Steerable antenna for spacecraft, and process of deploying the steerable antenna

Abstract

Steerable antenna for spacecraft, in particular satellites, comprising a main reflector (2) configured to reflect a beam of radio frequency signals in predefined transmission directions, a sub-reflector (3), a signal feeder (4) configured to transmit the RF signal beam to the main reflector (2) via the sub-reflector (3), a first actuator (5) configured to rotate the main reflector (2) and the sub-reflector (3) about a first axis (A), and a second actuator (6) configured to rotate the main reflector (2) about a second axis (B). The first axis (A) is aligned with the signal feed (2) and the second axis (B) coincides with a line from the centre of the main reflector (2) to the focal point (FP) of the main reflector (2). Process of deploying the steerable antenna.

Description

TECHNICAL FIELD

[0001] The present invention relates to a steerable antenna for spacecraft, in particular for satellites that are part of a constellation providing global 5G coverage worldwide, and to a process for deploying the steerable antenna. Steerable antennas are also called steerable or tracking antennas. They are antennas adapted to transmit and / or receive an electromagnetic signal beam over a wide sweep.PRIOR ART

[0002] Steerable antennas in particular for the aerospace sector are well known, the aim being to obtain a wide sweep without defocusing, i.e. with minimal degradation in signal quality over the sweep.

[0003] Steerable antennas comprise a parabolically curved main reflector, a sub-reflector, a feed and a positioning mechanism configured to move the antenna and position it to receive and / or transmit a signal beam from the feed to the ground by the action of the reflector and sub-reflector. European application EP1213788A2 discloses a steerable antenna comprising a feed positioned at a first fixed preset position, a main reflector, a sub-reflector positioned at a second fixed preset position, said sub-reflector being stationary with respect to the feeder. The feeder, the sub-reflector and the main reflector together provide a beam, the main reflector directing the beam in a preselected direction. The main reflector is positioned via a gimbal, so that it can scan the beam over a preselected sweep.

[0004] Generally, steerable antennas known in the prior art include two axes of rotation with rotary joints to propagate the signal along each of the axes of rotation. A problem associated with wide-sweep steerable antennas is that they are sometimes subject to a singularity that affects the ability to track a target when the beam is aligned with one of the rotational axes. This effect or singularity is known as "key hole" and usually occurs when one of the rotation axes points to Nadir. To solve this problem, some solutions propose the use of a third axis that provides the antenna with an additional degree of freedom that allows it to avoid the "key hole".

[0005] European applications EP2996197A1 and EP2584650A1 disclose steerable antennas without the "key hole" effect. To this end, in EP2996197A1, the steerable antenna comprises a support structure on which the feeder is mounted, a main reflector and a sub-reflector both mounted on the support structure with the possibility of relative movement between the main reflector and the sub-reflector. The antenna also comprises actuators configured so that the main reflector rotates around two axes of rotation and the sub-reflector rotates around only one of the two axes of rotation. The actuators are fixedly mounted on the support structure.

[0006] European application EP2584650A1 discloses a steerable antenna where the sub-reflector and the main reflector define a focal point arranged adjacent to a reflecting surface of the main reflector. The feeder is arranged adjacent to the focal point defining an axis pointing to the intersection point of the sub-reflector with the main reflector generating a signal defining an angle with the feed axis. The steerable antenna comprises two actuators, one for rotating the feeder, the sub-reflector and the main reflector about a first axis of rotation perpendicular to the feeder axis and not intersecting with the predetermined coverage area of the antenna, and another for rotating the main reflector or the assembly formed by the main reflector and the sub-reflector with respect to the feeder about a second axis of rotation aligned with the feeder axis, the second actuator being rotated by the first actuator.DISCLOSURE OF THE INVENTION

[0007] The object of the invention is to provide a steerable antenna for space vehicles, in particular for satellites, and a deployment process of the steerable antenna as defined in the claims.

[0008] A first aspect of the invention relates to a steerable antenna comprising a main reflector configured to reflect a beam of radio frequency signals in predefined transmission directions, a sub-reflector, a signal feeder configured to transmit the beam of radio frequency signals to the main reflector via the sub-reflector, a first actuator configured to rotate the main reflector and the sub-reflector with respect to a first axis, and a second actuator configured to rotate the main reflector with respect to a second axis. The first axis is arranged aligned with the signal feed and the second axis is arranged aligned with a line running from the centre of the main reflector to the focal point of the main reflector.

[0009] A second aspect of the invention relates to a steerable antenna deployment process comprising the following steps: release of the boom and main reflector from the platform; tilting of the arm around the first axis driven by the first actuator, and rotation of the main reflector relative to the arm driven by the spring element until a locked position of the main reflector relative to the arm is reached; tilting of the arm to a nominal deployment position.

[0010] The result is a steerable antenna with a compact and robust design that enables a wide sweep without defocusing and achieves optimum radiation performance. It enables a wide signal direction range of up to + / - 65° without key-holes over the entire signal direction range.

[0011] These and other advantages and features of the invention will become apparent in view of the figures and the detailed description of the invention.DESCRIPTION OF THE DRAWINGS

[0012] Figure 1 shows a perspective view of a part of a spacecraft with a steerable antenna according to the invention in a nominal deployed position. Figure 2 shows a detail view of the steerable antenna shown in Figure 1. Figure 3 shows a detail view of the steerable antenna shown in Figure 1 in an anchored position. Figure 4(a)-(f) sequentially shows some stages of the deployment process of the steerable antenna shown in figure 1. Figure 5 shows a detail of a deployment system of the steerable antenna shown in figure 1, in a position prior to locking. Figure 6 shows a detail of a deployment system of the steerable antenna shown in figure 1, in a locked position. DETAILED DISCLOSURE OF THE INVENTION

[0013] Figure 1 shows a part of a spacecraft, in particular of a satellite 40 comprising a plurality of steerable antennas 1 according to the invention (in figure 1 only one steerable antenna 1 is shown).

[0014] During the launch of the satellite 40, each steerable antenna 1 is anchored to the satellite 40 by means of a respective retention and release system 20. Once the satellite 40 arrives at the desired location in space, the retention and release system 20 releases each steerable antenna 1 so that each steerable antenna 1 is deployed from an anchor position (shown in Figure 3) to a position referred to as the nominal deployed position (shown in Figures 1 and 2) from which each steerable antenna 1 can be steered to emit or receive a beam of electromagnetic signals, in particular radio frequency signals (hereinafter RF signals) in predefined directions towards the Earth or towards another satellite or from the Earth or from another satellite. Preferably, the steerable antennas 1 according to the invention are configured to provide a global 5G signal coverage area around the Earth.

[0015] Each steerable antenna 1 according to the invention comprises a main reflector 2, a sub-reflector 3, and a signal feed 4 configured to transmit the RF signal beam towards the sub-reflector 3, the sub-reflector 3 being configured to reflect the RF signal beam from the signal feed 4 towards the main reflector 2, and the main reflector 2 being configured to reflect the RF signal beam in the predefined transmission directions towards the Earth or towards another satellite. The main reflector 2 can also reflect the beam of RF signals received from the Earth or another satellite towards the sub-reflector 3, the sub-reflector 3 being also configured to reflect the beam of RF signals towards the signal feeder 4.

[0016] Feeder 4 is fixed to a platform 30 of spacecraft 40, with no possibility of relative displacement with respect to platform 30. This avoids the need for rotating couplings. For this purpose, the feeder 4 comprises a bracket 9 which is fixed to the platform 30 of the spacecraft 40. The feeder 4 is separate from the main reflector 2 and the sub-reflector 3, not being connected or coupled to them. The signal feeder 4 comprises a signal source 15 connected to a feed horn 14 and the signal feeder 4 can operate simultaneously in dual band Tx and Rx with dual RHCP and LHCP polarisation. The feed horn 14 is supported in the bracket 9 of the feeder 4.

[0017] Each steerable antenna 1 further comprises a first actuator or azimuth actuator 5 configured to rotate the main reflector 2 and the sub-reflector 3 with respect to a first axis or azimuth axis A, and a second actuator or elevation actuator 6 configured to rotate the main reflector 2 with respect to a second axis or elevation axis B. Both actuators 5 and 6 are coupled to each other via a movable bracket 7 which rotates when actuated by the first actuator 5. Both actuators 5 and 6 are coupled to each other via a movable support 7 which rotates when actuated by the first actuator 5. The first actuator 5 has a fixed part 5a attached to the spacecraft 40 via a fixed support 8, and a rotating part 5b attached to the movable support 7. The second actuator 6 is coupled to the fixed bracket 8 via the movable bracket 7. The main reflector 2 is coupled to the second actuator 6 via an arm 10, the arm and the main reflector 2 being hinged with respect to the first actuator 5 between the nominal deployed position (shown in Figures 1 and 2) and the anchored position (shown in Figure 3). The second actuator 6 has a fixed part 6a mounted on the movable bracket 7 and a rotatable part 6b coupled to the main reflector 2 via an arm 10.

[0018] The sub-reflector 3 is fixed to the movable support 7 and rotates with respect to the first axis A, together with the main reflector 2, when actuated by the first actuator 5. The second actuator 6 can only actuate the arm 10 by rotating it around the second axis B but does not rotate the movable support 7 and thus the sub-reflector 3. In the embodiment shown in the figures, the movable support 7 is an L-shaped support. One end of the movable support 7 is coupled to the rotating part 5b of the first actuator 5, and the other end of the movable support 7 supports the second actuator 6.

[0019] The first actuator 5 and the second actuator 6 each comprise respectively a motor on the respective fixed parts 5a and 6a, and a gearbox on the respective movable parts 5b and 6b, through which the first actuator 5 is coupled to the movable support 7 and the second actuator 6 is coupled to the arm 10. Each respective motor and gearbox are of coaxial shafts.

[0020] The first actuator 5 has an axis of rotation coinciding with the first axis A and the second actuator 6 has an axis of rotation coinciding with the second axis B.

[0021] In particular, the first axis A coincides with an axis of a feed horn 14 of the feed 4. The second axis B passes through a geometrical centre CG of the main reflector 2 and through a focal point PF of the main reflector 2. In the embodiment shown in the figures, the first axis A forms an angle with the second axis B of 90° as shown in figures 1 and 2.

[0022] In the embodiment shown in the figures, the main reflector 2 is parabolic, and the sub-reflector 3 is substantially flat.

[0023] As indicated above, in the anchored position, each steerable antenna 1 is anchored to the platform 30 of the spacecraft 40 via the respective retention and release system 20. The hold-down and release system 20 comprises hold-down and release mechanisms 21 (also known as HDRMs), and passive couplings 22. Each hold-down and release mechanism 21 comprises a fixed part 21a which is arranged fixed to the satellite platform 30 of the satellite 40 and includes non-explosive release means, and a movable part 21b which is arranged fixed to the antenna 1 and comprises a screw attached to the non-explosive release means in the anchor position. The retention and release mechanisms 21 are known in the prior art and a detailed description of them is not considered necessary. On the other hand, each passive coupling 22 comprises a fixed part 22a attached to the satellite platform 30 of the satellite 40, and / or a movable part 22b attached to the steerable antenna 1, each fixed part 22a being coupled to the movable part 22b so as to form a spherical coupling or each fixed part 22a or each movable part 22b including a spherical coupling 22c at one end.

[0024] The main reflector 2 is pivotably mounted with respect to the arm 10 via a coupling 12 so that it can pivot with respect to the arm 10 until it is substantially horizontal in the anchoring position. Furthermore, it is pivotable with respect to the first actuator 5. The coupling 12 is a hinge-type coupling. In the anchoring position, the arm 10 is hinged, and anchored to the platform 30 of the spacecraft 40 via the retention and release system 20. This results in a compact and optimized steerable antenna 1, which can be hinged to facilitate anchoring of the main reflector and the sub-reflector during launch of the spacecraft.

[0025] In the embodiment shown in the figures, the main reflector 2 is anchored to the platform 30 through the corresponding retention and release mechanism 21 and corresponding passive couplings 22, the retention and release mechanism 21 being arranged between two passive couplings 22. On the other hand, the arm 10 is anchored to the platform 30 through the corresponding retention and release mechanism 21 arranged approximately in the centre of the arm 10 and passive couplings 22 arranged approximately at each end of the arm 10. In the anchored position, the bolt of the retention and release mechanisms 21 is pretensioned by the non-explosive release means, generating controlled compressive forces at each spherical coupling 22c which are sufficient to keep the arm 10 and the main reflector 2 anchored to the platform 30 during launch of the spacecraft 40. When the non-explosive release means are actuated, the preload in the passive couplings 22 and in the retention and release mechanisms 21 disappears, so that the arm 10 and the main reflector 2 are released from their attachment to the platform 30.

[0026] The steerable antenna 1 further comprises deployment means 23 configured to deploy the main reflector 2 to the nominal deployed position once the retention and release systems 20 have released the main reflector 2 and the arm 10 from their attachment to the platform 30. The deployment means 23 comprise at least one elastic element 24 in the coupling 12 of the arm 10 to the main reflector 2, and a spherical support 25 attached to the main reflector 2, the elastic element 24 and the spherical support 25 collaborating together with the first actuator 5 to deploy the arm 10 and the reflector 2 from the anchored position to the nominal deployed position. The elastic element 24 is configured to rotate the main reflector 2 with respect to the arm 10, from the anchoring position (shown in figure 5) until a stop surface 12b of the coupling 12 prevents further release (locking position shown in figure 6), the elastic element 24 being preloaded so that thereafter the main reflector 2 does not rotate with respect to the arm 10. Preferably, the elastic element 24 is a torsion spring. The preload of the spring element 24 is lower than the torque of the first actuator 5, thus avoiding the need to use dampers to control the rotational speed of the main reflector 12 relative to the arm 10.

[0027] In the embodiment shown in the figures, an end 24a of the elastic element 24 is coupled to a bracket 29 such that the bracket 29 rotates solidly with the end 24a of the elastic element 24, the bracket 29 including a projection 29b configured to abut against the abutting surface 12b of the coupling 12 in the locked position and prevent the reflector 2 from rotating relative to the arm 10.

[0028] In the embodiment shown in the figures, the coupling 12 of the arm 10 to the main reflector 2 and the elastic element 24 are arranged in a perimeter zone of the main reflector 2, the spherical support 25 being arranged in the perimeter zone of the main reflector 2, preferably at 90° to the coupling 12. The deployment means 23 further comprise a sliding surface 26 attached to the platform 30 of the spacecraft 40, configured so that the spherical support 25 slides on the sliding surface 26 during deployment to avoid tribological problems in the contact between the spherical support 25 and the sliding surface 26.

[0029] Figures 4(a) to 4(f) sequentially show the deployment process of the steerable antenna 1 comprising the following stages: releasing the arm 10 and the main reflector 2 from the platform 30; tilting the arm 10 about the first axis A driven by the first actuator 5, and rotating the main reflector 2 relative to the arm 10 until a locked position of the main reflector 2 relative to the arm 10 is reached; and tilting of the arm 10 until the nominal unfolded position is reached.

[0030] When the main reflector 2 rotates with respect to the arm 10 driven by the elastic element 24, the rotation is initially controlled by one of the passive couplings 22 attached to the main reflector 2 in contact with the sliding surface 26 through the spherical coupling 22c (position shown in Figure 4(b)). Before said passive coupling 22 stops contacting the sliding surface 26, the spherical support 25 of the deployment means 23 comes into contact with the sliding surface 26, said spherical support 25 then controlling the rotation of the main reflector 2 with respect to the arm 10 (position shown in figure 4(c)).

[0031] Finally, the spherical support 25 loses contact with the sliding surface 26 before the elastic element 24 is locked in the locked position (position shown in figure 4(e)). From the locking position of the main reflector 2 relative to the arm 10, the elastic element 24 remains preloaded and the position of the main reflector 2 relative to the arm 10 is maintained during subsequent tilting of the arm 10.

[0032] To move from the anchored position to the nominal deployed position, the first actuator 5 rotates the movable support 7 with respect to the first axis A, and with it, the arm 10 by an angle a of approximately 90°, and the main reflector 2 rotates by an angle b of approximately 40° with respect to the arm 10 driven by the elastic element 24. The second actuator 6 rotates the main reflector 2 with respect to the second axis B by an angle of approximately 15° until it is positioned in the nominal deployed position. Starting from the nominal deployed position, the first actuator 5 can rotate the main reflector 2 and the sub-reflector 3 in a range of approximately + / - 65° and the second actuator 6 can rotate the main reflector 2 in a range of approximately + / - 67°. This results in a steerable antenna 1 with a boresight axis capable of covering 360° of a semi-angle cone of 63°.

[0033] The described features and / or aspects of the steerable antenna are also valid and applicable to the process for deploying the steerable antenna, and the described features and / or aspects of the process are also valid and applicable to the steerable antenna.

Claims

1. Steerable antenna for spacecraft, in particular satellites, comprising a main reflector (2), a sub-reflector (3), a signal feeder (4) adapted for attachment to the spacecraft (40), the signal feeder (4) being configured to transmit or receive a beam of radio frequency signals to or from the main reflector (2) through the sub-reflector (3), the main reflector (2) being configured to reflect the beam of radio frequency signals in a predefined direction, a first actuator (5) configured to rotate the main reflector (2) and the sub-reflector (3) about a first axis (A), and a second actuator (6) configured to rotate the main reflector (2) about a second axis (B), characterised in that the second axis (B) passes through a geometric centre (CG) of the main reflector (2) and a focal point (FP) of the main reflector (2).

2. Steerable antenna according to the preceding claim, wherein the first axis (A) is aligned with an axis of a feed horn (14) of the feeder (4) and intersects with the second axis (B) in the sub-reflector (3).

3. A steerable antenna according to any one of the preceding claims, wherein the first axis (A) coincides with the axis of rotation of the first actuator (5), and the second axis (B) coincides with the axis of rotation of the second actuator (6).

4. A steerable antenna according to any one of the preceding claims, wherein the main reflector (2) is coupled to the second actuator (6) via an arm (10), wherein said arm (10) and the main reflector (2) can be swung with respect to the first actuator (5) between a nominal deployed position and an anchored position.

5. A steerable antenna according to the preceding claim, comprising a retention and release system (20) configured to anchor in the anchoring position the main reflector (2) and / or the arm (10) folded down to the spacecraft (40).

6. A steerable antenna according to any one of claims 4 and 5, wherein the first actuator (5) comprises a fixed part (5a) adapted to be attached to the spacecraft (40), and a rotating part (5b) attached to a movable support (7), and the second actuator (6) has a fixed part (6a) mounted on the movable support (7) and a rotating part (6b) coupled to the arm (10), the movable support (7) being configured to rotate about the first axis (A) when actuated by the first actuator (5).

7. Steerable antenna according to the preceding claim, wherein the sub-reflector (3) is fixed to the mobile support (7).

8. Steerable antenna according to any one of claims 4 to 6, wherein the main reflector (2) is pivotably coupled to the arm (10) via a coupling (12) allowing horizontal positioning of the main reflector (2) in the anchoring position.

9. Steerable antenna according to the preceding claim, comprising deployment means (23) configured to deploy the main reflector (2) and the arm (10) from the anchored position to the nominal deployed position, the deployment means (23) comprising at least one elastic element (24) in the coupling (12) between the main reflector (2) and the arm (10), and a spherical support (25) on the main reflector (2), the elastic element (24), the spherical support (25) and the first actuator (5) cooperating with each other to deploy the main reflector (2) to the nominal deployed position in a controlled manner.

10. Steerable antenna according to the preceding claim, wherein the spring element (24) is a torsion spring.

11. A steerable antenna according to any one of claims 9 and 10, wherein the deployment means (23) comprises a sliding surface (26) that cooperates with the spherical support (25) to control the deployment of the main reflector (2) with respect to the arm (10).

12. A steerable antenna according to any one of the claims, wherein each actuator (5,6) comprises a motor and a gear unit coupled to the respective motor, each motor and gear unit having coaxial shafts.

13. A steerable antenna according to any one of the preceding claims, wherein the main reflector (2) is parabolic and the sub-reflector (3) is flat.

14. A spacecraft comprising at least one steerable antenna (1) according to any one of the preceding claims, the signal feeder (4) of each steerable antenna (1) being attached to a platform (30) of the spacecraft (40).

15. The process of deploying the steerable antenna (1) according to any one of the preceding claims, comprising the following steps: o releasing the arm (10) and the main reflector (2) from the platform (30); o tilting of the arm (10) about the first axis (A) driven by the first actuator (5), and rotation of the main reflector (2) relative to the arm (10) until a locked position of the main reflector (2) relative to the arm (10) is reached; and o tilting the arm (10) until the nominal deployment position is reached.

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