laser Retro Reflector Technical Description" (V. Shargorodsky/2002).
    366366PROBA-2 Retroreflector Information Form (29 August 2008, PDF):
    367367
     
    3783782. Array manufacturer: Scientific Research Institute for Precision Instruments, Moscow
    3793793. Link (URL or reference) to any ground-tests that were carried out on the array:
    380 4. Other missions using this LRA design and/or type of cubes: CryoSat-1/2, GOCE
     3804. Other missions using this LRA design and/or type of cubes: !CryoSat-1/2, GOCE
    381381
    382382For accurate orbital analysis it is essential that full information is available in order that a model of the 3-dimensional position of the satellite centre of mass may be referred to the location in space at which the laser range measurements are made. To achieve this, the 3-D location of the LRA phase centre must be specified in a satellite fixed reference frame with respect to the satellite's mass centre. In practice this means that the following parameters must be available at mm accuracy or better.
     
    418418Subsequently, Edisoft has reached an agreement with OAR to implement an RTEMS maintenance centre (see related link) in Europe. Edisoft has complemented the validation and the toolset associated with RTEMS for the specific needs of the European space industry. Gaisler Research also provides services based around RTEMS on ERC32 and Leon.
    419419
    420 RTEMS has already been used in several space applications, in particular FedSat (a scientific Research and Development microsatellite), the Surrey's Solid State Data Recorder (a component used in the Disaster Monitoring Constellation), ChipSat (a System-on-Chip architecture), the Electra UHF antenna of the Mars Reconnaissance Orbiter and in the Galileo GIOVE-A and Herschel-Planck satellites.
     420RTEMS has already been used in several space applications, in particular !FedSat (a scientific Research and Development microsatellite), the Surrey's Solid State Data Recorder (a component used in the Disaster Monitoring Constellation), !ChipSat (a System-on-Chip architecture), the Electra UHF antenna of the Mars Reconnaissance Orbiter and in the Galileo GIOVE-A and Herschel-Planck satellites.
    421421= Providing flight opportunities =
    422422
     
    442442 *  Combined carbon-fibre and aluminium structural panels, developed by Apco Technologies SA (CH)
    443443 *  New models of reaction wheels from Dynacon (CA), startrackers from DTU (DK) and GPS receivers from DLR (DE)
    444  *  An upgraded telecommand system with a decoder largely implemented in software by STT- SystemTechnik GmbH (DE)
     444 *  An upgraded telecommand system with a decoder largely implemented in software by STT- !SystemTechnik GmbH (DE)
    445445 *  A digital Sun-sensor, developed by TNO (NL)
    446446 *  A dual-frequency GPS receiver, developed by Alcatel Espace (FR)
    447447 *  A fibre-sensor system for monitoring temperatures and pressures around the satellite, developed by MPB Communications Inc. (CA)
    448  *  A new startracker development being test-flown before use on the BepiColombo mission, developed by Galileo Avionica (IT)
     448 *  A new startracker development being test-flown before use on the !BepiColombo mission, developed by Galileo Avionica (IT)
    449449 *  A very high-precision flux-gate magnetometer, developed by DTU (DK)
    450450 *  An experimental solar panel with a solar flux concentrator, developed by CSL (BE)
     
    551551- Estimations of remaining spacecraft magnetic dipole to reduce pointing error
    552552
    553 - Autonomous avoidance of star tracker Earth/Sun blinding
     553- Autonomous avoidance of star tracker !Earth/Sun blinding
    554554
    555555- Inertial mode with fixed scanning rate for moon calibration.
     
    569569Beyond the technology demonstration through the PROBA program, it is also noted that the AOCS software technology developed in the course of this program is now the baseline of the AOCS of a major operational mission of the GMES (Global Monitoring for Environment and Security) program: Sentinel-3. NGC Aerospace Ltd (NGC) of Sherbrooke, (Québec), Canada was responsible for the design, implementation and validation of the autonomous GNC (Guidance, Navigation and Control) algorithms implemented as part of the AOCS software of PROBA-1 and PROBA-2. NGC has the same responsibilities for the PROBA-V mission (Ref. 16).
    570570
    571 EPS (Electric Power Subsystem): The PVA (Photo-Voltaic Array) uses GaAs triple junction cells with an of efficiency of 28%. To obtain the operating voltage of 31.5 V, 18 cells are included in each string in series with a blocking diode. The PVA consists of a total of 25 solar strings taken into account the loss of one string on the most contributing PVA panel. The average solar string power under EOL conditions (summer solstice and T = 40°C) yields 12.8 W. The maximal incoming power at EOL during an orbit is 144 W. The energy budget for PROBA-V is derived for a bus power consumption of 140 W assuming a worst case day in the summer and while not taken into account the effect of albedo. A worst case power budget analysis indicated a maximum capacity discharge of 1.66 Ah. Use of a Li-ion battery. The battery cells provide a capacity of 1.5 Ah per string. The PROBA-V battery is sized to 12 Ah taking into account capacity fading and loss of a string.
     571EPS (Electric Power Subsystem): The PVA (Photo-Voltaic Array) uses !GaAs triple junction cells with an of efficiency of 28%. To obtain the operating voltage of 31.5 V, 18 cells are included in each string in series with a blocking diode. The PVA consists of a total of 25 solar strings taken into account the loss of one string on the most contributing PVA panel. The average solar string power under EOL conditions (summer solstice and T = 40°C) yields 12.8 W. The maximal incoming power at EOL during an orbit is 144 W. The energy budget for PROBA-V is derived for a bus power consumption of 140 W assuming a worst case day in the summer and while not taken into account the effect of albedo. A worst case power budget analysis indicated a maximum capacity discharge of 1.66 Ah. Use of a Li-ion battery. The battery cells provide a capacity of 1.5 Ah per string. The PROBA-V battery is sized to 12 Ah taking into account capacity fading and loss of a string.
    572572
    573573Launch: A launch of the PROBA-V spacecraft as a secondary payload is planned for Q4 2012. The primary launcher is currently assumed to be Vega with the VESTA adapter.
     
    612612|| ||||Launch||||Mass||||Size||||Orbit||||Launcher||||Power consumption||||RF||||Nominal Life||||Ground Station||||Developed by||
    613613||Proba II||||11/2009||||130 Kg||||600 x 700 x 850 mm||||Altitude between 700 km and 800 km, Sun-synchronous, Inclination 98.298 degrees||||Rockot||||53–86 Watts||||S-band, 64 kbit/s uplink; 1 Mbit/s downlink||||2 years||||Redu (Belgium)||||Consortium led by QinetiQ Space nv of Belgium||
    614 ||Proba V ||||Planned in 2012||||160 Kg||||800 x 800 x 1000 mm||||Sun-synchronised polar orbit, 820 km, with a 10:30 AM local time at the descending node ||||To be decided – designed to be compatible with Vega, Soyuz or Falcon 1E launchers.||||131.2 Watts||||S-band TxRx: 5W BPSK; X-band Tx: 6 W filtered OQPSK; MMU= 16 Gbit||||2.5 years||||Satellite’s mission control centre in Redu, Belgium||||OIP Space Systems||
     614||Proba V ||||Planned in 2012||||160 Kg||||800 x 800 x 1000 mm||||Sun-synchronised polar orbit, 820 km, with a 10:30 AM local time at the descending node ||||To be decided – designed to be compatible with Vega, Soyuz or Falcon 1E launchers.||||131.2 Watts||||S-band !TxRx: 5W BPSK; X-band Tx: 6 W filtered OQPSK; MMU= 16 Gbit||||2.5 years||||Satellite’s mission control centre in Redu, Belgium||||OIP Space Systems||
    615615= External links =
    616616