Deployment control system for spacecraft having to fly in formation using simultaneous high precision determination of their positions

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DESCRIPTION

Control system for the deployment of spacecraft that must fly in formation, by simultaneous and highly precise determination of their positions

The invention relates to spacecraft, such as satellites, whose object is to move in formation to collectively guarantee a mission, and, more specifically, to control the deployment of said devices in space with a view to forming a formation that present a selected configuration.

In some formation flight missions, the distance between spacecraft (for example, satellites (or “flyers”), after deployment should be short, even very little (for example, a dozen meters) .

Deployment is initially envisaged to bring all spacecraft "into position" under dynamic conditions conducive to relative autonomous control. This deployment can be initiated by launching one or more modular systems, with one or more different launchers. In the present document, "modular system" is understood as a module temporarily transporting several spacecraft, comprising propulsion means, and from which the spacecraft are separated as soon as the propulsion means is no longer useful. Deployment can also be done by direct launch of all spacecraft into a transfer orbit.

All these deployment initiation techniques involve at least two common drawbacks: the distances between satellites can be such that visibility between satellites is not assured, and the "meeting" of spacecraft can be done with final relative movements that can carry risks of collision. In this document, "encounter" is understood to be the moment in which all spacecraft meet again inside a sphere, called an encounter space, of equal diameter to the maximum range of the relative positioning metrology system, for example, the radio frequency (RF) link between satellites (defined by its integrated RF relative metrology means), i.e. typically 8 km for the present example.

Today, no training deployment technique has been effectively applied. Only the technique of positioning a spacecraft with respect to a star is well known and mastered. This consists in measuring, by means of a station implanted on the ground, the successive positions of the spacecraft, and in determining, also on the ground, the maneuvers that aim to position the device with respect to the star as a function of the position measurements. These maneuvers are then (tele-) transmitted to the spacecraft in order to carry them out with the aid of the integrated displacement means (nozzle or nozzles and / or actuator or actuators).

Indeed, some deployment methods have been proposed, but none of them take into account the possibility of losing an Rf link between satellites, between launch and geostationary position (final position within the formation). However, during a single launcher cast, compliance with the RF range sphere (or meeting space) does not have to be guaranteed. On the other hand, spacecraft do not have to have sufficient maneuverability to compensate for the force differentials exerted on each one of them.

Furthermore, for a launch by several launchers, the communication of the elements of different types (modular systems and spacecraft) is never taken into account. Now, this communication necessarily requires recourse to an external means, because in this case the distances between satellites are much greater than the RF range. This external medium can only be the ground station that is responsible for taking position measurements, determining the maneuvers of spacecraft and transmitting these maneuvers through waves in the form of trajectory corrections so that it positions them in their respective nominal trajectories, while ensuring their encounters with an accuracy compatible with the RF range.

American patent US4375697 describes a satellite formation that includes a control device that allows maintaining the position of each satellite in the formation already formed. A difference with the object of the present application is that said control device is implanted in a control satellite that belongs to the satellite array and that constitutes a reference.

The American patent US5979830 describes a method that allows to preserve a formation of geostationary satellites in a given position, using an optical link between satellites.

The publication “Navigation data definitions and conventions, CCSDS 500.0-G-2, November 2005” describes a set of techniques for measuring speed, distance and angular direction for aerospace navigation.

Therefore, the invention aims to improve the situation.

For this purpose, it proposes a specific system for controlling the deployment of at least two spacecraft, provided with maneuvering means and whose object is to move according to a selected formation, and including a control device implanted on the ground comprising, for one side, a first means of

ES 2 397 646 T3 measurement responsible for determining substantially simultaneously and with high precision the orbital positions of spacecraft and, on the other hand, first calculation means responsible for determining for each of these spacecraft, as a function of their orbital positions, maneuvers that aim to position each of them substantially at a selected instant, at a selected position relative to a reference path, taking into account an hourly law of a reference device on this reference path, in order to position the formation in a selected configuration. The first measurement means comprise at least two measurement stations implanted on the ground in selected places and each one in charge of measuring substantially simultaneously at least the distances that separate them from the different spacecraft, as well as processing means in charge of determining orbital positions of spacecraft at least from substantially simultaneous distance measurements.

The control system, according to the invention, may include other features that can be considered separately or in combination, and in particular:

- the reference device is either one of the spacecraft of the formation (in this case the other spacecraft are called follower devices), or a fictitious device representative of the entire formation (for example, located at the center of gravity of is);

- its first calculation means can be responsible for determining maneuvers that aim to substantially position each device at a selected instant in a selected position with respect to the reference path and with a selected relative speed (for example, substantially zero) with respect to the reference device, taking into account the time law of the latter in its reference trajectory;

- it may comprise, on the one hand, second relative metrology means, for example, of the radio frequency (RF) type, implanted in each of the devices and responsible for taking at least measurements of the position of its own device with respect to at least some of the other spacecraft, when the latter are located in a so-called meeting space and, on the other hand, second calculation means implanted in each of the devices and in charge of estimating the risks of collision with at least some of the other spacecraft at least from the measurements of the relative position and, in the event of a risk of collision, that exceed a selected value, determine the local and temporary evasion maneuvers to be carried out;

> the second means of relative metrology can be responsible for taking at least measurements, as well as possibly estimations, of the relative speed of your spacecraft at least with respect to some of the other spacecraft. In this case, the second calculation means are responsible for estimating the risks of collision with at least some of the other spacecraft at least from the relative position measurements and / or relative speed measurements or estimates;

> the second relative metrology means and the second calculation means can be put into operation at the latest during the separation of their respective devices with respect to a launcher;

- Your first means of measurement can be responsible for determining the orbital positions of spacecraft using the technique called "Delta DOR" (for its acronym in English, "Delta Differential One way Range" (Differential Variation of unidirectional distances).

The invention also proposes a control device whose object is to form part of a deployment control system of the type as presented above.

The invention also proposes a spacecraft (such as a satellite) whose object is to move in formation within a group of spacecraft of the same type, and which comprises maneuvering means, and second relative metrology means and second means for calculating a deployment control system of the type as presented above.

Other characteristics and advantages of the invention will become apparent through examination of the following detailed description, and the attached drawing, in which the only figure illustrates in a very schematic way an example of deployment of two spacecraft by means of a control system according to the invention. The attached drawing may not only serve to complete the invention, but also to contribute to its definition, if applicable.

The object of the invention is to allow control of the deployment of spacecraft in space with a view to constituting a formation having a selected configuration.

Below, it is considered, by way of an illustrative and non-limiting example, that spacecraft are observation satellites (possibly of the "flyer" type) that fly (or are intended to fly) in formation in order to fulfill a mission. of space or terrestrial observation.

But the invention is not limited to this type of spacecraft. Indeed it refers to all spacecraft

ES 2 397 646 T3 whose object is to fly in formation according to a selected configuration (optionally modifiable).

The invention intervenes once the (space) devices have been separated from their launcher or launchers, as well as possibly from one or more modular systems (or transport module or modules with autonomous means of propulsion).

The invention proposes a system whose object is to control the deployment of at least two spacecraft ESi provided with maneuvering means MDi and whose object is to move according to a selected formation. In the present document, "deployment" is understood as the phase that allows positioning ESi spacecraft with respect to each other in such a way that they constitute a formation that presents a selected configuration (or geometry), in a durable (although modifiable) way.

In the example illustrated in the only figure, only two spacecraft ES1 and ES2 (i = 1 and 2) have been represented, which are intended to fly in formation. But the invention is not limited to this number. Indeed, it refers to any number of spacecraft as long as this number is greater than or equal to two (2).

A deployment control system according to the invention comprises at least one control device implanted on the ground and comprising first measurement means MM1j and MT and a first calculation module MC1.

In the following, it is considered, by way of illustrative and non-limiting example, that the control device is implanted on the ground. But this is not mandatory. Indeed, it is possible to contemplate the possibility that the location is made with respect to a constellation of satellites in flight (for example). The control device is then implanted in satellites of the constellation, which have synchronized clocks.

The first measuring means MM1j and MT are in charge of determining substantially simultaneously and with high precision the orbital positions of the ESi spacecraft that are in the deployment phase (in particular during the sub-phase called cruise (or “cruising ”) That should take them inside a meeting space). In this document, "meeting space" is understood as a sphere of diameter substantially equal to the radio frequency (RF) range between two spacecraft. This range is currently approximately 8 km.

Herein, high precision is understood as an accuracy that is typically about (or less than) a few tens, even a few hundred, of meters for a distance of about one million kilometers.

To achieve this type of precision, for example, the so-called Delta DOR (Delta Differential One way Range) technique, developed by ESOC and JFL and, for example, described in an article accessible at the address Internet http://www.issfd.dir.de/papers/P0118.pdf).

Such a technique makes it possible to obtain an accuracy of approximately 25 m for a distance of approximately 1.5 million kilometers (which corresponds to the point of Lagrange 2 (or L2) with respect to the Earth). The application of this technique requires at least two measuring stations implanted on the ground STj (in this case j = 1 and 2), for example on the ground, and a treatment module MT coupled to the stations Stj.

Any other technique, known to the person skilled in the art, can be used to determine the orbital position, which is based on the observation of devices by remote measuring means, oriented differently and synchronously.

Each (measurement) station STj comprises a first measurement module MM1j capable of emitting signals (for example, radio frequencies (RF) - RF is used for both metrology and communication links) destined for ESi spacecraft, of collect the reflected signals and deduce substantially simultaneously the distances (so-called “range” measurements) that separate it from the different ESi spacecraft, as well as preferably the radial velocities (for example, by Doppler measurements) of the ESi spacecraft. The simultaneity of the measurements taken by the stations STj in the ESi spacecraft makes it possible to avoid errors in the absolute measurement common to all devices.

The MT treatment module is in charge of determining the orbital positions of the different ESi spacecraft from distance measurements, and possible simultaneous radial velocity measurements (with the precision of the synchronized clocks of ground stations). STj) taken by the first measuring modules MM1j from the STj stations. The orbital positions are determined, for example, by means of a recurrent filtering, in time, of the set of the position / velocity measurements, of the Kalman filtering type, based on a prediction model of the trajectories of the ESi spacecraft provided by orbital mechanics and the model of measurements taken by ground stations STj.

As illustrated in the only figure, the treatment module MT, for example, can be implemented in a computer center CC physically separated from the stations STj. But, as a variant, it can be implanted in one of the

ES 2 397 646 T3 stations STj.

The first calculation module MC1, of the control system, is in charge of determining the maneuvers (or maneuver plans) of each spacecraft ESi based on the orbital positions determined simultaneously by the first measuring means MM1j and MT. More specifically, the first calculation module MC1 determines for each spacecraft ESi the maneuvers intended to position it at a substantially selected instant in a selected position (or meeting point) PR with respect to a reference path TR, taking into account the time law of a reference device on this reference path TR.

In this document, "hourly law" is understood to be the law that governs the displacement of the reference device in its reference path TR and, therefore, that allows knowing its theoretical position (as well as possibly its vector velocity at any moment) ) with respect to this reference path TR.

On the other hand, in this document "maneuvers" is understood as an order or orders (tele) transmitted from the ground to an ESi spacecraft and that are intended to correct its current trajectory, in order to direct it towards the point meeting.

As illustrated in the figure, the reference device is, for example, one of the spacecraft (ES1) in the array. In this case, the other ES2 spacecraft are called follower devices. The first calculation module MC1 then first determines the maneuvers aimed at positioning the reference device ES1, substantially on a selected reference path TR. It then determines for each follower device ES2, the maneuvers that are intended to position it at a selected instant substantially in a selected position (or meeting point) PR with respect to the reference path TR, taking into account the time law of the device. ES1 reference path on its TR reference path.

But, in a variant, the reference device can be a fictitious device that represents the whole of the formation. For example, this fictitious device is located at the center of gravity of the formation and its trajectory is the reference trajectory TR. In this case, all ESi spacecraft in the formation are follower devices. The first calculation module MC1 then first determines the reference path TR of the dummy (reference) device. It then determines for each spacecraft ES1, ES2, the maneuvers intended to position it at a substantially selected instant in a selected position (or meeting point) PR with respect to the reference path TR, taking into account the device's time law. dummy (reference) on its TR reference path.

In the following, it is considered by way of illustrative and non-limiting example, that the reference device is the spacecraft ES1 of the formation.

It is important to note that each ES2 tracker device has a selected position PR that substantially constitutes its final position with respect to the ES1 reference device and, therefore, within the selected formation (in the eventual local and near final corrections on the which we will come back to later). The objective is therefore to determine for each follower device ES2 a specific meeting point PR with respect to the follower device ES1, taking into account its reference trajectory TR, then the maneuvers that will allow it to go to this meeting point PR. A meeting point PR is a point in a space of five dimensions defined, for example, by the first five orbital parameters (the sixth being free, for example, the position in the orbit).

Thanks to the "maneuver plans" defined in this way by the first calculation module MC1 for each of said spacecraft ESi, it is possible to constitute the formation according to a selected configuration.

The first calculation module MC1, can, for example, be implemented in the calculation center CC, as illustrated in the only figure. But, as a variant, it could be implanted in one of the STj stations.

The calculation center CC (when it contains the first calculation module MC1 and has adapted transmission means) or one of the ground stations STj, for example, retransmit the maneuver plans to the different spacecraft ESi.

It should be noted that eventually, several successive maneuvering plans may be necessary for an ES2 follower device to reach the level of its specific PR rendezvous point. Consequently, the first measuring means MM1j can be configured to take measurements of the distances, as well as preferably of the radial velocities, either periodically, or in order, for example, of the first calculation module MC1.

Preferably, the maneuvers (or maneuver plan) that the calculation module MC1 determines for a follower device ES2 are intended to position it at a selected instant not only substantially at the level of a selected position (or meeting point) PR with respect to the reference path TR, but also with a relative speed, selected with respect to the reference device ES1. This relative velocity, for example, is selected substantially equal to zero (0), in order to fix the formation in the selected configuration.

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This is done taking into account the extrapolated hourly law, possibly corrected, of the reference device ES1, that is, the total estimate of the positions and speeds. The Kalman filter estimates everything at the same time, thanks to an accumulation of measurements of the distances, as well as preferably of the radial velocities in a given space of time. The prediction also provides the future hourly law, with a variation that diverges with time since the last measurement.

Conventionally, when an ESi spacecraft receives instructions from the ground defining the maneuvers that have been determined for it, it communicates them to an orbit correction module MCOi with which it is provided. The latter, an integral part of the semi-autonomous on-board control program, better known by the acronym GNC (for its acronym in English “Guidance Navigation and Control”), is in charge of converting in each time passage of the GNC the maneuvering instructions received from the ground into force and / or torque, and after converting this force and / or torque into order or orders that it transmits to the maneuvering means MDi of its ESi spacecraft in an understandable way .

The MDi maneuvering means can be of any type. In particular, it can be a nozzle or nozzles and / or an actuator or actuators, for example chemical, in particular for impulse or plasmic operations, by cold gas, by emission of electric or electromagnetic field, in particular for continuous impulse maneuvers.

It should be noted that as a variant the relative positions can be determined from the differentials of the measurements of each one of the ground stations STj instead of by estimating the absolute positions of each one of the spacecraft ESi. As the covariance of this differential magnitude is strong, the precision of the estimate can be improved, once it is entered in the constant data of the Kalman filter. The prediction model used can then be that of the differential disturbances between two given ESi spacecraft, assuming that the absolute position measurement of one of the ESi spacecraft has been previously taken, for example, the reference one.

The control system preferably contains, as a complement to its control device implanted on the ground (MM1j, MT, MC1), a second relative metrology means MM2i and a second calculation modules MC2i implanted at least in each of the ES2 trackers. (as well as eventually in the reference device ES1).

The second means of relative metrology MM2i are responsible for taking at least a few sets of instantaneous measurements of the relative position of the ESi spacecraft in which they have been implanted at least with respect to some of the other ESi spacecraft 'when they are in the Inside the meeting space, that is, within reach of the radio when the measurements are of the radio frequency (RF) type.

The second relative metrology means MM2i can be of another type than radio frequency.

They can also and eventually take relative velocity measurements that complete the sets of instantaneous relative positions measurements and allow at least some of the relative positions of the other ESi 'spacecraft to be determined with more precision with respect to their ESi spacecraft.

Each second MC2i calculation module is responsible for estimating the collision risks of its ESi spacecraft with each of the other ESi '(i #) spacecraft, object of measurements, at least from the measurements on the relative position and of the eventual relative speed measurements taken by the associated second relative metrology means MM2i. For example, the second calculation module MC2i supplies, for each spacecraft ESi 'object of measurement, a value representative of the probability that it will collide with its spacecraft ESi. The second calculation module MC2i can then compare each probability value with a threshold value, and when a probability value is higher than the threshold value (i.e. in case of collision risk), it determines local and transient avoidance maneuvers that They are intended to allow your ESi spacecraft to bypass at least one other ESi 'spacecraft.

In order to improve the precision of the collision risk estimates, the second relative metrology means MM2i can also take care of estimating (predicting) the relative positions and possibly the relative vector velocities of your ESi spacecraft with respect to each of the other spacecraft ESi 'object of a measure of distance and relative direction. For this, for example, they can include a Kalman filter in charge of giving the relative positions and speeds from the beginning of the measurement campaign to the last measurement moment, that is, the hourly law, thus allowing their values ​​to be extrapolated in the future. , with the condition of knowing the effect of the actuators and differential disturbances. It should be noted that the deviation from the estimate in the future (extrapolation) grows over time. A priori, the Kalman filter is not used in backup mode, but this could be raised.

The second calculation module MC2i is in charge of estimating the collision risks of your ESi spacecraft with each of the other ESi spacecraft 'object of measurement, based on measurements and / or estimates of relative position and possibly from measurements and / or relative speed estimates.

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The second relative metrology means MM2i and the second calculation module MC2i, which are implemented in at least each follower device ES2, constitute what the person skilled in the art generally calls an anti-collision (or evasion) device. In an ESi spacecraft, any type of anti-collision (or avoidance) device known to the person skilled in the art can be used.

The evasion maneuvers decided by a second calculation module MC2i have priority over the maneuvers (or maneuver plan) determined on the ground for its ESi spacecraft. Consequently, each time a second calculation module MC2i decides that its ESi spacecraft must perform evasion maneuvers, it communicates them to the associated orbit correction module MCOi (for example, of the GNC type). The latter converts them into a trajectory correction order or orders, then transmits them to the maneuvering means MDi of its spacecraft ESi in order for it to execute them.

Preferably, the second relative metrology means MM2i and the second calculation means MC2i, which are implemented in at least each follower device ES2, are arranged in such a way that they can be put into operation (or activated) at the latest when their space device is separated. of the launcher or the eventual modular system that transports them momentarily. In this way, each ESi spacecraft can be subject to collision detection from the beginning of the deployment phase.

When a tracker device ES2 has reached the level of its specific meeting point PR, it may happen that the reference device ES1 has slightly varied its reference trajectory TR and / or that it has not exactly followed its hourly law and / or that its speed relative to the follower device ES2 is not the one selected on the ground. In this case, the specific meeting point PR and / or the relative speed of the follower device ES2 must (must) be corrected. The measurements taken on the ground as they are not sufficiently precise, the second relative metrology means MM2i and the second calculation module MC2i, which are implanted in the ESi spacecraft, can then be in charge of positioning the latter precisely, taking into account the actual position of the reference device ES1. In effect, they are able to determine the relative position (and relative speed) of their tracking device ES2 with respect to the reference device ES1, and then compare them with the predicted relative position (and / or the predicted relative speed), and to determine the corrective maneuvers to be carried out by your ES2 tracker device to position it correctly and accurately with respect to the ES1 reference device, with the expected relative speed. This final correction can be done completely autonomously on board, but is activated either by a request from the ground, or automatically once the follower device ES2 has reached the level of the meeting point PR initially determined by ground.

The first MC1 and second MC2i calculation module, the orbit correction modules MCOi and the treatment module MT, can be made in the form of electronic circuits, computer (or computer) program modules, or a combination of circuits and programs. .

It should be noted that the invention makes it possible to reintegrate into a formation a spacecraft (tracker) that has moved away from the reference device a distance greater than the range of the radio (typically 8 km).

On the other hand, it should be noted that when choosing to define each meeting point with respect to the final orbit of the reference device, control problems are avoided during the cruise (or transfer) sub-phase, except the eventual (and preferential) application of an anti-collision function. This makes it possible to simplify the work of the orbit correction module (GNC) during this sub-phase and, therefore, to optimize consumption since it has been reduced to the needs related to navigation and guidance.

The invention is not limited to the embodiments of the control system, the control device and the spacecraft described above only by way of example, but encompasses any variant that a person skilled in the art may consider in the field of the following claims.

In all of the above, an application of the invention has been described in which the first measurement means and the first calculation means of the control device are implemented on the ground (in ground stations). But, the first measurement means and / or the first calculation means of the control device can be implanted in satellites of a constellation in flight that have synchronized clocks. For example, the first measurement means of the control device can be implanted in the satellites of a constellation in flight, while the first calculation means of the control device has been implanted in a ground station. The location of the spacecraft, which must be located according to a selected configuration, is then made with respect to the constellation of satellites in flight.