Flight control methods, devices, electronic equipment and storage media

Abstract

This application provides a flight control method, apparatus, electronic device, and storage medium, relating to the field of computer technology. The method includes: acquiring the initial orbit and target orbit of a spacecraft orbiting a celestial body, each orbit having the same perigee; the perigee being the position on the orbit closest to the celestial body; each time the spacecraft reaches the perigee, determining the spacecraft's current first orbit; if the first orbit is not the target orbit, instructing the spacecraft to accelerate once until the first orbit becomes the target orbit; when a change in the spacecraft's orbit is required, continuously accelerating the spacecraft to change its orbit, comparing the spacecraft's current velocity after each acceleration with its velocity in the target orbit, and further adjusting the spacecraft's velocity based on the difference. This allows for control of the spacecraft's orbital changes even without knowing the spacecraft's mass, until the spacecraft's current orbit is adjusted to the target orbit.

Description

Technical Field

[0001] This application relates to the field of computer technology, and more specifically, to a flight control method, apparatus, electronic device, and storage medium. Background Technology

[0002] Once a spacecraft is launched from the ground into space, it will travel to its predetermined orbit according to a pre-programmed procedure. Then, it will orbit the corresponding central celestial body according to the flight path indicated by the predetermined orbit. However, when the spacecraft receives a relevant mission, it often needs to change its current flight path in order to carry out the mission.

[0003] Currently, there are methods for changing the flight trajectory of satellites in existing technologies. These methods mainly utilize the known mass of the satellite to precisely control its orbital changes. However, for ordinary spacecraft, a considerable portion of their propulsion fuel is consumed between entering space and entering their operational orbit. Although the precise mass of the spacecraft before leaving the ground is known, the precise mass remaining after fuel consumption when it enters its operational orbit is unknown. Therefore, methods for changing the flight trajectory of satellites cannot be used on spacecraft. Summary of the Invention

[0004] The purpose of this application is to at least solve one of the aforementioned technical defects. The technical solution provided by the embodiments of this application is as follows:

[0005] In a first aspect, embodiments of this application provide a flight control method, including:

[0006] Obtain the initial orbit and target orbit of the spacecraft orbiting the star. Each orbit has the same periapsis; the periapsis is the position on the orbit closest to the star.

[0007] Whenever the spacecraft reaches its perigee, its current first orbit is determined. If the first orbit is not the target orbit, the spacecraft is instructed to accelerate until the first orbit becomes the target orbit.

[0008] The duration of the first acceleration during the initial acceleration is preset, while the duration of the second acceleration during subsequent accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time.

[0009] The speed at which the spacecraft reaches perihelion after each acceleration is determined based on the second orbit in which the spacecraft is located after acceleration.

[0010] Secondly, embodiments of this application provide a flight control device, including:

[0011] The orbit acquisition module is used to acquire the initial orbit and target orbit of the spacecraft orbiting the star. All orbits have the same periapsis; the periapsis is the position on the orbit that is closest to the star.

[0012] The acceleration indication module is used to determine the spacecraft's current first orbit whenever the spacecraft reaches the perigee. If the first orbit is not the target orbit, the module instructs the spacecraft to accelerate once until the first orbit becomes the target orbit.

[0013] The duration of the first acceleration during the initial acceleration is preset, while the duration of the second acceleration during subsequent accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time.

[0014] The speed at which the spacecraft reaches perihelion after each acceleration is determined based on the second orbit in which the spacecraft is located after acceleration.

[0015] Thirdly, embodiments of this application provide an electronic device, including a memory, a processor, and a computer program stored in the memory;

[0016] The processor executes a computer program to implement the method provided in the first aspect embodiment or any alternative embodiment of the first aspect.

[0017] Fourthly, embodiments of this application provide a computer-readable storage medium storing a computer program, which, when executed by a processor, implements the method provided in the first aspect embodiment or any optional embodiment of the first aspect.

[0018] The beneficial effects of the technical solutions provided in this application are:

[0019] The solution provided in this application embodiment allows for continuous acceleration of the spacecraft to change its orbit when a change is needed. The speed of the spacecraft in its current orbit after each acceleration is compared with its speed in the target orbit, and the speed of the spacecraft is further adjusted based on the difference. This allows for control of the spacecraft's orbital changes even when the mass of the spacecraft is unknown, until the spacecraft's orbit is adjusted to the target orbit. Attached Figure Description

[0020] To more clearly illustrate the technical solutions in the embodiments of this application, the accompanying drawings used in the description of the embodiments of this application will be briefly introduced below.

[0021] Figure 1 A flowchart illustrating a flight control method provided in an embodiment of this application;

[0022] Figure 2This is a graphical example of a spacecraft orbit switching process in one embodiment of this application;

[0023] Figure 3 This is a graphical example of any orbit of a spacecraft in one embodiment of this application;

[0024] Figure 4 This is a schematic diagram of the overall flow of a flight control method in one example of an embodiment of this application;

[0025] Figure 5 A structural block diagram of a flight control device provided in an embodiment of this application;

[0026] Figure 6 This is a schematic diagram of the structure of an electronic device provided in an embodiment of this application. Detailed Implementation

[0027] The embodiments of this application are described below with reference to the accompanying drawings. It should be understood that the embodiments described below with reference to the accompanying drawings are exemplary descriptions for explaining the technical solutions of the embodiments of this application, and do not constitute a limitation on the technical solutions of the embodiments of this application.

[0028] Those skilled in the art will understand that, unless specifically stated otherwise, the singular forms “a,” “an,” “the,” and “the” used herein may also include the plural forms. It should be further understood that the terms “comprising” and “including” as used in embodiments of this application mean that the corresponding feature can be implemented as the presented feature, information, data, step, element, and / or component, but do not exclude implementation as other features, information, data, step, element, component, and / or combinations thereof supported by the art. It should be understood that when we say that an element is “connected” or “coupled” to another element, the one element can be directly connected or coupled to the other element, or it can mean that the one element and the other element establish a connection relationship through an intermediate element. Furthermore, “connected” or “coupled” as used herein can include wireless connection or wireless coupling. The term “and / or” as used herein indicates at least one of the items defined by the term; for example, “A and / or B” can be implemented as “A,” or as “B,” or as “A and B.”

[0029] To make the objectives, technical solutions, and advantages of this application clearer, the embodiments of this application will be described in further detail below with reference to the accompanying drawings.

[0030] The technical solutions of this application and their effects are described below through several exemplary embodiments. It should be noted that the following embodiments can be referenced, borrowed from, or combined with each other. Identical terms, similar features, and similar implementation steps in different embodiments will not be repeated.

[0031] Figure 1 This is a flowchart illustrating a flight control method provided in an embodiment of this application. The subject performing this method can be a terminal (such as a computer, mobile phone, etc.). Figure 1 As shown, the method may include:

[0032] Step S101: Obtain the initial orbit and target orbit of the spacecraft orbiting the star. Each orbit has the same perihelion. The perihelion is the position on the orbit closest to the star.

[0033] In the embodiments of this application, the spacecraft can be any type of aircraft that performs specific tasks such as exploring, developing, and utilizing space and celestial bodies, such as rockets, space stations, and orbital retention platforms (which, after completing a satellite launch mission, achieve orbital retention by activating an orbital retention system and reusing rocket final stage products, enabling the orbital retention platform to fully utilize the advantages of the rocket's final stage propulsion system and perform rapid orbital maneuvers in orbit), etc. The embodiments of this application are not limited here.

[0034] A celestial body can refer to a natural object existing in outer space, such as the Earth, the Sun, and the Moon. This application does not limit the specific type of celestial body. The initial orbit can be the orbit the spacecraft was in before this series of accelerations, while the target orbit can be the orbit the spacecraft should be in after completing this series of accelerations. Orbits are generally elliptical. The periapsis can be the position on the spacecraft's orbit closest to the celestial body it orbits. For example, when the celestial body is Earth, the periapsis can be the closest point to Earth; when the celestial body is the Sun, the periapsis can be the closest point to Earth.

[0035] Specifically, before changing the orbit of a spacecraft, it is necessary to know the spacecraft's current initial orbit and the target orbit to which it needs to switch. Therefore, it is necessary to obtain the initial orbit and the target orbit for this orbit change first.

[0036] Step S102: Whenever the spacecraft reaches the perigee, determine the first orbit in which the spacecraft is currently located. If the first orbit is not the target orbit, instruct the spacecraft to accelerate once until the first orbit is the target orbit.

[0037] The duration of the first acceleration during the initial acceleration is preset, while the duration of the second acceleration during subsequent accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time.

[0038] The speed at which the spacecraft reaches perihelion after each acceleration is determined based on the second orbit in which the spacecraft is located after acceleration.

[0039] In embodiments of this application, the first orbit can be the orbit in which the spacecraft is located before each acceleration in a series of acceleration processes. The second orbit can be the orbit in which the spacecraft is located after each acceleration in some acceleration processes. The first acceleration duration can be the acceleration duration of the first acceleration in the same series of acceleration processes. The second acceleration duration can be the acceleration duration of each acceleration in the same series of acceleration processes other than the first acceleration.

[0040] Specifically, based on the Hohmann transfer orbit, it is known that accelerating the spacecraft at periapsis minimizes fuel consumption. Therefore, in this embodiment, the spacecraft is accelerated at periapsis during a series of acceleration processes. Acceleration can be achieved using pulse maneuvers (i.e., sending a signal to the spacecraft so that its plasma pulse thrusters consume a certain amount of fuel to generate instantaneous thrust, thereby achieving acceleration). Generally, given the spacecraft's mass, the velocity of the spacecraft at periapsis in each orbit can be calculated using the energy formula. Then, the acceleration can be calculated based on the required increase in velocity change. Finally, the required thrust can be calculated using Newton's second law: F (thrust) = m (spacecraft mass) / a (spacecraft acceleration). A suitable signal can then be sent to achieve precise acceleration, thus realizing precise orbital changes.

[0041] In practice, ground controllers can only know the overall mass of the spacecraft before takeoff. The spacecraft's fuel is continuously consumed during flight (including multiple processes such as entering space, entering orbit, and operating in orbit for a period of time). Although the mass of fuel consumed in a short period of time is small and negligible compared to the mass of the spacecraft, when the spacecraft operates for a long time, the continuous consumption of fuel will make the mass of fuel consumed not negligible. Therefore, it is often impossible to obtain a relatively accurate real-time mass of the spacecraft, and thus it is impossible to send accurate signals to achieve precise orbit changes.

[0042] Since the fuel consumed by the spacecraft in a short period is negligible, its mass can be considered constant. Therefore, if a signal of the same strength is sent during this period, the acceleration indicated by the signal can be assumed to be constant. Thus, this embodiment allows for multiple accelerations of the spacecraft without knowing its mass. The velocity at perihelion before and after each acceleration is determined based on the spacecraft's trajectory, and the acceleration duration is recorded. The acceleration for each acceleration is then determined based on the change in velocity at perihelion and the acceleration duration. Because each acceleration is short, the fuel consumed is considered negligible. Therefore, the acceleration of the current acceleration can be used as the acceleration for the next acceleration to calculate the required acceleration duration, and so on. This ensures a relatively accurate velocity at perihelion after each acceleration until the spacecraft switches to its target orbit.

[0043] It should be noted that, according to the Hohmann transfer orbit (e.g., Figure 2 As shown in the diagram, when a spacecraft switches orbits, it typically first switches from its initial orbit (orbit 1 in the diagram) to an intermediate orbit between the initial and final orbits (orbit 2 in the diagram), and then switches from the intermediate orbit to the final orbit (orbit 3 in the diagram). This switching process involves adding a velocity change to the spacecraft at its perigee or apogee (i.e., as shown in the diagram). The acceleration process in the embodiments of this application refers to each series of acceleration processes, specifically the process of switching from the initial track to the intermediate track or from the intermediate track to the final track, and not the entire process of switching from the initial track to the final track. (That is, during the process of switching from the initial track to the intermediate track, the initial track remains the initial track, and the intermediate track becomes the target track. The process of switching from the intermediate track to the final track is summarized, with the intermediate track considered as the initial track and the final track considered as the target track).

[0044] It should be noted that, according to the Hohmann transfer orbit, the acceleration locations for the two series of acceleration processes are different but complementary. For example, if acceleration is performed at the periapsis in the first series of acceleration processes, then acceleration will be performed at the apoapsis (the position farthest from the star in orbit) in the second series of acceleration processes. Optionally, in this embodiment, acceleration may be performed at the apoapsis first, followed by acceleration at the periapsis; this embodiment is not limited to this.

[0045] The solution provided in this application involves continuously accelerating the spacecraft to change its orbit when a change in the spacecraft's orbit is required. The velocity of the spacecraft in its current orbit after each acceleration is compared with its velocity in the target orbit, and the spacecraft's velocity is further adjusted based on the difference. This allows for control of the spacecraft's orbital changes even when the spacecraft's mass is unknown, until the spacecraft's orbit is adjusted to the target orbit.

[0046] Based on the above embodiments, as an optional embodiment, when the spacecraft passes the perigee for the second time, the duration of the second acceleration is determined in the following way:

[0047] Based on the first orbit, the initial orbit, and the target orbit, the velocities of the spacecraft when they reach the perihelion of the first orbit, the initial orbit, and the target orbit are determined accordingly.

[0048] Based on the spacecraft's velocity at perihelion when it is in the first orbit and the initial orbit, the change in the spacecraft's velocity at perihelion after the first acceleration compared to its velocity at perihelion before the first acceleration is determined as the first change.

[0049] Based on the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit respectively, the change between the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit is determined as the second change.

[0050] Based on the first change and the first acceleration duration, determine the first acceleration during the initial acceleration process;

[0051] The change in the spacecraft’s velocity when it reaches the perihelion before the second acceleration is determined based on at least one of the first change, the second change, and the first difference, and is used as the third change; the first difference is the difference between the spacecraft’s velocity when it reaches the perihelion of the target orbit and the velocity when it reaches the perihelion of the first orbit.

[0052] Based on the third change and the first acceleration, determine the second acceleration duration of the second acceleration.

[0053] Specifically, since this application requires continuous calculation of the acceleration for each acceleration during the acceleration process, and the calculation of acceleration is inseparable from the change in velocity of the spacecraft before and after acceleration, the change in velocity at the periapsis before and after acceleration can be calculated based on the spacecraft's orbit before and after acceleration. The specific calculation method can use the following formula:

[0054]

[0055] In this formula, v is the velocity at any point on the orbit, G is the gravitational constant, M is the mass of the star, and r is the straight-line distance from that point on the orbit to the star (e.g., ...). Figure 3 As shown in the figure, when the spacecraft is at its periapsis, r is the straight-line distance from the periapsis to the star, and when the spacecraft is at its apoapsis, r is the distance from the apoapsis to the star, and a is the length of the semi-major axis of the orbit.

[0056] After calculating the spacecraft's velocity at perihelion in each orbit, the acceleration during the first acceleration process is calculated. Since there is no data for the first acceleration, the acceleration information is completely unknown. Therefore, a random signal is often given during the first acceleration. Then, based on the change in the spacecraft's velocity at perihelion before and after the first acceleration (i.e., the first change) and the recorded first acceleration duration, the first acceleration during the first acceleration process is calculated.

[0057] Generally, when the second acceleration begins, the first acceleration during the initial acceleration can be used as a basis to calculate the spacecraft's velocity at perihelion after switching to the target orbit (hereinafter referred to as the target velocity). However, before the second acceleration begins, the second change needs to be compared with the first change. The purpose of this comparison is to determine if the difference between the first and second changes is too large, indicating that the acceleration time required for the second acceleration will be longer. If the acceleration is still based on the initial acceleration, the spacecraft's fuel consumption will be significant. Under the same signal conditions, the spacecraft's propulsion force is constant. According to Newton's second law, with a constant propulsion force, a decrease in spacecraft mass will lead to an increase in acceleration. In other words, during this acceleration process, the actual acceleration of the spacecraft in the latter half will be significantly greater than the first acceleration. However, the acceleration time is still calculated based on the first acceleration. Therefore, the actual velocity obtained after acceleration will be significantly greater than the target velocity, causing the spacecraft's orbit to exceed the target orbit.

[0058] Therefore, before the second acceleration begins, the second change needs to be compared with the first change. Based on the comparison result, the velocity change that the second acceleration needs to bring to the spacecraft is calculated. Specifically, if the comparison result indicates that the difference between the second and first changes is too large, the spacecraft's velocity at the periapsis needs to be increased to an appropriate value first, and then gradually accelerated to the target velocity according to the method described above. If the comparison result indicates that the difference between the second and first changes is not too large, the first acceleration can be used as a reference and gradually accelerated to the target velocity according to the method described above.

[0059] Based on the above embodiments, as an optional embodiment, the change in the spacecraft's velocity at the perihelion after the second acceleration compared to its velocity at the perihelion after the second acceleration is determined according to at least one of the first change, the second change, and the first difference. Specifically, this includes:

[0060] Determine the magnitude relationship between the second change and the first change, which is a first preset multiple;

[0061] If the second change is not less than the first change which is a first preset multiple, then the change in the spacecraft’s velocity at the point of near star after the second acceleration is determined based on the first change and the second change.

[0062] If the second change is less than the first change, which is a multiple of the first preset value, then the first difference is taken as the change in the spacecraft’s velocity at the point of near star after the second acceleration compared to the velocity at the point of near star after the second acceleration.

[0063] Specifically, generally speaking, since spacecraft often need to change speed significantly when switching orbits, and the speed change brought by the initial acceleration (i.e., the first change) is usually too small compared to the total speed change required (i.e., the second change), this embodiment compares the difference between the first and second changes by comparing the second change with the first change by a first preset multiple. Specifically, when the second change is not less than the first change by a first preset multiple, the difference between the first and second changes is considered too large; when the second change is not less than the first change by a first preset multiple, the difference between the first and second changes is considered not too large. For example, the initial velocity of the spacecraft at perihelion before the first acceleration is V1 = 100 m / s, and the velocity at perihelion after the first acceleration is V2 = 150 m / s, while the target velocity is V = 400 m / s. When the first preset multiple is set to 4, the first change... =V2-V1=50m / s, the second change =V-V1=300m / s. At this time, the second change is greater than 4 times the first change, so the difference between the first change and the second change is considered to be too large.

[0064] When the difference between the first and second changes is too large, the purpose of the second acceleration is to increase the spacecraft's speed at the perihelion to a suitable value. The second acceleration time in this process will not be too long, so the fuel consumption can be ignored, and the acceleration value in the first acceleration process can be used as the basis for calculation. When the difference between the first and second changes is not too large, the purpose of each acceleration after the second acceleration begins is to adjust the spacecraft's speed at the perihelion to the target speed. At this time, the first difference is directly used as the speed change that the second acceleration needs to bring to the spacecraft at the perihelion.

[0065] Based on the above embodiments, as an optional embodiment, the change in the spacecraft's velocity at the perihelion after the second acceleration compared to its velocity at the perihelion after the second acceleration is determined according to the first change and the second change, specifically including:

[0066] The second difference between the second change and the first change of the second preset multiple is used as the change in the spacecraft's velocity when it reaches the perihelion after the second acceleration compared to the velocity when it reaches the perihelion after the second acceleration. The second preset multiple is determined based on the first preset multiple.

[0067] In this embodiment of the application, the second preset multiple can be obtained by subtracting one from the first preset multiple.

[0068] Specifically, when the difference between the first and second changes is too large, the purpose of the second acceleration is to first increase the spacecraft's velocity at perihelion to an appropriate value. Extensive experiments have shown that this appropriate value can be determined by calculating the second difference between the second change and the first change multiplied by a second preset factor. Using the example mentioned earlier, the first preset factor is 4, and based on the first preset factor, the second preset factor is determined to be 3. =50m / s, second change =300m / s, at which point the second difference is... -3 If V1 = 150 m / s, then the second acceleration requires increasing the spacecraft's velocity by 150 m / s at the perihelion. During the second acceleration, the first acceleration from the first acceleration phase will be used as a reference for the duration of the second acceleration. For example, if the duration of the first acceleration, t1, is 10 s, then the first acceleration a1 = (V2 - V1) / t1 = 5 m / s². 2 Therefore, the second acceleration time of the second acceleration process is the second difference / a1=30s.

[0069] Based on the above embodiments, as an optional embodiment, starting from the spacecraft's third pass near the star point, the second acceleration duration for each acceleration is determined in the following manner:

[0070] Obtain the third orbit where the spacecraft was before the last acceleration and the second acceleration duration of the last acceleration, and determine the speed of the spacecraft when it reaches the perihelion of the third orbit based on the third orbit;

[0071] Based on the spacecraft's velocity when it reaches the perihelion of the first and third orbits respectively, the change in the spacecraft's velocity when it reaches the perihelion after the last acceleration compared to its velocity when it reaches the perihelion before the last acceleration is determined as the third change.

[0072] Based on the spacecraft's speed when it reaches the perihelion of the target orbit and the first orbit respectively, the change in the spacecraft's speed when it reaches the perihelion after this acceleration compared to its speed when it reaches the perihelion before this acceleration is determined as the fourth change.

[0073] The second acceleration duration for this acceleration is determined based on the third change, the second acceleration duration of the previous acceleration, and the fourth change.

[0074] Specifically, after the second acceleration is completed, the spacecraft's velocity at the periapsis has reached an appropriate value. Subsequent calculations can determine the duration of the second acceleration for each acceleration process based on the target velocity and the spacecraft's current velocity at the periapsis. Specifically, starting from the spacecraft's third passage through the periapsis, the duration of each acceleration will be determined based on the acceleration details of the previous acceleration process and the fourth change in velocity at the periapsis that this acceleration needs to bring to the spacecraft. The acceleration details of the previous acceleration process include the third change in velocity at the periapsis brought about by the previous acceleration process and the duration of the second acceleration process. Based on this information, the duration of the second acceleration required for this acceleration can be determined.

[0075] Optionally, in practice, to reduce computational load, the third acceleration can be performed similarly to the first. Each time the spacecraft reaches its periapsis, the same signal is sent to it. After each acceleration, the spacecraft's second orbit is obtained, and its current velocity at the periapsis is determined based on this orbit. This velocity is then compared to the target velocity. If the difference between the two velocities is less than a certain multiple (e.g., 2 times), the acceleration value from the previous acceleration process is determined based on the acceleration duration and the velocity change it brought to the spacecraft at the periapsis. The duration of the next acceleration is then calculated based on this acceleration value. Compared to the method described above, this method, except for the final acceleration, does not require information from the previous acceleration, thus reducing computational load.

[0076] Based on the above embodiments, as an optional embodiment, the second acceleration duration of the current acceleration is determined according to the third change amount, the second acceleration duration of the previous acceleration, and the fourth change amount, specifically including:

[0077] The second acceleration during the previous acceleration process is determined based on the third change and the second acceleration duration of the previous acceleration. The second acceleration duration of the current acceleration is determined based on the second acceleration and the fourth change.

[0078] Specifically, since the acceleration referenced from the third acceleration onwards is the second acceleration from the previous acceleration, the second acceleration from the previous acceleration can be calculated first. This second acceleration can be calculated based on the change in the spacecraft's velocity at perihelion before and after the previous acceleration (i.e., the fourth change) and the recorded duration of the second acceleration in the previous acceleration. After obtaining the second acceleration from the previous acceleration, it can be used as the spacecraft's acceleration in this acceleration process, and combined with the fourth change, the required second acceleration duration for this acceleration process can be calculated.

[0079] Based on the above embodiments, as an optional embodiment, the spacecraft is instructed to perform an acceleration, specifically including:

[0080] Based on the acceleration duration, a corresponding acceleration command is generated and sent to the spacecraft so that the spacecraft can accelerate once according to the acceleration command.

[0081] In embodiments of this application, the acceleration command can be used to instruct the spacecraft to perform a pulse maneuver.

[0082] Specifically, since most spacecraft are unmanned, the process of instructing a spacecraft to accelerate needs to be conducted at a ground-based spacecraft control center. The control center can observe in real time the shape of the spacecraft's orbit and its relative position to other celestial bodies, satellites, and other spacecraft. It can also send instructions to the spacecraft to perform tasks or adjust its orbit. Therefore, in this embodiment, an acceleration command can be sent to the spacecraft to instruct it to perform corresponding pulse maneuvers, thereby achieving acceleration at its periapsis. It is understood that each acceleration command includes information such as thrust and acceleration duration. After receiving the acceleration command, the spacecraft can analyze this information and, based on this information, cause the plasma pulse thruster to generate the corresponding thrust within the corresponding acceleration duration.

[0083] The following is combined Figure 4 The overall process of the flight control method provided in the embodiments of this application will be introduced, such as... Figure 4As shown, when a spacecraft needs to make orbital adjustments, it is first instructed to perform an initial acceleration. After the initial acceleration is completed, its velocity at the perihelion can be calculated based on the spacecraft's orbit before and after acceleration, and the duration of the initial acceleration can be recorded.

[0084] Next, it is necessary to determine how to perform the second acceleration. If the difference between the first change (i.e., the spacecraft's current velocity at perihelion and its velocity at perihelion in the target orbit) and the second change (i.e., the difference between the spacecraft's velocity at perihelion in the initial orbit and its velocity at perihelion in the target orbit) is too large, the velocity change brought to the spacecraft by the second acceleration will be further determined based on the first and second changes. Then, the second acceleration duration of the second acceleration will be calculated in conjunction with the first acceleration of the first acceleration process. If the difference between the first and second changes is not too large, the velocity change brought to the spacecraft by the second acceleration will be determined based on the spacecraft's current velocity at perihelion and its velocity at perihelion when in the target orbit. Then, the second acceleration duration of the second acceleration will be calculated in conjunction with the first acceleration of the first acceleration process.

[0085] For the acceleration process after the third acceleration, firstly, the second acceleration of the previous acceleration process is determined based on the change in the spacecraft's velocity at the periapsis before and after the previous acceleration process and the second acceleration duration of the previous acceleration process. Then, the velocity change brought to the spacecraft by this acceleration is determined based on the spacecraft's current velocity at the periapsis and the spacecraft's velocity at the periapsis when it is in the target orbit. Finally, the second acceleration duration of this acceleration is calculated by combining the second acceleration of the previous acceleration process.

[0086] For each subsequent acceleration, after each acceleration, it will be determined whether the spacecraft's current orbit is the target orbit. If it is, it means that the spacecraft's current series of orbit adjustment operations has been completed. If not, the acceleration process will be repeated until the spacecraft's current orbit is the target orbit.

[0087] Figure 5 A structural block diagram of a flight control device provided in an embodiment of this application is shown below. Figure 5 As shown, the flight control device 500 may include: an orbit acquisition module 501 and an acceleration indication module 502, wherein,

[0088] The orbit acquisition module 501 is used to acquire the initial orbit and target orbit of the spacecraft orbiting the star, and each orbit has the same perihelion; the perihelion is the position on the orbit that is closest to the star.

[0089] The acceleration indication module 502 is used to determine the first orbit currently occupied by the spacecraft whenever the spacecraft reaches the perigee. If the first orbit is not the target orbit, the spacecraft is instructed to accelerate once until the first orbit becomes the target orbit.

[0090] The first acceleration duration for the first acceleration is preset, while the second acceleration duration for non-first accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time.

[0091] The speed at which the spacecraft reaches the periapsis after each acceleration is determined based on the second orbit in which the spacecraft is located after the acceleration.

[0092] The solution provided in this application embodiment allows for continuous acceleration of the spacecraft to change its orbit when a change is needed. The speed of the spacecraft in its current orbit after each acceleration is compared with its speed in the target orbit, and the speed of the spacecraft is further adjusted based on the difference. This allows for control of the spacecraft's orbital changes even when the mass of the spacecraft is unknown, until the spacecraft's orbit is adjusted to the target orbit.

[0093] Based on the above embodiments, as an optional embodiment, the device further includes an acceleration duration determination module, specifically used for:

[0094] Based on the first orbit, the initial orbit, and the target orbit, the velocities of the spacecraft when they reach the perihelion of the first orbit, the initial orbit, and the target orbit are determined accordingly.

[0095] Based on the spacecraft's velocity at perihelion when it is in the first orbit and the initial orbit, the change in the spacecraft's velocity at perihelion after the first acceleration compared to its velocity at perihelion before the first acceleration is determined as the first change.

[0096] Based on the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit respectively, the change between the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit is determined as the second change.

[0097] Based on the first change and the first acceleration duration, determine the first acceleration during the initial acceleration process;

[0098] The change in the spacecraft’s velocity when it reaches the perihelion before the second acceleration is determined based on at least one of the first change, the second change, and the first difference, and is used as the third change; the first difference is the difference between the spacecraft’s velocity when it reaches the perihelion of the target orbit and the velocity when it reaches the perihelion of the first orbit.

[0099] Based on the third change and the first acceleration, determine the second acceleration duration of the second acceleration.

[0100] Based on the above embodiments, as an optional embodiment, the acceleration duration determination module is further configured to:

[0101] Determine the magnitude relationship between the second change and the first change, which is a first preset multiple;

[0102] If the second change is not less than the first change which is a first preset multiple, then the change in the spacecraft’s velocity at the point of near star after the second acceleration is determined based on the first change and the second change.

[0103] If the second change is less than the first change, which is a multiple of the first preset value, then the first difference is taken as the change in the spacecraft’s velocity at the point of near star after the second acceleration compared to the velocity at the point of near star after the second acceleration.

[0104] Based on the above embodiments, as an optional embodiment, the acceleration time determination module can also be used for:

[0105] The second difference between the second change and the first change of the second preset multiple is used as the change in the spacecraft's velocity when it reaches the perihelion after the second acceleration compared to the velocity when it reaches the perihelion after the second acceleration. The second preset multiple is determined based on the first preset multiple.

[0106] Based on the above embodiments, as an optional embodiment, the acceleration time determination module can also be used for:

[0107] Obtain the third orbit where the spacecraft was before the last acceleration and the second acceleration duration of the last acceleration, and determine the speed of the spacecraft when it reaches the perihelion of the third orbit based on the third orbit;

[0108] Based on the spacecraft's velocity when it reaches the perihelion of the first and third orbits respectively, the change in the spacecraft's velocity when it reaches the perihelion after the last acceleration compared to its velocity when it reaches the perihelion before the last acceleration is determined as the third change.

[0109] Based on the spacecraft's speed when it reaches the perihelion of the target orbit and the first orbit respectively, the change in the spacecraft's speed when it reaches the perihelion after this acceleration compared to its speed when it reaches the perihelion before this acceleration is determined as the fourth change.

[0110] The second acceleration duration for this acceleration is determined based on the third change, the second acceleration duration of the previous acceleration, and the fourth change.

[0111] Based on the above embodiments, as an optional embodiment, the acceleration time determination module can also be used for:

[0112] The second acceleration during the previous acceleration process is determined based on the third change and the second acceleration duration of the previous acceleration. The second acceleration duration of the current acceleration is determined based on the second acceleration and the fourth change.

[0113] Based on the above embodiments, as an optional embodiment, the device further includes an acceleration command sending module, specifically used for:

[0114] Based on the acceleration duration, a corresponding acceleration command is generated and sent to the spacecraft so that the spacecraft can accelerate once according to the acceleration command.

[0115] The following is for reference. Figure 6 It illustrates an electronic device suitable for implementing embodiments of this application (e.g., performing...). Figure 1 The diagram shows the structure of the terminal device or server 600 of the method shown. The electronic devices in the embodiments of this application may include, but are not limited to, mobile terminals such as mobile phones, laptops, digital broadcast receivers, PDAs (personal digital assistants), PADs (tablet computers), PMPs (portable multimedia players), vehicle terminals (e.g., vehicle navigation terminals), wearable devices, etc., as well as fixed terminals such as digital TVs, desktop computers, etc. Figure 6 The electronic device shown is merely an example and should not impose any limitation on the functionality and scope of use of the embodiments of this application.

[0116] The electronic device includes a memory and a processor. The memory stores a program for performing the methods described in the various method embodiments above; the processor is configured to perform the program stored in the memory. The processor may be referred to as processing device 601 as described below, and the memory may include at least one of read-only memory (ROM) 602, random access memory (RAM) 603, and storage device 608 as described below, as specifically shown below:

[0117] like Figure 6As shown, electronic device 600 may include a processing device (e.g., a central processing unit, a graphics processor, etc.) 601, which can perform various appropriate actions and processes according to a program stored in read-only memory (ROM) 602 or a program loaded from storage device 608 into random access memory (RAM) 603. RAM 603 also stores various programs and data required by electronic device 600. Processing device 601, ROM 602, and RAM 603 are interconnected via bus 604. Input / output (I / O) interface 605 is also connected to bus 604.

[0118] Typically, the following devices can be connected to I / O interface 605: input devices 606 including, for example, touchscreens, touchpads, keyboards, mice, cameras, microphones, accelerometers, gyroscopes, etc.; output devices 607 including, for example, liquid crystal displays (LCDs), speakers, vibrators, etc.; storage devices 608 including, for example, magnetic tapes, hard disks, etc.; and communication devices 609. Communication device 609 allows electronic device 600 to communicate wirelessly or wiredly with other devices to exchange data. Although Figure 6 Electronic devices with various devices are shown, but it should be understood that it is not required to implement or have all of the devices shown. More or fewer devices may be implemented or have alternatively.

[0119] Specifically, according to embodiments of this application, the processes described above with reference to the flowcharts can be implemented as computer software programs. For example, embodiments of this application include a computer program product comprising a computer program carried on a non-transitory computer-readable medium, the computer program containing program code for performing the methods shown in the flowcharts. In such embodiments, the computer program can be downloaded and installed from a network via communication device 609, or installed from storage device 608, or installed from ROM 602. When the computer program is processed by processing device 601, the functions defined above in the methods of embodiments of this application are performed.

[0120] It should be noted that the computer-readable storage medium described above in this application can be a computer-readable signal medium, a computer-readable storage medium, or any combination of the two. A computer-readable storage medium can be, for example,—but not limited to—an electrical, magnetic, optical, electromagnetic, infrared, or semiconductor system, apparatus, or device, or any combination thereof. More specific examples of a computer-readable storage medium may include, but are not limited to: an electrical connection having one or more wires, a portable computer disk, a hard disk, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fiber, portable compact disk read-only memory (CD-ROM), optical storage device, magnetic storage device, or any suitable combination thereof. In this application, a computer-readable storage medium can be any tangible medium containing or storing a program that can be instructed to be used by or in connection with a system, apparatus, or device. In this application, a computer-readable signal medium can include a data signal propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals can take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. A computer-readable signal medium can be any computer-readable medium other than a computer-readable storage medium, which can send, propagate, or transmit a program for use by or in connection with a system, apparatus, or device by instructions. The program code contained on the computer-readable medium can be transmitted using any suitable medium, including but not limited to: wires, optical fibers, RF (radio frequency), etc., or any suitable combination thereof.

[0121] In some implementations, clients and servers can communicate using any currently known or future-developed network protocol such as HTTP (Hypertext Transfer Protocol) and can interconnect with digital data communication (e.g., communication networks) of any form or medium. Examples of communication networks include local area networks (“LANs”), wide area networks (“WANs”), the Internet (e.g., the Internet of Things), and peer-to-peer networks (e.g., ad hoc peer-to-peer networks), as well as any currently known or future-developed networks.

[0122] The aforementioned computer-readable medium may be included in the aforementioned electronic device; or it may exist independently and not assembled into the electronic device.

[0123] The aforementioned computer-readable medium carries one or more programs that, when executed by the electronic device, cause the electronic device to:

[0124] The system acquires the spacecraft's initial orbit and target orbit around the star, with each orbit sharing the same perigee. The perigee is the closest point on the orbit to the star. Whenever the spacecraft reaches the perigee, its current first orbit is determined. If the first orbit is not the target orbit, the spacecraft is instructed to accelerate until the first orbit becomes the target orbit. The duration of the first acceleration is preset, while the duration of the second acceleration (for subsequent accelerations) is determined based on the change between the spacecraft's velocity at its previous perigee and its velocity at the current perigee. The spacecraft's velocity at the perigee after each acceleration is determined based on its new second orbit.

[0125] Computer program code for carrying out this application can be written in one or more programming languages ​​or a combination thereof, including but not limited to object-oriented programming languages ​​such as Java, Smalltalk, and C++, as well as conventional procedural programming languages ​​such as "C" or similar programming languages. The program code can be executed entirely on the user's computer, partially on the user's computer, as a standalone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In cases involving remote computers, the remote computer can be connected to the user's computer via any type of network—including a local area network (LAN) or a wide area network (WAN)—or can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0126] The flowcharts and block diagrams in the accompanying drawings illustrate the architecture, functionality, and possible implementations of systems, methods, and computer program products according to various embodiments of this application. In this regard, each block in a flowchart or block diagram may represent a module, segment, or portion of code containing one or more executable instructions for implementing the specified logical function. It should also be noted that in some alternative implementations, the functions indicated in the blocks may occur in a different order than those indicated in the drawings. For example, two consecutively indicated blocks may actually proceed substantially in parallel, and they may sometimes proceed in reverse order, depending on the functions involved. It should also be noted that each block in the block diagrams and / or flowcharts, and combinations of blocks in the block diagrams and / or flowcharts, may be implemented using a dedicated hardware-based system to perform the specified function, or using a combination of dedicated hardware and computer instructions.

[0127] The modules or units described in the embodiments of this application can be implemented in software or hardware. The names of modules or units do not necessarily limit the specific unit; for example, a first constraint acquisition module can also be described as a "module for acquiring the first constraint".

[0128] The functions described above in this document can be performed at least in part by one or more hardware logic components. For example, exemplary types of hardware logic components that can be used, without limitation, include: Field Programmable Gate Arrays (FPGAs), Application-Specific Integrated Circuits (ASICs), Application Standard Products (ASSPs), System-on-Chip (SoCs), Complex Programmable Logic Devices (CPLDs), and so on.

[0129] In the context of this application, a machine-readable medium can be a tangible medium that may contain or store a program for use by or in conjunction with instructions to a system, apparatus, or device. A machine-readable medium can be a machine-readable signal medium or a machine-readable storage medium. A machine-readable medium can be, but is not limited to, electronic, magnetic, optical, electromagnetic, infrared, or semiconductor systems, apparatus, or devices, or any suitable combination of the foregoing. More specific examples of machine-readable storage media include, based on electrical connections of one or more wires, portable computer disks, hard disks, random access memory (RAM), read-only memory (ROM), erasable programmable read-only memory (EPROM or flash memory), optical fibers, portable compact disc read-only memory (CD-ROM), optical storage devices, magnetic storage devices, or any suitable combination of the foregoing.

[0130] It should be understood that although the steps in the flowcharts of the accompanying figures are shown sequentially as indicated by the arrows, these steps are not necessarily performed in the order indicated by the arrows. Unless explicitly stated herein, there is no strict order constraint on the execution of these steps, and they can be performed in other orders. Moreover, at least some steps in the flowcharts of the accompanying figures may include multiple sub-steps or multiple stages. These sub-steps or stages are not necessarily completed at the same time, but can be performed at different times, and their order of execution is not necessarily sequential, but can be alternated or performed in turn with other steps or at least some of the sub-steps or stages of other steps.

[0131] The above description is only a partial embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the principle of the present invention, and these improvements and modifications should also be considered within the scope of protection of the present invention.

Claims

1. A flight control method, characterized in that, include: Obtain the initial orbit and target orbit of the spacecraft orbiting the star, with each orbit having the same perigee; The nearest point to the star is the position on the orbit closest to the star. Whenever the spacecraft reaches the perigee, the first orbit in which the spacecraft is currently located is determined. If the first orbit is not the target orbit, the spacecraft is instructed to accelerate once until the first orbit is the target orbit. The first acceleration duration for the first acceleration is preset, while the second acceleration duration for non-first accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time. The speed at which the spacecraft reaches the perigee after each acceleration is determined based on the second orbit in which the spacecraft is located after the acceleration; The duration of the second acceleration during the second passage of the spacecraft past the perigee is determined in the following manner: Based on the first orbit, the initial orbit, and the target orbit, the velocities of the spacecraft when they reach the perihelion of the first orbit, the initial orbit, and the target orbit are determined accordingly. Based on the speed of the spacecraft when it reaches the periapsis in the first orbit and the initial orbit, the change in the speed of the spacecraft when it reaches the periapsis after the first acceleration compared to the speed when it reaches the periapsis before the first acceleration is determined as the first change. Based on the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit, respectively, the change between the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit is determined as a second change. Based on the first change and the first acceleration duration, determine the first acceleration during the initial acceleration process; The change in the spacecraft’s velocity when it reaches the perihelion before the second acceleration compared to its velocity when it reaches the perihelion after the second acceleration is determined based on at least one of the first change, the second change, and the first difference, and is taken as the third change; the first difference is the difference between the spacecraft’s velocity when it reaches the perihelion of the target orbit and its velocity when it reaches the perihelion of the first orbit. The second acceleration duration of the second acceleration is determined based on the third change and the first acceleration.

2. The method according to claim 1, characterized in that, Determining the change in the spacecraft's velocity at perihelion after the second acceleration compared to its velocity at perihelion after the second acceleration, based on at least one of the first change, the second change, and the first difference, includes: Determine the magnitude relationship between the second change and the first change by a first preset multiple; If the second change is not less than the first change by the first preset multiple, then the change in the spacecraft’s velocity when it reaches the perihelion after the second acceleration is determined based on the first change and the second change. If the second change is less than the first change by the first preset multiple, then the first difference is taken as the change in the spacecraft's velocity at perihelion after the second acceleration compared to its velocity at perihelion after the second acceleration.

3. The method according to claim 2, characterized in that, The step of determining the change in the spacecraft's velocity at perihelion after the second acceleration compared to its velocity at perihelion after the second acceleration, based on the first and second changes, includes: The second difference between the second change and the first change by a second preset multiple is taken as the change in the spacecraft's velocity at perihelion after the second acceleration compared to the velocity at perihelion after the second acceleration. The second preset multiple is determined based on the first preset multiple.

4. The method according to claim 1, characterized in that, Starting from the third pass of the spacecraft near the star point, the duration of the second acceleration for each acceleration is determined in the following manner: Obtain the third orbit where the spacecraft was located before the last acceleration and the second acceleration duration of the last acceleration, and determine the speed of the spacecraft when it reaches the perihelion of the third orbit based on the third orbit; Based on the velocities of the spacecraft when it reaches the perihelion of the first orbit and the third orbit respectively, the change in the spacecraft's velocity when it reaches the perihelion after the last acceleration compared to its velocity when it reaches the perihelion before the last acceleration is determined as the third change. Based on the speeds of the spacecraft when it reaches the perihelion of the target orbit and the first orbit respectively, the change in the speed of the spacecraft when it reaches the perihelion after this acceleration compared to the speed when it reaches the perihelion before this acceleration is determined as the fourth change. The second acceleration duration for this acceleration is determined based on the third change, the second acceleration duration of the previous acceleration, and the fourth change.

5. The method according to claim 4, characterized in that, Determining the second acceleration duration for the current acceleration based on the third change, the second acceleration duration of the previous acceleration, and the fourth change includes: The second acceleration during the previous acceleration process is determined based on the third change and the second acceleration duration of the previous acceleration, and the second acceleration duration of the current acceleration is determined based on the second acceleration and the fourth change.

6. The method according to claim 1, characterized in that, The instruction to the spacecraft to perform an acceleration includes: Based on the acceleration duration, a corresponding acceleration command is generated and sent to the spacecraft so that the spacecraft can accelerate once according to the acceleration command.

7. A flight control device, characterized in that, include: The orbit acquisition module is used to acquire the initial orbit and target orbit of the spacecraft orbiting the star. All orbits have the same perigee. The nearest point to the star is the position on the orbit closest to the star. An acceleration indication module is used to determine the first orbit currently occupied by the spacecraft whenever the spacecraft reaches the perigee. If the first orbit is not the target orbit, the module instructs the spacecraft to accelerate once until the first orbit becomes the target orbit. The first acceleration duration for the first acceleration is preset, while the second acceleration duration for non-first accelerations is determined based on the change between the spacecraft's velocity when it last reached the perihelion and its velocity when it reaches the perihelion this time. The speed at which the spacecraft reaches the perigee after each acceleration is determined based on the second orbit in which the spacecraft is located after the acceleration; The acceleration duration determination module is used to determine the speed of the spacecraft when it reaches the perihelion of the first orbit, the initial orbit, and the target orbit, respectively, based on the first orbit, the initial orbit, and the target orbit. Based on the speed of the spacecraft when it reaches the periapsis in the first orbit and the initial orbit, the change in the speed of the spacecraft when it reaches the periapsis after the first acceleration compared to the speed when it reaches the periapsis before the first acceleration is determined as the first change. Based on the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit, respectively, the change between the velocities of the spacecraft when it reaches the perihelion of the initial orbit and the target orbit is determined as a second change. Based on the first change and the first acceleration duration, determine the first acceleration during the initial acceleration process; The change in the spacecraft’s velocity when it reaches the perihelion before the second acceleration compared to its velocity when it reaches the perihelion after the second acceleration is determined based on at least one of the first change, the second change, and the first difference, and is taken as the third change; the first difference is the difference between the spacecraft’s velocity when it reaches the perihelion of the target orbit and its velocity when it reaches the perihelion of the first orbit. The second acceleration duration of the second acceleration is determined based on the third change and the first acceleration.

8. An electronic device comprising a memory, a processor, and a computer program stored in the memory, characterized in that, The processor executes the computer program to implement the method of any one of claims 1-6.

9. A computer-readable storage medium having a computer program stored thereon, characterized in that, When the computer program is executed by the processor, it implements the method of any one of claims 1-6.

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