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1.INTRODUCTION1.1Mission objectivesThe Comet Interceptor mission was selected by ESA in June 2019 as ESA’s new fast-class mission in its Cosmic Vision Programme in cooperation with the Japanese Space Agency JAXA. Comprising three spacecraft, it will be the first to visit a Long Period Comet (LPC) or even an interstellar object that is only just starting its journey into the inner Solar System[1]. All comets that have been encountered by spacecraft so far are short-period comets: objects that have approached the Sun many times, and have thus undergone changes on their surfaces, hiding their original appearance and make-up. A true, pristine comet has yet to be encountered and explored. Such objects are difficult to target because they can only be discovered when approaching the Sun for the first time, leaving little time to plan and launch a mission to them. The only way to encounter dynamically new comets or interstellar objects is to discover them inbound with enough warning to direct a spacecraft to them. The time between their discovery, perihelion, and departure from the inner Solar System has until recently been very short, historically months to a year: far too little time to prepare and launch a spacecraft. This timescale is, however, lengthening rapidly, with recent advances allowing observational surveys to cover the sky more deeply, coherently, and rapidly, such as the current Pan-STARRS and ATLAS surveys, and the Large Synoptic Survey Telescope under construction in Chile, LSST[2][3][4],. Long Period Comets are now discovered much further away, considerably more than a year pre-perihelion; e.g. C/2017 K2 (Pan-STARRS) was discovered beyond Saturn’s orbit in 2017, and will pass perihelion in 2022. From 2023, LSST will conduct the most sensitive search for new comets ever, providing a true revolution in understanding their populations, and making this mission possible. Comet Interceptor will be a new type of mission, launched before its primary target has been found. The mission will indeed travel to an as-yet undiscovered comet, making a flyby of the chosen target when it is on the approach to Earth’s orbit. Its three spacecraft will perform simultaneous observations from multiple points around the comet, creating a 3D profile of a ‘dynamically new’ object that contains unprocessed material surviving from the dawn of the Solar System. The Comet Interceptor mission was selected to combine this breakthrough in comet discoveries with a compact, agile set of spacecraft that can reveal to us a huge amount about a long period comet, ideally one is truly pristine, entering the inner Solar System for the first time. Although far rarer than long-period comets, Comet Interceptor will also have the capability of encountering an interstellar object passing through our Solar System if such an object is found on a suitable trajectory. Comet Interceptor is planned to be launched with the ESA ARIEL spacecraft in 2029, and delivered to the Sun-Earth Lagrange Point L2. It will be a multi-element spacecraft comprising a primary platform which also acts as the communications hub, and sub-spacecraft, allowing multi-point observations around the target. All spacecraft will be solar powered. The spacecraft will remain connected to each other at L2, where they will reside until directed to their target. The mission cruise phase will last months to years. Before the encounter, the spacecraft will split into its separate elements, probably a few weeks pre-flyby. For very active comets, separation will be earlier, to maximize separation of the spacecraft elements, whilst for low activity targets, separation will occur only a few days before the encounter takes place. The Comet Interceptor team comprises an international group of experts led by Geraint Jones (Mission Principal Investigator, UCL Mullard Space Science Laboratory, UK) and Colin Snodgrass (Mission Deputy Principal Investigator, University of Edinburgh, UK). Comet Interceptor was adopted by ESA during the Agency’s Science Programme Committee meeting on 8 June 2022, meaning the study phase is complete and, following selection of the spacecraft prime contractor, work will soon begin to build the mission. 1.2The CoCa instrumentThe Comet Camera system on-board Spacecraft A, CoCa, is required to provide detailed imaging of the nucleus and the innermost coma of the target. The design uses previous heritage to establish a baseline performance that surpasses that of previous fly-by missions to comets. The instrument is based upon two elements. Firstly, it uses the telescope of the Colour and Stereo Surface Imaging System (CaSSIS)[5] that is successfully operating at Mars on the European Space Agency’s ExoMars Trace Gas Orbiter (TGO). Secondly, the CoCa design uses the detector system of the JANUS instrument from ESA’s JUICE mission. By integrating these two elements, CoCa can achieve an angular scale of 8 μrad px-1, which is nearly a factor of three superior to that of the Halley Multicolour Camera on-board Giotto. The detector system uses a rolling shutter technique to allow rapid image read-out with a minimum possible exposure time of 220 μs to avoid motion smear at closest approach for even the highest velocity fly-bys. A major difference here is that, unlike Giotto which was a spinning spacecraft, Comet Interceptor is a three-axis stabilised system implying that the exposure times can be selected. The detector allows saturation of the nucleus without blooming of charge. This, in turn, implies that the exposure times of selected images can be programmed to provide high signal to noise observations of the dust coma while saturating on the nucleus. This capability will increase the flexibility of the mission if targets are eventually found that have only weak dust emission. The CoCa system will be provided by a highly experienced team from Switzerland, Germany, Hungary and Spain. Nicolas Thomas is the CoCa instrument Science Lead. 1.3The Rotating Mirror Assembly (RMA)Description & DefinitionsThe Rotating Mirror Assembly (RMA) is the complete assembly including the scanning mirror assembly and the control electronics. Each element is separated mechanically and thermally interfaced with the S/C. The Scanning Mirror Assembly (SMA) is the mechanism to rotate the mirror. It is composed of:
The Scanning Mirror Drive Electronics (SME) is the required control electronics to drive and monitor the scanning mechanism. FunctionConsiderable effort has been invested in protecting CoCa from hyper-velocity dust impacts during the fly-by. It is to be recalled that HMC was damaged severely during the 1P/Halley encounter despite being mostly behind the Whipple shield of the spacecraft[6][7]. In the case of Comet Interceptor, a rotating mirror assembly (RMA) has been developed which will allow CoCa to be mounted behind the protection shield while still providing a continuous view of the nucleus. The RMA has two elements - the SMA (Scan Mirror Assembly) and the SME (Scan Mirror Electronics). The SMA (see Figure 4) is a mechanism holding the folding mirror and that will rotate this mirror in order to orient the field of view of CoCa towards the comet during encounter. It is based on a brushless DC motor moving the mirror via a gear system and an optical position sensor in order to allow closed loop control. The mechanism will be driven by the SME that will take care of powering the motor to position the folding mirror based on encounter parameters provided by the spacecraft platform combined with the read-out of the position sensor. The SMA includes a protection system that will hide the mirror from incoming dust particles during the most critical part of the encounter, when the spacecraft is closest to the nucleus. The objective of the RMA is to rotate the CoCa Field of View (FoV) in order to follow the object during fly-by without requiring moving the spacecraft. During initial approach, while the object is angularly close to the relative velocity vector, the scan mirror is orienting the FoV along this velocity vector. The optics will be designed in such a way to minimize its optical degradation during this period. When the spacecraft will be flying-by the object (meaning the object will be angularly going away from the velocity vector up to 180 degrees), the scan mirror will rotate the CoCa field of view in order to follow the object. The Dust Protection System will protect the SMA optics from damages due to dust environment as far as possible. RMA consortiumThe RMA development project is led by the Centre Spatial de Liège (CSL), a R&D Center of the University of Liege (ULiege) and an ESA-certified Test Centre, located in Liege, Belgium. CSL has a long heritage and experience in designing, producing, verifying and delivering optical calibration subsystems and space mechanisms. The RMA development is based on experience gained on past and running programmes like COROT, CHEOPS, PACS, Sentinel-3, Sentinel-4,… The current CSL involvement in the RMA project is funded through an ESA EXPRO contract while future Phases CD will be implemented under the PRODEX Program of ESA, funded by BELSPO. The SME responsibility is shared with the Swiss Company Thales Alenia Space Switzerland (TAS-CH). This activity is covered via Swiss national funding and ESA funding independent of CSL funding. Work with TAS-CH is in close collaboration with CSL team and is also supervised by University of Bern (UBE) as Swiss institute and ESA/Prodex who placed the industrial contract. UBE is the CSL partner as part of the consortium (UBE is PI of the CoCa instrument). 2.INSTRUMENT DESIGN2.1Overview and design driversRequirementsIn terms of optical performances, the following topics are considered critical:
Among the technical constraints on the assembly design and development, the following ones are considered as the main drivers:
For reasons of robustness and reliability, the design philosophy is mainly based on design heritage and lessons learnt from S3/OLCI Calibration Assembly and S4/UVN Calibration Assembly as well as from UVS scan mirror on the JUNO mission[9][10][11][12]. OLCI Calibration Assembly was also based on the heritage of the MERIS Calibration Assembly, successfully flown on Envisat mission. Design philosophy and early tradeoff on Dust Protection System (DPS) needThe role of the DPS is to protect the SMA optics from degradation during approach to the Comet. Indeed, the S/C will enter a zone named dust zone while progressing towards the comet. This means that the S/C will intercept dusts and particles with a potentially high relative velocity (up to 70 km/s). The exact environment naturally remains unknown and the protection design shall be based on available experience. During this period, the CoCa instrument will look along the relative velocity vector to observe the comet, exposing the first optics to this dust/particle flux and potentially to an intense degradation. In order to maintain the CoCa instrument input optics as clean as possible for the fly-by observation, a protection system was initially planned to be placed at the front of the RMA, looking to the forward direction of the system. This system shall:
In addition, the DPS will:
Three solutions were envisaged:
The periscope option were preferred on previous mission, for example the Stardust mission from NASA[13][14]. The drawbacks of this solution are:
The first impression on the window solution is that there is an increased risk of break-out following an impact due to the brittle nature of transparent glasses. Moreover, break-out would eventually mean large loss of optical transmission but also protection. Some way forward could be implemented to limit these risks:
However the major drawback of these two solutions occurs when the rotation starts. Indeed, the CoCa FoV will cover the edges of the mirror/window and disturb the observations. A solution to this issue would be to increase the distance between the SMA and the DPS so that the disturbed angle is reduced. For a realistic distance of 400 mm (for accommodation on S/C) and a pupil of about 150 mm, the angle would be between 25 and 30 degrees, which is reached about 25 s before closest approach. Assuming that the clear view through the DPS would be lost at rotation angle of about 1 deg (815 s before closest approach), 790 s of data would be deteriorated, i.e. most of the science data during approach. In order to limit the duration of the degradation to about 15% of the observing time (which is the general approach of allowable science time loss), an angle of a few degrees would be required, which leads to distance of about one meter between SMA and DPS. This configuration is not affordable for S/C accommodation. Another solution to reduce the effect of the DPS on the FoV when the movement starts would be to move the DPS out of the FoV. However, this solution would also come along with several drawbacks:
Following the above discussion, it has been decided to consider the complete removal of the DPS and replacing it by a baffle/shield on the SMA. Indeed, the SMA would anyway need a baffling solution in front of it to protect, first partially, then completely the scan mirror from the incoming dust. In additional to this, it is known that the maximum density of dust will be received very close to the encounter point. In other words, the DPS would efficiently protect the SMA only when the environment is the less damaging. The idea of the DPS-less solution was to design the scan mirror front baffle not only as an optical baffle but also as a dust shield. When the mirror and its shield start rotating, the baffle is hiding partially the mirror up to complete protection from dust impacting along the relative velocity vector. This solution would imply defining the evolution of the baffle length as a function of the desired angle at which we require full protection. The selected angle, combined to a dust model will then define the amount of particles we can allow reaching the mirror. The evolution of the baffle length as a function of the full protection angle is depicted in Figure 2. In order to evaluate the amount of dust that would reach the mirror, the Engineering Dust Coma Model of the Comet Interceptor mission (issue 4.1) is used. From this model, only the particles with an impact probability higher than 1 % are taken into account (trials have been made with other threshold probability without a real impact on the baffle length conclusion). The value covering the 50th percentile is taken into account. The number of particles impacting the mirror takes considers the projected surface of the mirror and the evolving partial protection. The top edge of the shield is assumed to be linear and scans a circular mirror projection. Figure 3 shows the amount of protection for the different bins of the dust model relative to no protection (value at 90 deg). It can be seen that we have a mostly linear behaviour. An evaluation of the damage as a function of protection angle has been performed based on review of literature assessing the relation between particle size, particle impact speed and damage (crater) size[15][16][17][18]. Figure 4 testifies of an approximately liner behaviour of the obscuration with respect to the protection angle. Please also note that no protection on the mirror implies an obscuration of 2.3 %. A last important factor to be taken into account is the added torque the baffle will impinge on the mechanism. Indeed the shield, when rotating and capturing dust, will accumulate some energy and convert it into a perturbation torque on the mechanism. This torque will be at its highest when the mirror rotated at 90 deg. Figure 5 shows the evolution of the torque with the full angle protection. As expected, the longer the baffle, the larger the torque. A fast decrease is to be noted at about 25-30 deg. This torque seems to remain acceptable at first sight but shall still be taken into account in the mechanism dimensioning. As a conclusion to this design trade-off, Table 1 introduces the different solutions that shall be compared. Table 1.Trade-off description of the solutions
Based on Table 1 the no-DPS option is the preferred solution. This latter will largely solve the mass problem and maximize the scientific observation time. The negative point is a perturbation torque on the mechanism (but it may not be completely absent from the other solutions if the optical baffle is long) but this can be taken into account in the mechanism design. The degradation of the mirror shall be further refined but a first realistic hypothesis is to assume it will remain below 1-2 % of obscuration. 2.2Subsystems descriptionScanning Mirror Assembly (SMA)The SMA will be split into two parts:
Such a distinction is made in a way to separate development and manufacturing of the “dirty” part that is the mechanism that will go through a preliminary bake-out after assembly and the “clean” part, the optics that will be separately pre-baked to ensure cleanliness. Optics assemblyThe Optics assembly is composed of the scan mirror and its mount on which will be integrated some baffling to protect the system from straylight and an additional plate that will be used as dust shield in conjunction with the optical baffle. Thicknesses, distances and materials are optimised to resist at best to the estimated dust impacts while limiting the mass. As defined in the Section 2.1, this baffle length is such that the mirror will be fully protected from particles when the SMA has rotated by 30deg. The SMA Optics Assembly is depicted in Figure 6. The goal of the scan mirror is to flip the line of sight of CoCa towards the velocity vector during approach and to rotate the line of sight of CoCa during fly-by to ensure that the Comet image remains stable on the detector. This mirror will be exposed to the harsh dust environment and as such shall be optimized in order to minimize the damages during approach. The baffle will surround the scan mirror, ensure light tightness at interface on the mechanism, and reduce straylight when observing the comet. Mechanism AssemblyThe purpose of the SMA is to rotate the CoCa Line of sight during fly-by. The states of the mechanism are the following:
The concept of the driving assembly is driven by the fact that the optical beam is coming through the interface plate. The driving assembly is based on an O-configuration bearing doublet (as for OLCI and UVN mechanisms) surrounding the optical beam. As for UVN mechanism, a custom solution will be developed for the bearings with the provider. The pair of bearing will share a single outer ring that will include a fixation flange to our structure. This way, the differential thermal expansion between the ball bearings (stainless steel) and the surrounding structure (aluminium alloy) will be controlled and taken away from the bearing tracks to avoid potential high friction. The inner rings remain separated in order to allow hard preload at the right value. The shaft (interfacing with inner rings) will be made of titanium allow to limit differential expansion. A large spur gear of similar diameter is placed on the rotor and is activated by a smaller spur gear on the rotating shaft of a motor. The number of teeth of these spur gears shall be refined but current concept is 273/21, giving a ratio of 13:1. A trade-off for the motorisation type has led to the selection of BLDC motor in order to fulfil the stability/jitter requirements. In order to avoid being lost in case of any glitch in the system during fly-by and to avoid performing a homing to redefine the actual mechanism position, an absolute position sensor shall be selected. In order to be absolute, the position sensor cannot be placed on a secondary pinion on the main gear (because of the multiple rotation that would occur on the position sensor). In order to limit risks of damages on the motor and on position sensor read-out head, they have been placed behind the structure with respect to the encounter relative velocity vector. The motor is also very low on the structure meaning that it will be partially hidden in the thickness of the mounting panel. Scanning Mirror Electronics (SME)The SME architecture includes the following functional blocks:
The main functional chains of the SME are
The SME functions are distributed over the following main components: The SMA contains redundant position sensors (encoders), redundant temperature sensors but a non-redundant motor. In order to save mass the SME is not fully redundant on unit level, but provides limited functional redundancy: Fully-redundant components: Functionally redundant components:
Non-redundant components: Performance simulationsTo evaluate the performances of the mechanism, a MatLab Simulink model was established. It simulates the dynamic mechanic behavior of the system. This model takes into account for the PID control of the motor, the bearings, the gears and the friction in each element. This model is used to evaluate the performances of the mechanism and once it is correlated with a physical model, it allows also extracting some useful parameters that cannot be obtained by direct measurements. Rotating mechanism accuracyThe mechanism accuracy is driven by three different error budget: the Absolute Pointing Error (APE), the Relative Pointing Error (RPE) and the Absolute position Knowledge Error (AKE). The APE refers to the error in position with respect to the target to be tracked. A large error means that the target might be out of the Field of View of the instrument. The evaluated APE is lower than 0.05° in plane and 0.07° out of plane. The RPE refers to the error in position due to instabilities during the tracking. A large error means that the images taken by the CoCa will be degraded. The current RPE is lower than 1.24”. The AKE refers to the error in position on the knowledge. Following the encounter, the images taken will be aggregated with the tracking data. A large error means that the images will not be referenced correctly wrt. the spacecraft. The current AKE is lower than 0.03°. Structural analysisThe structural analysis is composed of the usual analysis for space systems: Modal analysis, Quasi-static load, Sine load, Random load, Shock load, Thermo-elastic deformation. The computations are done in the SAMCEF software which is the usual Finite Element Analysis software used by CSL on all projects. The current design has passed the Preliminary Design Review (to be closed within the following month). Some design modifications will have to be included for the next phase. Thermal analysisA thermal model was created in NASTRAN to evaluate the temperature that will be reached on the instrument. The mechanism is exposed on one side to the direct space environment and to the other side it is interface with the spacecraft and the CoCa. The environment is then very cold for the system and temperatures as low as -90°C are currently anticipated on the structure. The need and feasibility of survival heaters is under evaluation to maintain an acceptable temperature around the driving assembly. Optical analysisThe optical analyses for the RMA consist in a straylight analysis. A model was made in ASAP and used in two cases : during the approach and during the fly-by. To limit the straylight in the system, the baffle was truncated. This allows reducing significantly the reflections at the entrance of the tube. The different sources taken into account are the reflections of the Sun and scattering of the surfaces. During the approach, the worst case will be having the Sun at 45° from Line of Sight (LoS). During the fly-by, the mirror will be rotating so that the Sun will have less impact. Nevertheless, the Sun can go up to 135° from LoS which means that when the RMA will reach 180° at the end of the fly-by, the aperture may be exposed. 3.BREADBOARD ACTIVITIESThe breadboarding activities consist in dust impact tests on mirror samples, the development of a SME breadboard, a SMA breadboard and performance tests on both. The different activities are detailed hereunder. 3.1Breadboarding activities general objectivesSMA and SME BBThe SMA BB is used to validate the design and driving concepts of the mechanism. It will follow a development sequence including the following steps:
Mirror coating samplesThe Mirror coating samples goal is to select and validate the best concept for resistance against dust environment that will be encountered in the vicinity of the comet.
3.2Dust impact testsThe rotating mirror will be exposed to (relatively) incoming flux of dust when entering the Comet dust coma. Models, as previously introduced, have been built to predict the amount of particles of various sizes that the S/C should encounter. During the fly-by, the scan mirror will be rotated and a baffle will progressively hide the mirror from incoming particles flux. After rotation of about 30°, the mirror is fully hidden for the rest of the fly-by. Baseline coatings are based on state-of-the-art coatings already qualified for space and documented in papers about similar missions (Giotto, Stardust) and on experience of qualification of mirror coatings for space at CSL[19][20]. Coatings shall be reflective in the visible and near infrared wavelength range from 400 nm to 975 nm. The reflectivity shall be higher than 85% and a degradation of the order of 1-2 % of the transmission after exposure to dust environment is considered acceptable. The three coatings described in Table 3 were selected for the breadboarding tests. The aluminium coating (type 1 in Table 2) was chosen on basis of the heritage of the Giotto mission. The MgF2 protection layer thickness was adjusted to have a sufficient protection to space environment while being thin enough to limit the reflectivity reduction. Unprotected aluminium has a minimum reflectivity slightly lower than 85% around 825 nm at 45° (~86% at 8°). With the selected thin MgF2 thickness, a reflectivity curve (at 8°) similar to the one presented in the article about the reflectivity of the Giotto mirror1[21][22]. With this protected aluminium coating, it is however difficult to obtain a reflectivity higher than 85% around 825 nm and an incidence of 45° which is the requirement for the mission. The silver coatings (type 2 & 3 in Table 2) were chosen to have a better reflectivity on all the wavelength range of interest (90-98%, minimum at 400 nm) and two different protection layer were tested. Table 2.Coating types
Using ESA EDCM model and taking into account the movement of the scan mirror, the amount of particles expected on the mirror is shown in Table. Table 3.Dust particles expected on mirror
Computations are based on 50th percentile of the model. The dust impact velocity defined by the relative velocity between S/C and Comet is expected between 10 km/s and 70 km/s. Dust density is expected between 0.3 g/cc and 1 g/cc. Two different facilities were used to perform the dust impacts:
At the University of Colorado, two different tests were conducted:
For these tests, no impact can be seen with naked eyes. Using a microscope, small craters with limited radius can be seen. The roughness and the reflectivity of the samples are unchanged compared to measurements performed prior the impact tests. At the University of Cranfield, six shots with the LGG were performed. For each coating type, a shot with 30 μm particles and another with 100 μm particles was done at a velocity of 5.1 km/s. For the shots with 30 μm particles, a lot of degradation was seen on the sample. It seems that the particles aggregated together and made large craters in the sample. For the shots with 100μm, the degradation can also be clearly seen but the craters are smaller and more spread. The samples are currently being returned to CSL for further analysis (crater dimension, reflectivity and roughness measurement). 3.3Mechanism testsDynamic measurementOne of the most important performance parameter is the target tracking accuracy. This parameter can be evaluated by comparing the commanded trajectory to the achieved trajectory. As the EGSE is able to record the encoder data at a frequency of 1 kHz, the dynamic behavior of the CM will be characterized based on the encoder measurements during the motion. These measurements allows characterizing:
The simulation model will be adjusted to fit with test observations and will be used to provide data that cannot be directly obtained by test, like the micro-vibration torque. An example of trajectory tracking result is shown in Figure 12 along with the measured position error. The maximum error is currently around 30 μrad which is higher than the requirement of 3 μrad. Further analyses and optimizations are required to reduce this error. Static measurementStatic angular measurements are made to verify the accuracy and repeatability of the mechanism. For the breadboard, only theodolite measurements will be used. The measurement accuracy for this kind measurement is around 4” which is much larger than the resolution of the encoder (0.3”). This measurement will only allow identifying large deviation from the commanded position. The measurement is made using an optical cube mounted on the rotor of the mechanism. When a transfer is done between cube faces, a total error of 12” applies. For small motions of +-0.5 and +-1° around fixed points, the accuracy of the motion was within the error of the theodolite measurement except for some points that were on the edges of the mirror and difficult to measure accurately. Motorization marginsTo evaluate the motorization margins, the maximum input current is limited in the EGSE. A trajectory is then injected in the mechanism and the tracking performance is evaluated. The minimum input current which allows a good tracking of the trajectory is then compared to the nominal current to extract the motorization margins. In ambient conditions, the minimum current to reach a good tracking of the trajectory is 50% (margin of 1). In Figure 14, a comparison between 40% and 50% limit current is shown. For the first, the trajectory is not followed completely. For the second, the trajectory is well followed. It is expected to have larger margin for the following models because the motor used for this breadboard is not the nominal one. Because of procurement delays, a commercial motor with less torque was used. Vibration testThe breadboard was not designed to withstand the vibration loads. Hence only a low level sine test was performed for each axes. The comparison with a Finite Element Model (FEM) is not possible because the model is not in the same configuration as the tested model. The vibration test showed a first eigenfrequency around 266 Hz for the main peak. The expected first eigenfrequency for the complete mechanism is 125Hz. Thermal vacuum testThe thermal vacuum test is used to verify the functionality of the mechanism in the cold environment. Different temperature steps were made at 20, 0°C, -20°C, -50°C. As the breadboard was using a different motor than the nominal one, it was expected to have difficulties moving the mechanism at the lower temperature. For each step, a measurement of the trajectory tracking performance is done at different speeds and different current. The minimum current allowing for a motion gives an indication on the motorization margin. The following results were obtained: Table 4.Minimum current during TVAC
The test was stopped after the -50°C step as it was not possible to activate the mechanism anymore. On the previous step at -20°C, the mechanism already showed some misbehavior at all current levels. The tracking of trajectory was not possible with the highest values. It is also noted that the performances of the mechanism were already degraded at 20°C when under vacuum and that the current could only be reduced to 90% when subjected to an environment of 0°C. The decrease in performance is quite sharp and is a good lesson learned from this breadboard. Once back at ambient, the performances were measured again and the nominal levels were achieved. Lessons learnedFrom the dust impact test, the biggest difficulty was to find a facility capable of performing the required tests. On top of this difficulty, once the tests are performed, some additional limitations are found and compromises had to be made in terms of particles velocities, sizes and counts. From the static and dynamic measurements, the lessons learned were mainly on the usage of the SME to control the SMA. Being the first time both were connected together, some parameter tuning had to be made. Additionally, the SME is controlled using scripts that needed to be developed for the specific use at CSL. During the Thermal Vacuum test, the mechanism showed some, larger than expected, performance degradation. Improvements of the mechanical design in terms of thermo-elastic behavior are already under evaluation to improve the performance at lower temperatures. 4.DEVELOPMENT PLANVideo and audio files can be included for publication. Table 3 lists the specifications for the mulitimedia files. Use a screenshot or another .jpg illustration for placement in the text. Use the file name to begin the caption. The text of the caption must end with the text “https://doi.org/doi.number.goes.here” which tells the SPIE editor where to insert the hyperlink in the digital version of the manuscript. 4.1General aspectsPreliminary development activities are made in Phase B, which includes the SMA, SME and mirror coating breadboarding phases. Major developments are performed in phase C/D, including:
Remaining developments are concerning recurrent productions: 4.2Model philosophyThe proposed model philosophy for the RMA is given in Table 4. Table 4.RMA model philosophy
4.3Verification approachVerification will be performed at a maximum of levels. Major components will be submitted to an independent verification by the supplier versus their own specifications. Major components are: All other parts will be verified by any of the methods mentioned here under before its integration on the assembly (metrology for mechanical parts, roughness and wavefront measurement on optics, …). An assembly level verification program is then run on the full assembly. This is performed at qualification level for the EQM to validate the design and at acceptance level for the flight models to validate the construction. 5.SUMMARYThis paper presents the current status of the RMA, a rotating mirror mechanism that will fly on the ESA’s mission Comet Interceptor. The RMA is a mechanism rotating a mirror which ensures that the Comet is kept within the FoV of the CoCa instrument during the closest part of the approach. This instrument takes advantage of the CSL’s experiencesgained on past and running programmes, both scientific and earth observation, like COROT, CHEOPS, PACS, Sentinel-3, Sentinel-4,… The Phase AB was successfully ended with the mission adoption, testifying of the satisfying results of the breadboard activities run through the entire program. At RMA level, the engineering analyses were validated by several tests including coating resistance to the harsh comet environment and an environmental campaign. The RMA is built by a European consortium including Belgium and Switzerland. 6.ACKOWLEDGEMENTSThe Comet Interceptor RMA project will be developed under the auspices of the ESA’s Prodex Programme thanks to the sponsorships of Belgium through the Belgian Science Policy (BELSPO) and Switzerland. REFERENCESSnodgrass, C., & Jones, G. H.,
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