A method and device for determining the deviation of the initial phase angle of a super-fast rendezvous and docking orbit injection

Abstract

The present invention provides a method and device for determining the initial phase angle deviation of a super-fast rendezvous and docking orbit insertion. The method includes: obtaining the guidance parameters of the chaser spacecraft during the two-pulse guidance at a long distance; determining the shortest flight time of the chaser spacecraft during the long-distance guidance section according to the preset engine startup time and guidance parameters; determining the longest flight time of the chaser spacecraft during the long-distance guidance section according to the preset rendezvous and docking time and the short-range flight time of the chaser spacecraft during the close-range autonomous control section; and determining the initial phase angle deviation of the chaser spacecraft at the time of orbit insertion according to the shortest flight time and the longest flight time. The method for determining the initial phase angle deviation of the super-fast rendezvous and docking orbit insertion provided by this solution can determine the initial phase angle deviation of the chaser spacecraft at the time of orbit insertion, and can ensure the safety and robustness of the long-distance guidance section.

Description

Technical Field

[0001] The present invention relates to the technical field of rendezvous and docking, and particularly relates to a method and device for determining the initial phase angle deviation of a super-fast rendezvous and docking during orbit injection. Background Art

[0002] The rendezvous and docking strategy is the way to transport astronauts from the ground to the International Space Station. This method has been adopted by the Soyuz spacecraft, the Space Shuttle with the International Space Station, and the Shenzhou spacecraft. The traditional rendezvous and docking plan takes 2 - 3 days, and the super-fast rendezvous and docking with a shorter time has become the current development trend.

[0003] The traditional rendezvous and docking plan includes a long-distance guidance section, a close-range autonomous control section, and a docking section. For a 2-hour super-fast docking, the flight times of these three stages have very strict requirements. To reduce the total flight time, the flight time of the long-distance guidance section must also be reduced, so the number of its executed pulses cannot be many. And the long-distance guidance section must guide the tracking spacecraft to a position close enough to the target spacecraft so that the rendezvous and docking can be quickly completed after entering the autonomous control section. Thus, the safety of the long-distance guidance section must be ensured. In addition, due to fewer pulses, the initial phase angle suitable for orbit injection is relatively small. Therefore, sufficient analysis must be carried out to solve the problems of ensuring the safety and robustness of the long-distance guidance section under the condition of ultra-short flight time. Summary of the Invention

[0004] The embodiments of the present invention provide a method and device for determining the initial phase angle deviation of a super-fast rendezvous and docking during orbit injection. This method can determine the initial phase angle deviation of the tracking spacecraft during orbit injection and ensure the safety and robustness of the long-distance guidance section.

[0005] In a first aspect, the embodiments of the present invention provide a method for determining the initial phase angle deviation of a super-fast rendezvous and docking during orbit injection, including:

[0006] Obtaining the guidance parameters of the tracking spacecraft under two-pulse guidance at a long distance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration when guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration when guiding the second pulse;

[0007] Determining the shortest flight time of the tracking spacecraft in the long-distance guidance section according to the preset engine start time and the guidance parameters;

[0008] Determining the longest flight time of the tracking spacecraft in the long-distance guidance section according to the preset rendezvous and docking time and the short-range flight time of the tracking spacecraft in the close-range autonomous control section;

[0009] Determining the initial phase angle deviation of the tracking spacecraft during orbit injection according to the shortest flight time and the longest flight time.

[0010] Optionally, determining the shortest flight time of the tracking spacecraft in the long-distance guidance section according to the preset engine start-up time and the guidance parameters includes:

[0011] Determining the first pulse time of the first pulse according to the preset engine start-up time, the first velocity increment, and the first acceleration;

[0012] Performing two-pulse guidance solution using the quasi-Newton method according to the first pulse time, the first velocity increment, and the second velocity increment to obtain the second pulse time of the second pulse;

[0013] Calculating the shortest flight time according to the second pulse time, the second velocity increment, and the second acceleration.

[0014] Optionally, the first pulse time is determined by the following formula:

[0015]

[0016] wherein, the t 1min is used to represent the first pulse time; the V p1 is used to represent the first velocity increment; the a1 is used to represent the first acceleration; the preset engine start-up time is 900 s;

[0017] The shortest flight time is determined by the following formula:

[0018]

[0019] wherein, the t yuan_min is used to represent the shortest flight time; the t 2min is used to represent the second pulse time; the V p2 is used to represent the second velocity increment; the a2 is used to represent the second acceleration.

[0020] Optionally, determining the longest flight time of the tracking spacecraft in the long-distance guidance section includes:

[0021] Performing a difference operation on the preset rendezvous and docking time and the short-range flight time to obtain the longest flight time.

[0022] Optionally, after determining the longest flight time of the tracking spacecraft in the long-distance guidance section, it further includes:

[0023] Calculating the latest start-up time of the second pulse according to the longest flight time, the second velocity increment, and the second acceleration;

[0024] The latest startup time is determined by the following formula:

[0025]

[0026] wherein, the t 2max is used to represent the latest startup time; the t yuan_max is used to represent the longest flight time; the V p2 is used to represent the second speed increment; the a2 is used to represent the second acceleration.

[0027] Optionally, determining the initial phase angle deviation of the tracking spacecraft at the time of orbit injection includes:

[0028] Performing a difference operation on the longest flight time and the shortest flight time to obtain a fluctuation time;

[0029] Calculating the initial phase angle deviation based on the fluctuation time;

[0030] The initial phase angle deviation is determined by the following formula:

[0031]

[0032] wherein, the θ bodong is used to represent the initial phase angle deviation; the t yuan_max is used to represent the longest flight time; the t yuan_min is used to represent the shortest flight time.

[0033] Optionally, after determining the initial phase angle deviation of the tracking spacecraft at the time of orbit injection, it further includes:

[0034] S1: Judging whether the guidance parameters are the same as the guidance parameters pre-designed for the two-pulse guidance at a long distance; if so, execute step S2, if not, terminate the ultra-fast rendezvous and docking mission;

[0035] S2: Obtaining the third speed increment executed by the first pulse in the two-pulse guidance at a long distance;

[0036] S3: Judging whether the absolute error increment between the third speed increment and the first nominal calculated pulse of the first pulse is greater than the first error threshold; if not, execute step S4, if so, exit the ultra-fast rendezvous and docking mission;

[0037] S4: The tracking spacecraft executes the second pulse and obtains the fourth speed increment executed by the second pulse in the two-pulse guidance at a long distance;

[0038] S5: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is greater than a second error threshold; if not, execute step S6, if so, exit the ultra-fast rendezvous and docking mission;

[0039] S6: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is less than a third error threshold; wherein, the second error threshold is greater than the third error threshold; if so, execute step S7; if not, execute step S8;

[0040] S7: Continue to execute the ultra-fast rendezvous and docking mission;

[0041] S8: The tracking spacecraft adopts an active braking strategy to ensure the safety of the ultra-fast rendezvous and docking mission.

[0042] In a second aspect, an embodiment of the present invention further provides an apparatus for determining the initial phase angle deviation of ultra-fast rendezvous and docking into orbit, including:

[0043] An acquisition module, configured to acquire the guidance parameters of the tracking spacecraft in the long-distance two-pulse guidance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration when guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration when guiding the second pulse;

[0044] A flight time determination module, configured to determine the shortest flight time of the tracking spacecraft in the long-distance guidance section according to a preset engine startup time and the guidance parameters; and determine the longest flight time of the tracking spacecraft in the long-distance guidance section according to a preset rendezvous and docking time and the short-range flight time of the tracking spacecraft in the short-distance autonomous control section;

[0045] A phase angle deviation determination module, which determines the initial phase angle deviation of the tracking spacecraft when entering orbit according to the shortest flight time and the longest flight time.

[0046] Optionally, the phase angle deviation determination module is further configured to perform the following operations:

[0047] Perform a difference operation on the longest flight time and the shortest flight time to obtain a fluctuation time;

[0048] Calculate the initial phase angle deviation according to the fluctuation time;

[0049] The initial phase angle deviation is determined by the following formula:

[0050]

[0051] wherein, the θ bodongfor representing the initial phase angle deviation; the t yuan_max for representing the longest flight time; the t yuan_min for representing the shortest flight time.

[0052] In a third aspect, an embodiment of the present invention further provides a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, the method for determining the initial phase angle deviation of the ultra-fast rendezvous and docking orbit insertion as described in any one of the above is implemented.

[0053] In a fourth aspect, an embodiment of the present invention further provides a computer-readable storage medium, on which a computer program is stored. When the computer program is executed on a computer, the computer is made to execute the method for determining the initial phase angle deviation of the ultra-fast rendezvous and docking orbit insertion as described in any one of the above.

[0054] An embodiment of the present invention provides a method and device for determining the initial phase angle deviation of the ultra-fast rendezvous and docking orbit insertion. The method calculates the shortest flight time and the longest flight time of the tracking spacecraft during the long-distance guidance section after orbit insertion, and uses the flight time as a constraint. By analyzing the shortest flight time and the longest flight time, the initial phase angle deviation of the tracking spacecraft at orbit insertion is obtained. In this way, by ensuring that the initial phase angle at orbit insertion is within this initial phase angle deviation, the safety and robustness of the long-distance guidance section can be further improved and guaranteed. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] In order to more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the following will briefly introduce the drawings required for use in the description of the embodiments or the prior art. Obviously, the following drawings are some embodiments of the present invention. For those of ordinary skill in the art, without creative efforts, other drawings can also be obtained based on these drawings.

[0056] Figure 1 is a flowchart of a method for determining the initial phase angle deviation of the ultra-fast rendezvous and docking orbit insertion provided by an embodiment of the present invention;

[0057] Figure 2 is a graph showing the in-plane trajectory change trend of the abnormal execution of the first pulse at a long distance provided by an embodiment of the present invention;

[0058] Figure 3 is an enlarged view near the long-distance guidance terminal in the graph showing the in-plane trajectory change trend of the abnormal execution of the first pulse at a long distance provided by an embodiment of the present invention;

[0059] Figure 4 is a graph showing the in-plane trajectory change trend of the abnormal execution of the second pulse at a long distance provided by an embodiment of the present invention;

[0060] Figure 5 An enlarged view near the long-distance guidance terminal in the in-plane trajectory change trend diagram of the abnormal execution of the second pulse provided by an embodiment of the present invention;

[0061] Figure 6 It is a hardware architecture diagram of a computing device provided by an embodiment of the present invention;

[0062] Figure 7 It is a structural diagram of a device for determining the initial phase angle deviation of a super-fast rendezvous and docking into orbit provided by an embodiment of the present invention. Detailed implementation manners

[0063] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. Apparently, the described embodiments are some, but not all, of the embodiments of the present invention. All other embodiments obtained by those of ordinary skill in the art based on the embodiments of the present invention without creative efforts shall fall within the protection scope of the present invention.

[0064] The following is the concept of the present invention. As Figure 1 shown, an embodiment of the present invention provides a method for determining the initial phase angle deviation of a super-fast rendezvous and docking into orbit. The method includes:

[0065] Step 100: Obtain the guidance parameters of the tracking spacecraft in the long-distance two-pulse guidance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration when guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration when guiding the second pulse;

[0066] Step 102: Determine the shortest flight time of the tracking spacecraft in the long-distance guidance section according to the preset engine startup time and the guidance parameters;

[0067] Step 104: Determine the longest flight time of the tracking spacecraft in the long-distance guidance section according to the preset rendezvous and docking time and the short-range flight time of the tracking spacecraft in the close-range autonomous control section;

[0068] Step 106: Determine the initial phase angle deviation of the tracking spacecraft at the time of orbit injection according to the shortest flight time and the longest flight time.

[0069] In an embodiment of the present invention, based on the guidance parameters of two-pulse guidance, the preset engine startup time, and the preset rendezvous and docking time, the shortest flight time and the longest flight time of the tracking spacecraft in the long-distance guidance section after entering the orbit are calculated, and the flight time is used as a constraint. By analyzing the shortest flight time and the longest flight time, the initial phase angle deviation of the tracking spacecraft at the time of entering the orbit is obtained. Thus, by ensuring that the initial phase angle at the time of entering the orbit is within the initial phase angle deviation, the safety and robustness of the long-distance guidance section can be further improved and guaranteed.

[0070] The present invention analyzes the time fluctuation range of the long-distance guidance flight from the mission constraints, and innovatively gives a rough analytical solution of the pulse execution time and the orbit entry phase fluctuation, so as to solve the safety and robustness problem of the long-distance guidance section under the ultra-short flight time.

[0071] The following describes Figure 1 the execution manners of the various steps shown.

[0072] In step 100, after entering the orbit, the tracking spacecraft calculates two guidance pulses and their corresponding accelerations of the tracking spacecraft in the long-distance guidance section according to its orbit after entering the orbit and the orbit where the target spacecraft is located.

[0073] In step 102, according to the preset engine startup time and the guidance parameters, the shortest flight time of the tracking spacecraft in the long-distance guidance section is determined, including:

[0074] According to the preset engine startup time, the first velocity increment, and the first acceleration, the first pulse time of the first pulse is determined;

[0075] According to the first pulse time, the first velocity increment, and the second velocity increment, the quasi-Newton method is used to solve the two-pulse guidance to obtain the second pulse time of the second pulse;

[0076] According to the second pulse time, the second velocity increment, and the second acceleration, the shortest flight time is calculated.

[0077] It should be noted that a combined control strategy of semi-major axis, eccentricity, argument of perigee, and argument of latitude is adopted to achieve long-distance guidance, and the quasi-Newton method is used to solve the two-pulse guidance to obtain the first velocity increment, the first pulse time, the second velocity increment, and the second pulse time. The second pulse time is the second pulse time corresponding to the first pulse time. This solution method adopts the prior art and will not be elaborated in detail.

[0078] In the present invention, a two-pulse guidance is adopted in the long-distance guidance section. The turn-on time of the first pulse is restricted by factors such as the establishment of the initial attitude, orbit, and platform after the rocket separates and enters the orbit, and thus cannot be too early. After the tracking spacecraft enters the orbit and eliminates the attitude deviation, it will not immediately enter the rapid rendezvous and docking. After entering the rapid rendezvous and docking, it also needs to wait for a period of time before starting to calculate the guidance parameters for the long-distance guidance. Generally speaking, after the tracking spacecraft enters the orbit, time system setting, sensor status confirmation, orbit confirmation, engine status setting, etc. need to be carried out. Therefore, it is necessary to determine the earliest engine turn-on time (i.e., the preset engine turn-on time) after entering the orbit to minimize the total time of the rendezvous and docking.

[0079] In a specific embodiment, the first pulse moment is determined by the following formula:

[0080]

[0081] wherein, the t 1min is used to represent the first pulse moment; the V p1 is used to represent the first velocity increment; the a1 is used to represent the first acceleration; the preset engine turn-on time is 900 s;

[0082] The shortest flight time is determined by the following formula:

[0083]

[0084] wherein, the t yuan_min is used to represent the shortest flight time; the t 2min is used to represent the second pulse moment; the V p2 is used to represent the second velocity increment; the a2 is used to represent the second acceleration.

[0085] In the present invention, after the tracking spacecraft enters the orbit and eliminates the attitude deviation, it can enter the rapid rendezvous and docking after 5 minutes. After entering the rapid rendezvous and docking, it takes 10 minutes for orbit determination, etc. before starting to calculate the guidance parameters for the long-distance guidance. Therefore, it is determined that 15 minutes after entering the orbit is the earliest engine turn-on time, that is, the preset engine turn-on time is 900 s. Considering that the guidance pulse requires thrust execution, the minimum value of the first pulse moment is t 1min , which is calculated by formula (1); since the pulses of the two-pulse guidance appear in pairs, the corresponding shortest flight time of the long-distance guidance section is calculated by formula (2) at this time.

[0086] In step 104, the determination of the longest flight time of the tracking spacecraft in the long-distance guidance section includes:

[0087] Perform a difference operation on the preset rendezvous and docking time and the short-range flight time to obtain the longest flight time.

[0088] Specifically, the longest flight time is determined by the following formula:

[0089] t yuan_max = t zong - t jin (3)

[0090] Wherein, the t yuan_max is used to represent the longest flight time; the t zong is used to represent the preset rendezvous and docking time; the t jin is used to represent the short-range flight time.

[0091] In the present invention, the preset rendezvous and docking times corresponding to different ultra-fast rendezvous and docking missions are different. When determining an ultra-fast rendezvous and docking mission, its preset rendezvous and docking time t zong has been determined under the influence of TT&C constraints, the short-range flight time t jin of the close-range autonomous control section is basically determined, and the tracking spacecraft immediately switches to the close-range autonomous control section after the second pulse ends, then the longest flight time can be calculated by the above formula (3).

[0092] In a specific embodiment, the preset rendezvous and docking time does not exceed 2 h.

[0093] After step 104, it further includes:

[0094] Calculate the latest start-up time of the second pulse according to the longest flight time, the second velocity increment, and the second acceleration;

[0095] The latest start-up time is determined by the following formula:

[0096]

[0097] Wherein, the t 2max is used to represent the latest start-up time; the t yuan_max is used to represent the longest flight time; the V p2 is used to represent the second velocity increment; a2 is used to represent the second acceleration.

[0098] In the present invention, after obtaining the longest flight time of the long-range guidance section, in order to ensure the smooth execution of the long-range guidance section, the latest start-up time of the second pulse can also be determined based on this longest flight time, and further time constraints are imposed on the execution of the second pulse, further improving the safety and robustness of the long-range guidance section.

[0099] In step 106, determining the initial phase angle deviation of the tracking spacecraft at the time of orbit injection includes:

[0100] Performing a difference operation on the longest flight time and the shortest flight time to obtain a fluctuation time;

[0101] Calculating the initial phase angle deviation based on the fluctuation time;

[0102] The initial phase angle deviation is determined by the following formula:

[0103]

[0104] where, the θ bodong is used to represent the initial phase angle deviation; the t yuan_max is used to represent the longest flight time; the t yuan_min is used to represent the shortest flight time.

[0105] It should be noted that the unit of time involved in the above formulas (1) to (5) is seconds, s.

[0106] Specifically, after obtaining the initial phase angle deviation, the nominal phase angle θ biaocheng can be further determined. Then, the initial phase angle fluctuation range is [θ biaocheng - 0.5θ bodong , θ biaocheng + 0.5θ bodong . For example, when the fluctuation time is 450 s, the initial phase angle deviation is 1 deg, and the initial phase angle fluctuation range is [θ biaocheng - 0.5, θ biaocheng + 0.5].

[0107] The present invention determines the flight time under engineering constraints, designs pulses under the constrained time, determines the initial phase angle deviation of orbit injection, and applies it to the 2-hour ultra-fast rendezvous and docking mission of Tianzhou-5 cargo spacecraft, achieving good results. The flight process is safe and reliable, with good robustness and good engineering practicability. At the same time, the method for determining the initial phase angle deviation of ultra-fast rendezvous and docking orbit injection can also be extended and applied to missions such as space-based attack and defense and on-orbit services.

[0108] After step 106, it further includes:

[0109] S1: Judging whether the guidance parameters are the same as the pre-designed guidance parameters for two-pulse guidance at a long distance; if so, execute step S2, if not, terminate the ultra-fast rendezvous and docking mission;

[0110] S2: Obtaining the third velocity increment executed by the first pulse in the two-pulse guidance at a long distance;

[0111] S3: Determine whether the absolute error increment between the third velocity increment and the first nominal calculated pulse of the first pulse is greater than a first error threshold; if not, execute step S4, and if so, exit the ultra-fast rendezvous and docking mission;

[0112] S4: The tracking spacecraft executes the second pulse and obtains a fourth velocity increment executed by the second pulse in the long-distance two-pulse guidance;

[0113] S5: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is greater than a second error threshold; if not, execute step S6, and if so, exit the ultra-fast rendezvous and docking mission;

[0114] S6: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is less than a third error threshold; wherein, the second error threshold is greater than the third error threshold; if so, execute step S7; if not, execute step S8;

[0115] S7: Continue to execute the ultra-fast rendezvous and docking mission;

[0116] S8: The tracking spacecraft adopts an active braking strategy to ensure the safety of the ultra-fast rendezvous and docking mission.

[0117] Specifically, in the present invention, it also includes an analysis of the safety of the rendezvous and docking mission for the long-distance two-pulse guidance. First, the safety analysis for the execution deviation of the first pulse in the long-distance guidance: Through simulation analysis, when the first pulse is too large, greater than the nominal value, at the end moment of the long-distance guidance, the altitude of the tracking spacecraft is higher than that of the target spacecraft, and the two spacecraft will not collide, and the mission is safe; if the first pulse of the long-distance guidance is too small, less than the nominal value, at the end moment of the long-distance guidance, the altitude of the tracking spacecraft is lower than that of the target spacecraft, and the two spacecraft will not collide, and the mission is safe. Secondly, the safety analysis for the execution deviation of the second pulse in the long-distance guidance: Through simulation analysis, when the second pulse is too large (for example, greater than the nominal value of 2 m / s), at the end moment of the long-distance guidance, the altitude of the tracking spacecraft is higher than that of the target spacecraft, and the two spacecraft will not collide, and the mission is safe; if the second pulse of the long-distance guidance is too small (for example, less than the nominal value of 2 m / s), at the end moment of the long-distance guidance, the altitude of the tracking spacecraft is lower than that of the target spacecraft, and the two spacecraft will not collide, and the mission is safe; when the execution value of the second pulse velocity is between the nominal values (for example, between -2 m / s and 2 m / s), a collision may occur, but the collision time is 30 minutes after the end of the second pulse. It should be noted that less than 2 m / s of the nominal value means that the difference between the nominal value and the actual value is less than 2 m / s.

[0118] It should be noted that the judgment in step S1 can be to judge whether the guidance parameters are the same as the pre-designed guidance parameters for two-pulse guidance at a long distance, or to judge whether the error between the guidance parameters and the pre-designed guidance parameters for two-pulse guidance at a long distance is less than a preset error threshold; if the two are the same or the error is less than the preset error threshold, step S2 is executed. Among them, the pre-designed guidance parameters are calculated by the ground measurement and control platform based on relevant orbit injection parameters, orbital altitude and other information. The third velocity increment in step S2 is the actual velocity increment when the first pulse is actually executed, while the first velocity increment is the theoretical velocity increment of the first pulse; the fourth velocity increment in step S4 is the actual velocity increment when the second pulse is actually executed, while the second velocity increment is the theoretical velocity increment of the second pulse.

[0119] In the present invention, after determining the initial phase angle deviation, deviation analysis is performed on the long-distance guidance pulse, and the safety strategy during long-distance guidance is executed as follows: First, when it is determined that the guidance parameters are the same as the pre-designed guidance parameters, the third velocity increment executed by the first pulse during the actual rendezvous and docking is obtained. If there is an error in the execution of the first pulse, its trajectory is still safe. However, if the absolute error increment between the third velocity increment and the first nominal calculation pulse is greater than the first error threshold, the ultra-fast rendezvous and docking mission cannot be completed; otherwise, the second pulse is continued to be executed. Then, the fourth velocity increment executed by the second pulse during the actual rendezvous and docking is obtained. When the absolute error increment between the fourth velocity increment and the second nominal calculation pulse of the second pulse is less than the third error threshold, it is confirmed that the mission can be carried out normally, and the ultra-fast rendezvous and docking mission is continued; if the absolute error increment is not less than the third error threshold and not greater than the second error threshold, it indicates that the trajectory is unsafe, and an active braking strategy is adopted to ensure the safety of the current mission; if the absolute error increment is greater than the second error threshold, the ultra-fast rendezvous and docking mission needs to be exited.

[0120] In a specific embodiment, the active braking strategy is to execute an active evacuation pulse of -1 m / s within 20 minutes.

[0121] For example, in a specific embodiment, a scheme design is carried out for a 2-hour ultra-fast rendezvous and docking mission, including:

[0122] A1: Determine the shortest flight time

[0123] The tracking spacecraft needs to eliminate attitude deviation and other factors for 5 minutes after orbit injection before entering the ultra-fast rendezvous and docking. After entering the ultra-fast rendezvous and docking, it takes 10 minutes for orbit determination and other operations to start calculating the guidance parameters for long-distance guidance, and time system settings, sensor status confirmation, orbit confirmation, engine status settings, etc. need to be carried out. Therefore, 15 minutes after orbit injection is the earliest engine startup time.

[0124] The earliest end time of the second long-range pulse calculated based on the earliest engine start time is 4150 s, so the shortest flight time is 4150 s.

[0125] A2: Determine the longest flight time

[0126] According to the TT&C constraints, the longest flight time of the tracking spacecraft during the entire mission cannot exceed 114 minutes, that is, the preset rendezvous and docking time is 6840 s. When the short-range flight time is determined to be 2250 s, the longest flight time in the long-range guidance section is 4590 s.

[0127] A3: Determine the initial phase angle deviation

[0128] The time fluctuation range in the long-range guidance section is 4150 s to 4590 s, and the fluctuation time is 440 s. After calculation, the initial phase angle deviation is 1 deg, that is, ±0.5 deg.

[0129] A4: Analysis of the safety of the first pulse execution in case of abnormality

[0130] The guidance parameters in the long-range guidance section are the same as those pre-designed for the two-pulse guidance in the long range. The deviation of the first pulse execution is simulated. In the simulation, the pulse deviation is ±2 m / s. After the execution of the first long-range pulse, the second pulse is not executed. Figure 2 It is half an orbital period after entering the orbit until the second pulse time. Figure 3 It is the enlarged part of the long-range guidance terminal. From Figure 2 and Figure 3 It can be seen that the closest distance does not enter the error ellipsoid and the trajectory is passively safe. Therefore, if the first pulse is calculated correctly but there is an error in the execution, no pulse correction is required.

[0131] A5: Analysis of the safety of the second pulse execution in case of abnormality

[0132] The execution of the first long-range pulse is normal, but there is a deviation in the execution of the second pulse. After the execution of the second pulse, it drifts freely for about half an orbital period. Considering the deviation of ±2 m / s, the in-plane relative motion trajectory is as shown in Figure 4 The enlarged part of the long-range guidance terminal is as shown in Figure 5 As shown in Figure 4 and Figure 5 It can be seen that if the second pulse is too small and the altitude is insufficient, no collision will occur; if the second pulse is too large and it drifts backward, the trajectory is safe. A collision may occur under the trajectory where the execution of the second pulse is abnormal and slightly larger than the normal pulse, but the collision generally occurs after 30 minutes, providing an opportunity for ground processing.

[0133] As shown in Figure 6 、 Figure 7As shown in the figure, an embodiment of the present invention provides a device for determining the initial phase angle deviation of a super-fast rendezvous and docking for orbit insertion. The device embodiment can be implemented by software, or by hardware, or by a combination of software and hardware. At the hardware level, as Figure 6 shown, it is a hardware architecture diagram of a computing device where the device for determining the initial phase angle deviation of a super-fast rendezvous and docking for orbit insertion provided by an embodiment of the present invention is located. In addition to Figure 6 the processor, memory, network interface, and non-volatile memory shown, the computing device where the device in the embodiment is located usually may also include other hardware, such as a forwarding chip responsible for processing packets, etc. Taking software implementation as an example, as Figure 7 shown, as a logically meaningful device, it is formed by the CPU of its computing device reading the corresponding computer program in the non-volatile memory into the memory and running it. A device for determining the initial phase angle deviation of a super-fast rendezvous and docking for orbit insertion provided by this embodiment includes:

[0134] An acquisition module 700, configured to acquire the guidance parameters of the tracking spacecraft during long-distance two-pulse guidance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration when guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration when guiding the second pulse;

[0135] A flight time determination module 702, configured to determine the shortest flight time of the tracking spacecraft during the long-distance guidance section according to a preset engine startup time and the guidance parameters; and determine the longest flight time of the tracking spacecraft during the long-distance guidance section according to a preset rendezvous and docking time and the short-range flight time of the tracking spacecraft during the short-distance autonomous control section;

[0136] A phase angle deviation determination module 704, configured to determine the initial phase angle deviation of the tracking spacecraft when entering the orbit according to the shortest flight time and the longest flight time.

[0137] In some specific implementation manners, the acquisition module 700 may be used to execute the above step 100, the flight time determination module 702 may be used to execute the above steps 102 and 104, and the phase angle deviation determination module 706 may be used to execute the above step 106.

[0138] In some specific implementation manners, the flight time determination module 702 is further configured to perform the following operations:

[0139] Determine the first pulse moment of the first pulse according to the preset engine startup time, the first velocity increment, and the first acceleration;

[0140] The first pulse moment is determined by the following formula:

[0141]

[0142] Among them, the t 1min is used to represent the first pulse moment; the V p1 is used to represent the first speed increment; a1 is used to represent the first acceleration; the preset engine start-up time is 900 s;

[0143] According to the first pulse moment, the first speed increment, and the second speed increment, the quasi-Newton method is used to solve the two-pulse guidance to obtain the second pulse moment of the second pulse;

[0144] According to the second pulse moment, the second speed increment, and the second acceleration, the shortest flight time is calculated;

[0145] The shortest flight time is determined by the following formula:

[0146]

[0147] Among them, the t yuan_min is used to represent the shortest flight time; the t 2min is used to represent the second pulse moment; the V p2 is used to represent the second speed increment; a2 is used to represent the second acceleration.

[0148] In some specific embodiments, the flight time determination module 702 is further configured to perform the following operations:

[0149] Perform a difference operation on the preset rendezvous and docking time and the short-range flight time to obtain the longest flight time.

[0150] In some specific embodiments, the flight time determination module 702 is further configured to perform the following operations:

[0151] According to the longest flight time, the second speed increment, and the second acceleration, calculate the latest start-up moment of the second pulse;

[0152] The latest start-up moment is determined by the following formula:

[0153]

[0154] Among them, the t 2max is used to represent the latest start-up moment; the t yuan_max is used to represent the longest flight time; the V p2 is used to represent the second speed increment; a2 is used to represent the second acceleration.

[0155] In some specific embodiments, the phase angle deviation determination module 704 is further configured to perform the following operations:

[0156] Perform a difference operation on the longest flight time and the shortest flight time to obtain a fluctuation time;

[0157] Calculate the initial phase angle deviation based on the fluctuation time;

[0158] The initial phase angle deviation is determined by the following formula:

[0159]

[0160] where, the θ bodong is used to represent the initial phase angle deviation; the t yuan_max is used to represent the longest flight time; the t yuan_min is used to represent the shortest flight time.

[0161] In some specific embodiments, the device further includes a safety module, and the safety module is configured to perform the following operations:

[0162] S1: Determine whether the guidance parameter is the same as the guidance parameter pre-designed for two-pulse guidance at a long distance; if so, execute step S2, if not, terminate the ultra-fast rendezvous and docking mission;

[0163] S2: Obtain the third velocity increment executed by the first pulse in the two-pulse guidance at a long distance;

[0164] S3: Determine whether the absolute error increment between the third velocity increment and the first nominal calculated pulse of the first pulse is greater than a first error threshold; if not, execute step S4, if so, exit the ultra-fast rendezvous and docking mission;

[0165] S4: The tracking spacecraft executes the second pulse and obtains the fourth velocity increment executed by the second pulse in the two-pulse guidance at a long distance;

[0166] S5: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is greater than a second error threshold; if not, execute step S6, if so, exit the ultra-fast rendezvous and docking mission;

[0167] S6: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is less than a third error threshold; where, the second error threshold is greater than the third error threshold; if so, execute step S7; if not, execute step S8;

[0168] S7: Continue to execute the ultra-fast rendezvous and docking mission;

[0169] S8: The tracking spacecraft adopts an active braking strategy to ensure the safety of the ultra-fast rendezvous and docking mission.

[0170] It can be understood that the structure illustrated in the embodiments of the present invention does not constitute a specific limitation on a device for determining the deviation of the initial phase angle of ultra-fast rendezvous and docking into orbit. In some other embodiments of the present invention, a device for determining the deviation of the initial phase angle of ultra-fast rendezvous and docking into orbit may include more or fewer components than those shown in the figure, or combine certain components, or split certain components, or have different component arrangements. The illustrated components can be implemented in hardware, software, or a combination of software and hardware.

[0171] Regarding the information interaction, execution process, etc. between the various modules within the above device, since they are based on the same concept as the method embodiments of the present invention, the specific content can be referred to the description in the method embodiments of the present invention, and will not be elaborated here.

[0172] The embodiments of the present invention further provide a computing device, including a memory and a processor. A computer program is stored in the memory. When the processor executes the computer program, it implements a method for determining the deviation of the initial phase angle of ultra-fast rendezvous and docking into orbit in any embodiment of the present invention.

[0173] The embodiments of the present invention further provide a computer-readable storage medium. A computer program is stored on the computer-readable storage medium. When the computer program is executed by a processor, the processor is enabled to execute a method for determining the deviation of the initial phase angle of ultra-fast rendezvous and docking into orbit in any embodiment of the present invention.

[0174] Specifically, a system or device equipped with a storage medium can be provided. Software program codes for implementing the functions of any one of the above embodiments are stored on the storage medium, and the computer (or CPU or MPU) of the system or device reads and executes the program codes stored on the storage medium.

[0175] In this case, the program code read from the storage medium itself can implement the functions of any one of the above embodiments. Therefore, the program code and the storage medium storing the program code constitute a part of the present invention.

[0176] Embodiments of the storage medium for providing program codes include floppy disks, hard disks, magneto-optical disks, optical disks (such as CD-ROM, CD-R, CD-RW, DVD-ROM, DVD-RAM, DVD-RW, DVD+RW), magnetic tapes, non-volatile memory cards, and ROMs. Optionally, the program codes can be downloaded from a server computer via a communication network.

[0177] A computer-readable signal medium may include a data signal propagated in a baseband or as part of a carrier wave, in which computer-readable program code is carried. Such propagated data signals may take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination of the foregoing. A computer-readable signal medium may also 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 an instruction execution system, apparatus, or device.

[0178] The program code contained on a computer-readable medium may be transmitted using any appropriate medium, including but not limited to wireless, wireline, optical fiber cable, RF, etc., or any suitable combination of the foregoing.

[0179] The computer program code for performing the operations of the present invention may be written in one or more programming languages or combinations thereof. The programming languages include object-oriented programming languages such as Java, Smalltalk, C++, and also include conventional procedural programming languages such as the "C" language or similar programming languages. The program code may be executed entirely on the user's computer, partially on the user's computer, executed as a stand-alone software package, partially on the user's computer and partially on a remote computer, or entirely on a remote computer or server. In the case of a remote computer, the remote computer may be connected to the user's computer through any type of network, including a local area network (LAN) or a wide area network (WAN), or may be connected to an external computer (e.g., through the Internet using an Internet service provider).

[0180] In addition, it should be clear that not only can the functions of any one of the above embodiments be realized by executing the program code read by a computer, but also by causing an operating system or the like operating on the computer based on the instructions of the program code to complete some or all of the actual operations.

[0181] In addition, it can be understood that the program code read from the storage medium is written into the memory provided in an expansion board inserted into the computer or into the memory provided in an expansion module connected to the computer, and then based on the instructions of the program code, a CPU or the like installed on the expansion board or expansion module is caused to execute some and all of the actual operations, thereby realizing the functions of any one of the above embodiments.

[0182] It should be noted that in this text, relational terms such as first and second are only used to distinguish one entity or operation from another entity or operation, and do not necessarily require or imply any actual relationship or order between these entities or operations. Moreover, the terms "comprising", "including" or any other variant thereof are intended to cover non-exclusive inclusion, such that a process, method, article or device comprising a series of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such process, method, article or device. Without further limitation, an element defined by the statement "comprising an..." does not exclude the presence of additional identical elements in the process, method, article or device comprising said element.

[0183] Those of ordinary skill in the art can understand that all or part of the steps of implementing the above method embodiments can be completed by hardware related to program instructions. The foregoing program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps including those of the above method embodiments; and the foregoing storage medium includes various media that can store program codes, such as ROM, RAM, magnetic disks or optical discs.

[0184] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention and are not intended to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those of ordinary skill in the art should understand that they can still modify the technical solutions described in the foregoing embodiments, or perform equivalent replacements for some of the technical features; and these modifications or replacements do not cause the essence of the corresponding technical solutions to deviate from the spirit and scope of the technical solutions of the embodiments of the present invention.

Claims

1. A method for determining the deviation of the initial phase angle of a super-fast rendezvous and docking orbit insertion, characterized in that, Including: Obtaining the guidance parameters of the tracking spacecraft in the two-pulse guidance at a long distance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration when guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration when guiding the second pulse; Determining the shortest flight time of the tracking spacecraft in the long-distance guidance section according to the preset engine start time and the guidance parameters; Determining the longest flight time of the tracking spacecraft in the long-distance guidance section according to the preset rendezvous and docking time and the short-range flight time of the tracking spacecraft in the close-range autonomous control section; Determining the initial phase angle deviation of the tracking spacecraft at the time of orbit injection according to the shortest flight time and the longest flight time, including: Performing a difference operation on the longest flight time and the shortest flight time to obtain the fluctuation time; Calculating the initial phase angle deviation according to the fluctuation time; The initial phase angle deviation is determined by the following formula: wherein, the θ bodong is used to represent the initial phase angle deviation; the t yuan_max is used to represent the longest flight time; the t yuan_min is used to represent the shortest flight time.

2. The method according to claim 1, characterized in that, The determining the shortest flight time of the tracking spacecraft in the long-distance guidance section according to the preset engine start time and the guidance parameters includes: Determining the first pulse moment of the first pulse according to the preset engine start time, the first velocity increment, and the first acceleration; Performing two-pulse guidance solution by using the quasi-Newton method according to the first pulse moment, the first velocity increment, and the second velocity increment to obtain the second pulse moment of the second pulse; Calculating the shortest flight time according to the second pulse moment, the second velocity increment, and the second acceleration.

3. The method according to claim 2, characterized in that, The first pulse moment is determined by the following formula: wherein, the t 1min is used to represent the first pulse moment; the V p1 is used to represent the first speed increment; a1 is used to represent the first acceleration; the preset engine start-up time is 900 s; The shortest flight time is determined by the following formula: wherein, the t yuan_min is used to represent the shortest flight time; the t 2min is used to represent the second pulse moment; the V p2 is used to represent the second speed increment; the a2 is used to represent the second acceleration.

4. The method according to claim 1, characterized in that, The determining the longest flight time of the tracking spacecraft in the long-distance guidance section includes: Performing a difference operation on the preset rendezvous and docking time and the short-range flight time to obtain the longest flight time.

5. The method according to claim 1, characterized in that, After the determining the longest flight time of the tracking spacecraft in the long-distance guidance section, it further includes: Calculating the latest start time of the second pulse according to the longest flight time, the second velocity increment, and the second acceleration; The latest start time is determined by the following formula: wherein, the t 2max is used to represent the latest startup time; the t yuan_max is used to represent the longest flight time; the V p2 is used to represent the second speed increment; the a2 is used to represent the second acceleration.

6. The method according to any one of claims 1 to 5, characterized in that, After the determining the initial phase angle deviation of the tracking spacecraft at the time of orbit injection, it further includes: S1: Judging whether the guidance parameters are the same as the pre-designed guidance parameters in the two-pulse guidance at a long distance; if so, execute step S2, if not, terminate the ultra-fast rendezvous and docking mission; S2: Obtaining the third velocity increment executed by the first pulse in the two-pulse guidance at a long distance; S3: Judging whether the absolute error increment between the third velocity increment and the first nominal calculated pulse of the first pulse is greater than the first error threshold; if not, execute step S4, if so, exit the ultra-fast rendezvous and docking mission; S4: The tracking spacecraft executes the second pulse and obtains the fourth velocity increment executed by the second pulse in the two-pulse guidance at a long distance; S5: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is greater than a second error threshold; if not, execute step S6, if so, exit the ultra-fast rendezvous and docking mission; S6: Determine whether the absolute error increment between the fourth velocity increment and the second nominal calculated pulse of the second pulse is less than a third error threshold; wherein, the second error threshold is greater than the third error threshold; if so, execute step S7; if not, execute step S8; S7: Continue to execute the ultra-fast rendezvous and docking mission; S8: The tracking spacecraft adopts an active braking strategy to ensure the safety of the ultra-fast rendezvous and docking mission.

7. An initial phase angle deviation determination device for ultra-fast rendezvous and docking orbit insertion, characterized in that Comprising: An acquisition module, configured to acquire the guidance parameters of the tracking spacecraft during long-distance two-pulse guidance; wherein, the guidance parameters include the first velocity increment of the first pulse, the first acceleration during guiding the first pulse, the second velocity increment of the second pulse, and the second acceleration during guiding the second pulse; A flight time determination module, configured to determine the shortest flight time of the tracking spacecraft in the long-distance guidance section according to a preset engine startup time and the guidance parameters; and determine the longest flight time of the tracking spacecraft in the long-distance guidance section according to a preset rendezvous and docking time and the short-range flight time of the tracking spacecraft in the short-distance autonomous control section; A phase angle deviation determination module, configured to determine the initial phase angle deviation of the tracking spacecraft at the time of orbit insertion according to the shortest flight time and the longest flight time; The phase angle deviation determination module is further configured to perform the following operations: Perform a difference operation on the longest flight time and the shortest flight time to obtain a fluctuation time; Calculate the initial phase angle deviation according to the fluctuation time; The initial phase angle deviation is determined by the following formula: Among them, the θ bodong is used to represent the initial phase angle deviation; the t yuan_max is used to represent the longest flight time; the t yuan_min is used to represent the shortest flight time.

8. A computing device, including a memory and a processor, wherein a computer program is stored in the memory, and when the processor executes the computer program, the method described in any one of claims 1-6 is implemented.

9. A computer-readable storage medium, on which a computer program is stored, and when the computer program is executed in a computer, the computer is made to execute the method described in any one of claims 1-6.

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