Solar Orbiter Mission
Solar Orbiter Mission
Solar Orbiter is a satellite mission of ESA (in the footsteps of Helios, Ulysses, SOHO and the Cluster missions) to explore the inner regions of the sun and the heliosphere from a near-sun orbit. Solar Orbiter is part of the ESA's Science Program Cosmic Vision 2015-2025. The Solar Orbiter project was initially selected by ESA's Science Program Committee in Oct. 2000 and re-confirmed as part of the ESA program in 2003. The Solar Orbiter mission of ESA and the SPP (Solar Probe Plus) mission of NASA (launch scheduled for 2018) are part of the common GHO (Great Heliophysics Observatory) program.
In 2011, the Solar Orbiter mission has undergone extensive study over a period of more than 10 years, both internally in ESA and in industry. This has resulted in a mature, detailed design that satisfies the requirements placed on the mission by the science objectives and addresses the key risk areas. - ESA's Science Program Committee selected the Solar Orbiter mission for implementation on October 4, 2011 with a launch scheduled for 2017. 1) 2) 3) 4)
ESA-NASA collaboration: NASA and ESA have a mutual interest in exploring the near-Sun environment to improve the understanding of how the Sun determines the environment of the inner solar system and, more broadly, generates the heliosphere itself, and how fundamental plasma physical processes operate near the Sun. A NASA-ESA MOU (Memorandum of Understanding) for a Solar Orbiter mission cooperation was signed in March 2012. 5)
For Solar Orbiter, also referred to as SolO in the literature, ESA is providing the spacecraft bus, integration of the instruments onto the bus, mission operations, and overall science operations. NASA is providing an EELV (Evolved Expendable Launch Vehicle) that will place the Solar Orbiter spacecraft into an inner heliospheric orbit with perihelia ranging from 0.28 to 0.38 AU and aphelia from 0.73 to 0.92 AU. The SolO nominal science mission will begin with a series of perihelion passes where the spacecraft is nearly co-rotating with the Sun. It will then use multiple Venus gravity assist maneuvers to move its orbital inclination to progressively higher helio latitudes, reaching 25° by the end of the nominal prime mission phase and around 34° by the end of the extended mission.
The overall objective is to provide close-up views of the sun's high latitude regions - to study fundamental physical processes common to solar, astrophysical and laboratory plasmas. The Solar Orbiter will, through a novel orbital design and its state-of-the-art instruments, provide exactly the observations required. 6) 7) 8) 9) 10) 11) 12) 13) 14)
• During the nominal operational lifetime, the Solar Orbiter operational orbit shall have the following parameters:
- Minimum perihelion radius larger than 0.28 AU to maximize the reuse of BepiColombo technology
- Perihelion radius within 0.30 AU in order to guarantee multiple observations close to the Sun
- Inclination with respect to solar equator increasing to a minimum of 25o (with a goal of 35o in the extended operational phase).
• At minimum perihelion passage, the spacecraft shall maintain a relative angular motion with respect to the solar surface such that individual solar surface features can be tracked for periods approaching one solar rotation.
• The Solar Orbiter system lifetime shall be compatible with a launch delay of 19 months (launch window locked to the next Venus gravitational assist opportunity).
Scientific requirements: The overarching objective of the Solar Orbiter mission is to address the central question of heliophysics: How does the Sun create and control the heliosphere? Achieving this objective is the next critical step in an overall strategy to address one of the fundamental questions in the Cosmic Vision theme: How does the Solar System work? To this end, the Solar Orbiter will use a carefully selected combination of in-situ and remote-sensing instrumentation, a unique orbit and mission design, and a well-planned observational strategy to explore systematically the region where the solar wind is born and heliospheric structures are formed.
The broad question that defines the overarching objective of the Solar Orbiter mission is broken down into four interrelated scientific questions:
1) How and where do the solar wind plasma and magnetic field originate in the corona?
2) How do transients drive heliospheric variability?
3) How do solar eruptions produce energetic particle radiation that fills the heliosphere?
4) How does the solar dynamo work and drive connections between the Sun and heliosphere?
Common to all of these questions is the requirement that Solar Orbiter make in-situ measurements of the solar wind plasma, fields, waves, and energetic particles close enough to the Sun that they are still relatively pristine and have not had their properties modified by dynamical evolution during their propagation. The Solar Orbiter must also relate these in-situ measurements back to their source regions and structures on the Sun through simultaneous, high-resolution imaging and spectroscopic observations both in and out of the ecliptic plane.
Basic mission requirements of Solar Orbiter:
- Total cruise phase duration < 3 years (goal) with valuable science during the cruise phase
- Orbital period in 3:2 resonance with Venus
- At least one orbit with perihelion radius < 0.25 AU and > 0.20 AU (science phase)
- Inclination with respect to solar equator increasing to a minimum of 30o
- During the extended operational lifetime, the Solar Orbiter operational orbit shall reach an inclination with respect to solar equator not lower than 35o (goal)
- Support a payload of 180 kg and 180 W (including 20% maturity margins) with a data rate of 100 kbit/s
- Provide onboard mass memory and communications with a single ESA deep-space ground station (New Norcia, Western Australia) in support of the science observations
- Fail-safe onboard autonomous operations during the perihelion passages (15 days without ground contact, in extremely harsh thermal environment).
The mission includes a nominal mission phase and a potential extended mission phase (corresponding to 6 solar orbits). The spacecraft consumables and radiation sensitive units shall be sized to meet the duration with extended phase: 9.5 years.
Table 1: Historical overview of the Solar Orbiter program (Ref. 12)
As with all spacecraft, mass and volume are at a premium due to launch vehicle constraints; however, the Solar Orbiter main spacecraft body is further constrained due to the fact that a sizable portion of the budget is taken up by the heat-shield, along with the fact that the spacecraft must be optimized to fit behind the heat-shield with sufficient margin to cover off-pointing cases, e.g. due to spacecraft anomalies. The Solar Orbiter spacecraft main body is approximately 2 m3 (with stowed appendages). With 33 instrument units to accommodate on-board, the allowable volume of each instrument unit must be tightly controlled.
The Solar Orbiter spacecraft configuration is dominated by the presence of the heat shield located at the top of the spacecraft in order to protect the spacecraft from the intense direct solar flux when approaching perihelion. The heat shield is over-sized to provide the required protection to the spacecraft box and externally-mounted units, in combination with the attitude-enforcement function of the FDIR (Failure Detection, Isolation and Recovery). The mechanical platform revolves around a robust, reliable and conventional concept with a central cylinder, four shear walls and six external panels. This concept is inspired by Astrium's Eurostar 3000 spacecraft platform. The design meets the Solar Orbiter's mission requirements according to a low-risk and low-cost philosophy.
In April 2012, ESA awarded a contract to build its next-generation Sun explorer to Airbus DS (former Astrium UK, Stevenage). Astrium UK will lead a team of European companies who will supply various parts of the spacecraft. 15) 16) 17) 18)
Figure 1: Artist's view of a baseline spacecraft in solar orbit (image credit: ESA)
The spacecraft and mission PDR (Preliminary Design Review) was completed on March 7, 2012.Following contract negotiations with the prime contractor, Phase-B2/C/D is proceeding with subsystem-level procurement and several lower-tier procurements.
The spacecraft is three-axis stabilized and always sun-pointed. Given the extreme thermal conditions at 45 solar radii (or 0.22 AU), equivalent to about 20 solar constants or approximately 28 kW/m2, a phenomenal amount of power from which the majority of the spacecraft must be protected. The thermal design of the spacecraft has been considered in detail. Accordingly, the bulk of the spacecraft is protected from the sun by a local heatshield (also referred to as sunshield) on the +X panel face of the spacecraft, combined with a stringent maintenance of a sun-pointing attitude for the spacecraft at all times during periods close to the sun (below ~0.7 AU). 19) 20) 21)
The spacecraft configuration is based on a square structure housing a simple mono-propellant propulsion system with no main engine . Due to the stringent environment encountered on the heliocentric orbits, the spacecraft is always sun-pointed and protected from solar irradiation by a heatshield. This heatshield covers the spacecraft bus and some of the external components such as in-situ instruments. It contains aperture openings providing the required field of view (FOV) for the remote sensing instruments. 22)
The avionics architecture is based on segregated processing functions of the platform and the payload data. The OBMU (On-Board Management Unit) is in charge of the spacecraft command / control, running the DHS (Data Handling Subsystem), AOCS and mission software component, and housing the interfaces with the platform equipments and the payload support unit. The PDPU (Payload Data Processing Unit) supports all functions of the sensor complement. Onboard communications is based on one MIL-STD-1553B bus for the data communications between the OBMU and the platform units, and on a SpaceWire network. 23)
SpaceWire has been selected as the sole communication interface between each of the instruments and the spacecraft DHS (Data Handling Subsystem). It also provides a key interface within the DHS itself, between are the OBC (On-Board Computer ) and the SSMM (Solid State Mass Memory).
Figure 2: SpaceWire network architecture (image credit: Astrium, ESA)
AOCS (Attitude and Orbit Control Subsystem): The AOCS employs an autonomous star tracker, gyros, and sun sensors for attitude acquisition and safe mode sensing; actuation is provided by reaction wheels and thrusters. The AOCS baseline architecture also includes a hard-wired safe mode using a sun sensor and a coarse gyro aimed at recovering as fast as possible the sun-pointed attitude in case of contingency, which is essential for the spacecraft thermal safety.
The AOCS constitutes a suite of components that in close interaction with the rest of the spacecraft controls the orientation and stability of the spacecraft, and executes the ground requested velocity changes for adjustment of the otherwise ballistic trajectory. This function includes the monitoring of its own health, as well as the provision of a reference on selected data related to trajectory and orientation, in order to support control of mechanisms.
A set of primary requirements to the AOCS are:
- Maximum 6.5o off-pointing from the Sun, with maximum 50s off-pointing over 2.3o
- Capacity of fine pointing without star tracker measurements for at least 24 hours
- A fine pointing Absolute Pointing Error of 42 arcsec, with an Attitude Knowledge Error of 25 arcsec. The Pointing Drift Error is specified at 13 arcsec over 24 hours, using 10s integration windows. All figures are applicable to Line of Sight to the Sun, 95% confidence.
The AOCS consists of most of the classical elements found on interplanetary missions, but with the special feature that the onboard computer handles all tasks, such as data handling, thermal control, AOCS and FDIR (Failure Detection, Isolation and Recovery), on a single processing module. The equipment used are two pairs of Fine Sun Sensors, two Inertial Measurement Units, two Star Trackers, four Reaction Wheels, and a redundant bi-propellant propulsion system consisting of 9 thrusters per branch. The Inertial Measurement Units consist of one nominal branch featuring high performance rate measurements from four tetrahedron oriented gyroscopes and a contingency branch providing reduced rate measurement performance. The nominal branch also includes four tetrahedron oriented accelerometer channels. All units are communicated with via two MIL-1553B redundant busses. The units are synchronized to the onboard time reference at a minimum of 8Hz data acquisition, corresponding to the attitude control frequency. 24)
EPS (Electric Power Subsystem): The solar panel design relies on a carbon/carbon substrate with triple junction GaAs cells. The operational temperature of this new solar array technology is expected to be 230oC. On SolO this can be achieved by implementing a large enough OSR (Optical Surface Reflector) ratio and solar array tilt angle such that the sun incidence angle is high enough to limit the incident solar flux. The EPS architecture employs a regulated power bus. One Li-ion battery is foreseen to cover mission needs during LEOP and Venus gravity assists. - The solar arrays can be rotated about their longitudinal axis to avoid overheating when close to the Sun.
The S/C dimensions are: 2.5 m x 3.0 m x 2.5 m. The pointing stability is better than 3 arcsec/15 min. The total spacecraft wet mass is about 1800 kg, the maximum power demand is ~ 1100 W. The payload suite mass budget is ~190 kg with a payload power consumption of 180 -250 W (depending on the mission phase).
Table 2: Overview of Solar Orbiter mission parameters 25)
The Solar Orbiter thermal control is based on using a sun pointed, flat heat shield to limit the sun flux on the spacecraft structure. By using this approach the elements behind the heat shield will be in a more benign thermal environment. 27)
All external components are shielded from direct solar illumination by the heat shield except for the instruments requiring direct view of the sun and the spacecraft appendages, i.e. the solar arrays, the RPW (Radio & Plasma Wave Analyzer) antennas and the HGA (High Gain Antenna). The heat shield is sized to prevent direct solar illumination on any of the shaded components during nominal pointing and for safe mode events of spacecraft off-pointing up to 6.5o from sun-center. However, the spacecraft must also withstand reflected solar flux and high IR flux from appendages outside of the heat shield shadow cone. In addition, the remote sensing instruments will all receive additional IR flux from the feedthroughs which allow them to view through the heat shield.
The design allocates the heatshield at the top of the spacecraft to free all four lateral walls for high efficiency radiators with good viewing factors towards cold space. A key strategy in the restriction of the Solar Orbiter mission cost is to reuse technology from other programs, primarily of course the BepiColombo program given the environmental similarities. The heatshield requirements call for:
• The heatshield must protect the majority of the spacecraft, including the payload, from the punishing incident solar flux (28 kW/m2 at perihelion)
• At the same time the heatshield must incorporate cut-outs to allow the RS (Remote Sensing)-instrumentation, and the sun sensors, access to the sun.
The definition of ‘protection' is that the heatshield will:
• Limit the overall radiative heat flux to the spacecraft to no more than 30 W in total
• Limit the overall conductive heat flux at all attachment points to the spacecraft to no more than 15 W in total.
The technological challenges of the heatshield were addressed through parallel contracts awarded to TAS-I and Airbus DS (former EADS Astrium) with the goal of design and production of thermal breadboards to demonstrate the concepts.
The essential function of the heatshield was identical in both cases. Each heatshield presents a planar surface to the sun, and relies on using multiple layers with large gaps in-between to facilitate lateral heat rejection to cold space. However the two resulting breadboard concepts were different in a number of key aspects:
- Materials: The choice of material for the outer layer (sunshade) is obviously critical as this effectively sets the temperature of the outer layer and the subsequent performance of the entire heatshield. The TAS-I design employed Carbon-Carbon fabric with an additional Nickel light blocking layer; the Astrium design used Keplacoat© on a Titanium foil.
In the meantime, the initial choice – carbon-fiber fabric – was ruled out. Instead the sunshade team began looking for the answer outside the space business. They found it in the shape of Irish company Enbio and its CoBlast technique, originally developed to coat titanium medical implants. The process works for reactive metals like titanium, aluminum and stainless steel, which possess a surface oxide layer. The team sprays the metal surface with abrasive material to grit-blast this layer off; also included is a second ‘dopant' material possessing whatever characteristics are needed. This simultaneously takes the place of the oxide layer being stripped out. The big advantage is that the new layer ends up bonded, rather than only painted or stuck on. It effectively becomes part of the metal.
- Support Panel: The Astrium concept used a separate Aluminum support panel for the heatshield in addition to the +X spacecraft panel upon which it is mounted, which allows the heatshield to be treated as a separate item to the spacecraft (highly desirable for programmatic reasons); this is in contrast to the TAS-I design in which the support panel of the heatshield is part of the spacecraft primary structure.
- Number of gaps: The Astrium design utilized a single gap between the sunshade and the support panel, with an additional gap between the support panel and the +X panel of the spacecraft. In contrast the TAS-I design employed 3 equidistant space gaps between 4 layers.
- Layer support: The Astrium design relied on pretensioned lashes in order to provide stiffness to the sunshade layer and a high degree of planarity (this improves thermal performance). In contrast the TAS-I design favored loose support of the layers by rigid Star Brackets – although the planarity of the layers is reduced, the mechanical performance of this approach is considerably better.
Figure 4: Astrium heatshield design, incorporating 2 lateral layers separated by tensioned Titanium lashes to provide rigidity and a high degree of planarity (image credit: Airbus DS)
Figure 5: TAS-I heatshield design incorporating multiple lateral layers separated by Star Brackets which loosely hold the layers (image credit: TAS-I)
Feedthrough doors and mechanisms: A critical component of the overall heatshield design is the feedthrough and door arrangement that allows the RS-instruments to see through the heatshield. The generic design is applicable for all the RS-instrument feedthroughs: a cylindrical feedthrough with internal vanes to specify the FOV of the instrument. Each feedthrough is mechanically supported by an interface to the support panel of the heatshield, and in turn the feedthroughs provide local support to the sunshade (uppermost) layer of the heatshield through a second interface.
The doors are made of Titanium, with a ‘duck-foot' design incorporating radial spars. The door does not provide any contamination control, it has only a light-blocking function, and consequently does not touch the structure underneath. Instead it is displaced above the feedthrough by ~1 mm, a sizing which ensures non-interaction of the door and feedthrough during launch. The accuracy of the door operation is not critical, as long as the door completely covers the aperture when it is required to do so. A launch lock is present at the door to constrain rotation during launch.
RF communications: The subsystem consists of a redundant set of transponders using X-band for the uplink, and X-band and Ka-band for the downlink. Depending on the mission phases, the transponders can be routed via RF switches to different antennas. The telecommunication subsystem provides hot redundancy for the receiving function and cold redundancy for the transmitting function. One steerable HGA (High Gain Antenna) is being used to support the X-band services for engineering data, and the Ka-band for the science data transmissions.
The X-DST (X-band Deep Space Transponder) is designed and developed by TAS-I (Thales Alenia Space, Italy). The digital platform (whosedesign is inspired by the software-defined radio concept) features a system-on-chip based DSP core, implementing on the same chip all the X-DST signal processing algorithms. 29)
Figure 6: Block diagram of the RF communications system (image credit: TAS-I)
The operations concept is such that the instrument data will be stored in a SSMM (Solid State Mass Memory), for later downlink during daily ground station passes of 8 hours. The science data is downlinked in X-band via the high gain antenna. During the 10 day science windows, the allocation for the nominal average data generation rate of the full payload is 120 kbit/s. This is also controlled via an allocation of the average per instrument. For the remote sensing instruments in particular, their allocation is insufficient to downlink the full raw data and therefore their designs are such as to allow pre-processing, data reduction, selection and associated internal data storage in order to ensure that optimum use is made of the TM bandwidth to downlink the best data. This is not only important for each instrument individually, but for the mission as a whole, as the overriding science objectives rely on combining observations of the same phenomenon from different instruments.
Thermal architecture of spacecraft:
The TCS (Thermal Control Subsystem) of the spacecraft represents the main design challenge, a critical element for spacecraft integrity and performance for a large proportion of the mission duration. The fundamental Solar Orbiter thermal requirement stipulates that the TCS will support payload and spacecraft subsystems such that it is designed to withstand all thermal environments encountered during the entire life of the mission. The selected approach is to rely on a sun-pointed spacecraft with the spacecraft protected from solar flux by the heatshield, and on specific technologies for the remaining exposed parts, such as the solar panels, communication antennas, and the heatshield. 30)
The heat rejection efficiency of the heatshield permits a quasi-decoupling of the spacecraft body from the direct sun irradiation (flux density of up to 28 kW/m2 at 0.2 AU).The heatshield is made with a highly reflecting/emissive external layer to dissipate the incident flux as much as possible radiatively.
Figure 7: Schematic of thermal architecture (image credit: EADS Astrium)
Several payload instrument apertures are implemented through the heatshield to let the remote sensing instruments observe the sun through baffles, and acquire the incident rays on their sensitive detectors. The instruments are either mounted directly on spacecraft lateral walls (in-situ instruments), and use dissipation transferred from the base plate of the unit to the external radiator, or mounted on the spacecraft shear walls (remote-sensing instruments) and use a conductive link from the instruments to the radiators viewing cold space, mounted on external walls, or use dedicated fluid loop pipes. Other radiators accommodated on the lateral walls of the spacecraft are used to cool down internal equipment that dissipate heat or receive solar flux (Figure 8).
The heatshield itself is an innovative and the most sophisticated piece of hardware on SolO. A flat heatshield design is selected and accommodated on top of the spacecraft whose side is always facing to the sun. The heatshield is supported by a structure decoupled from the spacecraft. This structure carries the remote sensing instrument baffles. The baffles cannot be supported by the spacecraft wall since they contain high temperature points or regions. The load-carrying structure is thermally decoupled from the spacecraft wall to minimize conduction loads. The mechanically autonomous heatshield design with respect to the spacecraft is very user-friendly to all AIV (Assembly, Integration and Verification) activities.
The preliminary design of the heatshield outside reflecting layer consists of a white ceramics coating on a titanium (Ti) plate, with an ?/? (absorption/emission) ratio as low as 0.4 - 0.6 at EOL. A multi-layer concept made of polished Ti foils and VDA/VDA (Vapor Deposited Aluminum) kapton foils is proposed for the next layer of insulation to efficiently dissipate the heat and maintain the spacecraft wall at room temperature.
The different layers are held through regularly spaced Ti stand-offs made with limited conductivity towards an Al honeycomb structure to which they are attached. This plate acts as the support structure of the heatshield and is mounted onto the spacecraft wall through a few stand-offs, with a classical MLI (kapton + Dacron) in-between for insulation.
Figure 8: Conceptual design of the heatshield structure (image credit: EADS Astrium)
The interfaces between the remote sensing instruments and the heatshield mainly comprise the baffles and instrument shutters aimed at protecting them from contamination and solar flux when they are not operated. The baseline concept is to thermally decouple the baffle from the instrument by attaching it to the heatshield support structure, and to dissipate their heat by conductive coupling through a radiator installed at the edge of the heatshield. Baffles of optical instruments are assumed to be in SiC, while baffles of particle detection instruments (SWA) could be in the same material as the heatshield first layer in order to lower their temperature.
Figure 9: Overview of the TCS (image credit: Astrium Ltd., ESA)
Figure 10: Illustration of the deployed Solar Orbiter spacecraft (image credit: Airbus DS, ESA)
• March 16, 2015: The Structural and Thermal Model (STM) for the Solar Orbiter mission will leave the Airbus Defence and Space premises in Stevenage (UK) towards the end of March for mechanical testing at IABG in Ottobrunn, Germany. This test is a crucial stage in the development of the Solar Orbiter. At the end of this three-month test campaign, the STM will be shipped back to Stevenage for further building before being shipped again to IABG for thermal testing. 31) 32) 33)
Figure 11: Photo of the STM during testing at Airbus Defence and Space (image credit: UKSA, Max Alexander)
• On March 13, 2015, the Solar Orbiter passed a major milestone when the heat shield was attached to the engineering model (Figure 12, left). MSSL (Mullard Space Science Laboratory) of UCL (University College London) has a significant involvement in the Solar Orbiter mission on both the remote sensing and in-situ sides. On the remote sensing side, the electronics box for the EUI (Extreme Ultraviolet Imager) is being built at MSSL, while on the in-situ side, MSSL is the PI (Principle Investigator) institute for the SWA (Solar Wind Analyzer) suite of instruments. 34)
Figure 12: The photo shows the heat shield on the left and on the right, Prof. Louise Harra (EUI Co-PI) and Prof. Chris Owen (SWA PI) at Airbus DS in Stevenage to see the Solar Orbiter heat shield being attached (image credit: UCL/MSSL)
• Fall 2014: Six instruments have passed CDRs: SolOHI, MAG, SPICE, EUI, PHI and EPD. The STIX and SWA CDRs were performed and awaiting outcome. Two CDRs (RPW, METIS) will follow. Various instrument STM (Structural/Thermal Model) items have been delivered to prime contractor facilities and integrated on the STM spacecraft. 35)
• June 3, 2014: ESA's Solar Orbiter mission has undergone its latest major test: its protective shield has been subjected to concentrated sunlight to prove it can cope with the fierce temperatures close in to our parent star. The outcome ensures it will balance solar illumination, the cold of deep space and internal heat sources to maintain the perfect operating temperature. 36)
• May 2014: Members of ESA's Solar Orbiter team watch expectantly as an essential part of the spacecraft is lowered into Europe's largest vacuum chamber: the multi-layered shield (of size 3.1 m x 2.4 m) that will protect their probe from the Sun's remorseless glare. 37)
- ENBIO, a surface treatment technology enterprise based on the Belfield Campus in Nova UCD (University College Dublin, Ireland), was awarded the contract to coat the main heatshield for ESA's Solar Orbiter mission. After undergoing and completing an extreme test process at the ESA/ESTEC in the Netherlands, the ENBIO solution was approved for use on flight hardware. The solution combines some old and new thinking: a pigment used in 30,000-year-old cave paintings and ENBIO's patented CoBlast process. 38)
- To provide such a system, ENBIO has been collaborating with ESA and Airbus Defence & Space since 2011, to develop a novel protective CoBlast Skin, called SolarBlack. It is critical that the Skin maintains its thermo-optical properties, despite years of exposure to extreme infrared and ultraviolet radiation, while not shedding material or outgassing vapor, which would risk contaminating Solar Orbiter's highly sensitive instruments. Additionally, the Skin needs to be conductive to avoid the build-up of static charge which might threaten a disruptive or destructive discharge to the craft. 39)
- SolarBlack is a CoBlast Skin of black calcium phosphate, which will be applied to the outermost titanium sheet of Solar Orbiter's multi-layered heatshield. It will be deployed via ENBIO's patented CoBlast process, which replaces, in one process step, a metal's natural oxide surface layer with a desired functional Skin – in this case SolarBlack. What makes CoBlast unique is the direct bond produced between the desired Skin and the underlying metal, without a troublesome oxide layer, providing the durability and adhesion required for skin integrity under such extreme conditions. CoBlast is also an environmentally friendly process, requiring no chemical, vacuum or thermal inputs. - SolarBlack has been qualified to meet the demands of this mission and is being specified on an increasingly wide variety of additional applications including sensor internals and heatshields.
Figure 13: On May 2, 2014, the engineering model of the sunshield, sandwiched together from multiple layers of titanium and outermost carbon coating, was placed into the 15 m-high and 10 m-diameter Large Space Simulator at ESA/ESTEC (image credit: ESA)
Legend to Figure 13: The Large Space Simulator is Europe's largest vacuum chamber for the sunshield's trial by sunlight. As its crucial test begins, all air will be extracted to produce space-quality vacuum, while the chamber walls are pumped with –190°C liquid nitrogen to mimic the extreme cold of deep space. - Then the light from 19 xenon lamps, each consuming 25 kW, will be tightly focused by mirrors into a concentrated beam of artificial sunlight upon the sunshield for a number of days.
• The CDR (Critical Design Review) for the spacecraft and the payload started in September 2013.
Launch: The Solar Orbiter spacecraft is projected to be launched in October 2018 by a NASA-provided EELV (Evolved Expendable Launch Vehicle) from KSC (Kennedy Space Center), Cape Canaveral, FL, USA.
- The launch of Solar Orbiter is now planned to take place in October 2018. The launch was previously targeted for July 2017. The decision to postpone the launch was taken in order to ensure that all of the spacecraft's scientific goals will be achieved, with all the system's components adequately tested prior to sending the spacecraft to the launch site. 40)
- In March 2014, NASA selected ULS (United Launch Services) LLC of Centennial, CO, to launch the Solar Orbiter Collaboration (SOC) mission to study the sun in July 2017. The Solar Orbiter will launch on an Atlas V 411 rocket from Space Launch Complex 41 at Cape Canaveral Air Force Station, Florida. 41)
Orbit: The Solar Orbiter will use Venus gravity assists to obtain the high inclinations reaching 35o with respect to the sun's equator (inclined ecliptic orbit) at the end of the cruise phase mission (the cruise phase will last about 3.4 years).
Using SEPM (Solar Electric Propulsion Module) in conjunction with multiple planetary swing-by maneuvers, it will take the Solar Orbiter only two years to reach a perihelion of 45 solar radii with an orbital period of 149 days. Within the nominal 5 year mission phase, the Solar Orbiter will perform several swing-by manoeuvres at Venus, in order to increase the inclination of the orbital plane to 30o with respect to the solar equator. During an extended mission phase of about two years, the inclination will be further increased to 38o.
- Elliptical orbit around the Sun with a perihelion as low as 0.28 AU and with increasing inclination up to more than 30o with respect to the solar equator.
- Aphelion between 0.8 AU and 0.9 AU
- Co-rotation pass: duration 10 days, with a maximum drift of 50o
- Period about 150 days
- Inclination evolving from 0o-30o (with respect to solar equator), 34o in the extended mission.
Figure 14: Solar Orbiter trajectory to orbit the Sun (image credit: ESA)
Figure 15: January 2017 launch: solar distance (image credit: ESA) 42)
Figure 16: January 2017 launch: solar latitude (image credit: ESA)
From its launch early in 2017, the Solar Orbiter will reach the nominal orbit around the Sun in 2020, operating in its near-Sun environment for at least 6 years, including the extended mission phase. During this period, the spacecraft will carry the science payload through 14 perihelion passages. At the same time, the heliocentric latitude will be gradually increased through repeated Venus gravity assist maneuvers, providing information about the behavior of the Sun at high latitudes.
The instruments have been selected jointly by ESA and NASA as part of the collaboration to provide the in situ and remote observations. The sensor complement consists of six remote sensing instruments operating at wavelength ranges from visible to X-ray, as well as four in-situ instruments covering all attributes of the interplanetary medium. It will acquire simultaneous spectra and images of the photosphere and corona; images of the photospheric magnetic field and gas velocity as well as measurements of the magnetic field and in-situ plasma properties at the location of the spacecraft. 43)
The challenging nature of the Solar Orbiter mission, along with tough constraints in the area of volume, mass and data rates, have led to a range of innovative design solutions for the payload, involving new technologies never previously used in space. In the thermal and optical domains, heat rejecting windows, limiting the bulk of the solar flux while allowing the wavelength(s) of interest to pass through, are being developed for two instruments. Several multi-layer coatings are also being designed for internal lenses and mirrors. For the three deployable antennas, the harsh thermal environment has had an important influence on the design of the deployment mechanism. For the purpose of polarization measurements, newly space qualified Liquid Crystal Variable Retarder technology is being applied and several instruments are also making advances in detector design, including newly designed CdTe X-ray detectors and back-illuminated Extreme UV CMOS detectors.
The remote sensing instruments opto-mechanical assemblies are mounted at the periphery of the main structure, by means of isostatic mounts thermally insulating the instruments from the rest of the spacecraft. This technique permits a direct FOV for the instruments and their baffle to cold space behind the heatshield, while keeping the main structure unaffected by the thermal control of the instruments; hence, insuring the very stable thermoelastic behavior needed for the line of sight (LOS) co-alignment.
Table 3: Summary of payload pointing and accommodation requirements
Figure 17: Payload accommodation onboard the Solar Orbiter (image credit: ESA) 44)
Figure 18: Mounting locations of some instruments on the SolO spacecraft (SolO consortium)
Although the Solar Orbiter payload is based on heritage from previous solar science missions, the very nature of this mission, along with its constraints, has triggered considerable innovation covering the full range of technical domains (Ref. 43).
ESA spearheaded the investigation of promising new technologies which were considered to hold benefits for the Solar Orbiter payload via several activities in the early phase of the mission (Phase A/B). These included TDAs (Technology Development Activities) for the concept of the PHI HREW (Heat Rejecting Entrance Window), developed under contract with Selex Galileo, and the LCVRs, developed under contract with INTA. Further examples include early UV detector radiation testing and the DRPM (Dynamically Reconfigurable Processing Module) study, which demonstrated the feasibility of the concept of routine partial reconfiguration of the Virtex-4 FPGAs as foreseen in the PHI design.
The instrument teams were involved from the early stages of the requirements specifications for these contracts and were able to follow the progress throughout. Following the end of the activity, responsibility for further development of these items was handed over to the instrument team with reduced risk.
Common procurement: Although the instruments are diverse, wherever possible a common approach to new technologies has been encouraged and facilitated. This is exemplified by the establishment of a CPPA (Central Parts Procurement Agency), which was established by ESA on behalf of the instrument teams for procurement of EEE (Electronic, Electrical and Electromechanical) parts. The service provided to the instrument teams includes the qualification of the EEE components themselves. This provides the opportunity for harmonization and reduces the cost to the instruments. It also lowers the schedule risk associated to qualification at mission level by coordinating procurement milestones and providing visibility to ESA of the qualification progress.
A) Entrance windows and filters:
Due to the high incident flux, the remote sensing instruments, which require a direct view of the Sun, must reject the majority of the incident energy, while allowing the wavelength range of interest to pass through. This is done in a variety of ways. PHI and STIX have windows located in the spacecraft heat-shield, EUI has entrance filters inside the instrument, while SPICE and the chronographs METIS and SoLOHI have internal mirror systems.
Optical Heat Rejecting Entrance Windows (PHI): Of these the most technologically challenging is the PHI Heat Rejecting Entrance Window (HREW), due to the very narrow requirement for the PHI science wavelength (617.3 nm ± 1.5A), the large aperture (140 mm for the HRT) and the need to minimize the thermal flux entering the instrument. Although a filtergraph inside the instrument filters the science wavelength to the required level, it is a delicate component which cannot tolerate the full solar load and thus the GREW is the first stage in this filtering process and restricts the incident light to a ~20nm band, allowing only 4% of the total incident energy through.
The concept of the PHI GREW consists of a series of coatings on a glass substrate (Sprawl 300), which filter the wavelength of interest in stages (see Figure 3). In addition to the strict requirements on wavelength pass band, as PHI is an optical polarimeter, there are strict requirements on the polarization, uniformity and wavefront distortion induced by the HREW.
These coatings must be shown to maintain their properties steadily throughout the lifetime of the mission under all operational conditions. This has been validated via a series of environment tests, both at sample and prototype level, involving thermal cycling and radiation tests. A further challenge has been to design a suitable mount for the HREW, which will withstand the vibrational loads while protecting the glass and ensuring that the mount does not disrupt the optical properties of the HREW. The current design has been qualified to 275°C and shown to survive and maintain its optical and thermal performance in all tests, with the exception of a simultaneous radiation test at high temperature in vacuum, which is yet to be performed. In sample level tests, the coated glass has been shown to survive up to a temperature of 350°C without degradation.
Figure 19: Schematic of PHI Heat Rejecting Entrance Window and its functional concept (image credit: ESA)
X-ray Windows (STIX): Although it's requirements are not as stringent as those of PHI, STIX needs a window which will be transparent to X-rays (above 4 keV), while minimizing the transmission of lower wavelength ranges (being opaque at wavelengths >300 nm). Due to the nature of the STIX instrument, which involves the generation of Moire shadowgrams on the detectors, the uniformity of the X-ray transmission is crucial, with a requirement of <4% rms variation in the transmission over the aperture (~200 mm diameter). These requirements lead to the need for a Be window with a thickness accuracy of 25 µm. The design involves two windows, one at the front and one at the back of the feed-through in the spacecraft heat-shield.
In order to lower the overall temperature of the window, a protective thermal coating of Al-SiOx is being qualified for use on the sun facing side of each window (Figure 20). For the STIX windows, along with the qualification of the coating, the processes and facilities for manufacture of the window are critical due to the nature of Be dust as a toxic material. Therefore, additional safety precautions are taken in planning and executing qualification tests, such as vibration testing, to ensure sufficient margins of safety against breaking the window.
Figure 20: STIX X-ray window coating (image credit: ESA)
Heat rejecting mirror system (SPICE): The SPICE optics unit (Figure 21) is composed of two parts: the telescope section and the spectrometer section. The primary purpose of the telescope is to reflect and focus as much as possible of the EUV solar radiation in the entire spectral range of SPICE onto the spectrograph entrance slits, while rejecting the unused solar flux (UV/visible/IR).
Figure 21: Illustration of the SPICE optics unit (image credit: ESA)
This is accomplished in several steps:
1) The "Solar Transparent" Primary Mirror (PM) is designed with a thin (~10 nm) boron carbide (B4C) coating applied on a fused silica substrate. This provides a VUV (Vacuum Ultraviolet) reflectance of >0.27 between 40 nm and 200 nm. This allows to reflect the EUV radiation of interest for science towards the spectrograph entrance slits.
2) Most of the solar visible and near-infrared radiation is transmitted with little absorption by the PM to the back of the instrument where it is reflected by the HRM (Heat Rejection Mirror) towards outer space. The HRM is a highly reflective vacuum deposited silver fold mirror accommodated inside a CFRP structure ("chimney"). The other internal surfaces of the chimney are uncoated and provide some radiative cooling due to their view to space.
3) The undesired the solar radiation reflected by the PM is reflected to a single heat dump by another set of mirrors in front of the slit. The flux is routed from this heat dump to S/C radiators. These mirrors are configured so that only the required science beam passes through to the slit. Baffles also intercept solar radiation that either diverges as it comes into the instrument or is off-axis due to spacecraft pointing.
Out of the 31.8 W solar load that enters the instrument (end-of-life hot operational case), 22.7 W are rejected by the HRM while only 9.1 W are absorbed internally by the instrument and by the heat dump, as illustrated in Figure 6.
Figure 22: SPICE thermal inputs and outputs (SDM: SPICE Door Mechanism; PM: Primary Mirror; HRM: Heat Rejection Mirror), image credit: ESA
EUV Entrance Filters (EUI): In order to filter the EUV science wavelength of interest for EUI, which is in a narrow passband of 15-31nm, entrance filters for each channel are located at the front of the instrument. These consist of 150 nm thick aluminum filters. The filters receive almost the full bulk of the solar flux (the 47.4 mm High Resolution Imager EUV filter will receive 17.44 kW/m2). While commercial space-qualified filters are available, these have not been qualified to the temperatures and incident solar flux levels which will be experienced by EUI. In addition, due to the very thin nature of these filters, a supporting mesh structure is required to survive the launch loads, however the mesh support in the standard commercial filters do not evacuate the IR load sufficiently.
Developments have therefore been undertaken to customize the filter design, introducing different Ni mesh patterns and also encapsulating the filter in a ribbed frame (Figure 23). Several permutations of the mesh and frame patterns and materials have been trialled to find the optimum combination of transmission performance, thermal conductivity and response to thermal cycling under high temperature and in vacuum. CSL have also investigated a prototype in which the supporting grid and frame are grown simultaneously on the Al filter ("grid-on-filter"). These investigations have led to a convergence on a custom design involving a spiral Al frame/rib pattern matching that of the underlying Ni mesh (Figure 23). This has been tested successfully to 13 SC.
Figure 23: Illustration of the EUV entrance filter (image credit: ESA)
B) Polarizers and other optics:
LCVR (Liquid Crystal Variable Retarder) technology introduction, (PHI and METIS): The polarization measurements in the PHI and METIS instruments are performed by means of LCVRs, a technology which has considerable heritage in ground- and balloon-borne instruments but has not yet flown in space. These polarizers consist of LC (Liquid Crystal) cells, which are electrooptical polarization modulators. By adjusting the voltage applied to the cells, the orientation of the LC molecules is changed and this changes the optical retardance (Figure 24). In the case of PHI, two LCVR cells will be oriented with their fast axes at 45° to each other as part of the PMP (Polarization Modulator Package), to enable differential retardance of the different polarization states. 45)
The advantages of this technology for space applications derive from the low resource demands (power, mass and volume) as well as the fact that it avoids the use of mechanisms and is easy to control and to synchronize with the detector readout.
The qualification activities performed for these cells included the investigation of several material types as well as the design of a mount assembly. The optimized design successfully passed mechanical, thermal and radiation tests such that TRL (Technology Readiness Level) 5 was reached in mid-2011.
The PMP for PHI consists of two APAN (Anti-Parallel Nematic) LCVRs oriented with their fast axes at 45° with respect to each other followed by a linear polarizer (the polarizing beam-splitter, the polarization analyzer) aligned with the fast axis of the first LCVR. The PMP generates four modulations of the polarization state in order to extract the Stokes parameters of the solar incoming light.
For METIS, only one LCVR is necessary in the PMP with a quarter-waveplate (not included in the PMP, but placed previously) in order to analyze the linear polarization. Nevertheless, an additional LCVR has been included with its fast axis parallel to the first cell, but with the pretilt angles of the liquid crystal molecules in opposite direction to obtain an extended and wider acceptance angle.
Obviously, in both instruments, a linear polarizer is included before the detector to be able to analyze the signal as change in the detected intensity.
Figure 24: LCVRs variable retardance with applied voltage (image credit: ESA)
Etalon (PHI): The fine filtering of the PHI science wavelength to the required narrow passband (optimum bandwidth in the range 80-120 mA) is performed in a Filtergraph consisting of Lithium Niobate (LiNbO3) electro-optic etalon. This is a single crystal with a clear aperture of 50 mm and fabrication finesse of ? 30. The etalon needs to be tuned in order to scan the spectral line as well as to compensate for the spacecraft radial velocity and this is performed by adjusting an applied high voltage. The temperature sensitivity of such an etalon is 2.9 pm/K, which means that the etalon must be kept in a temperature controlled oven, set to a temperature which can be maintained throughout the operational orbit.
As in the case of the LCVRs, the etalon wafers are commercially available and have been used on balloon-borne instruments, however they have not as yet flown on a space mission. A qualification for the Solar Orbiter is being undertaken, of which vibration testing, vacuum testing and proton irradiation testing up to 60 krad on a structurally representative sample have so far been performed successfully.
C) Detectors and Electronics:
Back-illuminated CMOS detectors: The detectors for each of the three channels of the EUI (Extreme UV Imager) instrument (Table 3) are based on CMOS APS technology. Prototype 1 k x 1 k front-thinned, double-gain detectors (APSOLUTE) have been developed for EUI, in parallel to similar 2 k x 2 k single-gain prototype development for PHI. These have successfully undergone testing to demonstrate suitability for the Solar Orbiter environment.
However further improvements on these designs are needed in order to fulfil the EUI science objectives:
- While front-thinned detectors are suitable for the EUI Lyman-? channel, the EUV channels require a new development of back-thinned detectors. This is needed in order to allow back-illumination of the detectors to achieve the required sensitivity in the EUV 10-40 nm range (Figure 25)
- Double gain, as developed in the APSOLUTE prototypes is needed in order to minimize the read-out noise and increase the speed and dynamic range.
- For the Full Sun Imager, the array size has to be increased to 3 k x 3 k in order to achieve the required FOV and angular resolution.
Each of these new developments presents its own challenges and full qualification of the new design for Solar Orbiter must be performed. This activity is currently underway.
Figure 25: Back-thinning of APS detectors (image credit: ESA)
IDeF-X ASIC (STIX and EPD): The IDeF-X (Imaging Detector Front-end) ASIC is a new development in Front End Electronics which may be flown for the first time on Solar Orbiter in the STIX and EPD (Energetic Particle Detector) instruments. STIX will use the IDeF-X HD (High-Dynamic) version, while the IDeF-X BD (Bi-directional) version will be used in EPD-STEP.
For STIX, the ASIC is integrated in the Caliste-SO detectors, which are hybrid components integrating a pixelated 10 mm x 10 mm CdTe sensor with its HV power supply, the analog front end read-out electronics in the form of the IDeF-X HD, passive components and a 20 pin surface mount interface (SOP interface). The µPCBs are stacked perpendicular to the CdTe sensor in the Caliste-SO body (Figure 26).
The IDef-X performs the read-out of up to 32 channels in terms of pixel location, energy and rise time. The design involves a CSA (Charge Sensitive Amplifier) stage followed by a pulse shaper, with several tunable parameters, including CSA bias, peaking time and low level discriminator. Different (partial) channel read-out modes can be configured in order to optimize power consumption.
The advantages of this ASIC are primarily in terms of savings on volume, mass and power. The power consumption per channel is 800 µW. As the detectors have to be cooled in order to minimize the noise and leakage current, savings in power consumption and thus heat dissipation are very beneficial.
The characterization of the Caliste-SO prototype has shown its ability to meet the STIX performance requirements and full qualification of this device is underway. 46)
Figure 26: Illustration of the STIX Caliste-SO hybrid (image credit: ESA)
Virtex-4 FPGA (PHI): The full data processing chain of PHI requires to perform an inversion of the RTE (Radiative Transfer Equation) to derive the magnetic vectors and line of sight velocity. Due to the limited telemetry bandwidth on Solar Orbiter, this must be performed onboard as there is insufficient bandwidth to downlink the raw PHI data. Such onboard processing can only be performed by an FPGA such as the Xilinx Virtex-4, while remaining within the instrument power budget. PHI are thus planning to use two Virtex-4 CF1140 devices (1140 pins) in their DPU (Data Processing Unit).
While the Virtex-4 component itself has been qualified for use on space missions, there is currently no assembly house in Europe which is space-qualified for soldering the CF1140 package of this component. Thus, the PHI project is undertaking a dedicated qualification program for the assembly of this device, involving the manufacture of dedicated verification boards which will undergo vibration and thermal cycling tests to simulate the mechanical strain to the solder joints which may be experienced during the mission.
D) On-board data processing:
Routine FPGA reconfiguration (PHI): In addition to the processing required for the RTE inversion, PHI require considerable on-board processing capability for instrument functionality such as control of the Image Stabilization System and data pre-processing.
To optimize mass, volume and power, the PHI project has developed a design making use of the ability to re-program the Virtex-4 FPGAs to routinely switch the FPGA code in order to perform different functions at different points in the orbit, as follows (Figure 27):
- In one configuration, FPGA1 performs ISS (Image Stabilization System) control and FPGA2 performs the data accumulation.
- In the second configuration, FPGA1 performs the RTE inversion and FPGA2 performs the data pre-processing.
While reprogrammable FPGA technology has been flown in space on several missions, the PHI architecture involving routine changes of the FPGA code throughout the mission is novel and relies on the robustness of not only of the FPGA code protection but also that of the SoCWire network between the components. This is also under verification as part of the ESA DRPM (Dynamically Reconfigurable Processing Module) study.
Figure 27: PHI use of routine FPGA reconfiguration (image credit: ESA)
E) Deployment Mechanisms:
Antenna deployment mechanism (RPW): The RPW (Radio and Plasma Wave) antenna (ANT) system consists of three identical antennas which are oriented at angles of roughly 120o in the operational configuration. Each antenna comprises a 5 m long sensor and an almost 1 m long boom that keeps the sensor at a distance from the spacecraft to reduce disturbances and supports the preamplifier close to the sensor. Due to the antenna length it is obvious that the antennas need to be launched in stowed configuration and deployed in orbit. As most of the length of the antennas sticks out from the heat shield, material selection must take into account direct sun illumination.
The general antenna design is based on several pipe segments connected by "Maeva" elastic hinges, the deployment of which is controlled by SMA (Shape Memory Alloy) components. In stowed configuration each antenna is fixed by three HDRMs (Hold Down and Release Mechanisms) which are opened in sequence to release the antenna. For each HDRM, two stowing flanges are guided by journal bearings and driven to their open position by springs, once a Frangibolt non-pyrotechnical device (SMA based) is actuated. The stowed antenna assemblies are protected by MLI shells in order to keep the temperature of the SMAs in the Frangibolts and hinges below the transformation temperature, thus avoiding premature and uncontrolled deployment. The MLI shells are mounted on carbon tubes that are attached to the HRDM flanges.
Figure 28: RPW antenna overview in stowed configuration (image credit: ESA)
The Maeva hinges, controlled by the SMA components, were developed by CNES. The Maeva hinges constitute Carpentier elastic blades that ensure both deployment and locking of the hinges. To control the deployment, each hinge is equipped with two SMA strings around which Cerafil heater wires are wound. For each antenna, all SMA heaters are electrically connected in series.
Figure 29: Deployment hinges (image credit: ESA)
Mechanical, thermal and functional (in particular deployment) tests are being performed on full antenna models and subsets. High temperature materials are also being qualified separately.
Solar remote-sensing instrument package:
Figure 30: Solar Orbiter payload summary and locations on the spacecraft (image credit: Airbus DS) 47)
PHI (Polarimetric and Helioseismic Imager):
The PHI instrument is a diffraction limited, wavelength tunable, quasi-monochromatic, polarization sensitive imager which is being developed by a consortium with major contributions from institutes in Germany, Spain and France, and smaller contributions from Sweden, Norway and Switzerland.
The objective of PHI is to measure the magnetic vector and the LOS (Line-of-Sight) velocity fields in the solar photosphere. It will thus probe the deepest layers of the Sun (including the solar interior using helioseismology) of all the instruments on Solar Orbiter or the Inner Heliospheric Sentinels. Since the magnetic field anchored at the solar surface produces most of the structures and energetic events in the upper solar atmosphere and significantly influences the heliosphere, PHI plays a key role in reaching the science goals of Solar Orbiter. Extrapolations of the magnetic field observed by PHI into the Sun's upper atmosphere and heliosphere will provide the information needed for other optical and in-situ instruments to analyze and understand the data recorded by them in a proper physical context. 48) 49) 50)
The instrument concept uses two telescopes:
• HRT (High Resolution Telescope) with an aperture diameter of 125-160 mm which achieves an angular resolution of 1.00 arcsec (150 km on the Sun at minimum perihelion distance of 0.22 AU). Its FOV (Field of View) will be 1000 x 1000 arcsec. The off-axis Ritchey-Chrétien HRT will image a fraction of the solar disk at a resolution reaching 150 km at perihelion (the same resolution as the Extreme Ultraviolet Imager's high resolution channels will have).
• FDT (Full Disk Telescope) with an aperture diameter of 15 mm providing an image of 2k x 2k pixels. The refractor FDT will be able to image the full solar disk at all phases of the orbit. It incorporates an off-pointing capability. 51) 52)
Each telescope will have its own PMP (Polarization Modulation Package) located early in the optical path in order to minimize polarization cross-talk effects. Polarimetry at a signal to noise level of 103 is baselined for PHI. The HRT and the FDT will sequentially send light to a Fabry-Perot filtergraph system (~ 100 mA spectral resolution) and on to a 2048 x 2048 pixel CMOS detector. PHI will have its own ISS (Image Stabilization System ) that will compensate spacecraft jitter or other disturbances. This system will be composed of a limb sensor and separate rapid tip-tilt mirrors for the FDT and the HRT.
Figure 31: Conceptual view of the PHI instrument (image credit: MPS)
MPS (Max Planck Institute for Solar System Research) of Lindau, Germany has system responsibility for the implementation of the PHI instrument. This comprises the development of a stable housing and its connections to the spacecraft's main body. Due to the close solar proximity during Solar Orbiter's perihelion passages, all instrument components are subject to stringent conditions regarding radiation hardness and thermal stability.
The MPS is responsible for the full optomechanical development of the HRT (High Resolution Telescope). The prospective optical concept is based on an off-axis Ritchey-Cretien telescope with an aperture diameter of 160 mm and an effective focal length of approximately 2500 mm. An off-axis design is preferred since it causes reduced thermal stabilization complexities which pose the main challenges of the instrumental design. In addition, a feed select mechanism for choosing between the two telescopes will be developed under authority of the MPS. In addition, MPS provides the focal plane assembly of the PHI instrument. This comprises the development of a fast 2 k x 2 k APS-detector system and the associated front-end electronics which can be qualified for the space mission.
The Spanish institute IAC (Instituto de Astrofísica de Canarias) is a partner of MPS. The consortium includes several other Spanish and German laboratories and the CNRS/IAS (Institut d'Astrophysique Spatiale) of France.
PHI will allow high-resolution and full-disk measurements of the photospheric vector magnetic field and line-of-sight velocity, as well as the continuum intensity in the visible wavelength range. For these objectives the instrument will consist of two telescopes, one to image a fraction of the solar disk at high spatial resolution and the second to image the full solar disk at all phases of the orbit. Both telescopes will be used sequentially to carry out narrowband measurements. Figure 32 shows the Solar Orbiter orbit with the radial velocity (i.e. relative to the Sun) as a function of the heliocentric distance. 53)
Figure 32: Solar Orbiter radial velocity as a function of the heliocentric distance time, for the January 2017 launch. The nominal, cruise and extended phases are represented by black, grey and blue curves, respectively (image credit: IAS)
FG (Filtergraph): PHI will carry out the measurements observing the Doppler- and Zeeman-effects in the FeI 6173 A absorption line (sensitive to magnetic field). This line will be scanned with a narrowband FG. At different spectral positions, the polarization state of the incoming light will be analyzed. The FG will provide a tuning range of ±0.6 A to follow the wavelength shift produced by the spacecraft radial velocity (i.e. up to 26 km/s along the orbit), plus the range required to scan the absorption line and the continuum. It will be tuned at six wavelength positions plus one for the continuum (Ref. 53).
The FG has been designed by IAS using the optical design provided by MPS. The design is based on previous experiences and includes a solid Fabry-Perot etalon and a narrowband prefilter. It extracts a spectral portion of the FeI 6173 A absorption line and a nearby continuum point. To achieve this objective, it will select a passband of 0.1 A. The etalon is made of Lithium Niobate (LiNbO3). This type of etalon is smaller and lighter than the piezo-stabilized ones. The temperature stabilization is very important since the tuning of the etalon can be lost as a consequence of temperature variations. Given that the performance of the etalon is optimum at normal incidence, the FG has a telecentric optical design. A prefilter blocks the secondary transmission peaks from the etalon. It is formed by two components: PF1 and PF2. The first is the narrow passband and the second one is the blocker for all the spectral range. The combination of both filters achieves the maximum transmission inside the passband and the maximum out-of-band rejection. PF1 has a passband width less than 3 A. It is very sensitive to the temperature and to the AOI (Angle of Incidence).
Figure 33: Photo of assembled FG. The first telecentric lens is seen on top (image credit: IAS)
As it can be seen from Figure 32, the heliocentric distance and the radial velocity will vary significantly along the observations. The distance to the Sun will go from 0.28 AU to 0.9 AU and will have a direct effect on the instrument temperature, therefore the FG is designed to keep the etalon and prefilter at a constant temperature. Figure 33 shows the picture of the FG prototype completely assembled. All the components, with the exception of the lenses are placed inside a thermally-controlled oven.
The fine tuning of the FG is essential for the instrument scientific exploitation. The tuning will rely on the ground characterization and on the on-board calibration. Both constitute an important milestone in the development of the instrument. The ground characterization is currently being performed at IAS.
The FG characterization includes optical, electrical and mechanical tests. Its objective is to assess the spectral response, the image quality, the stability and the operation life. It will be used to build the look-up tables included in the on-board calibration program. These tables are the optical transmission maps as a function of wavelength and the spectral response as a function of the tuning voltage. As summarized in Table 4, each optical component is being tested separately and, once integrated in the FG as a complete system. The tests include measuring the transmission profile all along the useful surface, the AOI sensitivity, the thermal gradients across the useful surface, the tuning stability and the thermal stability. These measurements have been ranged according to their priority and complexity. They are done in air and in vacuum conditions. A special setup has been prepared for the tests in a 10000-class clean room and in a solar telescope (i.e. the Tour Solaire at Meudon, France).
Table 4: Characterization tests
The test setup for these tests includes one "in-air" chamber and three vacuum chambers of different sizes, all of them with temperature control. Different light sources are in use: Sun light, monochromatic and white light. One CCD and two calibrated photodiodes are used for the visible and the near-infrared measurements. In addition, the test setup comprises several high-precision mechanisms to align and tilt the components. The spectrometers in the test setup have a resolving power (R) of 15 x 103 for the broadband measurements and 3 x 105 for the narrowband ones.
Figure 34: FG passband scan obtained from the convolution (black plus) of the solar spectrum (orange), the theoretical etalon transmission profile (red line) and the prefilter measured transmission profile (green). The secondary Fabry-Perot passband is shown as a red dotted line (image credit: IAS)
Figure 34 shows the measured PF1 narrowband transmission and the simulation of the spectral scan that will be obtained by the FG. It is the result from the convolution of the solar spectrum with the prefilter measurement and with the etalon profile being tuned step by step in the wavelength range shown. Figure 34 shows the dependency on the prefilter transmission profile, thus the importance to measure it with high precision. Currently, the prototype of the prefilter has been measured to 0.06% (1? error) with the spectroscopic test setup at the solar telescope.
The PHI instrument has a mass of 34 kg, power consumption of ~31 W, and a date rate of 20 kbit/s.
Figure 35: Schematic view of the PHI instrument and its components (image credit: MPS)
SPICE (Spectral Imaging of the Coronal Environment):
SPICE is an EUS (Extreme-Ultraviolet Spectrograph) imager for the Solar Orbiter mission. The SPICE instrument was selected selected as a NASA contribution to SolO mission under the PI-ship of SwRI (Southwest Research Institute). This instrument is designed to make observations in the far and extreme ultraviolet and will provide plasma diagnostics of the solar atmosphere where the temperature ranges from tens of thousands to several million degrees.
MPS (Max Planck Institute for Solar System Research) of Lindau, Germany is actively participating in the design of the SPICE instrument. The main MPS contribution is the design and development of the primary telescope mirror. The primary mirror plays a key role in the optical and thermal design of the SPICE spectrograph. The off-axis mirror with a parabolic figure will have a novel reflective coating of boron carbide that will be efficient in the extreme ultraviolet but let the solar visible and thermal radiation pass largely unaffected. In this way the heat load on the mirror is reduced. 54)
The scientific goal is to determine the plasma density, temperature, element/ion abundances, flow speeds and the structure of the solar atmosphere using spectroscopic observations of emission lines in the UV/EUV spectral region. Spectroscopic observations of emission lines in the UV/EUV region provide important plasma diagnostics of the solar atmosphere, providing the necessary tools for probing the wide range of solar plasma temperatures. These may range from tens of thousands to several million K. The analysis of emission lines, mainly from trace elements in the sun's atmosphere, provides information on plasma density, temperature, element/ion abundances, flow speeds and the structure and evolution of atmospheric phenomena.
The SPICE instrument consists of a single element off-axis parabolic telescope and a TVLS (Toroidal Variable Line Spaced) grating spectrograph with two IAPS (Intensified Active Pixel Sensor) detectors. SPICE also includes a DPU to control each of the mechanisms, perform data compression and provide the SpaceWire interface to the spacecraft. The off-axis parabola mirror forms an image of the Sun onto the entrance slit assembly containing three interchangeable slits of differing widths. The slit selects a portion of the solar image and passes it to a concave TVLS grating which re-images the spectrally dispersed radiation onto two array detectors. Beyond 0.35 AU, off-limb observations are made by inserting a quartz filter to reduce UV scattered light in the instrument and allow observations of the outer corona beyond >0.30 RS. The two spectral passbands cover the same spatial field of view simultaneously with no scanning of the detectors or grating. The detectors are solar blind, IAPS sensors, and require no visible light rejection filters. The stigmatic spectra produced are magnified, yet maintain high spectral resolution in one dimension and high spatial resolution in the other. The SPICE observing strategy is to produce 2D spectro-heliograms (spectral images) of selected line profiles and line intensities only. The wavelengths covered by SPICE are 702-792 A (Band 1), 972-1050 A (Band 2) and 485-525 A (2nd order). The selected lines represent the full range of temperatures and heights in the solar atmosphere, from the chromosphere to the flaring corona. SPICE derives heritage from SOHO/CDS, SOHO/SUMER, as well as the RAISE and EUNIS sounding rocket programs.
Figure 36: Conceptual view of the SPICE instrument (image credit: MPS, SwRI)
EUI (Extreme UV Imager):
The exceptional orbits of the Solar Orbiter mission permit to achieve breakthroughs in plasma astrophysics by imaging the magnetized solar atmosphere in the extreme-ultraviolet (EUV) emission.
• To provide EUV imagery with at least a factor 2 higher spatial resolution than currently available, to reveal the fine-scale structure of coronal features
• To provide full-disc EUV imagery of the sun to reveal the global structure and irradiance of inaccessible regions such as the "far side" of the sun and the polar regions
• To study the connection between in-situ and remote-sensing observations.
The EUI instrument suite on board of Solar Orbiter is composed of two HRI (High Resolution Imagers), one at Lyman ? and one dual-band in the extreme UV, and one dual-band FSI (Full Sun Imager) working alternatively at the two 174 Ä and 304 Ä EUV passbands. 58)
The EUI instrument is developed in a collaboration which includes CLS (Centre Spatial de Liege) and Royal Observatory of Belgium, Belgium, the IAS (Institut d'Astrophysique Spatiale) and Institut d'Optique, France, the UCL Mullard Space Science Laboratory (MSSL), London, UK, and MPS (Max Planck Institute for Solar System Research), Katlenburg-Lindau, Germany.
The quiescent, large-scale corona can be imaged in the spectral bandpass centered at 174 A and dominated by Fe IX/X emission lines formed at temperatures around 0.8–1.1 MK (million Kelvin). The lower transition region can be imaged in the 304 A bandpass dominated by the He II emission line formed at a temperature around 0.08 MK (or lower). Simultaneity of images in two bandpasses is not required. Both bandpasses can thus be combined inside a single telescope. The telescope aperture of 0.5 cm allows achieving acceptable signal-to-noise ratios with exposure times between 1 - 10 s allowing a cadence of around 10 minutes in each bandpass (to allow the observations of solar eruptive events such as CMEs and flares).
Figure 37 shows the EUI optical unit with the main components of the three channels. The spectral selection is obtained by multilayer coating deposited on the mirrors and by a set of transmission filters rejecting the visible and infrared light. The detectors feature a 2 k x 2 k array for the HRI and a 3 k x 3 k array for the FSI channel. Back-thinned silicon CMOS-APS with 10 µm pixel pitch, sensitive in the EUV wavelength range are considered for the FSI and the HRIEUV channels. A front side illuminated CMOS-APS is considered for the HRILy? channel (Ref. 58).
Figure 37: EUI optical bench configuration (image credit: CSL)
Figure 38: Functional block diagram of FSI and HRIs (image credit: CLS, IAS, MSSL, MPS)
Table 5: Main characteristics of the EUI channels
METIS (Multi-Element Telescope for Imaging and Spectroscopy)
The METIS instrument suite is conceived to perform both EUV (Extreme Ultraviolet) spectroscopy of the solar disk and off-limb, and near-sun coronagraphy and spectroscopy in FUV. METIS observations are crucial for answering some fundamental solar physics questions concerning the origins of the fast and slow wind, the sources of solar energetic particles, and the eruption and early evolution of coronal mass ejections.
1) The origin of heating/acceleration of the solar wind streams
2) The origin , acceleration and transport of the solar energetic particles
3) The transient ejection of coronal mass and its evolution in the inner heliosphere.
The METIS instrument suite consists of three different elements, sharing the same optical bench: electronics, power supply, and heat shield aperture. In particular METIS will provide for the first time:
- simultaneous imaging of the full corona in polarized visible light (590-650 nm) and narrow-band ultraviolet HI Lyman ? (121.6 nm line)
- monochromatic imaging of the full corona in the extreme ultraviolet HeII Lyman ? (30.4 nm line)
- spectroscopic observations of HeII Lyman ? in corona.
The annular FOV (Field of View) ranges between 1.4 and 3.0 solar radii, when the perihelion is 0.28 AU, and the attained spatial resolution is 20 arcsec.
These measurements will allow a complete characterization of the three most important plasma components of the corona and the solar wind (electrons, protons, helium).
The three elements are identified as:
• COR (VIS and EUV Coronagraph)
• EUS (EUV Spectrometer)
• SOCS (Solar Orbiter Coronal Spectrometer).
The payload includes two core instrument packages, optimized to meet the solar and heliospheric science objectives.
METIS consists of a single optical head which uses a single aperture on the spacecraft sun facing thermal shield. The optical head is mounted on an optical bench (Figure 39), plus a main electronics and power supply box. The instrument front end consists of the visible-light and UV/EUV coronagraphic imager, which comprises the optics, detectors, proximity electronics and electrical interface. Internally to this coronagraph, in a suitable position, a dispersion grating intercepts a small portion of the solar corona (so that the obtained corona image is actually not complete) to perform UV and EUV spectroscopy. This additional channel does not need any mechanism, and shares one of the imaging coronagraph detectors.
Optical design: METIS is designed around an innovative concept for an externally occulted solar coronagraph, based on an IEO (Inverted External-Occulter). The IEO is a small circular aperture which replaces the classical annular aperture of the standard externally occulted solar coronagraph design. A boom connects the IEO to the M0, a small spherical mirror that rejects back the disk-light through the IEO. Many are the advantages of this novel design with respect to the classical one. Considering its application to this instrument, they can be summarized in the following points:
- smaller external occulter diameter
- thermal load on M0 greatly reduced
- on-axis telescope configuration
- more compact, cylindrical structure.
The smaller external occulter diameter implies a smaller aperture on the S/C thermal shield; the reduced thermal load produces a lower temperature inside the instrument and a better stability and control of the optical bench; the on-axis configuration gives a better optical performance, and has the advantage of a simpler mechanical structure; the compactness of the structure allows to optimize the available resources, and the cylindrical structure gives a symmetric configuration and an easier baffling, which is always an extremely critical point for solar coronagraphy. Beyond M0, an on-axis annular shape Gregorian telescope focuses the solar corona on the focal plane assembly. The Gregorian configuration of the primary and the secondary mirror, respectively M1 and M2, gives access to the primary focal plane for the placement of the IO (Internal Occulter). The IO blocks the light diffracted by the edge of the IEO and reflected by M1.
Figure 39: 3D view of the METIS layout (image credit: METIS consortium)
A conceptual scheme of the METIS coronal imager is shown in Figure 40. This is a novel design for an externally occulted all-reflective coronagraph, in which the external occulter has also the function of system entrance pupil. In such a way, the size of the external aperture is minimized and the thermal flux entering the instrument is greatly reduced, thus allowing to meet the stringent thermal requirements of SO. This novel optical configuration takes the name of ICOR (Inverted Coronagraph) because the diaphragms are inverted with respect to a standard coronagraph design.
Figure 40: Optical layout of the ICOR instrument (only UV-EUV channel), image credit: METIS consortium
The UV/EUV and VL detectors are, respectively, an intensified active pixel sensor (IAPS) with scale factor 20 arcsec/pixel and image size: 30.7 mm (1024 x 1024) with 30 µm equivalent pixel size; and an APS with scale factor 10 arcsec/pixel and image size: 20.5 mm (2048 x 2048) with 10 µm pixel size.
Table 6: METIS performance parameters
As shown in Figure 40, the first optical element is the IEO (Inverted External Occulter), which is in this design, a small circular aperture (40 mm O) in the Solar Orbiter thermal shield. Coronal light entering the IEO is then collected by a simple on-axis Gregorian telescope (aspherical mirrors M1 and M2), which makes the coronal image on the telescope focal plane where the detector is located. An IO (Inverted Occulter) is located close to the telescope prime focus with the function of blocking the light diffracted by the edges of the IEO. When METIS is pointing at the Sun's center, direct disk light impinges on a small (69 mm O) spherical mirror (M0) which backreflects it through the IEO. The portion between the IEO and M0 is called the "boom" and consists, optically, of three stops: IEO, on the front face of the S/C heat shield, and the SEA (Shield Entrance Aperture), on the back face of the S/C heat shield, and the annular aperture deliminated internally by M0 and externally by the hole inside M2.
Figure 41: Optomechanical structure of METIS (image credit: METIS consortium, Ref. 63)
Some of the advantages resulting from this scheme with respect to the classical external occulter design are the following:
- Smaller diameter boom (~3 times) through the S/C thermal shield: this implies a smaller aperture on the S/C heat shield and an improved mechanical stability
- The thermal load on M0 is greatly reduced (by more than 90%) with respect to a standard externally occulted configuration: this means that there is a lower temperature inside the instrument, with a consequent easier control of the optical bench
- On-axis telescope configuration, which yields a better optical performance.
The telescope mirrors are coated with a suitable multilayer coating, optimized to enhance the reflectivity in the EUV He line, but still having good reflectivity at 121.6 nm and in the visible range. Hence, it is possible, by means of dedicated bandpass filters, to obtain the solar corona imaging at the three mentioned different wavelength bands.
A Lyot stop ,positioned slightly after the IO, blocks the light diffracted by M0 and reflected by M1. A filter wheel positioned just in front of the UV detector allows to select either visible and/or UV/EUV (HI 121.6 nm or HeII 30.4 nm line) images of the corona. When the thin aluminum filter is in the path, only HeII observations are performed. When the Al/MgF2 interference filter is in the path, the HI light is transmitted to the UV detector, while the broadband visible light (VL) is reflected towards the liquid crystal polarimeter, in order to perform measurements of the linear polarization of the visible solar corona.
To improve the scientific return of this instrument, a spectroscopic channel has also been included in the METIS optical path. Essentially, in the prime focus of the Cassegrain telescope, a three slit system is located in correspondence of an equatorial region of the solar corona; this slit system inhibits the possibility of doing imaging in this portion of the corona, so the actual coronal images of METIS will have a small sector missing. Light passing through the slits is collected by a diffraction grating located in a sector of the Cassegrain telescope secondary mirror. UV light is then dispersed and focused on the same UV/EUV detector used for imaging.
As an externally occulted coronagraph (Figure 42) the telescope has an occulting system between the Sun and the primary mirror, M1, that puts M1 in the shadow of the Sun disk. The rejection of solar disk radiation is the most critical issue in a coronagraph design: solar disk radiation that impinges on the coronagraph detectors, no matter how, is called stray light. Traditionally, the external occultation concepts consists of an occulting disk that prevents solar disk radiation from reaching the entrance pupil aperture. METIS operates with an inverted occulted system concept: light from the solar disk and from the corona enters a round external aperture that acts as the entrance pupil and is called IEO (Inverted External Occulter). Solar disk light is then blocked and rejected through IEO by a Sun disk rejecting mirror, M0. This configuration has the advantage, very important for this mission, of reducing the thermal load inside the instrument. In order to reduce the amount of stray light, the METIS Sun disk radiation is rejected by means of standard techniques used in externally occulted coronagraphs:
• a Sun disk rejection mirror, M0, that reflects back the disk radiation through the entrance pupil
• an internal occulter, IO, that is placed in the conjugate plane of the IEO relative to the primary mirror M1, that blocks light diffracted by the edge of IEO
• a Lyot stop, LS, placed in the conjugate plane of M0 relative to M1, that blocks secondary diffracted light from the edge of M0.
The stray light sources inside the coronagraph are due to two main contributors:
1) Sun disk light diffracted by the edges of the entrance aperture IEO Light diffracted by IEO is mainly taken care of by the internal occulter, IO, and by the Lyot stop, LS. Nevertheless, diffracted light impinging on the primary mirror, M1, is partially scattered by its surface and becomes the major contributor to the final stray light level.
2) Sun disk light entering the entrance aperture IEO Disk light entering IEO is taken care by the disk light rejection mirror, M0, that re-images the Sun disk back through IEO itself. Scattering from the M0 surface will contribute to the stray light level via diffusion off the telescope components: mounts and walls.
Additional contributions to the stray light will be produced by particulate contamination on those components illuminated directly by the Sun (IEO and M0) or illuminated by the light diffracted by IEO (M1, mounts and coronagraph walls).
An extensive theoretical analysis shows that the stray light levels in all three METIS wavelength bands meet the requirements. The analysis is based on an ideal instrument, i.e., it does not include manufacturing and alignment tolerances and also the cleanliness of the optical components, and includes approximations on the shapes on the occulters.
In order to characterize the coronagraph stray light, it is necessary to build a laboratory model of the stray light rejection subsystem and test it inside a special facility with the capability of reproduce as closely as possible the solar irradiance conditions on the coronagraph. Laboratory tests are important to better characterize the occulting system and find and improve its limits and defects (Ref. 62). Table 7 summarizes the sources of stray light and how METIS is dealing with them and Figure 43 shows the comparison between the stray light estimate at 0.28 AU of S/C heliospheric distance with the expected coronal signal.
Table 7: Stray light sources and METIS stray light rejection
METIS in-flight off-pointing:
The Solar Orbiter mission profile foresees three 10 day length observing periods per orbit for the remote sensing instruments which are demanding in terms of resources, included telemetry. Each period is called RSW (Remote Sensing Window) and includes the spacecraft pass through perihelion and maximum and minimum heliolatitude.
Outside the RSWs, the RS instruments do not operate and the METIS door will be closed. During the RSWs, the Solar Orbiter will be sun-center pointed with the exception of planned off-pointing windows to allow high resolution instruments to point targets on the disk up to the limb, and of S/C maneuvers such as wheel off-loading, high gain antenna maneuvers, breaks in fine guidance polynomials.
METIS, as a coronagraph, is designed to operate sun-centered. When the spacecraft is off-pointed, METIS can operate with degraded performance up to a limit of off-pointing which is a function of the instrument parameters and of the heliocentric distance of the Solar Orbiter.
Two limit angles, ?max and ßmax, are defined, respectively, as the maximum angle of off-pointing at which, with degraded performance, it is possible to acquire scientific data, and as the maximum angle at which safety procedure are needed to protect the instrument.
As shown in Figure 44, ?max is the angle at which the solar disk radiation starts to illuminate the lateral surface of the IEO cone, while ßmax is the angle at which M0 does not block the full disk light. Figure 44 provides also the numerical values of ?max and ßmax, while Figure 45 shows the dependence of these angles from the S/C heliocentric distance. Both angles take into account the APE (Absolute Pointing Error) of both the S/C and METIS. The Figure 45 displays also the size of the solar disk as a function of distance, showing that about 0.55 AU, the maximum allowed off-pointing exceeds the size of the Sun.
Figure 44: Definition of the angles ?max and ßmax (image credit: METIS consortium)
Figure 45: The angles ?max and ßmax as a function of the S/C heliocentric distance, with the size of the Sun disk (image credit: METIS consortium)
METIS operating as a COR (Coronagraph):
Co-rotation during the helio-synchronous phases of the orbit will freeze coronal structures in the plane of the sky for many days. This allows the study of the evolution of the magnetic configuration of streamers to test the hypothesis of magnetic reconnection as one of the main processes leading to the formation of the slow solar wind. The out-of-ecliptic vantage point will also allow a unique view of the plasma distribution and solar wind expansion in the coronal low-latitude/equatorial belt. Therefore, it will be possible to measure the longitudinal extent of coronal streamers and coronal mass ejections. These parameters, that at present are unknown, are essential to determining the magnetic flux carried by plasmoids and coronal mass ejections in the heliosphere.
The prime objectives of COR are:
• To investigate the evolution of the magnetic configuration of streamers in order to test the hypothesis of magnetic reconnection as one of the main processes leading to the formation of the slow solar wind during the quasi helio-synchronous phases of the orbit
• To measure the longitudinal extent of coronal streamers and coronal mass ejections from an out-of-ecliptic advantage point. These data are essential to determine the magnetic flux carried by plasmoids and coronal mass ejections in the heliosphere
• To investigate the large-scale structure of the F-corona (the dust) and the cometary sources of the dust near the Sun. This will provide important information for the in-situ instruments in the payload that measure plasma and dust.
Instrument concept: COR is an externally occulted telescope designed for broad-band polarization imaging of the visible K-corona and for narrow-band imaging of the UV corona in the H I Lyman-?, 121.6 nm, line in an annular field of view between 1.2 and 3.5 solar radii, when the Solar Orbiter perihelion is 0.22 AU.
STIX (Spectrometer / Telescope for imaging X-rays)
STIX is a state-of-the art X-ray imaging telescope for solar observations designed and built at FHNW (Fachhochschule Nordwestschweiz, Windisch - University of Applied Sciences and Arts Northwestern Switzerland). STIX provides imaging spectroscopy of solar thermal and non-thermal X-ray emission, measurements at the highest ever spatial resolution and sensitivity. The PI (Principal Investigator) of STIX is Arnold Benz of ETH Zürich. 66) 67) 68)
Figure 46: Illustration of the STIX instrument (image credit: STIX consortium)
Bursts of hard X-rays (> 20 keV) are the most common signature of the impulsive phase of a solar flare. In fact, the X-ray continuum is the most direct signature of energetic electrons at the Sun. The X-rays are bremsstrahlung, produced by accelerated electrons colliding with the ambient solar atmosphere. STIX is an X-ray imaging spectrometer, operating from 3 to 150 keV, which determines the location of X-ray emission from the sun as a function of time and energy to a spatial precision of 1 arcsec over a 38 arcmin imaging FOV and 2 arcmin over a wider 5o FOV.
The prime scientific objectives of STIX are: To establish the timing, location and spectra of energetic electrons near the sun. This will enable these electrons to be related to subsequent observations by the in-situ solar energetic particle and radio instruments. In this way, STIX serves as a high-energy link between imaging and in-situ observations. In conjunction with in situ instruments onboard both SolO and SPP, STIX will analyze the magnetic link between the heliospheric electron populations and their host regions of injection, close to the Sun. Additional STIX science objectives are: 69)
1) To determine the size and morphology of hot (>10 MK, formed by solar flaring activity) thermal plasma and non-thermal hard X-ray sources on the Sun with a spatial resolution of 400 km at closest perihelion, about 5 times better than previously achieved
2) To compare with co-temporal observations of other space-based instruments in order to measure the directivity of solar X-ray emission
3) To compare observations of partially occulted, "behind-the-limb" solar flares with those of other possible spaceborne instruments in order to isolate weak coronal components of hard X-ray emission from the bright footpoint sources and hence determine the relationship between energetic electrons that lose their energy in the corona and those that impact the solar chromosphere.
Instrument concept: STIX imaging uses the same indirect imaging technique as employed on Yohkoh / HXT. Imaging information is encoded in the relative count rates in separate detector elements located behind pairs of grids, each of which absorbs a distinct directionally-sensitive fraction of the incident flux. The telescope employs a set of 64 sub-collimators, each of which consists of a pair of widely separated, X-ray opaque grids with an X-ray detector element located behind the rear grid. Front and rear grid pairs have identical pitch and orientation, whose choice determines the spatial frequency to be measured.
As was demonstrated by Yohkoh/HXT, the relative count rates of a pair of sub-collimators, one of whose grids is displaced by one quarter of its pitch, can be used to accurately measure both the real and imaginary parts of one Fourier component of the angular distribution of the source. With 64 sub-collimators, the imaging system then measures 32 different Fourier components. This data can then be used to reconstruct the source image, using well-established techniques employed by radio astronomy, Yohkoh/HXT and RHESSI.
Figure 47: Conceptual view of the STIX instrument (image credit: ESA)
The two main parts (i.e., the imager and the electronics box) of STIX are mounted independently on the spacecraft. The imager consists of 32 pairs of X-ray opaque grids mounted in front of 32 solid-state CZT (Cadmium Zinc Telluride) X-ray detectors which are located on the electronics box and make up the spectrometer. 70)
STIX will cover the energy range between 4 keV and 150 keV with an energy dependent resolution of 1 keV to 15 keV. Its finest angular resolution and temporal resolution are 7 arcsec and 0.1 seconds, respectively. The transmission of each grid pair is very sensitive to the direction of incidence of the X-ray flux. The relative count rates of the detectors behind the different subcollimators (different pitches & orientations) encode the spatial information.
Table 8: Imaging performance of STIX
Table 9: Specification of STIX detectors
Figure 48: Illustration of the CZT detector elements (image credit: STIX consortium)
Legend to Figure 49: The attenuators are encircled. They are movable and are automatically inserted between the rear grids and the detectors during large flares to prevent excessive count rates at low X-ray energies. The attenuators will be developed by the Greek team (Ref. 69).
STIX attenuators: The attenuators, or shutters, which will be developed by the Greek team, are a critical component of the STIX instrument (Figure 49). Attenuators are dictated by the substantial dynamical range of incident X-ray fluxes from solar flares: the largest X-ray flares can provide as many as 105 more count rates in X-ray photons compared to those of the smallest micro flare that STIX can detect. For steep flare spectra and for fluxes between 3 - 150 keV, count rates can be higher by factors of 107 to 109. - An additional factor of 20 must be accommodated because of SolO's varying heliocentric distance from successive aphelia to perihelia. The attenuators must be movable, enabled at high count rates and disabled at lower ones to protect the STIX detectors from excessive low-energy signal and at the same time allowing them to retain full sensitivity to small flare events. This will make STIX responsive to the entire dynamical range of solar X-ray flux (Ref. 69).
STIX will include two movable attenuators, mounted on the sunward side of the spectrometer module. Each of them will consist of a circular aluminum sheet with different widths (thick and thin). Both sheets are known to preferentially absorb a known fraction of the low-energy X-ray flux - the flux mostly responsible for the excessive signal - incurring a negligible effect at higher X-ray energies.
The movable attenuators will be complemented by a fixed insertion mechanism that will dictate their motions. The insertion mechanism will be triggered into action by an electronic signal depending on the incident count rate of X-ray photons. The signal will rely on internal logic and will generate an electric pulse without spacecraft-generated commands for the insertion or removal of the shutters. The shutters will be ordered to move one at a time by a software-generated command when the incident X-ray flux exceeds a preset threshold. This software, allowing the STIX processor to command one or both shutters to shift from one fixed position to another to attenuate the low-energy X-rays will be contributed by a Czech team participating to the STIX consortium.
STIX science (Ref. 66):
1) Field line tracking: STIX X-ray observations, in combination with other SolO instruments, in particular radio and >keV electrons measurements, will be used to magnetically link the heliospheric region where SolO is making in-situ observations back to regions at the Sun. This directly addresses a prime SolO science question: "What are the origins of the solar wind streams and the heliospheric magnetic field?"
2) X-ray imaging spectroscopy: STIX measurements of the characteristics of the energetic electrons at the Sun can be directly compared to EPD's (Energetic Particle Detector) measurements of the electrons in-situ at SolO, to address another prime science question: "What are the sources, acceleration mechanisms, and transport processes of solar energetic particles?"
• Stereo observations between STIX and other X-ray instruments
- Direct measurement of source height (by comparing positions)
- Directivity of electrons (by comparing intensities)
- Isolate and characterize coronal contributions and relate to footpoint emission (difficult to do with normal imaging)
3) X-rays from CMEs (Coronal Mass Ejections):
• X-ray observation of disk-occulted events: STIX's high sensitivity measurements of X-rays from the very high coronal sources associated with fast CMEs can help to address another prime question: "How do coronal mass ejections evolve in the inner heliosphere?"
• HXR (Hard X-Ray) emission from electrons in magnetic structures related to coronal mass ejections.
4) High sensitivity studies:
• Effective area ~ 0.25 RHESSI › ~ 5 times RHESSI at 0.23 AU
• Lower background than RHESSI › 2.5 times smaller events can be seen
• Up to 15 times more sensitive to faint emission
- HXR micro/nanoflare observations (coronal heating)
- HXR emissions from quiet Sun
- HXR emissions related to CMEs (in occulted events)
- HXR emission from escaping electron beams
Table 10: Parameters of the STIX instrument
Figure 50: Imaging in STIX (image credit: STIX consortium)
Figure 51: Illustration of the STIX imager (image credit: STIX consortium)
Figure 52: STIX instrument processing unit (image credit: STIX consortium)
SolOHI (Solar Orbiter Heliospheric Imager)
SolOHI is an NRL (Naval Research Laboratory), Washington D.C. optical telescope which was chosen by ESA in January 2012 to fly on the SolO mission. SolOHI's revolutionary measurements will allow scientists to identify CME's (Coronal Mass Ejections), which are space weather events. These solar eruptions can travel from 120 to 3,500 km/s and have masses greater than a few billion tons. CMEs can affect electromagnetic fields on Earth impacting power lines, satellite communications, and cell phone service. 71)
SolOHI is an evolution of the HI (Heliospheric Imager) developed for the STEREO SECCHI instrument, but with a new light-weight camera system designed to handle the increased radiation levels that are expected to be seen. The FOV (Field of View) of SolOHI is 40o, twice that of SECCHI/HI-1. At perihelion it will have resolution comparable to LASCO/C2 with the C3 FOV is a visible-light telescope that images the light scattered from free electrons in the solar wind plasma. Thus it observes solar wind structures – streamers or plasma sheets, CMEs, density fluctuations, comets, etc. 72) 73)
SolOHI's detector is a 4 k x 4 k APS (Advanced Pixel Sensor) using CMOS technology. This will be the first time such a large format APS detector is flown. The instrument will make high-resolution images of the corona and solar wind, including CMEs, to determine how they propagate and interact with the background solar wind. With its large FOV, SolOHI will be able to connect the remote sensing observations of the corona to the plasma being measured in-situ at the spacecraft.
The instrument will incorporate a mosaic of SRI-designed complementary metal-oxide semiconductor (CMOS) image sensors. The full-flight SolOHI focal plane will use four of these 2k x 2k side-buttable imagers providing a 4 k x 4 k, or 16 megapixel, format. This will be the first flight for such a large-format CMOS detector. 74)
Figure 53: Conceptual view of the SolOHI instrument (image credit: NRL)
Table 11: Overview of SolOHI instrument parameters
Figure 54: Photo of the SolOHI instrument baffle assembly qualification model (image credit: ESA, NRL)
SolOHI is funded by NASA under its Living with a Star Program, which is designed to understand how and why the sun varies, how planetary systems respond and the effect on human space and Earth activities. NASA's Goddard Space Flight Center in Greenbelt, MD, manages the program for the agency's Heliophysics Division of NASA's Science Mission Directorate.
Heliospheric in-situ instrument package:
SWA (Solar Wind Plasma Analyzer)
The SWA investigation is a major international hardware collaboration, led by UCL/MSSL (PI: Christopher J. Owen). In addition to the overall leadership of the suite, UCL/MSSL will provide the bulk of the hardware for the EAS (Electron Analyzer System), one of the 3 sensor systems within the suite. The PAS (Proton-Alpha Sensor) and the HIS (Heavy Ion Sensor) are led by partners in France and the USA, respectively. The central data processing unit, to be built in Italy, serves all 3 sensors and completes the suite. 75) 76)
The prime scientific objectives of SWA are:
- To provide observational constraints on kinetic plasma properties for a fundamental and detailed theoretical treatment of all aspects of coronal heating
- To investigate charge- and mass-dependent fractionation processes of the solar wind acceleration process in the inner corona
- To correlate comprehensive in-situ plasma analysis and compositional tracer diagnostics with spaceborne and ground-based optical observations of individual stream elements.
In addition, the SWA instrument will enable the investigation of:
- 3He and "unusual" charge states in CME-related flows
- The interaction of solar wind ions on dust grains in the heliocentric distance range associated with the "inner source." Freshly produced pick-up ions from this inner source are specially suited as test particles for studying the dynamics of incorporation of these particles into the solar wind or their further re-energization.
Instrument concept: SWA is a compromise between sensitivity, mass/charge- and mass- and time resolution parameters - due to the limited resources allocated for the instrument. SWA has to cover a large dynamic range in ion fluxes. Since there is an enormous difference between the proton fluxes at perihelion (typically 1014 m-2 s-1) and the fluxes of relevant minor ion tracers at 1 AU (e.g. Fe10+ at typically 108 m-2 s-1 etc.), it is suggested to implement three different sensors:
• PAS (Proton/?-particle Sensor). The objective is to investigate the velocity distribution of the major ionic species at a time resolution equivalent to the ambient proton cyclotron frequency. The sensor is sun pointing. SWA/PAS consists of an electrostatic analyzer with an ion steering (IS) system at the aperture. It is designed to measure the full 3D velocity distribution functions of the major solar wind species, protons and ?-particles, in the energy range ? 0.2–20 keV/q. 77)
The SWA/PAS instrument was built by IRAP (Institut de Recherche en Astrophysique et Planétologie), Toulouse, France with with technical and hardware contributions from UCL/MSSL and Charles University, Prague. Key challenges: These include development of a sensor heat shield to deal with the Sun-facing thermal environment, engineering the structure of the sensor to adequately protect against the 150 krad radiation environment over mission lifetime, and to design the sensor to handle the large dynamic range of the expected ion fluxes.
• EAS (Electron Analyzer System). The instrument consists of two (three optional if resources allow) sensors to cover nearly 4? steradian of viewing space and to allow the determination of the primary moments of the electron velocity distribution with high temporal resolution.
• HIS (Heavy Ion Sensor). The instrument of SwRI (Southwest Research Institute) allows the independent determination of the major charge states of oxygen and iron (up to 9) and a coarse mapping of the three-dimensional velocity distribution of some prominent minor species. Also, pick-up ions of various origins, such as weakly-ionized species (C+, N+, Ne+, Mg+, Si+, etc.), should be measured. The sensor is sun pointing.
Figure 55: Illustration of the SWA/EAS instrument (image credit: UCL/MSSL)
Figure 56: Illustration of the SWA/PAS (Proton Alpha Sensor) instrument (image credit: IRAP)
Figure 57: Illustration of the SWA/HIS (Heavy Ion Sensor) instrument (image credit: SwRI)
RPW (Radio & Plasma Wave Analyzer)
The prime objective of RPW is to provide measurements of both the electric field and magnetic field in a broad frequency band (typically from a fraction of a Hertz up to several tens of MHz) covering characteristic frequencies in the solar corona and interplanetary medium. The measurements of both electrostatic and electromagnetic waves provide different diagnostics:
- Electrostatic waves provide in-situ information in the vicinity of the spacecraft
- Electromagnetic waves provide extensive remote-sensing of energetic phenomena in the solar corona and interplanetary medium.
In addition, the RPW instrument on Solar Orbiter will enable the investigation of:
- Waves and turbulence that occur much closer to the sun than previously measured
- The north-south symmetry of the radio radiation in the solar corona using, for the first time, viewing angles from well out of the ecliptic plane.
Instrument concept: RPW comprises two main sub-systems: PWS (Plasma Waves System) covering in-situ measurements, and the RAD (Radio Astronomy Detector) for remote-sensing.
Principal Investigator of RPA: Milan Maksimovic, LESIA, Observatoire de Paris, France. Collaborating countries (hardware): France, Sweden, Czech Republic, Austria.
MAG will provide in-situ measurements of the heliospheric magnetic field. Scientific topics to be addressed by MAG include the way the Sun's magnetic field links into space and evolves over the solar cycle; how particles are accelerated and propagate around the solar system, including to the Earth; how the corona and solar wind are heated and accelerated. The magnetometer instrument will comprise two sensors, both in shadow and mounted on a boom behind the spacecraft body.
The prime objective of MAG is to provide vector measurements of the solar wind magnetic field with high resolution (better than 1 nT) at sub-second sampling. The MAG instrument will enable the investigation of: 78)
- The link between coronal structures and their signatures in the solar wind
- Kinetic effects in the solar wind plasma
- Large-scale structures in the solar wind, e.g., coronal mass ejections
- MHD waves and turbulence.
The reference MAG instrument consists of dual 3-axis fluxgate sensors mounted on a deployable boom, which is positioned in the shadow of the Orbiter body. One sensor is at the end of the boom and the other at an intermediate distance from the spacecraft body. It measures fields in several gain-ranges, which are automatically selected by the DPU according to the in-situ magnetic field strength. This configuration allows compensation for spacecraft generated stray fields.
Principal Investigator: Tim Horbury, ICSTM, London, United Kingdom. Collaborating countries (hardware): United Kingdom.
EPD (Energetic Particle Detector):
The EPD suite contains five different sensors to measure the composition, timing, and distribution functions of suprathermal and energetic particles, such as, electrons and protons. Principal Investigator: Javier Rodríguez-Pacheco, University of Alcala, Spain. Collaborating countries (hardware): Spain, Germany, USA, ESA.
The prime objectives of EPD are:
- To determine in-situ the generation, storage, release and propagation of different species of solar energetic particles in the inner heliosphere
- To identify the links between magnetic activity and acceleration on the sun of energetic particles, by virtue of combined remote-sensing of their source regions and in-situ measurements of their properties
- To characterize gradual (typically CME-related) and impulsive (typically flare-related) particle events and trace their spatial and temporal evolution near the sun
- To measure energetic pick-up particles originating from the interaction of the solar wind with near-sun dust.
The EPD will determine chemical and charge composition and energy spectra of ions in a wide energy range, from about the typical solar wind energies of a few keV to several 100 MeV/nucleon for protons and heavy ions. Electrons should be measured from 10 keV to 10 MeV. The combination of electrostatic E/Q-analysis with time-of-flight E/M-determination and subsequent direct energy measurement in a solid state detector has been employed in many EPDs in the past and is also a possible design option for the Solar Orbiter.
The EPD suite consists of five sensors measuring electrons, protons, and ions from helium to iron, and operating at partly overlapping energy ranges from 2 keV up to 200 MeV/n. The EPD sensors and a common data processing unit with low voltage power supply are: 79)
1) STEIN (SupraThermal Electrons, Ions, and Neutrals)
2) SIS (SupraThermal Ion Spectrograph)
3) EPT (Electron Proton Telescope)
4) LET (Low Energy Telescope)
5) HET (High Energy Telescope)
6) ICU (Instrument Control Unit).
The EPD sensors share the ICU (Instrument Control Unit) that is composed by the Common Data Processing Unit and the Low Voltage Power Supply (CDPU/LVPS), which is the sole power and data interface of EPD to the spacecraft. 80)
• The Institute for Experimental and Applied Physics (IEAP) of the Christian-Albrechts-University (CAU) in Kiel, Germany, is developing four sensors for the EPD (Energetic Particle Detector): STEIN, EPT, HET, and SIS. The development and manufacturing of STEIN, EPT, and HET is funded through DLR (German Space Agency), and SIS is funded by ESA as a facility instrument.
• The AIP (Astrophysikalisches Institut Potsdam) contributes to the development of two of these sensors: EPT (Electron Proton Telescope) and HET (High Energy Telescope).
Figure 58: Functional block diagram of the EPD instrument (image credit: University of Kiel)
STEIN consists of a single unit having two view cones in opposite directions. SIS consists of two sensor heads with roughly opposite (160Â°) view directions sharing a common electronics box. EPT -HET has multiple view cones sharing a common electronics box. There are two identical EPT-HET units. LET has multiple view cones and consists of two separate identical units.
EPT consists of two dual double-ended telescopes: EPT1 points in the ecliptic plane along the Parker spiral in both the solar and anti-solar directions; EPT2 points 45o out of the ecliptic plane. Each sensor unit consists of a dual double-ended magnet/foil solid state detector particle telescope. One of the major scientific goals of SolO will be the investigation of particle acceleration and propagation at the Sun. Electrons, as measured by the EPT, play a crucial role in the study of energetic processes on the Sun as they provide a direct link to the sites of particle acceleration. Due to the high sensitivity of the EPT it will measure small energetic particle events and even nano flare particles (Ref. 70).
HET covers the high-energy particle range for protons (up to 100 MeV) and heavier ions, thus providing information on the largest solar energetic particle events, which can produce high energy, damaging interplanetary radiation levels. HET has two oppositely directed field of views with 50o full angle and resolves 3Helium and multiple heavy ion species. HET's large collecting power allows fast cadence for high-energy heavy ions. During the study phase, the possibility of obtaining limited information on the neutron flux intensity from HET will be investigated.
HET consists of two sensor double-ended heads, one pointing sun/anti-sunward, the other out of the ecliptic. Thus, HET has a total of four viewing directions. Both HET sensors are identical and consist of a double-ended set of solid state detectors and a high-density calorimeter scintillator. Current trades still involve BGO and GSO as scintillator materials. 81)
Figure 59: Illustration of the EPD instrument (image credit: AIP)
LET (Low Energy Telescope) is designed to identify elements from H to Ni with excellent element and energy resolution over the range from ~1.5 to ~60 MeV/nucleon where SEP spectral breaks typically occur. It will also separate 3He from 4He down to levels of ~1% and resolve Ne and Mg isotopes. In addition, the broad dynamic range will provide measurements of trans-Fe elements with 30 < Z < 83, which are often enriched by factor of ~100 in impulsive SEP events. LET accomplishes these measurements using the well-known E-E method of particle detection in six small stacks of silicon detectors. Using the energy-loss patterns in successive detector layers, the charge, mass, and kinetic energy of species from H to Ni can be uniquely identified onboard at event rates up to ~10 kHz. The LET design has heritage from SOHO/ERNE, which has more than 14 years of successful operation in space. 82)
The basic structure of LET is shown in Figure 60. Two circular front detectors, D1 and D2, define the view cones of the three telescopes. The thickness of the D1 detector is 20 µm, while D2 is 80 µm thick. The thickness of D1 is the main factor determining the lower limit of the operational energy range of LET and it also determines the proton upper energy range due to the decreasing signal with increasing energy. For particle identification, it is required that the particle penetrates the thin foil (e.g., 8 µm polyimide with aluminum coating) above D1, the detector D1 itself, and gives a signal (energy loss of ~150 keV) from D2. Below D2, the three silicon detectors, D3, D4, and D5 have sufficient thickness to stop all particles with energies below the defined higher limit of the operational energy range. At the bottom of each telescope, a detector (AC) is used in anti-coincidence with the other detectors to reject penetrating particles. The front detectors D1 and D2 are both divided in three active areas, a center area and a surrounding annular area with two segments. This division gives angular resolution within the 40o view cone of the three telescopes. More importantly, the field of view for proton and helium measurements in each of the three directions can be dynamically limited to a cone with a full opening angle of 14o, while simultaneously maintaining the full opening angle for heavy ion measurements. This feature allows high-sensitivity measurements of even small solar particle events, while ensuring unsaturated operation during very high flux conditions of large SEP (Solar Energetic Proton) events. The total and center area geometric factors are 0.21 cm2sr and 0.0024 cm2 sr, respectively, giving total geometry factors of 1.26 and 0.014 cm2 sr for six telescopes. The collecting power of the first detector alone, determined by the active surface of D1 and the edge of the collimator structure, is 1.68 cm2 sr.
Figure 60: Structure of the LET instrument (image credit: University of Kiel)
The analyzed data are transferred to the ICU once per second, which allows implementation of a burst mode with 1 second cadence. However, the basic time resolution of LET is 10 s. Proton, 3He, and 4He energy spectra and samples of PHA data are accumulated by the ICU software with this cadence. A longer integration time of 100 s is used for the spectra of heavier ions.
SIS (SupraThermal Ion Spectrograph): SIS identifies particles by time-of-flight (TOF) mass spectrometry. It covers the energy range from ~0.008 MeV/nucleon to 10 MeV/nucleon for He to Fe. SIS consists of two units, each with one particle telescope (sunward and antisunward) located near a single electronics box with power and data interface to the ICU. Particles are detected when they pass through the entrance foil and trigger one of the solid state detectors (SSDs) in the array in the back. Secondary electrons emitted when the ion passes through the entrance-, mid-, and detector foils are accelerated to ~1 kV and directed via isochronous mirrors onto MCPs (Micro-Channel Plates). Fast signals from the MCPs provide Start-1, Start-2, and STOP signals that are analyzed to obtain the time-of-flight measurements for the ions. The SSD measures the residual energy of the ion. By combining the two parameters, and the known flight path of the telescope, the mass of the ion can be identified individually. SIS's high-mass resolution (m/sm ~ 50) allows the project to identify 3He/4He ratios down to <1 %. SIS is now being developed under our European lead in collaboration with Spain's SENER and the Johns Hopkins University Applied Physics Laboratory(JHU/APL) in Laurel, MD, USA.
Figure 61: Energy coverage of the EPD instrument suite (image credit: University of Kiel)
ICU (Instrument Control Unit): ICU is composed by the CDPU (Common Data Processing Unit) and the LVPS (Low Voltage Power Supply). It provides a single point digital interface to the spacecraft and to all EPD sensors (STEIN, HET-EPT, LET and SIS). It also shares information with the STIX, MAG and SWA investigation to allow synchronized high data rate burst-mode operations following on-board identification of predefined triggering events in the EPD data. The ICU is designed to manage sensors control and monitoring, the sensors timing clock, and sensor data collection, compression, and packetization for telemetry. ICU is also responsible of the S/C telecommand reception and delivery, if necessary, to the sensors.
The ICU box is mounted inside the spacecraft body. The box dimensions are 186 mm x 152 mm x 100 mm. The CDPU contains a LEON 2 microprocessor, RAM, EEPROM and PROM memories, two hot redundant SpaceWire interfaces, and six identical serial sensor interfaces. The CDPU PROM contains boot code that can load flight code from either the EEPROM or from the spacecraft interface via telecommand. The EEPROM contains TBD copies of the flight code. The RAM (up to 1 Gbit) is used for code, variables, and buffers. With 1 Gbit memory, at 3.100 Mbit/s download rate, up to four days of data can be stored in this memory. Most of the CDPU RAM is used for a burst memory to save high time resolution interval snapshots from STEIN, HET-EPT, LET and SIS, which are then played out slowly via telemetry. In case of power on/off cycling, the CDPU will reset and resume operations from a defined mode.
The LVPS board is responsible of filtering, monitoring and switching the spacecraft primary power, ranging from 26 -29 V (in accordance with EID-A [AD 1] limits), to the sensors. It also provides the power supply to both CDPUs, nominal and redundant. In the latter case, +3.3 V and +1.5 V are provided. The LVPS will be designed to operate with nominal performance within a power bus voltage of 28 V with variations from 26 -29 V. The ICU includes following protection mechanisms for the sensors:
• Under-voltage protection: The LCLs (Low Capacitive Limit) switches, responsible of switching on/off the sensors, are only enabled if the input power bus voltage is above certain level and are automatically switched off if this power bus voltage drops below it.
• Short-circuit protection: If a short-circuit is produced in the primary power of any sensor, the corresponding LCL limits its current during some milliseconds. After that, this LCL is switched off.
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The information compiled and edited in this article was provided by Herbert J. Kramer from his documentation of: "Observation of the Earth and Its Environment: Survey of Missions and Sensors" (Springer Verlag) as well as many other sources after the publication of the 4th edition in 2002. - Comments and corrections to this article are always welcome for further updates (email@example.com).