Space platform and method for deploying such a platform

Abstract

Space platform and method for deploying such a platform. The present invention relates to a space platform (10) formed of a plurality of identical space modules (20) assembled together to form an elongated structure along a platform axis (X), said structure having a substantially homogeneous mass distribution along the platform axis (X). Figure for the abstract: Figure 2

Description

Title of the invention: Space platform and method for deploying such a platform

[0001] The present invention relates to a space platform formed of a plurality of space modules and a method for deploying such a platform.

[0002] The invention falls within the field of large orbital structures, usable for different purposes.

[0003] The prior art known in the field of such orbital structures is based mainly on functional architectures of the ISS type ('International Space Station' in English or 'Station Spatiale Internationale' in French).

[0004] In such an architecture, the various subsystems composing it are assembled on a supporting structure. The attitude of this structure is generally controlled by CMG (Control Momentum Gyroscope) type actuators or reaction wheel type actuators, which have the disadvantage of representing a significant mass, or by thrusters, which have the disadvantage of consuming propellants.

[0005] This type of architecture therefore quickly reveals its limitations when the dimensions become significant, on the order of several hundred meters. These limitations are expressed particularly in terms of controllability and mass.

[0006] More specifically, the ISS-type architecture presents a functional and centralized structure that has drawbacks in terms of mass optimization and the distance between the different functions. This then leads to losses, for example, due to heat dissipation in the power chains. This type of architecture also presents difficulties in attitude control, mainly due to significant deflections related to the large dimensions exposed to solar flux.

[0007] As a result, existing architectures cannot be used when it is necessary to increase these dimensions in order to provide functions not currently provided.

[0008] The present invention aims to remedy these drawbacks and to propose a solution allowing easy increase in the dimensions of a space platform while ensuring its attitude controllability.

[0009] To this end, the invention aims at a space platform formed of a plurality of identical space modules assembled together to form an elongated structure along a platform axis, said structure having a substantially homogeneous mass distribution along the platform axis.

[0010] According to other advantageous aspects of the invention, the platform comprises one or more of the following features, taken individually or in all technically possible combinations:

[0011] - each space module comprises a payload and is self-sufficient in energy and in heat dissipation;

[0012] - each module can be deployed independently of the other modules;

[0013] - each module comprises a plurality of solar panels forming a first plan of the corresponding module;

[0014] the modules being assembled so that their first planes form a useful plane of the platform;

[0015] - the platform being intended to be placed in orbit such that its useful plane is permanently oriented towards the Sun;

[0016] - the platform being intended to be placed in an orbit such that the axis of platform is oriented with the Earth's nadir;

[0017] - the platform having a structure in "1";

[0018] - the platform further comprising a robot capable of moving on a surface of the platform and configured to deploy elements forming each module and / or to perform platform maintenance operations;

[0019] - the platform further comprising a mass capable of moving on a surface of the platform is configured to control the platform's attitude along at least one axis;

[0020] - said mass is configured to control the attitude of the platform at least in shoelace;

[0021] - said mass is a robot;

[0022] - the platform comprising at least one fixed or movable actuator / propeller by a robot;

[0023] - said structure has a substantially homogeneous mass distribution in a useful plane formed by the platform axis and an axis perpendicular to this platform axis.

[0024] The invention also relates to a method for deploying a space platform as defined above, comprising the following steps:

[0025] - deployment of each space module constituting the platform;

[0026] - assembly of the space modules together.

[0027] The invention also relates to a space module intended to be deployed in space to form a space platform, the space module being formed of a predetermined number of elements all having the same flat shape in an extension plane of the corresponding element, each element being of a predetermined type, at least one element forming a payload.

[0028] According to other advantageous aspects of the invention, the module comprises one or more of the following features, taken individually or in all technically possible combinations:

[0029] - the type of each element is chosen from the group comprising:

[0030] - solar panel;

[0031] -radiator;

[0032] - payload;

[0033] - the module comprising a plurality of solar panels arranged around the payload in the foreground;

[0034] preferably, the first plane being intended to be oriented towards the Sun;

[0035] - the module comprising a plurality of radiators arranged in a second plane perpendicular to the foreground;

[0036] preferably, the second plane being intended to be aligned with the nadir of the Earth;

[0037] - the radiators are arranged substantially opposite the payload and away from this one;

[0038] - said flat shape is suitable for being inscribed in a circle having a cross-section transverse section of a cargo compartment of a space launcher;

[0039] - said flat shape has identical dimensions for all elements;

[0040] - the payload includes a data storage and processing server and means of communicating this data with another space module of the same space platform and / or with an external communication station;

[0041] - the payload further includes a heat pump to dissipate heat;

[0042] - each solar panel comprises a frame and a veil stretched over this frame and featuring printed solar cells;

[0043] preferably, the frame being made of carbon-epoxy;

[0044] preferably, the veil being made of polyetheretherketone (PEEK);

[0045] - at least some of the elements are connected by connection interfaces, hinges and / or springs;

[0046] - the module being devoid of an energy storage device.

[0047] The invention also aims at a space platform comprising a plurality of modules as described above, the modules being assembled together.

[0048] The invention also aims at a method of piloting a space platform elongated along a platform axis, the method comprising a phase of active control of the attitude of the platform by displacement of a mass on a control surface of the platform, the control surface extending along the platform axis.

[0049] According to other advantageous aspects of the invention, the method comprises one or more of the following features, taken individually or in all technically possible combinations:

[0050] - active attitude control is carried out with respect to a single axis;

[0051] - active attitude control is yaw control;

[0052] - the method further comprising a phase of passive attitude control by the Earth's gradient;

[0053] - passive attitude control is pitch and / or roll control;

[0054] - active control of the platform's attitude is further achieved by actuation of one or more actuators / propellers;

[0055] - said mass is a deployment and / or maintenance robot;

[0056] - the method further comprising a movement of one or more actuators / propelled by the robot;

[0057] - the control surface has a substantially flat surface extending over the entire platform length.

[0058] The invention also aims at a control system for a space platform, comprising technical means for implementing the process as defined above.

[0059] The invention will become clearer upon reading the following description, given solely by way of non-limiting example, and made with reference to the accompanying drawings in which: - [Fig.1] [Fig.1] is a schematic view of several positions of a space platform according to the invention with respect to the Earth; - [Fig.2] [Fig.2] is a detailed schematic front (part A) and side (part B) view of the space platform of [Fig.1], the space platform comprising a plurality of space modules; - [Fig.3] [Fig.3] is a detailed front (part A) and side (part A) view of one of the space modules of [Fig.2]; - [Fig.4] [Fig.4] is a schematic view illustrating an example of the composition of the spatial module of [Fig.3]; - [Fig.5] [Fig.6] [Fig.7] Figures 5 to 7 are different views illustrating the implementation of a deployment method for the space module of [Fig.3] according to different embodiments of the invention; - [Fig.8] [Fig.8] is a flowchart of a process for controlling the space platform of [Fig.2].

[0060] Several positions of a space platform 10 according to the invention have indeed been illustrated on [Fig.1].

[0061] These examples illustrate in particular the positioning of the space platform 10 in relation to the earth T.

[0062] Thus, as illustrated in this figure, the space platform 10 is intended to be positioned in an Earth orbit such that at least one of the surfaces of this platform, hereafter referred to as the useful surface, is permanently exposed to the Sun. An example of such an orbit is a 6-18 hour SSO orbit at an altitude of 1400 km.

[0063] This platform 10 is, for example, intended to form a global server for storing and processing data in space, also called in English "data center".

[0064] To this end, platform 10 is capable of communicating with terrestrial communication systems to receive and transmit data, as will be explained in more detail later. In some examples, this communication can be achieved via intermediate communication systems, such as a relay constellation between platform 10 and Earth. This solution minimizes the ground infrastructure required and allows for direct access to end users (mobile devices, for example).

[0065] Alternatively, the space platform 10 presents any other space application requiring the use of large-scale structures in space. In some examples, the space platform 10 may present several applications combined within the same structure.

[0066] The space platform 10 is illustrated in more detail in [Fig.2].

[0067] With reference to this [Fig.2], the platform 10 extends along a platform axis X which is aligned along the nadir direction N of the earth T.

[0068] More particularly, as can be seen in part A of figure 2 illustrating the front view of the space platform 10, this platform has a J-shaped structure. In other words, the structure of the space platform 10 has a main part 12 extending along the platform axis X and two secondary parts 14, 16 extending perpendicularly to the platform axis X and positioned at each end of the main part 12.

[0069] In addition, the secondary parts 14, 16 are positioned symmetrically with respect to a plane passing through the main part 12 at the center and extending along the platform axis X.

[0070] By way of example, the space platform 10 has a span L1 along the platform axis X of between 200 and 300 m, preferably approximately equal to 210 m. Also by way of example, the space platform 10 has a span perpendicular L2 to the platform axis X which is between 60 and 90 m, for example approximately equal to 80 m.

[0071] The space platform 10 forms a flat structure. Also by way of example, the extent L3 exhibiting the thickness of this platform is between 10 and 20 m and for example substantially equal to 11 m.

[0072] Of course, other dimensions are also conceivable for this platform 10. In addition, as will be explained in more detail later, these dimensions can be modified during the operation of the platform 10 for the addition and / or removal of space modules.

[0073] In general, the ratio between the thickness (i.e. the extent L3) of this platform 10 and its axial extent (i.e. the extent L1) is between 1 / 100 and 1 / 10.

[0074] The space platform 10 is formed of a plurality of identical space modules 20. These modules 20 are assembled together to form the structure of the platform as illustrated in [Fig. 2]. For example, the main part 12 of the space platform 10 is formed of five space modules 20 aligned along the platform axis X. Each of the secondary parts 14, 16 is formed of three space modules 20 aligned along axes perpendicular to the platform axis X.

[0075] The space modules 20 also have substantially the same mass, so that the platform 10 has a substantially homogeneous mass distribution along the platform axis X and advantageously, in a useful plane PU as defined below. By "substantially homogeneous," we mean here a homogeneous mass distribution at the scale of the space modules 20, i.e., considering these space modules 20 within the platform 10 as mass points.

[0076] The space modules 20 are interconnected to form a substantially rigid structure of the platform 10. The means of connection may include clipping different parts forming these space modules 20 or tightening or screwing, as will be explained in more detail later.

[0077] Advantageously, the space modules 20 are interconnected only structurally.

[0078] Alternatively, the space modules 20 are also interconnected by electrical cables and / or data exchange cables.

[0079] Advantageously, each space module 20 is self-sufficient in energy and heat dissipation. In other words, each space module 20 exhibits autonomy from the other modules within the same platform 10.

[0080] In some cases, the space modules 20 can be interconnected with each other by means of wireless link in order to exchange data with each other.

[0081] Even more advantageously, the various space modules 20 can be added, removed and / or replaced autonomously after the platform has been placed in orbit spatial 10. In other words, these operations do not affect the functioning of the other space modules 20.

[0082] An example of a space module 20 is illustrated in more detail in [Fig.3].

[0083] With reference to [Fig. 3], the spatial module 20 is formed of a predetermined number of elements 21, 22, 23 assembled together to form a first plane PI visible when the module 20 is viewed from the front (part A of [Fig. 3]) and a second plane P2 visible when the spatial module 20 is viewed from the side (part B of [Fig. 3]). The second plane P2 is perpendicular to the first plane PI.

[0084] When the space modules 20 are assembled within the space platform 10, their first planes PI form a single plane, also called the useful plane PU. This useful plane then corresponds to the useful surface of the space platform 10 as described above. In particular, the useful plane PU formed by the PI planes is permanently oriented towards the Sun.

[0085] As regards the second planes P2, as illustrated in [Fig. 3], those belonging to the space modules 20 of the main part 12 and to the space modules 20 located at the center of the secondary parts 14, 16, also form a single plane that is perpendicular to the useful plane PU. This plane is hereafter referred to as the secondary plane PS. The secondary plane PS thus remains permanently in the shadow of the useful plane PU. The second planes P2 of the space modules 20 belonging to the extremities of the secondary parts 14, 16, on the other hand, form planes parallel to the secondary plane PS. These parallel planes also remain in the shadow of the useful plane PU.

[0086] Returning to the description of [Fig. 3], the space module 20 comprises a predetermined number of elements all having the same flat shape in the first plane PL

[0087] Each element 21, 22, 23 is of a predetermined type. In other words, the space module 20 comprises only elements of which each type is predetermined. Advantageously, the type of each element is chosen from among three predetermined types, namely a solar panel, a radiator, and a payload.

[0088] In the example of [Fig. 3], the element 21 arranged in the center of the first plane PI forms a payload. The elements 22 arranged around the payload 21 form solar panels 22. Finally, the elements 23 arranged in the second plane P2 form radiators.

[0089] With reference to [Fig.4], by way of example, the space module 20 may include a single payload 21, eighteen solar panels 22 and four radiators 23.

[0090] The elements 21, 22, 23 of the spatial module 20 all form the same shape, which is for example the hexagonal shape.

[0091] More generally, these elements 21, 22, 23 have a shape that can be inscribed in a circle C representing a cross-section of a cargo compartment of a space launcher, such as the fairing of this launcher.

[0092] Advantageously, this shape is adapted to fit optimally within this circle C while having straight faces that allow these elements to be attached to the other elements. Preferably, each element 21, 22, 23 has the shape of a regular polygon with a number n of sides greater than or equal to 5. Advantageously, the number n of sides is equal to 6. This shape facilitates the assembly of the elements together to form a spatial module and the assembly of the modules together to form the platform 10.

[0093] Furthermore, the elements 21, 22, 23 can all have the same dimensions in the first plane PI. In other words, only the thickness of these elements can vary.

[0094] Advantageously, elements of the same type have the same thickness as illustrated in [Fig. 4]. In particular, the solar panels 22 and the radiators 23 can have a thickness of approximately 3 cm, while the payload 21 can have a thickness of approximately 30 cm.

[0095] The payload 21, for example, has a housing comprising a plurality of electronic components.

[0096] The electronic components are chosen in particular according to the application given to the space platform 10.

[0097] For example, the electronic components include a data storage server and communication means. These communication means include, for example, local communication means with the payload of other space modules 20 or external communication means, for example with a ground station or any other external communication station such as a satellite.

[0098] The payload 21 advantageously includes a heat pump enabling the dissipation of heat accumulated in the casing forming the payload 21.

[0099] In some embodiments, the payload 21 may include any other element implementing the operation of the space module 10 or of the space platform 10.

[0100] The solar panels 22 are arranged around the payload 21 as illustrated in [Fig. 3]. In particular, the solar panels 22 can form two circles Cl, C2 around the payload 21. The first circle Cl, forming an inner circle, can consist of six solar panels 22, and the second circle C2, forming an outer circle, can consist of twelve solar panels 22.

[0101] Each solar panel 22 comprises a frame made for example of carbon epoxy and a veil stretched over the frame and made for example of polyetheretherketone (PEEK).

[0102] The frame may have a predetermined profile shape (for example, an "H" shape) allowing, for example, the guidance of power supply cables around the periphery of each solar panel. The frame may also have fastening means to the other elements of the same space module 20 or of another space module 20, i.e. to the other solar panels and / or possibly to the payload 21. The means of attachment may include clipping means (using for example bayonet systems), screwing or hinges.

[0103] The veil stretched over the frame advantageously presents solar cells or forms a support for these cells, for example printed on it.

[0104] These solar cells can be of a predetermined type (for example, of the Perovskite type) allowing the generation of electrical energy by methods known per se. The solar cells are connected by cables guided by the structural frame to the other solar panels in order to power the payload 21. In some cases, the payload 21 further includes a power transformer allowing the electrical energy delivered by the solar panels 22 to be converted into usable electrical energy to power the various electrical components of this payload 21. For example, this power transformer can convert the power from 400 V to 48 V.

[0105] The frame of the solar panels 22 allows the sail to be tensioned to the desired level, for example by means of a system of linkages, pulleys and cables, possibly passively by exploiting the differential expansion of the materials to generate the tension force. Advantageously, the sails can be relaxed during the launch of the launcher so as not to be damaged by acoustic waves during launch.

[0106] The sails stretched within the frame of the solar panels 22 advantageously form the first plane PI of the space module 20, as described previously.

[0107] The radiators 23 are arranged in the second plane P2 advantageously away from the payload 21.

[0108] Furthermore, these radiators 23 are positioned opposite the payload 21 in order to conduct heat generated by this payload. To distance the radiators 23 from the payload 21, a spacer frame 26 can be used.

[0109] The arrangement of the radiators 23 in the second plane P2, aligned with the Earth's nadir (N) and perpendicular to the first plane PI, minimizes the exposure of these radiators 23 to the Earth's infrared radiation. Furthermore, the spacing of the radiators 23 from the payload 21 minimizes infrared irradiation by the payload 21. This ensures their optimal operation.

[0110] With reference to [Fig. 3], the space module 20 has in the first plane PI a first extent L3 aligned with the platform axis X and a second extent L4 perpendicular to this axis. Each of these extents is, for example, between 20 and 30 m. Advantageously, the first extent L3 is substantially equal to 29 m and the second extent L4 is advantageously equal to 27 m.

[0111] Furthermore, the space module 20 also has a thickness L5 which is, for example, between 10 and 20 m and is, for example, approximately 11 m. In other words, the thickness L5 of the space module 20 forms the thickness of the space platform 10.

[0112] Fig. 4 also gives approximate masses of the different elements forming the space module 20.

[0113] Thus, in the example of [Fig.4], the mass of the payload 21 can be substantially equal to 1200 kg, the mass of each solar panel 22 can advantageously be equal to 18 kg and the mass of each radiator 23 can be substantially equal to 274 kg.

[0114] The various elements of the space module 20 are advantageously configured to be stacked one on top of the other in the space launcher and then deployed in space. Thus, the elements constituting the space modules 20 of the space platform 10 can be stacked one on top of the other so that the stack can be carried in a single carrier and then deployed in space.

[0115] Advantageously, the heaviest elements such as the payload 21 are stacked first (in the lower position). Then, the lighter elements such as the radiators 23 are stacked on top, and then the lightest elements such as the solar panels 22 are stacked on top of that in order to minimize stack deflection during launch and thus facilitate launching.

[0116] In general, the stacking method depends on the way these different elements are deployed in space.

[0117] Subsequently, a method for deploying the space modules 20 will be described with reference to Figures 5 to 7 showing different embodiments of this deployment method.

[0118] According to a first embodiment, all the elements or at least the payload 21 and the solar panels 22 are linked together to be deployed in a spiral as illustrated in [Fig.5].

[0119] In particular, according to this embodiment, the payload 21 constitutes the first element of the stack which is then positioned at the beginning of the deployment process.

[0120] The payload 21 can be connected to one of the adjacent solar panels 22, for example by hinges or springs.

[0121] Thus, the stack made up of the solar panels 22 can be rotated to position this first solar panel 22 adjacent to the payload 21. The deployment of the solar panels 22 can thus be continued in a spiral by first forming the first circle Cl around the payload 21 and then the second circle C2 around this payload 21.

[0122] Furthermore, during deployment, the adjacent elements can be connected to each other for example by clipping or screwing or tightening or by any other possible means of fixing.

[0123] This type of deployment can for example be carried out autonomously by the space module 20 itself using motorized hinges or by a robot as will be explained above.

[0124] The radiators 23 can be hung behind the solar panels 22 and the payload 21 in a similar manner.

[0125] According to a second embodiment of the deployment process illustrated in [Fig.6], a robot 30 is used.

[0126] This robot 30 allows, for example, the stack of various elements constituting the space modules 20 to be unfolded by taking each element of the stack one by one and positioning them appropriately. In this case, the assembly of each space module 20 can also be carried out in a spiral, starting this time with the outer circle C2 and moving towards the center to position the payload 21. Alternatively, the deployment can begin with the payload 21 and then by positioning the solar panels 22 in a spiral as in the previous case.

[0127] The robot 30 can also hook the different space modules 20 together or after finishing the assembly of a space module 20, proceed to the assembly of an adjacent space module 20.

[0128] Advantageously, the robot 30 is configured to move for example on a surface of the different space modules 20, for example on the first plane PI of these modules 20.

[0129] For this purpose, the robot 30 has means for attaching itself to this first plane PI, consisting, for example, of magnets or any other available means. Alternatively, the robot 30 uses rails formed, for example, by the frames of the solar panels 22 to move along the first planes PI of the different space modules 20.

[0130] In addition or alternatively, the robot 30 has several arms (for example, 3 or 4). Each arm has grippers at its end for grasping structural elements. Advantageously, the robot 30 is powered by these structural elements, which have contactors positioned periodically along the path taken by the robot on the platform 10. The grippers of the robot 30 therefore have electrical contactors opposite them. Thus, in some examples, the robot 30 does not need batteries.

[0131] After deployment, the robot 30 can be used for the maintenance of the space platform 10 or for piloting this platform 10 as will be explained in more detail later.

[0132] Fig. 7 illustrates a third embodiment of the deployment method for space module 20.

[0133] According to this embodiment, the stack P forming the space module 20 has a particular stacking order and connections.

[0134] According to this embodiment, the solar panels 22 are classified into two groups and each group is connected to the payload 21. The stack P thus formed comprises first of all the radiators 23 at the bottom of the stack with a connection to the payload 21, then the first group of solar panels 22 also connected to the payload 21, then the payload 21 itself and finally, at the top of the stack P, the second group of solar panels 22 also connected to the payload 21.

[0135] Each group of solar panels 22 comprises a plurality of triplets. Each triplet comprises solar panels connected consecutively, thus defining solar panels of 3 rows. The solar panels of row 3 are connected directly to the payload 21, the solar panels of row 2 are connected to those of row 3, and the solar panels of row 1 are connected directly to those of row 2. These connections are, for example, formed by hinges, advantageously motorized.

[0136] In the example of [Fig.7], each group comprises 3 triplets of solar panels 22.

[0137] During a first SI step of the deployment process, the radiators 23 are removed from the stack using the corresponding link.

[0138] Then, during a second step S2 of the process, the radiators 23 are partially opened and a first pair of triplets from the different groups are opened on either side of the payload 21.

[0139] Then, during a third step S3, the radiators 23 are opened completely and a second pair of triplets from the different groups are opened on either side of the payload 21.

[0140] Then, during the fourth step S4, a third pair of triplets from the different groups are opened on either side of the payload 21.

[0141] Thus, at the end of this step, the 3rd rank solar panels 22 form the first circle Cl around the payload 21.

[0142] During a fifth step S5, each triplet is partially unfolded, thus positioning the second-row solar panels adjacent to the third-row panels.

[0143] Finally, during a sixth step S6 of this process, each triplet is fully unfolded, thus positioning the rank 1 solar panels adjacent to those of rank 3 and rank 2.

[0144] Thus, the second circle C2 consisting of solar panels of ranks 1 and 2 is closed.

[0145] Of course, during this deployment process, the deployment of the radiators 23 and the solar panels 22 can be carried out independently.

[0146] Furthermore, in certain embodiments of the deployment method, different space modules 20 can be deployed independently. In such a case, the method also includes a step of connecting different modules together. For this purpose, the robot 30 as defined above can be used.

[0147] A method for piloting the space platform 10 will now be explained with reference to [Fig.8] showing an organizational chart of its steps.

[0148] In particular, this method includes an active attitude control phase of the space platform 10 and a passive attitude control phase.

[0149] The active control phase is implemented for example by a control system 40.

[0150] This control system 40 has, for example, an electronic component embedded in the platform 10. For example, this control system 40 is part of one of the payloads 21 of the space modules 20 constituting the space platform 10.

[0151] Alternatively, the control system 40 is remote. It is implemented, for example, in a ground station or any other remote station, and the corresponding commands are transmitted to the space platform 10 using means of communication known per se.

[0152] The active attitude control phase of the space platform 10 includes a first step 110 of moving a mass on a control surface of the space platform 10.

[0153] Advantageously, the control surface is formed from the useful plane PU of this platform, as explained previously.

[0154] The mass is, for example, formed by the robot 30 as explained previously. This robot 30 is then able to move along the useful plane PU of the space platform and thus modify the mass distribution on this space platform 10. This movement creates a control torque by actively positioning the center of mass relative to the point of application of the forces (essentially from the solar flux).

[0155] In some examples, the mass may be formed or comprise elements movable by a pulley system, or a fluid movable by a fluidic system. In these cases, such a system may be complementary to the robot 30.

[0156] Furthermore, in certain cases, the robot 30 can also be configured to, if necessary, move peripheral solar panels 22 to create larger torques, particularly in the event of repairs which would otherwise require it to be moved and / or prevent it from creating a control torque.

[0157] According to different embodiments of the invention, the robot 30 can be used solely for piloting the space platform or, in addition to this piloting, for the maintenance and / or deployment of the platform 10, as explained previously.

[0158] Advantageously, active attitude control is performed with respect to a single axis. For example, active attitude control is performed only in yaw, thanks in particular to the structure in 1 of the platform 10.

[0159] Thus, the control system 40 can measure the rotation of the space platform around said axis and compensate for this yaw rotation by moving the robot 30 on the useful plane PU of the space platform 10.

[0160] In some embodiments, the active attitude control phase of the space platform 10 may also include a second step 120 which includes the actuation of one or more actuators or thrusters disposed on the space platform 10.

[0161] Prior to the implementation of this step 120, the control module 40 can also actuate the robot 30 to move appropriately one or more actuators / propulsors for example in the useful plane PU of the space platform 10. Then, the control system 40 can determine a propulsion value required for each actuator / propulsor according to its positioning relative to the space platform 10.

[0162] The passive attitude control phase of the space platform 10 is carried out inherently through the local gravitational gradient of the Earth's potential.

[0163] In particular, this passive attitude control includes pitch and roll control of the space platform 10. In other words, the space platform 10 exhibits natural pitch and roll stability.

[0164] Of course, the piloting method may include additional active attitude control along two or three axes.

[0165] For this purpose, one or more robots can be used simultaneously to move in different planes of the space platform 10.

[0166] It is therefore understood that the present invention has a number of advantages.

[0167] First, the invention offers an optimization of the space occupied by the launcher fairing. This is achieved through the specific shape of the elements constituting the space platform, and in particular the space modules forming this platform. This shape is identical for all elements and also has the same dimensions in their respective extension planes. Thus, a single heavy launcher can carry all the elements necessary to deploy the space platform.

[0168] The invention also makes it possible to achieve an optimization of the mass of the space platform.

[0169] In particular, the space platform does not require a battery for its operation because the solar panels point permanently towards the Sun thanks to a special orbit.

[0170] In addition, the solar panels are particularly light given the particular choice of materials constituting them and the architecture of a sail stretched by a frame.

[0171] Furthermore, the use of a heat pump dissipates the heat generated by the payload. This solution minimizes the cooling surface area required by the radiators. Thus, the number of radiators can be reduced while still ensuring the necessary cooling of the payload.

[0172] Furthermore, the radiators are optimally positioned in the shade of the solar panels and are minimally exposed to terrestrial infrared radiation. This allows the radiators to operate under optimal conditions and their surface area to be minimized.

[0173] Finally, the structural bending stresses induced by the solar wind are minimized by distributing the wing load over the entire surface exposed to the Sun of this large platform. This then minimizes the mass of structural reinforcements between the different space modules constituting the space platform.

[0174] The invention also makes it possible to minimize the different types of losses.

[0175] In particular, the space platform according to the invention is made up of identical modules, self-sufficient in power and dissipation. Furthermore, the dimensions of each module are optimized to minimize electrical and thermal power losses.

[0176] The invention also offers various deployment options. These options include, for example, deploying the elements constituting each space module using motorized hinges. Another option involves using an indigenous robot. Finally, another deployment option involves assembling the hexagonal elements individually.

[0177] The invention also has internal advantages of modularity and extensibility.

[0178] In particular, the various space modules can be added later to increase the size of the platform.

[0179] Furthermore, the platform can be piloted in a particularly simple manner. In particular, stability along at least two axes is ensured by the Earth's gravity gradient.

[0180] This allows for a reduction in the mass of the actuators / propellants required for control and avoids their consumption of propellant.

[0181] Furthermore, the invention proposes active yaw control by adjusting the center of mass relative to the center of solar pressure. Thus, a movable mass such as a The indigenous robot adjusts its position to generate or neutralize torques by adjusting the distances between the center of mass and the point resulting from the application of solar radiation pressure forces.

Claims

Demands

1. Space platform (10) formed of a plurality of identical space modules (20) assembled together to form an elongated structure along a platform axis (X), said structure having a substantially homogeneous mass distribution along the platform axis (X).

2. Platform (10) according to claim 1, wherein each space module (20) comprises a payload (21) and is self-sufficient in energy and heat dissipation.

3. Platform (10) according to claim 1 or 2, wherein each module (20) is deployable independently of the other modules (20).

4. Platform (10) according to any one of the preceding claims, wherein each module (20) comprises a plurality of solar panels (22) forming a first plane (FP) of the corresponding module (20); the modules (20) being assembled so that their first planes (FP) form a useful plane (UP) of the platform (10).

5. Platform (10) according to claim 4, intended to be placed in orbit such that its useful plane (PU) is permanently oriented towards the Sun.

6. Platform (10) according to any one of the preceding claims, intended to be placed in an orbit such that the platform axis (X) is oriented with the nadir of the Earth.

7. Platform (10) according to any one of the preceding claims, having a structure in "1".

8. Platform (10) according to any one of the preceding claims, further comprising a robot (30) capable of moving on a surface of the platform (10) and configured to deploy elements (21, 22, 23) forming each module (20) and / or to perform maintenance operations on the platform (10).

9. Platform (10) according to any one of the preceding claims, further comprising a mass (30) capable of moving on a surface of the platform (10) and configured to control the attitude of the platform (10) along at least one axis.

10. Platform (10) according to claim 9, wherein said mass (30) is configured to control the attitude of the platform at least in yaw.

11. Platform (10) according to claim 9 or 10, wherein said mass (30) is a robot (30).

12. Platform (10) according to any one of the preceding claims, comprising at least one actuator / propellant fixed or movable by a robot (30).

13. Platform (10) according to any one of the preceding claims, wherein said structure has a substantially homogeneous mass distribution in a useful plane (PU) formed by the platform axis (X) and an axis perpendicular to this platform axis (X).

14. A method for deploying the platform (10) according to any one of the preceding claims, comprising the following steps: - deployment of each space module (20) constituting the platform (10); - assembly of the space modules (20) together.

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