Satellite, in particular a microsatellite

Abstract

A satellite, particularly a microsatellite, comprises a plurality of compartments selected from an avionics compartment (2a), a propulsion compartment, a payload compartment. Each compartment of the plurality of compartments (2) comprises at least one module formed by a plurality of bays (6) arranged laterally with respect to one another. Each bay (6) comprises a structure (8) having at least a base wall (8A), a back wall (8B) and possibly two side walls (8C) extending laterally from the back wall, particularly with opposite inclinations, a front panel (8D) being removably connectable to the structure (8). Each bay (6) further comprises at least one onboard unit (10), a plurality of mechanical connection elements, configured to allow mutual fastening of different bays (6) arranged adjacent to each other, and at least one electrical connection element for transmission of power and / or signal. The at least one electrical connection element is configured to establish power and / or signal connection with at least one different bay (6), so as to guarantee a power and / or signal communication between onboard units (10) belonging to different bays (6).

Description

[0001] Satellite, in particular a microsatellite'

[0002] DESCRIPTION

[0003] Field of the invention

[0004] The present invention relates generally to the construction of satellites. More in detail, the invention has been developed with reference to microsatellites with masses between ten and two hundred kilograms.

[0005] Prior art

[0006] In recent decades, the advent of the space economy has facilitated and intensified access to space, with a constantly growing number of annual launches. This trend is set to continue in the near future, driven in particular by increased launch accessibility, the growth of commercial and governmental applications, and the proliferation of new satellite constellations, which will further accelerate the increase in the number of launches.

[0007] On the other hand, small satellites, specifically microsatellites having masses between ten and two hundred kilograms, have become increasingly complex, aiming to achieve ever greater functionalities. Such small satellites are traditionally produced in limited numbers, often as single units, following a prototyping approach that clashes, for example, with the need to reduce time to launch and to build large constellations, which can comprise tens, hundreds, or even thousands of satellite units, for example for Earth observation or telecommunications.

[0008] The traditional prototyping approach essentially aims to optimize the mass, volumes, and performance of the single satellite, without however taking into account aspects such as scalability, industrialization, modification, or updating of the satellite's onboard components. These are fundamentally monolithic designs, whose internal space is not rationalized and whose mechanical and electrical interfaces are not standardized. Furthermore, the related production, assembly, and testing processes are complex, lengthy, and require the use of highly specialized personnel. In addition, all this also limits the possibility of outsourcing the construction of parts of the satellite, due to integration difficulties. In other words, these small satellites of fundamentally monolithic structure allow minimizing mass and volumes, but with the consequence of imposing extremely complicated architectures.

[0009] Consequently, the reusability of microsatellites across multiple missions is also severely limited. The monolithic architecture and the heterogeneity of the adopted technical solutions makeany design modification, such as the replacement of the payload or the updating of any type of onboard component, a complex and costly operation, involving high redesign costs. Furthermore, as already described earlier, the assembly methods are highly specific to each satellite model, reducing the adaptability of already developed solutions. This approach limits the possibility of exploiting economies of scale and introduces significant inefficiencies in all phases of the MAIT (Manufacturing, Assembly, Integration & Test) process, slowing down production times and increasing overall costs.

[0010] US 6206327 B1 discloses a modular satellite bus structure which is configurable depending on the mission, composed of different scalable modules, comprising a propulsion module and one or more stackable space modules. Each module includes a central hub of generally cylindrical shape from which a plurality of radial panels extend outward, to form a plurality of bays intended to contain subsystem equipment. The hub includes a central cavity sized to receive a propulsion module as a fully assembled unit. Between adjacent radial panels, frames are mounted to support removable external panels, which frontally close the bays. The external panels have supports for mounting equipment on their inner surface, as well as radiators on their outer surface, to provide thermal cooling to the equipment.

[0011] Object and summary of the invention

[0012] The present invention essentially aims to solve the aforementioned drawbacks.

[0013] In this context, an aim of the present invention is to provide a satellite, particularly a microsatellite, having a highly flexible and easily assemblable structure.

[0014] An auxiliary aim of the invention is to provide a small-sized satellite, in particular a microsatellite, whose structure allows a more rational use of the available space, and the possible accommodation of bulky components, compared to known solutions for this class of satellites.

[0015] At least one of the aforementioned aims is achieved, according to the invention, by a satellite, in particular a microsatellite, having the features forming the subject of the claims that follow, which form an integral part of the technical teaching provided herein in relation to the invention.

[0016] Brief description of the drawings

[0017] Further objects, features, and advantages of the invention will become clear from the detailed description that follows, made with reference to the accompanying drawings, provided by way of non-limiting example only, in which:

[0018] - figure 1 is a perspective view of a first satellite according to a possible embodiment of the invention,- figure 2 is a perspective view of a part of a satellite according to a possible embodiment of the invention,

[0019] - figures 3 and 4 are perspective views, from different angles, of the supporting structure of a bay used in the construction of a satellite according to possible embodiments of the invention, - figure 5 is a partially exploded perspective view of a module of a satellite according to possible embodiments of the invention, formed by a plurality of bays according to figures 3 and 4, - figure 6 is a perspective view of the module of figure 5, with the respective bays assembled together,

[0020] - figure 7 is a perspective view of parts of the satellite of figure 1 assembled together, - figure 8 is a front elevation view of a bay of a satellite according to possible embodiments of the invention,

[0021] - figures 9 and 10 are perspective views of a bay of a first type, respectively with some side cover panels present and removed, used in the construction of a satellite according to possible embodiments of the invention,

[0022] - figures 11 and 12 are perspective views of a further bay of a first type, respectively with some side cover panels present and removed, used in the construction of a satellite according to possible embodiments of the invention,

[0023] - figures 13 and 14 are respectively a rear perspective view and a bottom view of a bay of the type shown in figures 9-10 or 11 -12,

[0024] - figures 15 and 16 are perspective views, respectively front and rear, of part of a structure of a bay of a second type, used in the construction of a satellite according to possible embodiments of the invention in combination with bays of a first type as shown in figures 1 -10 or 11 -12,

[0025] - figures 17 and 18 are plan views, respectively from above and below, of the part of the bay structure of figures 15-16; and

[0026] - figures 19 and 20 are perspective views from different angles of a module of a satellite according to possible embodiments of the invention, formed by a plurality of first-type and second-type bays according to figures 9-14 and 15-18.

[0027] Detailed description of the invention

[0028] The reference to an embodiment in this description is intended to indicate that a particular configuration, structure, or feature described in relation to the embodiment is comprised in at least one embodiment. Therefore, phrases such as "in one embodiment", "in various embodiments" and the like, possibly present in different places in this description, are not necessarily referring to thesame embodiment. Furthermore, particular configurations, structures, or features described or illustrated may be combined in any suitable way in one or more embodiments, even different from those depicted. Some configurations, structures, or features described or illustrated in reference to "one embodiment" or "in various embodiments" may be at least partially present or absent in other embodiments.

[0029] The references used herein are for convenience only and therefore do not define the scope of protection or the extent of the embodiments. The spatial references (such as "upper", "lower", "top", "bottom", etc.) used herein are for convenience only and refer to the examples as shown in the figures.

[0030] Reference numeral 1 in figure 1 - and in general in the figures - designates as a whole a satellite according to possible embodiments, in particular a satellite with dimensions between 10 and 200 kg, hereinafter also identified as a microsatellite. As will become clearer later, according to an important aspect of the invention, the satellite 1 has a substantially modular basic structure, preferably distinguished by a partitioning of the satellite itself into compartments or decks.

[0031] The satellite 1 can comprise one or more compartments, in particular one or more main compartments. In various embodiments, the satellite according to the invention includes a plurality of main compartments selected at least from an avionics compartment, a propulsion compartment and a compartment intended for the payload.

[0032] In particular, in the non-limiting embodiment shown in figure 1, the satellite 1 comprises the aforementioned three main compartments, namely an avionics compartment designated by 2a, a propulsion compartment designated by 2b and a payload compartment designated by 2c. Note that the propulsion compartment 2b is an optional, and therefore non-essential, compartment of the satellite 1. As will be seen, according to possible variant embodiments, the payload compartment 2c can also be considered optional, and therefore not indispensable for the satellite 1.

[0033] In various embodiments the avionics compartment 2a houses the onboard components necessary to ensure the correct functioning of the satellite and its payload. Inside it, there may be present one or more onboard computers, a telecommunications-related system, an electrical power management system and a satellite attitude control system, implemented for example by means of Reaction Wheels and / or Control Moment Gyroscopes (CMG).

[0034] In various embodiments the propulsion compartment 2b comprises the propulsion system, which can be of any known type according to the needs of the mission for which the satellite 1 is intended. The compartment 2b preferably also includes the supply tanks for the propulsion system, which are scalable according to the specific requirements of the mission and according to the sizeof the satellite 1 itself. In the embodiment shown in figure 1, the propulsion compartment 2b faces the avionics compartment 2a and is interconnected therewith. On a side opposite the avionics compartment 2a, the propulsion compartment 2b preferably includes the interface configured to provide the mechanical and electrical coupling with a launcher. As already indicated, the presence of the compartment 2b is not essential.

[0035] In various embodiments, the compartment 2c comprises in its entirety a volume prearranged for housing the payload. This is preferably a volume variously configurable according to the characteristics of the payload itself, guaranteeing maximum flexibility during integration. In the example shown, reference 2c' indicates an external cover of the compartment 2c, which represents the maximum volume potentially occupiable by the payload. As already indicated, the presence of the compartment 2c may not be necessary.

[0036] According to an aspect of the invention, the satellite 1 can comprise one or more modules for each compartment, wherein at least one module comprises a plurality of bays arranged so as to form a radial configuration. Said bays can be structurally equivalent to each other, i.e., have respective supporting structures that are substantially similar, or combinations of bays having different shapes or supporting structures can be provided.

[0037] Figure 2 exemplifies the case of a satellite architecture, shown only partially, which includes a propulsion compartment 2b, at the lower part, and an avionics compartment 2a, the latter being formed by two modules stacked on each other, each designated by 4. In this example, the central module 4 and the upper module 4 are formed by the integration of four bays and two bays, each designated as a whole with 6.

[0038] As will be seen, in various preferred embodiments, the bays 6 have an overall substantially trapezoidal cross-section and are arranged in a peripheral sense around a central volume 16, as highlighted for example for the central module 4 of figure 2. This is not, however, an essential characteristic, when a module - such as the upper module 4 of figure 2 - is formed by only two bays 6.

[0039] In preferred embodiments, at least the avionics compartment 2a, or at least one module 4 of the avionics compartment 2a, is obtained by assembling together several bays 6.

[0040] Referring in particular to figures 3 and 4, each bay 6 comprises a supporting structure 8, which is preferably formed in a number of parts assembled together. In various preferential embodiments, the supporting structure 8 comprises a lower or base wall 8A, a rear or back wall 8B, and two side walls 8C. Note that the base wall 8A and / or each side wall 8C (and / or an upper 8E wall described later) could be omitted, in variant embodiments in which a compartment 2 is provided witha common base plate on which all its related bays 6 are fixed, or in variants in which, for example, two opposed first bays of a first type, provided with side walls, are associated in an orthogonal direction with two opposed second bays of a second type, not provided with side walls. The illustrated configurations are in fact exemplary, since one or more bays may be free of one or more side walls and / or share walls with an adjacent bay.

[0041] The side walls 8C extend from two opposite longitudinal sides of the back wall 8B, particularly with opposite inclinations, and are preferably fixed at their lower edge to the base wall 8A (and / or the upper wall 8E). The back wall 8B is preferably fixed between the side walls 8C (and possibly also to the base wall 8A and / or the upper wall 8E). The fixing can occur according to any known method, preferably via threaded fastening elements.

[0042] The base wall 8A preferably has a peripheral profile substantially in the shape of a trapezoid, particularly an isosceles trapezoid, so that even the single bay 6 closed frontally by a panel 8D has an overall cross-section (according to a plane parallel to the base wall 8A) that is substantially trapezoidal. The supporting structure 8, i.e., the walls 8A-8C that compose it, is preferably formed of metal or metal alloy, for example aluminum. Preferably, the same back wall 8B has a cross-section substantially in the shape of an isosceles trapezoid, i.e., its edges closest to the side walls 8C are inclined in opposite directions. The trapezoidal cross-section of the bays 6 constitutes an enabling geometry for modular peripheral assembly, as it allows obtaining, by varying the number of assembled bays, modules having different external polygonal cross-sections and / or a central volume 16 with a correspondingly polygonal cross-section.

[0043] To the structure 8 of each bay 6, at least a part of a front panel can be removably connected, some of these panels being designated 8D for example in figure 2. The fixing of each panel 8D to a structure 8 can also occur according to any known method, preferably via threaded fastening elements. For example, the outer sides of the side walls 8C and the base wall 8A (i.e., their frontal sides relative to the back wall 8B) can be provided with holes for respective fastening screws of a panel 8D. The panel 8D is also preferably formed of metal or metal alloy, for example aluminum.

[0044] Further external paneling elements can also be provided at the top of bays 6 intended to be exposed directly to the outside of the satellite. Some of these panels or upper walls, designated 8E, are visible in figure 2. The panels or walls 8E can be configured similarly to the base walls 8A, and be connected to the back wall 8B and / or to the side walls 8C, i.e., also have structural functions.

[0045] As shown for example in figures 4-7, which illustrate an example of an avionics compartment 2a, each bay 6 is equipped with at least one onboard unit 10. The unit 10 can be for example selected from sensing elements and / or onboard computers and / or telecommunications devices and / orelectrical power storage and distribution devices or other. In other words, therefore, each bay 6 defines a subset of the module 4 provided with its own structure 8, for example a chassis made of aluminum, and containing part of the onboard avionics, such as onboard computers, radios, attitude and orbit control systems, power conversion systems or other. As particularly noted in figure 5, in various embodiments, at least one onboard unit 10 comprises one or more components mounted on the inner side of the back wall 8B of the respective bay 6.

[0046] Preferably, each subsystem of the satellite 1 is housed in a distinct bay 6. This approach allows each bay to be developed, tested and possibly updated independently from the rest of the satellite, offering significant advantages in terms of scalability and modularity.

[0047] As already mentioned, in a particular embodiment, the structure 8 including the front paneling of the bays 6 has a cross-sectional shape that is substantially trapezoidal, wherein by crosssection is meant a section made parallel to the plane identified by the base wall 8A. This configuration facilitates assembly and improves accessibility to the bay by specialized technicians. Furthermore, these advantages open the possibility, in a future perspective, of performing maintenance, updates, or even replacements of bays 6 directly in orbit, during the operational life of the satellite, possibly after removal of the front panel 8D of the respective bay 6.

[0048] In addition, the radial composition of the satellite 1 and, in particular, the associated arrangement of the bays 6 allows providing modules 4 in which the external paneling, constituted by the union of the front panels 8D of the bays 6, can be partially open to the outside: this allows the possible integration of onboard components 10 or payload that require specific pointing and exposure conditions to space. Preferably, the external paneling of the module 4 is obtained in a segmented manner, via front panels (8D) associated with the individual bays (6), each removable and / or openable independently from the others, so as to allow selective access to the contained subsystems.

[0049] In a preferred embodiment, each bay 6 is provided with a plurality of mechanical connection elements and at least one electrical connection element for the transmission of power and / or signal. Preferably, to facilitate integration and allow the possibility of commissioning third parties to manufacture different components, said connection elements are standardized, guaranteeing maximum compatibility with the rest of the satellite. In various embodiments, the position and the hole pattern of the mechanical connection elements (12a, 12b, 12c) are maintained conforming among bays 6, so as to allow the replacement of one bay 6 with another without structural modifications to the module 4 or the compartment 2. This makes it possible, as needed, to rapidly reconfigure the module 4 according to the mission, keeping the external dimensions and the mainmechanical interfaces unchanged.

[0050] In various embodiments, mechanical connection elements, some of which are indicated with 12a in figures 3-4, are located at the two side walls 8C of each bay 6 and are configured to guarantee at least the coupling between the respective side walls 8C of different bays arranged adjacent to each other, via suitable fastening elements. The mechanical connection elements 12 can comprise for example through holes, with associated threaded fastening elements.

[0051] In addition or alternatively, mechanical connection elements, some of which are indicated with 12b in figures 3-4, are substantially located at the two sides of the back wall 8B corresponding to the side walls 8C, and are configured to guarantee at least the coupling between lateral edges of respective back walls 8B of different bays arranged adjacent to each other, via suitable fastening elements. The mechanical connection elements 12b can also comprise for example through holes, with associated threaded fastening elements.

[0052] Further mechanical connection elements, some of which are indicated with 12c in figures 3-4, can be provided at the lower and upper sides of the back wall 8B, and be used to enable coupling between different modules 4 or compartments 2, via suitable fastening elements. The mechanical connection elements 12c can also comprise through holes with associated threaded fastening elements.

[0053] Specifically, in various preferential embodiments, such as those illustrated in the figures, the mechanical connection elements 12c are configured to allow the coupling and / or mutual fastening between modules 4 or compartments 2 arranged superimposed on each other, as for example shown in figure 2 (superposition of two modules 4) or in figure 7 (superposition of two compartments 2a and 2b). In a preferred embodiment, the superimposed arrangement of modules and / or compartments is adopted to optimize the axial development of the satellite, making it more compatible with the launcher and improving the efficiency in the use of internal volumes.

[0054] The above cited at least one electrical connection element that equips each bay 6, for the transmission of power and / or signal, is configured to guarantee communication between different onboard units 10 present in the satellite 1. Said electrical connection elements can be aimed at allowing both connections within the same bay 6, connections within the same compartment 2, and connections between different compartments.

[0055] Said electrical connection elements can consist of multipolar electrical connectors, to which corresponding complementary connectors associated with the wiring that equips the various onboard units 10 are intended to be coupled. In various embodiments, the distribution of power and / or signals can be implemented as a bus distributed among bays 6, via standardized connectorspositioned in dedicated locations (for example brackets 14a, 14b and / or seats 14c, 14d) and via passages (8B', 8C) configured to route any wiring between adjacent bays and / or between stacked modules. This architecture can be useful as needed to allow the replacement of a bay 6 while keeping the main wiring unchanged, via the uncoupling and recoupling of said connectors.

[0056] For clarity, in figures 1-7 neither the said wiring with the corresponding connectors, nor the electrical connection elements, which as said can be of a type known per se , are shown. However, the positioning zones of such electrical connection elements are shown in the figures, said zones being able to be obtained for example by suitable brackets and / or by passages associated with the structure 8 of the respective bay 6.

[0057] Referring for example to figure 4, 14a and 14b indicate two brackets for supporting corresponding standardized connectors for transporting signals (data) and power, respectively, where in the example said brackets 14a, 14b are cantilever-fixed to the inner side of the back wall 8B. Preferably, the back wall 8B is also provided with passages, indicated with 8B' for example in figures 4 and 8, preferably in a position close to the brackets 14a, 14b, which allow the passage of wiring coming from, or directed to, a different bay 6 or module 4. Again in figure 4, 14c and 14d indicate through seats defined at the base wall 8A of the bay 6, for the positioning of respective electrical connectors used for the mutual connection between two stacked compartments 2, or for the connection between more superimposed modules 4 that obtain a same compartment 2 (as exemplified in figure 2).

[0058] In addition or alternatively, the side walls 8C of each bay 6 can be provided with lateral passages having functions similar to those indicated for the passages 8B': some of such lateral passages are indicated with 8C for example in figures 3-4.

[0059] Figure 5 exemplifies an assembly phase of an avionics compartment 2a composed of four similar peripheral bays, while figure 6 exemplifies the same compartment assembled. In these figures, as well as in the subsequent figure 7, the paneling of the compartment 2a is removed to show the exemplified avionics component housed in each bay, in particular at the inner side of the respective back wall 8B. Figure 7 exemplifies the architecture of a satellite composed of an avionics compartment 2a, obtained according to figures 5-6, and a propulsion compartment 2b.

[0060] As appreciable particularly from figures 5 and 6, the composition of four bays 6 within a single module 4 is such as to obtain a central volume 16, which is defined by the back walls 8B of each bay 6, once these are coupled and interconnected via the mechanical connection elements 12a and / or 12b of figures 3 and 4. In other words, the central volume 16 is generated directly by the peripheral coupling of substantially planar back walls 8B of distinct bays 6, resulting in a polygonalsection (for example quadrangular) devoid of continuous cylindrical surfaces.

[0061] Specifically, the central volume 16, which can also extend along more stacked modules 4, allows stowing and employing particularly bulky components. Compared to conventional solutions, which usually foresee a cylindrical volume, this configuration therefore allows a more rational and optimized use of the available space, while maintaining a fixed and standardized distribution of the avionics component. This guarantees greater flexibility in the integration of any additional components without compromising the structural organization. The fact that the back walls 8B are substantially planar, i.e., that the central volume 16 has a substantially polygonal, in particular quadrangular, section, also avoids the need to foresee specific mechanical adapters, which in known solutions must be used to install on a curved-surface body avionics components typically having a flat-faced chassis.

[0062] Each bay 6, in addition to serving as a housing volume for the onboard units 10, performs a structural function. In particular, the satellite, not being provided with a single, shared skeleton for all its elements, uses the bays 6 as integrated structural elements. Unlike architectures based on a central body (hub) and on radial ribs or panels, a satellite module according to the invention does not require a unique central structural element: the function of primary structure is in fact performed directly by the set of bays assembled together. In particular, the main load paths are transferred between adjacent bays via their respective mechanical connections, reducing the need for central frames or skeletal elements.

[0063] For this purpose, preferably, the structure 8 of each bay also presents reinforcement elements. In various embodiments, the reinforcement elements comprise ribs, some of which are indicated with 18 for example in figure 3, present on one or more of the walls of the structure 8 (in the example, on the outer faces of walls 8B and 8C), configured to confer to the satellite a predetermined mechanical resistance and stiffness. The reinforcement elements 18 can assume different forms based on the type of load to which the satellite will be subjected, and consequently can define predefined load paths. This configuration is advantageous for guaranteeing the satellite 1 the capacity to withstand the loads that develop in any phase of its operational life, particularly the launch phase with the space launcher.

[0064] It should also be emphasized that, advantageously, the proposed composable structure contributes to attenuating electromagnetic interactions between systems located in distinct bays, which in particular reduces the risk of interference between avionics components and facilitates the characterization of the internal electromagnetic environment of the satellite. This attenuation is also favored by the physical segmentation into bays provided with metallic walls and by the possibility ofrouting wiring through dedicated passages (8B1, 8C) at appropriate points.

[0065] From what has been previously described, it is clear how at least one compartment of the satellite according to the invention, and in particular at least the avionics compartment 2a, has a structure that is composable in a peripheral direction, based in particular on a subdivision into trapezoidal bays, and that is also possibly composable in an axial direction, via superposition of multiple modules 4.

[0066] Each compartment 2 of the satellite, as well as the internal bays 6 that compose at least the avionics compartment 2a, is independent, i.e. , it can be produced and tested autonomously, to then be integrated into the satellite assembly. This approach allows offering modularity, as the set of the various assembled bays 6 constitutes the satellite, while simultaneously allowing the updating or replacement of individual bays according to mission needs. Furthermore, as stated, the proposed architecture allows adding additional modules in the axial direction as well, to scale the satellite's performance.

[0067] As described previously, according to the invention, the bays 6 represent a sub-assembly of at least one main assembly represented by a compartment of the satellite 1 , particularly at least its avionics compartment, which is provided with its own structural component (represented by the structure 8) and contains part of the components, i.e., the onboard avionics of the satellite (comprising for example computers, radios, power conversion systems, etc.).

[0068] The structural component of each bay 6, i.e., its structure 8, is substantially standardized, adopting reference interfaces, both mechanical and electrical.

[0069] The substantially trapezoidal geometry of the individual structures 8 of the bays 6 implies various advantages, and in particular:

[0070] - allows an optimization of the internal volume of the compartment of interest, particularly the avionics compartment: once integrated, the bays 6 in fact create the central volume 16, which can be used for housing bulky components (for example, the payload); as stated, this central volume 16 is delimited by the minor bases (i.e., by the walls 8B) of the trapezoidal structures, and this guarantees a more rational use of space compared to typical configurations of small satellites, in which a central space is traditionally delimited by a cylindrical wall;

[0071] - allows, as needed, the creation of compartments partially open to the outside, in cases where mission requirements include particular pointing and exposure conditions to space for payload;

[0072] - the radial layout of the bays 6, arranged around the central volume 16, makes each module 4 or bay easily accessible, following removal of the respective front panels 8D;- the union of the trapezoidal bays 6 leads to the creation of preferential paths for the launch loads to which the structure 8 is subjected, increasing the stiffness of the entire integrated system.

[0073] Preferably each bay 6 is reserved for a particular function of the satellite 1 and, as stated, is provided with standardized connectors in terms of electrical and data interfaces. These connectors can be installed on one or more walls of the structure 8 of the bays 6, and allow the electrical / electronic / data connection of multiple bays 6 with each other, facilitating their integration, replacement, and updating.

[0074] Inside each bay 6 are housed, in addition to a respective portion of the satellite's onboard components, the wiring necessary to connect the various components with the peripheral connectors of the bay itself, as exemplified for conductors C1, C2 and C3 in figure 8, which exemplifies the internal wiring of a generic bay (in said figure, with C3' are exemplified the conductors of wiring coming from another bay 6).

[0075] Each bay 6 is therefore designed with independent fastenings and connections, and this approach facilitates the integration, replacement, and updating of the individual modules, therefore of the individual bays, without compromising the entire assembly and without impacts on the other bays. The partitioning of the avionics into bays, moreover, allows testing its components independently during the development phases. This circumstance, combined with the ease of access determined by the radial layout, opens the possibility, in a future perspective, of performing modular replacements in orbit during the operational life of the satellite 1, to allow maintenance and updating services.

[0076] The bays 6 can be of various types, specialized for functionality, including for example all those subsystems that participate in providing a specific function. By way of non-limiting example, in preferred embodiments, at least the following three main types of bays can be identified:

[0077] a) an avionics / onboard computer and telecommunications bay, which includes the onboard computer and management part of the satellite 1 , as well as the part dedicated to data management and sending of this data to ground (telecommunications subsystem);

[0078] b) an attitude determination and control system (ADCS) bay, which includes the instruments and systems for providing the attitude determination and control function, and therefore the onboard sensors and the related electronics that allows determining and controlling the three axes of the satellite 1 ;

[0079] c) a power management system (EPS) bay, which includes the electronics necessary for managing power supply and conversion on board the satellite 1; this bay can also be directly interfaced to the solar panels of the satellite 1 (said panels being indicated with 30 in figure 1).The electrical connection between the different bays 6 and between the different compartments 2 is based on standardized connectors and protocols. It is preferable to separate the connectors dedicated to the payload compartment 2c from the connectors dedicated to the avionics compartment 2a, so as to avoid design adaptations in the platform wiring, reducing non-recurring costs and guaranteeing repeatability among the various satellites of the same family. This constitutes a significant advantage, as it establishes a standard, simplifying the integration phase, and opening up the development of bays also by third parties. This solution also allows flexibility in the use of already available and qualified avionics and technologies, not necessarily requiring the development of components dedicated to this configuration.

[0080] The compartment 2c is a space of the satellite 1 reserved for the payload. This space can be completely customizable according to the characteristics of the payload itself, allowing it to host multiple different types of components, such as for example cameras, hyperspectral detectors, radars or antennas. These types of payload naturally fall within a housing with characteristics defined according to the satellite class. At the interface level, the payload compartment 2c, when provided, is nonetheless standardized in terms of connectors and electrical interfaces with the rest of the satellite 1 and this allows hosting different types of components without impacts on the avionics bus architecture of the satellite.

[0081] The propulsion compartment 2b, when provided, is reserved for housing the propulsion system of the satellite 1 in a respective housing. The related standard interfaces are nonetheless maintained towards the adjacent avionics compartment. As already indicated, within the propulsion compartment 2b it is possible to install different classes of propulsion systems according to mission needs, scaling the capacity of the tanks but always maintaining, however, the same structural and avionics interfaces. Preferably the propulsion compartment also includes externally the part of the interface with the launcher.

[0082] As previously indicated, in possible embodiments, a satellite module or a compartment thereof, for example the one designated 2a, can comprise bays having different shapes or supporting structures. In the following, the case of a compartment formed by the composition of bays of a first type and a second type is described, distinguished by different base structures which however present similar components.

[0083] In various embodiments, the features described below with reference to figures 9-20 can be adopted individually or in any combination with each other in a same bay, in a same composition of bays, or in one or more modules and / or compartments of the satellite, it being understood that such features are compatible with the embodiments already described with reference to figures 1-8. Inparticular, the solutions described below relating to an intermediate upright (8F), to a segmented frontal closure of the bays (8D1, 8D"), to thermal bridge elements (23) and to the presence of bays of a second type (6'), can cooperate with each other to optimize, according to the mission, structural robustness, accessibility, integration of external components and thermal management.

[0084] Figures 9-14 refer to first-type bays 6, substantially similar to those already described previously, whose supporting structure 8 includes a base wall 8A, a back wall 8B and two side walls. In the exemplified case, the structures 8 further comprise upper walls 8E, of the type previously indicated, having structural functions, which present a peripheral profile substantially similar to that of the base walls 8A.

[0085] In various embodiments, each bay comprises an intermediate upright arranged between the respective base wall and upper wall, in an intermediate region relative to the lateral ends of the bay, and in a position generally opposite the back wall, where preferably the intermediate upright is fixed to said walls by means of threaded fastening members that engage respective through holes, and where preferably at least one of the base wall and the upper wall of the considered bay presents a reinforcement, for example a thickening, at said through holes.

[0086] In various embodiments, said intermediate upright cooperates with the base wall and with the upper wall to define an internal structural reference of the bay, suitable for stabilizing and repeating over time the positioning of elements mounted frontally and / or externally to the bay. In this way, the intermediate upright contributes not only to the stiffness of the supporting structure, but also to the alignment precision of components and interfaces associable with the frontal closure of the bay and / or to components intended to be faced outward.

[0087] Referring in particular to the example of figure 9, the supporting structure 8 includes a further structural element, represented by an intermediate upright 8F, preferably formed of metal or metal alloy, for example aluminum. The intermediate upright 8F extends between the base wall 8A and the upper wall 8E, in a region intermediate to the respective lateral ends, in a position generally opposite the back wall 8B; in the example, the two ends of the intermediate upright 8F are fixed between facing or internal surfaces of the walls 8A and 8E, in particular near their front edge. Uprights of this type are also usable in the bays 6 described with reference to figures 2-8.

[0088] The fixing of the intermediate upright 8F is preferably obtained at the two ends thereof via threaded members, indicated with 15 in figures 9 and 13, at respective through fastening holes defined in the walls 8A and 8E, such as the holes indicated with 17 in figure 16. For this purpose, at least one of the base wall 8A and the upper wall 8E can be provided with a structural reinforcement at such holes 15: in the case of figure 16, such reinforcement - indicated with 17a - is constitutedby a thickening of the base wall 8A.

[0089] In various embodiments the intermediate upright 8F, in addition to performing structural functions, is also configured as a support element for components of the compartment or module satellite, or of a corresponding bay 6, in particular avionics and / or communication components intended to be faced outward of the module itself, such as at least one from among an antenna, a sun sensor, a star sensor. For example, referring to the case of the bay 6 represented in figures 9-10, an antenna, indicated with 20, and a sun sensor, indicated with 21, are fixed to the outer side of the intermediate upright 8F.

[0090] In the example of figure 9, moreover, active components of the satellite module or of the related bay 6 can also be associated with the inner side of the upper wall 8E, such as one or more star-trackers 22, i.e., cameras used - according to known methods - to ascertain the satellite's orientation in space. In the example, two star-trackers 22 are fixed, with opposite orientation, to the upper wall 8E, so that their respective imaging optics face outward of the bay 6: for this purpose, the front panels 8D' of the bay 6, which extend between a respective side wall 8C and the intermediate upright 8F, have a height less than the distance between the walls 8A and 8E, precisely to allow at least partial exposure of the sensors 22.

[0091] Figure 10 shows the same bay 6 of figure 9, but with said front panels 8D' removed. Also from this figure it is noted how, preferably, at least one onboard unit 10 of the bay 6 comprises one or more components mounted on the inner side of the back wall 8B, and how at the base wall 8A there can be provided electrical connectors 14c', 14d' usable for example for the mutual connection between two stacked compartments of the satellite, or for the connection between more stacked modules that form a same compartment of the satellite (as already exemplified in figure 2).

[0092] In various embodiments, radiating surfaces of the satellite module are obtained by the corresponding front panels of the bays 6, to which heat is transferred by means of one or more elements suitable for heat transfer coupled to the back wall 8B and / or to other internal components, allowing effective cooling even with segmented and removable external paneling.

[0093] In particular, in various preferential embodiments, the use of removable and segmented front panels is made compatible with efficient thermal management by means of one or more thermal bridges, which define a repeatable thermal coupling between internal heat sources and radiating surfaces. This configuration allows maintaining the accessibility and replaceability of the bays without penalizing the heat evacuation capacity, since the thermal transfer towards the front panels can be restored by simple recoupling of the panel and / or the thermal bridge.

[0094] In various embodiments, one or more bays comprise at least one thermal bridge elementfixed to the back wall, which extends, preferably cantilevered, towards the front of the bay itself and is configured to come into contact, preferably by a front portion thereof in the form of a plate, with the inner side of a front panel of the same bay, so as to define a preferential path for thermal conduction from the internal heat sources of the module towards radiating surfaces obtained from said front panels.

[0095] In the example of figure 10, two thermal bridge elements 23 are fixed (for example screwed) to the back wall 8B of the bay 6, particularly on the two sides relative to the intermediate upright 8F, which are preferably formed with a thermally conductive material, such as a metal or a metal alloy. Each element 23 extends in the direction of the front of the bay 6, so that its front part, preferably in the form of a plate, is in contact with the inner side of a respective front panel 8D'. Preferably the front part of the element 23 and the panel 8D' are secured to each other, for example by means of threaded members, to ensure contact between the parts, in particular a relatively extensive contact.

[0096] The thermal bridge element 23, or each thermal bridge element 23, is configured for the transmission or dissipation of heat from the back wall 8B of the bay 6 towards a respective frontal closure panel 8D ' of the bay itself.

[0097] The thermal bridge element 23, or each thermal bridge element 23, has the function of creating a preferential conduction path between the internal heat sources of the satellite module and the front panels 8D' exposed to space, so as to transfer the thermal energy towards radiating surfaces obtained from such panels and dissipate it. In this way, the element 23 allows evacuating heat, for example heat possibly present in the central volume 16 of the module (which can be for example in communication with the central volume of a propulsion module that generates heat) or heat generated by the avionics equipment present inside the considered bay, thereby contributing to maintaining the temperatures of the subsystems within operational limits.

[0098] Figures 11-14 show with different views another embodiment of a bay of a first type, with walls 8A, 8C, 8C, 8E and with an intermediate upright 8F. In this case, a star-tracker 22 and a support 24 for a connector for connection of a solar panel mounted on the satellite body (in particular, a fixed panel) are associated with the outer side of the upright 8F.

[0099] In various embodiments, the frontal closure of a bay is obtained in a segmented manner and comprises at least one reduced-height front panel arranged between a side wall and the intermediate upright, and at least one further front panel that completes the frontal closure of the bay, wherein said further panel can be configured for supporting external components and / or at least one external interface.

[0100] The segmented frontal closure allows, in various embodiments, selective access to thesubsystems contained in the bay, permitting the removal or opening of only one panel without the need to completely disassemble the frontal paneling of the module 4 or compartment. Furthermore, the segmentation allows reserving dedicated frontal portions for the integration of external components and / or interfaces, while simultaneously maintaining the structural continuity and modularity of the bay.

[0101] In the case of figures 11-14, between each side wall 8C and the intermediate upright 8F extend reduced-height front panels 8D ', similar to those previously described, as well as further front panels 8D" that complete the frontal closure of the bay 6, and which are configured for supporting external components of the module, such as, for example, one or more antennas 25 - for example for telemetry / telecommand (TM / TC) and / or for payload data transmission (PDT) - and one or more supports 26 for connectors for connection of one or more deployable wing-type solar panels (not shown) with which the satellite may be equipped. In various embodiments, therefore, one or more front panels (8D', 8D") also perform a support function for external components and / or for interfaces, maintaining the modularity of the module or compartment.

[0102] In figure 11 , some of the threaded fasteners 27, 28 used to fix the back wall 8B and the side walls 8C to the base wall 8Afrom below are visible (similar fasteners can be used for fixing the upper wall 8E to the walls 8B and 8C), while in figure 12 some onboard components 10 mounted on the back wall 8B are visible, as well as the connection connectors 14c' and 14d' at the base wall 8A.

[0103] From figure 12 it is also noted how, in possible embodiments, some onboard components 10' can also be mounted on the inner side of at least one of the side walls 8C. Moreover, as noted in figure 13, components of various nature 10" (for example avionics or electrical connection) can also be mounted at the outer side of the side walls 8C (for example in the case of using bays of a second type as described later).

[0104] In various embodiments, each side wall of a first-type bay comprises, substantially at a lower and / or upper end, a fastening flange protruding outward and preferably provided with through holes, said flange obtaining a mechanical interface for fastening an adjacent second-type bay.

[0105] In particular from figures 13 and 14 it is noted how, in preferred embodiments, each side wall 8C has, substantially at the lower end, a fastening flange 8C", protruding outward, provided with respective through holes. Possibly, similar flanges or flanges having a similar function could be provided substantially at the upper end of each side wall 8C. As clarified later, in various embodiments the flanges 8C" can be used to secure a bay of a second type 6' to a bay of a first type 6, the bay of the second type 6' being free from side walls.

[0106] In various embodiments, in fact, a satellite module or compartment comprises at least onebay whose supporting structure includes at least a base wall and a back wall, but is devoid of side walls. Such a second-type bay is configured to be arranged adjacent to at least one first-type bay provided with side walls, such that at least one side wall of the first-type bay also laterally delimits the second-type bay, obtaining a shared lateral delimitation between adjacent bays of different types.

[0107] Figures 15-18 show precisely with different views an example of such a second-type bay, designated by 6', whose supporting structure 8' includes in the depicted example only the base wall 8A and the back wall 8B, it being understood that the structure 8' preferably also includes a respective upper wall 8E, as well as a respective intermediate upright 8F of the types previously described. The combined use of first-type bays 6 and second-type bays 6' allows reducing mass and structural complexity, since one or more lateral delimitations can be shared between adjacent bays, while still maintaining the possibility of mounting onboard units, components and standardized connectors in the second-type bays 6'.

[0108] Also in the bays 6' various onboard units or components 10 can be mounted on the inner side of the backwall 8B, and on the base wall 8A connectors of the type already indicated with 14c' and 14d' can be mounted. In the example shown, moreover, one or more brackets 14a for supporting corresponding standardized connectors 14e, 14f for transporting signals (data) and / or power are cantilever-mounted on the inner side of the back wall 8B.

[0109] The base wall 8A of a structure 8' of a second-type bay 6' is configured for fastening to an adjacent first-type bay 6, particularly at least at the base wall of the latter.

[0110] In particular, in various embodiments, the base wall 8A of the supporting structure 8' of a second-type bay 6' comprises, at the oblique sides of its peripheral profile, a plurality of through apertures for fastening members, wherein said apertures are configured to cooperate with the through holes of a flange belonging to a side wall of a first-type bay 6, so that a portion of the base wall 8A of the second-type bay 6' can overlap onto said flange and be fastened thereto via threaded fastening members.

[0111] In the particular case illustrated in figures 15-18, at the oblique sides of the trapezoidal profile of the base wall 8A, along its edges, through apertures 8A' for threaded fasteners are therefore provided, via which said base wall 8A can be fastened to one of the flanges 8C" of figures 13-14, which protrude outward from the oblique sides of the respective base wall. In practice, for fastening purposes, the portion of the base wall 8A of a structure 8' of a bay 6' is overlapped onto a respective flange 8C" of the base wall 8A of the structure 8 of a bay 6, so as to axially align the respective through holes, into which the necessary threaded fasteners are then engaged. It will be appreciated that a similar construction can also be provided for fastening the upper wall 8E of a structure 8' tothe upper wall 8E of a structure 8. Again in figures 15 and 16, 12a' designate some seats or through holes substantially at the lateral edges of the back wall 8B of the structure 8', which are intended to receive threaded members used for fastening to said back wall 8B the lateral edges of respective side walls 8C of a structure 8 of a contiguous bay 6.

[0112] In various embodiments, a module or compartment comprises four bays arranged peripherally, wherein two first-type bays 6 are set in mutually opposite positions and two second-type bays 6' are set in mutually opposite positions, so that each second-type bay 6' is arranged adjacent to a corresponding first-type bay 6 and is laterally delimited by a side wall 8C of said first-type bay. Such a case is shown in figures 19 and 20, which show with different views an example of a four-bay avionics module, obtained from the composition of two first-type bays 6 according to figures 9-14 and two second-type bays 6' according to figures 15-18.

[0113] As noted, in the depicted example, bays of similar structure are in opposite positions, i.e., the two bays 6 are opposite each other, and similarly the bays 6' are opposite each other. As can be understood, in this way, each bay 6' is arranged adjacent to a bay 6, with a side wall 8C of the latter also laterally delimiting a bay 6'. The arrangement in opposing pairs of structurally similar bays allows, in various embodiments, obtaining a more balanced distribution of masses and loads, as well as making the configuration of the mechanical and electrical interfaces between the various bays more repeatable. This symmetry also facilitates the reconfiguration of the module or compartment for different missions, since a pair of opposite bays can be replaced or varied while keeping the overall architecture of the compartment and the coupling conditions with the other bays unchanged.

[0114] The combined presence of first-type bays 6 and second-type bays 6' allows locally modulating the level of lateral 'closure' of the structure according to integration and mass requirements. In particular, a second-type bay 6' can be employed when the lateral delimitation can be shared with an adjacent first-type bay 6, thereby reducing the number of structural parts while still maintaining the possibility of housing onboard assemblies and obtaining standardized connections. In this way, the shared lateral delimitation constitutes a common partition between adjacent bays obtained without introducing dedicated separate elements.

[0115] As can be seen, in the case of embodiments of the type exemplified in figures 15-20 the satellite has at least one compartment 2a, particularly an avionics compartment, which comprises at least one module formed by a plurality of bays arranged laterally one relative to the other, wherein the plurality of bays comprises first-type bays 6 and one or more second-type bays 6'. The first-type bays 6 have a respective first structure 8 which comprises at least a base wall 8A, a back wall 8B and two side walls 8C that extend laterally from the back wall 8B, particularly with oppositeinclinations, and optionally an upper wall 8E generally parallel to the base wall 8A, at least a front panel 8D, 8D' being removably connectable to said first structure 8. The second-type bays 6' have a respective second structure 8' which comprises at least a base wall 8A, a back wall 8B, and optionally an upper wall 8E generally parallel to the base wall 8A, at least a front panel 8D, 8D' being removably connectable to said second structure 8'. Also in this case the bays of the plurality of bays 6, 6' each comprise:

[0116] - at least one onboard unit 10, for example a sensing unit and / or a computing unit and / or an electronic unit,

[0117] - a plurality of mechanical connection elements 12a, 12b, 12c, configured to allow mutually fixing together different bays of the plurality of bays 6, 6' arranged adjacent to each other, and - at least one electrical connection element for the transmission of power and / or signal. Also in this case the at least one electrical connection element of a bay of the plurality of bays 6, 6' can be configured to provide power and / or signal connection with at least one different bay of the plurality of bays 6, 6', so as to guarantee a power and / or signal communication between onboard units 10 belonging to different bays of the plurality of bays 6, 6'.

[0118] In this case the bays of the plurality of bays 6, 6' are mutually fixed via the plurality of mechanical connection elements 12a, 12b, 12c in such a way that, when a second-type bay 6' is arranged adjacent to a first-type bay 6, a side wall 8C of said first-type bay 6 also laterally delimits said second-type bay 6', providing a common lateral delimitation.

[0119] The characteristics and advantages of the present invention are clear from the description provided.

[0120] The organized architecture of one or more compartments of the satellite via a generic number of bays coupled in a peripheral direction facilitates the preparation of functional systems and subsystems that can be assembled in a simple and rapid manner. Each bay is an independent subassembly, specialized for functionality, which can be integrated and tested independently.

[0121] The arrangement of the bays in a peripheral sense allows defining a central volume, delimited by the back walls of the bays, increases volumetric efficiency and allows a more rational use of space. Each bay can contribute both to the housing of systems and subsystems, and to the transmission of structural loads. The proposed architecture allows obtaining compartments with the same or a different number of bays that can be superimposed on different levels, allowing an expansion of the satellite in an axial direction.

[0122] Each bay has a defined shape and volume, defining a general standard adaptable to already available avionics and technology. The mechanical interface means, for the mutual fasteningbetween the bays and / or for the mutual fastening between multiple compartments, are standardized, simplifying and making the integration phase unambiguous. Each bay can be provided with electrical / electronic interface means that are standardized, which minimizes wiring and simplifies the interconnection between bays and their integration, as well as the development of new third-party bay models.

[0123] Naturally, the construction details and the forms of execution can be widely varied with respect to what has been described and illustrated without thereby departing from the scope of the present invention, as defined by the attached claims.

[0124] As previously indicated, the presence of a specific compartment 2c is not strictly necessary for the purposes of implementing the invention. In possible variant embodiments, in fact, the payload could be installed on the avionics compartment 2a, particularly within its central volume 16 and / or housed in at least one bay 6 of the avionics compartment 2a dedicated to the payload. However, nothing in principle prevents also obtaining the payload compartment via a plurality of bays of the described type.

[0125] In possible embodiments, at least one bay, or each bay, can comprise at least one voltage regulator device (exemplified as VR in figure 7), configured to guarantee the integrability of components present on the satellite that operate at different supply voltages.

[0126] In possible embodiments, within a same bay, or each bay, or more generally in the satellite, at least one communication protocol converter device (exemplified as CD in figure 7) can be present, in order to guarantee the possibility of communication between all onboard components.

[0127] In possible embodiments, at least one bay, or each bay, can be provided with a device for electrical isolation of power and / or signal (exemplified as ID in figure 7), configured to guarantee the isolation of the bay itself both in predetermined circumstances, such as in the case of an intrusion of the satellite, and in the event of a malfunction of one or more onboard components that could affect the integrity of the satellite itself.

Claims

CLAIMS1. A satellite (1), in particular a microsatellite, having one or more compartments (2a, 2b, 2c),wherein at least one compartment, in particular an avionics compartment (2a), comprises at least one module (4) formed by a plurality of bays (6, 6') arranged laterally with respect to one another, wherein said plurality of bays comprises at least some first bays (6) and, optionally, at least one or more second bays (6'),wherein at least one first bay (6) of the plurality of bays (6, 6') comprises:- a structure (8) having at least a base wall (8A), a back wall (8B) and two side walls (8C) that extend laterally from the back wall (8B), in particular with opposite inclinations, a front panel (8D, 8D') being removably connectable to said structure (8),- at least one onboard unit (10), such as a sensing unit and / or a computing unit and / or an electronic unit,- a plurality of mechanical connection elements (12a, 12b, 12c), configured to enable different bays of the plurality of bays (6, 6') arranged adjacent to each other to be mutually fixed together, and- at least one electrical connection element for the transmission of power and / or signal,wherein the at least one electrical connection element of a bay of the plurality of bays (6, 6') is configured to establish power and / or signal connection with at least one different bay of the plurality of bays (6, 6'), so as to ensure power and / or signal communication between onboard units (10) belonging to different bays of the plurality of bays (6, 6'), andwherein the bays of the plurality of bays (6, 6') are mutually fixed via the plurality of mechanical connection elements (12a, 12b, 12c) in such a way that a side wall (8C) of a first bay (6) of the plurality of bays (6, 6') is adjacent to a side wall (8C) of a further first bay (6) of the plurality of bays (6, 6'), or else, when a second bay (6') free of side walls (8C) is arranged adjacent to one said first bay (6), the side wall (8C) of the first bay (6) also laterally delimits said second bay (6'), providing a common lateral delimitation.

2. The satellite according to claim 1 , wherein the bays of the plurality of bays (6, 6') are fixed to each other in a perimetric direction of the satellite (1), so as to define a central volume (16) that is delimited by the back walls (8B) of the plurality of bays (6, 6').

3. The satellite according to claim 1 or claim 2, wherein the structure (8) of each first bay (6) of the plurality of bays (6, 6') comprises a base wall (8A), with the two side walls (8C) extending substantially each from a respective longitudinal side of the back wall (8B), preferably fixed at their lower edge to the base wall (8A), and with the back wall (8B) extending substantially between the two side walls (8C).

4. The satellite according to any one of claims 1 -3, wherein the structure (8, 8') or the base wall (8A) of each bay of the plurality of bays (6, 6') has a substantially trapezoidal cross-section in a plane parallel to the base wall (8A).

5. The satellite according to any one of claims 1 -4, wherein the structure (8, 8') of each bay of the plurality of bays (6, 6') is provided with reinforcement elements (17a, 18), configured to achieve a predetermined stiffness of the same structure (8, 8').

6. The satellite according to any one of claims 1-5, wherein the one or more compartments (2a, 2b, 2c) comprises / comprise at least two compartments arranged stacked on each other.

7. The satellite according to any one of claims 1-6, wherein the at least one module (4) comprises at least a first module (4) and a second module (4) stacked on each other, the first module (4) and the second module (4) each being formed by a respective plurality of bays (6, 6'), the pluralities of bays (6, 6') of the first module (4) and the second module (4) including the same number of bays or a different number of bays.

8. The satellite according to any one of claims 1 -7, wherein at least one bay of the plurality of bays (6, 6') comprises a voltage regulator device (VR) configured to ensure integrability between different components of said at least one onboard unit (10) which operate at different voltages.

9. The satellite according to any one of claims 1 -8, wherein at least one bay of the plurality of bays (6, 6') comprises a communication protocol converter device (CD) configured to ensure compatibility between different components of said at least one onboard unit (10) of a same bay or of different bays.

10. The satellite according to any one of claims 1 -9, wherein at least one bay of the plurality of bays (6, 6') comprises a device for electrical isolation of power and / or signal (ID).

11. The satellite according to any one of claims 1-10, wherein the at least one electrical connection element of a bay of the plurality of bays (6, 6') is compliant with, or standardized with respect to, the at least one electrical connection element of each other bay of the plurality of bays (6, 6').

12. The satellite according to any one of claims 1-11, wherein the plurality of mechanical connection elements (12a, 12b, 12c) of a bay of the plurality of bays (6, 6') is compliant with, or standardized with respect to, a corresponding plurality of mechanical connection elements (12a, 12b, 12c) of each other bay of the plurality of bays (6, 6').

13. The satellite according to claim 6, wherein the at least two compartments have respective mechanical connection elements (12c) that are compliant with, or standardized with respect to, one another, to enable fixing of the at least two compartments (2) in a stacked condition.

14. The satellite according to claim 7, wherein the first module (4) and the second module (4) have respective mechanical connection elements (12c) that are compliant with, or standardized with respect to, one another, to enable fixing of the first module (4) and the second module (4) in a stacked condition.

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