Patent Description:
Nowadays, a spacecraft is usually designed with a limited lifespan of around <NUM>-<NUM> years, depending on the mission. The main factor that limits the lifetime of a spacecraft is usually the lack of fuel to perform the orbital and attitude maneuvers. The mass and the volume of these vehicles are limited by the launcher characteristics, so an increase of the propellant mass to extend the lifespan is not a solution.

This is a problem that face most of the vehicles, at least the ones that do not use electric propulsion. However, this fact is much more limiting when talking about telecommunication satellites, more specifically the ones placed in the geostationary orbit.

The telecommunication satellites that perform in the geostationary orbit have been assigned a specific slot inside this orbit, in order to develop their missions without interfering with other vehicles. This occurs because the geostationary orbit is very strategical for telecommunication satellites, since they can always point to the same sub-satellite point to develop their activities.

Moreover, they have a considerable mass (about some thousands of kilograms) and volume, so the launch cost of these satellites is too high.

As a consequence, in order to increase the economic profitability of these satellites, it is necessary to think in a way to extend the lifespan of these vehicles.

In order to support all these, it has been discussed the use of an external vehicle that would dock to the target satellite (usually, a telecommunication satellite), providing an increase in the lifespan. The type of support given to the satellite could be very wide, being the principal one the required propulsion to perform the orbital and attitude maneuvers, but it could also be the telecommunication or navigation support. The vehicle proposed could be categorized as a Commercial Servicing Vehicle (CSV), although there are many kind of vehicles that could fit on this group.

The first type of CSV that has been considered is the Mission Extension Vehicle (MEV). This vehicle is an independent vehicle designed and fabricated from the beginning, in order to perform the rendezvous and the docking with several target satellites and extend the lifespan of these satellites.

Until <NUM>, all the satellites that were designed had a finite lifespan, mainly determined by the amount of fuel that could be integrated inside the vehicles. However, in October <NUM>, the first Mission Extension Vehicle was launched, in order to perform the docking with the satellites that were in their last phase of their lifespan, giving operational support as propulsion maneuvers or attitude and orbit control, in order to extend their operational lifespan.

This was a very relevant milestone in space missions, as it changed the way the satellites will be designed in the following decades. On one hand, it will be possible to perform longer missions using the same satellites, increasing the return on invested capital (mainly for geostationary satellites). On the other hand, the limitation of the fuel in the satellite design is reduced, as now it is able to perform smaller satellites that will be assisted periodically to perform their missions.

However, the design, fabrication and launch of a MEV also represents a significant cost, apart from the fact that these vehicles should be capable to reach to the geostationary orbit, with the amount of propellant mass that it involves.

Currently there are systems in which a service satellite has an on-orbit docking and locking device to perform the docking and locking of the service satellite and a target satellite.

<CIT> discloses a satellite-to-satellite on-orbit docking and locking device that is used for docking a service satellite with a non-cooperative target satellite, and comprises a guide structure, a locking device, and a positioning pin, wherein the guide structure is mounted on the service satellite, the guide structure comprises a tapered surface for eliminating the deviation in the axial tilt angle of docking posture with the non-cooperative target satellite, the guide structure further comprises a sectioned notch with a bevel edge, and the sectioned notch is used for correcting the deviation in circumferential rotation angle with respect to a docking ring of the non-cooperative target satellite; the locking device is mounted on the service satellite, the locking device comprises a cam wheel, two clamping claws stretch out from the edge of the cam wheel, and a bayonet is formed between the two clamping claws, so as to restrain the inner edge of the docking ring of the non-cooperative target satellite. The positioning pin is arranged on the guide structure and is matched with a positioning pin hole on the docking ring of the non-cooperative target satellite, so as to precisely guide and restrain the relative position between the service satellite and the non-cooperative target satellite. The satellite-to-satellite on-orbit docking and locking device is applicable to all satellites with the docking ring; and the satellite-to-satellite on-orbit docking and locking device is substantially wide in application range, including large satellites, small satellites, and micro satellites.

<CIT> discloses a fully androgynous, reconfigurable closed loop feedback controlled low impact docking system with load sensing electromagnetic capture ring. The docking system of the present invention preferably comprises two docking assemblies, each docking assembly comprising a load sensing ring having an outer face, one of more electromagnets, one or more load cells coupled to said load sensing ring. The docking assembly further comprises a plurality of actuator arms coupled to said load sensing ring and capable of dynamically adjusting the orientation of said load sensing ring and a reconfigurable closed loop control system capable of analyzing signals originating from said plurality of load cells and of outputting real time control for each of the actuators. The docking assembly incorporates an active load sensing system to automatically dynamically adjust the load sensing ring during capture instead of requiring significant force to push and realign the ring.

These devices satisfactorily perform the docking function and are suitable for satellites with an interface ring. However, the service spacecraft on which the docking system is assembled are spacecraft specifically designed and launched to perform the docking with satellites that were in their last phase of their lifespan.

Thus, it is an object of the invention to provide a docking system for spacecraft suitable for using in a spacecraft that has already been used in a previous mission and can be reused for another one in orbit as a continuation task.

The invention provides a docking system for spacecraft, comprising an end ring and a docking assembly, the docking assembly comprising:.

The invention also provides a spacecraft with a docking system that comprises a docking system for spacecraft of the invention on one end, such that one end of the linear actuators is attached to the inner surface of the spacecraft.

This configuration of the docking system for spacecraft allows its integration in the volume of the spacecraft when the docking assembly is in its stowed position.

The use of a satellite adapter as a spacecraft with a docking system has a plurality of advantages. On the one hand, as the design and fabrication cost of the satellite structure will be reduced, the economical saving will be considerable. Moreover, the cost of the launch would disappear, since the satellite adapter would already be in-orbit once the first mission is completed (payload separation). On the other hand, the time needed to develop this mission for the extension of the satellite will be considerably lower, since many steps of the development of the satellite would be skipped.

Other advantages of the invention are listed below:.

Other features and advantages of the present invention will become apparent from the following detailed description of an illustrative embodiment and not limiting its purpose in connection with the accompanying figures.

The docking system <NUM> for spacecraft according to an embodiment of the invention is shown in a stowed position in <FIG> and in a deployed position in <FIG>.

The docking system <NUM> comprises an end ring <NUM> and a docking assembly <NUM>.

The docking assembly <NUM> comprises the following elements:.

The embodiment shown in the figures includes three docking platforms <NUM>.

The support ring <NUM> and the end ring <NUM> are connected by elastic means <NUM> (in <FIG> and <FIG> the elastic means <NUM> are longitudinal springs), such that in a stowed position (see <FIG>) the docking assembly <NUM> is placed inside the end ring <NUM> and in a deployed position (see <FIG>) the docking assembly <NUM> is placed out of the end ring <NUM>.

The longitudinal springs can have one of its ends on a fixed element <NUM> in the end ring <NUM> (see <FIG>).

As for the initial engagement means <NUM>, they can be latch mechanisms, as the ones shown in <FIG>, <FIG> and <FIG>. They can comprise a latch bracket <NUM>, a spring and a latch <NUM> as a movable element.

The docking platforms <NUM> can comprise an end stop and torque springs <NUM> with one pin <NUM> attached to the support ring <NUM> and one pin <NUM>' attached to the main body <NUM> of the docking platform <NUM> (see <FIG> and <FIG>).

As for the docking mechanism <NUM>, it is represented in <FIG> and can comprise:.

The docking system <NUM> can be integrated in a spacecraft <NUM>, such that one end of the linear actuators <NUM> is attached to the inner surface of the spacecraft <NUM>, as seen in <FIG> and <FIG>.

The spacecraft <NUM> can be a satellite adapter in a launcher (see <FIG>), and also a satellite with an interface ring.

The satellite adapter can be equipped with the launcher last stage <NUM> that performs the control and guidance functions.

This docking system <NUM> is especially applicable to a satellite adapter in a launcher. When the launch phase finishes, the last stage of the launcher places the payload (a customer satellite, for example) in the geostationary orbit. <FIG> shows the satellite adapter <NUM>, the last stage <NUM>, the customer satellite <NUM> and the docking system <NUM> for spacecraft before the satellite separation.

<FIG> shows the separation of the customer satellite <NUM> from the satellite adapter <NUM>.

<FIG> shows the deployment of the docking assembly <NUM> after the separation of the customer satellite <NUM> from the satellite adapter <NUM>.

When the separation of the customer satellite <NUM> from the satellite adapter <NUM> and the deployment of the docking assembly <NUM> have finished, the first phase of the mission for life extension of the satellite adapter starts (it is called the rendezvous).

<FIG> shows the approximation of the satellite adapter <NUM> to the target satellite <NUM>. It can be seen that the target satellite <NUM> is completely deployed. The target satellite <NUM> has an interface ring <NUM> that remains in it after the separation when it was launched into orbit, and there is no need for any special adaptations in the interface ring <NUM> for performing the docking.

<FIG> shows the docking between the target satellite <NUM> and the satellite adapter <NUM>.

Once the docking is completed (see <FIG>) the satellite adapter <NUM> (service satellite) could give support to the target satellite <NUM>, for example in orbit and attitude maneuvers, navigation, communication, etc..

At the beginning of the mission, the docking assembly <NUM> is stowed, as it needs to be integrated inside the internal volume of the satellite adapter <NUM> (see <FIG>).

The first phase of the docking is the deployment of the docking assembly <NUM>, shown in <FIG>.

There can be a joint element between the docking platforms <NUM> and the satellite adapter <NUM> (for instance, a stop, not shown in the figures) that prevents the docking platforms <NUM> from unfolding. The joint element can also be an electrovalve.

When the customer satellite <NUM> is at a certain distance away from the satellite adapter <NUM>, the joint between the docking platforms <NUM> and the satellite adapter <NUM> is cut-off, so the docking platforms <NUM> do not have the opposition of the joint and the torque springs <NUM> attached to the docking platforms <NUM> make them rotate.

In <FIG> the configuration of these torque springs <NUM> can be seen. In this <FIG> there are two torque springs <NUM> attached to each docking platform <NUM>, one at each end. The aim of the torque springs <NUM> is to apply a torque to the articulation that joins the docking platform <NUM> with the support ring <NUM>, so one pin <NUM> needs to be fixed to the docking platform <NUM> and the other pin <NUM>' needs to be fixed to the support ring <NUM>.

Accordingly, after the customer satellite <NUM> is separated from the satellite adapter <NUM> the required distance to avoid the interference with the satellite adapter <NUM>, the docking platforms <NUM> start rotating. The docking platforms <NUM> rotate due to the action of the torque springs <NUM> until the lower surface of the main body <NUM> of the docking platforms <NUM> touches the upper surface of the support ring <NUM>. In order to avoid the rotation of the docking platform <NUM>, a latch mechanism <NUM> can be introduced in the articulation between the docking platform <NUM> and the support ring <NUM>, such that when the docking platform <NUM> touches the support ring <NUM>, the latch <NUM> can be activated by a spring.

The next phase is the shock absorbing position. <FIG> shows the position in which the support ring <NUM> is displaced vertically to achieve a gap between the docking assembly <NUM> and the upper surface of the satellite adapter <NUM>. The aim of this gap is to protect the satellite adapter <NUM> and the target satellite <NUM> when the contact between them takes place. The shock energy is absorbed by the support ring <NUM>.

For that purpose several longitudinal springs <NUM> (for instance, three ones, in <FIG>) are arranged, fixed to the end ring <NUM> of the satellite adapter <NUM> on one end and to the support ring <NUM> on the other end. The longitudinal springs <NUM> are positioned around the support ring <NUM> circumference with the same angular distance between each other.

When the support ring <NUM> is in the retracted position, the longitudinal springs <NUM> are compressed and the support ring <NUM> is held by the linear actuators <NUM>. In order to deploy the support ring <NUM>, the actuators <NUM> stop applying force and the longitudinal springs <NUM> are elongated, displacing the support ring <NUM> in the axial direction.

As the longitudinal spring <NUM> needs to be attached to the structure of the satellite adapter <NUM>, a fix element <NUM> has been designed to serve as a platform to fix the longitudinal springs <NUM>. The other end of the spring <NUM> is attached to the bottom surface of the support ring <NUM>, as seen in <FIG>.

The next phase, called soft docking, starts with the approximation of the satellite adapter <NUM> to the target satellite <NUM>. The guide members <NUM> of the docking platform <NUM> correct the misalignments of the docking, and the interface ring <NUM> of the target satellite <NUM> touches the upper surface of the main body <NUM> of the docking platform <NUM>. In that instant the first docking phase takes place, in which the initial engagement means <NUM> on the main body <NUM> make the initial engagement. In the embodiment of <FIG>, <FIG> and <FIG> the initial engagement means <NUM> are latches. After this initial engagement the target satellite <NUM> is partially constrained by the initial engagement means <NUM> (i.e., the latches in <FIG>, <FIG> and <FIG>).

The latch <NUM> is outspread from its support element before the approach of the target satellite <NUM>. The position of the latch <NUM> is achieved using a spring attached both to the latch <NUM> and to the latch support. Then, when the interface ring <NUM> of the target satellite <NUM> touches the latch <NUM>, the spring <NUM> is compressed and the latch <NUM> is inserted progressively in the latch support. The target satellite <NUM> will continue approaching the satellite adapter <NUM> and finally, the ring <NUM> of the target satellite <NUM> makes contact with the docking platform <NUM>. At that instant, the ring <NUM> stops making contact with the latch <NUM> and the latch <NUM> rotates to its initial position pushed by the force of the spring (see the steps in <FIG>). Finally the latch <NUM> presses the inner flange of the interface ring <NUM> of the target satellite <NUM>.

At this point, the soft docking phase finishes. The satellite adapter <NUM> and the target satellite <NUM> are prevented from axial relative displacement between them. Then, the shock produced by the contact will compress the longitudinal springs <NUM> and will retract the support ring <NUM> until the maximum compression of the springs <NUM> is achieved. After that, the springs <NUM> will apply the same compression force in the opposite direction to the support ring <NUM> and an oscillatory movement between them will take place.

This oscillatory movement is attenuated by the pull and push forces that the linear actuators <NUM> will apply, so, after some time, the relative movement between the satellite adapter <NUM> and the target satellite <NUM> will stop.

However, in order to achieve a rigid joint between the satellite adapter <NUM> and the target satellite <NUM> and to be able to perform the mission in a correct way, it is necessary to perform the final phase of the docking.

This phase, called the hard docking, is performed by retracting the support ring <NUM> until it reaches its initial position (see <FIG>). As a result of this axial movement, the pressure member <NUM> of the docking mechanism <NUM> rotates, applying a force on the inner flange of the interface ring <NUM> of the target satellite <NUM> (<FIG> and <FIG>), thus completing the docking.

This docking mechanism <NUM> is not only fixed to the docking platform <NUM>, but is also articulated to the end ring <NUM> in the satellite adapter <NUM>. When the support ring <NUM> is displaced upwards to place in the shock absorbing position, the docking mechanism <NUM> makes the pressure member <NUM> (for instance, in the shape of a bridle) open.

Conversely, when the actuators <NUM> are retracted to move the support ring <NUM> downwards, the pressure member <NUM> is moved in the opposite way, that is, the pressure member <NUM> closes. In this way, when the support ring <NUM> reaches its final position, the pressure member <NUM> applies the maximum force on the inner flange of the interface ring <NUM> of the target satellite <NUM> necessary to achieve a rigid docking.

<FIG> shows the configuration of the docking mechanism <NUM> in its final position.

When the actuators <NUM> retract the support ring <NUM>, the springs <NUM> attached to this support ring <NUM> will compress. In this way, the actuators <NUM> will be placed in a configuration of continuous attraction in order to ensure the hard docking during the mission.

However, it could be possible to introduce a latch system between the support ring <NUM> and the end ring <NUM> of the satellite adapter <NUM>, releasing the continuous actuation of the actuators <NUM>. As there are several longitudinal springs <NUM> placed on the support ring <NUM>, once the hard docking is completed, the springs <NUM> will apply an upwards force to the support ring <NUM> and the hard docking will be rigid, avoiding possible clearances.

Claim 1:
Docking system (<NUM>) for spacecraft, comprising an end ring (<NUM>) and a docking assembly (<NUM>), the docking assembly (<NUM>) comprising a support ring (<NUM>) connected to a plurality of linear actuators (<NUM>) for deployment and retracting, the support ring (<NUM>) and the end ring (<NUM>) being connected, such that in a stowed position the docking assembly (<NUM>) is placed inside the end ring (<NUM>) and in a deployed position the docking assembly (<NUM>) is placed out of the end ring (<NUM>),
characterized in that the docking assembly comprises at least three docking platforms (<NUM>) hinged to the support ring (<NUM>), each docking platform (<NUM>) comprising:
- a main body (<NUM>),
- a guide member (<NUM>),
- initial engagement means (<NUM>), and
- a docking mechanism (<NUM>) fixed to the main body (<NUM>) and articulated to the end ring (<NUM>),
the support ring (<NUM>) and the end ring (<NUM>) being connected by elastic means (<NUM>).