Solar Wind Magnetosphere Ionosphere Link Explorer Spacecraft

SMILE (Wang and Branduardi-Raymont 2020) is a joint ESA and CAS space science mission, which aims at deepening our understanding of the interaction of the solar wind with the Earth’s magnetosphere by capturing global images of the dayside magnetosheath and cusps of the magnetosphere, and the aurorae at the North Pole simultaneously, while monitoring the in-situ plasma environment (see Fig. 1). The mission was adopted by CAS in November 2016 and by ESA in March 2019.

Artistic view of the SMILE mission

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The SMILE Mission will use novel soft X-ray imaging technology to obtain, for the first time, the global image of the large-scale structures of the geospace. This is critical for the quantitative analysis and understanding of the global features of the magnetosphere.

The interaction of the solar wind with the Earth’s atmosphere leads to the formation of the large structures of the magnetosphere, including the bow shock, magnetopause, and the cusp regions. The positions and shapes of the magnetopause and cusps change constantly as the Earth’s magnetosphere responds to the varying solar wind dynamic pressures and interplanetary magnetic field orientations. Both the fast and slow solar wind can be interrupted by large, fast-moving bursts of plasma called interplanetary Coronal Mass Ejections (CMEs). When a CME impacts the Earth’s magnetosphere, it temporarily deforms the Earth’s magnetic field, changing its direction and strength, and induces large electrical currents; this is called a geomagnetic storm, and it is a global phenomenon. The southward interplanetary magnetic fields, as presented in a CME event, could induce magnetic reconnection in the Earth’s magnetotail; this process accelerates protons and electrons downward toward the Earth’s atmosphere, where they form the aurora, resulting in substorms.

The Scientific objectives of the SMILE mission are summarized as: (1) explore the fundamental modes of the dayside solar wind/magnetosphere interaction; (2) understand the substorm cycle; (3) determine how CME-driven storms arise and their relationship to substorms. (Branduardi-Raymont et al. 2018)

The responsibilities of the Project among CAS and ESA have been defined in the early stages in parallel to the phase 0, which were formalized and agreed upon through the SMILE Implementing Arrangement signed by ESA Director of Science and NSSC Director General in May 2017, and more details were specified at the Project-level in the SMILE Joint Project Implementation Plan. Detailed interfaces at Instrument-level are further clarified in the respective Interface documentation, see Sect. 7.1. The high-level responsibilities are defined as shown in the figure below and summarized hereafter (see Fig. 2).

CAS-ESA responsibilities in the Joint SMILE Mission

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CAS is responsible for the development of the satellite Platform (PF), the Spacecraft system-level testing during the Assembly, Integration, Verification and Test (AIT) campaign, In-orbit Tracking and Telemetry and Command (TT&C) in S-band, Science Application System (SAS) as well as Ground Support System (GSS). It also provides the MAGnetometer (MAG), Light Ion Analyzer (LIA) and Ultraviolet Imager (UVI) instruments. ESA is responsible for the development of Payload Module (PLM), the Spacecraft test facilities and services during the Spacecraft AIT campaign, the launch vehicle, launch site, and science operations and ground receiving station in X-band and support in S-band when necessary, as well as the development of the Soft X-ray Imager (SXI) through a Consortium of European institutions led by the University of Leicester (UK), and the image intensifier and mirrors of the UVI instrument.

3.1 Key Design Driver

The key design drivers for the SMILE satellite are the science orbit and requirements from scientific instruments. The scientific orbit is a Highly Elliptical Orbit (HEO) with an apogee of 121,000 km(18.99Re) and a perigee of 5000 km(0.78Re). The launch vehicle injects the spacecraft into a Low-Earth Orbit. Then, the spacecraft uses its own Propulsion Module (PM) to reach the final scientific orbit. The maximum mass capability of Vega-C is about 2250 kg for the required injection of 700 km × 700 km LEO and nominal 70-degree inclination.

Among the key requirements of the scientific instruments are the pointing targets and maximum allowed pointing errors of SXI and UVI, the magnetic cleanliness requirement of MAG which propagate to the rest of the spacecraft and instruments and a rotation manoeuvre of the satellite for the calibration of MAG. The satellite surface remnant electric field could affect the plasma direction measurement of LIA. Therefore the absolute potential of the satellite surface should be less than +30 V and differential potential should be less than 1.5 V.

• Science-driven selection of the final orbit: SMILE seeks to investigate the dynamic response of the Earth’s magnetosphere-ionosphere to the solar wind. This is achieved through simultaneous in-situ measurements of solar wind, magnetosheath plasma, remote X-ray images of the magnetosheath and cusp region, and remote UV images of the northern aurora. The in-situ measurements identify the upstream drivers in the solar wind and magnetosheath; The X-ray and UV imaging provides a monitor of the downstream response in the magnetosphere-ionosphere. To fulfill its scientific objectives, the SMILE satellite is designed to orbit in a highly inclined trajectory. Several reasons dictate the need for a high-inclination orbit. First, this permits simultaneous imaging of the magnetopause at the dayside point and high-latitude auroras. A larger apogee results in an extended duration from a vantage point above. Second, the science objectives require an orbit that traverses the magnetopause into the upstream magnetosheath and solar wind. Such orbit enables in-situ measurements at the adjacent upstream of the magnetosphere, increasing the timing accuracy of geomagnetic driver to within 1 minute (Dai et al. 2023). Accurate timing is crucial because global-scale magnetosphere convection responds to the solar wind within ten minutes (Dai et al. 2024). The ___location of the magnetopause and bowshock boundary can be roughly estimated from an empirical model (Lin et al. 2010; Jelínek et al. 2012). For instance, we could consider a simple parabolic model of the magnetopause and bowshock (Kobel and Flückiger 1994). Assuming a standoff distance of 10Re and 12Re at the subsolar point respectively, the geocentric distance of magnetopause and bowshock at X = 0 are 14.1Re and 18.3Re. The geocentric distance should be larger when the spacecraft is near the apogee in the nightside. This indicates that the ideal apogee should exceed 14-18Re. However, the apogee should not be excessively large as it would decrease the spatial resolution of imaging measurements. Consequently, the scientific detection orbit of SMILE is defined as a high-inclination, high elliptical orbit (HEO) with an apogee of approximately 19 Re. This carefully chosen orbit configuration maximizes SMILE’s scientific observation capabilities while considering factors such as orbit stability, observational time windows, and payload constraints.

• Final science orbit and Vega-C maximum mass performance: The key driver for the design of the satellite is its dry mass and the maximum propellant mass that can be embarked on the PM as required for the orbit transfer phase. The baseline operational final orbit satisfying the science requirements is a Highly Elliptical Orbit (HEO) around the Earth with an apogee of 121,000 km above the North pole and a perigee of 5000 km below the South pole. The S/C of around 2250 kg total mass will be injected into an orbit of 700 × 700 km at 188° RAAN (Right Accession of Ascending Node) with an inclination of 70° by Vega-C, a launcher evolved from Vega.

• SXI Pointing Requirements: The subsolar magnetopause is the region where the solar wind first encounters the Earth’s magnetic field. Observing the direct response of magnetopause to solar wind driving is the main target of the SMILE mission. Therefore, the centre of SXI is mainly oriented to the subsolar magnetopause. Normally, the magnetopause is about 10 Re away from the centre of the Earth (X = 10 Re) under average solar wind conditions. In a compressed state, the subsolar magnetopause is around X ∼ 8 Re with solar wind flux ∼ 4.9 × 10⁸ /cm² s. the Dayside magnetopause can retreat to the geosynchronous orbit or even closer to the Earth during strong storm events. To better observe the solar wind-magnetosphere interaction, the SXI near apogee mainly points to X ∼ 8 Re. This gives an offset of 24° from the SXI line of sight to the UVI line of sight (pointing to the centre of the Earth) in the spacecraft X_S/Z_S plane and towards the spacecraft +X_S direction.

• UVI Pointing Requirements: According to the scientific objectives and requirements of UVI, UVI will image the entire auroral oval in a characteristic wavelength overcoming the contamination by the dayglow with spatial resolution of ∼100 km or better. According to the design of UVI, the angular resolution is 0.04°, equivalent to 35 km to 85 km on the Earth’s surface when the altitude of SMILE satellite increases from 30,000 km to 121,000 km. A pointing accuracy better than half pixel, i.e. 0.02°, is targeted. The pointing accuracy of UVI is determined mainly by two factors including the measuring errors of the transformation matrix from the UVI reference cube to the satellite reference cube and the errors of satellite attitude. The measuring errors of the transformation matrix are determined by the theodolite, which is usually smaller than 10 arcseconds, or 0.003°. Therefore, to realize the final pointing accuracy of better than 0.02°, the measuring errors of satellite attitude should be smaller than 0.017°.

• In-flight Calibration of MAG: The zero-measurement of the fluxgate magnetometer is subject to drift in long-term operation. The in flight calibration is the only way to ensure long-term data accuracy. Operating in orbit, the spacecraft generates magnetic fields which also manifest themselves as an offset lumped together with the instrument zeros in the magnetometer measurements. In-flight calibrations are designed to find the magnetic field measurement offsets, which is the sum of the sensor zero and the static spacecraft magnetic field. Sensor rotation in an environmental magnetic field is a common practice for calibrating the magnetic field instrument zeros. To calibrate Smile MAG in flight, the mission is scheduled to perform magnetic calibration maneuvers (roll maneuvers) once every ∼2 month. A set of rotations about two different axes separated greater than 45 degrees is sufficient to uniquely determine the sensor zeros plus spacecraft field as long as the spacecraft magnetic field does not vary during the maneuvers. The periodic MAG calibration sequences are designed to provide ∼5 rotations about one axis, followed by another ∼5 rotations about the second spacecraft axis, with a rotation rate not slower than 0.5 degrees per second, and each set of 5 rotations requires about 60 minutes to complete.

• Orbit Transfer: The initial orbit of the SMILE satellite is a Low Earth Orbit (LEO), while the target orbit is a Highly Elliptical Orbit (HEO) with an apogee altitude of 121,000 km and a perigee altitude of 5000 km. The PM will transfer the satellite from the initial orbit to the target HEO. The design of orbit transfer should follow the principles below: (1) The fuel consumption should be within the fuel budget. (2) The orbit design needs to balance the orbit maneuver times and the fuel loss. (3) The orbit design should meet the constraints of the ground stations. According to the design principles, the orbit transfer strategy is to raise the apogee first and then the perigee, and no active control is carried out for the Argument of Perigee (AoP) and the RAAN. Besides, to balance the number of orbit maneuvers and the fuel loss, the orbit transfer is decomposed into eleven maneuvers. To increase the solvability, the idea of multi-step approaching the target has been adopted. After iterations, the orbit transfer scheme satisfying the observational constraints of the maneuvers by the ground stations is achieved, as shown in Fig. 3.

  

  Orbit transfer of the SMILE satellite

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• MAG Magnetic cleanliness Requirements: Magnetic cleanliness of the spacecraft is essential for the accurate measurement of the weak ambient magnetic field. According to the mission science requirements, at the outboard Mag sensor, the spacecraft generated static magnetic field must be less than 10 nT; magnetic disturbances in the frequency range (1 ∼ 40 Hz) should be less than 100 pT peak-peak; and magnetic transients should be less than 1 nT. A program has been implemented to control the remnant magnetism of the spacecraft. During the design phase, a magnetic model of the spacecraft has been built up, which accounts for the layout and magnetic moments of the platform and the payload module. Fed by the magnetic and geometric parameters from all subsystems, the model has been used to evaluate the spacecraft generated magnetic field and to identify the major sources contributing to the magnetic field of the spacecraft. Magnetic field control requirements were issued to every subsystem, as one of the guidance for the subsystem design, for example, the use of non-magnetic materials or special design of circuits to lessen the contribution to the spacecraft generated magnetic field. The magnetic model is maintained, verified, and updated throughout the mission development at subsystem level. The final achievement of the magnetic field control program is to be verified by the magnetic compatibility test on the flight model of the spacecraft. Ground magnetic field tests at spacecraft level aim at characterizing the spacecraft magnetic field both in DC and AC. Mitigation measures, such as magnetic compensation, would be adopted in case of non-compliances. The ground test data will also be used as references for the future in orbit calibration and data processing.

• LIA Requirements to the Surface Electric Field of SC: The S/C will be charged when it is immersed in the plasma environment. The S/C surface charging potential will deflect the ambient ion trajectories and change the ion energy measured by LIA instruments. These effects bring measurement errors on the plasma moments calculation. So, LIA requires the S/C to do specific potential control design to minimize the surface potential in flight operation. From the science requirement, LIA should measure the ions with the energy range of 50 eV to 20 keV. Since the S/C does not have an active potential control instrument and the positive potential is more severe on ions measurement, the surface charging requirement focus on the positive charging and the surface differential charging. So, LIA requires the S/C absolute potential (the potential between S/C surface and the absolute ground at infinity) to be less then +30 V and the differential potential between any two conductive S/C surface points to be less than 1.5 V.

• Platform and PLM interfaces: The SMILE platform is developed and manufactured in IAMC in China (Shanghai) and PLM is developed and manufactured in Airbus in Spain (Madrid). Due to the inherent difficulties of such a set-up, the interfaces were defined from the start to be as simple as possible. More details on this can be found in Chaps. 5 and 7.1. SMILE configuration on space as showed as Fig. 4.

  

  Artistic view of the SMILE spacecraft in flight configuration

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3.2 Spacecraft Overall System Design

As already mentioned, SMILE consists of a single spacecraft, made up of two main elements, a PlatForm (PF) and a PayLoad Module (PLM). The PLM is accommodated on top of the PF through a shared base plate, part of the PLM, to reduce the mass requirement and ensure a mechanically simple interface between PLM and PF shown as in Fig. 5.

Elements of the SMILE spacecraft

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The 3-axis stabilized PF provides the S/C resources, and raises the orbit from the insertion orbit (low Earth orbit, LEO) to the science orbit (highly elliptical orbit, HEO). It also hosts one science instrument, the Light Ion Analyzer (LIA), which will determine the basic moments and the distribution function of the ions.

The PF consists of two modules, a Service Module (SVM) and a Propulsion Module (PM). The PF provides the following functions:

1. (1)

   Electrical power supply for the S/C;

2. (2)

   Attitude and orbit control for the S/C;

3. (3)

   Tracking, Telemetry and command (TT&C) through the S-band transponder for the communication between S/C and ground stations;

4. (4)

   On-board data handling (OBDH) for the S/C;

5. (5)

   Structure support for the PF;

6. (6)

   Thermal control for the PF and the PLM survival;

7. (7)

   Structure interfaces with Launch Vehicle (LV), PLM and LIA’s;

8. (8)

   Thermal and electrical interfaces with the PLM and the LIA’s;

9. (9)

   Electrical interface with Launch Vehicle (LV).

The PLM is equipped with a PLM extended control unit (PLM-ECU), an X-Band subsystem (XBS), a thermal control subsystem (TCS), and a Harness Subsystem, to support, operate and manage the instruments and their scientific data. The PLM hosts the other three instruments including:

The Soft X-ray Imager (SXI), which will map spectrally the Earth’s magnetopause, magnetosheath and magnetospheric cusps; The Ultraviolet Imager (UVI), which will capture the images the auroral regions; The Magnetometer (MAG), with its deployment boom, which will measure the magnetic field simultaneously with the LIA.

The SMILE S/C layout is designed as shown in figures below (see Figs. 6 and 7). To obtain the best FoV performance, two units of the LIA are placed on the PM structure.

PF layout

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PF layout (PM accommodation)

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The main characteristics are as below (see Table 1).

3.3 Working/Operational Modes

The S/C operation modes are defined based on the agreement on the CAS-ESA Joint Mission Requirements Document, summarized in Fig. 8.

S/C Operation Mode and Transitions Between the Modes

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The Launch & Acquisition Mode is a mode used from Pre-launch phase up to S/C Sun acquisition.

The Safe Mode is a mode automatically entered when the S/C detected some pre-defined FDIR or by TC. In Safe mode, the PLM will be OFF and PF will provide the survival power for the PLM thermal control and the PF will also keep a minimum number of units ON to save power.

The Stand-by Mode is a mode used in the LEO phase (including SSO storage phase, if necessary) where the PLM is OFF. It’s also an engineering mode for the PF operation test, configuration and minor failures treatment.

The Idle Mode is a mode defined when not performing any science operation. On the other hand, the PLM safe mode or any unit’s level safe mode will trigger the S/C to enter Idle mode.

The Science Mission Mode is driven by the science mission operation requirements and can be split into 4 sub-modes: Full Science Mode (FSM) and Reduced Science Modes (RSM1/RSM2/RSM3).

The Calibration Mode allows specific calibration activities to be performed. The Calibration Mode is driven by each of the instruments and has a strong interaction with the PF activities (e.g. pointing).

The Communication Mode is the mode driven by the communication requirement. In this mode, the S/C pointing will be programmed to ground station tracking mode which allows the X-band antenna pointing to the ground station and ensures a high-speed communication to download the scientific data and S/C housekeeping data.

The Orbit Maneuver Mode is the mode driven by the orbit raising requirement. In this mode, the S/C shall turn attitude pointing to the velocity vector of orbit and fire the 490 N thruster to raise the orbit to the science operation orbit.

3.4 Product Tree

The S/C-level product tree is shown in Fig. 9. and explosive structure as Fig. 10.

Spacecraft product tree

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Spacecraft structure

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4.1 Structure & Mechanism Subsystem

The S/C structure is composed of the PLM structure and PF structure. The PF structure is a stack structure composed of the PM structure which features a large carbon fiber reinforced Polymer (CFRP) cylinder, and the SVM structure made of several honeycomb panels and CFRP frames.

The SVM main panels and frames are combined following a ‘#’ type interface with PLM, which effectively support the PLM. The S/C interface ring with LV is designed as a stranded 1194 ring made of high strength aluminium.

There are two solar arrays mounted on the SVM +Y and −Y lateral panels. The solar arrays are installed and deployed by two solar array mechanisms. The mechanism is driven by two scroll springs in the root joint and six scroll springs in the middle joints. The fixation is provided by four preloaded bolts which are separated by four pyro-cutters. The solar array mechanism is a very mature design which has been qualified in various CAS satellites.

To reduce the shock on PLM instruments and platform (show as Fig. 11) units, the solar arrays are supported by a shock attenuation structure, which can dampen the shock from the pyros very efficiently.

SMILE FM Platform

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The PF FM integration and environmental tests have been completed in September 2024.

4.2 Thermal Control Subsystem

In each orbit environment and working mode, the performance of the thermal control subsystem should meet the following requirements: (1) temperature of the battery pack 10 ∼ 30 °C, and temperature of most of the remaining equipment −15 ∼ 45 °C; (2), the propulsion system temperature range of 10-30 °C, the temperature difference between tanks of the same type less than 3 °C, the temperature difference between tanks of different types less than 5 °C; (3) decoupling of payload and platform thermal design, such that the interface operating temperature range is −30 ∼ 25 °C.

The thermal design principle of the service module mainly adopts passive thermal control method, and adds active thermal control in local areas. The propulsion system mainly adopts active thermal control method. The composition and design of the thermal control subsystem is shown in the following figures (see Figs. 12 and 13).

Composition of Thermal Control Subsystem

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Thermal Design Diagram

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The following two pictures show the satellite temperature distribution in HEO_HOT (FSM) case and HEO_COLD (PLM OFF) case, the values indicate that the temperatures are within the defined range (see Fig. 14).

SMILE Temperature Distribution Simulation

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4.3 Electrical and Power System (EPS) Design

The main purpose of EPS design is to make sure that the system energy is balanced over a single orbit and there is an adequate design margin. The design target is to ensure that the DoD of single orbit is no more than 20% for SSO and 60% for HEO in normal mode. At the same time, the output power of SA should maintain a 10% design margin.

EPS is a non-regulated bus with direct transmission of energy. In this way, EPS will lighten the weight of the subsystem while considering the requirement of residual magnetization. The subsystem will provide the required energy and make an efficient reduction in weight without BDR (Battery Discharging Regulation) or BCR (Battery Charging Regulation) component. At the same time, it makes use of 8 batteries pack in series, and the voltage of the main bus will be clamped by a battery working in the 24 to 33.6 V range.

EPS will supply the power for the load of the satellite and meet the power requirement in all working modes. EPS on the platform includes Solar array (SA), battery and power control distribution unit (PCDU). SA is the source of energy, battery is the energy storage unit. In the sun light, the SA converts light energy into electrical energy, while in the shadow, battery supplies the power for each unit on the satellite.

While charging the Li-battery, it adopts constant current and constant voltage through S3R. The constant-current charging of battery is realized under the condition of abundant power of the solar array, while battery package constant voltage charging is realized under MEA control. It will be considered that the battery is full, and the charge ends when the charging current is less than 0.03C. And the charging current can be adjusted by instruction to achieve flexible control on orbit.

4.4 Attitude & Orbit Control Subsystem

The attitude and orbit control subsystem (AOCS) is responsible for monitoring and controlling the spacecraft attitude and orbit during all mission phases after separation from the launcher, in particular for high pointing accuracy during science observation. The major requirements are as follows: (1) Stabilization of the spacecraft after separation from the launcher; (2) Orbital maneuver from LEO to HEO; (3) Autonomous attitude acquisition and determination; (4) Stabilization and control of the attitude in 3 axes with science operation pointing requirements; (5) Attitude maneuvers to allow the calibration of MAG.

The performance requirements for AOCS subsystem are as follows (see Table 2).

The configuration of the AOCS is as follows (see Table 3).

During the science operation phase, the PF attitude is driven by the guidance law and controlled by RWs with PID algorithm under inertial coordinates. The simulation results (see Fig. 15 and Fig. 16) show that PF AKE is 0.02°(3\(\sigma \)) and RPE is 0.03° per minute (3\(\sigma \)), which meets the science requirements.

Simulation Results of PF RPE During Science Operation Phase

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Simulation Results of PF APE During Science Operation Phase

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4.5 Propulsion Subsystem

The propulsion subsystem adopts bipropellant and dual-pressure mode system. The propellant tank capacity is up to 1650 kg. Oxidizing agent is MON-1 and fuel is MMH. The same propellant is stored in two 380 L surface tension tanks. Pressurant gas is helium that is stored in a 90 L composite gas cylinder. The system configuration is 1 × 490 N engine for orbit maneuvering and 12 × 10 N thruster for attitude control. The performance of the propulsion subsystem is as follow (see Table 4).

The propulsion subsystem provides the force and torque required for the AOCS of the satellite. The main functions of the propulsion system are as follow: (1) After separation, the 10 N thrusters provide torque capability to carry out rate damping. (2) During orbit maneuvering phase, the 490 N engine provides force capability for orbit transfer. The 10 N thrusters allow to maintain the right attitude during the period of main engine ignition. (3) Throughout the mission, the 10 N thrusters provide the necessary torques to unload the momentum of RWs.

4.6 TT&C Subsystem

The onboard TT&C subsystem architecture is shown in the figure above and consists of: (1) Two USB transponders, working in the same frequency bands, each with 8 W High Power Amplifier (HPA) inside; (2) One Microwave switch; (3) One 4-port RF network and one 6-port RF network; (4) Two LGA TX/RX antennas and two HPA TX antennas; (5) One GNSS receiver with one low gain RX antenna and one high gain RX antenna (see Fig. 17).

Onboard TT&C subsystem Diagram

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In LEO orbit, the LGA TX/RX antenna and the low power transmitter (0.5 W) are used to perform TT&C communication at TM data rates of 8 kbps. The S/C-ground communication period is 30 min/day and usually divided into 4 passages (see Fig. 18).

TT&C Passage Plan in HEO orbit

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In HEO orbit, the LGA TX/RX antennas and the receivers are used to perform TC uplink communication, while the HPA(8 W) and the HPA TX antenna are used to perform TM downlink communication with varying TM data rates (2 kbps, 4 kbps, 8 kbps, 16 kbps) depending on the S/C-ground distance, shown in table below. The S/C-ground communication period is 60 min/orbit and divided into 8 arcs located at different S/C-ground distances as shown in Table 5. A red dot represents a communication arc. The TM data rate is switched by telecommands planned by the ground system.

In transfer orbit, it will follow LEO operation mode when orbit height is below 9000 km, while HEO operation mode is triggered above 9000 km.

All the FM units passed the acceptance tests and have been delivered. All the electrical function tests and the TT&C working mode tests were completed. S-band board to ground RF Compatibility Tests (RFCT) (see Fig. 19) with both Chinese and European stations were successfully completed.

TT&C Units Delivery Test and S-band RFCT with ESA Ground Station

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4.7 On-Board System Management Unit (SMU)

SMILE on-board Space Management Unit consists of an On-board Computer (OBC) and the software running on it. The processor of OBC adopts SPARC V8 architecture AT697, with a main frequency of 80 MHz, 128 KB of Bootloader Processor PROM, 4 MB of EEPROM, 4 MB of SRAM (EDAC), and 256 MB of SDRAM (EDAC). The architecture of OBC is shown below (see Fig. 20).

Hardware Architecture of OBC

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The SMU subsystem has completed its functional tests (see Fig. 21) and is now integrated with all other SC units.

S/C Functional Test on Table

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4.8 Spacecraft Flight Software (FSW)

The spacecraft flight software manages primarily the platform functions based on hardware design. It adopts a centralized management method with the PF SMU software as the core to perform functions, including comprehensive management of satellite, space-ground communication, energy management, thermal control, orbit calculation, the attitude and orbit control, orbit transfer control, and payload task management. PLM software, however, separately manages the payload module elements. The data flow diagram of the spacecraft Flight Software System is shown in the following figure (see Fig. 22).

The data flow diagram of the spacecraft Flight Software System

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The functions of the FSW include: (1) Comprehensive management of onboard products, measure external information with various hardware components, obtain relevant analogue and digital quantities, and process data with corresponding software to obtain instructions and data output; management of the satellite operation modes according to the satellite flight program, and autonomous powering on and off of onboard equipment, safety mode management, etc; onboard time management, which will be managed by the On-board data handling subsystem. (2) Space-ground communication: Complete uplink signal demodulation, synchronization, data parsing and instruction execution, as well as downlink data packaging, modulation, and transmission. (3) Energy management: Achieving the acquisition and evaluation of the working status of battery and power control system, over discharge protection, and corresponding fault handling. (4) Thermal control: Collect temperature within the satellite, execute the heater switch commands, and complete loop control of heaters. (5) Orbit calculation: Perform orbit data extrapolation based on GNSS output orbit data, ground injected orbit data, or theoretical orbit data. (6) Attitude and orbit control: Obtain the satellite attitude and perform attitude and orbit control according to the mission and platform requirements. (7) Orbit transfer control: The satellite enters the orbit transfer phase according to the orbit transfer requirements and performs orbit control according to TC from ground stations. (8) Payload task management: The platform software and payload software collaborate to complete observation tasks: SXI imaging of the interaction area between the solar wind and the magnetopause, UVI observation of the auroral region, in-situ LIA and MAG observations. The platform software provides the resources required for satellite observation tasks.

The procurement, development, verification and delivery of the Payload Module (PLM) is under the responsibility of ESA. An industrial phase A and B1 took place between 2016 and 2018 involving three parallel and competitive PLM contracts, leading to the PLM Preliminary Design Review (PDR) in early 2019. Following the PDR, the ESA Science Programme Committee adopted the ESA contribution to the SMILE mission, which includes the provision of the PLM, the launch vehicle, the spacecraft FM AIT facilities and services, part of the operations and the management of the Project. Airbus S.A.U. (Madrid, Spain) has subsequently been selected at PDR, for the PLM implementation phase (phase B2, C, D, E1). The following describes the PLM design and Assembly, Integration and Test plans.

5.1 PLM Thermo-Mechanical Sub-System

The SMILE PLM is designed to be physically the top panel of the spacecraft. It is highly integrated as the top panel of the Service Module with approximately 120 bolted interfaces. The PLM total mass allocation is restricted to a 126 kg maximum including three PLM-mounted instruments: MAG, SXI, and UVI. The panel, an aluminium honeycomb composite structure approximately 1 m², employs a high-density core (130 kg/m3) that, along with block and edge inserts, provides sufficient stiffness for accommodating the instruments and providing the required alignment stability. This stiffness and the thermal behaviour were tested in 2022 with a Structural Thermal Model (STM) to qualification levels. A view of the testing is provided in the figure below (see Fig. 23).

PLM STM during vibration testing (Courtesy of Airbus)

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The accommodation density of the PLM is quite high with eleven physical units and the associated data, power, and RF harnessing all accommodated on the top panel of the spacecraft (see Fig. 24). The PLM employs 2 heat pipe networks and an MLI tent supported by secondary structures to maintain the thermal environment within a range of −20 ° to +55 °C. The 2 heat pipe networks are logically split between PLM electronics units (+X) and instrument electronics units (−Y), each connected to a dedicated radiator for evacuating heat and with a dedicated set of heaters to mitigate cold scenarios. The SXI and UVI are each accommodated along the −X edge of the PLM to provide these instruments with radiator surfaces which allow them to provide the necessary thermal conditions for their detectors. The MAG boom is accommodated on the +X edge to avoid any proximity conflicts with the spacecraft solar arrays. This ___location also provides sufficient room for the 3 mounting fixtures for the launch configuration. The X-band antenna is mounted towards the −X edge to ensure that communications do not violate the solar illumination constraints imposed by the instrument thermal environments.

PLM layout (without MLI) (Courtesy of Airbus)

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The thermal environment is maintained through 3-tier thermistor monitoring of the PLM by the Platform computer, the PLM computer (Extended Control Unit or ECU), as well as the SXI and UVI instruments all of which have full redundancy in monitoring and heater control. The Platform ensures that the lower survival limits are never exceeded, while the PLM provides control to make sure the units’ operating conditions to be met. The instruments provide the fine control for their detector environments to ensure stability throughout the operational orbits as needed.

Another function of the PLM thermo-mechanical system is to ensure the PLM-level alignment and pointing requirements for the SXI, UVI and MAG. The SXI and UVI lines of sight are required to be offset from each other by 24° in the Z/X plane, to allow the SXI and UVI to observe their respective targets as illustrated in Fig. 25, while respecting the required sun exclusion angles defined by the SXI baffle and the thermal requirements on the radiators. The platform AOCS ensures that these sun exclusion angles are fulfilled during observations.

Illustration of SMILE pointing targets and Sun exclusion angle

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The PLM is an interface point between the instruments and the spacecraft. Therefore it aggregates the requirements on the spacecraft pointing in terms of modes and performances. These requirements are then synthesised across the instruments into a single hypothetical instrument based upon the spacecraft target in each pointing mode. There are 5 possible pointing modes and their targets have been considered:

1. 1.

   SXI-driven: 20.3° ahead of the earth as shown in Fig. 25 above,

2. 2.

   UVI-driven: pointing towards +80° Earth latitude,

3. 3.

   Star Starring mode: the satellite points at a fixed target in the celestial sphere,

4. 4.

   Ground Station pointing: the X-band antenna is pointed at the relevant ground station, and

5. 5.

   Idle: the solar panels are pointed at the sun.

To meet the scientific objectives of the instruments, 3 performance values need to be considered: APE, AKE, and RPE. The absolute pointing error (APE) is the maximum (3-Sigma) acceptable error between the intended target and the actual spacecraft pointing. The Absolute Knowledge Error (AKE) is the difference between the estimated and the actual spacecraft pointing. The final element is the relative pointing error (RPE), which is the variation allowed over a specific time basis.

The following pointing performance are required by the instruments in order to meet their scientific objectives (see Table 6).

5.2 PLM Electrical, Software and Functional Sub-Systems

The SMILE PLM is designed to be as self-contained as possible, keeping the interfaces to a minimum, allowing as much independent verification as possible. This design philosophy is to simplify system testing and verification between the PLM and the platform, and streamline the testing in Europe and China as far as possible. The redundancy design philosophy is to provide full cold redundancy with no cross strapped functions, to reduce the possibilities of single point failures. The various electrical and functional interfaces were tested on an Electrical Functional Model (EFM).

The PLM receives all of its power from the platform, and the commands from the ground via the platform S-band link or directly from the platform with respect to spacecraft Fault Detection Identification and Recovery (FDIR). The PLM ensures its own safety by implementing its own FDIR including PLM safe mode.

The PLM manages its own mission timeline which is onboard independent from the platform timeline. To ensure system operation, this is coordinated on ground through the command planning process. The PLM Extended Control Unit (ECU) is a highly integrated computer unit which manages the PLM Mission Timeline, power distribution to the PLM units and instruments, stores the science and housekeeping data for the entire spacecraft, and manages the X-band downlink. All digital communications within the PLM are point to point using either RS-422 standard or spacewire. The following figure shows the data and power architectures (see Fig. 26).

PLM electrical architecture

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The PLM aggregates all the housekeeping data from spacecraft system, platform and PLM, as well as instruments (LIA, MAG, SXI, and UVI), stores them within the Mass Memory Module (MMM) flash storage in the ECU within dedicated packet stores. When in visibility of a ground station the PLM transmits this data at 150M Symbols per second to the ground via the X-band chain. The MMM has a storage capacity of approximately 260 Gbit, which is divided between various packet stores. This storage capacity is more than sufficient to store the design requirement of 40 Gbit for all spacecraft data collected during a 51-hour orbit, and it is sufficient to recover from one missed transmission pass without loss of data and there are significant margins beyond these requirements.

The ECU also houses the General Processing Module (GPM), which houses the computer core of the PLM. This computer core loads following a power-on, and its basic configuration is assured with a pair of high-powered command lines which ensure that the correct configuration with the platform is established. This occurs within the Boot Software, which hands off the operations of the PLM to the Service Software which is responsible for its basic functions, and its higher-level functions through an Application Software (ASW). The ASW is based around the simple architecture of a cyclic scheduler, which runs on a one-eighth of a second cycle and manages all the functions of the PLM. The PLM software operates on the basis of 3 modes, i.e. transition to safe, Safe, and Operation. The software modes, interactions and transitions can be seen in the following figure (see Fig. 27).

PLM software mode transitions

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The PF and PLM ensure that the On-Board time is synchronised between all units to within 100 ms through the exchange of a precise timing signal transmitted via a dedicated RS-422 standard link at a frequency of once per second, called a pulse per second (PPS) line. The ASW manages all the operational functions of the PLM including the PLM Mission Timeline, commands loaded onto the PLM for execution when the spacecraft is not in direct communication with the ground.

The PLM ECU is responsible for distributing the power and monitoring consumption via current measurements for all units on the PLM via the Power Distribution Module of the ECU. The switching and protection functions are implemented through Latching Current Limiters (LCLs) which ensure that any short circuit faults do not endanger the PLM or the entire spacecraft.

The PLM power consumption is monitored by the ECU and is limited by requirement to be less than 155 W on average throughout an orbit. While peak power demand is driven by the short duration deployment of the MAG boom, it is less than the required 500 W over a 100-second duration. In general, no heater power is needed when all the units are on, as their dissipations are sufficient for keeping the PLM warm. When units are powered off, there is capacity to replace the missing heat via heaters on the respective heat pipe network.

5.3 PLM Model Philosophy, Assembly, Integration and Testing

The PLM model philosophy allows verification of the Payload Module as independently as possible with respect to the satellite PF and up to a certain level. The extensive use of additional development models and simulators at all levels focuses on the early verification of critical interfaces, within which extensive verification of Instrument and PF interfaces will be performed. Taking this objective into account, together with the level of maturity of the proposed solution, the following PLM model philosophy has been established:

• Structural and Thermal Model (STM), including instruments STM’s. This model was primarily used to verify the mechanical and thermal design of the PLM. PLM STM activities are completed during which PLM-level thermal and mechanical tests took place. The PLM STM was also integrated on the PF QM in China and took part in the satellite-level QM test campaign in 2022, thereby verifying the thermo-mechanical interfaces with the platform.

• Electrical and Functional Model (EFM), including instruments simulators and EM’s. This model was primarily used to verify the electrical interfaces and functional interface within the PLM, including instruments, but also with the PF QM units, which were tested at Airbus on two occasions for a system-level functional test. The PLM EFM activities were almost entirely completed and the remaining tests will be conducted in parallel to the PLM PFM AIT.

• Proto-Flight Model (PFM)

Of course, PLM-level GSE’s were developed as necessary. It is also worth noting that a PLM simulator was developed, which allows functional representation of the PLM, and was delivered to all instruments and platform to anticipate interface tests with those entities. An RF suitcase, using the X-band transmitter was also produced and used for all tests with the relevant X-band ground stations, including at DLR (Germany), responsible through ESA of the X-band main station O’Higgins in Antarctica, and at Sanya (China), used to complement O’Higgins.

The PLM PFM AIT is now going on smoothly. All flight elements have been assembled and integrated mechanically and electrically, including PLM primary and secondary structures, (see Fig. 28) ECU, X-band chain, antenna, radiators, heat pipes. The last version of the PLM SW is available and the final validation at PLM-level ongoing. The MAG FM and UVI QM are integrated and going through functional testing. The SXI flight model is delivered in June 2024 and finally the UVI flight model in November 2024.

PLM flight model status (June 2024)

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Following integration of SXI FM, final functional testing will be performed. The PLM thermal test will be conducted thereafter. This will be done with UVI QM due to late arrival of UVI FM, ensuring high representativity of the configuration. Finally, the UVI FM will be integrated at the end of the PLM PFM AIT campaign, just before the start of the spacecraft-level AIT campaign at ESTEC. It is also worth noting that the originally scheduled PLM-level mechanical tests will not be carried out to mitigate a schedule risk. The PLM mechanical workmanship is verified via other means, such as independent verification of the torques of the various units at regular intervals throughout the PLM AIT campaign.

6.1 Soft X-ray Imager: SXI

The soft X-ray imager (SXI) is developed by the University of Leicester with contributions from other institutes in the UK, Austria, Spain, Norway, Switzerland, the Unites States and China. SXI will observe the ___location, shape, and motion of dayside magnetospheric boundaries, including the bow shock, magnetopause, and cusp regions in the soft X-ray band ranging from 0.2 to 2.5 keV. SXI will also measure the time-dependent solar wind composition from the SWCX X-ray diagnostic emission lines.

The SXI instrument (see Fig. 29) is a wide field lobster-eye telescope with X-ray optimized CCD detectors. The focusing optics is formed by two 4 × 4 arrays of square pore micro channel plates (Micro-Pore Optics, MPO) that are slumped to a common nominal radius of curvature of 600 mm. The individual 40 × 40 mm² MPOs are 1.2 mm thick, with a pore cross section of 40 × 40 \(\mu \)m² and a wall thickness of 12 \(\mu \)m. The internal surfaces of the individual micropores are coated with 25 nm iridium to enhance X-ray reflectivity. A total of 32 MPOs are mounted on a spherical frame with a radius of curvature of 600 mm which, by the basic geometrical principles of the lobster-eye telescope, results in a nominal focal length of 300 mm.

Left: Photograph of the SXI PFM. The Front-End Electronics box will be mounted under the telescope. Right: Cross section of the SXI telescope showing the main components of the instrument

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The SXI focal plane consist of 2 large format, back-illuminated CCDs. Each CCD, procured by ESA and supplied by Teledyne e2v (UK) contains 4510 × 4510 native pixels with a pixel size of 18 \(\mu \)m. A fraction of 1/6 of the active area near the read-out end of each CCD array is used as frame store area through which an “asymmetric” frame transfer read-out with a factor of 36 on-chip pixel binning is realized. In conjunction with the geometry of the micro-pore optics, this configuration provides a field of view of 15.5° × 26° for the SXI lobster-eye telescope.

Due to the wide field of view of SXI and the requirement for a pointing direction relatively close to Earth, the instrument requires an optical baffle extending beyond the actual telescope to minimize the straylight potentially reaching the focal plane. The telescope design also features optical blocking filters of 100 nm Al placed on top of the individual MCP arrays to block local as well as astrophysical background emission. Despite the CCD arrays being optimized to reduce the sensitivity to ionizing radiation, the expected radiation damage still necessitates the integration of a shielding door in front of the sensors to protect the CCDs during the passage of the radiation belts, which occurs once every orbit. The associated door mechanism will be operated before and after each radiation belt crossing.

6.2 Ultraviolet Imager: UVI

Simultaneously with SXI observations of the solar wind interaction zones, the UVI instrument (see Fig. 30) will provide ultraviolet images of Earth’s Northern aurora (Aurora Borealis) in the 160–180 nm section of the Lyman-Birge-Hopfield (LBH) band of molecular nitrogen. Implementation of new filter technologies and an innovative telescope design enables the instrument to provide images of the full auroral oval throughout the year at all local times even in fully sunlit conditions. UVI is developed by the National Space Science Center of Chinese Academy of Sciences (NSSC, CAS) and the Beijing Institute of Space Mechanics and Electricity (BISME) with contributions from Belgium.

Photograph of the UVI Qualification Model. Electronics box (UVI-E) to the left, UVI Camera (UVI-C) to the right

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The image-forming section of the camera comprises a four mirror on-axis, all-reflecting telescope. This type of reflecting telescope is very compact and has excellent light gathering power. Moreover, it has excellent resolution across its circular field of view of 9.9-degree diameter and has a high throughput in the far-ultraviolet part of the spectrum. Thin film filter coatings are deposited on each of the imaging mirrors to provide signal filtering and definition of the science wavelength band. Further filtering is accomplished via the individual components of the detector.

The solar-blind detector system comprises a micro-channel plate (MCP) based image intensifier, optically coupled to a CCD detector. The telescope optics take images of the light onto a MCP through an input window coated with a suitable photocathode material (Cesium Iodide, CsI). The emitted photoelectrons are amplified within the MCP and converted back to light again at the exit of the MCP by a layer of P43 phosphor. These photons, now at visible wavelengths (emission peak 545 nm), are then re-imaged onto the CCD sensor by a compact relay lens optics.

The detector array itself is a 1024 × 1024-pixel CCD with a pixel size of 13 \(\mu \)m. It is operated in either 2 × 2 or 4 × 4 binning mode resulting in final science images of 512 × 514 or 256 × 258 pixels respectively, with a special resolution compliant with the required 150 km from spacecraft apogee over the entire 9.9-degree field of view.

The combination of the four filtering surfaces, the MgF₂ entrance window and the photocathode of the image intensifier can accomplish an out-of- band rejection required to take the images of the aurora in fully sunlit conditions, and provide the required sensitivity of ∼20 Rayleigh in a 60 s exposure.

6.3 Magnetometer: MAG

The magnetometer MAG (see Fig. 31) is developed by the National Space Science Center of CAS. The scientific goal of MAG is to establish the orientation and measure the magnitude of the solar wind magnetic field. This information is of critical importance for understanding and interpreting the X-ray imaging data. The magnetometer will be used in combination with the in-situ plasma measurements to detect interplanetary shocks and solar wind discontinuities passing over the spacecraft. The interfaces that MAG will measure are typically moving at hundreds of km/s in the solar wind, and consequently it only takes a few seconds to pass over the spacecraft. The magnetometer is therefore designed to measure the magnetic field at a sampling rate of up to 40 Hz and with an accuracy of 0.1 nT.

MAG boom flight model in folder configuration before integration on the PLM

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MAG consists of two individual sensor heads, each of which comprises three ring-core fluxgate sensors oriented along three orthogonal axes, thus sampling the three spatial components of the local magnetic field. The fluxgate sensor design is known to exhibit good stability and has extensive space heritage. Both sensor heads are mounted on a 3 m, three segment deployable boom, at a distance of 2.2 m and 3.0 m from the root flange of the boom respectively. The boom itself, heritage of the Chinese FengYun-4 mission, is primarily made of CFRP and other non-magnetic materials, to minimize residual magnetic fields. It will be launched in folded configuration with two thermal knife-based hold-down and release mechanism launch locks. A set of two closed cable loop mechanisms will synchronize the three boom segments during deployment.

The front-end electronics for each of the two sensor heads, including sensor drive and read-out circuits, as well as the digital processing, power supply and communication systems, are located in a dedicated electronics box, mounted on the PLM. For deployment of the boom, the boom release devices will be powered by the PLM directly.

6.4 Light Ion Analyzer: LIA

The in-situ plasma analyzer LIA (see Fig. 32) is developed by the National Space Science Center in China with contributions from the UK and France. LIA will determine the basic movements of the solar wind and magnetosheath ion distributions, such as density, velocity, temperature tensor, and heat flux vector under different solar wind conditions over 4\(\pi \) steradians.

Left: Operating principle of the LIA sensor head. 1- Electrostatic Analyzer, 2 - MCP Anode Array, 3 - Field of View Deflection System, 4 - Variable Geometric Factor System. A potential ion path is indicated in red. Right: The two LIA instrument flight models

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The LIA is designed as a top-hat type electrostatic analyzer (ESA), consisting of two concentric, hemispherical shell sections, which only permits particles of a selected energy and charge to reach the detector. The selection of the particle energy is realized by applying a negative high-voltage to the inner hemisphere of the analyzer. Charged particles of selected energy that pass through the analyzer section are detected though an annular micro-channel plate (MCP) in front of a set of 48 uniformly segmented anodes. The maximum angular resolution in azimuth therefore is 7.5° but can be reduced to a minimum of 30° by summarizing the signals from four adjacent anodes.

By design, the ion analyzer is sensitive to an instantaneous field of view of 360° × 3°. The LIA sensor head is also equipped with a field of view deflector system (FDS): by applying a high voltage to either the upper or lower deflector electrodes of the FDS, the incoming plasma ions can be steered from a selected arrival direction into the analyzer section. The instantaneous field of view can in this way be scanned in the elevation direction by varying the high voltage applied on the deflector electrodes, and a total field of regard of 2\(\pi \) steradians can be achieved. To realize the required full 4\(\pi \) sky coverage, two identical electrostatic analyzers are mounted at opposite sides of the spacecraft platform.

In addition to the field of view deflection system, the LIA design also includes a “top-cap” electrode to provide a variable geometric factor system (VGFS). The VGFS allows the variation of the instrument sensitivity and increases the dynamic range of the sensor by up to two orders of magnitude, and therefore allows the optimization of its response to the ions in the different plasma populations encountered in the SMILE orbit.

6.5 Instruments Summary and Budgets

A summary of the main instrument parameters is provided in the following Table 7.

7.1 Joint Design, PA Activities and Reviews

The overall set-up is rather innovative due to the nature of the collaboration which is unprecedented between CAS and ESA on that scale. However, the overall Project Management principles follow a very classical and well-proven scheme for Space projects (see Fig. 33). A small-size Project team is present both at CAS/NSSC and ESA/ESTEC. Those two project teams manage their respective contributions independently through contracts or agreements with the companies or institutes in charge of the development, but communicate on a very frequent basis to discuss the necessary matters, anomalies, issues, long-term plans, etc. through progress meetings and teleconferences. Remote ways of communications have obviously increased and improved efficiency throughout the Covid period. Typical space project management tools were used, allowing to establish and control the mission-level schedule (see example below) and risks. A Steering Board has been defined at the onset of the project to resolve key issues, but thus far, it has never been solicitated.

Extract of SMILE high-level Mission schedule

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In the early stage of the mission development (phase 0), the joint responsibilities were defined at high-level with the objective to define preliminarily but clearly the interfaces between the PLM and the Platform as well as all other interfaces between CAS and ESA as simple and independent as possible to enable an efficient development by both Parties of their respective elements. This led to the joint responsibilities defined in Chap. 2.

The phase 0 consisted of concurrent engineering sessions involving both ESA and CAS, partly at ESTEC (Concurrent Design Facility: CDF) and partly online, where the overall mission design was born, and the principles of the mission were established based on the initial scientific mission proposal submitted to both CAS and ESA. Following this process, the CAS-ESA Implementing Arrangement and Joint Project Implementation Plan were established to govern the management, delineate responsibilities between the two Parties for development of the mission, define the documentation to be exchanged at each step of the development up to launch and in-orbit operations, and finally specify the review process. The documentation and reviews were streamlined in a way to be consistent with a fast-track approach to ensure that both Parties preserve their know-how and only exchange the strictly required technical interfaces-related information. At the end of phase 0, a first draft of the science requirements was produced, as well as a first mission concept.

The preliminary system design definition then took place throughout the phases A and B, during which both parties elaborated separately on the preliminary design of their elements. This process, of course, was conducted while iterating jointly at mission-level the overall mission design to ensure consistency in the elements and to develop the mission and system requirements, as well as the detailed interfaces across those elements. At an intermediate step at the end of phase A, a joint CAS-ESA System Requirements Review took place, thereby confirming the overall mission and system requirements.

The preliminary design phase eventually led to two major blocks of activities and documentation that governed the remainder of the development, up to this day:

• System engineering: mission requirements document is based on the consolidated science requirements, the system requirements, and all other system engineering related documents, e.g. system engineering plan, Joint spacecraft AIT plan, mathematical model exchanges, coordinate requirements, margins requirements, etc.

• Interface engineering: intensive co-engineering in phase B2 took place, after the selection of the PLM Prime Contractor, where the joint interfaces were defined in detail between PLM and platform as well as between PLM and instruments, between spacecraft and ESTEC test facilities (including GSE) and finally between spacecraft and ground segment, as both Parties are involved in all mission elements. Those interfaces were formalized in relevant Interface Requirements Documents (IRD’s) and Interface Control Documents (ICD’s) or similar documents as represented by AIT plan etc. All interfaces’ changes were then managed in the next stages of development when necessary through the relevant Change Request process, to be approved by the other Party.

The above preliminary system engineering phase concluded with the successful Mission Preliminary Design Review (M-PDR) in January 2020.

Following PDR, the detailed design phase started, also known as Phase C. In that phase, the detailed design of all elements was further consolidated on the basis of the agreed mission and system requirements, documents and system interfaces. Detailed design down to material and component levels was frozen and test articles were produced and tested. The objectives and achievements of different models are addressed in this document, such as the Platform Qualification Model, the PLM Structural and Thermal Model, etc. The detailed design documentation, along with the test results of those models, enabled the spacecraft and Mission Critical Design Review (M-CDR) in June 2023 (see Fig. 34).

Joint spacecraft/Mission CDR in Shanghai

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The overall mission phase D then officially started (although it had started at lower-level as normal practice in V-model system engineering) allowing the manufacturing and testing of the flight model to proceed. This phase is now nearing completion for the individual elements as all flight models have been produced, assembled and are in their final stage of testing. The last sub-phase of the Phase D is the complete spacecraft AIT, during which the Platform and PLM will be integrated as one of the very first steps. The entire spacecraft will undergo typical functional, electrical and environmental tests as required before launch, see Sect. 7.5. That will then lead to the spacecraft Qualification and Flight Acceptance Review, now foreseen in summer 2025.

The joint review process has proven its value and effectiveness throughout. It is however worth noting that each Party conducted their own equivalent reviews, i.e. SRR, PDR, CDR and QAR, internally for the elements under their responsibilities, following normal practices internal to each organization. Conversely, the joint mission reviews are carried out in a time-efficient and concise manner with the objective to wrap-up and inform the other party of the outcomes of the lower-level reviews, address the high-level mission and system aspects, and address any other element that is not reviewed separately, e.g. space debris, launcher interfaces.

Finally, a joint PA/QA/safety process was also established in a Joint PA plan for all activities which require exchanges and consistency between the PA and safety approaches of the Parties. Documents and information outlined in this plan include typical PA and QA aspects and a minimum amount of information and joint PA/QA-related meetings were identified which must be put in place by both Parties to provide the minimum transparency: Non-Conformance Reports (NCR) and associated Non-conformance Review Boards (NRB), Procedures for joint integration and test activities, Joint Verification/Test Reports, Red-tag/Green-tag Items List, Joint Inspections Reports, failure mode analyses, list of components and material if relevant to the other party (e.g. compatibility with TVAC), safety analysis, contamination budgets, etc.

7.2 Platform QM and PLM STM Testing

As already mentioned, a Satellite QM was built by CAS/IAMC and went through an AIT program in Phase C as per the following sequence (see Fig. 35). While this activity was led by CAS/IAMC who built the platform QM and performed the tests (see Fig. 36) (during Covid period when travels to China were not possible), this phase was a major joint undertaking involving both CAS and ESA in the QM which accommodated the ESA-provided PLM and the test runs (see Fig. 37).

Satellite QM AIT sequence overview. Note: the mechanical test includes both empty tank test and full tank tests

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Test bench of the QM S/C

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Integration steps of the platform QM and PLM STM on the overall satellite QM

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During SMILE tests on the electrical and functional test bench in Phase C, electrical interface and functions of each subsystem were verified. The tests covered the satellite’s TM/TC interface, S-band, and power supply interfaces, and verified the software and hardware interfaces and overall subsystem functions. It was confirmed that the status of each subsystem is correct, the data correct and reliable, and the test coverage comprehensive.

From May 2022 to December 2022, the satellite QM AIT tasks were completed, including construction of main structure, assembly and integration of instruments by July 2022, environmental tests, EMC, and magnetic test.

7.3 Platform Structural Model Fit-Check with Launcher Adaptor

To de-risk the mechanical interfaces, an early fit-check and clamp-band release test were conducted between an initial structural model of the platform, representing the interface ring with the launcher, and the launcher adaptor of Vega-C. This test took place at ESTEC (see Fig. 38), where the PF structural model and a representative launcher adapter were delivered in early 2022. This allowed us to confirm the validity of the platform ring flatness and geometry vs the launcher adaptor. It is also useful to derive the shock transfer function towards all satellite units induced by the launcher separation event, serving an input to the shock qualification of the spacecraft. The picture below depicts the operations of the fit-check.

Satellite-Launcher fit-check at ESTEC

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7.4 Platform and PLM Electrical and Functional Testing (Early Simulators and QM Tests), RF Compatibility Tests

Following sufficient but early testing of the Platform QM units with the PLM simulator – the latter being just software-based, two rounds of interface and system tests took place at Airbus in June 2023 and February 2024. These tests utilized the PLM EFM and Platform QM units (computer and PCDU) integrated together as they would be on the flight model, to verify the electrical and functional interfaces on real hardware. Furthermore, PF FM S-band space-to-ground test at ESOC (Germany) was completed in May 2023 (vs ESA-provided S-band ground stations), and S-band space-to-ground RF compatibility test at Weinan (China) was completed in April 2023 (see Fig. 39).

Radio Frequency Compatibility tests

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7.5 Joint Spacecraft AIT Activities in ESTEC

7.5.1 Overall AIT Campaign

Following PF-level AIT in China, and PLM-level AIT in Madrid and then ESTEC, the spacecraft-level joint AIT will take place at the ESTEC Test Centre in the Netherlands. During this phase, IAMC is the prime entity responsible for the AIT activities. ESA is providing the test services via ETS (European Test Services, the operator of the ESTEC Test Centre), and coordinates all the parties involved in the campaign.

Following the transportation of the PF and all Chinese GSE’s from China to Europe by plane, it will be unpacked in the ESTEC clean-room, and various checks will take place to ensure that no damage was caused during the journey. Meanwhile, the AIT of the PLM will be completed. The joint AIT starts with the electrical assembly of the PLM with the PF using extension cables. During this “open-heart” test, all the electrical and functional (i.e. software) interfaces between the PLM and the PF are tested with flight elements for the first time, including long duration operations and failure cases. When this test is successfully completed, the final assembly of the spacecraft can take place: mechanical integration of the PLM on top of the PF, integration of the solar arrays and finalisation of the MLI around the spacecraft.

The spacecraft will then be in full launch configuration for the first time, and a fit-check with the Launch Vehicle Adapter will be conducted to confirm the compatibility. Figure 40 depicts the AIT flow for these first two phases of the AIT campaign, namely PF preparation and SC final integration.

Phase 1 and 2 of the spacecraft AIT

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The Ground Support Equipment (GSE), supporting the AIT of the SMILE spacecraft, are provided both by IAMC and ESA. The ESA GSE are all re-used from previous science projects, namely Solar Orbiter, Bepi Colombo and JUICE, to make the best use of the resources (see Fig. 41). The compatibility of some of them was checked during the SC SM and QM campaigns.

Solar Orbiter Ground Handling Adapter (GHA) and Multi-Purpose Trolley (MPT) used during SM AIT campaign in ESTEC (left); Solar Orbiter Vibration Test Adapter (VTA) and AIT Clamp Band (ACB) used during QM AIT campaign in Shanghai (right)

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7.5.2 Environmental Test Campaign

After final assembly and integration, the SMILE spacecraft FM will undergo an environmental test campaign. The flow, depicted in Fig. 42 below, is standard with a few specificities. The traditional sequence is respected: mechanical tests with solar arrays mounted, followed by thermal vacuum test with solar array dismounted, followed by EMC test with solar array mounted once again. The EMC test is conducted at the beginning of the campaign to mitigate any potential EMC compatibility issues between the PF and the PLM as early as possible, since this aspect was not tested during the SC QM campaign, where the PLM STM was not used. Following the environmental tests, a deployment of the MAG-boom will be conducted. The campaign also includes a magnetic characterization test at the end of the flow. This test will be conducted in a standard clean room, making use of specific equipment as accurate as possible, as there is no dedicated magnetically clean test chamber at ESTEC.

Phase 3 of the spacecraft AIT

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The test facilities selected for the SMILE campaign are the Large Space Simulator (LSS) for the TBTV test (see Fig. 43), the Large European Acoustic Facility (LEAF) for the acoustic test, the 320 kN DUAL and 640 kN multi-shakers for the sine vibration, and the Maxwell EMC facility for the EMC test.

Spacecraft FM configuration in the LSS

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When the environmental tests are completed, the spacecraft will be finalized before shipment to the launch site. This phase includes final functional tests, global leak test, and re-assembly of the solar panels.

The table above summarizes the tests that will be conducted at spacecraft-level (see Table 8).

7.6 System Operation Validation Test (SOVT)

During spacecraft joint AIT (see Fig. 44), two slots are foreseen to perform the System Operation & Validation Test (SOVT). The main objective of the SOVT is to validate that all the ground segment interfaces are operational to control the spacecraft and receive telemetry from it (up- and down-link), thereby establishing a firm basis for the satellite’s orbit testing and subsequent operations. The scientific data exchange and archiving are also validated during the SOVT. Those tests involve the spacecraft which is “made available” to the ground segment.

Phase 4 (last phase) of the spacecraft AIT

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Several test cases, representative of various in-flight scenario, will be run to achieve the SOVT main objective which has been subsequently split in various sub-objectives. The list presented here below is non-exhaustive:

• Validate mission planning processes and interfaces, incl. scientific protocols.

• Validate data downlink from satellite to all involved ground entities.

• Validate data uplink to satellite from all involved ground entities.

• Verify the ability to control the spacecraft in various flight operational phases (LEOP, Science modes etc…), incl. command generation.

• Validate the data format, data products and the archiving process.

Those test cases will cover various post-launch real-world scenarios, including satellite initialization, routine scientific operations, instrument parameter configuration, on-board software updates, PLM data replay and operation scenarios, MAG calibration scenarios, critical operation rehearsals, emergency operations, and satellite power-off procedures.

As shown in Fig. 45 above, the SOVT organization is fully representative of the flight operational chain:

• Instrument teams will interface with SAS for determining the science plan.

• SAS will communicate the science plan to GSS which will in turn transfer this plan in Payload Module telecommands.

• GSS will transfer the generated PLM telecommands to CCC.

• CCC will then upload the PLM telecommands (along with the PF telecommands) to the PF S-band EGSE and ultimately to the PF. This transfer is possible thanks to the ESA-provided Network Data Interface Unit (NDIU) which, in simple terms, mimics the behaviour of a ground station uploading commands to a spacecraft through cabled connection.

• The spacecraft executes the telecommands and generates 2 types of telemetry: housekeeping telemetry and science telemetry.

• The housekeeping telemetry is routed back to CCC for monitoring via the same link used to upload the telecommands (i.e. S-band EGSE → NDIU → CCC).

• The science telemetry is routed to GSS via the X-band SCOE (PLM GSE) linked to a second dedicated NDIU.

• GSS then dispatches all necessary data to generate the science products (L0 data, auxiliary files, orbit data etc…) to ESAC (for SXI) and to SAS (for UVI, MAG and LIA).

• Dummy L1-L3 products are then generated by respective science team and the archiving system in GSS and ESAC is validated.

SOVT schematics

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The complete launch services procurement is under the responsibility of ESA, the formal interface with the launcher provider. Arianespace will provide the baseline launch vehicle Vega-C. CAS/IAMC, in charge of the spacecraft system and therefore leading spacecraft preparation before combined operations with the launcher, is however involved as necessary in technical interface discussions to ensure compatibility and feasibility of all aspects.

Following the spacecraft AIT and successful QFAR, the spacecraft will be shipped by sea to the launch site (Centre Spatial Guyanais, CSG) in Kourou, French Guyana (see Fig. 46). This transport will take about two weeks.

Vega-C launch vehicle

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All interfaces with the actual launcher Vega-C and the launch site are already well consolidated through iterations conducted under a Preliminary Mission Analysis Contract between ESA and Arianespace, involving IAMC as necessary. This among others leads to the first issue of the Launcher-Spacecraft Interface Control document, trajectory analyses, Launcher-Spacecraft couple load analyses, launch campaign operations plan, first rounds of safety reviews.

All those aspects are now under finalisation through the launch services procurement contract in place since early 2024 with the objective to start the launch campaign in Q3 2025 for the launch at end of 2025. In particular, the operations plan for the launch campaign is described in the following figures (see Fig. 47).

Launch campaign sequence of operations

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At launch, the S/C will be in launch configuration where a minimum number of units are on to ensure the necessary autonomy of the satellite shortly after separation. Launch will take place and spacecraft will be injected into the targeted LEO 700 × 700 km circular 70-deg inclination orbit. After separation, the spacecraft will carry out the first acquisition – temporarily planned to be done by New Norcia (ESA ground station) –to stabilize its attitude. After verification of the health of the satellite and a number of low-Earth orbits, the transfer to the HEO scientific orbit will be performed by the PM which will raise the S/C from LEO to HEO through an estimated set of eleven manoeuvres. The transfer will take about one month. The complete in-orbit commissioning activity, including the instruments, will start after the S/C arrives at HEO. The in-orbit commissioning review will be conducted maximum 3 months after launch, which will evaluate and review the in-orbit performance of S/C and scientific instruments. The Scientific data will be archived in both ESA Space Astronomy Centre (ESAC, Madrid) and CAS National Space Science data Centre (NSSDC). Operations are nominally planned to last 3 years after launch, with a potential 2-year extension. The spacecraft is fully operated by CCC and GSS from CAS, and ESA will support the scientific data processing and archiving, provide the main X-band ground station for downloading all data, and offer S-band ground stations for support during LEOP in case of contingency.