Method and system for inspecting a constellation of satellites in low Earth orbit, and associated inspection satellite

Abstract

A method for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, characterized in that it employs an inspection satellite equipped with inspection data collection equipment, the launch of which is phased to meet the satellites to be inspected within said constellation.

Description

Title of the invention: Method and system for inspecting a constellation of satellites in low Earth orbit, and associated inspection satellite. FIELD OF THE INVENTION

[0001] The invention relates to an inspection method / system for a constellation operator in low Earth orbit. It also relates to an inspection satellite implemented in this system. STATE OF THE ART

[0002] Low Earth orbit (LEO) satellite networks have become an emerging and promising technology for providing access to wide-reach networks. Inter-satellite links (ISL) extend the reach of network service and reduce its dependence on terrestrial infrastructure.

[0003] High Earth Orbit Robotics Pty Ltd (HEO) uses sensors onboard satellites to inspect and image objects in orbit. Their growing database of flyover inspections includes observations such as the external shape, size, orientation of an object, signs of damage, color and what is on or off on a satellite.

[0004] There is thus a strong demand for the inspection of satellite constellations. However, the solutions currently available do not allow for both remote inspection of these satellites and the essential close-range inspection needed to assess their health at a lower cost. The current need to use several inspection satellites for a single constellation is hindering the development of inspection services.

[0005] The object of the present invention is to provide a new solution to the inspection needs of satellite constellations, by remedying the drawbacks of currently available solutions. Description of the invention

[0006] This objective is achieved with a method for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, characterized in that it implements an inspection satellite equipped with inspection data collection equipment, the launch of which is phased to meet the satellites to be inspected within said constellation.

[0007] The method according to the invention may further advantageously include, from the inspection satellite, a collection of local space weather data and / or electromagnetic data.

[0008] This inspection method provides both from an inspection satellite a remote inspection service of satellites to be inspected within the constellation and a proximity inspection service with respect to satellites to be inspected within said constellation.

[0009] The method is particularly suitable for inspecting a constellation of satellites with an inclination of 88°.

[0010] It further includes a step for choosing an altitude of the inspection satellite so as to phase a quasi-encounter between the inspection satellite and a satellite of the constellation at the pole of intersection of orbital planes distributed in different longitudes of an ascending node.

[0011] Once the inspection satellite is phased with a small altitude differential, a step is planned in which the inspection satellite crosses the target satellite to be inspected within the constellation, in the vicinity of the intersection pole, and then a step to take, during the crossing, a capture of a plurality of images of the satellite to be inspected.

[0012] According to another aspect of the invention, a system is proposed for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, implementing the method according to the invention, characterized in that it comprises an inspection satellite equipped with inspection data collection equipment and whose launch is phased to meet the satellites to be inspected within said constellation, and a ground station equipped with a communication system with the inspection satellite to receive the collected inspection data and means for processing said collected and received data.

[0013] According to yet another aspect of the invention, a satellite is proposed for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, characterized in that it includes equipment for measuring and collecting inspection data obtained when said inspection satellite crosses a satellite of the constellation, and a propulsion system.

[0014] The propulsion system of this inspection satellite may advantageously include an electric propulsion stage.

[0015] In-orbit inspection, beyond simply observing and imaging a target at a distance or close range, allows for the collection of "in vivo" orbital data, ranging from direct camera observation to the gathering of space weather and electromagnetic data. This inspection activity is carried out by a stellite orbiter (a nanosat-type spacecraft) equipped with electric propulsion for approach and rendezvous maneuvers, providing efficiency in terms of mass consumption and thus enabling multiple missions.

[0016] The inspection satellite according to the invention is equipped with a solid fuel storage, and a dedicated radar for the proximity approach.

[0017] The system according to the invention thus performs a "distant inspection" up to 1 km (resolution 10-50 cm), without rendezvous maneuvers and with an approach in "absolute" maneuvers of the Formation Flying type (for Formation Flying), then a "proximity inspection" which includes Rendez-Vous and Proximity Operations maneuvers to approach to 100m with a centimeter resolution.

[0018] This invention makes it possible to carry out a complete health status of a satellite, and then to assist in the deorbiting of a failing satellite carried out either by the operator or by a third-party deorbiting service.

[0019] The mission is carried out on a satellite of the OneWeb (OW) low orbit (1200 km) telecommunications satellite constellation. It is suitable for civilian commercial projects (OneWeb, Iris2), civilian institutional projects (CNES) and military projects ("watchdog" mission).

[0020] The advantage for a constellation operator is having both a remote inspection service (first level of observation and diagnosis) and a near-Earth inspection service for more specific cases. In both cases, there is a collection of local space weather and electromagnetic data, which allows both for updating local data and for potentially anticipating the design of future generations of satellites. The first generation (GEN1) of this constellation has an inclination of 88°, and the second generation (GEN2) a mix of inclinations of 88° and 55° at an altitude of 1200 km.

[0021] Access to the orbital planes for each of these altitudes was analyzed and resulted in the following strategy: - for future inclinations of 55°, the natural drift of the ascending node (-3.68° per day) allows a sweep of the orbital planes in less than 100 days, making inspection possible with an inspection satellite according to the invention, designated as the Orbiter. - For inclinations of 88° (current GNE1, 60% of GEN2), the J2 effect, linked to perturbations in the gravitational field due to the Earth not being perfectly spherical, is too weak due to the high inclination. Hence, a short-range "meeting" strategy is proposed in the present invention.

[0022] The altitude of the Orbiter inspection satellite is chosen to phase a "quasi-meeting" at the pole of intersection of the orbital planes distributed in different longitudes of the ascending node.

[0023] With a small altitude differential once phased, the Orbiter inspection satellite will cross its target in the vicinity of this "pole," with visibility first in front of it, then above, then behind it, and will therefore be able to take pictures for each orbital plane, modulo the chosen landscape strategy. This strategy makes it possible to observe all satellites with a single Orbiter inspection satellite, depending on the fuel capacity to take advantage of the altitude differentials.

[0024] Whether for 55° or 88°, the present invention therefore proposes a strategy enabling the observation of potentially all the satellites in the constellation with a single satellite, thanks in particular to the choice of electric propulsion. This strategy is crucial for the viability of the economic model minimizing the number of Orbiter inspection satellites. DESCRIPTION OF THE FIGURES

[0025] - [Fig. 1] [Fig. 1] illustrates the respective trajectories of two satellites in a model for calculating geometric parameters of inter-satellite link (ISL); - [Fig.2] [Fig.3] [Fig.4] [Fig.5] Figures 2 to 5 schematically illustrate a example of the realization of an inspection satellite according to the invention; - [Fig.6] [Fig.6] illustrates an example of the implementation of a Mission Software Interface; - [Fig.7] [Fig.7] illustrates an example of the architecture of the ground segment put in works in an inspection system according to the invention; - [Fig.8] [Fig.8] represents an example of the integration of a CSP communication protocol implemented in an inspection system according to the invention. DETAILED DESCRIPTION

[0026] The implementation of the inspection method according to the invention is carried out in two deployment phases: - a first Orbiter_V 1 satellite dedicated to a long approach (a few kilometers, compatible with "classic" maneuvers coordinated according to the position of the target) and based on a platform as common as possible with a camera but without a guidance, navigation and control (GNC) system or Radar: - then a second Orbiter_V2 satellite capable of making long approaches like Orbiter_Vl but also rendezvous maneuvers in close proximity to the target (up to a few tens of meters) based on a platform equipped with Radar and GNC autonomous rendezvous software.

[0027] The advantage for a satellite constellation operator is having both a remote inspection service (first level of observation and diagnosis) and a near-Earth inspection service for more specific cases. In both cases, there is a collection of local Space Weather data which will allow both the updating of local data and the potential for anticipating the design of future generations of satellites. This primarily involves measuring and predicting the radiation level in the constellation's orbital plane.

[0028] A possible inspection strategy within the scope of the present invention may include: - access to the satellites targeted during launch by phasing-type drift, playing on the altitude differential, - A close-range "encounter" strategy is used because the J2 effect on the natural drift of the ascending node from the other orbital planes is too weak due to the high inclination. The orbiter's altitude is chosen to phase a "near-encounter" at the pole of intersection of the orbital planes distributed at different longitudes of the ascending node. With a small altitude differential once phased, the Orbiter satellite will cross its target near this "pole" (with visibility first in front of it, then above, then behind) and can therefore take images for each orbital plane, depending on the chosen landscape strategy. This strategy allows observation of all satellites with a single Orbiter, provided there is a good compromise between timing and altitude delta to optimize the phasing. - a conventional backup strategy exists by playing on the differential altitude of the circular orbit to create a differential J2 effect on the longitude of the ascending node and aim for an adjacent plane, but this is costly in delta-V and requires time (typically 1 month of drift).

[0029] The inspection method / system according to the invention therefore proposes a strategy for potentially observing all the satellites in the constellation with a single satellite, thanks in particular to the choice of electric propulsion and the crossing of orbital planes at the pole. This strategy is crucial for the viability of the economic model, which minimizes the number of Orbiter satellites.

[0030] The paper DI “Modeling and Analysis of Inter-satellite Link in LEO Satellite Networks”, by Quan Chen et al., published in 2021 at the 13th International Conference on Communication Software and Networks (ICCSN), 10.1109 / ICCSN52437.2021.9463648, discloses a computational model for the geometric parameters of inter-satellite links (ISLs), including range and pointing direction, which are key parameters for communication performance. The variation models ISLs with satellite phase are theoretically derived and verified by numerical simulations. The propagation delay of multi-ISL paths is also studied.

[0031] In a practical application of the method / system according to the invention, the inclination and altitude of interest for launches are respectively 88° and 1200 km.

[0032] Initially, a satellite in a plane of the constellation is selected as the target, for which both distant and then close-range inspections can be performed. Other satellites in the plane can be inspected, including a passive / non-cooperative satellite, depending on the constellation's capabilities.

[0033] These inspections will be preceded by an approach around a virtual point to demonstrate the in-orbit capabilities and validate the edge-to-ground interface and the operational safety elements.

[0034] In a second step, the crossing scenario “at the pole” described above is demonstrated in orbit with a target-satellite of another orbital plane: phasing, “formation flying” type maneuvers, attitude control capability to “spot” (point) the target according to the viewing angles.

[0035] A demonstration of the drift of the ascending node's longitude (LNA) by differential altitude of the fighter and of the distant inspection once in the adjacent plane can also be provided.

[0036] Finally, the phasing-crossing and remote inspection strategy with snapshots is implemented in the vicinity of orbital plane crossings. Depending on the fuel reserves for electric propulsion, the mission continues as a commercial mission for the constellation operator. End-of-life operations can also be planned. Injection orbit and target orbit

[0037] Several launch vehicle options are considered in order to reach an injection orbit. This involves defining an altitude and a probable injection plan to be provided as input to preliminary mission analysis tools.

[0038] For its launch, the Orbiter inspection satellite can be integrated as a secondary payload and must therefore be able to reach the target orbit while respecting the constraints related to the injection of the primary payload, which will have priority. The injection orbits of the constellation satellites for recent launches are given in Table 1, as well as the typical injection for an Arianespace Vega launcher. Vega Launcher, Arianespace, Typical Dispersion Capacity: circular orbit at 700 km altitude for a 1500 kg payload. At 3°: a: ± 15 km e: +0.0012 i: +0.15° RAAN: +0.2° Falcon 9 SpaceX OneWeb 17 - 03 / 2023: injection into a 589 x 600 km orbit, at an inclination of 86.51°. Perigee: ±10 km Apogee: ±10 km i: +0.1° RAAN: +0.15° LVM-3 ISRO OneWeb 18 - 03 / 2023: injection into a circular orbit at an altitude of 450 km, inclination of 87.4°. No information found. Soyuz Arianespace OneWeb VS27 - 02 / 2020: injection into a circular orbit at an altitude of 475 km, inclination of 87.4°. At 3o: a: + 12 km e: +0.0012 i: + 0.12° RAAN: + 0.2°

[0039] Table 1 - Possible injection orbits and launcher dispersions for the inspection satellite

[0040] The injection altitudes are therefore between 450 km and 700 km with an inclination ranging from 86.51° to 87.4°. Since the altitude of the target orbits is much higher than the recorded injection orbits, it is decided to consider a pessimistic case beyond the 3° dispersions, with an injection altitude of 400 km.

[0041] Consider a constellation operating at an altitude of 1200 km, on planes inclined at 88°. The launches reported in Table 1 having targeted only quasi-polar orbits, the exact / operational injection inclination retained here is 87.4°.

[0042] Calculation of the transfer strategy to the target orbit

[0043] As illustrated previously, the injection orbit may differ significantly from the operational orbit. This necessitates the establishment of a feasible station-placement strategy, taking into account the time available for the mission and the capabilities of the platform. Since the orbiting satellite is equipped with a low-thrust engine, the transfer to the target orbit must be calculated using appropriate optimization methods.

[0044] First, the MIPELEC software (https: / / www.connectbycnes.fr / mipelec) was used to determine an initial transfer strategy without operational constraints (indirect optimization method). This tool, developed by CNES (National Centre for Space Studies), was specifically designed for optimizing low-thrust transfers in minimum time.

[0045] It should be noted that only the transfer to the target orbit is considered in this preliminary study, and that the constraints related to phasing with the target object are not taken into account. Continuous transfer without operational constraints

[0046] The assumptions of the transfer are as follows: Circular injection orbit Circular target orbit Propellant characteristics Altitude: 400 km Inclination: 87.4° Right ascension of the ascending node: 0.0° Altitude: 1200 km Inclination: 87.9° Right ascension of the ascending node: 0.0° Maximum thrust: 0.8 mN ISP: 3000 s Total mass of the inspector: 24 kg

[0047] We assume a continuous and constant thrust at the maximum value. We require that the transfer take place in minimum time.

[0048] With these parameters, the following minimum time transfer is obtained: - Mass of propellant consumed (kg): 0.345 - Transfer time (days): 146 - Equivalent AV (m / s): 423 Impact on the satellite's altitude

[0049] The amplitude of variation of the thrust vector in declination is approximately 40°. For a transfer involving an inclination increment of 0.5°, the satellite must therefore be able to vary the thrust direction by 40° in half an orbit, i.e. approximately 32 minutes at an altitude of 400 km (minimum period on the transfer).

[0050] Transfer taking into account operational constraints

[0051] It is now necessary to add to the maneuvers thus calculated the consideration of operational constraints that require reducing or stopping thrust during the transfer (energy constraints, solar panel pointing, star-tracker glare, engine rest time, etc.). A preliminary solution consists of simulating the transfer using the Orekit library https: / / www.orekit.org by activating thrust on the allowed segments, after MIPELEC optimization at the start of each segment.

[0052] Thrust is inhibited on each prohibited segment (e.g., eclipse phase or maximum thrust duration reached on the preceding segment). For this preliminary analysis, the developed analysis tools allow for the consideration of eclipse phases. New constraints may be added on this basis in the future. Phasing with the approach to the target satellite

[0053] In a phasing step with the target, this involves establishing an approach strategy once the inspection satellite has reached an altitude close to, but still slightly below, that of the target, so as to maintain a drift in A true relative anomaly. The safety clearances around the target, as outlined below, must be observed. For the initial approach after launch, the altitude gain maneuver can be combined with phasing, thus conserving propellant.

[0054] Station keeping and remote inspection in the plan

[0055] The station-keeping can take place between two satellites of the constellation (approximately 7 m / s / year) or a few kilometers from a target satellite in co-orbital offset (watchdog type mission), which allows remote inspection.

[0056] For this latter case, the observation capabilities towards and around the target of the Orbiter inspection satellite are studied. One possible scenario for remote inspection (RI) is to position oneself below the target (safe trajectory) and drift slowly, taking images before, below, and after passing in the vicinity of the target. This technique, similar in principle to the pole crossing described later below, is a delta-V-efficient and perfectly safe alternative to remote station-keeping. Avoiding collision with the target

[0057] A collision avoidance strategy with the target can be implemented for both a distant inspection (IE) and a proximity inspection (IP).

[0058] Safety volumes are defined around the target, the potential violation of these volumes to be monitored on board: collision volume (the smallest, which must never be crossed), emergency volume (which triggers an avoidance maneuver) and safety volume (which allows to reject maneuvers calculated on board).

[0059] Relative state choices are planned between the Orbiter inspection satellite and the target (Constellation satellite formation) ensuring a safe distance between the orbiter and the target even in the event of a system failure.

[0060] In onboard trajectory monitoring, the risk of collision is estimated using the distance from Mahalanobis. This provides an upper bound on the probability of collision.

[0061] An emergency maneuver is also planned to be triggered if the distance is less than the emergency volume, or if the relative propagation predicts a risk of collision, or if an alert / failure is detected on a Fault Detection, Isolation and Recovery (FDIR) equipment. Appointment Scenario (Local Inspection)

[0062] Once the Orbiter_V2 inspection satellite has been brought by phasing to a distance of 10 km–3 km from the target, the radar can be acquired and thus all the GNC guidance, navigation, and control elements using radar relative navigation can be initialized. The scenario for a rendezvous mission for proximity inspection can be initiated.

[0063] A scenario corresponds to a phase of flight in relative formation (Rendez-vous Proximity Operations), during which the fighter jet maneuvers autonomously using radar-based relative navigation, within a reference frame relative to the target. A TM / TC link is established (monitoring, ground command capability), but the entire relative formation flight plan is transmitted on board before the start of the flight. The scenario is composed of several relative flight phases. Each phase is defined according to parameters defined relative to the target, allowing for the easy creation of a trajectory by guidance, regardless of the target's orbit.

[0064] For an inspection scenario, the trajectories are defined with the offset on V-bar and R-bar (Station keeping typically at -1000m with 100m offset) and for inspections (100m).

[0065] To transition from one of these three trajectories to another, reconfiguration maneuvers are necessary. These maneuvers can be initiated from the ground or automatically on board via maximum time management per phase, as defined in the scenario. The scenario also includes station-keeping phases, as well as phases dedicated to filming. After completing the inspection, a disengagement maneuver is performed, allowing the aircraft to move safely away from the target. Safety and emergency volumes are taken into account during maneuver calculations throughout the scenario, in which case the maneuver can be aborted, modified, or interrupted.

[0066] A nominal inspection scenario, with the current parameterization, lasts approximately 23h30m in total. Phase Duration in Maneuver Mode Duration in Station Keeping Co-orbital Formation 3h45 lh50 Inspection 1 2h40 lh50 Inspection 2 lh40 lh50 Retreat 2hl0

[0067] The first phase consists of a reconfiguration maneuver and station-keeping, to reach a co-orbital state 1000m behind the target along the along-track axis.

[0068] The second phase consists of joining an inspection orbit centered on the target. Images are also taken when the orbit is stabilized.

[0069] The third phase is similar, joining a second inspection orbit around the target.

[0070] When the operations are completed, a parting maneuver allows one to move away from the target in a safe manner.

[0071] In this scenario, the reconfiguration or initial approach phase starts at the Radar acquisition at 3km.

[0072] Orbit transfers once in operational orbit (multi-target case)

[0073] Several targets can possibly be inspected in the same plane. Possible orbital transfers are analyzed. The insertion of the inspection satellite takes place in a plane of the satellite constellation, for example the OneWeb constellation operating at an altitude of 1200 km on circular orbits, whose planes are separated by 15° in longitude from the ascending node, and whose inter-satellite angle is 7.2° (50 satellites per plane, i.e. a distance of approximately 953 km at this altitude). Transfer between two satellites on the same plane

[0074] Inspecting several satellites in the same plane is a viable option that is less costly in terms of propellant mass. This simply involves modifying the inspector's altitude to create a drift in relative true anomaly. An estimate of the cost of such a transfer can be provided to define an order of magnitude. The VA required for a 6-hour transfer between two satellites separated by 953 km is approximately 15 m / s. Transfer between constellation planes

[0075] A direct transfer between two planes (15° difference in longitude of the ascending node) for a constellation of this type is obviously far too costly to be considered. Thus, a MIPELEC simulation gives a transfer time of 974 days for an AV of 3000 m / s.

[0076] Another solution presented could be to play on the differential altitude (or semi-major axis) to obtain a differential effect of the J2, in other words a relative drift of the ascending node could be generated by modifying the attitude of the inspector satellite.

[0077] This possibility is feasible but remains costly in terms of delta-V for an acceptable drift time (approximately 70 m / s for one month of drift to recover the 15° delta-RAAN between 2 planes, see [Fig. 7]). It is considered here as a backup, depending on the delta-V budget.

[0078] An innovative alternative consists of observing satellites from different planes by phasing with them so as to cross them in the vicinity of the “top” or “pole” of the point of junction of the orbits at 88° where the crossing distances can be minimized by choosing the right phasing between the hunter and the target.

[0079] For this analysis, we rely on the expression for this distance identified in the context of work on the Inter-Satellite Link. For two satellites placed in circular orbits with equal inclinations and semi-major axes, it is indeed possible to analytically calculate the distance separating them, with reference to the aforementioned DI publication. This therefore makes it possible to predict the inter-satellite distance during a pole crossing, and Therefore, the inspector's phasing relative to the target is adjusted to minimize this distance for potential image acquisition. An analytical expression for the azimuth and elevation of the target satellite in a local orbital frame is also identified, thus allowing the pointing to be predicted in order to keep the object within the field of view.

[0080] After verifying the advanced expressions in D1, a preliminary visibility analysis at the pole can be performed. An example for a target (called satellite B) on an adjacent plane (15°), with a true anomaly difference of 7.2° from the inspector satellite, is given in [Fig. 6]. The inspector-satellite B distance, as well as the azimuth and elevation coordinates in a local orbital frame on an orbit, are presented there (starting point at the equator). During pole crossings, the minimum inter-satellite distance is approximately 1 km for an azimuth of approximately 0° and an elevation of 3.8°. An overview of the azimuth and elevation variations is given in [Fig. 8], with a longitude difference of the ascending node AQ = 15° and a true anomaly difference Anu transformed into an inter-satellite distance difference ranging from 10 km to 150 km. This approach makes it possible to identify the expected scrolling speeds for satellites navigating on other planes.We observe maxima of angular velocity when approaching the poles, which can be on the order of degrees / second with regard to the azimuth.

[0081] The preliminary mission analysis for observing inter-plane satellites shows that the cross-observation strategy at the “pole” by the orbiter is possible (distance, apparent angular velocity) and it offers the advantage of a possible revisit of the satellites in different planes at a lower cost in delta-V (compared with a variation of the semi-major axis to have a differential effect of the J2).

[0082] End of life of an inspection satellite

[0083] An end-of-life strategy must be defined to comply with the Space Operations Act (LOS), which requires decommissioned satellites to re-enter the atmosphere within a maximum of 25 years. The passivation of onboard systems must also be considered to avoid any inherent risk of creating new debris. a. Active desorbation

[0084] A first approach consists of reducing the satellite's altitude to minimize atmospheric reentry time and reduce the risk of collision with other satellites orbiting in the same area. Considering a circular operational orbit at an altitude of 1200 km reduced to 150 km, the altitude required for this maneuver is as follows: a MIPELEC simulation gives a transfer time of approximately 193 days for a delta-V of 561 m / s, corresponding to 454 grams of propellant consumed, with continuous thrust. This therefore necessitates extending the operational duration by at least 7 months to complete this transfer.

[0085] A transfer to a higher altitude would reduce the delta-V and the associated duration, but would increase the atmospheric reentry delay. An analysis of the maximum altitudes allowing reentry in 25 years with average solar activity is summarized below and would require targeting 715 km, corresponding to a delta-V of 243 m / s.

[0086] A passive deorbiting analysis for an initial circular orbit at an altitude of 550 km was provided in DRl, which deviates from the new operational orbit considered in the response to this tender (1200 km). An analysis of the maximum altitudes allowing reentry in 25 years was provided, for different values ​​of ballistic coefficients (controlled attitude with maximum cross-section, uncontrolled attitude and "worst-case" with minimum ballistic coefficient):

[0087] Illustration of the Ballistic Coefficient at Maximum Altitude

[0088] Maximum section 0.0449 m2 / kg 715 km

[0089] Random rotation 0.0275 m2 / kg 674 km

[0090] Minimum section 0.0046 m2 / kg 540 km

[0091] The maximum altitude allowing reentry in 25 years is 715 km in the case where The attitude is controlled. It will therefore be impossible to ensure a passive deorbiting at an altitude of 1200 km respecting a reentry time of 25 years, given the ballistic coefficients involved.

[0092] An active deorbiting with transfer to a maximum altitude of 715 km is therefore a more feasible strategy (with post-transfer passivation of on-board systems). a. Passivation of embedded systems

[0093] According to a preliminary analysis of the passivation of the onboard systems, the NanoPower BP8 type batteries developed by GOMSpace must be discharged. Since the propulsion system contains no reactive liquid or any pressurized component (DR1), no specific action on this system should be necessary for the end of the satellite's life. Preliminary Propellant Assessment

[0094] With the satellite / propulsion characteristics of the Orbiter_V2, the total available velocity due to the indium mass included at baseline in the Micro-ENPULSION microsatellite propulsion system is 1630 m / s)

[0095] The following phases are to be considered for the propellant balance:

[0096] Launch and Early Operations (LEOP) and Positioning (MIP)

[0097] Rendezvous and IP

[0098] Job Maintenance (MAP)

[0099] Orbit transfer and phasing for IE

[0100] Transfer to end-of-life orbit

[0101] Collision avoidance

[0102] As mentioned previously, the LEOP and MIP phases require an AV of approximately 423 m / s in order to reach the target orbit in a "pessimistic" injection scenario.

[0103] The air velocity (AV) required for a close-range inspection is 0.65 m / s (result from the CNG simulator). As a non-limiting example, a minimum of 10 inspections can be set for the demonstration mission, which brings the AV requirement to 6.5 m / s. Before the margin policy is applied, this figure is rounded up to 10 m / s to account for the "under construction" nature of the CNG simulator.

[0104] The requirements for orbit transfers will need to be determined more precisely depending on the targets. As previously explained, the VA required for a 6-hour transfer between two satellites in the same plane separated by 953 km is approximately 15 m / s, which provides a first order of magnitude. Indeed, it will always be possible to lengthen the transfer time to reduce the VA cost. A transfer between two planes of the constellation is unlikely. The alternative of observing certain satellites during polar crossings requires phasing the inspector with the target to reduce the distance at the crossing. Here again, the VA cost can be reduced by lengthening the phasing time.

[0105] If we consider a capacity of 5 remote inspections, we obtain for phasing in the plane or for phase for at the crossing at the “pole” a budget of 85 m / s.

[0106] Regarding avoidance maneuvers, a GNC avoidance simulation gives an AV of 0.083 m / s. The number of avoidance maneuvers at 550 km is estimated at 3 per year, using the DRAMA (Debris Risk Assessment and Mitigation Analysis) software if the probability of triggering is set at 10⁶. This figure can be considered conservative for 1200 km, hence a budget of 0.25 m / s.

[0107] The end-of-life budget is 415 m / s to “descend” from a circular orbit at 1200 km to a circular orbit at 715 km. Flight Phase Delta-V Delta-V with Margin (20%) Positioning 423 m / s 510 m / s Position Hold 8 m / s 10 m / s IE and Subsequent Phasings 85 m / s 100 m / s IP (Rendezvous) 10 m / s 12 m / s Avoidance 0.25 m / s 0.3 m / s End of Life (Active then Passive) 415 m / s 500 m / s Total 941.25 m / s 1132.3 m / s

[0108] In total, the LEOP / MIP, MAP, inspection, end-of-life and phases Collision avoidance requires a margined AV of 1132 m / s at this stage of project development.

[0109] There therefore remains a significant margin in relation to the INDIUM capacity of the Micro-ENPULSION engine (1630 m / s) of the orbiter_V2. Ground network and station visibility

[0110] A simulation of station visibility was carried out at an altitude of 550 km using the KA-SAT satellite network. This simulation was repeated at an altitude of 1200 km and an inclination of 87.9° over a period of 5 years, which shows that, as one might intuitively anticipate, the situation is much better with the following data: - Daily crossings: 53.92 - Maximum duration of non-visibility: 5164 s - Average cumulative visibility time per day: 43488 s - Maximum visibility time: 976 s - Average visibility time per pass: 806.49 s

[0111] Passages are more frequent, visibility windows are wider, and periods of non-visibility are reduced. Extended visibility windows can thus be selected for critical operations, separated by safe Station Keepings awaiting the next window.

[0112] The inspection system according to the invention, intended for carrying out remote inspection (RI) and close inspection (CI) missions, comprises: - An Orbiter V2 satellite based on a 16U platform from GomSpace, extrapolated from their recurring 12U platform. The 16U platform is complemented by the following payload elements: Radar (for GNC during Rendezvous), Camera (to capture images during IE and IP inspections), and a space weather instrument. The platform's payload controller hosts the onboard GNC software, camera image processing, and FDIR / safety management for formation flying and Rendezvous. - a ground segment that contains both the classic control center of a LEO satellite: Telemetry / Remote Control (TM / TC), flight dynamics, monitoring, the specific part for formation flying coordination with the OneWeb constellation Control Center, the specific part for GNC Rendezvous (planning, monitoring, ground FDIR, takeover capability, safety) and the Specific section on data exploitation: images, space weather data - a network of ground stations to ensure telemetry / TM / TC remote control coverage.

[0113] In a practical implementation method, the platform, payload integration, satellite integration, and ground functional validation (Flatsat) can be carried out by GomSpace. The Mission Analysis and GNC components are performed by CS Group.

[0114] TNG develops and delivers the Radar, GomSpace and Simeria Sense deliver the camera and CS Group develops the ground segment. Product Platform - Satellite

[0115] A platform is planned with its flight software and an integration of three payloads necessary for the mission: camera, radar and SWE (Space Weather Electrons) instrument to measure the 3D distribution of high-energy electrons present in the solar wind.

[0116] This platform will be capable of storing at least 16 gigabytes of payload data. The payloads exchange data with the onboard computer via two SpaceWire links developed by the European Space Agency (one for the camera and one in anticipation of future development), two SPI (Serial Peripheral Interface) links: one for the radar and one in anticipation of future development, and a CAN (Controller Area Network) link for the SWE instrument.

[0117] The target orbit of the constellation is at an altitude of 1200 km, without necessarily having a Sun-synchronous orbit (SSO). At 1200 km, the satellite's lifespan is estimated at 2.5 years due to radiation. The payload has a mass of approximately 5 kg. The platform provides a maximum power of 60 W to the payloads.

[0118] The platform includes a redundant Star Tracker, each star tracker pointing in different directions. The satellite shall also accommodate a redundant gyroscope and a redundant magnetorquer. The Attitude Knowledge Error (AKE) shall be less than 0.1 degrees on each of the three axes at 3°, and the Absolute Pointing Error (APE) shall be less than 0.5 degrees on each of the three axes at 3°.

[0119] In a preferred embodiment, an electric propulsion system is provided for the operations of the payload. The propulsion system has a delta-V capacity of at least 1200 m / s and a continuous acceleration towards the satellite of at least 2.5 x 10A-5 m / s2 (i.e. a thrust of 0.6 mN for a 24 kg satellite).

[0120] Communication with the ground is ensured via telemetry and remote control equipment (TMTC) in S-band, S-band and X-band HSL. All communications between the satellite and the ground segment (remote controls, platform telemetry, payload telemetry) as well as within the ground segment are encrypted. Design of the Inspection Satellite

[0121] The satellite consists of a standard platform produced by the company GomSpace: 16U with two deployable tracking solar panels (TSP for Tracking Solar Panels) and three 8-cell battery packs (NanoPower BP8) offering a total capacity of 259 Wh.

[0122] The satellite is capable of delivering a maximum power of 60 W to the payload, but after a "payload" mode, it is necessary to allow the batteries to return to their nominal depth of discharge (DoD). According to data from extensive battery tests, the satellite is also capable of delivering an average power of 80 W to the payload and propulsion system during a 2-hour maneuver with two 35-minute eclipses (resulting in a DoD of 45%), assuming that such a depth of discharge does not occur more than once a week over a 5-year lifespan.

[0123] The platform comprises an advanced Attitude Determination and Orbit Control (ADCS) subsystem, including thin solar sensors, three magnetorquers (one for redundancy), two gyroscopes, four NanoTorque GSW-1200-4P reaction wheels (one for redundancy), two star trackers (one for redundancy), and a magnetometer. The proposed telemetry / remote control TMTC solution is S-band.

[0124] The GNSS (Global Navigation Satellite System) subsystem consists of a Tallysman TW1322 antenna and an onboard Tallysman TW3867 dual-band GNSS receiver.

[0125] The proposed High-Speed ​​Link (HSL) solution is in S-band for uplink and downlink, using the NanoCom ANT2150 antenna. In addition, a GomSpace NanoCom ANT8250 antenna in X-band is provided for downlink at very high data rates of 50 Mbps, with an upgrade up to 225 Mbps if required.

[0126] The design and technical characteristics of an orbiting satellite according to the invention are illustrated in Figures 2 to Attitude and Orbit Control System (AOCS)

[0127] Different operational and fallback modes of the SCAO attitude and orbit control system are provided: - Initial and fallback mode: Use of a minimal set of sensors and actuators - 3-axis pointing mode, with selectable reference frames: - ECI (Earth's Center of Inertia) - LVLH (Local Vertical Local Horizontal) - SUN (Pointing towards the sun) - Ground station tracking (Autonomous start of LLA coordinate tracking when user-defined thresholds are reached).

[0128] For this mission, additional SCAO modes are required to meet the requirements.

[0129] Payload attitude reference input: - The SCAO must be able to be controlled to track a target of interest. The direction of the target of interest is determined by the payload. - The proposed solution would be for the payload to define the desired mode and provide the SCAO with attitude and angular velocity references. - The payload must provide the reference attitude at an interval to be determined. - The SCAO remains responsible for safety functions and the control of withdrawal modes. All collision avoidance and guidance maneuvers are performed by the CNG Guidance, Navigation and Control system in the Payload Controller computer. - The SCAO has integrated monitoring systems to ensure that if the payload or the Guidance, Navigation and Control (GNC) system ceases to provide attitude references to SCAO, then the latter should switch to other nominal pointing modes. Additional security mode#:

[0130] Use of magnetorches only for pointing towards the sun and the ground station. This specific guidance law is integrated into a payload / CNG partition (in the payload controller). The payload can then interact directly with the propulsion system. Additional interfaces and backup managers are implemented to allow control of the propulsion system, for example, from the onboard computer. Mission software

[0131] With reference to [Fig. 6], which illustrates an example of a typical mission configuration, a development kit, such as that offered by GomSpace, includes key features and can be used directly as flight software for the inspection satellite's onboard computer: - The CubeSat Space Protocol (CSP) - A small network layer delivery protocol designed for CubeSats, used by most GomSpace products, with the exception of simple products such as simple sensors. - Open Standard: www.libcsp.org - The GomSpace shell: - A command-line interface for configuration and debugging - Customers / orders for standard GomSpace products - The GomSpace settings system - Versatile parameter system, used for configuration and telemetry - Supports file system (FLASH) and persistent memory (FRAM) storage - Modifying parameters without compilation - Remote access (settings can be changed directly from the ground) - Ease of use, same concept on most GomSpace products - GomSpace File Transfer Protocol (FTP) - Loading and downloading files (telemetry, software, payload data) - Resumption of interrupted downloads - “GomSpace Housekeeping” - Telemetry data collection from all components - Transmission of telemetry data to the ground (beacons) - Configuration performed via a JSON file (modifiable on the fly) - “GomSpace Flightplanner” - Executes GomSpace shell commands at a specific time - SCAO pointing configuration - Activation of the payload power channel - Payload data collection has begun

[0132] Satellite autonomy and fault detection, isolation and recovery (FDIR) concept

[0133] The purpose of the FDIR system is to rapidly detect the presence of faults during operation and to preserve the health and operational capability of the satellite. The The FDIR system helps isolate faults so that they do not propagate and affect other subsystems or components.

[0134] The fault protection functions of the inspection satellite according to the invention are implemented at different levels: - Watchdog functions are implemented at various subsystem levels, particularly for the Electrical Power System (EPS) - operational modes of the electrical power system. - Higher-level FDIR parameter monitoring.

[0135] In general, subsystems can autonomously manage minor faults and anomalies thanks to integrated monitoring functions. Environmental conditions

[0136] The 16U platform of the inspection satellite according to the invention is designed for typical LEO orbits up to a maximum altitude of 800 km. Generally speaking, increasing the altitude from 800 km to 1200 km has a significant impact on the components and materials and can affect the lifespan.

[0137] Optical camera onboard the inspection satellite

[0138] The optical camera must allow observation of the target during both distant (1 km - 30 km) and close-range (100 m) inspections with a resolution of 50 cm and 1 cm respectively. Since the target of interest has a dimension of approximately 1 m, the minimum field of view is therefore that which allows observation of the entire target during close-range inspection, i.e.:

[0139] ep] =180 / 77.1 / 100 = 0.75 0

[0140] Considering that the solar panels are deployed over a total span of 6 m, the total field of view required to image the target in the direction and the panels in the transverse direction is 3.4°. However, given that the close-up inspection will be conducted via a secure orbit around the target, the entire target will be imaged through the accumulation of photographs. Thus, a field of view between 0.75° and 3.4° will be suitable for the camera.

[0141] One possible choice is the xScapelOO camera (model TriScapelOO) developed by Simeria Sense. This camera is initially intended for ground inspection from low Earth orbits and is therefore designed for acquiring photographs at very high frame rates (microsecond exposures) and resolutions (approximately 5 m at 500 km). It has a field of view of 2.22° x 1.66°, making it suitable for close-range inspection, and its mass, volume, and power budget is compatible with GomSpace's 16U and 12U platforms. The CMOS sensor allows for low read noise and thus very short exposure acquisition.

[0142] In the present invention, this type of sensor offers the opportunity to adjust the exposure in order to optimize the signal-to-noise ratio by considering the relative speed with the target and its illumination.

[0143] For digital image processing, maximizing angular resolution afterward is possible by restoring spatial frequencies that are not lost due to defocusing and noise pollution (Gaussian noise from the detector and Poisson noise from the target). A first approach consists of deconvolving the images, that is, eliminating the instrumental response (PSF for Point Spread Function) which convolves the initial light distribution of the observed object. Assuming that the PSF is stationary over time during observation, it can be calibrated by observing a bright star before or after (or even during if a star is in the field of view). Deconvolution can thus allow a gain of almost a factor of 2 or more in the final resolution [Fétick et al. 2020, Yan et al. 2022, Petit et al. 2019, Hickman et al. 2020].Also, the use of prior information about the target, such as very high-resolution images calibrated in the laboratory before launch, can also be used in the process [Conan et al. 98, Lau et al. 2023].

[0144] In addition, super-resolution approaches can also be deployed to combine photographs acquired during inspection, with a potential gain of a factor of 5 on resolution, theoretically allowing 1cm resolution to be achieved for close inspection at 100 m with the xScapelOO camera. CNG system available by appointment

[0145] The Rendez-vous Guidance, Navigation and Control (GNC) system is based on the following points: - relative navigation with the Radar which provides Range, Range-rate and Los (2 angles) information - guidance to calculate the optimal trajectory to reach the target point using electric propulsion (low thrust) - control to ensure trajectory tracking via thrust commands and to transmit attitude commands to the platform's SCAO - Specific FDIR to monitor satellite equipment and ensure the monitoring of rendezvous trajectories versus collision risk

[0146] . Relative navigation is based on a Kalman filter that uses the Clohessy-Wiltshire equations for the transition matrix and to calculate the relationship between the radar measurement and the state. The state vector propagation is based on double absolute propagation with the orbiter's GNSS measurements in order to be more accurate in the absence of measurements.

[0147] The guidance system calculates in real time on board the optimal trajectory to reach the next point in the local orbital frame. The use of low electric thrust makes this onboard calculation necessary, unlike a pulsed Homing-Closing approach such as that used in ATVs (Automatic Transfer Vehicles). The problem of finding the optimal trajectory that minimizes power consumption is solved on board, using a transition matrix derived from the Clohessy-Wiltshire equations (fighter-target proximity of a few kilometers), over typical horizons of 100 seconds. Constraints such as maximum thrust, safety volumes, power limits, DoD limits for the batteries, and maximum duration are taken into account to validate the solution.The GNC simulator shows that convergence is good provided there is a minimum thrust, hence the choice of the Micro Enpulsion thruster of 0.8 mN for an Orbiter satellite mass of 24kg.

[0148] The control modes are as follows: - Sun-pointing mode: The satellite orients its solar panels with a solar pointing to maximize battery charging. - Radar acquisition mode (Tracking). The hunter orients the radar axis towards the estimated position of the target. In addition, since the panels can rotate around their axis of rotation, solar pointing optimization can be performed when the satellite is in radar acquisition mode, while keeping the target within the radar's FOV. - Photo mode: The camera is pointed towards the target and the conditions related to the stability of the pointing are taken into account to ensure the quality of the shot. - Maneuvering mode: The satellite aligns itself with the thrust direction controlled by the guidance system. The remaining degree of freedom allows for optimization of radar pointing or solar panel pointing.

[0149] The GNC takes into account the operational constraints related to the platform: eclipse management, thrust attitude and radar measurement, battery management (maximum DoD monitoring and triggering of Sun Pointing mode for recharging), latency time for performing attitude maneuvers, sun in the camera's field of view. Embedded software

[0150] The additional embedded software for the 16U platform is located in the “Payload Controller” computer (CNC, FDIR, camera processing). Particular attention is paid to the interfaces between the OBC (On-Board Controller) itself, which consists of two computers for the Attitude and Orbit Control System (AOCS), and the Payload Controller. useful). Commands to the platform equipment (including propulsion) go through the OBC on-board computer.

[0151] The air-to-ground interface is particularly critical for the Rendezvous, since the ground segment must monitor rendezvous operations and take over if necessary. It is therefore necessary both to send Rendezvous-related Telemetry data to the ground and to be able to receive Telemetry (TC) data from the ground. The platform's TM-TC system meets these requirements, complemented by the KA-SAT ground network. Ground Segment Architecture

[0152] The architecture of the ground segment, illustrated by [Fig.7], proposed for the inspection system according to the invention is an assembly of three GOSMIC components belonging to the CS GROUP ground segment product line: CSNano, SPnode and OREFLIDS.

[0153] The CS Nano component is composed of two subsystems:

[0154] The GSnode ground station manager is the single interface point between the control center and the ground stations. It performs three main functions: -> Manage real-time data during transit: receiving Telemetry and remote control transmission. It acts as a proxy for the rest of the Control Center. Currently available interfaces with ground station providers are CORTEX, the KSat Protocol, and the CCSDS SLE interface.

[0155] The TM / TC core is responsible for processing the telemetry data received from the satellite and for generating the commands to be sent to it. The main functions provided by this subsystem are as follows: -> Acquisition of telemetry frames, packet extraction and processing parameters. - >Encoding, generation and formatting of commands. - >Production of information on the model and aggregation of telemetry data usage reports. - > Execution of operational flight and ground procedures. Communication protocol

[0156] The communication protocol used between the satellite, the ground station(s) and the control center is the CubeSat Space Protocol (CSP). CSP is a layer 3 / 4 (network / transport) protocol, comparable to the IP protocol for the internet, with a concept of address and port for each network actor, as shown in [Fig. 8].

[0157] To interface with this network, a CSP gateway must be implemented in the GSnode ground station manager. This gateway, having its own address at within the CSP network, offers a TCP / IP interface allowing CSNano components to communicate with the GOMSPACE platform. CSP protocol integration

[0158] A CCnode module provides remote control sending and telemetry processing services. Sending a remote control involves sending one (or more) CSP packets (request) to a network device; the device's (possible) response to this request is processed as telemetry at the CCnode level.

[0159] Beyond simply sending individual remote controls, the platform generally offers higher-level services, such as file transfer, which require more complex TM / TC exchange sequences. For services of this type, specific applications can be developed in CCnode, providing a simplified interface for the operator in operational procedures. Automation#: SPnode

[0160] An SPnode module is a generic, cloud-native scheduler for managing sequences of activities. It offers a user-friendly web interface, designed to manage a very large number of activities, consisting of a time-series diagram view and a monitoring view. It can be used to automate a control center with several hundred satellites. It can be interfaced to automate certain tasks during operational implementation. Flight Dynamics System (FDS): OREFLIDS

[0161] Designed to meet the requirements of New Space, OREFLIDS is a generic product that allows for the easy and rapid development of a Flight Dynamics System (FDS) for any type of mission (LEO, GEO, nanosatellites, interplanetary, satellite constellations, etc.). Based on OREKIT and Hipparchus, which are proven open-source solutions, OREFLIDS natively provides the following basic space mechanics services: - Conversions of dates, reference points and orbital elements including TLEs, - Analytical and numerical orbit propagation, - -> Detection and generation of multiple orbital events, - -> Orbit determination based on GNSS measurements and / or measurements stations, - Importing, exporting and processing files in CCSDS format, - Generation of simulated measurements including GNSS measurements and station measurements such as range, range rate, radec and angular measurements. - Comparison of orbital elements between ephemerides. Use of the FDS for the mission

[0162] CS GROUP has significant experience on missions involving orbital rendezvous, particularly on the ATV (Automated Transfer Vehicle – cargo sent to the International Space Station). CS GROUP is also in charge of developing the FDS for the ADRAS-J mission, which aims to inspect a piece of debris. The proposed solution is therefore largely based on the experience gained from these two missions.

[0163] It is assumed here that the GNC simulator and the specific procedures related to relative orbital dynamics (Guidance, relative navigation by Extended Kalman filtering with Radar) will be included in the FDS framework.

[0164] The FDS may also benefit from certain HOOP modules when deemed more efficient and validated, particularly for the conventional orbital transfer phases.

[0165] The FDS will be employed during the following mission phases: - Launch and Early Orbit Phase (LEOP): survival exit, first orbit restoration. - Deployment to Station (MIP) and phasing: calculation of the transfer strategy to the target orbit, approach to the target satellite. - Rendezvous: transfer to a point closer to the target. - Remote and close-up inspection: approaching the target and taking photographs, Relative navigation via onboard radar measurements, real-time guidance (link with the CNG hi-fi simulator) - Maintenance at Station (MAP): calculation of the strategy for maintaining orbital parameters within the allocated windows during inter-inspection phases - Orbit transfer: calculation of the transfer strategy for inspecting different targets. - Collision avoidance: calculation of a collision avoidance strategy if necessary - Deorbiting: transfer to an orbit allowing atmospheric reentry in less than 25 years.

[0166] While the non-inspection phases are standard from an FDS perspective, the inspection phases require special attention. Indeed, since attitude and position control relative to the target are calculated on board, the information available to the FDS must be as up-to-date as possible regarding past and future strategy: - Past maneuvers / attitude: updated via retrieval of TM from the propulsion system and SCAO. - Future maneuvers / attitude: the predictions made on board must be brought down so that the FDS has the maneuver strategy (Guidance, timing and thrust direction, raw measurements, filter innovation) navigation,....) and the predicted attitude, so that propagation and calculation are as accurate as possible.

[0167] Regarding attitude data, the two points mentioned above can be extended to other phases of the mission. Attitude can indeed be predicted on board (autonomous attitude mode, for example). It will therefore be necessary to retrieve the attitude time constant (TM) as well as the predicted attitude, i.e., past and future data respectively.

[0168] Emphasis is placed here on two critical functions for proximity and rendezvous operations (in addition to the CNG simulator) which will be integrated into OREFLIDS: Collision risk assessment

[0169] The collision risk assessment component is responsible for retrieving conjunction alert messages, made available by USSSTRATCOM in the form of files in CDM (Conjunction Data Message, CCSDS standard) format, and processing them in order to provide the decision elements necessary for calculating a collision avoidance maneuver.

[0170] The component retrieves information from the CDM file, which can be in JSON, XML, or key / value format. The conjunction date, minimum distance, and collision probability are compared with the permissible thresholds entered in the FDS. An alert is generated if any of the thresholds are exceeded.

[0171] There are no plans to integrate a collision probability calculation component, as this is a complex function with several possible calculation methods. Furthermore, the results are difficult to validate, which inspires little confidence. For these reasons, it seems more reasonable to use the collision risk data reported in the CDM (conjunction date, minimum distance, and collision probability). The minimum distance can, however, be recalculated by the FDS using the latest TLE (Two-line Element) of the encountered object (or the orbit provided in the CDM) and the most up-to-date orbit of our satellite.

[0172] Collision avoidance maneuver / debris avoidance

[0173] A component of the onboard system is responsible for generating a collision avoidance maneuver (CAM) when necessary, in order to reduce the risk of collision with debris in the event of a dangerous approach.

[0174] The decision to initiate a CAM is made by a committee of experts, after analysis of one or more files in CDM (Conjunction Data Message) format. The main criteria are the distance to the TCA (Time of Closest Approach) and the probability of collision. Other criteria are also important, such as having several CDMs over several consecutive days, the reliability of the data (covariance), the size of the objects encountered (if known), etc. Such a decision cannot be part of an automated sequence.

[0175] The generation of a CAM is a special case of maneuver generation. The same principles will apply: an impulse delta-V is first calculated, then spread according to the properties of the propulsion model.

[0176] A two-maneuver avoidance strategy is developed: the first aims to deflect the satellite's trajectory, the second is necessary to return to the nominal trajectory.

[0177] Several strategies are possible. A tangential maneuver approach is incorporated into the FDS: a tangential maneuver induces an altitude differential and a drift in relative true anomaly, thus reducing the risk of collision. In terms of propellant consumption, such a maneuver is more effective when performed earlier. Typically, if the time before the TCA is doubled, the delta-V required is halved for a similar result.

[0178] Other strategies (radial or out-of-plane thrusts) are possible and may be more or less effective depending on the geometry of the conjunction. The tangential strategy is considered sufficient in the case of the FDS used in the inspection system according to the invention.

[0179] The minimum distance will also be calculated to verify that the avoidance maneuver has indeed increased the distance to the TCA.

[0180] Of course, the present invention is not limited to the embodiment just described and many other embodiments can be considered without departing from the scope of the present invention.

Claims

Claims

1. Method for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, characterized in that it implements an inspection satellite provided with inspection data collection equipment, the launch of which is phased to meet the satellites to be inspected within said constellation.

2. Inspection method according to the preceding claim, characterized in that it further comprises, from the inspection satellite, a collection of local space weather data and / or electromagnetic data.

3. Method according to one of the preceding claims, characterized in that it provides both from an inspection satellite a remote inspection of satellites to be inspected within the constellation, this remote inspection being carried out without rendezvous maneuver up to 1 km with a resolution of 10 to 50 cm, and a proximity inspection with respect to satellites to be inspected within said constellation, this proximity inspection including rendezvous maneuvers and proximity operations to approach at 100m with a centimetric resolution.

4. Method according to any one of the preceding claims, characterized in that it is arranged to inspect a constellation of satellites having an inclination of 88°.

5. Method according to the preceding claim, characterized in that it further comprises a step for choosing an altitude of the inspection satellite so as to phase a near encounter between the inspection satellite and a satellite of the constellation at the pole of intersection of orbital planes distributed in different longitudes of an ascending node.

6. Method according to the preceding claim, characterized in that it further comprises, once the inspection satellite is phased with a low altitude differential, a step in which the inspection satellite crosses the target satellite to be inspected within the constellation, in the vicinity of the intersection pole.

7. Method according to the preceding claim, characterized in that it further comprises, during the crossing, taking a plurality of images of the satellite to be inspected.

8. System for inspecting a plurality of satellites within a constellation of satellites in low Earth orbit, implementing the method according to any one of the preceding claims, characterized in that it comprises an inspection satellite provided with inspection data collection equipment and the launch of which is phased to meet the satellites to be inspected within said constellation, and a ground station provided with a system for communicating with the inspection satellite to receive the collected inspection data and means for processing said collected and received data.

9. Satellite intended to inspect a plurality of satellites within a constellation of satellites in low Earth orbit, implemented in the inspection method according to any one of claims 1 to 7, characterized in that it comprises equipment for measuring and collecting inspection data obtained when said inspection satellite crosses a satellite of the constellation, and a propulsion system.

10. Satellite according to the preceding claim, characterized in that the propulsion system comprises an electric propulsion stage.

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