wiki:UserApp/Proba-V

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FUGUYS

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ProbaV

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The ‘V’ in its name stands for Vegetation: Proba-V will fly a reduced-mass version of the Vegetation instrument currently on board the Spot satellites to provide a daily overview of global vegetation growth.

{| align="center" |+PROBA V

|450px? |}

Bearing a different designation from its predecessors, Proba-V is an operational as well as experimental mission, designed to serve an existing user community.

The aim is to guarantee data continuity for the Vegetation dataset once the current Spot missions end.

{| border="1" |+ Proba-V facts and figures

|Launch date: |mid-2012

|Mass: |160 kg

|Orbit: |Sun-synchronised polar orbit, 820 km, with a 10:30 AM local time at the descending node

|Instrument: |Newly designed version of the Vegetation instrument flown on the Spot series

|Guest technology payloads: |Gallium Nitride amplifier incorporated in communication subsystem; Energetic Particle Telescope and one other payload to be decided at a later stage

|Prime contractor: |Qinetiq Space Belgium

|Payload developer: |OIP Space Systems

|Ground Station: |Satellite’s mission control centre in Redu, Belgium complemented by a data reception station to be located in the north of Europe.

|Launcher: |To be decided – designed to be compatible with Vega, Soyuz or Falcon 1E launchers. |}

PROBA-V (Project for On-Board Autonomy - Vegetation)

The PROBA-V (Vegetation) mission definition is an attempt, spearheaded by ESA and CNES, to accommodate an improved smaller version of the large VGT (Vegetation) optical instrument of SPOT-4 and SPOT-5 mission heritage on a small satellite bus, such as the one of PROBA-2.

As of 2008, small satellite technologies have reached a level of maturity and reliability to be used as a platform for an operational Earth observation mission. Furthermore, advancements in the techniques of detectors, optics fabrication and metrology are considered sufficiently mature to permit the design of a compact multispectral optical instrument.

The C/D Phase started in July 2010. The system CDR (Critical Design Review) took place in the spring of 2011. The acceptance review is planned for Dec. 2011 and the flight acceptance review is planned for the spring of 2012.. ESA is responsible for the overall mission, the technological payloads and for the launcher selection.

Background:

The VGT instruments (VGT1 & VGT2), each with a mass of ~160 kg and fairly large size, have provided the user community with almost daily global observations of continental surfaces at a resolution of 1.15 km on a swath of ~2200 km. The instruments VGT1 on SPOT-4 (launch March 24, 1998) and VGT2 on SPOT-5 (launch May 4, 2002) are quasi similar optical instruments operating in the VNIR (3 bands) and SWIR (1 band) range.

The Vegetation instruments were jointly developed and funded by France, Belgium, Italy, Sweden, and the EC (European Commission). The consortium of CNES, BelSPO (Federal Public Planning Service Science Policy), SNSB (Swedish National Space Board) and VITO (Flemish Institute for Technological Research) is providing the user segment services (data processing, archiving, distribution). Vegetation principally addresses key observations in the following application domains:

  • General land use in relation to vegetation cover and its changes
  • Vegetation behavior to strong meteorological events (severe droughts) and climate changes (long-term behavior of the vegetation cover)
  • Disaster management (detection of fires and surface water bodies)
  • Biophysical parameters for model input devoted to water budgets and primary productivity (agriculture, ecosystem vulnerability, etc.).

As of 2008, a Vegetation archive of 10 years of consistent global data sets has been established permitting researchers access on a long-term basis. The SPOT-5 operational lifetime is estimated to expire in 2012. Pleiades, the next French satellite for Earth Observation, is solely dedicated to high-resolution imaging (on a fairly narrow swath) and will not embark any instrument providing vegetation data.

Since the SPOT series spacecraft will not be continued and the SPOT-5 spacecraft will eventually fail — there is of course a great interest in the EO user community to the Vegetation observation in the context of a smaller mission, affordable to all concerned.

PROBA-V will continue the production of Vegetation products exploiting advanced small satellite technology. However, this implies in particular a redesign of the Vegetation payload into a much smaller unit to be able to accommodate it onto the PROBA bus.

Overview of key requirements of the PROBA-V mission

  • Data and service continuity: filling the gap between SPOT-VGT and the Sentinel-3 mission
  • Spectral and radiometric performance identical to VGT
  • GSD: 1 km mandatory, improved GSD is highly disirable: 300 m (VNIR bands), 600 m (SWIR band). Image quality and geometric accuracy, equal to or better than SPOT-VGT
  • Provision of daily global coverage of the land masses in the latitudes 35º and 75º North and in the latitudes between 35° and 56° South, with a 90% daily coverage of equatorial zones - and 100% two-daily imaging, during day time, of the land masses in the latitudes between 35º North and 35º South..

{| align="center" |+Artist's view of the PROBA-V spacecraft (image credit: ESA)

|File:ProbaV Auto12.jpeg? |}

An extensive feasibility study and trade-off work was undertaken to identify a solution that could meet not only the technical challenges, but that could also be developed and tested within a tight budget of a small satellite mission.

The PROBA-V project of ESA includes the Space Segment (platform contract award to QinetiQ Space NV of Kruibeke, Belgium - formerly Verhaert), the Mission Control Center (Redu, Belgium) and the User Segment (data processing facility) at VITO NV. VITO (Vlaamse instelling voor technologisch onderzoek - Flemish Institute for Technological Research) is located in northern Belgium. VITO’s processing center of VGT1 and 2 data (SPOT-4 and SPOT-5) is operational since 1999. VITO is also the prime investigator and data service provider of PROBA-V for the user community including product quality control.

Implementation schedule:

  • The Phase B of the project started in January 2009
  • SRR (System Requirements Review) is in Q4 of 2009
  • PDR (Preliminary Design Review) in Q2 of 2010
  • HMA (Heterogeneous Mission Access) and QA4EO (Quality Assurance for Earth Observation) implementation for user data. Planned interoperability with GSCDA V2 (GMES Space Component Data Access Version 2).

{| align="center" |+PROBA-V project organization (image credit: ESA)

|File:ProbaV Auto11.jpeg? |}

Spacecraft:

An industrial team, led by QinetiQ Space NV (Belgium), is supported by several European subcontractors and suppliers, and is responsible for the development of the flight satellite platform, the vegetation payload and the Ground Segment.

The spacecraft bus (fully redundant) is of heritage from the PROBA-1 and PROBA-2 missions (structure, avionics, AOCS, OBS with minor modifications). The PROBA-V spacecraft has a total mass of ~160 kg, and a volume of 80 cm x 80 cm x 100 cm. The three-axis stabilized platform is designed for a mission lifetime of 2.5 years.

The spacecraft resources management is built around ADPMS (Advanced Data and Power Management System), which is currently flying on PROBA-2. The data handling part of ADPMS is partitioned using compact PCI modules. A cold redundant mass memory module of 16 Gbit is foreseen for PROBA-V. The newly developed mass memory will use NAND flash technology.

The power distribution and conditioning part of ADPMS supplies an unregulated bus, with each equipment having its internal DC/DC converter. The power conditioning system is designed around a Li-ion battery.

Overview of PROBA-V subsystems

{| border="1" |+ Overview of PROBA-V subsystems

|Avionics |ADPMS (cold redundant),MPM (Main Processor Module): LEON2-E Sparc V8 processor, 50 MHz,42 MIPS, 10 FLOPS,Mass memory Module: 16 Gbit Flash, EDAC protected |PROBA-2 ; New development

|EPS (Electric Power Subsystem |Photo-Voltaic Array : Triple junction GaAs? cells; Cover glass CMG 100AR coating, 25 strings, 18 cells per string;Battery:12 Ah Li-ion (7s8p) ABSL 18650HC cells |Herschel; PROBA-1

|Bus structure |Aluminum (AA2024-T3);Aluminum (AA7075-T7351);3 CFRP (EX-1515/M55J + Redux 312L) outer panels |New development

|AOCS actuators |3 magnetotorquers (internally cold redundant); 4 reaction wheels (3 + 1 for redundancy); 2 magnetometers (cold redundant |ROBA-2/PROBA-1

|AOCS sensors |2 star trackers; 2 GPS (cold redundant); AOCS IF box (internally redundant); RW Power Supply box (internally redundant) |PROBA-1/-2; New development

|Onboard SW |Operating System: RTEMS (Real-Time Executive for Multiprocessor Systems) |PROBA-2

|Thermal |Passive (MLI and paint) |

|RF communications |S-band TxRx?: 5W BPSK; X-band Tx: 6 W filtered OQPSK; MMU (Mass Memory Unit) = 16 Gbit |PROBA-1/-2; ; New development |}

{| align="center" |+PROBA-V spacecraft accommodation (image credit: QinetiQ Space)

|File:ProbaV Auto10.jpeg? |}

AOCS (Attitude and Orbit Control Subsystem)

AOCS (Attitude and Orbit Control Subsystem) provides three-axis attitude control including high accuracy pointing and maneuvering in different spacecraft attitude modes. The AOCS SW is an extension of the one of PROBA-2, including the following algorithms required by the on-board autonomous mission and payload management:

  • Prediction of land/sea transitions using a land sea mask to reduce the amount of data generated
  • Optimization of attitude in Sun Bathing mode to enhance incoming power while avoiding star tracker blinding
  • Momentum dumping without zero wheel speed crossings during imaging
  • Estimations of remaining spacecraft magnetic dipole to reduce pointing error
  • Autonomous avoidance of star tracker Earth/Sun? blinding
  • Inertial mode with fixed scanning rate for moon calibration.

The AOCS hardware selection for PROBA-V consists of a high accuracy double star tracker head, a set of reaction wheels, magnetotorquers, magnetometers and a GPS receiver.

The main AOCS modes are: Safe, Geodetic, Sun Bathing and Inertial mode.

  • The satellite Safe mode is used to detumble the spacecraft after separation from the launcher and it will be used to recover from spacecraft anomalies.
  • The Geodetic mode is used during nominal observation of the Earth’s vegetation. In this mode the payload is pointed towards the geodetic normal to the Earth’s surface. An extra steering compensation, yaw-steering, is added in this mode, to minimize the image distortion caused by the rotation of the Earth. This yaw-steering maneuver ensures that the spectral imagers are oriented such that the lines of pixels are perpendicular to the ground-trace at each moment. In this mode the star trackers and the GPS receiver are used as sensors and the reaction wheels as actuators.
  • On each orbit, the spacecraft enters the Sun Bathing mode from -56º latitude until entry of eclipse. This is to enhance the incoming power.
  • The Inertial mode coupled with an inertial scanning of the Moon at a fixed rate is used for monthly radiometric full moon instrument calibration purposes. The pointing towards the moon takes 2.5 min, 9 min for scanning the moon and 2.5 min to return to nominal observation mode. It is sufficient to have the moon in the FOV of the SI (Spectral Imager) for a number of pixels.

Beyond 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.

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.

Launch

Launch: 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.

Orbit: Sun-synchronous orbit, altitude = 820 km, inclination = 98.8º, LTDN (Local Time on Descending Node) = 10:30 hours (with a drift limited between 10:30 and 11:30 AM during the mission lifetime).

RF communications: S-band for TT&C transmissions and low-gain antennas with omni-directional up- and downlink capability. The uplink symbol rate will be fixed at 64 ks/s, while the downlink can be set to a high rate (< 2 Ms/s) for nominal imaging or to a low rate at 329 ks/s for off-nominal conditions. The CCSDS protocol is used for the TT&C transmissions.

X-band downlink of payload data is in X-band at a data rate of 35 Mbit/s. The onboard mass memory is 88 Gbit. The Redu station (Belgium) is being used for TT&C communication services. The X-band uses two cold redundant high-rate X-band transmitters and two nadir pointing isoflux antennas, both RHCP.

The S-band transceivers will be connected to RS422 outputs (cross strapped) of ADPMS while the X-band transmitters (8090 MHz) will be connected to the LVDS outputs not cross-strapped. The X-band link budget results in a link margin of 6 dB which will allow a reduction of the RF output power. Therefore the X-band transmitter will be designed (customer furnished item) to support various output power settings such that after commissioning, a lower output power might be selected.

Data compression: The massive amount of data produced by the instrument is beyond the capabilities of the bandwidth available on board of a small satellite. Data are reduced by using a lossless data compression algorithm implemented in a specific electronics. The data compression ratio obtained using standard CCSDS compression algorithms (CCSDS 133.0 B-1) is shown in Table 2.

Compression ratio

{| border="1" |+ Spectral band

|Blue |10.8

|Red |7.2

|NIR |SWIR 2

|Red |5.4 |}

Table 2: Overview of compression rates== S-band ==

The selection of an S-band transceiver and the development of an innovative and generic X-band transmitter for small satellites has been initiated in a collaborative program between CNES and ESA and is funded under GSTP-5 (General Support Technology Program-5). The X-band transmitter is a high-performance device optimized for the needs and constraints of small platforms for which small volume, low mass, low power consumption, and low cost cost are important parameters. Moreover, some key features such as modulation (filtered Offset-QSK), coding scheme (convolutional 7 ½), data and clock interfaces (LVDS packet wire serial interface) have been selected in compliance with CCSDS recommendations, but also to ease the interoperability with most of the existing on-board computers and ground station demodulators.

X-band

The development of the new X-band transmitter is based almost exclusively on COTS components to achieve at the same time high performances and low recurrent cost. The transmitter also features an innovative functionality with an on-board programmable RF output power from 1-10 W which allows to match finely with the chosen bit rate, and to reduce as much as possible the margins of the link budget and therefore the consumption power. PROBA-V is the first mission to use this newly developed transmitter. The transmitter has a mass of 1 kg, a size of 160 mm x 115 mm x 46 mm, an in-orbit life time of 5 years, and a radiation hardness of 10 krad. Data rates from 10-100 Mbit/s are available. The X-band transmitter was manufactured by TES Electonic Solutions of Bruz, France.

{| align="center" |+Overview of the transmitter architecture (CNES, TES)

|File:ProbaV AutoF.jpeg??|} |}

{| align="center" |+Photo of the X-band transmitter (image credit: CNES, ESA)

|File:ProbaV AutoE.jpeg? |}

Sensor complement: (VGT-P)

The PROBA-Vegetation payload is a multispectral spectrometer with 4 spectral bands and with a very large swath of 2285 km to guarantee daily coverage above 35 latitude. The payload consists of 3 identical SI (Spectral Imagers), each with a very compact TMA telescope. Each TMA, having a FOV of 34º, contains 4 spectral bands: 3 bands in the visible range and one band in the SWIR spectral range.

VGT-P is restricted to imaging land and dedicated calibration zones. On-board the spacecraft there is for each spectral imager a land sea mask that is provided by the PI (Principal Investigator). The land sea mask removes the pixels that contain only sea and it dictates when each SI should be in imaging mode.

OIP (Optronic Instruments & Products, Belgium) is the industrial prime contractor for the payload and is responsible for the design and development of the PROBA-V instrument and AMOS (Belgium) is responsible for the manufacturing and alignment of the telescope. The major payload challenge lies in the fact that the wide-swath imaging instrument has to fit into a small satellite with limited resources. The TMAs and the SWIR FPA have to be developed for the VGT-P since no COTS products are available

Each SI (Spectral Imager) contains one telescope, a beam splitter to separate the VNIR from the SWIR spectral bands, spectral bandpass filters to select the spectral bands, and the VNIR and SWIR focal plane arrays. The spectral bands will be realized by spectral bandpass filters centered on 460, 658, 834 and 1610 nm, with bandwidths of respectively 42, 82, 121 and 80 nm. The filters will be applied on the detector windows.

The optical axis of the central telescope will point to nadir and the two outer telescopes will point 34º from nadir. Together the three TMAs will cover a complete FOV of 102º. The optical system is telecentric, and the aperture is located at the position of the second (spherical) mirror.

{| align="center" |+Conceptual accommodation of the VGT-P inside the PROBA-V spacecraft (image credit: OIP, ESA)

|File:ProbaV AutoD.jpeg? |}

Figure 6 shows the payload mounted on the PROBA-V platform. Given the reduced size of the platform, a H-shape structure, the only practical location of the payload is on the anti-velocity panel. This accommodation, with respect to a solution with the payload in the middle of the structure, has the advantage of a very simple assembly and clean mechanical interface. The drawback is a larger temperature gradient due to the close vicinity of the payload to the solar panel.

{| align="center" |+Block diagram of the VGT-P (image credit: OIP)

|File:ProbaV AutoC.jpeg? |}

Legend to Figure 7:

  • ROE (Read Out Electronics)
  • PSU (Power Supply Unit)
  • DHU (Data Handling Unit)
  • PEU (Peripheral Electronics Unit)
  • MLI (Multi-Layered Insulation)

TMA telescope development

TMA telescope development: VGT-P makes use of a set of three such telescopes, identical to each other. The purpose of the related ESA GSTP (General Support Technology Program) development is to demonstrate the feasibility of one item of the set with respect to its required optical quality, and to secure the instrument development. The entire telescope is an athermal design made of the same aluminium material. The mirrors quality is achieved by SPDT and the alignment rely on the very precise matching of the mirrors with the mounting structure.

Taking into account the mission constraints and objectives, including the innovative features of the instrument, a full-aluminum design was selected. This choice allows taking benefit from the recent developments in ultra-precision milling and turning techniques, as well as in optical aluminum production. Furthermore, this leads to a homothetic telescope behavior. The optical performance requirement of the telescope with regard to MTF (Modulation Transfer Function) is given in Table 4.

SPDT (Single Point Diamond Turning): Diamond turning is a process of mechanical machining of precision elements using Computer Numerical Control (CNC) lathes equipped with natural or synthetic diamond-tipped cutting elements. The SPDT process is widely used to manufacture high-quality aspheric optical elements from crystals, metals, acrylic, and other materials. Optical elements produced by the means of diamond turning are used in optical assemblies in telescopes, scientic research instruments and numerous other systems and divices. Diamond turning is specifically useful when cutting materials that feature aspheric shapes such as TMA surfaces.

{| border="1" |+ Performance requirements of MTF

|Band |Nominal MTF (%) |2? MTF (%) |Max. frequency (lp/mm)

|Blue |68.1 |53 |38.5

|Red |68.5 |54 |38.5

|NIR |68 |53.7 |38.5

|SWIR |71 |62.4 |20.0 |}

{| align="center" |+Optical design concept of the TMA (ray tracing diagram), image credit: OIP

|File:ProbaV AutoB.jpeg? |}

Baffle design (Ref. 20): The aim of the baffle design is to block the out-of-field light which could enter the instrument and reach the detector, directly or through one or several reflections on the mirrors. This 1st order analysis didn’t consider vanes on the baffles and diffusion on M1 of out-of-field light.

The preliminary baffle layout is presented in Figure 9. It comprises 7 baffles: 1 at the entrance aperture of the instrument and 6 placed inside the instrument. An aperture stop is also placed at the level of the secondary mirror.

The baffle #1 is placed at the entrance of the instrument. Its role is to limit the out-of-field light that could directly reach the mirrors. The combination of the baffles #1 and #2 stops the direct view of the M3 mirror through the instrument entrance. The length of the upper side of the entrance baffle is defined to stop the light which could directly reach the M3 mirror and that could not be stopped by the lower side of the entrance baffle and by baffle #2. Some out-of-field light can also reach the M2 and M3 mirrors after reflecting on M1. This cannot be totally avoided but the length of the lower side of baffle #1 has been chosen in such a way that this straylight is stopped by the baffle #3 after reflecting on M3. The baffle #3 is placed below the M2 mirror and stops the direct view of the M1 mirror by the VNIR detector. The baffle #4 is a critical location where reflection or diffusion on the M2 structure can occur and bring stray light to the VNIR detector which is very close. Vanes will be placed at this location. The baffles #5 and #6 are placed near the focal planes to isolate the detectors from each other. The baffle #7 avoids a direct view to the SWIR detector from the M1 or M3 mirrors.

{| align="center" |+Proba-V TMA preliminary baffles layout (image credit: CSL, OIP, ESA/ESTEC)

|File:ProbaV AutoA.jpeg? |}

{| align="center" |+Figure 10: Illustration of the optical assembly of VGT-P and two star trackers on the optical bench (image credit: OIP)

|File:ProbaV Auto9.jpeg? |}

SWIR detector development: This development concerns the large format SWIR focal plane array containing at least 2704 pixels with 25 µm pitch. The solution selected uses the mechanical butting technique with 3 overlapping detectors of 1024 pixels and approximately 80 pixels in the overlap area. In Figure 11 the linear detector arrays are shown in green, while the ROICs (Readout Integrated Circuits) are presented in red. Xenics NV of Leuven, Belgium, is developing the InGaAs? SWIR detector array.

Several techniques were evaluated to realize the required alignment accuracy of the 3 PDA (Photo Diode Array) subarrays in the FPA. The requested alignment accuracies are:

  • In plane alignment accuracy, ?X and ?Y = ± 12.5 µm
  • Out of plane alignment accuracy, ?Z = ± 50.0 µm
  • Subarray PDA separation = < 1.5 mm.

{| align="center" |+Schematic view of of the mechanically butted SWIR detector array (image credit: OIP, Xenics)

|File:ProbaV Auto8.jpeg? |}

{| align="center" |+Figure 12: Drawing of the subarray alignment tools with the 3 PDAs (green) mounted on the mount (image credit: OIP, Xenics)

|File:ProbaV Auto7.jpeg? |}

{| align="center" |+Photo of the fully assembled FPA in its package (image credit: OIP, Xenics)

|File:ProbaV Auto6.jpeg? |}

Thermal design of the VGT-P instrument:

One of the major drawbacks of using multiple optical systems in parallel while imaging, is the effect of pointing inaccuracies due to thermo-elastic and mechanical deformations. It is obvious that such pointing errors can easily destroy the quality of the images. For the VGT-P, the stringent geo-location requirements demand the instrument to be thermally stabilized as much as possible to reduce any thermo-elastic disturbances.

Since the PROBA platform is fairly limited in the delivery of power, VGT-P needs to be very efficient in its power use. As a direct consequence, there is no possibility to have an active thermal control system to stabilize the instrument. The thermal design of the instrument must therefore be very carefully assessed and engineered.

Thermal isolation:

Firstly, as the surrounding satellite panels are heavily fluctuating in temperature during the orbit, it is of the utmost importance to shield the instrument thermally from these platform variations. To reduce the radiative heat loads from the environment, the instrument is completely wrapped in a 12 layer MLI. To reduce the conductive heat loads from the mounting plane, the instrument is mounted by means of titanium quasi isostatic mounting feet. These quasi isostatic mounting feet also play a major role in the transfer of the thermo-mechanical deformations from the underlying platform to the optical bench as they strongly reduce these deformations. Therefore, these titanium flexures as they are called not only serve as a thermal isolation, but also acts as a thermo-elastic isolator.

Power reduction: A natural step to reduce the thermo-elastic effects on the instrument is to reduce as much as possible the heat load on the optomechanics. Therefore, all non critical and heavy heat dissipating detector read-out electronics are separated from the optics. The FPAs of the telescope only contain the detector and electronic components which drive the radiometric performances of the instrument. These FPA electronics are connected through a flex rigid to the ROE (Read-Out Electronics) which is thermally and structurally disconnected from the optomechanics. All major heat dissipating components are located in there.

Obiously, also the central electronics (DHU and PSU) are separated from the optomechanical imaging system. By doing this, the total power dissipation on the optical bench is only 9W, which is less than ¼ of the total power dissipation of the complete VTG-P instrument.

Heat dissipation:

To dissipate this heat load, a radiator is needed. Several concepts were proposed and analyzed. The most efficient radiators point towards deep space which would enable us to cool down the complete instrument to very cold temperatures. This had a drawback that additional heaters would have been needed to stabilize the thermal regime of the instrument to normal working temperature. Moreover, as the instrument is always pointing downwards towards Earth, the radiator would have been located on the side of the instrument which naturally induces an asymmetry in the optomechanics. Such asymmetry is not desired in an imaging sensor with stringent pointing requirements. Moreover, heat pipes would have been mandatory to extract as efficient as possible all heat of the detectors towards the radiator which unnecessarily complicated the complete design.

From a thermo-elastic point of view, it was highly desirable to respect the symmetry of the instrument as much as possible and to symmetrically extract the heat from the FPA’s on the optical bench. Thus, it was chosen to locate the radiator in front of the instrument and point it towards the earth surface. As the earth is thermally quite stable at a fairly modest temperature and as the payload is always pointed nadir, this is the perfect heat drain for the instrument. The implementation of this concept reduces the complexity dramatically: the radiator, covered with aluminized Teflon, is connected through two thermal straps towards the front of the instrument without the need to install heat pipes.

Stability:

Stability is the key aspect of thermo-elastic performance. Of course, without the possibility of an active thermal control system, stability is quite difficult to achieve in a thermal environment which is constantly varying over the orbit.

To tackle this problem, the first stage was to avoid the randomness in the heat loads on the instrument and to have constant thermal regime along the orbit. As the payload is encircling the Earth with its radiator pointing at nadir, the heat load on the radiator is subjected to a varying regime from sunlit to eclipse and back. From the point of view of efficient power use, the imaging circuits on the instrument are switched off by the satellite if no imaging is needed (over the oceans, over the poles, during eclipse). This would induce different thermal regimes from one orbit to the other, which is not acceptable from pointing point of view. But leaving all electronics switched on during non operation is a no go considering the lack of power. As a compromise, during sunlit conditions and when the imaging electronics is powered off, a heater located on the detector with a heat load equal to the heat load of the detector and FPA is powered. In this way, the heat loads on the optical system remain constant during sunlit. During eclipse, all is switched off. - As a consequence, a constant thermal regime on the optics is established: during 1/3 of the orbit (eclipse) the radiator faces only IR and the instrument is switched off. During the 2/3 of the orbit, the radiator sees IR and albedo and the instrument is switched on.

Gradients:

The final challenge in the thermal design is to avoid thermal gradients in the instrument as gradients are hard to control and can severely affect the thermo-elastic performance. As already described, the heat extraction has respected the symmetry of the instrument. An unavoidable asymmetry is the location of the Star Trackers as they have their own limitations. The heat load from the FPA and the detectors on the telescopes is normally entering the instrument through the TMAs to the top skin of the optical bench. However, this would heavily distort and bend the optical bench as the top skin will expand more than the bottom. To reduce this effect, thermal straps are designed to extract most of the heat (4/5) from the detectors and the FPA towards the optical bench, the rest is still entering the TMA structure. To reduce the thermal bending, the heat straps are mounted on the side of the optical bench to avoid the bending of the bench.

For further info File:ProbaV Auto6.jpeg

PROBA missions

The following table provides an overview of existing PROBA missions performed by ESA.

{| border="1" |+ PROBA Missions overview ! !! Proba II !! Proba V

! Launch

| 11/2009
Planned in 2012

! Mass |130 Kg |160 Kg

! Orbit | Altitude between 700 km and 800 km, Sun-synchronous, Inclination 98.298 degrees | Sun-synchronised polar orbit, 820 km, with a 10:30 AM local time at the descending node

!Launcher |Rockot |To be decided – designed to be compatible with Vega, Soyuz or Falcon 1E launchers. |}

ESA and RTEMS validation and tools

Saab Space AB performed a validation of the real-time operating system RTEMS. Since it is available for many different targets and includes a multitude of functionality, ranging from I/O drivers to file-systems and beyond, it was agreed to only focus on the parts that were applicable for European space community applications. This implied that only the ERC32 target and a limited sub-set of the configurable RTEMS managers had to be considered.

Subsequently, 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.

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.

See also