System for releasing a satellite from a launch vehicle

Abstract

A release system (1) for releasing a satellite from a launch vehicle, comprising: (i) a torsion bar (2) having a first end (2a) fixed to the launch vehicle (3) by bearing means (4) and locked in rotation about a longitudinal axis (X) of the torsion bar (2), and a second end (2b) connected to the launch vehicle (3) by hinging means (5) and free to rotate about the longitudinal axis (X); (ii) at least one launching arm (6) extending perpendicularly from the torsion bar (2) and comprising: (a) a torsion rod (7) having a first end (7a) fixed in one piece to the torsion bar (2); and (b) a guide (8) having a first end (8a) connected to a second end (7b) of the torsion rod (7), and a second free end (8b); (iii) at least one sliding element (10) fixed in one piece to the satellite (18) to be launched and arranged to engage the guide (8) in a sliding manner; and (iv) a limit stop element (12) designed to act on the torsion rod (7) to interrupt the rotation of the launching arm (6) about the longitudinal axis (X).

Description

[0001] CROSS-REFERENCE TO RELATED APPLICATIONS

[0002] This patent application claims the benefit of Italian Application No. 102020000012415, filed on May 26, 2020, the disclosure of which is incorporated herein by reference in its entirety. Technical Field

[0003] The present invention relates to a system for releasing a satellite from a launch vehicle. Background Art

[0004] Currently, a widely used release system is provided with a plurality of coil springs arranged in an integral manner with the launch vehicle and adapted to generate thrust to eject the satellite. Such a release system involves launching the satellite in a direction perpendicular to the separation plane.

[0005] It will be obvious to those skilled in the art that the coil spring must be selected according to the mass of the satellite to be launched, which necessarily implies the possibility of having to replace the coil spring whenever the mass of the satellite to be launched changes.

[0006] However, despite the widespread use of this technical solution, it involves a number of problems.

[0007] In fact, once the satellite is launched by the action of the coil spring, it undergoes a rotational motion.

[0008] This rotational movement is primarily due to the fact that the springs are not identical to each other and, at the same time, cannot extend simultaneously relative to each other.

[0009] The latter aspect stems from the impossibility of activating the various release elements in a precisely synchronized manner. As those skilled in the art will readily appreciate, even a few microseconds (µs) difference between the actuation of the various release elements can lead to significant imbalances in the total thrust experienced by the satellite, thus causing it to spin. Other factors contributing to the uncertainty of the satellite during the release step and its subsequent rotation are the non-ideal knowledge of the position of its center of mass and the transient vibrations of the launch vehicle structure.

[0010] The rotation of the satellite after release necessarily delays the deployment of the solar panels and thus prolongs the use of energy from the satellite's internal batteries. This aspect represents a considerable risk factor, since energy from the satellite's batteries must always be available during the steps immediately after release and before the deployment of the solar panels in order to perform "emergency" operations.

[0011] In addition, in the case of a launch vehicle carrying a plurality of satellites arranged side by side, the release of the satellites along a direction perpendicular to the separation plane can entail the risk of collisions between the satellites. In fact, it is known that for release systems using the thrust of a helical spring, it is appropriate to consider a "uncertainty cone" of about 15° with respect to the intended release direction. This situation implies a lower occupation of the space available in the launch vehicle nose cone and limits the number of satellites to be installed on the same launch vehicle.

[0012] In addition to the above problems, there is a need in the art for a system capable of guaranteeing a release speed higher than that generated by current systems, using a limited number of helical springs as thrust means. In fact, a higher release rate would result in a faster movement of the satellite away from the launch vehicle, and therefore, it would be possible to deploy the solar panels more quickly. This would thus reduce the use of battery power, with the related benefits as mentioned above.

[0013] For safety reasons, the deployment of the solar panels must take place at a distance of a few hundred meters from the launch vehicle, and if this distance is not reached by the thrust of the release, it is necessary to rely on the battery energy.

[0014] Finally, another need in the art relates to the possibility of having a release system that can be adapted to satellites having different masses, without having to replace part of the components. In other words, there is a need for a release system whose thrust means is effective regardless of the mass of the satellite. As mentioned above, in fact, the solutions that provide for the use of helical springs must require the selection of the solution according to the mass of the satellite to be launched. SUMMARY

[0015] The object of the present invention is a system for releasing a satellite from a launch vehicle; said release system is characterized in that it comprises: (i) a torsion bar having a first end and a second end, said first end being fixed to the launch vehicle by bearing means and being locked in rotation about a longitudinal axis of said torsion bar, said second end being connected to said launch vehicle by hinging means and being free to rotate about said longitudinal axis; (ii) at least one launch arm extending perpendicularly from said torsion bar and comprising (a) a torsion bar having a first end fixed in a unitary manner to said torsion bar, and (b) a guide having a first end connected to a second end of said torsion bar, and a second free end; (iii) at least one sliding element fixed in a unitary manner to the satellite to be launched and arranged to engage said guide in a sliding manner; and (iv) a limit stop element designed to act on said torsion bar to interrupt the rotation of the launch arm about the longitudinal axis.

[0016] Preferably, said first end of said guide is connected to said second end of said torsion bar by means of a locking joint. Thus, by means of the locking joint, the release direction can be changed according to the angle between the guide and the torsion bar.

[0017] Preferably, said guide comprises an energy absorber.

[0018] Preferably, the system comprises torsion pre-loading means acting on said first end of the torsion bar. These torsion pre-loading means set the torsion of the bar, i.e. already upstream of the loading torsion due to the rotation of the launch arm to which the satellite to be launched is engaged. Thus, the elastic force required for the effective release can be adjusted according to the mass of the satellite to be launched.

[0019] Preferably, the system comprises a reversible locking element arranged to lock the launch arm to the limit stop element. BRIEF DESCRIPTION OF DRAWINGS

[0020] In the following, by means of the attached drawings, for illustrative and non-limiting purposes, an embodiment is described, in which:

[0021] Figure 1 is a general perspective view of the release system of the application according to an embodiment;

[0022] Figure 2 the movement steps of a satellite subjected to the action of the release system according to the application are shown; and

[0023] Figure 3 a launch vehicle carrying two satellites is shown, on each of which the release system according to the application is applied. DETAILED DESCRIPTION

[0024] In Figure 1 , the reference 1 indicates as a whole a system according to a preferred embodiment of the application.

[0025] 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 means of a support 4 and a second end 2b fixed to the launch vehicle 3 by means of a hinge 5. The first end 2a is locked in rotation, i.e. it cannot rotate about the longitudinal axis X of the torsion bar during the loading performed by the rotation of the launch arm, as will be described below. On the contrary, the second end 2b is free to rotate about the longitudinal axis X due to the hinge 5. As will be shown in the following, the torsion due to the fact that the first end 2a is locked and the second end 2b is free to rotate, obtains the elastic load required for the release of the satellite.

[0026] The system 1 comprises a launching arm 6 which is a means of transmitting the thrust generated by the torsion of the torsion bar to the satellite to be launched. The launching arm 6 consists of a torsion bar 7 and a guide 8 connected to the torsion bar 7 by a locking joint 9. In particular, the torsion bar 7 has a first end 7a connected in a unitary manner to the torsion bar 2 and a second end 7b engaged by the locking joint 9, while the guide 8 has a first end 8a engaged by the locking joint 9 and a second free end 8b.

[0027] The system 1 comprises a sliding element 10 which, in use, is fixed to the satellite to be launched and is arranged to slide on the guide 8 and then to exit the guide at the second free end 8b thereof.

[0028] The presence of the locking joint 9 makes it possible to vary the angle between the torsion bar 7 and the guide 8 and thus allows the direction of release to be selected which will be transmitted to the satellite.

[0029] The guide 8 comprises an energy absorber 11 located near the first end 8a thereof.

[0030] The system 1 comprises a limit stop element 12 which is arranged to lock the rotation of the torsion bar 7 about the axis X. In particular, the limit stop element 12 comprises an impact surface 13 on which a portion of the torsion bar 7 impacts.

[0031] The system 1 comprises a locking element 14 which, once a portion of the torsion bar 7 has impacted the impact surface 13, ensures that the torsion bar 7 is reversibly locked on the limit stop element 12. According to a preferred embodiment, the locking element 14 comprises a slot 15 obtained in a portion of the torsion bar 7 and a pin 16 which extends from the impact surface 13. A reversible male-female locking is thus achieved. Once the torsion bar 7 has finished its stroke due to the presence of the limit stop element 12, the torsion bar 7 will be locked on the impact surface 13. Obviously, unlike what has been disclosed, it is also possible to obtain the locking element 14 by forming a slot in the impact surface 13 and arranging a pin on a portion of the torsion bar 7.

[0032] Finally, the system 1 comprises a torsional preloading element, which is shown schematically and is indicated by the reference 17.

[0033] The torsional preloading element 17 acts on the first end 2a of the torsion bar 2 and sets the basic torsional level of the torsion bar. In fact, the torsional preloading element 17 rotates the first end 2a of the torsion bar 2 and then locks its position. Thus, it is possible to vary the elastic force of the torsion bar as a function of the mass of the satellite to be launched.

[0034] Preferably, the torsion bar 2 is made of steel, as are the other components subjected to high mechanical stresses, while the bulk of the system is made of aluminium. Preferably, the slide 10 is made of a polymer material in order to ensure a low level of friction between the slide 10 and the guide 8 .

[0035] In use, after the locking joint 9 has been set, the launch arm 6 is rotated, thereby causing the torsion bar 2 to be twisted (loaded), until its position is locked by a temporary locking device, which is removed before the satellite is launched and, for the sake of simplicity, is not shown or described herein. The satellite is then mounted by engaging the slide 10 attached thereto onto the guide 8 of the launch arm 6.

[0036] Once the launch vehicle with the satellite has reached separation conditions, the launch vehicle can be commanded to release the satellite locking system (for example, using a blasting belt tensioner system or using explosive bolts) and subsequently commanded to rotate freely on the launch arm 6 .

[0037] like Figure 2 As shown, once the satellite's lock to the launch vehicle is released, the torsion bar 2 forces the launch arm 6 and satellite 18 to rotate about axis X until the torsion bar 7 strikes the limit stop 12. During the push step, the constraint between the guide and the slider is achieved in such a way that no orbit other than the desired circular orbit is permitted for the satellite. Thus, the motion of the satellite 18 is converted from rotation to pure translation, while the slider 10 slides along the guide 8 and then leaves the guide at its second free end 8b. To this end, the guide and slider (prismatic constraint) are dimensioned to provide a satellite torque pulse in the three coordinate axes, eliminating all possible rotations and achieving only the desired pure translation at the exit of the guide.

[0038] It should be noted that the release system of the present invention does not release the satellite in a direction perpendicular to the separation plane.

[0039] This makes it possible to mount multiple satellites on a launch vehicle, even in close proximity to one another.

[0040] from Figure 3 It can be seen that the launch vehicle can accommodate two satellites and set them close to each other, without the risk of collision during the release process.

[0041] In effect, the release system of the present invention enables two satellites to be launched from opposite sides of a launch vehicle.

[0042] As can be seen from the above, the release system of the present invention includes a single thrust point on the satellite, which, combined with the stabilizing effect of the prism guide, does not produce rotational motion, which is a typical feature of release systems in the prior art.

[0043] The absence of rotational movements and the high release rate generated by the torsion bar enable stabilization on the three satellite axes and, subsequently, faster deployment of the solar panels, compared to the systems of the prior art. This effect inevitably entails important advantages related to lower energy consumption from the satellite batteries, thereby generating reserve energy to support possible initial emergencies.

[0044] In addition, the presence of the torsion preloading element 17 makes it possible to vary the thrust generated from the torsion bar as a function of the mass of the satellite to be launched and the desired release rate, without the need to change the components of the system.

[0045] Unlike what is disclosed above, the release system of the present application can comprise two launch arms instead of one. This change is necessary in the case of satellites of significant mass. However, the two launch arms are connected to the same torsion bar in such a way as to prevent even the slightest asynchronous release. This solution can be implemented by means of a connecting element between the first launch arm and the second launch arm, which are connected at their ends to the torsion bar. This connecting element can comprise a connecting tube arranged outside the torsion bar.

Claims

1. A release system (1) for releasing a satellite from a launch vehicle, characterized in that include: (i) a torsion bar (2) having a first end (2a) and a second end (2b), the first end (2a) being fixed to the launch vehicle (3) by means of support means (4) and being locked in rotation about the longitudinal axis (X) of the torsion bar (2), and the second end (2b) being connected to the launch vehicle (3) by means of articulation means (5) and being free to rotate about the longitudinal axis (X); (ii) at least one launching arm (6) extending perpendicularly from the torsion bar (2) and comprising: (a) a torsion bar (7) having a first end (7a) fixed to the torsion bar (2) in an integral manner; and (b) a guide member (8) having a first end (8a) connected to the second end (7b) of the torsion bar (7), and a second free end (8b); (iii) at least one slider (10) fixed in an integral manner to the satellite (18) to be launched and arranged to engage the guide (8) in a sliding manner; and (iv) a limit stop element (12) designed to act on the torsion bar (7) to interrupt the rotation of the firing arm (6) about the longitudinal axis (X).

2. The release system for releasing a satellite from a carrier rocket according to claim 1, characterized in that The first end (8a) of the guide (8) is connected to the second end (7b) of the torsion bar (7) by a locking joint (9) so as to be able to change the angle between the guide (8) and the torsion bar (7).

3. The release system for releasing a satellite from a carrier rocket according to claim 2, characterized in that The guide member (8) comprises an energy absorber (11).

4. The release system for releasing a satellite from a carrier rocket according to claim 1, wherein: The release system comprises a torsional preloading device (17) acting on the first end (2a) of the torsion bar (2).

5. The release system for releasing a satellite from a carrier rocket according to claim 1, wherein: The limit stop element (12) comprises an impact surface (13) against which a corresponding portion of the launching arm (6) impacts to end its travel.

6. The release system for releasing a satellite from a carrier rocket according to claim 1, wherein: The release system comprises a reversible locking element designed to lock the firing arm on the limit stop element.

Images