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Solar Orbiter—Heat shield and system technology
Acta Astronautica 65 (2009) 1076 – 1088 www.elsevier.com/locate/actaastro
Solar Orbiter—Heat shield and system technology J. Poncya,∗ , F. Jubineaua , F. D’Angelob , V. Perottob , J.J. Juilleta a Thales Alenia Space, 100 boulevard du midi, F-06100 Cannes, France b Thales Alenia Space Italia, Turin, Italy
Received 5 December 2007; accepted 9 March 2009 Available online 16 April 2009
Abstract Solar Orbiter will enhance our knowledge of the Sun by observations and in situ measurements as close as 0.22 AU from our star. Placed on an orbit with a period two-thirds the one of Venus, Solar Orbiter will use the many encounters with the planet to gradually incline its orbit and gain view on the Sun’s poles. The permanent in situ observations will be associated to remote-sensing observations over large parts of the orbits. ESA Science Directorate has awarded in parallel two Solar Orbiter Heat Shield and System Technology contracts to industry. This paper presents the achievements of Thales Alenia Space thanks to one of these two ESA contracts. It shows how the main technical challenge brought by the heat flux of 20 solar constants has been addressed by the system and heat shield design. The design and manufacturing of a breadboard of the heat shield in view of thermal test verification is then reported. The main technological developments and residual risks are assessed, paving the way for the definition phase of the program. © 2009 Elsevier Ltd. All rights reserved.
1. Introduction The most important of all stars, our Sun, requires a deeper understanding of its metabolism. The originality of the Solar Orbiter mission lies in its ability to combine in situ measurements in the inner system with closer remote-sensing observations of the Sun and the Corona. However, navigating in the inner system constrains the Abbreviations: AOCS, attitude and orbit control subsystem; FDIR, failure detection identification and recovery; FS, front shield; HGA, high gain antenna; HS, heat shield; HTHB, high temperature heat barrier; IR, infra red; MGA, medium gain antenna; MLI, multi- layer insulation; MPPT, maximum power point tracking; OSR, optical solar reflectors; PDHU, Payload Data Handling Unit; PPS, pulse per second; RF, radio frequency; S/C, SpaceCraft; SMU, Satellite Management Unit ∗ Corresponding author. E-mail address: [email protected] (J. Poncy). 0094-5765/$ - see front matter © 2009 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2009.03.031
mission and spacecraft design, both in terms of large delta-v in the Sun’s gravity well and of the very hostile thermal environment. An Assessment Study was carried out in 2005, with Thales Alenia Space as one of ESA’s two contractors. It identified the heat shield (HS), the high gain antenna (HGA) and the solar array as the main elements requiring anticipation of technological activities. The latter two being covered by ESA’s funded BepiColombo development, ESA subsequently complemented the development effort by awarding two contracts for a Heat Shield and System Technology study [1]. This paper presents the achievements of Thales Alenia Space in the frame of one of these contracts. Thales Alenia Space succeeded in finding technical solutions that limit the technological developments to the heat shield and the instruments doors and baffles. Our achievements increased the degree of maturity of
J. Poncy et al. / Acta Astronautica 65 (2009) 1076 – 1088 Table 1 Baseline mission overview. Timeline
Flight time (years)
Event date (calendar)
Launch Venus GAM-1 Deep Space Maneuver-1 Earth GAM-1 Earth GAM-2 Venus GAM-2 Venus GAM-3 Venus GAM-4
0.0 0.51 1.02 1.38 3.21 3.39 4.62 5.85
2015/05/22 2015/11/26 2016/05/28 2016/10/08 2018/08/08 2018/10/09 2020/01/02 2021/07/08
End of Nominal Mission
6.13
2021/07/08
Venus GAM-5 Venus GAM-6
7.08 8.31
2022/06/19 2023/09/11
End of Extended Mission
8.64
2024/01/31
this complex mission, both in terms of heat shield and of system, enabling to start the definition phase upon ESA’s final decision on Solar Orbiter mission. 2. Solar Orbiter mission By approaching as close as 45 solar radii, the Solar Orbiter will view the solar atmosphere with unprecedented spatial resolution circa 100 km pixel size. Over extended periods Solar Orbiter will deliver images and data of the polar regions and the side of the Sun not visible from Earth. The scientific goals of the Solar Orbiter are ([2]): • to determine in situ the properties and dynamics of plasma, fields and particles in the near-Sun heliosphere, • to survey the fine detail of the Sun’s magnetized atmosphere, using a camera capable of detecting solar features as small as 100 km across, • to identify the links between activity on the Sun’s surface and the resulting evolution of the corona and inner heliosphere, using solar co-rotation passes, • to observe and fully characterize the Sun’s polar regions and equatorial corona from high latitudes. The current baseline trajectory for the mission (direct escape and transfer based on a tour of planetary flybys and a Deep Space Maneuver) is shown in Table 1. The mission will rely on a chemical propulsion system for maneuver performance. The “in situ” set of instruments will be commissioned shortly after LEOP and, nominally, will be continuously switched on, while the
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“remote-sensing” instruments are to be commissioned after the first Earth GAM and operated during certain windows around perihelion passages. Science operations should be possible during the initial part of the trajectory (i.e. before Venus GAM 2); therefore it can be considered that the nominal mission will start shortly after LEOP and initial instrument commissioning. After Venus GAM 2 the spacecraft is injected in a resonant orbit (the period of revolution of the Solar Orbiter around the Sun is 23 the period of Venus) in order to increase the inclination of the orbit with respect to the ecliptic. The following timeline is indicative of the mission and will be confirmed during the course of the Solar Orbiter development phase. During the nominal mission, the Solar Orbiter performance requirements shall be fully met with all specified margins according to the Mission Requirements Document [NR1] for the Solar Orbiter mission. The duration of the Solar Orbiter extended operational lifetime shall be about 2.5 years. Approval of an extension of the mission is dependent on the spacecraft health and status of expendables such as fuel. Alternative mission profiles are considered in case of common launch with NASA’s Solar Sentinels. Thales Alenia Space has studied them along with the nominal mission. 3. Solar Orbiter payload The payload combines in situ instruments (field analyses and particles collectors) with extensive allspectrum remote observation of the Sun and the Corona. Table 2 lists the baseline (core) payload considered for the purpose of this study. The payload complement may be further modified or pared down by ESA in the near future. In addition, some augmentation payload is being traded-off, in particular a Wide-Field Coronagraph or Heliospheric Imager type of instrument. Thales Alenia Space has successfully implemented all the instruments, and the subsequent interface requirements have been inferred. 4. Design drivers and criteria The major design drivers specific to the Solar Orbiter mission are: • The cost and schedule of the mission, leading to privilege recurring units and to reuse the specific developments of BepiColombo, especially on solar array and high gain antenna.
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Table 2 Core payload. Instrument
Acron.
Science goals
In situ instruments Solar Wind Plasma Analyzer
SWA
Radio and Plasma Wave Analyzer Magnetometer Energetic Particle Detector Dust Particle Detector Neutron Gamma ray Detector
RPW MAG EPD DPD NGD
Investigation of kinetic properties and composition (mass and charge states) of solar wind plasma Investigation of radio and plasma waves including coronal and interplanetary emissions Investigation of the solar wind magnetic field Investigation of the origin, acceleration and propagation of solar energetic particles Investigation of the flux, mass and major elemental composition of near-Sun dust Investigation of the characteristics of low energy solar neutrons, and solar flare processes
Remote-sensing instruments Visible Imager and Magnetograph EUV Spectrometer EUV Imager Visible Coronograph Spectrometer Telescope Imaging X-ray
VIM EUS EUI COR STIX
Investigation Investigation Investigation Investigation Investigation
• The close perihelion (0.22 AU in baseline mission scenario), generating high thermal fluxes up to 28,500 W/m2 , high radiations, UV ageing and solar disturbance torques. • The large variation in solar array illumination, from 0.5 to 20 solar constants, with the cruise phase at 1.5 AU sizing the power generation. • The very large data transmission capability required, due to a variable distance between Earth and Spacecraft, generating needs for a powerful radio frequency (RF) system and a large mass memory. • The numerous instruments requiring reliable thermal rejection, crossing of the heat shield, high pointing accuracy, many appendages or fields of view and very stringent cleanliness levels (leading to add doors in many cases). • The launcher compatibility, limiting the mass and volume of the spacecraft: Soyuz launch, with Atlas as a possible alternative. The drivers have been processed during the tradeoffs carried out by Thales Alenia Space, so as to provide for the largest scientific return under the strict necessity of mastered cost and schedule. This resulted in the following criteria for the selection of the technical solutions: • maximize recurrence and control of technical and schedule risks; • control mass and volume; • get robust payload accommodation; • maximize thermal robustness; • minimize interactions between radiators, appendages and fields of view;
of of of of of
the magnetic and velocity fields in the photosphere properties of the solar atmosphere the solar atmosphere using high resolution imaging in the EUV coronal structures using polarized brightness measurements in Visible energetic electrons near the Sun, and solar X-ray emission
• minimize instruments misalignments; • maximize communication link and rate; • maximize uninterrupted observation periods, with minimization of disturbance torques. 5. Solar Orbiter spacecraft design Thales Alenia Space has conducted extensive tradeoffs to select a configuration that best fits these mission drivers. The team included members of TAS Italy (thermal control, heat shield, TTC) and TAS France (system and other subsystems). A special emphasis has been put to provide configurations that combine lowcost low-risk approach, strong reuse of BepiColombo heritage, resilient-schedule modular design and maximum resources for science. The mission is proposed to be fulfilled by a parallelepiped body shadowed by a heat shield. The high gain antenna and solar arrays are managed in a series of configurations that increase the data link and power capabilities and protect the BepiColombo heritage. The satellite carries the 11 primary instruments of the core payload specified for the purposes of this study. 5.1. An optimal configuration In launch configuration (Fig. 1), the satellite is in a compact configuration that controls the loads on the units and minimizes the out-of-plane coupling with the heat shield. The satellite has its South face up. The high gain antenna is stowed on the anti-Sun face on its backside.
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Fig. 1. Solar Orbiter in launch configuration. The compact configuration decreases mass and minimizes loads on heat shield, HGA.
Fig. 3. Solar Orbiter in cruise or science orbit—warm conditions. As the distance to Sun decreases, the solar flux compensates for the tilt of the panels, resulting into an increased power, while keeping the temperature below BepiColombo qualification status.
Fig. 2. Solar Orbiter in cruise configuration—cold conditions. In cruise configuration all appendages are deployed. The solar array presents a face fully equipped with cells towards Sun, providing power for in situ instruments and spacecraft heating. Deploying HGA towards the launcher interface face frees the opposite face.
Fig. 4. Solar Orbiter in science orbit. Hot conditions—tilted solar array. Closer to the Sun, the solar array presents a back face equipped with 25% of cells towards Sun rotated by 65◦ . The low cells/optical reflectors ratio keeps the temperature of the array within qualification status.
After the deployment of the appendages soon after separation from launcher, the spacecraft enters its cruise phase, pointing the Sun with the configuration of Fig. 2. As shown in Fig. 3, when need arises to protect from heat, the solar array presents its 100%-cell face towards Sun rotated by an adjustable angle up to about 65◦ , providing enough power for all instruments in science orbit. The HGA points the Earth, essentially from North face, with its solar torque opposite to solar array, enabling to target 0 thruster activation during the 10-day observation periods. When getting closer to the Sun, attention shall then be paid to protect the solar array and the HGA. These two items must inherit their design from BepiColombo to limit risks and costs. At the present stage, the detailed design is not frozen on BepiColombo and the actual qualification status will result from years of development still ahead. It has therefore been considered compulsory that the system design is made robust to some
variations of the final qualification status achieved by BepiColombo. First, there is a distance to Sun beneath which the 100%-cell face of the solar array can no longer face the Sun. This threshold is assessed to be about 0.35 AU and will be refined as BepiColombo progresses. Note that it is not limiting the mission in a sensitive way in the proposed solution. Beneath about 0.35 AU, thanks to a rotation about its long direction as depicted by Fig. 4, the solar array presents its back face equipped with 25% of cells, towards Sun rotated by 65◦ . Even closer to the Sun, the thermal environment on the HGA may exceed the one of BepiColombo. The configuration offers the possibility to protect the HGA from excessive heat by placing it in the shadow like in Fig. 5. The HGA can then operate towards most of the anti-Sun half-space providing valuable interactivity for science observation program during many perihelia. The solar array attachment is such that its solar torque
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Fig. 5. Solar Orbiter in science orbit—hot conditions—tilted solar array and shadowed HGA. By offering a second operational position for the HGA that can be placed in full shadow of the spacecraft at any time the configuration further secures the adequacy of technological heritage.
is minimized, enabling to target zero thruster activation during the perihelion 10-day observation periods. This second HGA position also allows some pointing in Sun half-space for large elevations, canceling interferences with solar array or Magnetometer boom in cold/warm conditions. In shadowed position, the HGA is cold enough not to impair the heat rejection by the anti-Sun face radiators. The selected system external configuration results in pluses for science objectives: • The communication with the spacecraft is optimized, with maximized interactivity with payload at perihelia. • Thanks to two-sided array, the power margin is increased on all science orbit, not limiting payload on-time. • The management of solar torque to balance array and HGA, and the robust attitude and orbit control subsystem (AOCS) enable to target zero thruster activation during the 10-day observation campaigns. %vspace*12pt 5.2. Accommodation: a secured thermal behavior The selected face allocation results from the following rationale. As there is obviously one Sun face, the four adjacent faces must include one face with high gain antenna able to scan the ecliptic, hence either on North or South. To get long undisturbed observations periods close to Sun, the solar array shall not induce a significant torque, so be symmetrical with two wings. The East and West faces have one wing each. To efficiently reject heat from Sun-facing instruments, one lateral face must be kept free from coupling with wide hot appendages,
in addition to anti-Sun face. North is better than South for the HGA given the mission profile. BepiColombo recurrence on solar array leads to benefit from the two-sided concept to create a dual capacity by having different cell coverage ratios on the two faces. A full-covered face then becomes adapted to low illumination at 1.5 AU, while a face with more reflectors and less cells can withstand thermally the 20 solar constants of Solar Orbiter perihelion. The criticality of the thermal control is such that its robustness and subsequently its simplicity shall be privileged in the trade-offs. To limit the number of Peltier coolers and to maximize the heat rejection capability, the areas free from interaction with hot appendages shall be as large as possible. By using the North face as the Launcher Interface, the selected configuration provides for two usable very cold faces: the South face and the anti-Sun face. The heat shield is subsequently on a lateral face, which presents as an additional advantage a milder out-of-plane dynamic environment. Another valuable spin-off of the HGA deployment towards launcher interface is that its mechanisms and RF-rotary joints can be placed in permanent shadow, further increasing robustness versus BepiColombo qualification status. The solar array is implemented so as to minimize the solar torque at perihelion. This results in zero thruster activation during the 10-day observation periods. Hence no interruption of the observations, since pulsing the thrusters would trigger the closure of the instruments doors to protect optics from contamination. The RPW antennas are accommodated on the South face. Their thermal coupling with the payload radiators decreases the rejection capability of the South face in a limited way. 5.3. A modular design that decouples deformations and schedule The proposed architecture includes two modules (see Fig. 6): • The main module: Structure (excepted Sun face), thermal control, propulsion, harness. • The heat shield module: Heat shield with brackets, Sun face panel, baffles, doors, doors mechanisms, filters, etc. The platform units and the instruments are mounted onto the main module. The platform units are placed in anti-Sun half. The remote-sensing instruments are accommodated on both sides of a South-facing panel as
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Fig. 6. Spacecraft modularity. The modularity ensures both mechanical decoupling and secures the schedule by parallelizing activities and risks.
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way, being close to their radiators. In addition, gathering the remote-sensing instruments on one structure results into a very stable pointing and excellent co-registration performance, as demonstrated by the detailed thermoelastic analyses. The heat shield module can undergo an independent activities path, including a specific high temperature thermal test. The integration schedule and the launch date are thus secured by the parallel paths of main module, heat shield module, instruments, solar arrays and HGA. 5.4. Thermal design: a robust and payload-friendly solution
Fig. 7. Instruments accommodation. Gathering the remote-sensing instruments and Star Trackers on one structure results into a very stable pointing and excellent co-registration performance.
As reported above, the robustness of the thermal control has received a very high priority in all system architecture choices. The strong will to decouple the heat shield, its baffles and its doors from the instruments materializes in the concept of an independent heat shield module. This implies long titanium parts at baffles interfaces, thermal washers at heat-shield-structure to Sun panel connections and multi-layer insulations to close baffle-to-instruments gaps. It is also benefited from the extensions of the heat-shield-bearing Sun panel to add a thermal control capacity further securing the evacuation of heat from the heat shield module. The spacecraft then provides for all the heaters and the dedicated radiators needed by the payload on the South/East/West faces. In situ instruments have local regulation with dedicated heaters and radiators. The Star Trackers are mounted on the South face on the same structure as the remote-sensing instruments, and have a dedicated radiator. The other platform units are thermally controlled by evacuating their heat on the very efficient anti-Sun face, thus preserving the units heritage. Additional high temperature multi-layer insulation (MLI) is implemented on the parts of the satellite that will receive solar flux in case of depointing. The heating budget and the radiators have been sized after thorough modeling and analyses, that included heat shield modeling, without evidencing shortage of power resources or radiators areas. 5.5. RF system: a powerful data flow with secured qualification status
illustrated by Fig. 7, along with the alignment-critical Star Trackers. The instruments on top are protected from radiations by a secondary structure that is also used as a support for the radiators. All remote instruments can therefore reject their heat to space in a very efficient
The critical parts of the dual band X + Ka BepiColombo RF system, especially the HGA and the medium gain antenna (MGA) are reused. The MGA recurrence is complete, including its arms and
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Fig. 8. RF system block diagram. The recurrence is preserved from BepiColombo, with increased power in science orbit.
mechanisms. The HGA diameter has to be strictly identical to BepiColombo not to generate significant development risks and costs. The compliance with the BepiColombo heritage for the mechanisms and the RF rotary joint is secured by placing them in permanent shadow. The rest of the HGA is protected, if needed, by the second operational position in the spacecraft shadow cone. The RF system diagram is provided in Fig. 8. A trade-off has been conducted concerning Ka-band: this latter has been found necessary due to the very demanding data flow across large Earth-Spacecraft distances. To provide for better transmission of science data and to keep a reasonable mass memory size, the data rate has even been further improved beyond specification. This is particularly adapted to the Thales Alenia Space system configuration, which presents an excess of power generation in science orbit allowing for a higher data transmission, while having a solar array sized by the sunlight-deprived cruise phase. This generates a valuable improvement of the availability of science data on ground during science orbit, without penalizing the BepiColombo technological recurrence for the solar array panels.
product lines. The data handling subsystem will be based on the following major elements:
5.6. Avionics: a recurring design
• One cold face, 100% cells, 0% optical solar reflector (OSR), covering power needs in cruise phase, and usable with rotation about wing long dimension up to about 0.35 AU thus providing ample power margins in science orbit.
The data handling units and subsystem architecture are assumed as baseline recurring from GMES Sentinel 3, benefiting from the Thales Alenia Space
• A computer, called “Satellite Management Unit” (SMU), managing all the satellite platform and payload elements. • A computer, called “Payload Data Handling Unit” (PDHU), collecting and storing the science packets issued by the instruments. The Payload Mass Memory is included in the PDHU. It has been sized with Solar Orbiter mission profiles and data link budget. The number of standard boards is compatible with the modular capability of the generic PDHU. The resulting functional architecture is sketched in Fig. 9. 5.7. Power: preserving the BepiColombo heritage The proposed solar array configuration is a twowinged array, with two symmetrical wings on E/W faces balancing the disturbance torque and substrate dimensions inherited from BepiColombo. It is two-sided, as on BepiColombo, but with a differentiation of faces:
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Fig. 9. Functional block diagram. The recurring functional and hardware designs guarantee the control of the functional validation phases, while providing for a choice of standard SpaceWire and 1553 interfaces.
• One hot face, 25% cells, 75% OSR, used at perihelion with rotation about wing long dimension, with margins versus the limit angle of 70◦ to secure the power regulation, and large overlap domain with the 100%-face. By rotating along its long dimension, it minimizes the interference with fields of view, and enables the recurrence with the BepiColombo two-sided design and drive mechanisms. The attachment of the wings to the spacecraft body is located so as to minimize the solar torque, targeting zero wheels desaturation during the 10-day observation periods. The Power Conditioning and Distribution Unit (including its maximum power point tracking, MPPT, management to optimize the effect of the variation of temperature on voltage and current), and the Lithiumion batteries are fully inherited from BepiColombo.
5.8. AOCS: guaranteeing the sun pointing thanks to the Herschel–Planck design Thales Alenia Space experience on Herschel–Planck is vital for Solar Orbiter. The two programs share indeed the absolute obligation to keep Sun pointing and be able to recover it extremely rapidly in case of anomaly, or risk of losing the spacecraft. In the case of Solar Orbiter, dispositions are taken in the thermal design of the body of the satellite to be robust to transient exposure to the 20 Solar Constant flux. However, the thermal flux is then so high that the satellite survivability relies on the swift reaction from the spacecraft. In this context Thales Alenia Space has considered compulsory to reuse the AOCS and safe mode designs from Herschel–Planck. Given the outstanding performance of the Star Tracker-pointed Herschel–Planck AOCS, this is also beneficial to Solar Orbiter in the
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nominal modes. Some units have to be adapted due to the difference in thermal environment but off-theshelf hardware could always be found that satisfied the needs of Solar Orbiter without at worst more than a delta-qualification. Nominal modes include a science mode with all payload on, use of four reaction wheels to minimize unloading and zero crossings, fine gyros and star trackers, and an in situ mode with only in situ instruments on. This latter is less demanding in terms of wheel off-loading, and subsequently uses three reaction wheels and coarse gyros. The orbit correction mode relies on star trackers, coarse gyros and 1 IN thrusters. The safe modes include an acquisition mode, a Sun-spin mode and a Sun inertial mode. The high level failure detection identification and recovery (FDIR) anomaly detection relies on a separate set of hardware to guarantee the independence of the off-pointing evaluation. Slit Sun sensors are used as anomaly detectors. The thrusters accommodation privileges force-free attitude control and minimization of contamination towards heat shield, remote-sensing instruments. 5.9. An extensive justification The spacecraft design presented above results from the thorough trade-offs carried out by Thales Alenia Space. It combines the imperative use of low-risk proven designs and the fulfillment of the challenging needs of Solar Orbiter mission in terms of stringent environment, power and data transmission. It succeeds in limiting the technological development to the heat shield itself, its baffles and doors and results into a minimized amount of delta-qualification for all the rest. Table 3 summarizes how the main features of the baselined Solar Orbiter design answer the design criteria. This design has been successfully analysis-proven. It is now consolidated by a comprehensive analytical effort addressing: • • • • • • • • • • • •
mass and power system budgets, link budget, mass memory sizing, mechanical resistance and loads, thermal sizing, thermoelastic distortions, molecular contamination, particulate contamination, magnetic contamination, radiation environment, AOCS modes simulations, wheels sizing,
• FDIR and safe mode analyses, • solar array power analysis, etc. The degree of justification has enabled to generate not only the compliance matrix and the system requirements but also the instruments interface requirements, the subsystem specifications and the units environmental requirements. The level of the study has been pushed to a degree of detail that now allows starting the definition phase upon ESA’s decision. 6. Design and development of the Solar Orbiter heat shield The design of the heat shield has been elaborated in conjunction with the system design. The evolution towards a consistent heat shield module integrating all items crossing the shield has been a natural result of the system optimization and of the maximization of the robustness. It enables the full control and validation of the thermal behavior of this critical area. The selected design has obeyed to the same criteria as for the system, with special emphasis on thermal robustness. 6.1. Heat shield architecture Its sound and simple design consists of several heat barriers separated by gaps that radiate the heat to space. It includes the following main components: • Front shield (FS), first layer that receives direct solar light. • Several high temperature heat barriers (HTHB), used behind the front shield. • Structural elements to support the FS and the HTHBs and maintain gaps between them. It is integrated and validated along with the Sun panel of the spacecraft, all doors, doors mechanisms and baffles. Different geometries, composition and combinations of the FS, HTHBs, have been envisaged. The geometry has been investigated concluding that a simple heat shield with planar front shield, separated from spacecraft by gaps, provides for very good protection. To improve this lateral heat rejection, the inner layers are highly reflective in the infrared. The optical properties of the surface facing sun become less important, and for a suitable number of gaps the conditions at the rear of the heat shield are little affected by the properties of the front. The number of gaps has been optimized
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Table 3 The baselined spacecraft design addresses the challenges of the Solar Orbiter mission. Criteria
Major features
Maximize recurrence, control of costs and risks
• Extensive heritage from BepiColombo (TTC, HGA, solar array), GMES (avionics), Herschel–Planck (AOCS design, sensors) • Modular design, with independent heat shield module • HGA deployable on North face, with mechanisms in the shadow, optimizing recurrence with BepiColombo definition while securing thermal environment of mechanisms • Platform units on anti-Sun face, favoring avionics recurrence with mild environment
Control mass, volume
• Sun face is lateral in launch configuration, minimizing heat shield mass and mechanical environment, and reducing the central area of shield
Get robust payload accommodation
• Remote-sensing instruments on South face, with high rejection capability • Additional capability on lateral faces and growth potential along launcher axis
Maximize thermal robustness
• Attitude control inherited from Herschel–Planck, ensuring swift and reliable Sun de-pointing recovery • PAS, HIS on North face, minimizing impact of large fields of view • Sun face lateral in launch configuration, reducing the central area of shield • Heat shield module with panel extension provides for extra-rejection capability
Minimize interactions radiators/appendages/fields of view
• Two operational positions of HGA • Rotating solar array reduces thermal rejection capability on East/West faces but minimizes interaction with in situ fields of view
Minimize instruments misalignments
• Remote-sensing instruments on one South facing structure, decoupled from heat shield
Maximize communication link
• Additional operational position of HGA, in shadow, guaranteeing science downlink during many close perihelia, and escape from solar array and Magnetometer interferences at any time • Increased data flow in Science orbit
Maximize uninterrupted observation periods
• Two symmetrical solar arrays on East/West faces balance thedisturbance torque • Wheels sized to avoid desaturation
subsequently, so as to provide for reasonable mass while coupling the layers and the instruments baffles to space. The selected design minimizes the heat load into the spacecraft and allows a great flexibility to trim the heat load using the sun panel extension as radiator. By minimizing the heat transfer to the spacecraft and making it weakly dependent on the front shield temperature, it becomes possible to focus on design aspects other than optical properties, such as outgassing, charge build-up, and mechanical aspects. A weak dependence of shield performance on the optical properties also provides more flexibility in choice of materials and increases the robustness of the design. 6.2. Heat shield thermal layers The thermal design of the HS has evolved through several trade-offs, including flexible versus rigid layers.
The selected design employs flexible layers. The use of flexible layers is of great advantage, as it allows easy adaptation to the many instruments apertures and doors, and decoupling with the spacecraft in terms of loads and thermoelastic distortions. The baselined design enables to select the best material for the front shield without altering the rest of the heat shield. The baseline front shield has a high absorptivity. It is insensitive to variations of optical properties, robust to temperatures beyond 700 ◦ C, electrically conductive, generates no molecular contamination and enables to control particulate contamination. Thales Alenia Space has also developed a high-performance white and electrically conductive front shield (patented), further demonstrating the robustness of the selected heat shield concept. The HTHB layers use materials selected for their resistance to high temperature and favorable optical properties.
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Mastering the effect of thermal layers attachment is key to control the performance in a high thermal load environment. Thales Alenia Space has acquired a great knowledge of the effects of such junctions in the context of very high heat fluxes. Environments up to 60 kW/m2 on transients have been experienced on Pleiades and provide for valuable lessons learned that benefit Solar Orbiter. The detailed design has been defined for flight definition and implemented in the heat shield breadboard. 6.3. Heat shield structure The trade-offs have scanned different structural concepts including frames or independent brackets. The design is not purely mechanical: if not well controlled, the structure elements can produce an important contribution to the heat load into the spacecraft. The following drivers have been considered and analyzed in the mechanical design optimization: • Ensure adequate thermal performance. • Provide mechanical interfaces sustaining the thermal layers. • Maintain gaps between layers accounting for thermal expansion. • Provide the necessary areas for instruments baffles and doors accommodation. • Decouple baffles. • Maximize accessibility. • Enable integration and maintenance of the entire heat shield. • Minimize mass. • Minimize design complexity. • Get stiffness/strength best compromise. The selected design is an excellent answer to all these needs. The structure can be easily adapted to the configuration of the instruments baffles that cross the shield, so that no constraint is transferred to the instruments and to their accommodation. The integrability is very strong, enabling to replace components in any phases of the program. 6.4. Instruments doors and baffles The remote-sensing instruments, as well as PAS and HIS analyzers, have fields of view or fields of collection that cross the heat shield. The apertures, sometimes large, require baffling. In addition many remote-sensing apertures are to be protected from contamination by doors. These latter have to be placed close to the front
shield and so be commanded by shafts that cross the layers of the heat shield. The selected design is highly adaptive to the multiple crossings. The baffles are mounted on the Sun panel so as to minimize thermal conduction. 6.5. Analytical justification The heat shield has undergone an extensive analytical evaluation, with detailed modeling and assessment of behavior under sine, acoustic, quasi-static, thermoelastic loads, and of course thermal environment. The thermal load has been assessed in a preliminary way with local thermal models. The detailed assessment has relied on an integrated thermal modeling of the whole satellite, mastering the fluxes and interfaces with instruments in a consistent way. 6.6. Development tests To secure the Solar Orbiter development prior to the confirmation of the mission, Thales Alenia Space designed a comprehensive test campaign on a full-scale representative breadboard and on samples. The chief motivation for the Solar Orbiter heat shield technology development is the harsh near Sun environment, with solar flux peaking at about 20 solar constants, and temperatures above 600 ◦ C, to be sustained in a long-duration mission, totaling about 200,000 equivalent Sun hours. Accordingly, the development has focused on demonstrating a thermal performance and thermo-mechanical stability, which can be reliably extrapolated to the desired lifetime. This is accomplished in a two-pronged approach: • Test of materials samples, to establish basic materials properties and their stability in time and environment. • Test of a representative breadboard to provide overall verification of the selected design. 6.7. Characterization testing Thales Alenia Space pre-selected the optimally suited materials from a shortlist established on the basis of our in-house experience and the experience ESA matured in the BepiColombo technology developments. The final selection has then been made in agreement with ESA, after our database has been further augmented by tests at sample level. The sample test program has been built to address the stability of the optical, thermal and mechanical properties of the materials, with measurements before and after long-duration exposure
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IR loading rig
Full-scaled Heat Shield
S/C simulator
Fig. 10. Full-scale heat shield breadboard. The breadboard includes all thermal layers of the Shield, its structure, the Sun panel and a representation of the apertures, with flight-definition junctions of all thermal elements.
to high-temperatures, along with outgassing characterization. 6.8. Breadboard testing For the overall verification, Thales Alenia Space has designed and manufactured in Turin a full-scale breadboard (Fig. 10), fully representative of the mechanical and thermal design. It includes apertures and baffles representing two types of crossings. Its fidelity to the flight design includes the representation of the junctions and attachments between all heat shield elements. The full scale is a major feature of Thales Alenia Space breadboard design. It enables: • to validate the evacuation of the heat with the real shield-to-spacecraft stand-off in a high-power thermal vacuum test; • to represent the effect of discontinuities such as instruments aperture, junctions or interface brackets in a non-disproportionate way; • to use the same breadboard to carry out representative vibration tests in later phases of the project. The test article includes a full-scale model of the heat shield mounted on a spacecraft simulator covered by MLI and controlled by heaters. The overall set-up is represented in Fig. 11. The verification of the HS thermal performance consists in measuring, in conditions representative of the extreme flight ones, the overall heat transfer to the spacecraft by all paths: conduction through the supports, radiation through the gap and
Fig. 11. Breadboard and test set-up. The breadboard is loaded in infra-red, validating not only the general concept, but also the effect of apertures and junctions, at a detailed level.
multilayer insulation and the instrument apertures according to their specific design. The determination of the optical properties of the HS in mission conditions is made at sample level, with reproduction of solar spectrum, while the overall functional performances, which are a function of the temperature, are determined using infra red (IR) sources. The front shield is exposed to a heat source that transfers heat by radiation. IR sources face the front shield. Their intensity is trimmed to give the expected in-flight front shield temperatures. Thales Alenia Space has designed and built the IR rig. Design drivers have been the uniformity of the illumination of the test article, tolerance of the high temperatures, minimization of the power dispersed inside the vacuum chamber. To reduce the lamp power, the lamp array is fitted with rear insulation to concentrate the emission on the test article, so that the power is essentially absorbed by the FS. The temperatures of the chamber shrouds are those of liquid nitrogen. The test article contains heaters for temperature control of spacecraft simulator as shown in Fig. 12. The 70m3 chamber in Thales Alenia Space Cannes is used for the test. 7. Solar Orbiter: a cost- and schedule-efficient concept In addition to the outstanding technical performance, the selected designs present an excellent cost-efficiency and a resilient schedule. Due to programmatic constraints, Solar Orbiter will indeed have a 5-year B2/C/D phase leading to a launch in 2015. This implies privileging recurrent designs and the early mitigation of the remaining technological risks.
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to minimize unexpected extra validation loops on both software and hardware test benches. The risks identified and the planning of all development, engineering, manufacturing and satellite integration activities lead to confirm the feasibility of the 2015 launch. 8. Conclusion
Fig. 12. Spacecraft simulator. The test set-up measures in a straightforward way the heat load transferred to the regulated spacecraft simulator.
The feasibility of the mission is also financial, with an important stress on the identification of cost drivers and on subsequent control and reduction of costs. Thales Alenia Space has addressed this by analyzing the cost and schedule drivers and selecting the lowestrisk design with recurring units and a robust heat shield. Technical requirements and design drive costs primarily via the robustness of pay load accommodation and mechanical–thermal configuration, the simplicity and the recurrence of avionics and AOCS, the TTC link sizing, and the solar array power demand. On all these drivers, the selected design trims costs, ensuring the financial feasibility. Robust designs, by providing margins, will also minimize potential additional iterations during the development and should be privileged as much as possible in this schedule-driven context. This motivated the selection by Thales Alenia Space of a safe system architecture protecting most parts in the shadow of the shield, with a simple stand-off-based design for the shield using robust end-of-life thermo-optical properties. The sun-pointing reliability inherited from Herschel–Planck AOCS and the recurrence of the avionics will be key
Solar Orbiter is a challenging mission, which covers a wide spread spectrum of scientific objectives with very high accuracy in a stringent thermal environment. The system configuration elaborated by Thales Alenia Space enables to reuse HGA, solar array technology from BepiColombo while satisfying the science objectives. Associated with the modular, robust and flexible heat shield design, it ensures the mastery of the upcoming phases of the development, maximizing the control of costs and schedule. The achievements in terms of heat shield development and the maturity gained on system and subsystem design and requirements enable to start the definition phase upon ESA’s final decision on Solar Orbiter mission. Acknowledgments The authors want to thank the whole study team at Thales Alenia Space for their excellent work. The work presented in this paper was performed under funding of the European Space Agency (ESA) in the frame of the Solar Orbiter Heat Shield/System Technology contract, coordinated by Mr. Ph. Kletzkine, and of Thales Alenia Space internal funding. We would like also to thank ESA’s Ph. Kletzkine, J. Marti Canales, D. Renton, Ph. Poinas and D. McCoy for the fruitful exchanges, very good relation and their support in our achievements. References [1] ESA: Solar Orbiter Heat Shield/System Technology SOW: SOLEST-SOW-56. [2] ESA: Solar Orbiter Technical Information Update Document— A: SOL-EST-IF-271.
Solar Orbiter—Heat shield and system technology J. Poncya,∗ , F. Jubineaua , F. D’Angelob , V. Perottob , J.J. Juilleta a Thales Alenia Space, 100 boulevard du midi, F-06100 Cannes, France b Thales Alenia Space Italia, Turin, Italy
Received 5 December 2007; accepted 9 March 2009 Available online 16 April 2009
Abstract Solar Orbiter will enhance our knowledge of the Sun by observations and in situ measurements as close as 0.22 AU from our star. Placed on an orbit with a period two-thirds the one of Venus, Solar Orbiter will use the many encounters with the planet to gradually incline its orbit and gain view on the Sun’s poles. The permanent in situ observations will be associated to remote-sensing observations over large parts of the orbits. ESA Science Directorate has awarded in parallel two Solar Orbiter Heat Shield and System Technology contracts to industry. This paper presents the achievements of Thales Alenia Space thanks to one of these two ESA contracts. It shows how the main technical challenge brought by the heat flux of 20 solar constants has been addressed by the system and heat shield design. The design and manufacturing of a breadboard of the heat shield in view of thermal test verification is then reported. The main technological developments and residual risks are assessed, paving the way for the definition phase of the program. © 2009 Elsevier Ltd. All rights reserved.
1. Introduction The most important of all stars, our Sun, requires a deeper understanding of its metabolism. The originality of the Solar Orbiter mission lies in its ability to combine in situ measurements in the inner system with closer remote-sensing observations of the Sun and the Corona. However, navigating in the inner system constrains the Abbreviations: AOCS, attitude and orbit control subsystem; FDIR, failure detection identification and recovery; FS, front shield; HGA, high gain antenna; HS, heat shield; HTHB, high temperature heat barrier; IR, infra red; MGA, medium gain antenna; MLI, multi- layer insulation; MPPT, maximum power point tracking; OSR, optical solar reflectors; PDHU, Payload Data Handling Unit; PPS, pulse per second; RF, radio frequency; S/C, SpaceCraft; SMU, Satellite Management Unit ∗ Corresponding author. E-mail address: [email protected] (J. Poncy). 0094-5765/$ - see front matter © 2009 Elsevier Ltd. All rights reserved. doi:10.1016/j.actaastro.2009.03.031
mission and spacecraft design, both in terms of large delta-v in the Sun’s gravity well and of the very hostile thermal environment. An Assessment Study was carried out in 2005, with Thales Alenia Space as one of ESA’s two contractors. It identified the heat shield (HS), the high gain antenna (HGA) and the solar array as the main elements requiring anticipation of technological activities. The latter two being covered by ESA’s funded BepiColombo development, ESA subsequently complemented the development effort by awarding two contracts for a Heat Shield and System Technology study [1]. This paper presents the achievements of Thales Alenia Space in the frame of one of these contracts. Thales Alenia Space succeeded in finding technical solutions that limit the technological developments to the heat shield and the instruments doors and baffles. Our achievements increased the degree of maturity of
J. Poncy et al. / Acta Astronautica 65 (2009) 1076 – 1088 Table 1 Baseline mission overview. Timeline
Flight time (years)
Event date (calendar)
Launch Venus GAM-1 Deep Space Maneuver-1 Earth GAM-1 Earth GAM-2 Venus GAM-2 Venus GAM-3 Venus GAM-4
0.0 0.51 1.02 1.38 3.21 3.39 4.62 5.85
2015/05/22 2015/11/26 2016/05/28 2016/10/08 2018/08/08 2018/10/09 2020/01/02 2021/07/08
End of Nominal Mission
6.13
2021/07/08
Venus GAM-5 Venus GAM-6
7.08 8.31
2022/06/19 2023/09/11
End of Extended Mission
8.64
2024/01/31
this complex mission, both in terms of heat shield and of system, enabling to start the definition phase upon ESA’s final decision on Solar Orbiter mission. 2. Solar Orbiter mission By approaching as close as 45 solar radii, the Solar Orbiter will view the solar atmosphere with unprecedented spatial resolution circa 100 km pixel size. Over extended periods Solar Orbiter will deliver images and data of the polar regions and the side of the Sun not visible from Earth. The scientific goals of the Solar Orbiter are ([2]): • to determine in situ the properties and dynamics of plasma, fields and particles in the near-Sun heliosphere, • to survey the fine detail of the Sun’s magnetized atmosphere, using a camera capable of detecting solar features as small as 100 km across, • to identify the links between activity on the Sun’s surface and the resulting evolution of the corona and inner heliosphere, using solar co-rotation passes, • to observe and fully characterize the Sun’s polar regions and equatorial corona from high latitudes. The current baseline trajectory for the mission (direct escape and transfer based on a tour of planetary flybys and a Deep Space Maneuver) is shown in Table 1. The mission will rely on a chemical propulsion system for maneuver performance. The “in situ” set of instruments will be commissioned shortly after LEOP and, nominally, will be continuously switched on, while the
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“remote-sensing” instruments are to be commissioned after the first Earth GAM and operated during certain windows around perihelion passages. Science operations should be possible during the initial part of the trajectory (i.e. before Venus GAM 2); therefore it can be considered that the nominal mission will start shortly after LEOP and initial instrument commissioning. After Venus GAM 2 the spacecraft is injected in a resonant orbit (the period of revolution of the Solar Orbiter around the Sun is 23 the period of Venus) in order to increase the inclination of the orbit with respect to the ecliptic. The following timeline is indicative of the mission and will be confirmed during the course of the Solar Orbiter development phase. During the nominal mission, the Solar Orbiter performance requirements shall be fully met with all specified margins according to the Mission Requirements Document [NR1] for the Solar Orbiter mission. The duration of the Solar Orbiter extended operational lifetime shall be about 2.5 years. Approval of an extension of the mission is dependent on the spacecraft health and status of expendables such as fuel. Alternative mission profiles are considered in case of common launch with NASA’s Solar Sentinels. Thales Alenia Space has studied them along with the nominal mission. 3. Solar Orbiter payload The payload combines in situ instruments (field analyses and particles collectors) with extensive allspectrum remote observation of the Sun and the Corona. Table 2 lists the baseline (core) payload considered for the purpose of this study. The payload complement may be further modified or pared down by ESA in the near future. In addition, some augmentation payload is being traded-off, in particular a Wide-Field Coronagraph or Heliospheric Imager type of instrument. Thales Alenia Space has successfully implemented all the instruments, and the subsequent interface requirements have been inferred. 4. Design drivers and criteria The major design drivers specific to the Solar Orbiter mission are: • The cost and schedule of the mission, leading to privilege recurring units and to reuse the specific developments of BepiColombo, especially on solar array and high gain antenna.
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Table 2 Core payload. Instrument
Acron.
Science goals
In situ instruments Solar Wind Plasma Analyzer
SWA
Radio and Plasma Wave Analyzer Magnetometer Energetic Particle Detector Dust Particle Detector Neutron Gamma ray Detector
RPW MAG EPD DPD NGD
Investigation of kinetic properties and composition (mass and charge states) of solar wind plasma Investigation of radio and plasma waves including coronal and interplanetary emissions Investigation of the solar wind magnetic field Investigation of the origin, acceleration and propagation of solar energetic particles Investigation of the flux, mass and major elemental composition of near-Sun dust Investigation of the characteristics of low energy solar neutrons, and solar flare processes
Remote-sensing instruments Visible Imager and Magnetograph EUV Spectrometer EUV Imager Visible Coronograph Spectrometer Telescope Imaging X-ray
VIM EUS EUI COR STIX
Investigation Investigation Investigation Investigation Investigation
• The close perihelion (0.22 AU in baseline mission scenario), generating high thermal fluxes up to 28,500 W/m2 , high radiations, UV ageing and solar disturbance torques. • The large variation in solar array illumination, from 0.5 to 20 solar constants, with the cruise phase at 1.5 AU sizing the power generation. • The very large data transmission capability required, due to a variable distance between Earth and Spacecraft, generating needs for a powerful radio frequency (RF) system and a large mass memory. • The numerous instruments requiring reliable thermal rejection, crossing of the heat shield, high pointing accuracy, many appendages or fields of view and very stringent cleanliness levels (leading to add doors in many cases). • The launcher compatibility, limiting the mass and volume of the spacecraft: Soyuz launch, with Atlas as a possible alternative. The drivers have been processed during the tradeoffs carried out by Thales Alenia Space, so as to provide for the largest scientific return under the strict necessity of mastered cost and schedule. This resulted in the following criteria for the selection of the technical solutions: • maximize recurrence and control of technical and schedule risks; • control mass and volume; • get robust payload accommodation; • maximize thermal robustness; • minimize interactions between radiators, appendages and fields of view;
of of of of of
the magnetic and velocity fields in the photosphere properties of the solar atmosphere the solar atmosphere using high resolution imaging in the EUV coronal structures using polarized brightness measurements in Visible energetic electrons near the Sun, and solar X-ray emission
• minimize instruments misalignments; • maximize communication link and rate; • maximize uninterrupted observation periods, with minimization of disturbance torques. 5. Solar Orbiter spacecraft design Thales Alenia Space has conducted extensive tradeoffs to select a configuration that best fits these mission drivers. The team included members of TAS Italy (thermal control, heat shield, TTC) and TAS France (system and other subsystems). A special emphasis has been put to provide configurations that combine lowcost low-risk approach, strong reuse of BepiColombo heritage, resilient-schedule modular design and maximum resources for science. The mission is proposed to be fulfilled by a parallelepiped body shadowed by a heat shield. The high gain antenna and solar arrays are managed in a series of configurations that increase the data link and power capabilities and protect the BepiColombo heritage. The satellite carries the 11 primary instruments of the core payload specified for the purposes of this study. 5.1. An optimal configuration In launch configuration (Fig. 1), the satellite is in a compact configuration that controls the loads on the units and minimizes the out-of-plane coupling with the heat shield. The satellite has its South face up. The high gain antenna is stowed on the anti-Sun face on its backside.
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Fig. 1. Solar Orbiter in launch configuration. The compact configuration decreases mass and minimizes loads on heat shield, HGA.
Fig. 3. Solar Orbiter in cruise or science orbit—warm conditions. As the distance to Sun decreases, the solar flux compensates for the tilt of the panels, resulting into an increased power, while keeping the temperature below BepiColombo qualification status.
Fig. 2. Solar Orbiter in cruise configuration—cold conditions. In cruise configuration all appendages are deployed. The solar array presents a face fully equipped with cells towards Sun, providing power for in situ instruments and spacecraft heating. Deploying HGA towards the launcher interface face frees the opposite face.
Fig. 4. Solar Orbiter in science orbit. Hot conditions—tilted solar array. Closer to the Sun, the solar array presents a back face equipped with 25% of cells towards Sun rotated by 65◦ . The low cells/optical reflectors ratio keeps the temperature of the array within qualification status.
After the deployment of the appendages soon after separation from launcher, the spacecraft enters its cruise phase, pointing the Sun with the configuration of Fig. 2. As shown in Fig. 3, when need arises to protect from heat, the solar array presents its 100%-cell face towards Sun rotated by an adjustable angle up to about 65◦ , providing enough power for all instruments in science orbit. The HGA points the Earth, essentially from North face, with its solar torque opposite to solar array, enabling to target 0 thruster activation during the 10-day observation periods. When getting closer to the Sun, attention shall then be paid to protect the solar array and the HGA. These two items must inherit their design from BepiColombo to limit risks and costs. At the present stage, the detailed design is not frozen on BepiColombo and the actual qualification status will result from years of development still ahead. It has therefore been considered compulsory that the system design is made robust to some
variations of the final qualification status achieved by BepiColombo. First, there is a distance to Sun beneath which the 100%-cell face of the solar array can no longer face the Sun. This threshold is assessed to be about 0.35 AU and will be refined as BepiColombo progresses. Note that it is not limiting the mission in a sensitive way in the proposed solution. Beneath about 0.35 AU, thanks to a rotation about its long direction as depicted by Fig. 4, the solar array presents its back face equipped with 25% of cells, towards Sun rotated by 65◦ . Even closer to the Sun, the thermal environment on the HGA may exceed the one of BepiColombo. The configuration offers the possibility to protect the HGA from excessive heat by placing it in the shadow like in Fig. 5. The HGA can then operate towards most of the anti-Sun half-space providing valuable interactivity for science observation program during many perihelia. The solar array attachment is such that its solar torque
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Fig. 5. Solar Orbiter in science orbit—hot conditions—tilted solar array and shadowed HGA. By offering a second operational position for the HGA that can be placed in full shadow of the spacecraft at any time the configuration further secures the adequacy of technological heritage.
is minimized, enabling to target zero thruster activation during the perihelion 10-day observation periods. This second HGA position also allows some pointing in Sun half-space for large elevations, canceling interferences with solar array or Magnetometer boom in cold/warm conditions. In shadowed position, the HGA is cold enough not to impair the heat rejection by the anti-Sun face radiators. The selected system external configuration results in pluses for science objectives: • The communication with the spacecraft is optimized, with maximized interactivity with payload at perihelia. • Thanks to two-sided array, the power margin is increased on all science orbit, not limiting payload on-time. • The management of solar torque to balance array and HGA, and the robust attitude and orbit control subsystem (AOCS) enable to target zero thruster activation during the 10-day observation campaigns. %vspace*12pt 5.2. Accommodation: a secured thermal behavior The selected face allocation results from the following rationale. As there is obviously one Sun face, the four adjacent faces must include one face with high gain antenna able to scan the ecliptic, hence either on North or South. To get long undisturbed observations periods close to Sun, the solar array shall not induce a significant torque, so be symmetrical with two wings. The East and West faces have one wing each. To efficiently reject heat from Sun-facing instruments, one lateral face must be kept free from coupling with wide hot appendages,
in addition to anti-Sun face. North is better than South for the HGA given the mission profile. BepiColombo recurrence on solar array leads to benefit from the two-sided concept to create a dual capacity by having different cell coverage ratios on the two faces. A full-covered face then becomes adapted to low illumination at 1.5 AU, while a face with more reflectors and less cells can withstand thermally the 20 solar constants of Solar Orbiter perihelion. The criticality of the thermal control is such that its robustness and subsequently its simplicity shall be privileged in the trade-offs. To limit the number of Peltier coolers and to maximize the heat rejection capability, the areas free from interaction with hot appendages shall be as large as possible. By using the North face as the Launcher Interface, the selected configuration provides for two usable very cold faces: the South face and the anti-Sun face. The heat shield is subsequently on a lateral face, which presents as an additional advantage a milder out-of-plane dynamic environment. Another valuable spin-off of the HGA deployment towards launcher interface is that its mechanisms and RF-rotary joints can be placed in permanent shadow, further increasing robustness versus BepiColombo qualification status. The solar array is implemented so as to minimize the solar torque at perihelion. This results in zero thruster activation during the 10-day observation periods. Hence no interruption of the observations, since pulsing the thrusters would trigger the closure of the instruments doors to protect optics from contamination. The RPW antennas are accommodated on the South face. Their thermal coupling with the payload radiators decreases the rejection capability of the South face in a limited way. 5.3. A modular design that decouples deformations and schedule The proposed architecture includes two modules (see Fig. 6): • The main module: Structure (excepted Sun face), thermal control, propulsion, harness. • The heat shield module: Heat shield with brackets, Sun face panel, baffles, doors, doors mechanisms, filters, etc. The platform units and the instruments are mounted onto the main module. The platform units are placed in anti-Sun half. The remote-sensing instruments are accommodated on both sides of a South-facing panel as
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Fig. 6. Spacecraft modularity. The modularity ensures both mechanical decoupling and secures the schedule by parallelizing activities and risks.
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way, being close to their radiators. In addition, gathering the remote-sensing instruments on one structure results into a very stable pointing and excellent co-registration performance, as demonstrated by the detailed thermoelastic analyses. The heat shield module can undergo an independent activities path, including a specific high temperature thermal test. The integration schedule and the launch date are thus secured by the parallel paths of main module, heat shield module, instruments, solar arrays and HGA. 5.4. Thermal design: a robust and payload-friendly solution
Fig. 7. Instruments accommodation. Gathering the remote-sensing instruments and Star Trackers on one structure results into a very stable pointing and excellent co-registration performance.
As reported above, the robustness of the thermal control has received a very high priority in all system architecture choices. The strong will to decouple the heat shield, its baffles and its doors from the instruments materializes in the concept of an independent heat shield module. This implies long titanium parts at baffles interfaces, thermal washers at heat-shield-structure to Sun panel connections and multi-layer insulations to close baffle-to-instruments gaps. It is also benefited from the extensions of the heat-shield-bearing Sun panel to add a thermal control capacity further securing the evacuation of heat from the heat shield module. The spacecraft then provides for all the heaters and the dedicated radiators needed by the payload on the South/East/West faces. In situ instruments have local regulation with dedicated heaters and radiators. The Star Trackers are mounted on the South face on the same structure as the remote-sensing instruments, and have a dedicated radiator. The other platform units are thermally controlled by evacuating their heat on the very efficient anti-Sun face, thus preserving the units heritage. Additional high temperature multi-layer insulation (MLI) is implemented on the parts of the satellite that will receive solar flux in case of depointing. The heating budget and the radiators have been sized after thorough modeling and analyses, that included heat shield modeling, without evidencing shortage of power resources or radiators areas. 5.5. RF system: a powerful data flow with secured qualification status
illustrated by Fig. 7, along with the alignment-critical Star Trackers. The instruments on top are protected from radiations by a secondary structure that is also used as a support for the radiators. All remote instruments can therefore reject their heat to space in a very efficient
The critical parts of the dual band X + Ka BepiColombo RF system, especially the HGA and the medium gain antenna (MGA) are reused. The MGA recurrence is complete, including its arms and
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Fig. 8. RF system block diagram. The recurrence is preserved from BepiColombo, with increased power in science orbit.
mechanisms. The HGA diameter has to be strictly identical to BepiColombo not to generate significant development risks and costs. The compliance with the BepiColombo heritage for the mechanisms and the RF rotary joint is secured by placing them in permanent shadow. The rest of the HGA is protected, if needed, by the second operational position in the spacecraft shadow cone. The RF system diagram is provided in Fig. 8. A trade-off has been conducted concerning Ka-band: this latter has been found necessary due to the very demanding data flow across large Earth-Spacecraft distances. To provide for better transmission of science data and to keep a reasonable mass memory size, the data rate has even been further improved beyond specification. This is particularly adapted to the Thales Alenia Space system configuration, which presents an excess of power generation in science orbit allowing for a higher data transmission, while having a solar array sized by the sunlight-deprived cruise phase. This generates a valuable improvement of the availability of science data on ground during science orbit, without penalizing the BepiColombo technological recurrence for the solar array panels.
product lines. The data handling subsystem will be based on the following major elements:
5.6. Avionics: a recurring design
• One cold face, 100% cells, 0% optical solar reflector (OSR), covering power needs in cruise phase, and usable with rotation about wing long dimension up to about 0.35 AU thus providing ample power margins in science orbit.
The data handling units and subsystem architecture are assumed as baseline recurring from GMES Sentinel 3, benefiting from the Thales Alenia Space
• A computer, called “Satellite Management Unit” (SMU), managing all the satellite platform and payload elements. • A computer, called “Payload Data Handling Unit” (PDHU), collecting and storing the science packets issued by the instruments. The Payload Mass Memory is included in the PDHU. It has been sized with Solar Orbiter mission profiles and data link budget. The number of standard boards is compatible with the modular capability of the generic PDHU. The resulting functional architecture is sketched in Fig. 9. 5.7. Power: preserving the BepiColombo heritage The proposed solar array configuration is a twowinged array, with two symmetrical wings on E/W faces balancing the disturbance torque and substrate dimensions inherited from BepiColombo. It is two-sided, as on BepiColombo, but with a differentiation of faces:
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Fig. 9. Functional block diagram. The recurring functional and hardware designs guarantee the control of the functional validation phases, while providing for a choice of standard SpaceWire and 1553 interfaces.
• One hot face, 25% cells, 75% OSR, used at perihelion with rotation about wing long dimension, with margins versus the limit angle of 70◦ to secure the power regulation, and large overlap domain with the 100%-face. By rotating along its long dimension, it minimizes the interference with fields of view, and enables the recurrence with the BepiColombo two-sided design and drive mechanisms. The attachment of the wings to the spacecraft body is located so as to minimize the solar torque, targeting zero wheels desaturation during the 10-day observation periods. The Power Conditioning and Distribution Unit (including its maximum power point tracking, MPPT, management to optimize the effect of the variation of temperature on voltage and current), and the Lithiumion batteries are fully inherited from BepiColombo.
5.8. AOCS: guaranteeing the sun pointing thanks to the Herschel–Planck design Thales Alenia Space experience on Herschel–Planck is vital for Solar Orbiter. The two programs share indeed the absolute obligation to keep Sun pointing and be able to recover it extremely rapidly in case of anomaly, or risk of losing the spacecraft. In the case of Solar Orbiter, dispositions are taken in the thermal design of the body of the satellite to be robust to transient exposure to the 20 Solar Constant flux. However, the thermal flux is then so high that the satellite survivability relies on the swift reaction from the spacecraft. In this context Thales Alenia Space has considered compulsory to reuse the AOCS and safe mode designs from Herschel–Planck. Given the outstanding performance of the Star Tracker-pointed Herschel–Planck AOCS, this is also beneficial to Solar Orbiter in the
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nominal modes. Some units have to be adapted due to the difference in thermal environment but off-theshelf hardware could always be found that satisfied the needs of Solar Orbiter without at worst more than a delta-qualification. Nominal modes include a science mode with all payload on, use of four reaction wheels to minimize unloading and zero crossings, fine gyros and star trackers, and an in situ mode with only in situ instruments on. This latter is less demanding in terms of wheel off-loading, and subsequently uses three reaction wheels and coarse gyros. The orbit correction mode relies on star trackers, coarse gyros and 1 IN thrusters. The safe modes include an acquisition mode, a Sun-spin mode and a Sun inertial mode. The high level failure detection identification and recovery (FDIR) anomaly detection relies on a separate set of hardware to guarantee the independence of the off-pointing evaluation. Slit Sun sensors are used as anomaly detectors. The thrusters accommodation privileges force-free attitude control and minimization of contamination towards heat shield, remote-sensing instruments. 5.9. An extensive justification The spacecraft design presented above results from the thorough trade-offs carried out by Thales Alenia Space. It combines the imperative use of low-risk proven designs and the fulfillment of the challenging needs of Solar Orbiter mission in terms of stringent environment, power and data transmission. It succeeds in limiting the technological development to the heat shield itself, its baffles and doors and results into a minimized amount of delta-qualification for all the rest. Table 3 summarizes how the main features of the baselined Solar Orbiter design answer the design criteria. This design has been successfully analysis-proven. It is now consolidated by a comprehensive analytical effort addressing: • • • • • • • • • • • •
mass and power system budgets, link budget, mass memory sizing, mechanical resistance and loads, thermal sizing, thermoelastic distortions, molecular contamination, particulate contamination, magnetic contamination, radiation environment, AOCS modes simulations, wheels sizing,
• FDIR and safe mode analyses, • solar array power analysis, etc. The degree of justification has enabled to generate not only the compliance matrix and the system requirements but also the instruments interface requirements, the subsystem specifications and the units environmental requirements. The level of the study has been pushed to a degree of detail that now allows starting the definition phase upon ESA’s decision. 6. Design and development of the Solar Orbiter heat shield The design of the heat shield has been elaborated in conjunction with the system design. The evolution towards a consistent heat shield module integrating all items crossing the shield has been a natural result of the system optimization and of the maximization of the robustness. It enables the full control and validation of the thermal behavior of this critical area. The selected design has obeyed to the same criteria as for the system, with special emphasis on thermal robustness. 6.1. Heat shield architecture Its sound and simple design consists of several heat barriers separated by gaps that radiate the heat to space. It includes the following main components: • Front shield (FS), first layer that receives direct solar light. • Several high temperature heat barriers (HTHB), used behind the front shield. • Structural elements to support the FS and the HTHBs and maintain gaps between them. It is integrated and validated along with the Sun panel of the spacecraft, all doors, doors mechanisms and baffles. Different geometries, composition and combinations of the FS, HTHBs, have been envisaged. The geometry has been investigated concluding that a simple heat shield with planar front shield, separated from spacecraft by gaps, provides for very good protection. To improve this lateral heat rejection, the inner layers are highly reflective in the infrared. The optical properties of the surface facing sun become less important, and for a suitable number of gaps the conditions at the rear of the heat shield are little affected by the properties of the front. The number of gaps has been optimized
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Table 3 The baselined spacecraft design addresses the challenges of the Solar Orbiter mission. Criteria
Major features
Maximize recurrence, control of costs and risks
• Extensive heritage from BepiColombo (TTC, HGA, solar array), GMES (avionics), Herschel–Planck (AOCS design, sensors) • Modular design, with independent heat shield module • HGA deployable on North face, with mechanisms in the shadow, optimizing recurrence with BepiColombo definition while securing thermal environment of mechanisms • Platform units on anti-Sun face, favoring avionics recurrence with mild environment
Control mass, volume
• Sun face is lateral in launch configuration, minimizing heat shield mass and mechanical environment, and reducing the central area of shield
Get robust payload accommodation
• Remote-sensing instruments on South face, with high rejection capability • Additional capability on lateral faces and growth potential along launcher axis
Maximize thermal robustness
• Attitude control inherited from Herschel–Planck, ensuring swift and reliable Sun de-pointing recovery • PAS, HIS on North face, minimizing impact of large fields of view • Sun face lateral in launch configuration, reducing the central area of shield • Heat shield module with panel extension provides for extra-rejection capability
Minimize interactions radiators/appendages/fields of view
• Two operational positions of HGA • Rotating solar array reduces thermal rejection capability on East/West faces but minimizes interaction with in situ fields of view
Minimize instruments misalignments
• Remote-sensing instruments on one South facing structure, decoupled from heat shield
Maximize communication link
• Additional operational position of HGA, in shadow, guaranteeing science downlink during many close perihelia, and escape from solar array and Magnetometer interferences at any time • Increased data flow in Science orbit
Maximize uninterrupted observation periods
• Two symmetrical solar arrays on East/West faces balance thedisturbance torque • Wheels sized to avoid desaturation
subsequently, so as to provide for reasonable mass while coupling the layers and the instruments baffles to space. The selected design minimizes the heat load into the spacecraft and allows a great flexibility to trim the heat load using the sun panel extension as radiator. By minimizing the heat transfer to the spacecraft and making it weakly dependent on the front shield temperature, it becomes possible to focus on design aspects other than optical properties, such as outgassing, charge build-up, and mechanical aspects. A weak dependence of shield performance on the optical properties also provides more flexibility in choice of materials and increases the robustness of the design. 6.2. Heat shield thermal layers The thermal design of the HS has evolved through several trade-offs, including flexible versus rigid layers.
The selected design employs flexible layers. The use of flexible layers is of great advantage, as it allows easy adaptation to the many instruments apertures and doors, and decoupling with the spacecraft in terms of loads and thermoelastic distortions. The baselined design enables to select the best material for the front shield without altering the rest of the heat shield. The baseline front shield has a high absorptivity. It is insensitive to variations of optical properties, robust to temperatures beyond 700 ◦ C, electrically conductive, generates no molecular contamination and enables to control particulate contamination. Thales Alenia Space has also developed a high-performance white and electrically conductive front shield (patented), further demonstrating the robustness of the selected heat shield concept. The HTHB layers use materials selected for their resistance to high temperature and favorable optical properties.
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Mastering the effect of thermal layers attachment is key to control the performance in a high thermal load environment. Thales Alenia Space has acquired a great knowledge of the effects of such junctions in the context of very high heat fluxes. Environments up to 60 kW/m2 on transients have been experienced on Pleiades and provide for valuable lessons learned that benefit Solar Orbiter. The detailed design has been defined for flight definition and implemented in the heat shield breadboard. 6.3. Heat shield structure The trade-offs have scanned different structural concepts including frames or independent brackets. The design is not purely mechanical: if not well controlled, the structure elements can produce an important contribution to the heat load into the spacecraft. The following drivers have been considered and analyzed in the mechanical design optimization: • Ensure adequate thermal performance. • Provide mechanical interfaces sustaining the thermal layers. • Maintain gaps between layers accounting for thermal expansion. • Provide the necessary areas for instruments baffles and doors accommodation. • Decouple baffles. • Maximize accessibility. • Enable integration and maintenance of the entire heat shield. • Minimize mass. • Minimize design complexity. • Get stiffness/strength best compromise. The selected design is an excellent answer to all these needs. The structure can be easily adapted to the configuration of the instruments baffles that cross the shield, so that no constraint is transferred to the instruments and to their accommodation. The integrability is very strong, enabling to replace components in any phases of the program. 6.4. Instruments doors and baffles The remote-sensing instruments, as well as PAS and HIS analyzers, have fields of view or fields of collection that cross the heat shield. The apertures, sometimes large, require baffling. In addition many remote-sensing apertures are to be protected from contamination by doors. These latter have to be placed close to the front
shield and so be commanded by shafts that cross the layers of the heat shield. The selected design is highly adaptive to the multiple crossings. The baffles are mounted on the Sun panel so as to minimize thermal conduction. 6.5. Analytical justification The heat shield has undergone an extensive analytical evaluation, with detailed modeling and assessment of behavior under sine, acoustic, quasi-static, thermoelastic loads, and of course thermal environment. The thermal load has been assessed in a preliminary way with local thermal models. The detailed assessment has relied on an integrated thermal modeling of the whole satellite, mastering the fluxes and interfaces with instruments in a consistent way. 6.6. Development tests To secure the Solar Orbiter development prior to the confirmation of the mission, Thales Alenia Space designed a comprehensive test campaign on a full-scale representative breadboard and on samples. The chief motivation for the Solar Orbiter heat shield technology development is the harsh near Sun environment, with solar flux peaking at about 20 solar constants, and temperatures above 600 ◦ C, to be sustained in a long-duration mission, totaling about 200,000 equivalent Sun hours. Accordingly, the development has focused on demonstrating a thermal performance and thermo-mechanical stability, which can be reliably extrapolated to the desired lifetime. This is accomplished in a two-pronged approach: • Test of materials samples, to establish basic materials properties and their stability in time and environment. • Test of a representative breadboard to provide overall verification of the selected design. 6.7. Characterization testing Thales Alenia Space pre-selected the optimally suited materials from a shortlist established on the basis of our in-house experience and the experience ESA matured in the BepiColombo technology developments. The final selection has then been made in agreement with ESA, after our database has been further augmented by tests at sample level. The sample test program has been built to address the stability of the optical, thermal and mechanical properties of the materials, with measurements before and after long-duration exposure
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IR loading rig
Full-scaled Heat Shield
S/C simulator
Fig. 10. Full-scale heat shield breadboard. The breadboard includes all thermal layers of the Shield, its structure, the Sun panel and a representation of the apertures, with flight-definition junctions of all thermal elements.
to high-temperatures, along with outgassing characterization. 6.8. Breadboard testing For the overall verification, Thales Alenia Space has designed and manufactured in Turin a full-scale breadboard (Fig. 10), fully representative of the mechanical and thermal design. It includes apertures and baffles representing two types of crossings. Its fidelity to the flight design includes the representation of the junctions and attachments between all heat shield elements. The full scale is a major feature of Thales Alenia Space breadboard design. It enables: • to validate the evacuation of the heat with the real shield-to-spacecraft stand-off in a high-power thermal vacuum test; • to represent the effect of discontinuities such as instruments aperture, junctions or interface brackets in a non-disproportionate way; • to use the same breadboard to carry out representative vibration tests in later phases of the project. The test article includes a full-scale model of the heat shield mounted on a spacecraft simulator covered by MLI and controlled by heaters. The overall set-up is represented in Fig. 11. The verification of the HS thermal performance consists in measuring, in conditions representative of the extreme flight ones, the overall heat transfer to the spacecraft by all paths: conduction through the supports, radiation through the gap and
Fig. 11. Breadboard and test set-up. The breadboard is loaded in infra-red, validating not only the general concept, but also the effect of apertures and junctions, at a detailed level.
multilayer insulation and the instrument apertures according to their specific design. The determination of the optical properties of the HS in mission conditions is made at sample level, with reproduction of solar spectrum, while the overall functional performances, which are a function of the temperature, are determined using infra red (IR) sources. The front shield is exposed to a heat source that transfers heat by radiation. IR sources face the front shield. Their intensity is trimmed to give the expected in-flight front shield temperatures. Thales Alenia Space has designed and built the IR rig. Design drivers have been the uniformity of the illumination of the test article, tolerance of the high temperatures, minimization of the power dispersed inside the vacuum chamber. To reduce the lamp power, the lamp array is fitted with rear insulation to concentrate the emission on the test article, so that the power is essentially absorbed by the FS. The temperatures of the chamber shrouds are those of liquid nitrogen. The test article contains heaters for temperature control of spacecraft simulator as shown in Fig. 12. The 70m3 chamber in Thales Alenia Space Cannes is used for the test. 7. Solar Orbiter: a cost- and schedule-efficient concept In addition to the outstanding technical performance, the selected designs present an excellent cost-efficiency and a resilient schedule. Due to programmatic constraints, Solar Orbiter will indeed have a 5-year B2/C/D phase leading to a launch in 2015. This implies privileging recurrent designs and the early mitigation of the remaining technological risks.
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to minimize unexpected extra validation loops on both software and hardware test benches. The risks identified and the planning of all development, engineering, manufacturing and satellite integration activities lead to confirm the feasibility of the 2015 launch. 8. Conclusion
Fig. 12. Spacecraft simulator. The test set-up measures in a straightforward way the heat load transferred to the regulated spacecraft simulator.
The feasibility of the mission is also financial, with an important stress on the identification of cost drivers and on subsequent control and reduction of costs. Thales Alenia Space has addressed this by analyzing the cost and schedule drivers and selecting the lowestrisk design with recurring units and a robust heat shield. Technical requirements and design drive costs primarily via the robustness of pay load accommodation and mechanical–thermal configuration, the simplicity and the recurrence of avionics and AOCS, the TTC link sizing, and the solar array power demand. On all these drivers, the selected design trims costs, ensuring the financial feasibility. Robust designs, by providing margins, will also minimize potential additional iterations during the development and should be privileged as much as possible in this schedule-driven context. This motivated the selection by Thales Alenia Space of a safe system architecture protecting most parts in the shadow of the shield, with a simple stand-off-based design for the shield using robust end-of-life thermo-optical properties. The sun-pointing reliability inherited from Herschel–Planck AOCS and the recurrence of the avionics will be key
Solar Orbiter is a challenging mission, which covers a wide spread spectrum of scientific objectives with very high accuracy in a stringent thermal environment. The system configuration elaborated by Thales Alenia Space enables to reuse HGA, solar array technology from BepiColombo while satisfying the science objectives. Associated with the modular, robust and flexible heat shield design, it ensures the mastery of the upcoming phases of the development, maximizing the control of costs and schedule. The achievements in terms of heat shield development and the maturity gained on system and subsystem design and requirements enable to start the definition phase upon ESA’s final decision on Solar Orbiter mission. Acknowledgments The authors want to thank the whole study team at Thales Alenia Space for their excellent work. The work presented in this paper was performed under funding of the European Space Agency (ESA) in the frame of the Solar Orbiter Heat Shield/System Technology contract, coordinated by Mr. Ph. Kletzkine, and of Thales Alenia Space internal funding. We would like also to thank ESA’s Ph. Kletzkine, J. Marti Canales, D. Renton, Ph. Poinas and D. McCoy for the fruitful exchanges, very good relation and their support in our achievements. References [1] ESA: Solar Orbiter Heat Shield/System Technology SOW: SOLEST-SOW-56. [2] ESA: Solar Orbiter Technical Information Update Document— A: SOL-EST-IF-271.