Geosynchronous orbit satellite position protection method with unconstrained thruster installation direction

Abstract

The invention discloses a geosynchronous orbit satellite position protection method with an unconstrained thruster installation direction, and belongs to the field of satellite attitude and orbit control. The method comprises the following steps: calculating an orbit parameter estimation value of a satellite according to a future orbit position obtained by recursion of a current orbit position; generating a corresponding satellite orbit position keeping strategy according to orbit parameters which do not meet a preset threshold range in the orbit parameter estimation values; and determining whether the satellite orbit position keeping strategy meets the sunlight avoidance constraint and the thruster layout constraint, and if not, adjusting the satellite orbit position keeping strategy according to the sun vector position, the sunlight avoidance demand and the thruster layout at the future position keeping moment. The method can greatly improve the position protection operation efficiency and guarantee the satellite load safety.

Description

Technical Field

[0001] This invention relates to the field of satellite attitude and orbit control technology, and in particular to a method for maintaining the position of a geosynchronous orbit satellite with unrestricted thruster installation direction. Background Technology

[0002] The Remote Sensing 41 satellite is designed with multiple functions, including Earth staring imaging, area array mosaic imaging, linear array pushbroom imaging, and motion tracking imaging, to meet users' needs in various scenarios. As the high-resolution high-orbit optical satellite to date, it also requires high orbital maintenance accuracy, necessitating frequent position-keeping operations in orbit.

[0003] In related technologies, the thrusters of the Remote Sensing 41 satellite are mainly located on the back floor. Using thrusters installed on the waist of the satellite on the east and west sides for position maintenance operations is too inefficient and cannot simultaneously meet the needs of data adjustment and sunlight avoidance, thus reducing the safety of the satellite payload.

[0004] Therefore, there is an urgent need for a geostationary orbit satellite position preservation method with unrestricted thruster installation direction to solve the above-mentioned technical problems. Summary of the Invention

[0005] This invention provides a position-keeping method for geostationary orbit satellites with unrestricted thruster installation direction, which can greatly improve the efficiency of position-keeping operations and ensure the safety of satellite payloads. The technical solution is as follows: On the one hand, a method for position maintenance of geosynchronous orbit satellites with unconstrained thruster installation direction is provided, the method comprising: The estimated orbital parameters of the satellite are calculated based on the future orbital position derived recursively from the current orbital position. Based on the orbital parameters in the estimated orbital parameters that do not meet the preset threshold range, a corresponding satellite orbital position holding strategy is generated. Determine whether the satellite orbital position holding strategy meets the sunlight avoidance constraint and thruster layout constraint. If not, adjust the satellite orbital position holding strategy according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

[0006] On the other hand, a geosynchronous orbit satellite position-keeping device with unrestricted thruster installation direction is provided, the device comprising: The calculation module is used to calculate the estimated orbital parameters of the satellite based on the future orbital position derived recursively from the current orbital position. The generation module is used to generate corresponding satellite orbit position holding strategies based on the orbit parameters in the estimated orbit parameters that do not meet the preset threshold range; The determination module is used to determine whether the satellite orbital position holding strategy meets the sunlight avoidance constraints and thruster layout constraints. If it does not meet the constraints, the satellite orbital position holding strategy is adjusted according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

[0007] On the other hand, a computer device is provided, the computer device including a memory and a processor, the memory for storing computer programs, and the processor for executing the computer programs stored in the memory to implement the steps of the above-described method for maintaining the position of geosynchronous orbit satellites with unconstrained thruster installation direction.

[0008] On the other hand, a computer-readable storage medium is provided, wherein a computer program is stored therein, and when the computer program is executed by a processor, the steps of the above-described method for maintaining the position of a geosynchronous orbit satellite with unconstrained thruster installation direction are implemented.

[0009] On the other hand, a computer program product is provided, including a computer program that, when executed by a processor, implements the steps of the above-described method for maintaining the position of a geosynchronous orbit satellite with unconstrained thruster installation direction.

[0010] The technical solution provided by this invention can bring at least the following beneficial effects: by predicting the satellite orbit one week from now and calculating the satellite's longitude, eccentricity, inclination, and corresponding drift law, a position-keeping strategy is autonomously generated for parameters exceeding error limits; simultaneously, under constraints of sunlight avoidance and thruster layout, the position-keeping attitude is autonomously planned based on the autonomously generated position-keeping strategy. This collaborative adjustment method enables onboard software to autonomously solve the position-keeping strategy, generate position-keeping attitude adjustment laws and programmable execution sequences under multiple constraints, greatly improving the efficiency of position-keeping operations and ensuring satellite payload safety. Attached Figure Description

[0011] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0012] Figure 1 This is a flowchart of a geostationary orbit satellite position maintenance method with unrestricted thruster installation direction provided by an embodiment of the present invention; Figure 2 This is a schematic diagram of an orbital position preservation strategy provided in an embodiment of the present invention; Figure 3 This is a flowchart of attitude-preserving adjustment under constraints provided by an embodiment of the present invention; Figure 4 This is a structural diagram of a geosynchronous orbit satellite position-keeping device with unrestricted thruster installation direction provided in an embodiment of the present invention; Figure 5 This is a hardware architecture diagram of a computer device provided in an embodiment of the present invention. Detailed Implementation

[0013] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0014] As mentioned earlier, existing satellite orbital position maintenance methods often cannot simultaneously meet the requirements of data adjustment and sunlight avoidance, which leads to a reduction in satellite payload safety.

[0015] Based on this, the concept of the present invention is to combine the real-time solar vector and the position-keeping operation requirements, and to use the back-floor thruster for position-keeping operation, which requires attitude adjustment, while also taking into account the sunlight avoidance constraints.

[0016] The following describes the specific implementation of the above concept.

[0017] Please refer to Figure 1 The present invention provides a method for position maintenance of a geosynchronous orbit satellite with unconstrained thruster installation direction, the method comprising: Step 100: Calculate the estimated orbital parameters of the satellite based on the future orbital position obtained recursively from the current orbital position; Step 102: Generate a corresponding satellite orbit position holding strategy based on the orbit parameters in the estimated orbit parameters that do not meet the preset threshold range; Step 104: Determine whether the satellite orbit position holding strategy meets the sunlight avoidance constraint and thruster layout constraint. If not, adjust the satellite orbit position holding strategy according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

[0018] In this embodiment of the invention, by predicting the satellite's orbit one week in advance and calculating its longitude, eccentricity, inclination, and corresponding drift law, a position-keeping strategy is autonomously generated for parameters exceeding error limits. Simultaneously, under constraints of sunlight avoidance and thruster layout, the position-keeping attitude is autonomously planned based on the generated strategy. This collaborative adjustment method enables onboard software to autonomously solve the position-keeping strategy, generate position-keeping attitude adjustment laws, and programmable execution sequences under multiple constraints, significantly improving the efficiency of position-keeping operations and ensuring satellite payload safety.

[0019] The following description Figure 1 The execution method of each step is shown.

[0020] First, for step 100, the estimated values ​​of the satellite's orbital parameters are calculated based on the future orbital position derived recursively from the current orbital position.

[0021] In this embodiment of the invention, the orbital parameters include the satellite's longitude, eccentricity, inclination, and drift law.

[0022] Specifically, the estimated longitude It is calculated using the following formula:

[0023] In the formula, Right ascension of the ascending node; Argument of perigee; The near-point of the horizon; ω is the orbital angular velocity.

[0024] The drift law estimate D is calculated using the following formula: The eccentricity estimate is calculated using the following formula: In the formula, The surface mass ratio of the satellite; The rate of change of tilt angle is .

[0025] Then, for step 102, a corresponding satellite orbit position holding strategy is generated based on the orbit parameters in the estimated orbit parameters that do not meet the preset threshold range.

[0026] like Figure 2As shown, firstly, it is determined whether the satellite is allowed to perform autonomous position holding. If the satellite is allowed to perform position holding, it is determined whether the longitude deviation of the satellite when it is in filter convergence and there is no unexecuted position holding strategy is greater than a preset first threshold. If so, the longitude correction is calculated based on the perturbation acceleration and longitude parameters of the fixed position. If the longitude deviation is less than the preset first threshold, it is determined whether the eccentricity of the satellite when it is in filter convergence and there is no unexecuted position holding strategy is greater than a preset second threshold. If so, the eccentricity correction is calculated based on the difference between the target eccentricity and the current eccentricity. If the eccentricity is less than the preset second threshold, it is determined whether the satellite performs east-west position holding. If so, it is ignited at perigee or apogee based on the calculated correction, and the filter convergence time is set to two days after ignition.

[0027] Furthermore, when the satellite is not allowed to maintain its position, the autonomous position-keeping flag and counter are reset to zero, and the filter convergence time is set to two days.

[0028] For step 104, determine whether the satellite orbit position holding strategy meets the sunlight avoidance constraint and thruster layout constraint. If not, adjust the satellite orbit position holding strategy according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

[0029] In this embodiment of the invention, for the generated position preservation strategy, it is first necessary to perform sunlight avoidance conflict detection. Specifically, the solar vector position, satellite orbit position and celestial attitude at the future position preservation time are predicted first. Based on the above information, the satellite flight state at the future position preservation execution time is estimated. If the autonomous position preservation strategy and the sunlight avoidance strategy conflict at any time, the execution time of the satellite orbit position preservation strategy is postponed, and the satellite orbit position preservation strategy is regenerated at the point where the position preservation efficiency of sunlight avoidance is the highest.

[0030] The deferred method can be implemented using the following code: Fire_ang = ((Fire_t )%86400) / 86400*2π; if Fire_ang>(SunAvoid.α FSLmt -SunAvoid.α SKMFSLmt ) and Fire_ang<(SunAvoid.α FSLm +SunAvoid.α SKMFSLmt ) IF2 Fire_ang>SunAvoid.α FSLmt Fire_tAddC =(SunAvoid.α SKMFSLmt - (Fire_ang - SunAvoid.α FSLmt ))*86400 / (2π); ELSE2 Fire_tAddC=-(SunAvoid.α SKMFSLmt -(SunAvoid.α FSLmt - Fire_ang) )*86400 / (2π); END2 END1 Fire_t = Fire_t + Fire_tAddC; Fire_t: is the bit-preserving ignition time; SunAvoid.α FSLmt : Sunlight avoidance starting angle; SunAvoid.α SKMFSLmt : Safety threshold angle.

[0031] Furthermore, when the satellite's orbital position-maintaining strategy is simultaneously constrained by sunlight avoidance and thruster layout, the satellite's attitude adjustment angle is calculated based on the solar vector position at the future position-maintaining moment, sunlight avoidance requirements, and thruster layout. Based on the calculation results, the satellite is then subjected to... Figure 3 The orbital attitude adjustment scheme is shown.

[0032] Specifically, it is determined whether the interval between the current time and the ignition time is less than a preset interval threshold. If so, the orbital position maintenance direction of the satellite is determined. When the satellite is in an eastward orbital position maintenance state, a first attitude adjustment scheme for the satellite is determined based on the relationship between the current time and the first maintenance time period. When the satellite is in a westward orbital position maintenance state, a second attitude adjustment scheme for the satellite is determined based on the relationship between the current time and the second maintenance time period. The satellite with the adjusted attitude is then ignited to achieve orbital control. After the orbital control is completed, the satellite is returned to normal mode to establish a ground-to-ground service attitude.

[0033] The first attitude adjustment scheme is as follows: When the satellite is in an eastward orbital position-maintaining state, if the position-maintaining time is between 18:00 and 22:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, -90, 0) in sequence; if the position-maintaining time is between 12:00 and 18:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, 90, 180) in sequence; if the position-maintaining time is between 8:00 and 12:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, -90, 180) in sequence; if the position-maintaining time is between 22:00 and 4:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, -90, 0) in sequence.

[0034] The second attitude adjustment scheme is as follows: When the satellite is in a westward orbital position-maintaining state, if the position-maintaining time is between 2:00 and 5:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, 90, 0) in sequence; if the position-maintaining time is between 6:00 and 12:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, 90, 180) in sequence; if the position-maintaining time is between 12:00 and 16:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, -90, 0) in sequence; if the position-maintaining time is between 20:00 and 2:00, the satellite's roll angle, pitch angle, and yaw angle will be adjusted to (0, -90, 180) in sequence.

[0035] In this embodiment of the invention, after adjusting the satellite's attitude using the above method, the satellite needs to be ignited to achieve orbit control. After orbit control is completed, the satellite is returned to normal mode to establish a ground-to-ground operational attitude.

[0036] It is worth noting that the method mentioned in the above embodiments has been used in the on-orbit programmable positioning and maintenance of the Yaogan-41 satellite. By utilizing the current orbital state of the satellite and spacecraft, as well as the constraints of sunlight avoidance and thruster layout, the positioning and maintenance operation strategy can be autonomously solved. This method can be directly applied to current spacecraft orbit change operations, greatly improving the autonomy of space missions and the timeliness in case of emergencies.

[0037] Please refer to Figure 4 This invention provides a geostationary orbit satellite position-keeping device with unrestricted thruster installation direction, the device comprising: The calculation module 400 is used to calculate the estimated orbital parameters of the satellite based on the future orbital position obtained recursively from the current orbital position; The generation module 402 is used to generate a corresponding satellite orbit position holding strategy based on the orbit parameters in the estimated orbit parameters that do not meet the preset threshold range; The determination module 404 is used to determine whether the satellite orbit position holding strategy meets the sunlight avoidance constraint and thruster layout constraint. If it does not meet the constraint, the satellite orbit position holding strategy is adjusted according to the solar vector position, sunlight avoidance requirement and thruster layout at the future position holding time.

[0038] In this embodiment of the invention, the orbital parameters include the satellite's longitude, eccentricity, inclination, and drift law.

[0039] In this embodiment of the invention, when the generation module 402 generates a corresponding satellite orbit position-keeping strategy based on orbit parameters that do not meet a preset threshold range in the estimated orbit parameters, it specifically performs the following operations: determining whether the satellite is allowed to perform autonomous position-keeping; if the satellite is allowed to perform position-keeping, determining whether the longitude deviation of the satellite when it is in filtered convergence and has no unexecuted position-keeping strategy is greater than a preset first threshold; if so, calculating the longitude correction amount based on the perturbation acceleration and longitude parameters of the fixed position; if the longitude deviation is less than the preset first threshold, determining whether the eccentricity of the satellite when it is in filtered convergence and has no unexecuted position-keeping strategy is greater than a preset second threshold; if so, calculating the eccentricity correction amount based on the difference between the target eccentricity and the current eccentricity; if the eccentricity is less than the preset second threshold, determining whether the satellite performs east-west position-keeping; if so, igniting at perigee or apogee based on the calculated correction amount, and setting the filtered convergence time to two days after ignition.

[0040] In this embodiment of the invention, after determining whether the satellite is allowed to perform autonomous position holding, the generation module 402 is further configured to: if the satellite is not allowed to perform position holding, then clear the autonomous position holding flag and counter to zero, and set the filtering convergence time to two days.

[0041] In this embodiment of the invention, adjusting the satellite orbital position holding strategy based on the solar vector position, sunlight avoidance requirements, and thruster layout at the future position holding time includes: when the satellite orbital position holding strategy conflicts with a preset sunlight avoidance strategy, delaying the execution time of the satellite orbital position holding strategy and regenerating the satellite orbital position holding strategy at the point of highest position holding efficiency for sunlight avoidance; when the satellite orbital position holding strategy is constrained by sunlight avoidance and thruster layout, calculating the satellite's attitude adjustment angle and adjusting the satellite's orbital attitude based on the calculation results.

[0042] In this embodiment of the invention, the step of adjusting the satellite's orbital attitude based on the calculation results includes: determining whether the interval between the current time and the ignition time is less than a preset interval threshold; if so, determining the satellite's orbital position maintenance direction; when the satellite is in an eastward orbital position maintenance state, determining a first attitude adjustment scheme for the satellite based on the relationship between the current time and the first maintenance time period; when the satellite is in a westward orbital position maintenance state, determining a second attitude adjustment scheme for the satellite based on the relationship between the current time and the second maintenance time period; igniting the satellite after attitude adjustment to achieve orbital control, and returning the satellite to normal mode after orbital control is completed to establish a ground-to-ground service attitude.

[0043] It should be noted that the geostationary orbit satellite positioning device with unrestricted thrust installation direction provided in the above embodiments is only an example illustrating the division of the above functional modules. In practical applications, the above functions can be assigned to different functional modules as needed, that is, the internal structure of the device can be divided into different functional modules to complete all or part of the functions described above. Furthermore, the geostationary orbit satellite positioning device with unrestricted thrust installation direction provided in the above embodiments and the geostationary orbit satellite positioning method embodiment with unrestricted thrust installation direction belong to the same concept. The specific implementation process is detailed in the method embodiment and will not be repeated here.

[0044] Embodiments of this application also provide a computer device, please refer to... Figure 5 The computer device includes a processor and a memory, the memory storing at least one instruction, at least one program, code set or instruction set, the at least one instruction, at least one program, code set or instruction set being loaded and executed by the processor to implement the geostationary orbit satellite position-keeping method with unconstrained thrust installation direction provided in the above-described method embodiments.

[0045] Embodiments of this application also provide a computer-readable storage medium storing at least one instruction, at least one program, code set, or instruction set, wherein the at least one instruction, at least one program, code set, or instruction set is loaded and executed by a processor to implement the geostationary orbit satellite position-keeping method with unconstrained thruster installation direction provided in the above-described method embodiments.

[0046] Embodiments of this application also provide a computer program product, which includes a computer program. A processor of a computer device reads the computer program from a computer-readable storage medium and executes the computer program, causing the computer device to perform the geostationary orbit satellite position-keeping method with unrestricted thrust installation direction as described in any of the above embodiments.

[0047] For ease of description, the above systems or devices are described separately as various modules or units based on their functions. Of course, in implementing this application, the functions of each unit can be implemented in one or more software and / or hardware components.

[0048] As can be seen from the above description of the embodiments, those skilled in the art can clearly understand that this application can be implemented by means of software plus necessary general-purpose hardware platforms. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, can be embodied in the form of a software product. This computer software product can be stored in a storage medium, such as ROM / RAM, magnetic disk, optical disk, etc., and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute the methods described in various embodiments or some parts of the embodiments of this application.

[0049] Finally, it should be noted that in this document, relational terms such as first, second, third, and fourth are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0050] The above description is only a preferred embodiment of this application. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of this application, and these improvements and modifications should also be considered within the scope of protection of this application.

Claims

1. A method for position maintenance of a geosynchronous orbit satellite with unrestricted thruster installation direction, characterized in that, The method includes: The estimated orbital parameters of the satellite are calculated based on the future orbital position derived recursively from the current orbital position. Based on the orbital parameters in the estimated orbital parameters that do not meet the preset threshold range, a corresponding satellite orbital position holding strategy is generated. Determine whether the satellite orbital position holding strategy meets the sunlight avoidance constraint and thruster layout constraint. If not, adjust the satellite orbital position holding strategy according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

2. The method as described in claim 1, characterized in that, The orbital parameters include the satellite's longitude, eccentricity, inclination, and drift law.

3. The method as described in claim 2, characterized in that, The step of generating a corresponding satellite orbit position-keeping strategy based on orbit parameters that do not meet a preset threshold range from the estimated orbit parameters includes: Determine whether the satellite is permitted to perform autonomous position maintenance; If the satellite is allowed to maintain its position, determine whether the longitude deviation of the satellite when it is in the filtering convergence state and there is no unexecuted longitude strategy is greater than a preset first threshold. If so, calculate the longitude correction amount based on the perturbation acceleration and longitude parameters of the fixed position. If the longitude deviation is less than a preset first threshold, determine whether the eccentricity of the satellite when it is in the filtering convergence and there is no unexecuted position-keeping strategy is greater than a preset second threshold. If so, calculate the eccentricity correction amount based on the difference between the target eccentricity and the current eccentricity. If the eccentricity is less than the preset second threshold, determine whether the satellite should maintain its east-west position. If so, ignite at perigee or apogee according to the calculated correction amount, and set the filtering convergence time to two days after ignition.

4. The method as described in claim 3, characterized in that, After determining whether the satellite is permitted to perform autonomous position holding, the process also includes: If the satellite is not allowed to maintain its position, then the autonomous position-keeping flag and counter will be cleared, and the filter convergence time will be set to two days.

5. The method as described in claim 2, characterized in that, The adjustment of the satellite orbital position-keeping strategy based on the solar vector position, sunlight avoidance requirements, and thruster layout at future position-keeping times includes: When the satellite orbit position holding strategy conflicts with the preset sunlight avoidance strategy, the execution time of the satellite orbit position holding strategy is delayed, and the satellite orbit position holding strategy is regenerated at the point where the position holding efficiency of sunlight avoidance is the highest. When the satellite orbital position maintenance strategy is constrained by sunlight avoidance and thruster layout, the satellite's attitude adjustment angle is calculated, and the satellite's orbital attitude is adjusted according to the calculation results.

6. The method as described in claim 5, characterized in that, The process of adjusting the satellite's orbital attitude based on the calculation results includes: Determine whether the interval between the current moment and the ignition moment is less than a preset interval threshold; if so, determine whether the satellite's orbital position maintains its orientation. When the satellite is in an eastward orbital position maintenance state, the first attitude adjustment scheme of the satellite is determined according to the relationship between the current time and the first maintenance time period. When the satellite is in a westward orbital position-holding state, the second attitude adjustment scheme of the satellite is determined according to the relationship between the current time and the second position-holding time period. The satellite, after attitude adjustment, is ignited to achieve orbit control. After orbit control is completed, the satellite is returned to normal mode to establish a ground-to-ground operational attitude.

7. A geostationary orbit satellite position-keeping device with unrestricted thruster installation direction, characterized in that, The device includes: The calculation module is used to calculate the estimated orbital parameters of the satellite based on the future orbital position derived recursively from the current orbital position. The generation module is used to generate corresponding satellite orbit position holding strategies based on the orbit parameters in the estimated orbit parameters that do not meet the preset threshold range; The determination module is used to determine whether the satellite orbital position holding strategy meets the sunlight avoidance constraints and thruster layout constraints. If it does not meet the constraints, the satellite orbital position holding strategy is adjusted according to the solar vector position, sunlight avoidance requirements and thruster layout at the future position holding time.

8. A computer device, characterized in that, The computer device includes a memory and a processor. The memory is used to store computer programs, and the processor is used to execute the computer programs stored in the memory to implement the steps of the method according to any one of claims 1-6.

9. A computer-readable storage medium, characterized in that, The storage medium stores a computer program, which, when executed by a processor, implements the steps of the method described in any one of claims 1-6.

10. A computer program product, characterized in that, Includes a computer program, which, when executed by a processor, implements the steps of the method according to any one of claims 1-6.

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