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)
Figure 3: Front view of the Solar Orbiter spacecraft configuration with the three RPW antennas, high-gain antenna, instrument boom and solar arrays deployed (image credit: ESA) 26)
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.
Figure 7: Photo of the EM (Engineering Model) deep space transponder (image credit: ESA, TAS) 30)
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. 31)
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
Figure 8: 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 9).
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 α/ε (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.
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 10: Overview of the TCS (image credit: Astrium Ltd., ESA)
Figure 11: Illustration of the deployed Solar Orbiter spacecraft (image credit: Airbus DS, ESA)
Launch: ESA's Solar Orbiter spacecraft was launched on 10 February 2020 (04:03 GMT) by a NASA-provided Atlas-V 411 vehicle of ULA, designated AV-87, from KSC (Kennedy Space Center) SLC-41 (Space Launch Complex), Cape Canaveral, FL, USA. 32) 33) 34)
- Solar Orbiter, an ESA-led mission with strong NASA participation, will provide the first views of the Sun's uncharted polar regions from high-latitudes, giving unprecedented insight into how our parent star works. This important mission will also investigate the Sun-Earth connection, helping us to better understand and predict periods of stormy space weather.
- Signals from the spacecraft were received at New Norcia ground station in Australia at 06:00 CET (Central European Time), following separation from the launcher upper stage in low Earth orbit.
Figure 12: Replay of the launch of the Solar Orbiter spacecraft (video credit: ESA)
Figure 13: The ESA Solar Orbiter team, along with ESA Director of Science Günther Hasinger (eighth from the right), at the launch pad at NASA's Kennedy Space Center in Florida, US, on Sunday 9 February 2020. In the background, the Atlas V 411 rocket that would lift the spacecraft into space several hours later(photo credit: ESA, P. Olivier)
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 maneuvers 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.
During the initial cruise phase, which lasts until November 2021, Solar Orbiter will perform two gravity-assist maneuvers around Venus and one around Earth to alter the spacecraft’s trajectory, guiding it towards the innermost regions of the Solar System. At the same time, Solar Orbiter will acquire in situ data and characterize and calibrate its remote-sensing instruments. The first close solar pass will take place in 2022 at around a third of Earth’s distance from the Sun. 35)
Figure 14: Animation showing the trajectory of Solar Orbiter around the Sun, highlighting the gravity assist maneuvers that will enable the spacecraft to change inclination to observe the Sun from different perspectives (video credit: ESA/ATG medialab)
The spacecraft’s orbit has been chosen to be ‘in resonance’ with Venus, which means that it will return to the planet’s vicinity every few orbits and can again use the planet’s gravity to alter or tilt its orbit. Initially Solar Orbiter will be confined to the same plane as the planets, but each encounter of Venus will increase its orbital inclination. For example, after the 2025 Venus encounter it will make its first solar pass at 17º inclination, increasing to 33º during a proposed mission extension phase, bringing even more of the polar regions into direct view.
Figure 15: Solar Orbiter trajectory to orbit the Sun (image credit: ESA)
Figure 16: January 2017 launch: solar distance (image credit: ESA) 36)
Figure 17: 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.
• February 17, 2020: First measurements by a Solar Orbiter science instrument reached the ground on Thursday 13 February providing a confirmation to the international science teams that the magnetometer on board is in good health following a successful deployment of the spacecraft’s instrument boom. 37)
Figure 18: Data collected with the Magnetometer (MAG) instrument during the deployment of the instrument boom of ESA's Solar Orbiter spacecraft show how the magnetic field decreases from the vicinity of the spacecraft to where the instruments are actually deployed (image credit: ESA)
- Ground controllers at ESOC in Darmstadt, Germany, switched on the magnetometer’s two sensors (one near the end of the boom and the other close to the spacecraft) about 21 hours after liftoff. The instrument recorded data before, during and after the boom’s deployment, allowing the scientists to understand the influence of the spacecraft on measurements in the space environment.
- “We measure magnetic fields thousands of times smaller than those we are familiar with on Earth,” says Tim Horbury of Imperial College London, Principal Investigator for the Magnetometer instrument (MAG). “Even currents in electrical wires make magnetic fields far larger than what we need to measure. That’s why our sensors are on a boom, to keep them away from all the electrical activity inside the spacecraft.”
- “The data we received shows how the magnetic field decreases from the vicinity of the spacecraft to where the instruments are actually deployed,” adds Tim. “This is an independent confirmation that the boom actually deployed and that the instruments will, indeed, provide accurate scientific measurements in the future.”
- As the titanium/carbon-fibre boom stretched out over an overall 30-minute period on Wednesday, almost three days after liftoff, the scientists could observe the level of the magnetic field decrease by about one order of magnitude. While at the beginning they saw mostly the magnetic field of the spacecraft, at the end of the procedure, they got the first glimpse of the significantly weaker magnetic field in the surrounding environment.
- “Measuring before, during, and after the boom deployment helps us to identify and characterize signals that are not linked to the solar wind, such as perturbations coming from the spacecraft platform and other instruments,” says Matthieu Kretzschmar, of Laboratoire de Physique et Chimie de l'Environnement et de l'Espace in Orleans, France, Lead Co-investigator behind another sensor located on the boom, the high frequency magnetometer of the Radio and Plasma Waves instrument (RPW) instrument.
- “The spacecraft underwent extensive testing on ground to measure its magnetic properties in a special simulation facility, but we couldn’t fully test this aspect until now, in space, because the test equipment usually prevents us from reaching the needed very low level of magnetic field fluctuations,” he adds.
• February 14, 2020: At 16:00 CET on Thursday, 13 February, the critical first 83 hours of Solar Orbiter’s unique mission to study our star came to an end. 38)
- “This early phase is like the birth of a child,” says Operations Director Andrea Accomazzo. “Engineers want to be sure that it can survive on its own in its new environment.”
- “In the case of a spacecraft, they need it to be powered by its solar arrays, able to communicate with Earth, and able to control its orientation in space.”
- After a thunderous launch at 05:03 CET on Monday, Solar Orbiter headed south east from Cape Canaveral Air Force Station in Florida, flying over the South Atlantic before being released far above the shores of Western Australia. In Darmstadt, ESA’s mission control team soon took control of the spacecraft and began several days of round-the-clock flight control activities.
- “We planned for years and trained for months for this and I’m proud of all the teams that worked through this critical period,” says Andrea. “The first stage of Solar Orbiter’s mission was a success, but it certainly threw us a few challenging moments!”
Sunny side up
- The solar arrays were deployed at 06:24 CET and unfolded from the spacecraft’s body like wings. They soon began to supply power to the craft, relieving the burden on the batteries.
- With communication established and the necessary tests carried out, the first command was sent to the craft and the mission officially began.
- Until then, Solar Orbiter had still been in its strobing phase – rotating to maximize the chance of pointing one of its two communication antennas at ground stations on Earth. With this command, the spacecraft fired its thrusters and stabilized its orientation.
Figure 19: Visualization showing the deployment of various boom/antennas on the Solar Orbiter spacecraft. Initially, the first Radio and Plasma Waves (RPW) antenna is deployed. Then the boom hosting a suite of scientific instruments is deployed (MAG, RPW, and SWA to measure the magnetic and electric fields, and solar wind around the spacecraft). Subsequently, the remaining two RPW antennas are deployed. Finally, the high gain antenna dish is unfurled. In reality this sequence is spaced out over a 24 hour period (image credit: ESA)
Rise and shine
- The mission control team then began to wake up vital parts of the craft. They switched on the ‘reaction wheels’ – used to more precisely control the orientation of the spacecraft than can be achieved with thrusters – and the ‘Mass Memory’ to store the science data collected by Solar Orbiter’s science instruments.
- “Following successful tests, Solar Orbiter was brought into its Nominal Control Mode,” says Andrea.
- “Then came the window for the deployment of the ‘Radio and Plasma Waves’ (RPW) antennas, the instrument boom, and the high-gain antenna. The antennas and the boom are used to move sensors away from the spacecraft body to prevent disturbances to their measurements, while the high-gain antenna is used to communicate with the spacecraft across astronomical distances.”
Cool under pressure
- Before each antenna or boom could be deployed, Solar Orbiter was tilted to point it towards the Sun and warm it up. After the deployment of the first RPW antenna, the instrument boom was due next.
- But as the team tilted the craft to warm up the instrument boom, they noticed something unexpected.
- “We saw that the pins that hold the doors of the remote observation instruments safely in place during launch were cooling down more rapidly than expected as we tilted them away from the Sun,” says Andrea.
- “If their temperature had fallen to below -40°C, they could have undergone ‘cold welding’, sticking them in place and preventing the doors of the remote sensing instruments from opening.”
- “The pins had already gotten too cold to move, and were in danger of sticking. To prevent this, we sent the command for Solar Orbiter to enter its ‘safe mode’, resetting its orientation and pointing the pins back towards the Sun.”
- “Every launch comes with a unique set of challenges. But our teams train for events like this and were able to quickly respond to this situation and ensure a safe start to Solar Orbiter’s mission.”
Out with the boom
- Once warm enough, the pins were safely moved (it had now proven necessary to move them at an earlier stage than originally planned) and the team continued with the deployment of the boom.
- Following the deployment of the second and third RPW antennas and the high-gain antenna, operations began to wind down as the radio communication link between Solar Orbiter and ESA’s ground stations was switched to the more powerful high-gain antenna.
- “We were thrilled to see that the instrument boom and all three electric antennas were correctly deployed,” says Yannis Zouganelis, ESA’s deputy project scientist for the Solar Orbiter mission.
- “These appendages will enable us to probe the solar wind and, together with the remote-sensing instruments, reveal our Sun and its behavior in unprecedented detail. We can’t wait to start taking our measurements.”
- Solar Orbiter will now spend approximately three months in its commissioning phase, during which the mission control teams will conduct a test maneuver and check that the spacecraft and its 10 scientific instruments are working as intended to achieve the mission’s ambitious goals.
- In roughly two years, Solar Orbiter will reach its primary science orbit where it will study our star’s polar regions like no spacecraft has before.
• February 11, 2020: Solar Orbiter is ESA’s latest mission to study the Sun up close. Launched in the early hours of 10 February from Cape Canaveral, Florida, the spacecraft is due to arrive at its fiery destination in approximately two years. 39)
Figure 20: At this very moment, a spacecraft is headed toward the brightly burning Sun, photographed here on an Antarctic summer day by ESA sponsored medical doctor Stijn Thoolen at Concordia research station (photo credit: ESA/IPEV/PNRA–S. Thoolen)
- Solar Orbiter will face the Sun from within the orbit of Mercury, approximately 42 million kilometers from the solar surface. This is an ideal distance: from here Solar Orbiter can take remote images and measurements that will provide the first views of the Sun’s uncharted polar regions.
- At the southern poles on Earth, in Antarctica, the Sun has an exceptional presence on people living at the remote Concordia research station. During the Antarctic summer, the sun shines 24 hours a day. It would be perfect for sunbathing, except for the fact that the average summer temperature is only –30ºC.
- Consequently, in the winter the Sun does not appear above the horizon for over three months and the crew stationed in Concordia live with outside temperatures of –80ºC in complete darkness.
- While Solar Orbiter is en route to observing the Sun up close, the crew in Concordia are preparing for life without and enjoying the last rays of sunlight while they can. This picture shows a halo that can occur when sunlight is refracted off ice crystals in the atmosphere.
- The mission will investigate how intense radiation and energetic particles being blasted out from the Sun and carried by the solar wind through the Solar System impact our home planet, to better understand and predict periods of stormy ‘space weather’.
- While this results in beautiful aurora seen in the Arctic and Antarctic circles, stormy space weather can be disastrous. Solar storms have the potential to knock out power grids, disrupt air traffic and telecommunications, and endanger space-walking astronauts, for example.
- A better understanding of how our parent star works is critical to our preparedness for these scenarios on Earth.
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. 40)
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.
Figure 21: Payload accommodation onboard the Solar Orbiter (image credit: ESA) 41)
Figure 22: 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. 40).
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 ± 1.5Å), 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 23: 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 24). 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.
Heat rejecting mirror system (SPICE): The SPICE optics unit (Figure 25) 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).
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 26: 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 27). 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 27). This has been tested successfully to 13 SC.
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 28). 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. 42)
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.
Etalon (PHI): The fine filtering of the PHI science wavelength to the required narrow passband (optimum bandwidth in the range 80-120 mÅ) 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-α 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 29)
- 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.
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 30).
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. 43)
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 31):
- 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.
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 32: 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 33: 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 34: Solar Orbiter payload summary and locations on the spacecraft (image credit: Airbus DS) 44)