WO2011110701A1 - System for adjusting the position and attitude of orbiting bodies using guide satellites - Google Patents

Abstract

System for adjusting the position and attitude of orbiting bodies using guide satellites that includes a satellite (1) fitted with a system for outputting a high-speed ion stream (3) and a secondary propulsion system (4), said satellite being located close to a body (2) the orbit and/or position of which is to be adjusted, and the position of the satellite being controlled such that the ion stream generated is directed against the body (2) generating the thrust necessary to adjust the orbit and/or orientation thereof. The system can be used to move active satellites, space debris or any other celestial body, transmitting the necessary thrust without physical contact with the body the orbit and/or orientation of which is to be adjusted. This helps to prevent problems related to collisions in orbit and enables complex rendezvous and docking manoeuvres.

Description

 SYSTEM OF MODIFICATION OF THE POSITION AND ATTITUDE OF BODIES IN ORBIT THROUGH GUIDE SATELLITES

Technical Field of the Invention

The present invention is framed in the aerospace field. More particularly, it belongs to the guidance, maneuver and control systems and mechanisms for satellites. Specifically, to those who employ propulsion through a directed flow of ions as a means to modify the position and / or orientation (attitude) of satellites or more generally bodies in orbit.

State of the Art

The prior art has already dealt with the problem of modifying orbits of bodies by means of satellites. In almost all cases, the proposed system has always included a physical docking system between the satellite and the orbiter body involved.

The problem with these methods is that, in general, the spatial object whose orbit is to be modified is non-cooperative, that is, it does not have any control system designed to engage with another satellite. In addition, the spatial object in the worst case will have an unstable position without a fixed orientation (for example, rotating chaotically around its center of mass) as often happens in the case of space debris. Performing a controlled coupling in these cases constitutes a very difficult and risky technological problem, for which advanced robotic tools and highly complex control systems are required.

Currently, a solution to this problem is addressed in US application US 2007/0285304. In it, it is proposed to transmit a thrust between satellites through a flow of gas ejected by a rocket. The expulsion rate of the particles of said gas is provided by a chemical combustion reaction. However, the gas thrust system suffers from important limitations. First, the angle of divergence of the gas expelled by the nozzle always exceeds 20-30 degrees, so that in order to effectively transmit the thrust the distance between the guide satellite and the body pushed by the gas must be very small. Second, the system has a very low specific thrust, when compared to modern ionic thrusters, so that fuel consumption is very high. These limitations may be overcome using radically different technology propellants that instead of gas expel high speed plasma. The ions that make up this plasma are accelerated by an electrical or electromagnetic system at speeds of 30 km / s and more. Although ionic propellants are technologically many more complex than gas propellants and require a power system to expel high-speed ions, they offer two fundamental advantages: a much higher specific thrust and a much smaller divergence angle.

One of the most urgent and interesting needs for the current market is the modification of geostationary satellite orbits that have reached the end of their useful life, usually due to fuel depletion. Due to orbital disturbances of various types, these satellites move away from their nominal orbit, with the consequent deterioration in performance and in their role. Various analyzes have shown that it would be economically convenient to correct the orbits of these satellites by means of guide satellites equipped with coupling systems; Some systems for this purpose have already been patented, for example, in US 6,945,500, which describes a system to extend the life of a satellite (Satellite Life Extension Spacecraft) that includes a mechanism for connecting with the satellite whose life you want to extend, together with a distributed rocket system to locate and control the center of mass of the set of the two satellites. It is noted that, in said invention, push transmission requires that the two satellites be mechanically connected. Making a connection of this style requires prior meeting and coupling maneuvers ("rendezvous and docking").

Brief Description of the Invention

The present invention relates to a system for modifying the position of a body for a guide satellite characterized in that it comprises:

 - primary means of propulsion by ejection of a flow of ions to influence and effect a thrust on the body;

 - a measuring module configured to estimate by means of the information obtained with a radar, the mass of the body, the effective section (A ^) of the body for impact and the relative distance (d) of the body with the guide satellite;

 - a control module coupled with the measurement module, said control module configured to activate and direct, in accordance with the orientation and relative distance (d) with the body, the primary means;

- the control module further configured to activate secondary propulsion means of the guide satellite to orient and move said guide satellite, the means Secondary controlled according to the estimated mass (m2) of the body and the propulsion force generated by the activated primary means, so that the secondary means compensate for variations in the relative distance (d) to keep it substantially constant.

This proposed system is suitable for the modification of orbits and / or position of a body that is orbiting. Specifically, this system allows to move active satellites (to correct orbit injection errors or to prolong their life), space debris or small celestial bodies (asteroids, meteoroids, etc.) thanks to the action of guide satellites. The necessary thrust is transmitted without physical contact with the body whose orbit and / or position is to be modified thus avoiding problems of collision in orbit and without requiring complex "rendezvous and docking" maneuvers. The prior art systems mentioned, perform very complicated maneuvers and present a high risk of collisions or damage to the participating satellites. The present invention drastically reduces the complexity and risks associated with this type of maneuver since any type of mechanical contact between the bodies involved is avoided, the thrust being transmitted indirectly through a high-speed ion stream.

The basic purpose of the system is to solve the fundamental problem of the displacement of satellites and, more generally, of space objects allowing the displacement to occur in the simplest way possible and minimizing the risks of collisions or damage of satellites involved in the maneuver.

It is proposed to transmit a thrust between two satellites or, more generally, between a satellite and an orbiting body, through a flow of high-speed ions avoiding direct physical contact between the two entities. In the current state of the art it is possible to accelerate ions produced by electric propulsion systems, up to speeds exceeding 30 km / s. These propulsion systems have never been used to produce directed thrust on another satellite located some distance from the source of the ion flow. In addition, modern ionic propellants can reach very small values of the flow divergence angle. The theoretical divergence value can be estimated using the following formula:

where v _E is the ion ejection velocity, m _i its mass, T _eV the thermodynamic temperature of the plasma before being accelerated (normally T _eV * 1 eV) and q _e = 1.6 ^χ 10 ^"I9 is the charge of the electron.

This allows the entire ionic flow to be transmitted at a maximum distance given by: d = Itanp ^(B)

A _eff being the effective section, with respect to the direction of the guide satellite, of the space object to be controlled.

Since it is possible to reach speeds v _E of 30 km / s and greater, the angle of divergence φ can reach a theoretical value, for example using xenon ions, less than 3 degrees, which corresponds to a distance of about \ ^ A _eff . Using mercury the angle of divergence drops to less than 2 degrees while with argon it rises to more than 4 degrees. Note that neon and krypton ions are also suitable for use in the invention.

In fact, and due to other factors, the value of φ is larger than indicated by equation (A) so it is reasonable to take 6 degrees and a maximum distance of action as the limit of current technology:

The thrust transmitted by the ion stream can be calculated in several ways. For example, a widely used formula is as follows:

Ρ _ρ1 * 2η— (D)

Where P is the electrical power available on board the satellite and dedicated to the propulsion system, and 77 is the performance of the propulsion system. The value of η depends on the type of propellant and exceeds 70% in ionic engines. The value of P depends on the mass dedicated to the power system on board the satellite. For example, a 1 kW system It normally has a mass cost of 10-20 kg. Using said system and assuming 7 = 0.7 and v _£ = 30 km / s, a thrust of about 0.05 N is reached. Although small, this thrust turns out to be several orders of magnitude greater than the atmospheric resistance force, acting on a conventional satellite , in low Earth orbit (400 to 1000 km) and can be used, for example, to clear congested orbits of space debris.

Once the technological limits of the thrust system have been defined, the detailed description of the invention can be passed.

Description of the figures

To complement the description and in order to help a better understanding of the features of the invention, the present specification of figures 1 to 4 is attached, as an integral part thereof.

The invention will be described in more detail below with reference to an exemplary embodiment thereof represented in said figures. Figure 1 represents an example of how the displacement of a non-active body or satellite (2), target of the ion flow (7) directed by the ion generation / expulsion system (primary propulsion system) (3) occurs. installed on a guide satellite (1) that includes a secondary propulsion system (4) to maintain a distance and a relative position between both objects and a radar system (5) to measure the distance and the effective section of the object.

Figure 2 represents an example of how the displacement of an active satellite (2) occurs. The key elements of the system are the same except for distance measurement that must be more sophisticated and can include not only a radar system but also a vision system to estimate the orientation of the active satellite and locate suitable points to direct the flow of ions.

Figure 3 schematically represents an embodiment of the invention, with inputs and outputs so that said system can be provided with the devices and subsystems responsible for calculating the parameters necessary to perform the maneuver properly. Figure 4 represents a block diagram, where the relationships between the different modules have been illustrated. They appear in the figure, the control module (9) that governs the ionic propeller (3) used to move external bodies, the propeller of the satellite itself (4), the measurement module (10) that collects information from a radar (5) ) and optionally a vision system (11). Optionally, a trigger device (12) is shown that allows the ionic propeller (3) to be directed towards a target without the need to maneuver the guide satellite as a whole. These non-essential modules are shown in dashed line. Detailed description of the invention

The known parameters at the input are the mass m, of the guide satellite, the maximum distance d _max that the ion flow can reach; The parameters estimated at the entrance are the mass m ₂ of the object to maneuver, its effective section A,. * and its position d with respect to the guide satellite. On the way out we have the forces F _pi and F _p2 that must be provided by the propulsion systems with the necessary guidance for the guide satellite so that the maneuver is correctly fulfilled, and the direction of firing of the ion flow. The operation of the invention according to an embodiment of the invention is explained below. For this, the devices that are involved in one way or another are listed.

 Space object to be maneuvered (2): It can be artificial (e.g. active or inactive satellite or part thereof, rocket etc.) or natural (asteroid, comet, meteoroid, etc.).

Guide satellite (s): 1 or more satellites that are responsible for transmitting the thrust to the object to be maneuvered without coming into physical contact with it.

 Electrostatic ionic propellant (3), that is to say a system that ionizes argon, mercury or xenon atoms by exposing electrons from a cathode, and accelerates the ions produced by passing them through charged grids before expelling them generating the desired thrust. This propellant system will always be used as a primary propellant to transmit the thrust the thrust to the object to be maneuvered and can optionally be used as a secondary propellant to control the relative distance between the bodies involved in the maneuver.

Power system (8): Provides the guidance satellite with the power needed to power the various propellers and measuring and control instruments. It can be any: solar panels, electrodynamic moorings, nuclear system, etc. Secondary propeller (4), that is to say a system, housed in the guide satellite, that provides the necessary thrust to control the relative distance between the bodies involved in the maneuver. It can be any type of propellant (ionic, chemical, electrodynamic mooring, solar sail, etc.).

Measuring radar (5): It is housed in the guide satellite and has the function of measuring the relative distance between the guide satellite and the object to maneuver and to estimate the size (ie the effective section) of the latter.

 Attitude control module (9): It is housed in the guide satellite and has the function of controlling the orientation of the latter so that it can correctly direct the flow of ions to the object to maneuver.

 Relative distance control module (10): Collect the necessary information from the measurement radar (5) to control the two propulsion systems (3,4) so that the relative distance between the bodies is kept constant. Optionally, if it exists, you can use the information provided by a vision module.

Optional devices:

 Vision module (11): It is housed in the guide satellite and has the function of measuring the relative distance between guide and object satellite to maneuver and estimate not only the size (effective section) of the latter but also its orientation and necessary physical characteristics to ensure that the ion stream is directed to non-sensitive parts of the controlled object avoiding possible damage to the operation of the latter. Said systems elaborate an image of the analyzed object which has to be processed by last generation recognition algorithms so that sensitive and non-sensitive parts of the controlled object can be identified.

Pointing device (12): it allows to vary the ejection of the ion flow without the need to maneuver the orientation of the guide satellite so that the control of the object to be maneuvered is more effective.

As can be seen from the drawings, the invention, which is implemented in a guide satellite (1), has a system for generating and expelling a flow of high-speed ions (3) directed towards the celestial body whose orbit and / or position you want to modify (2), and a secondary propulsion system (4). The system of generation and expulsion of high-speed ions produces a thrust F _pl on the body (2) and, by Newton's law of action and reaction, an equal and opposite thrust (-F ^) on the guide satellite. The secondary propulsion system (4) must provide an additional thrust F ₂ to balance the system and achieve relative acceleration (a _{re /} ) null between the two bodies according to what is indicated by the following equation:

Equation E is the control equation of the system to perform the movement maneuver of the celestial body (2). In this equation the mass m _x of the guide satellite is known with some precision while m ₂ is known only approximately (for example through spatial object databases). A measurement system based on a radar (5), or if available, in more precise vision systems (1 1), is able to estimate the distance d between the two bodies at different times.

The attitude control module (9) is responsible for performing the necessary correction and acting on the two thrusters so that said distance remains constantly equal to a desired distance (the latter cannot exceed the maximum limit established by equation C) .

Depending on the behavior of the system during the control maneuver, it will even be possible to refine the estimate of the mass m ₂ of the object to be moved.

Note that the secondary propulsion system (4) does not have to be another ionic propellant. Electrodynamic moorings, solar sails, aerobraking propulsion systems, and liquid or solid propellant rockets can be used. The relative distance control module (10) is necessary not only to keep the distance between the body and the guide satellite constant but also to locate the position of the body with sufficient precision and direct the flow of ions (7) towards it.

In the event that the body to be controlled is an active satellite, it will be necessary to prevent the flow of high-speed ions from impacting any sensitive part of the satellite. For this reason, the guide satellite can additionally have a more sophisticated vision module (1 1) capable of estimating the attitude of the active satellite and locating sensitive parts of it and which are robust. The vision module (11) produces an image of the satellite that has to be processed by recognition algorithms of latest generation so that the ion flow direction control system acts as established.

In this invention, the guide satellite is located in the vicinity of the geostationary satellite until it reaches a maximum distance d _max given by the formula C. A vision system locates the side of the satellite that corresponds to the nozzle of the apogee engine, that is to say the Primary propulsion system (3) used to maneuver these types of satellites throughout their lifespan. The position of the guide satellite is controlled so that the flow of high-speed ions can be directed against the nozzle of said motor (fig. 2) because it is a stronger part.

Next, the displacement maneuver begins in which the guide satellite provides the necessary thrust, while controlling the relative distance between the two satellites (which must be kept constant) and the alignment of the ion flow with the nozzle of the controlled satellite. The latter can be used not only to provide the thrust, but also to control, through minor modifications of the firing direction, the position of the geostationary satellite when necessary.

Claims

Claims 1. System for modifying the position of a body (2) for a guide satellite (1) characterized in that it comprises: - primary means (3) of propulsion by ejection of a flow of ions to influence and effect a thrust on the body (2); - a measuring module (10) configured to estimate by means of the information obtained with a radar (5), the mass (m2) of the body (2), the effective section (Α ^) of the body (2) for impact, the position and relative distance (d) of the body (2) with the guide satellite (1); - a control module (9) coupled with the measurement module (10), said control module (9) configured to activate and direct, according to the orientation and relative distance (d) with the body (2) , the primary means (3); - the control module (9) further configured to activate secondary means (4) for propulsion of the guide satellite (1) to orient and move said guide satellite (1), the secondary means (4) controlled according to the mass ( m2) estimated of the body (2) and the propulsion force generated by the activated primary means (3), so that the secondary means (4) compensate for variations in the relative distance (d) to keep it substantially constant. 2. System according to claim 1, characterized in that the measuring module (10) is configured to communicate with the control module (9) if the relative distance (d) exceeds a limit value (d _ma ) that depends on the effective section (Ae _ff ) of the body (2). 3. System according to claim 2, characterized in that the control module (9) is configured to deactivate the primary propulsion means (3) and to activate the secondary propulsion means (4) to orient and move the guide satellite (1) . System according to any one of the preceding claims, characterized in that the primary propulsion means (3) comprise a pointing device (12) for directing the flow of ions emitted by said primary means (3) in a specific direction. System according to any one of the preceding claims, characterized in that the measuring module (10) comprises a vision system (1 1) configured to estimate the relative distance (d), the effective section, the orientation and identify suitability of the area of impact on the body (2). System according to any one of the preceding claims, characterized in that the control module (9) accesses a database to collect information about the body (2) from the data obtained by the measurement module (0) to further determine other characteristics and confirm the data obtained. 7. System according to claim 5 or 6, characterized in that when the body (2) is an artificial satellite, the vision system (11) is configured to detect the nozzle of the apogee motor in said satellite and estimate its effective section ( A corresponding _eff ) to guide the flow of ions emitted by the primary means (3) only towards this area. System according to any one of the preceding claims, characterized in that the primary means (3) of propulsion eject a flow of argon ions. 9. System according to any one of the preceding claims, characterized in that the primary propulsion means (3) eject a flow of mercury ions. System according to any one of the preceding claims, characterized in that the primary propulsion means (3) eject a flow of xenon ions. 1 System according to any one of the preceding claims, characterized in that the primary propulsion means (3) eject a flow of krypton ions 12. System according to any one of the preceding claims, characterized in that the primary propulsion means (3) eject a neon ion flow. 13. System according to any one of the preceding claims, characterized in that the measurement system (10) allows to adjust the estimation of the body parameters (2) iteratively when the information supplied by the radar (5) and / or the system of Artificial vision (1 1) is updated.