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

Abstract

Disclosed is 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 secured 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 (θ) with an orbital speed axis (X), the predetermined angle (θ) 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 speed (dV1) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining secured 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 speed (dV2).

Description

Description Title: Method and space system for the ejection of at least two satellites from a launcher Technical field

[0001] The present disclosure relates to a method for ejecting at least two satellites from a launcher. The satellites are initially arranged side by side and retained on the launcher. The present disclosure also relates to a space system comprising a launcher and at least two satellites, suitable 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 retained on the launcher, minimizing the risks of collision between the satellites during ejection and then in orbit. Summary

[0003] This 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 coincident with a longitudinal axis of the launcher and secured to each other 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 being between 30° and 40°.The method comprises a first separation comprising the simultaneous separation from the launcher of said satellites along the longitudinal axis with a first predetermined differential speed relative to the launcher, said at least two satellites remaining integral with each other, then a second separation comprising the separation of said satellites from each other along an axis perpendicular to the central axis and included in the orbital plane, with a second predetermined differential speed.

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

[0006] The two separations are sequenced, with the second separation occurring after the first separation, although the time between the two separations may be very short.

[0007] The first separation takes place along a first ejection axis parallel to the longitudinal axis of the launcher. The second separation takes place 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 process has the advantages of being simple and minimizing the risk of collision.

[0009] In fact, thanks to this process, it is possible to avoid a short- and long-term collision between the satellites and with the launcher. In particular, the principle of the first separation followed by the second Separation avoids the ejection of satellites one after the other from the launcher. In such a solution of ejecting 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 rotation speed on it, which makes the solution impossible.

[0010] The features set out 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 on 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 on 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 norm. In the present 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 ratio of its length to its width at least equal to two. Each satellite is preferably initially retained at one longitudinal end on the launcher. The satellites are preferably identical, at least in terms of external shape.

[0014] It should be noted that, due to such a length of each satellite, the ejection cone, that is, the cone into 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. The number of satellites is preferably less than or equal to four.

[0016] The satellites may be initially connected to the launcher using at least one first hold-down and release device, also called HDRM for "hold down and release mechanism" in English. The first hold-down and release device may be configured to release the satellites for the implementation of the first separation. The implementation is triggered in particular by the launcher. At least one first ejector, in particular four first ejectors, may be arranged between the launcher and each of the satellites. Said at least one first ejector preferably extends substantially along the longitudinal axis. Said at least one first ejector is advantageously initially precharged so as to store a first quantity of energy adapted to confer the first differential speed during the first separation.

[0017] Thus, the or each first holding and releasing device makes it possible to separate, during the first separation, the satellites and the launcher, while the first ejector(s) makes it possible to confer the differential speed during the first separation.

[0018] The satellites may be initially connected to each other using at least one second hold and release device (HDRM), including at least two such HDRM devices. The second holding and releasing device may be configured to release the satellites for implementing the second separation. At least one second ejector, in particular four second ejectors, may be initially arranged between the satellites. Said at least one second ejector preferably extends substantially perpendicularly to the central axis, which was initially merged with the longitudinal axis, and in the plane of the orbit. Said at least one second ejector is advantageously initially precharged so as to store a second quantity of energy adapted to confer said second differential speed during said second separation.

[0019] Thus, the or each second holding and releasing device makes it possible to separate, during the second separation, the satellites from each other, while the second ejector(s) makes it possible to confer the differential speed 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 speed and the energy to achieve this depends in particular on the mass of the bodies, in this case the satellites, to be ejected. In a particular example, corresponding to a given satellite mass, the first quantity of energy can be between 20J and 30J while the second quantity of energy can be between 5J and 10J. The values of the first and second quantities of energy will of course have to be adapted according to the mass of the satellites and may then not be included in these ranges of values.

[0021] The first ejectors and / or the second ejectors are preferably constituted by springs which, when released from their constraint, push back along their axis. In particular, the first ejectors and / or the second ejectors are preferably distinct from the so-called tilted springs, capable of rotating the satellites. Indeed, the latter, although conceivable, are complex and risky because of the length of each satellite, the slightest rotation generated during the ejection risking generating a collision between the satellites.

[0022] The first and / or second holding and release devices are preferably constituted by systems generating few shocks. In addition, in the case where there are several holding and release devices per interface, it would preferably be necessary for the holding and release devices to be able to release with good simultaneity. The time lapse between the implementation of the first separation and the implementation of the second separation is preferably less than 10 s, more preferably less than 5 s. This time lapse is the one that is necessary to detect the first separation and trigger the second separation. During this time lapse, the satellites still attached to each other may possibly rotate on themselves because of the first ejection. We seek to avoid this rotation as much as possible, so we seek to maintain a reduced time lapse 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 above, there is provided a space system for ejecting at least two satellites, comprising a launcher receiving the initially connected satellites to the launcher, said satellites being oriented along a central axis coincident with the longitudinal axis of the launcher and secured to each other side by side. The space system comprises a module for controlling an initial orientation of the direction of the longitudinal axis of the launcher, so as to form a predetermined angle with an orbital speed axis, the predetermined angle being between 20° and 60°, in particular between 25° and 50°, preferably being between 30° and 40°.The space system also comprises a first module for controlling a first separation comprising the simultaneous separation from the launcher of the satellites along the longitudinal axis with a first predetermined differential speed relative to the launcher, the satellites remaining attached to each other, then a second module for controlling 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 speed.

[0024] The first control module may be configured to control the first separation upon remote command from a ground station. One of the satellites may be configured to detect the implementation of the first separation and to trigger, following this detection, the second separation controlled 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, particularly at the last stage of the launcher.

[0026] The launcher may comprise at least one first holding and releasing device (HDRM) for initially holding the satellites connected to the launcher. The first control module may be configured to actuate the first holding and releasing device so as to separate the satellites from the launcher. The launcher may comprise 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 precharged so as to store a first quantity of energy adapted to impart the first differential velocity during the first separation.

[0027] The satellites may comprise at least one second holding and releasing device (HDRM) for initially connecting them together. The second control module may be configured to actuate the second holding and releasing device so as to separate the satellites from each other. The satellites may comprise at least one second ejector, in particular four second ejectors, initially arranged between the satellites and extending substantially transversely, in particular perpendicularly, relative to the central axis. Said at least one second ejector is preferably initially precharged so as to store a second quantity of energy adapted to confer said second differential speed 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 analyzing the attached drawings, in which:

[0029] [Fig. 1] shows schematically 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 of 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 Figure 3.

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

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

[0035] The drawings and description below contain, for the most part, elements of a certain character. 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 references designate identical or similar elements. For the sake of brevity, only the elements which are useful for understanding the embodiment described are shown in the figures and are described in detail below.

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

[0038] Reference is now made to Figure 1. There we see a space system 1 for ejecting at least two satellites 2, in this example exactly two satellites referenced respectively 2a and 2b. The space system 1 comprises a launcher 3 receiving, at its last stage, the satellites 2 initially connected to the launcher 3. As visible in Figure 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 to each other side by side.

[0039] The satellites 2 are in this example identical, at least in terms of their external shape. Each satellite 2 has a length greater than its width. In the example illustrated, the satellites 2 each have a ratio of its length L1 to its width 12 at least equal to two. For example, the length L1 is equal to 4 m and the width 12 is equal to 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 last stage of the launcher 3.

[0040] In the illustrated example, the launcher 3 comprises at least one first holding and releasing 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 using a first holding and releasing device (HDRM) 5 arranged at the interface with the launcher. The launcher 3 also comprises at least one first ejector 6, in this example four first ejectors 6, arranged 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 comprise at least one second holding and releasing device (HDRM) 7 to initially connect them together, on either side of the central axis. In this example, the satellites 2 are initially connected to each other using two second holding and releasing devices 7 arranged between the satellites. The satellites 2 comprise 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 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 process 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 firstly comprises, as illustrated in Figure 2, the control of 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 speed 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 Figure 1, the space system 1 comprises a control 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 speed axis X.

[0045] The method then comprises, as illustrated in Figure 3, the triggering of a first separation comprising the simultaneous separation from the launcher 3 of the satellites 2 along the longitudinal axis A of the launcher 3, with a first predetermined differential speed relative to the launcher 3, the satellites 2 remaining integral with each other 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 may be configured to control the first separation on remote command from a ground station. The satellite comprises in particular a telecommunications module with the ground station. The first control module 11 is in particular configured to actuate the first holding and releasing device 5 so as to separate the satellites 2 from the launcher 3. The first holding and releasing device 5 is configured to release the satellites for the implementation of the first separation, this implementation being triggered in particular by the launcher 3.The first ejectors 6, formed by axial springs in this example, are initially preloaded so as to store a first quantity of energy suitable for imparting the first differential speed during the first separation.

[0047] The projection of the first differential velocity, named dVx1, between the launcher 3 and the satellites 2 on the orbital velocity axis X is preferably between 0.02 ms -1 and 0.6m.s -1 , in this example equal to 0.6 ms -1 , being represented by a double arrow. The first differential velocity dV1 is in fact projected onto the orbital velocity axis X according to the formula: dV1x = dV1*cos0.

[0048] Thus, when the first separation is triggered, the first control module 11 causes each holding and releasing device 5 to move from an initial holding state to a state of releasing the satellites 2 from the launcher 3. During this release, the presence of the first ejectors 6 provides a first differential speed between the launcher 3 and the satellites 2. This makes it possible to avoid collision between the launcher 3 and the group of satellites 2.

[0049] The method comprises, after the first separation, a period of time, illustrated in Figure 4, during which the satellites 2, still secured to each other, possibly rotate on themselves while moving away from the longitudinal axis A. The period of time is preferably less than 10s, in particular 5s. At the end of the period of time, the satellites 2 have, for example, rotated through an angle a which, in the example illustrated, is 10° at most. Thus, the central axis of all the satellites secured to each other can deviate from the longitudinal axis of the launcher. The central axis of all the satellites secured to each other can also remain aligned with the longitudinal axis of the launcher. The angles and energy levels stored are chosen to operate within a range of angles of deviation of the central axis from the longitudinal axis. In the following, we consider the case where the central axis remains aligned with the longitudinal axis.

[0050] The method then comprises, after this period of time, the triggering of a second separation illustrated in Figure 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 speed.

[0051] The space system 1 comprises, 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 the 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 holding and releasing devices 7 so as to separate said satellites 2a and 2b from each other. The second holding and releasing 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 preloaded so as to store a second quantity of energy adapted to confer the second differential speed during the second separation.

[0052] When the second separation is triggered, the second control module 12 causes each holding and releasing device 7 to move from an initial holding state to a state of releasing the satellites 2a and 2b from each other. During this release, the presence of the second ejectors 8 provide the second differential speed between the satellites 2. This makes it possible to avoid collision between the satellites 2 but also between the launcher 3 and each of the satellites 2.

[0053] The projection of the second differential velocity dVx2 between satellites 2 on the orbital velocity axis X is in this example between 0.02 ms -1 and 0.6 ms -1 . This dVx2 projection is illustrated by a double arrow in Figure 5, being in this example greater than 0.2 ms -1 The second differential velocity dV2 is in fact projected onto the orbital velocity axis X according to the equation: dVx2 = dV2*sin0.

[0054] The result of the combination of the first differential speed and the second differential speed is that launcher 3 and each of satellites 2a and 2b each have a different speed, greatly reducing or even eliminating the risk of collision between them. It should be noted that satellite 2a goes backwards while satellite 2b goes forwards, relative to launcher 3.

[0055] The combination of the two successive separations, also called ejections, makes it possible to respect a certain speed differential between each object, that is to say between each of the launcher and the satellites. This differential speed allows each object not to get closer than an acceptable risk of collision during the following orbits. We obtain a differential speed between launcher 3 and satellite 2a, called dV_3 / 2a, projected on the orbital speed axis X, which verifies the following equation: dV_3 / 2a=dVx1-dVx2 / 2. Furthermore, we obtain a differential speed between satellite 2a and satellite 2b, called dV_2a / 2b which verifies the following equation: dV_2a / 2b=dVx2.

[0056] The initial angle 0 can be optimized, subject to a minimum speed difference between each object, to minimize the energy required by each first or second ejector and therefore 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 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 speed and the energy to achieve this depends in particular on the mass of the bodies, in this case the satellites, to be ejected. In a particular example, corresponding to a given satellite mass, the first quantity of energy can be between 20J and 30J while the second quantity of energy can be between 5J and 10J. The values of the first and second quantities of energy will of course have to be adapted according to the mass of the satellites 2 and may then not be included in these ranges of values.

[0059] The first separation takes place 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 takes place along a second ejection axis which is axis B. The second axis ejection is in the plane of the orbit and perpendicular to the central axis defined in relation to all the satellites before their separation.

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

[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 the width 12 of each satellite 2, the depth p of each satellite 2, 2a or 2b, which, in this example, is of the same dimension as the width 12. The shape of each satellite 2 can vary. In particular, the section of 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 represented with the same dimensions according to 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 be of different masses.

[0062] The first and second ejectors 6 and 8 are axial springs, in particular compression springs. In particular, they are preferably not 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 for example arranged two by two 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 arranged in the center of the length of the satellites between the two satellites.

Claims

Claims

1. 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 coincident with a longitudinal axis (A) of the launcher and secured to each other 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 (6) with an orbital velocity axis (X), the predetermined angle (6) 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 (dV1) relative to the launcher (3), said at least two satellites (2; 2a, 2b) remaining secured to each other, 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 speed (dV2).

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

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

4. Method according to any one of the preceding claims, wherein said satellites (2; 2a, 2b) are initially connected to the launcher (3) using at least one first holding and releasing device (HDRM) (5) which is configured to release said satellites (2; 2a, 2b) for the implementation of the first separation, said implementation being triggered in particular by the launcher (3), and wherein at least one first ejector (6), in particular four first ejectors (6), is arranged 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 precharged so as to store a first quantity of energy adapted to confer said first differential velocity (dV1) during the first separation.

5. A method according to any one of the preceding claims, wherein said satellites (2; 2a, 2b) are initially connected to each other using at least one second holding and releasing device (HDRM) (7), in particular at least two, which is configured to release said satellites (2; 2a, 2b) for implementing the second separation, and wherein at least one second ejector (8), in particular four second ejectors (8), is initially arranged between said satellites (2; 2a, 2b) and extends perpendicularly to the central axis and in the plane of the orbit, said at least one second ejector (8) being initially precharged so as to store a second quantity of energy adapted to confer said second differential velocity (dV2) during said second separation.

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

7. 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 coincident with a longitudinal axis (A) of the launcher and secured to each other side by side, said space system (1) being characterized in that it comprises: a. a module for controlling 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 speed axis (X), the predetermined angle (9) being between 20° and 60°, b. a first module for controlling 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 speed (dV1) relative to the launcher (3), said satellites (2; 2a, 2b) remaining integral with each other, then c.a second module for controlling 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 speed (dV2).

8. Space system (1) according to claim 7, wherein the first control module is configured to control the first separation on remote order 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 controlled by the second control module.

9. Space system (1) according to any one of claims 7 to 8, in which each satellite (2; 2a, 2b) has a ratio of its length (L1) to its width ( / 2) at least equal to two, and in which 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 holding and releasing 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 holding and releasing device (5) so as to separate said satellites from the launcher (3), and wherein the launcher (3) comprises at least one first ejector (6), in particular four first ejectors (6), arranged 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 preloaded so as to store a first quantity of energy adapted to confer said first differential velocity (dV1) during 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 holding and releasing device (HDRM) (7) for initially connecting them together, said second control module being configured to actuate said second holding and releasing 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 arranged between said satellites (2; 2a, 2b) and extending substantially transversely relative to the central axis, said at least one second ejector (8) being initially precharged so as to store a second quantity of energy adapted to confer said second differential velocity (dV2) during the second separation.

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