Method and space system for ejecting at least two satellites from a launcher

Abstract

Method for ejecting at least two satellites (2; 2a, 2b) from a launcher (3), said at least two satellites (2; 2a, 2b) being oriented along a central axis coinciding with a longitudinal axis (A) of the launcher and joined together side-by-side, the method comprising: an initial orientation of the direction of the longitudinal axis (A) of the launcher, so as to form a predetermined angle (θ) with an orbital velocity axis (X), the predetermined angle (θ) being between 20° and 60°, a first separation comprising the simultaneous separation from the launcher (3) of said satellites (2; 2a, 2b) along the longitudinal axis (A) with a first predetermined differential velocity (dV1) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining joined together, then a second separation comprising the separation of said satellites (2;2a, 2b) relative to each other along an axis perpendicular to the central axis and contained in the orbital plane, with a second predetermined differential velocity (dV2). Figure from the abstract: Figure 5;

Description

Title of the invention: Method and space system for ejecting at least two satellites from a launcher technical field

[0001] This disclosure relates to a method for ejecting at least two satellites from a launch vehicle. The satellites are initially arranged side-by-side and retained on the launch vehicle. This disclosure also relates to a space system comprising a launch vehicle and at least two satellites, adapted for implementing the method. Prior art

[0002] There is a need to eject, from a launcher, at least two elongated satellites, initially arranged side-by-side and held on the launcher, minimizing the risks of collision between the satellites during ejection and then in orbit. Summary

[0003] The present disclosure proposes a solution to this problem.

[0004] A method is proposed for ejecting at least two satellites from a launcher, said at least two satellites being oriented along a central axis coinciding with a longitudinal axis of the launcher and secured together side-by-side. The method comprises an initial orientation of the direction of the longitudinal axis of the launcher, so as to form a predetermined angle with an orbital velocity axis, the predetermined angle being between 20° and 60°, in particular between 25° and 50°, preferably between 30° and 40°.The method comprises a first separation including the simultaneous separation of said satellites from the launcher along the longitudinal axis with a first predetermined differential velocity relative to the launcher, said at least two satellites remaining attached to each other, and then a second separation including the separation of said satellites from each other along an axis perpendicular to the central axis and contained in the orbital plane, with a second predetermined differential velocity.

[0005] Said at least two satellites are initially retained by the launcher, in particular on the last stage of the launcher.

[0006] The two separations are sequenced, the second separation occurring after the first separation, even if the time between the two separations can be very short.

[0007] The first separation occurs along a first ejection axis parallel to the longitudinal axis of the launcher. The second separation occurs along a second ejection axis. The second ejection axis is perpendicular to the central axis and in the plane of the orbit.

[0008] The ejection method has the advantages of being simple and minimizing the risk of collision.

[0009] Indeed, thanks to this method, a short- and long-term collision between the satellites and with the launcher can be avoided. In particular, the principle of the first separation followed by the second separation makes it possible to avoid the ejection of the satellites one after the other from the launcher. In such a solution of ejecting the satellites one after the other from the launcher, the ejection forces would act with a lever arm on the last stage of the launcher, generating a very high rotational speed on it, which makes the solution impossible.

[0010] The features described in the following paragraphs may optionally be implemented independently of each other or in combination with each other:

[0011] The projection of the first differential velocity between the launcher and said satellites onto the orbital velocity axis is preferably between 0.02 ms⁻¹ and 0.6 ms⁻¹. The projection of the second differential velocity between said satellites onto the orbital velocity axis is preferably between 0.02 ms⁻¹ and 0.6 ms⁻¹.

[0012] The first or second differential velocity is a vector defined by an angle and a magnitude. In this disclosure, the first or second differential velocity is characterized by its projection onto the orbital velocity axis, a value that is important for collision avoidance, the angle being adaptable depending on the specific case.

[0013] Each satellite preferably has a length greater than its width. In this case, the satellites are initially elongated along the longitudinal axis of the launcher. In particular, each satellite may have a length-to-width ratio of at least two. Each satellite is preferably initially secured at one longitudinal end on the launcher. The satellites are preferably identical, at least in external shape.

[0014] It should be noted that, because of such a length of each satellite, the ejection cone, that is to say the cone in which the satellite can be ejected without collision problems, is very small.

[0015] The number of satellites may be two. Alternatively, the number of satellites may be three, or even four. Preferably, the number of satellites is four or less.

[0016] The satellites can initially be connected to the launcher using at least one first hold-down and release mechanism (HDRM). The first hold-down and release mechanism can be configured to release the satellites for the implementation of the first Separation. The process is initiated, in particular, by the launcher. At least one first ejector, or in particular four first ejectors, may be positioned between the launcher and each of the satellites. This at least first ejector preferably extends substantially along the longitudinal axis. This at least first ejector is advantageously initially pre-charged to store a sufficient initial amount of energy to impart the initial differential velocity during the first separation.

[0017] Thus, the first holding and release device or devices allows the satellites and the launcher to be separated during the first separation, while the first ejector(s) allow the differential velocity to be imparted during the first separation.

[0018] The satellites can be initially linked together using at least one second hold-and-release device (HDRM), in particular at least two such HDRMs. The second hold-and-release device can be configured to release the satellites for the implementation of the second separation. At least one second ejector, in particular four second ejectors, can be initially arranged between the satellites. Said at least one second ejector preferably extends substantially perpendicularly with respect to the central axis, which was initially coincident with the longitudinal axis, and in the plane of the orbit. Said at least one second ejector is advantageously initially pre-charged so as to store a second quantity of energy suitable for imparting said second differential velocity during said second separation.

[0019] Thus, the second holding and releasing device allows the satellites to be separated from each other during the second separation, while the second ejector(s) allow the differential velocity to be imparted during the second separation.

[0020] The second quantity of energy is advantageously less than the first quantity of energy, so that the mass of the ejectors is lower on the satellites than on the launcher. The objective of the ejection is to provide a given differential velocity, and the energy required to achieve this depends, in particular, on the mass of the bodies, in this case the satellites, to be ejected. In a specific example, corresponding to a given satellite mass, the first quantity of energy can be between 20 J and 30 J, while the second quantity of energy can be between 5 J and 10 J. The values ​​of the first and second quantities of energy will, of course, need to be adjusted according to the mass of the satellites and may therefore fall outside these ranges.

[0021] The first and / or second ejectors are preferably made of springs which, when released from their tension, propel themselves along their axis. In particular, the first and / or second ejectors are preferably separate from the so-called tilted springs, which are used to rotate the satellites. Indeed, these The latter, although conceivable, are complex and risky because of the length of each satellite; the slightest rotation generated during ejection risks generating a collision between the satellites.

[0022] The first and / or second holding and release devices are preferably composed of systems that generate few shocks. Furthermore, if there are several holding and release devices per interface, it is preferable that these devices be able to release with good simultaneity. The time interval between the implementation of the first separation and the implementation of the second separation is preferably less than 10 seconds, and even more preferably less than 5 seconds. This time interval is the time required to detect the first separation and trigger the second separation. During this time interval, the satellites, still connected to each other, may rotate due to the first ejection. This rotation is minimized, therefore the aim is to maintain a short time interval between the first and second separations.The rotation angle of the satellites during this time period is preferably less than or equal to 10°.

[0023] According to another aspect, in combination with the foregoing, a space system for ejecting at least two satellites is proposed, comprising a launcher receiving the satellites initially attached to the launcher, said satellites being oriented along a central axis coinciding with the longitudinal axis of the launcher and secured together side-by-side. The space system includes a guidance module for an initial orientation in the direction of the longitudinal axis of the launcher, so as to form a predetermined angle with an orbital velocity axis, the predetermined angle being between 20° and 60°, in particular between 25° and 50°, preferably between 30° and 40°.The space system further comprises a first control module for a first separation comprising the simultaneous separation of the satellites from the launcher along the longitudinal axis with a first predetermined differential velocity relative to the launcher, the satellites remaining attached to each other, and then a second control module for a second separation comprising the separation of the satellites from each other along an axis transverse to the central axis with a second predetermined differential velocity.

[0024] The first control module can be configured to command the first separation upon remote command from a ground station. One of the satellites can be configured to detect the implementation of the first separation and, following this detection, to trigger the second separation commanded by the second control module.

[0025] Each satellite may have a length-to-width ratio of at least two. Each satellite is preferably initially retained at one longitudinal end on the launcher, in particular on the last stage of the launcher.

[0026] The launcher may include at least one first hold-and-release device (HDRM) to initially hold the satellites attached to the launcher. The first control module may be configured to actuate the first hold-and-release device so as to separate the satellites from the launcher. The launcher may include at least one first ejector, in particular four first ejectors, disposed between the launcher and each of the satellites and extending substantially along the longitudinal axis. Said at least one first ejector is preferably initially pre-charged so as to store a first amount of energy suitable for imparting the first differential velocity during the first separation.

[0027] The satellites may include at least one second hold-and-release device (HDRM) for initially connecting them together. The second control module may be configured to actuate the second hold-and-release device so as to separate the satellites from each other. The satellites may include at least one second ejector, in particular four second ejectors, initially arranged between the satellites and extending substantially transversely, in particular perpendicularly, with respect to the central axis. Said at least one second ejector is preferably initially pre-charged so as to store a second quantity of energy suitable for imparting said second differential velocity during the second separation. Brief description of the drawings

[0028] Other features, details and advantages will become apparent upon reading the detailed description below, and upon analysis of the accompanying drawings, on which:

[0029] [Fig.1] schematically shows in side view an example of a spatial system.

[0030] [Fig.2] schematically shows, in space, an initial orientation of the direction of the longitudinal axis of the launcher in the implementation of a method for ejecting two satellites from a launcher according to an example.

[0031] [Fig.3] schematically shows a first separation in the implementation of the process according to an example.

[0032] [Fig.4] schematically shows the behavior of the satellites after the first separation illustrated in [Fig.3].

[0033] [Fig.5] schematically shows a second separation in the implementation of the process according to an example.

[0034] [Fig.6] shows schematically, partially and in perspective, an example of a spatial system. Description of the implementation methods

[0035] The drawings and description below contain, essentially, elements of a definite nature. They may therefore not only serve to better understand this disclosure, but also contribute to its definition, if necessary.

[0036] In the various figures, the same reference numerals designate identical or similar elements. For the sake of brevity, only the elements that are useful for understanding the described embodiment are shown in the figures and are described in detail below.

[0037] In the following description, when reference is made to absolute positional qualifiers, such as the terms "front", "back", "top", "bottom", "left", "right", etc., or relative positional qualifiers, such as the terms "above", "below", "superior", "inferior", etc., or to orientational qualifiers, such as "horizontal", "vertical", etc., reference is made, unless otherwise specified, to the orientation of the figures.

[0038] Reference is now made to [Fig. 1]. This shows a space system 1 for ejecting at least two satellites 2, in this example exactly two satellites referenced respectively as 2a and 2b. The space system 1 comprises a launcher 3 receiving, on its upper stage, the satellites 2 initially connected to the launcher 3. As can be seen in [Fig. 1], the launcher 3 extends along a longitudinal axis A. Initially, the satellites 2 are oriented, around a central axis, along the longitudinal axis A and are secured together side-by-side.

[0039] In this example, the satellites 2 are identical, at least in external shape. Each satellite 2 has a length greater than its width. In the illustrated example, each satellite 2 has a length-to-width ratio (L1 / Z2) of at least two. For example, the length L1 is 4 m and the width Z2 is 1 m. The thickness of each satellite 2, not visible in this example, can be equal to the width, i.e., 1 m in this example. Each satellite 2 is initially retained at a longitudinal end 4 on the launcher 3, at the level of the launcher's final stage.

[0040] In the illustrated example, the launcher 3 includes at least one first hold-and-release device (HDRM) 5 for initially holding the satellites 2 connected to the launcher 3. In this example, each satellite 2 is initially connected to the launcher 3 by means of a first hold-and-release device (HDRM) 5 disposed at the interface with the launcher. The launcher 3 also includes at least one first ejector 6, in this example four first ejectors 6, disposed between the launcher 3 and each of the satellites 2 and extending along the longitudinal axis A.

[0041] Furthermore, still in this example, the satellites 2 include at least one second hold-and-release device (HDRM) 7 to initially connect them to each other, on either side of the central axis. In this example, the satellites 2 are initially connected to each other by means of two second holding and release devices 7 arranged between the satellites. The satellites 2 have at least one second ejector 8, in this example four second ejectors 8 initially arranged between the satellites 2 and extending transversely, in this example perpendicularly with respect to the central axis which coincides with the longitudinal axis A when the satellites are attached to the launcher.

[0042] We will now describe the method of ejecting said at least two satellites 2 from the launcher 3, into space above the earth T, with reference to figures 2 to 5.

[0043] The method first includes, as illustrated in [Fig.2], piloting an initial orientation of the direction of the longitudinal axis A of the launcher 3, so as to form a predetermined angle 0 with an orbital velocity axis X, the predetermined angle 0 being between 20° and 60°, in this example equal to 40° at + or - 5°.

[0044] To do this, as illustrated in [Fig.1], the space system 1 includes a steering module 10 for the initial orientation of the direction of the longitudinal axis A of the launcher 3, so as to form the predetermined angle 0 with the orbital velocity axis X.

[0045] The method then includes, as illustrated in [Fig.3], the triggering of a first separation comprising the simultaneous separation out of the launcher 3 of the satellites 2 along the longitudinal axis A of the launcher 3, with a first predetermined differential velocity relative to the launcher 3, the satellites 2 remaining together as illustrated.

[0046] The space system 1 comprises, in the example illustrated in [Fig. 1], a first control module 11 configured to control this first separation. The first control module 11 can be configured to control the first separation upon remote command from a ground station. The satellite includes, in particular, a telecommunications module for communication with the ground station. The first control module 11 is specifically configured to actuate the first hold-and-release device 5 so as to separate the satellites 2 from the launcher 3. The first hold-and-release device 5 is configured to release the satellites for the implementation of the first separation, this implementation being initiated, in particular, by the launcher 3.The first ejectors 6, formed by axial springs in this example, are initially pre-charged so as to store a first quantity of energy suitable for imparting the first differential velocity during the first separation.

[0047] The projection of the first differential velocity, named dVxl, between the launcher 3 and the satellites 2 onto the orbital velocity axis X is preferably between 0.02 ms⁻¹ and 0.6 m.s⁻¹*, in this example equal to 0.6 ms⁻¹, being represented by a double arrow. The first differential velocity dVl is indeed projected onto the orbital velocity axis X according to the formula: dVlx = dVl*cos0.

[0048] Thus, when the first separation is engaged, the first control module 11 causes each holding and release device 5 to transition from an initial holding state to a satellite release state from the launcher 3. During this release, the presence of the first ejectors 6 provides an initial differential velocity between the launcher 3 and the satellites 2. This prevents collision between the launcher 3 and the group of satellites 2.

[0049] The method includes, after the initial separation, a period of time, illustrated in [Fig. 4], during which the satellites 2, still connected to each other, may rotate about their own axis, moving away from the longitudinal axis A. This period is preferably less than 10 seconds, in particular 5 seconds. At the end of this period, the satellites 2 have, for example, rotated through an angle α, which, in the illustrated example, is a maximum of 10°. Thus, the central axis of the connected satellites may deviate from the longitudinal axis of the launcher. The central axis of the connected satellites may also remain aligned with the longitudinal axis of the launcher. The angles and energy levels stored are chosen to operate within a range of deviation angles of the central axis from the longitudinal axis. In the following discussion, we consider the case where the central axis remains aligned with the longitudinal axis.

[0050] The process then includes, after this time period, triggering a second separation illustrated in [Fig.5], comprising the separation of the two satellites 2 from each other along an axis B perpendicular to the central axis (before separation) and included in the orbital plane, with a second predetermined differential velocity.

[0051] The space system 1 includes, in the example illustrated in [Fig. 1], a second control module 12 which is configured to control the second separation. One of the satellites 2, in this case satellite 2a, is, for example, configured to detect the implementation of the first separation and to trigger, following this detection, the second separation controlled by the second control module 12. The second control module 12 is configured to actuate the second hold-and-release devices 7 so as to separate said satellites 2a and 2b from each other. The second hold-and-release devices 7 are configured to release the satellites 2 for the implementation of the second separation.The second ejectors 8, which are springs oriented along an axis B perpendicular to the central axis and belonging to the orbital plane in this example, are initially pre-charged so as to store a second quantity of energy suitable for conferring the second differential velocity during the second separation.

[0052] During the activation of the second separation, the second control module 12 causes each holding and release device 7 to transition from an initial holding state to a release state for the satellites 2a and 2b. During this release, the presence of the second ejectors 8 imparts the second differential velocity between the satellites 2. This prevents collisions between the satellites 2, as well as between the launch vehicle 3 and each of the satellites 2.

[0053] The projection of the second differential velocity dVx2 between the satellites 2 onto the orbital velocity axis X is, in this example, between 0.02 ms1 and 0.6 ms'. This projection dVx2 is illustrated by a double arrow in [Fig. 5], being greater than 0.2 ms' in this example. The second differential velocity dV2 is indeed projected onto the orbital velocity axis X according to the equation: dVx2 = dV2*sin0.

[0054] The combination of the first and second differential velocities results in the launcher 3 and each of the satellites 2a and 2b having a different velocity, significantly reducing or even eliminating the risk of collision between them. It should be noted that satellite 2a moves backward while satellite 2b moves forward, relative to launcher 3.

[0055] The combination of the two successive separations, also called ejections, ensures a certain velocity differential between each object, that is, between each of the launcher and the satellites. This differential velocity prevents each object from approaching beyond an acceptable collision risk during subsequent orbits. A differential velocity is obtained between the launcher 3 and satellite 2a, denoted dV_3 / 2a, projected onto the orbital velocity axis X, which satisfies the following equation: dV_3 / 2a = dV_x - dV_x² / 2. Furthermore, a differential velocity is obtained between satellite 2a and satellite 2b, denoted dV_2a / 2b, which satisfies the following equation: dV_2a / 2b = dV_x².

[0056] The initial angle 0 can be optimized, subject to a minimum velocity difference between each object, to minimize the energy required by each first or second ejector and thus minimize the mass of the first and second ejectors, as well as the residual rotations generated by the dispersions on them.

[0057] The ejectors can be positioned with a large gap between them for each ejection, which minimizes the residual rotations generated by force / energy dispersions between each ejector.

[0058] The second quantity of energy is advantageously less than the first quantity of energy, so that the mass of the ejectors is lower on the satellites 2 than on the launcher 3. The objective of the ejection is to provide a given differential velocity, and the energy required to achieve this depends in particular on the mass of the bodies, in this case the satellites, to be ejected. In a specific example, corresponding to a satellite mass Given the given values, the first energy quantity can be between 20J and 30J, while the second energy quantity can be between 5J and 10J. The values ​​of the first and second energy quantities will, of course, need to be adjusted according to the mass of the satellites and may therefore fall outside these ranges.

[0059] The first separation occurs along a first ejection axis parallel to the longitudinal axis A of the launcher 3, which coincides with the central axis when the satellites are attached to the launcher. The second separation occurs along a second ejection axis, which is axis B. The second ejection axis lies in the orbital plane and is perpendicular to the central axis defined with respect to all the satellites before their separation.

[0060] In this example, the number of satellites is two. The satellites are then arranged opposite each other on either side of the central axis. Alternatively, the number of satellites can be three, or even four. The satellites are then arranged around the central axis. The number of satellites is preferably four or less.

[0061] Figure 6 shows the space system 1 with two satellites 2, 2a and 2b. Figure 6 shows, in addition to the length L1 and width 12 of each satellite 2, the depth p of each satellite 2, 2a or 2b, which, in this example, is the same dimension as the width 12. The shape of each satellite 2 can vary. In particular, the cross-section with width 12 and depth p can be different from a square, being, for example, rectangular, trapezoidal, or any other shape. The satellites are, for example, shown with the same dimensions in terms of their length, width, and depth, but can also, for example, be of different dimensions. The satellites can, for example, have identical masses, but can also, for example, have different masses.

[0062] The first and second ejectors 6 and 8 are axial springs, in particular compression springs. In particular, they are not, preferably, tilted springs. The first ejectors 6, four in number for each satellite 2, are distributed at the base of each satellite 2, in particular in a regular manner, for example in a square, not all being aligned, as illustrated for example in [Fig. 6]. The second ejectors 8, for example four in number for each pair of satellites, are arranged, for example, in pairs near the longitudinal ends of the satellites 2. In each pair of second ejectors, the second ejectors 8 are, for example, arranged side-by-side in the depth of the satellites 2. A holding and releasing device 7 is, for example, located at the center of the length of the satellites between the two satellites.

Claims

Demands

1. A method for ejecting at least two satellites (2; 2a, 2b) from a launcher (3), said at least two satellites (2; 2a, 2b) being oriented along a central axis coinciding with a longitudinal axis (A) of the launcher and joined together side-by-side, the method comprising: a. an initial orientation of the direction of the longitudinal axis (A) of the launcher, so as to form a predetermined angle (0) with an orbital velocity axis (X), the predetermined angle (0) being between 20° and 60°, b. a first separation comprising the simultaneous separation from the launcher (3) of said satellites (2; 2a, 2b) along the longitudinal axis (A) with a first predetermined differential velocity (dVl) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining joined together, then c.a second separation comprising the separation of said satellites (2; 2a, 2b) from each other along an axis perpendicular to the central axis and included in the orbital plane, with a second predetermined differential velocity (dV2).

2. A method according to claim 1, wherein the projection of the first differential velocity (dV1) between the launcher (3) and said satellites (2; 2a, 2b) onto the orbital velocity axis (X) is between 0.02 ms1 and 0.6 ms1 and wherein the projection of the second differential velocity (dV2) between said satellites onto the orbital velocity axis (X) is between 0.02 ms1 and 0.6 ms1

3. A method according to any one of the preceding claims, wherein each satellite (2; 2a, 2b) has a ratio of its length (L1) to its width (Z2) of at least two, and wherein each satellite (2; 2a, 2b) is initially retained at a longitudinal end (4) on the launcher (3).

4. A method according to any one of the preceding claims, wherein said satellites (2; 2a, 2b) are initially connected to the launcher (3) by means of at least one first hold and release device (HDRM) (5) which is configured to release said satellites (2; 2a, 2b) for the implementation of the first separation, said implementation being initiated in particular by the launcher (3), and in which at least one first ejector (6), in particular four first ejectors (6), is disposed between the launcher (3) and each of said satellites (2; 2a, 2b) and extends substantially along the longitudinal axis (A), said at least one first ejector (6) being initially pre-charged so as to store a first quantity of energy suitable to confer said first differential velocity (dVl) at the first separation.

5. A method according to any one of the preceding claims, wherein said satellites (2; 2a, 2b) are initially linked together by means of at least one second holding and release device (HDRM) (7), in particular at least two, which is configured to release said satellites (2; 2a, 2b) for the implementation of the second separation, and wherein at least one second ejector (8), in particular four second ejectors (8), is initially disposed between said satellites (2; 2a, 2b) and extends perpendicularly with respect to the central axis and in the plane of the orbit, said at least one second ejector (8) being initially pre-charged so as to store a second quantity of energy suitable for imparting said second differential velocity (dV2) at said second separation.

6. A method according to any one of the preceding claims, wherein the time interval (dt) between the implementation of the first separation and the implementation of the second separation is less than 10s, preferably less than 5s.

7. A space system (1) for ejecting at least two satellites (2; 2a, 2b), comprising a launcher (3) receiving said satellites (2; 2a, 2b) initially connected to the launcher (3), said satellites (2; 2a, 2b) being oriented along a central axis coinciding with a longitudinal axis (A) of the launcher and secured together side-by-side, said space system (1) being characterized in that it comprises: a. a guidance module with an initial orientation in the direction of the longitudinal axis (A) of the launcher, so as to form a predetermined angle (0) with an orbital velocity axis (X), the predetermined angle (0) being between 20° and 60°, b. a first control module for a first separation comprising the simultaneous separation out of the launcher (3) of said satellites (2; 2a, 2b) along the longitudinal axis (A) with a first predetermined differential velocity (dV1) relative to the launcher (3), said satellites (2; 2a, 2b) remaining together, then c. a second control module for a second separation comprising the separation of said satellites (2; 2a, 2b) from each other along an axis transverse to the central axis with a second predetermined differential velocity (dV2).

8. Space system (1) according to claim 7, wherein the first command module is configured to command the first separation on remote command from a ground station and wherein one of said satellites (2; 2a, 2b) is configured to detect the implementation of the first separation and to trigger, following said detection, the second separation commanded by the second command module.

9. Space system (1) according to any one of claims 7 to 8, wherein each satellite (2; 2a, 2b) has a ratio of its length (L1) to its width (Z2) of at least two, and wherein each satellite (2; 2a, 2b) is initially retained at a longitudinal end (4) on the launcher (3).

10. Space system (1) according to any one of claims 7 to 9, wherein the launcher (3) comprises at least one first hold and release device (HDRM) (5) for initially holding said satellites (2; 2a, 2b) connected to the launcher (3), said first control module being configured to actuate said first hold and release device (5) so as to separate said satellites out of the launcher (3), and wherein the launcher (3) comprises at least one first ejector (6), in particular four first ejectors (6), disposed between the launcher (3) and each of said satellites (2; 2a, 2b) and extending substantially along the longitudinal axis (A), said at least one first ejector (6) being initially pre-charged so as to store a first quantity of energy suitable for imparting said first differential velocity (dVl) at the first separation.

11. Space system (1) according to any one of claims 7 to 10, wherein said satellites (2; 2a, 2b) comprise at least one second hold and release device (HDRM) (7) for initially linking them together, said second control module being configured to actuate said second hold and release device so as to separate said satellites (2; 2a, 2b) from each other, said satellites (2; 2a, 2b) comprising at least one second ejector (8), in particular four second ejectors (8), initially disposed between said satellites (2; 2a, 2b) and extending substantially transversely with respect to the central axis, said at least one second ejector (8) being initially pre-charged so as to store a second quantity of energy suitable for imparting said second differential velocity (dV2) at the second separation.

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