A long-range guidance method and device based on in-orbit orbit control data evaluation correction

Abstract

The application provides a long-distance guiding method and device based on in-orbit orbit control data evaluation correction, and relates to the technical field of rendezvous and docking. The method comprises the following steps: acquiring a pitch attitude value and an acceleration of a pulse execution process of a tracking spacecraft in a de-orbiting process of the tracking spacecraft from an orbit of a target spacecraft to an initial entry orbit; determining a first trace position error according to the pitch attitude value; determining a second trace position error according to the acceleration and a first pulse executed by the tracking spacecraft in a long-distance guiding section; and evaluating and correcting a terminal nominal position of long-distance guiding calculated by the tracking spacecraft from the initial entry orbit according to the first trace position error and the second trace position error. The long-distance guiding method based on in-orbit orbit control data evaluation correction provided by the scheme can realize evaluation correction of the terminal of the long-distance guiding section based on de-orbiting orbit control data, and obtain a high-precision terminal of the long-distance guiding section.

Description

Technical Field

[0001] This invention relates to the field of rendezvous and docking technology, and in particular to a long-distance guidance method and device based on on-orbit track control data evaluation and correction. Background Technology

[0002] The rendezvous and docking strategy is the only way to transport astronauts from the ground to the International Space Station. This method has been used by Soyuz spacecraft, the Space Shuttle and the International Space Station, as well as Shenzhou spacecraft. Traditional rendezvous and docking methods take 2 to 3 days, and faster rendezvous and docking is becoming the current trend.

[0003] Traditional rendezvous and docking schemes include a long-range guidance phase, a short-range autonomous control phase, and a docking phase. For a 2-hour ultra-fast docking, the flight time requirements for these three phases are very stringent. In order to reduce the total flight time, the long-range guidance phase must provide precise initial conditions for the short-range autonomous control phase while minimizing flight time. Therefore, determining a high-precision long-range guidance phase endpoint is a key challenge. Summary of the Invention

[0004] This invention provides a long-distance guidance method and device based on on-orbit track control data evaluation and correction. This method can correct the long-distance guidance segment terminal based on track control data during track descent, resulting in a high-precision long-distance guidance segment terminal.

[0005] In a first aspect, embodiments of the present invention provide a long-range guidance method based on on-orbit track control data evaluation and correction, comprising:

[0006] Acquire the pitch attitude values ​​and acceleration during the pulse execution process of the tracked spacecraft during its orbital change from the target spacecraft's orbit to its initial insertion orbit;

[0007] The first trajectory position error is determined based on the pitch attitude value;

[0008] The second trajectory position error is determined based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase.

[0009] The terminal nominal position of the tracking spacecraft, calculated from its initial orbital trajectory, is evaluated and corrected based on the first track position error and the second track position error.

[0010] Optionally, determining the first trajectory position error based on the pitch attitude value includes:

[0011] Based on the pitch attitude value, calculate the average pitch attitude value when the track control engine pulses are working during the track change process;

[0012] Based on the average pitch attitude value, determine the first trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0013] The average pitch attitude value and the first trajectory position error are determined by the following formula:

[0014]

[0015] δX1=3000θ ave

[0016] Wherein, the θ ave The t0 is used to represent the average pitch attitude value; the t0 is used to represent the start-up time of the orbit control engine; the t f The value is used to indicate the shutdown time of the track control engine; θ(t) is used to indicate the pitch attitude value at time t; and δX1 is used to indicate the first trajectory position error.

[0017] Optionally, the first pulse is determined by the following method:

[0018] Obtain the end-point acceleration of the tracking spacecraft at the end of the orbit change process;

[0019] The velocity increment and pulse timing of the first pulse are solved using a quasi-Newton method; wherein, the tracking spacecraft employs two-pulse guidance during the long-range guidance phase;

[0020] Based on the first pulse, determine the required amount of fuel to be consumed;

[0021] The actual acceleration at the pulse moment is determined based on the thrust of the orbit control engine, the fuel consumption, and the terminal acceleration.

[0022] The start-up time of the first pulse is calculated based on the actual acceleration, the velocity increment, and the pulse time.

[0023] Optionally, the power-on time is calculated using the following formula:

[0024]

[0025]

[0026]

[0027] Wherein, the a end The term t1 is used to represent the terminal acceleration; t1 is used to represent a terminal moment before the track control engine shuts down during the track change process; t fThe value a(t) is used to represent the shutdown time of the track control engine during the track change process; the value a(t) is used to represent the acceleration at time t; the value a impulse The term "t" is used to represent the actual acceleration; "F" represents the thrust of the orbital control engine; "Fuel" represents the fuel consumption; and "t" represents the amount of fuel used. begin Used to indicate the power-on time; the t impulse Used to indicate the pulse time; the V p Used to represent the speed increment.

[0028] Optionally, determining the second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase includes:

[0029] Based on the acceleration, calculate the average track-changing acceleration when the track control engine pulses are working during the track-changing process;

[0030] Obtain the theoretical pulse execution time and pulse time of the tracking spacecraft executing the first pulse in the initial orbit insertion orbit;

[0031] Based on the average orbital acceleration, the first pulse, the pulse time, and the theoretical pulse execution time, determine the second trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0032] The average orbital acceleration and the second trajectory position error are determined by the following formulas:

[0033]

[0034] δX2=290t err

[0035]

[0036] Wherein, the a ave The term t0 is used to represent the average orbital acceleration; t0 is used to represent the start-up time of the orbital control engine; t f The value a(t) represents the shutdown time of the track control engine; the value δX2 represents the acceleration at time t; the value δX2 represents the second trajectory position error; and the value t represents the acceleration at time t. err Used to represent the power-on time error; the t impulse Used to indicate the pulse time; the V p Used to represent the velocity increment of the first pulse; the t ini Used to indicate the execution time of the theoretical pulse.

[0037] Optionally, the step of evaluating and correcting the terminal nominal position of the tracking spacecraft, calculated from the initial orbital insertion trajectory, based on the first track position error and the second track position error includes:

[0038] The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal.

[0039] The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

[0040] Secondly, embodiments of the present invention also provide a long-range guidance device based on on-orbit track control data evaluation and correction, comprising:

[0041] The acquisition module is used to acquire the pitch attitude values ​​and acceleration during the pulse execution process of the tracking spacecraft during the orbital change process from the target spacecraft's orbit to the initial insertion orbit;

[0042] A position error determination module is used to determine a first trajectory position error based on the pitch attitude value; and to determine a second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft in the long-range guidance phase.

[0043] The evaluation and correction module is used to evaluate and correct the terminal nominal position of the tracking spacecraft, calculated from the initial orbital insertion trajectory, based on the first track position error and the second track position error.

[0044] Optionally, the evaluation and correction module is also configured to perform the following operations:

[0045] The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal.

[0046] The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

[0047] Thirdly, embodiments of the present invention also provide a computing device, including a memory and a processor, wherein the memory stores a computer program, and when the processor executes the computer program, it implements the long-distance guidance method based on on-orbit track control data evaluation and correction as described in any of the above claims.

[0048] Fourthly, embodiments of the present invention also provide a computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to execute the long-distance guidance method based on on-orbit track control data evaluation and correction as described in any of the preceding claims.

[0049] This invention provides a long-range guidance method and apparatus based on on-orbit orbit control data evaluation and correction. The method first acquires the pitch attitude and acceleration values ​​of the tracking spacecraft during the entire orbital maneuver from its initial orbital altitude to its initial orbital insertion orbit. Then, using the pitch attitude values, acceleration, and the first pulse executed by the tracking spacecraft during the long-range guidance phase, it determines the first and second trajectory position errors of the tracking spacecraft's terminal position as it enters long-range guidance from its initial orbital insertion orbit. These errors are then used to evaluate and correct the terminal position. Thus, by conducting an in-space rendezvous and docking test with a tracking spacecraft already docked with the target spacecraft in the same orbit, and analyzing relevant data during the orbital descent process, accuracy evaluation is achieved. Simultaneously, feedforward compensation for the nominal terminal position during the long-range guidance phase of the rendezvous and docking process can be implemented, thereby improving the terminal's accuracy. Attached Figure Description

[0050] To more clearly illustrate the technical solutions in the embodiments of the present invention or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0051] Figure 1 This is a flowchart of a long-distance guidance method based on on-orbit track control data evaluation and correction provided in an embodiment of the present invention;

[0052] Figure 2 This is a pitch attitude change curve during a single orbital descent provided by an embodiment of the present invention;

[0053] Figure 3 This is an attitude fitting curve for a rapid rendezvous, docking, and orbital change process provided in an embodiment of the present invention;

[0054] Figure 4 An acceleration variation curve for a single orbital descent process provided in an embodiment of the present invention;

[0055] Figure 5 This is a hardware architecture diagram of a computing device provided in an embodiment of the present invention;

[0056] Figure 6This is a structural diagram of a long-distance guidance device based on on-orbit track control data evaluation and correction, provided in an embodiment of the present invention. Detailed Implementation

[0057] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are some embodiments of the present invention, but not all embodiments. All other embodiments obtained by those skilled in the art based on the embodiments of the present invention without creative effort are within the scope of protection of the present invention.

[0058] The following is the concept of the present invention, such as Figure 1 As shown, this embodiment of the invention provides a long-range guidance method based on on-orbit track control data evaluation and correction. The method includes:

[0059] Step 100: Obtain the pitch attitude value and acceleration during the pulse execution process of the tracking spacecraft during the orbit change process from the target spacecraft's orbit to the initial insertion orbit;

[0060] Step 102: Determine the first trajectory position error based on the pitch attitude value;

[0061] Step 104: Determine the second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase;

[0062] Step 106: Evaluate and correct the terminal nominal position of the tracking spacecraft, calculated from the initial orbital insertion trajectory, based on the first track position error and the second track position error.

[0063] It should be noted that in this invention, the tracking spacecraft initially docks with the target spacecraft, at which point their orbits are identical. The tracking spacecraft then descends to its initial insertion orbit for the rendezvous and docking mission, thus obtaining the initial conditions for the mission. This allows the tracking spacecraft to be used again for a rapid rendezvous and docking mission, further refining the determination of the long-range guidance phase's endpoint. In this way, after completing its original mission, the tracking spacecraft can conduct another in-orbit rendezvous and docking test, further improving the endpoint accuracy and optimizing the rapid rendezvous and docking mission. Specifically, the rapid rendezvous and docking mission in this invention is a 2-hour ultra-fast rendezvous and docking mission.

[0064] In this embodiment of the invention, the method first acquires the pitch attitude values ​​and acceleration during the pulse execution process of the tracking spacecraft during its orbital maneuver from the target spacecraft's orbit to its initial insertion orbit. Then, using the pitch attitude values, acceleration, and the first pulse executed by the tracking spacecraft during the long-range guidance phase, it determines the first and second trajectory position errors of the tracking spacecraft's terminal as it enters long-range guidance from its initial insertion orbit. The terminal is then evaluated and corrected based on these first and second trajectory position errors. Thus, by analyzing relevant data from the tracking spacecraft's orbital maneuver before rendezvous and docking, feedforward compensation is achieved for the terminal during the long-range guidance phase of the rendezvous and docking process, thereby improving the accuracy of the terminal's nominal position.

[0065] The following description Figure 1 The execution method of each step is shown.

[0066] In step 100, during the orbital change process of the tracking spacecraft descending from the orbit of the target spacecraft to the initial orbit, at least one pulse is used to descend to the initial orbit, and the pitch attitude value of the tracking spacecraft and the acceleration during the pulse execution process are obtained during the orbital change process.

[0067] In step 102, determining the first trajectory position error based on the pitch attitude value includes:

[0068] Based on the pitch attitude value, calculate the average pitch attitude value when the track control engine pulses are working during the track change process;

[0069] Based on the average pitch attitude value, determine the first trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0070] The average pitch attitude value and the first trajectory position error are determined by the following formula:

[0071]

[0072] δX1=3000θ ave (2)

[0073] Wherein, the θ ave The t0 is used to represent the average pitch attitude value; the t0 is used to represent the start-up time of the orbit control engine; the t f The value is used to indicate the shutdown time of the track control engine; θ(t) is used to indicate the pitch attitude value at time t; and δX1 is used to indicate the first trajectory position error.

[0074] It should be noted that, for a trajectory change process that executes only one pulse, the pulse determines the start-up and shutdown times of the orbit control engine, and the average pitch attitude value is the average change in pitch attitude of the tracking spacecraft caused by the pulse. For a trajectory change process that executes at least two pulses, the average pitch attitude value can be the average change in pitch attitude of the tracking spacecraft caused by all pulses, or the average change in pitch attitude caused by any pulse during the trajectory change process, preferably the average change in pitch attitude caused by the last pulse during the trajectory change process.

[0075] Specifically, if the average pitch attitude value can be considered as the average change in pitch attitude of the tracked spacecraft caused by all pulses, then

[0076]

[0077] Where, θ ave Used to represent the average pitch attitude value; n is the number of pulses executed during the trajectory change process; t 0i Used to indicate the start-up time of the track control engine corresponding to the i-th pulse; t fi θ(t) represents the shutdown time of the track control engine corresponding to the i-th pulse. i ) is used to represent the pitch attitude value corresponding to the i-th pulse at time t.

[0078] In this invention, for on-orbit testing, establishing the initial conditions for rendezvous and docking requires tracking the spacecraft to perform a large-pulse descent. This descent process should have a similar attitude control effect to the large-pulse orbit change in the long-range guidance phase. Therefore, this invention uses on-orbit data analysis during the descent pulse execution to statistically analyze the average pitch attitude deviation (i.e., the average pitch attitude value) during the orbit change process, and determines the first trajectory position error δX1 according to formula (2) to correct the long-range guidance. Meanwhile, since there are many factors affecting orbit control accuracy, and these factors may be coupled, the method of this invention can avoid analyzing the influencing factors causing the pitch attitude deviation, thus simplifying the correction method and improving the correction efficiency.

[0079] In step 104, the first pulse is determined by the following method:

[0080] Obtain the end-point acceleration of the tracking spacecraft at the end of the orbit change process;

[0081] The velocity increment and pulse timing of the first pulse are solved using a quasi-Newton method; wherein, the tracking spacecraft employs two-pulse guidance during the long-range guidance phase;

[0082] Based on the first pulse, determine the required amount of fuel to be consumed;

[0083] The actual acceleration at the pulse moment is determined based on the thrust of the orbit control engine, the fuel consumption, and the terminal acceleration.

[0084] The power-on time is calculated based on the actual acceleration, the velocity increment, and the pulse time.

[0085] In a preferred embodiment, the power-on time is calculated using the following formula:

[0086]

[0087]

[0088]

[0089] Wherein, the a end The term t1 is used to represent the terminal acceleration; t1 is used to represent a terminal moment before the track control engine shuts down during the track change process; t f The value a(t) is used to represent the shutdown time of the track control engine during the track change process; the value a(t) is used to represent the acceleration at time t; the value a impulse The term "t" is used to represent the actual acceleration; "F" represents the thrust of the orbital control engine; "Fuel" represents the fuel consumption; and "t" represents the amount of fuel used. begin Used to indicate the power-on time; the t impulse Used to indicate the pulse time; the V p Used to represent the speed increment.

[0090] Specifically, to avoid errors in the instantaneous acceleration obtained at the end of the orbit change process, the terminal acceleration of the tracking spacecraft at the end of the orbit change process is selected. The average acceleration is then calculated at the end of the orbit change process, where the mass change of the tracking spacecraft is relatively small, to further reduce this data error. For example, the terminal acceleration can be the average of the acceleration of the tracking spacecraft in the last 10 seconds of the orbit change process, or it can be the terminal acceleration in the last 5 seconds of the orbit change process.

[0091] In this invention, since pulse guidance requires thrust, executing the first pulse consumes fuel. To further accurately calculate the acceleration at the pulse moment, formula (5) is used to take the consumed fuel into account and calculate a more accurate actual acceleration a at the pulse moment. impulse At this point, based on formula (6), the start-up time of the first pulse can be calculated more accurately, thereby obtaining the terminal of the high-precision long-distance guidance segment.

[0092] It should be noted that after the tracking spacecraft enters orbit, based on its orbital trajectory and the target spacecraft's orbit, the two guidance pulses and their start times for the long-range guidance phase are calculated. This includes the velocity increment and pulse timing of the first pulse, and the velocity increment and pulse timing of the second pulse. Specifically, the guidance parameters include, but are not limited to, using a combined control strategy of semi-major axis, eccentricity, perigee argument, and latitude argument to achieve long-range guidance, and employing a quasi-Newton method to solve for the two-pulse guidance. The methods for obtaining the guidance parameters utilize existing technology and will not be elaborated further.

[0093] In step 104, determining the second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase includes:

[0094] Based on the acceleration, calculate the average track-changing acceleration when the track control engine pulses are working during the track-changing process;

[0095] The velocity increment and pulse timing of the first pulse are solved using a quasi-Newton method; wherein, the tracking spacecraft employs two-pulse guidance during the long-range guidance phase;

[0096] Obtain the theoretical pulse execution time when the tracking spacecraft executes the first pulse in the initial orbit insertion orbit;

[0097] Based on the average orbital acceleration, the first pulse, the pulse time, and the theoretical pulse execution time, determine the second trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0098] The average orbital acceleration and the second trajectory position error are determined by the following formulas:

[0099]

[0100] δX2=290t err (8)

[0101]

[0102] Wherein, the a ave The term t0 is used to represent the average orbital acceleration; t0 is used to represent the start-up time of the orbital control engine; t f The value a(t) represents the shutdown time of the track control engine; the value δX2 represents the acceleration at time t; the value δX2 represents the second trajectory position error; and the value t represents the acceleration at time t. err Used to represent the power-on time error; the t impulse Used to indicate the pulse time; the V p Used to represent the velocity increment of the first pulse; the t iniUsed to indicate the execution time of the theoretical pulse.

[0103] It should be noted that the average orbit change acceleration is similar to the average pitch attitude value. For an orbit change process that executes only one pulse, the activation and deactivation times of the orbit control engine are determined by that pulse, and the average orbit change acceleration is the average of the orbit change accelerations of the tracking spacecraft caused by the pulse. For an orbit change process that executes at least two pulses, the average orbit change acceleration can be the average of the orbit change accelerations of the tracking spacecraft caused by all pulses, or the average of the orbit change accelerations caused by any pulse in the orbit change process, preferably the average of the orbit change accelerations caused by the last pulse in the orbit change process.

[0104] It should be noted that the theoretical pulse execution time includes, but is not limited to, the theoretical pulse execution time obtained based on prior knowledge, and the theoretical pulse execution time obtained through multiple simulation experiments.

[0105] In one specific implementation, the pulses used to calculate the average orbit change acceleration and the average pitch attitude value are the same. For example, the last pulse is used to calculate both the average orbit change acceleration and the average pitch attitude value.

[0106] In this invention, the long-range guidance phase employs a two-pulse guidance strategy. The first pulse is executed some time after the tracked spacecraft enters orbit, and the second pulse immediately transitions to the near-range autonomous control phase. Therefore, the terminal accuracy of the long-range guidance phase depends on the first pulse. However, pulse guidance requires finite thrust, and inaccurate activation timing can lead to terminal trajectory position errors. Extensive experimental analysis shows that near the execution time of the first pulse, the relative trajectory velocity of the two spacecraft is on the order of approximately 290 m / s. Therefore, advancing or delaying the pulse by t... err When the first pulse is executed at the start of the time, the trajectory position error caused by the end of the long-range guidance segment will be δX2.

[0107] In an embodiment of the present invention, t err When t = 0, the power-on time of the first pulse will not cause an error in the terminal of the long-range guidance segment, therefore at t err≠ When the value is 0, it will affect the terminal accuracy of the long-range guidance segment.

[0108] It should be noted that the pitch attitude values ​​in the above formulas (1) to (9) are in degrees, deg; and the time is in seconds, s.

[0109] In step 106, the terminal nominal position of the tracking spacecraft, calculated from its initial orbital insertion trajectory, is evaluated and corrected based on the first track position error and the second track position error, including:

[0110] The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal.

[0111] The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

[0112] In this invention, feedforward compensation is performed on the long-range guidance segment based on the track control data during the on-orbit orbit change process. The terminal of the long-range guidance segment is directly compensated for the trajectory position through the first trajectory position error to obtain the corrected final terminal position. At the same time, when the second trajectory position error is zero, it is determined that the start-up time of the first pulse calculated by formula (6) will not affect the terminal accuracy. Therefore, the accuracy level of the terminal nominal position is the highest. When the second trajectory position error is not zero, the larger the absolute value of the second trajectory position error, the lower the accuracy level of the terminal nominal position. Thus, the accuracy assessment of the terminal nominal position is achieved through the second trajectory position error.

[0113] It should be noted that the terminal's nominal position is calculated based on the orbital insertion orbit, the target spacecraft's orbit, the two-pulse guidance used by the tracking spacecraft in the long-range guidance phase, and the requirements of the rendezvous and docking mission.

[0114] For example, taking the 2-hour rapid rendezvous and docking on-orbit data analysis of the Tianzhou-2 cargo spacecraft as an example, during the orbital maneuver from the same orbit as the target spacecraft to the initial insertion orbit, a pulse is executed, and the pitch attitude change of the Tianzhou-2 cargo spacecraft during one pulse execution is as follows: Figure 2 As shown, Figure 2 The average pitch attitude value is -0.3817 degrees. Meanwhile, ground-based simulations of pitch attitude changes for rapid rendezvous and docking missions show... Figure 3 The average pitch attitude value is -0.3808deg, and the difference between the two is within the error threshold range. Therefore, it is confirmed that the attitude control effect of the descent process is basically the same as that of the large pulse orbit change in the long-distance guidance section. The method is effective. The long-distance guidance can use this average pitch attitude value as a reference to set attitude-related parameters and correct the terminal.

[0115] In this embodiment, the acceleration change during a single orbital control process prior to the 2-hour rapid rendezvous and docking test is as follows: Figure 4 As shown, the average orbital acceleration is approximately 0.095 m / s². Considering the acceleration a in the final stage... end Take 0.096 m / s². Further considering the acceleration change due to fuel consumption during the first pulse of the rapid rendezvous and docking, determine a. impulse The value is 0.097 m / s². Therefore, when the pulse time t is obtained...impulse When the first pulse is activated, the activation time is t. impulse -V p / 2 / 0.097.

[0116] Thus, by taking the above two points into account, this specific embodiment ensures that the trajectory position deviation of the terminal of the Tianzhou-2 cargo spacecraft during the 2-hour rapid rendezvous and docking test does not exceed 200m, provided that other stages are completed smoothly. This successfully provides high-precision initial conditions for the near-range autonomous control phase.

[0117] like Figure 5 , Figure 6 As shown, this embodiment of the invention provides a long-range guidance device based on on-orbit track control data evaluation and correction. The device embodiment can be implemented through software, hardware, or a combination of both. From a hardware perspective, as... Figure 5 The diagram shown is a hardware architecture diagram of a computing device for a long-range guidance device based on on-orbit track control data evaluation and correction, provided by an embodiment of the present invention. (Except for...) Figure 5 In addition to the processor, memory, network interface, and non-volatile memory shown, the computing device in the embodiment may also include other hardware, such as a forwarding chip responsible for processing packets. Taking software implementation as an example, such as... Figure 6 As shown, a device in a logical sense is formed by the CPU of its computing device reading the corresponding computer program from the non-volatile memory into memory and running it. This embodiment provides a long-range guidance device based on on-orbit track control data evaluation and correction, comprising:

[0118] The acquisition module 300 is used to acquire the pitch attitude value and acceleration during the pulse execution process of the tracking spacecraft during the orbit change process from the target spacecraft's orbit to the initial insertion orbit;

[0119] The position error determination module 302 is used to determine a first trajectory position error based on the pitch attitude value; and to determine a second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft in the long-range guidance phase.

[0120] The evaluation and correction module 304 is used to evaluate and correct the terminal nominal position of the tracking spacecraft, which is calculated from the initial orbital insertion orbit, based on the first track position error and the second track position error.

[0121] In some specific implementations, the acquisition module 300 can be used to perform the above step 100, the position error determination module 302 can be used to perform the above steps 102 and 104, and the evaluation and correction module 304 can be used to perform the above step 106.

[0122] In some specific implementations, the position error determination module 302 is also used to perform the following operations:

[0123] Based on the pitch attitude value, calculate the average pitch attitude value when the track control engine pulses are working during the track change process;

[0124] Based on the average pitch attitude value, determine the first trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0125] The average pitch attitude value and the first trajectory position error are determined by the following formula:

[0126]

[0127] δX1=3000θ ave (2)

[0128] Wherein, the θ ave The t0 is used to represent the average pitch attitude value; the t0 is used to represent the start-up time of the orbit control engine; the t f The value is used to indicate the shutdown time of the track control engine; θ(t) is used to indicate the pitch attitude value at time t; and δX1 is used to indicate the first trajectory position error.

[0129] In some specific embodiments, the device further includes a pulse calculation module, which performs the following operations:

[0130] Obtain the end-point acceleration of the tracking spacecraft at the end of the orbit change process;

[0131] The velocity increment and pulse timing of the first pulse are solved using a quasi-Newton method; wherein, the tracking spacecraft employs two-pulse guidance during the long-range guidance phase;

[0132] Based on the first pulse, determine the required amount of fuel to be consumed;

[0133] The actual acceleration at the pulse moment is determined based on the thrust of the orbit control engine, the fuel consumption, and the terminal acceleration.

[0134] The power-on time is calculated based on the actual acceleration, the velocity increment, and the pulse time.

[0135]

[0136]

[0137]

[0138] Wherein, the a endThe term t1 is used to represent the terminal acceleration; t1 is used to represent a terminal moment before the track control engine shuts down during the track change process; t f The value a(t) is used to represent the shutdown time of the track control engine during the track change process; the value a(t) is used to represent the acceleration at time t; the value a impulse The term "t" is used to represent the actual acceleration; "F" represents the thrust of the orbital control engine; "Fuel" represents the fuel consumption; and "t" represents the amount of fuel used. begin Used to indicate the power-on time; the t impulse Used to indicate the pulse time; the V p Used to represent the speed increment.

[0139] In some specific implementations, the position error determination module 302 is also used to perform the following operations:

[0140] Based on the acceleration, calculate the average track-changing acceleration when the track control engine pulses are working during the track-changing process;

[0141] Obtain the theoretical pulse execution time and pulse time of the tracking spacecraft executing the first pulse in the initial orbit insertion orbit;

[0142] Based on the average orbital acceleration, the first pulse, the pulse time, and the theoretical pulse execution time, determine the second trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working;

[0143] The average orbital acceleration and the second trajectory position error are determined by the following formulas:

[0144]

[0145] δX2=290t err (8)

[0146]

[0147] Wherein, the a ave The term t0 is used to represent the average orbital acceleration; t0 is used to represent the start-up time of the orbital control engine; t f The value a(t) represents the shutdown time of the track control engine; the value δX2 represents the acceleration at time t; the value δX2 represents the second trajectory position error; and the value t represents the acceleration at time t. err Used to represent the power-on time error; the t impulse Used to indicate the pulse time; the V p Used to represent the velocity increment of the first pulse; the t ini Used to indicate the execution time of the theoretical pulse.

[0148] In some specific implementations, the evaluation and correction module 304 is also used to perform the following operations:

[0149] The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal.

[0150] The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

[0151] It is understood that the structures illustrated in the embodiments of the present invention do not constitute a specific limitation on a long-range guidance device based on on-orbit track control data evaluation and correction. In other embodiments of the present invention, a long-range guidance device based on on-orbit track control data evaluation and correction may include more or fewer components than illustrated, or combine some components, or split some components, or have different component arrangements. The illustrated components may be implemented in hardware, software, or a combination of software and hardware.

[0152] The information interaction and execution process between the modules in the above-mentioned device are based on the same concept as the method embodiment of the present invention, and the specific details can be found in the description of the method embodiment of the present invention, and will not be repeated here.

[0153] This invention also provides a computing device, including a memory and a processor. The memory stores a computer program, and when the processor executes the computer program, it implements a long-distance guidance method based on on-orbit track control data evaluation and correction according to any embodiment of this invention.

[0154] This invention also provides a computer-readable storage medium storing a computer program. When executed by a processor, the computer program causes the processor to perform a long-range guidance method based on on-orbit track control data evaluation and correction according to any embodiment of this invention.

[0155] Specifically, a system or apparatus equipped with a storage medium may be provided, on which software program code implementing the functions of any of the embodiments described above is stored, and the computer (or CPU or MPU) of the system or apparatus may read and execute the program code stored in the storage medium.

[0156] In this case, the program code read from the storage medium can itself implement the function of any of the above embodiments, and therefore the program code and the storage medium storing the program code constitute part of the present invention.

[0157] Examples of storage media used to provide program code 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. Alternatively, program code can be downloaded from a server computer via a communication network.

[0158] Computer-readable signal media may include data signals propagated in baseband or as part of a carrier wave, carrying computer-readable program code. Such propagated data signals may take various forms, including but not limited to electromagnetic signals, optical signals, or any suitable combination thereof. Computer-readable signal media may also be any computer-readable medium other than computer-readable storage media, which can send, propagate, or transmit programs for use by or in conjunction with an instruction execution system, system, or device.

[0159] The program code contained on a computer-readable medium may be transmitted using any suitable medium, including, but not limited to, wireless, wire, optical fiber, RF, etc., or any suitable combination thereof.

[0160] Computer program code for performing the operations of this invention can be written in one or more programming languages ​​or a combination thereof, including 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 it can be connected to an external computer (e.g., via the Internet using an Internet service provider).

[0161] Furthermore, it should be clear that not only can the program code read by the computer be executed, but also the operating system or other components operating on the computer can be instructed based on the program code to perform some or all of the actual operations, thereby realizing the function of any of the embodiments described above.

[0162] Furthermore, it is understood that the program code read from the storage medium is written to the memory set in the expansion board inserted into the computer or to the memory set in the expansion module connected to the computer. Then, based on the instructions of the program code, the CPU or other components installed on the expansion board or expansion module execute some and all of the actual operations, thereby realizing the function of any of the above embodiments.

[0163] It should be noted that, in this document, relational terms such as "first" and "second" are used only to distinguish one entity or operation from another, and do not necessarily require or imply any such actual relationship or order between these entities or operations. Furthermore, the terms "comprising," "including," or any other variations thereof are intended to cover non-exclusive inclusion, such that a process, method, article, or apparatus that comprises a list of elements includes not only those elements but also other elements not expressly listed, or elements inherent to such a process, method, article, or apparatus. Without further limitations, an element defined by the phrase "comprising one..." does not exclude the presence of other identical elements in the process, method, article, or apparatus that includes said element.

[0164] Those skilled in the art will understand that all or part of the steps of the above method embodiments can be implemented by hardware related to program instructions. The aforementioned program can be stored in a computer-readable storage medium. When the program is executed, it performs the steps of the above method embodiments. The aforementioned storage medium includes various media that can store program code, such as ROM, RAM, magnetic disk, or optical disk.

[0165] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, and not to limit them; although the present invention has been described in detail with reference to the foregoing embodiments, those skilled in the art should understand that modifications can still be made to the technical solutions described in the foregoing embodiments, or equivalent substitutions can be made to some of the technical features; and these modifications or substitutions 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 long-range guidance method based on on-orbit track control data evaluation and correction, characterized in that, include: Acquire the pitch attitude values ​​and acceleration during the pulse execution process of the tracked spacecraft during its orbital change from the target spacecraft's orbit to its initial insertion orbit; The first trajectory position error is determined based on the pitch attitude value; The second trajectory position error is determined based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase. The terminal nominal position of the tracking spacecraft, calculated from its initial orbital trajectory, is evaluated and corrected based on the first track position error and the second track position error.

2. The method according to claim 1, characterized in that, Determining the first trajectory position error based on the pitch attitude value includes: Based on the pitch attitude value, calculate the average pitch attitude value when the track control engine pulses are working during the track change process; Based on the average pitch attitude value, determine the first trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working; The average pitch attitude value and the first trajectory position error are determined by the following formula: δX1=3000θ ave Wherein, the θ ave The t0 is used to represent the average pitch attitude value; the t0 is used to represent the start-up time of the orbit control engine; the t f The value is used to indicate the shutdown time of the track control engine; θ(t) is used to indicate the pitch attitude value at time t; and δX1 is used to indicate the first trajectory position error.

3. The method according to claim 1, characterized in that, The first pulse is determined by the following method: Obtain the end-point acceleration of the tracking spacecraft at the end of the orbit change process; The velocity increment and pulse timing of the first pulse are solved using a quasi-Newton method; wherein, the tracking spacecraft employs two-pulse guidance during the long-range guidance phase; Based on the first pulse, determine the required amount of fuel to be consumed; The actual acceleration at the pulse moment is determined based on the thrust of the orbit control engine, the fuel consumption, and the terminal acceleration. The start-up time of the first pulse is calculated based on the actual acceleration, the velocity increment, and the pulse time.

4. The method according to claim 3, characterized in that, The power-on time is calculated using the following formula: Wherein, the a end The term t1 is used to represent the terminal acceleration; t1 is used to represent a terminal moment before the track control engine shuts down during the track change process; t f The value a(t) is used to represent the shutdown time of the track control engine during the track change process; the value a(t) is used to represent the acceleration at time t; the value a impulse The term "t" is used to represent the actual acceleration; "F" represents the thrust of the orbital control engine; "Fuel" represents the fuel consumption; and "t" represents the amount of fuel used. begin Used to indicate the power-on time; the t impulse Used to indicate the pulse time; the V p Used to represent the speed increment.

5. The method according to claim 1, characterized in that, The determination of the second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft during the long-range guidance phase includes: Based on the acceleration, calculate the average track-changing acceleration when the track control engine pulses are working during the track-changing process; Obtain the theoretical pulse execution time and pulse time of the tracking spacecraft executing the first pulse in the initial orbit insertion orbit; Based on the average orbital acceleration, the first pulse, the pulse time, and the theoretical pulse execution time, determine the second trajectory position error of the terminal for long-range guidance of the tracking spacecraft when the pulse is working; The average orbital acceleration and the second trajectory position error are determined by the following formulas: δX2=290t err Wherein, the a ave The term t0 is used to represent the average orbital acceleration; t0 is used to represent the start-up time of the orbital control engine; t f The value a(t) represents the shutdown time of the track control engine; the value δX2 represents the acceleration at time t; the value δX2 represents the second trajectory position error; and the value t represents the acceleration at time t. err Used to represent the power-on time error; the t impulse Used to indicate the pulse time; the V p Used to represent the velocity increment of the first pulse; the t ini Used to indicate the execution time of the theoretical pulse.

6. The method according to any one of claims 1 to 5, characterized in that, The evaluation and correction of the terminal nominal position of the tracking spacecraft, calculated from its initial orbital insertion trajectory, based on the first track position error and the second track position error includes: The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal. The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

7. A long-range guidance device based on on-orbit track control data evaluation and correction, characterized in that, include: The acquisition module is used to acquire the pitch attitude values ​​and acceleration during the pulse execution process of the tracking spacecraft during the orbital change process from the target spacecraft's orbit to the initial insertion orbit; A position error determination module is used to determine a first trajectory position error based on the pitch attitude value; and to determine a second trajectory position error based on the acceleration and the first pulse executed by the tracking spacecraft in the long-range guidance phase. The evaluation and correction module is used to evaluate and correct the terminal nominal position of the tracking spacecraft, calculated from the initial orbital insertion trajectory, based on the first track position error and the second track position error.

8. The apparatus according to claim 7, characterized in that, The evaluation and correction module is also used to perform the following operations: The nominal position of the terminal is compensated for based on the first trace position error in order to correct the nominal position of the terminal. The nominal position of the terminal is evaluated based on the second trace position error to determine the accuracy level of the nominal position of the terminal; wherein the absolute value of the second trace position error is inversely proportional to the accuracy level.

9. A computing device comprising a memory and a processor, wherein the memory stores a computer program, and the processor, when executing the computer program, implements the method as described in any one of claims 1-6.

10. A computer-readable storage medium having a computer program stored thereon, which, when executed in a computer, causes the computer to perform the method of any one of claims 1-6.

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