Satellite release system from launch vehicle

Abstract

The release system (1) for releasing a satellite from a launch vehicle comprises: (i) a torsion bar (2) fixed to the launch vehicle (3) by support means (4) and having a first end (2a) locked against rotation about a longitudinal axis (X) of the torsion bar (2), and a second end (2b) connected to the launch vehicle (3) by hinge means (5) and free to rotate about the longitudinal axis (X); (ii) a torsion bar (2) extending perpendicularly from the torsion bar (2) and having a first end (7a) fixed to the torsion bar (2); (b) at least one launch arm (6) having a bar (7), (b) a guide (8) having a first end (8a) connected to a second end (7b) of the torsion lever (7) and a second free end (8b); (iii) at least one slider (10) integrally fixed to the satellite (18) to be launched and arranged to slidably engage the guide (8); and (iv) a limit stop element (12) designed to act on the torsion lever (7) to stop rotation of the launch arm (6) about the longitudinal axis (X).

Description

[Technical field]

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS This patent application claims priority to Italian Patent Application No. 102020000012415, filed May 26, 2020, the entire disclosure of which is incorporated herein by reference.

[0002] Technical Field The present invention relates to a system for deploying a satellite from a launch vehicle. [Background technology]

[0003] Currently, commonly used ejection systems include a number of coil springs that are integral with the launch vehicle and are adapted to generate thrust to eject the satellite, such an ejection system including launching the satellite along a direction perpendicular to the separation plane.

[0004] As will be appreciated by those skilled in the art, the coil spring must be selected according to the mass of the satellite to be launched, and it will inevitably become necessary to replace the coil spring whenever the mass of the satellite to be launched changes.

[0005] However, despite its popularity, such solutions are associated with a number of problems.

[0006] In fact, the satellite undergoes a rotational movement when launched due to the action of a coil spring.

[0007] Such rotational motion is primarily due to the fact that the springs are never perfectly identical to one another and never stretch relative to one another at the same time.

[0008] The latter aspect stems from the impossibility of precisely synchronizing the actuation of the different ejection elements. As will be clear to those skilled in the art, even a difference of a few microseconds in the actuation of the various ejection elements will result in significant imbalances in the overall thrust experienced by the satellite and thus in its rotation. Other elements of the uncertainty of the satellite in the ejection step and the resulting rotation are the imperfect knowledge of the position of the center of mass and the temporary vibrations of the launch vehicle structure.

[0009] The rotations that the satellite undergoes after being released inevitably lead to a delay in the deployment of the solar panels and therefore to a prolonged use of energy from the batteries on board the satellite. This aspect is a considerable element of risk, since in order to perform "emergency" operations, the energy of the satellite batteries must always be available for the step immediately after release and before the deployment of the solar panels.

[0010] Moreover, if a launch vehicle carries several satellites side-by-side, there is a risk of the satellites colliding with each other if they are released perpendicular to the separation plane. Indeed, it is well known that for release systems using coil spring thrust, it is appropriate to consider a "cone of uncertainty" of about 15° with respect to the intended release direction. Such a condition implies a low occupation of the available space in the launch vehicle's nose cone, limiting the number of satellites that can be carried on the same launch vehicle.

[0011] In addition to the above problems, there is a need for a system that can ensure faster release speeds than those provided by current systems that use a limited number of coil springs as a propulsion means. Indeed, a faster release speed would result in faster movement of the satellite away from the launch vehicle, and therefore potentially faster deployment of the solar panels, which would result in reduced battery power usage and the associated benefits mentioned above.

[0012] For safety reasons, the deployment of the solar panels must take place at a distance of several hundred meters from the launch vehicle, and if the thrust of the ejection does not allow this distance to be reached, resort will be made to battery energy.

[0013] Finally, another need recognized in the art is the ability of the release system to be adaptable to satellites of different masses without the need to replace some of the components, i.e., a need exists for a release system in which the propulsion means is effective regardless of the mass of the satellite. As noted above, the actual coil spring solution must be selected depending on the mass of the satellite to be launched. Summary of the Invention [Means for solving the problem]

[0014] The object of the invention is a system for releasing a satellite from a launch vehicle, characterized in that the release system comprises (i) a torsion bar fixed to the launch vehicle by support means and having a first end locked in rotation about the longitudinal axis of the torsion bar and a second end connected to said launch vehicle by hinge means and free to rotate about said longitudinal axis, (ii) at least one launch arm extending perpendicularly from said torsion bar and comprising (a) a torsion lever having a first end integrally fixed to said torsion bar, (b) a guide having a first end and a second free end connected to the second end of said torsion bar, (iii) at least one slider integrally fixed to the satellite to be launched and arranged to slidably engage with said guide, and (iv) a limit stop element designed to act on said torsion lever to interrupt the rotation of the launch arm about its longitudinal axis.

[0015] Preferably, the first end of the guide is connected to the second end of the torsion lever by a locking joint, which allows the release direction to be changed depending on the angle between the guide and the torsion lever.

[0016] Preferably, the guide includes an energy absorber.

[0017] Preferably, the system comprises a torsion preload element acting on said first end of the torsion bar. These means set the bar start torsion already upstream of the load torsion due to the rotation of the launch arm engaging the satellite to be launched. This makes it possible to adapt the elastic force required for an effective release according to the mass of the satellite to be launched.

[0018] Preferably, the system includes a reversible locking element configured to lock the firing arm to the limit stop element.

[0019] Embodiments will now be described, by way of example and not limitation, with reference to the accompanying drawings, in which: [Brief description of the drawings]

[0020] [Figure 1] FIG. 1 is an overall perspective view of a release system of the present invention according to one embodiment. [Diagram 2] 3A-3D illustrate steps of movement of a satellite subjected to the action of an ejection system according to the invention; [Diagram 3] 1 shows a launch vehicle carrying two satellites, each of which is applied with a release system according to the invention. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

[0021] In FIG. 1, 1 indicates an overall system according to a preferred embodiment of the present invention.

[0022] The system 1 comprises a torsion bar 2 fixed to a launch vehicle 3. In particular, the torsion bar 2 has a first end 2a fixed to the launch vehicle 3 by a support 4 and a second end 2b fixed to the launch vehicle 3 by a hinge 5. The first end 2a is locked in rotation, i.e. it cannot rotate around the longitudinal axis X of the torsion bar during the load carried out by the rotation of the launch arm, as will be explained below. Conversely, the second end 2b is free to rotate around the X axis by means of the hinge 5. The torsion resulting from the fact that the first end 2a is locked and the second end 2b is free to rotate, as will be shown below, obtains the elastic load required to release the satellite.

[0023] The system 1 includes a launch arm 6, which is a member that transmits the thrust generated by the torsion of the bar to the satellite to be launched. The launch arm 6 is composed of a torsion lever 7 and a guide 8 connected to the torsion lever 7 by a locking joint 9. In particular, the torsion lever 7 has a first end 7a integrally connected to the torsion bar 2 and a second end 7b engaged by the locking joint 9, and the guide 8 has a first end 8a engaged by the locking joint 9 and a second free end 8b.

[0024] The system 1 comprises a slider 10 which, in use, is fixed to a satellite to be launched and which is arranged to slide over the guide 8 and then exit from its second end 8b.

[0025] The presence of the locking joint 9 makes it possible to modify the angle between the torsion lever 7 and the guide 8 and therefore to select the ejection direction transmitted to the satellite.

[0026] The guide 8 comprises an energy absorber 11 arranged near its first end 8a.

[0027] The system 1 comprises a limit stop element 12 arranged to lock the rotation of the torsion lever 7 about the axis X. In particular, the limit stop element 12 comprises a striking surface 13 against which a part of the torsion lever 7 strikes.

[0028] The system 1 comprises a locking element 14 which ensures a reversible locking of the torsion lever 7 to the limit stop element 12 when a part of the torsion lever 7 strikes the striking surface 13. According to a preferred embodiment, the locking element 14 comprises a slot 15 obtained in a part of the torsion lever 7 and a pin 16 extending from the striking surface 13. A reversible male-female lock is thereby achieved. The torsion lever 7 is locked to the striking surface 13 when it has finished its stroke due to the presence of the limit stop element 12. Obviously, it is also possible to obtain the locking element 14, different from the one disclosed, by making a slot in the striking surface 13 and arranging a pin in a part of the torsion lever 7.

[0029] Finally, the system 1 includes a torsional preload element, shown diagrammatically and designated 17 .

[0030] The torsion preload element 17 acts on the first end 2a of the torsion bar 2 and sets the base torsion level of the bar. In effect, the torsion preload element 17 rotates the first end 2a of the torsion bar 2 and subsequently locks it in position. This makes it possible to vary the elastic force of the torsion bar depending on the mass of the satellite to be launched.

[0031] While most of the system is made from aluminium, the torsion bar 2 is preferably made from steel, as are the other components subjected to high mechanical stresses. Preferably, the slider 10 is made from a polymer material to ensure a low level of friction between the slider 10 and the guide 8.

[0032] In use, after the locking joint 9 has been set, the launch arm 6 is rotated and twisted (loaded) the torsion bar 2 until its position is locked by the temporary locking means, which are removed before the satellite is launched and for simplicity are not shown or described. The satellite is then loaded by engaging the slider 10 attached to it with the guide 8 of the launch arm 6.

[0033] When the launch vehicle carrying the satellite reaches the separation condition, the launch vehicle can be commanded to release the satellite's locking system (e.g., using a pyrotechnic belt tensioner system or an explosive bolt), resulting in the launch arm 6 being free to rotate.

[0034] As shown in Figure 2, when the satellite is unlocked from the launch vehicle, the torsion bar 2 forces the launch arm 6 and the satellite 18 to rotate around the X-axis until the torsion lever 7 strikes the limit stop element 12. During the thrust step, a constraint between the guide and the slider is obtained so that no orbit other than the desired circular one is allowed for the satellite. The movement of the satellite 18 is converted from a rotational motion to a pure translational motion by the slider 10, which slides along the guide 8 and exits from the second end 8b. For this purpose, the guide and the slider (prismatic constraint) are sized in such a way that they provide torque pulses to the satellite on the three coordinate axes so that at the exit of the guide all possible rotations are cancelled and only the required pure translation is obtained.

[0035] It should be noted that the release system of the present invention does not release satellites in a direction perpendicular to the separation plane.

[0036] This would even allow a launch vehicle to carry multiple satellites very close to each other.

[0037] As can be seen in Figure 3, the launch vehicle can accommodate two satellites and bring them close to each other during release without the risk of a collision.

[0038] In effect, the release system of the present invention is such that the two satellites are launched from opposite sides of the launch vehicle.

[0039] From the foregoing it can be derived that the inventive ejection system, which includes a single thrust point on the satellite combined with the stabilizing effect of the prism guide, does not generate the rotational motion typical of ejection systems of the known art.

[0040] The absence of rotational movements and the high ejection speeds generated by the torsion bars provide stabilization on the three satellite axes, allowing the subsequent deployment of the solar panels to occur faster than with prior art systems. This effect inevitably leads to important advantages related to the reduction of energy consumption from the satellite batteries, thus providing reserve energy to support a possible early emergency situation.

[0041] Furthermore, the presence of the torsional preload element 17 makes it possible to vary the thrust force available from the torsion bar according to the mass of the satellite to be launched and the desired release velocity, without modifying any of the system's components.

[0042] Unlike the one disclosed above, the ejection system of the invention may comprise two launch arms instead of one. This variant is necessary if the mass of the satellite is very large. However, the two launch arms are connected to the same torsion bar, thus preventing even the slightest asynchronous ejection. This solution can be implemented by a connection element between the first and second launch arms, the ends of which are connected to the torsion bar. Such a connection element can consist of a connection tube arranged outside the torsion bar. [Explanation of symbols]

[0043] 1 System 2 Torsion bar 2a First end 2b Second end 3 Launch vehicle 4 Support 5 Hinge 6 Launch Arm 7 Torsion Lever 7a First end 7b Second end 8 Guide 8a First end 8b Second end 9. Locking Joint 10 Slider 11 Energy absorber 12. Restriction suspension elements 13 Striking surface 14 Rock Elements 15 Slots 16-pin 17 Torsional preload element 18 Satellite

Claims

1. A deployment system (1) for deploying an artificial satellite from a launch vehicle, the deployment system (1) comprising: (i) a torsion bar (2) having a first end (2a) fixed to the launch vehicle (3) by support means (4) and locked against rotation about a longitudinal axis (X) of the torsion bar (2), and a second end (2b) connected to said launch vehicle (3) by hinge means (5) and free to rotate about said longitudinal axis (X); (ii) at least one firing arm (6) extending perpendicularly from said torsion bar (2) and comprising: (a) a torsion lever (7) having a first end (7a) integrally fixed to said torsion bar (2); and (b) a guide (8) having a first end (8a) connected to a second end (7b) of said torsion lever (7), and a second free end (8b); (iii) at least one slider (10) fixed integrally to the satellite (18) to be launched and arranged to slidably engage said guide (8); and (iv) a limit stop element (12) designed to act on the torsion lever (7) to interrupt the rotation of the firing arm (6) about the longitudinal axis (X); A deployment system (1) for deploying an artificial satellite from a launch vehicle, comprising:

2. 2. A release system (1) for releasing a satellite from a launch vehicle as claimed in claim 1, characterized in that the first end (8a) of the guide (8) is connected to the second end (7b) of the torsion lever (7) by a locking joint (9), allowing a change of the angle between the guide (8) and the torsion lever (7).

3. A release system (1) for releasing a satellite from a launch vehicle according to claim 2, characterized in that the guide (8) comprises an energy absorber (11).

4. A release system (1) for releasing a satellite from a launch vehicle according to any one of claims 1 to 3, characterized in that the release system comprises torsional preload means (17) acting on the first end (2a) of the torsion bar (2).

5. A release system (1) for releasing a satellite from a launch vehicle as claimed in any one of claims 1 to 4, characterized in that the limit stop element (12) comprises a striking surface (13) against which a part of each launch arm (7) impacts to end its stroke.

6. A release system (1) for releasing a satellite from a launch vehicle according to any one of claims 1 to 5, characterized in that it comprises a reversible locking element designed to lock the launch arm to the limit stop element.

Images