A method for autonomous rendezvous with small celestial bodies based on integrated design of line-of-sight measurement and trajectory maneuver

Abstract

The application provides a method for autonomous rendezvous with small celestial bodies based on integrated design of line-of-sight measurement and trajectory maneuver, comprising: obtaining planned nominal maneuver trajectory and orbit transfer information on the ground; keeping the direction of the probe to the small celestial bodies after each orbit transfer, and measuring the central line-of-sight direction of the target small celestial bodies; according to the line-of-sight measurement results and the measurement results of the orbit transfer velocity increment, solving the initial relative position and initial velocity of the probe to the small celestial bodies, and the current relative position and velocity; according to the relative position at the orbit transfer time, the state transition matrix and the flight time, determining the required velocity increment to reach the next nominal orbit transfer position; controlling the probe to autonomously adjust the posture to the orbit control velocity increment direction, and after reaching the nominal orbit control time, controlling the probe to autonomously start the orbit control process; after completing the orbit transfer, autonomously judging whether the probe reaches the vicinity of the terminal position of the rendezvous section.

Description

TECHNICAL FIELD

[0001] The present application belongs to the field of navigation and guidance in deep space small celestial body exploration, and particularly relates to a method for autonomously rendezvousing with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver. BACKGROUND

[0002] For a small celestial body exploration mission, considering the limited determination orbit accuracy of the target small celestial body and the influence of the determination orbit accuracy of the explorer, the position determination orbit error of the explorer relative to the small celestial body can reach hundreds of kilometers or more. For safety reasons, the ground can only directly guide the explorer to a distance range of hundreds to thousands of kilometers away from the small celestial body, and then the relative navigation of the on-board GNC subsystem can be carried out according to the measurement results of the relative navigation sensor, and the explorer can autonomously approach to a distance of tens of kilometers away from the small celestial body. Generally, such a process is called a rendezvous process with the small celestial body.

[0003] Due to the weight and power limitations, the on-board range-finding radar can generally only reach a measurement range of tens of kilometers at most. On the other hand, even if the range-finding radar can reach a range-finding range of hundreds of kilometers, it is also difficult to accurately point to the target small celestial body which has a diameter of only hundreds of meters. Therefore, for a long-distance range of tens of kilometers, the explorer generally only has the means to measure the central line-of-sight direction of the target small celestial body by using an optical camera, and cannot directly obtain the distance information of the target.

[0004] According to the observability theory of relative navigation, complete information of the relative position cannot be obtained by using only line-of-sight measurement; in order to obtain complete observability of the relative position, a certain trajectory maneuver needs to be combined. For trajectory maneuver, the greater the maneuver amplitude, the stronger the navigation observability obtained, but on the other hand, the more fuel consumed, which requires comprehensive consideration of the optimization of fuel consumption and observability.

[0005] Therefore, it is necessary to provide a method for autonomously rendezvousing with a small celestial body, which solves the problem of rendezvousing with a target small celestial body only by relying on line-of-sight measurement in a range of tens to thousands of kilometers. SUMMARY

[0006] In view of the condition that only the central line-of-sight of the target small celestial body can be measured in a range of tens to thousands of kilometers in a small celestial body exploration mission, the present application provides a method for autonomously rendezvousing with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver, which solves the problem of rendezvousing with a target small celestial body only by relying on line-of-sight measurement.

[0007] The technical scheme provided by the present application is as follows:

[0008] In a first aspect, a method for autonomously rendezvousing with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver, comprising:

[0009] Before the rendezvous flight, a planned nominal maneuver trajectory, a velocity increment of each maneuver and a nominal orbit transfer time are obtained on the ground;

[0010] After each orbit transfer, the probe keeps pointing to the small celestial body, and the optical camera measures the direction of the line of sight to the center of the target small celestial body at a set time interval;

[0011] After each measurement of the direction of the line of sight to the center of the target small celestial body, the least squares method is used to solve the relative navigation based on the accumulated line of sight measurements and the measured velocity increments of the previous two orbit transfer arcs, or based on the accumulated line of sight measurements from the handover point and the measured velocity increments of each orbit transfer, to obtain the initial relative position and initial velocity of the probe relative to the small celestial body, and then the state transition matrix is used to obtain the relative position and velocity of the probe at the current time;

[0012] At the first time before each orbit transfer, the velocity increment required to reach the next nominal orbit transfer position is determined based on the relative position at the orbit transfer time, the state transition matrix and the flight time;

[0013] When the second time before the orbit transfer is reached, the probe is controlled to autonomously adjust the attitude to the direction of the orbit control velocity increment, and after the nominal orbit control time is reached, the probe is controlled to autonomously start the orbit control process;

[0014] After the orbit transfer is completed, it is autonomously judged whether the probe has reached the vicinity of the terminal position of the rendezvous segment, if so, the probe is controlled to autonomously brake and stop the rendezvous process, otherwise the pointing attitude of the probe to the target small celestial body is repeatedly adjusted and the subsequent rendezvous flight process continues.

[0015] In a second aspect, a control device for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver is located on a probe, comprising:

[0016] One or more processors;

[0017] A storage device for storing one or more programs,

[0018] When the one or more programs are executed by the one or more processors, the one or more processors implement the method for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver according to the first aspect.

[0019] In a third aspect, a readable storage medium has a computer program stored thereon, which is executed by a processor to implement the method for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuver according to the first aspect.

[0020] Fourthly, a computer program product comprising: a computer program that, when run, executes the autonomous rendezvous method for small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering as described in the first aspect.

[0021] The method for autonomous rendezvous with small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering provided by the present invention has the following beneficial effects:

[0022] This invention provides a method for autonomous rendezvous with a small celestial body based on integrated line-of-sight measurement and trajectory maneuvering. Before conducting the rendezvous flight, the method acquires the planned nominal maneuver trajectory, the velocity increment for each maneuver, and the nominal orbit change time as indicated on the ground. After each orbit change, the probe maintains its orientation towards the small celestial body, and at set intervals, the line-of-sight direction of the target small celestial body's center is measured using an optical camera. After each measurement of the target small celestial body's center line-of-sight direction, based on the accumulated line-of-sight measurement results within the previous two orbit change arcs and the measured velocity increments of the previous two orbit changes, or, if the number of orbit changes is less than two, based on the accumulated line-of-sight measurement results from the handover point and the measured velocity increments of each orbit change, the least squares method is used to perform relative navigation calculations to obtain the initial relative position and initial velocity of the probe relative to the small celestial body. Then, the state transition matrix is ​​used to obtain... The method of this invention determines the relative position and velocity of the probe at the current moment; at the first moment before each orbit change, based on the relative position at the time of the orbit change, the state transition matrix, and the flight time, the velocity increment required to reach the next nominal orbit change position is determined; when the second moment before the orbit change is reached, the probe is controlled to autonomously adjust its attitude to the direction of the orbit control velocity increment, and after reaching the nominal orbit control time, the probe is controlled to autonomously start the orbit control process; after completing the orbit change, the probe is autonomously determined to be near the rendezvous segment terminal position. If it is, the probe is controlled to autonomously brake and stop the rendezvous process; otherwise, the attitude adjustment is repeated to the pointing attitude of the target small celestial body and the subsequent rendezvous flight process continues. This invention solves the problem of rendezvous with target small celestial bodies relying solely on line-of-sight measurements within a range of tens to thousands of kilometers. It can be extended to near-Earth spacecraft rendezvous scenarios relying solely on line-of-sight, greatly reducing the number and cost of hardware configurations for rendezvous sensors. Attached Figure Description

[0023] Figure 1 A flowchart of a method for autonomous rendezvous with small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering design;

[0024] Figure 2 A schematic diagram defining observability;

[0025] Figure 3 This is a schematic diagram comparing the trajectories before and after optimization. Detailed Implementation

[0026] The features and advantages of the present invention will become clearer and more apparent from the following detailed description.

[0027] The term “exemplary” as used herein means “serving as an example, embodiment, or illustration.” Any embodiment illustrated herein as “exemplary” is not necessarily to be construed as superior to or better than other embodiments.

[0028] This invention provides a method for autonomous rendezvous with small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering, such as... Figure 1 As shown, it includes the following steps:

[0029] Step 1: Before conducting the rendezvous flight, obtain the planned nominal maneuver trajectory, the speed increment for each maneuver, and the nominal orbit change time from the ground to the satellite.

[0030] The ground system plans the maneuver trajectory for the rendezvous process based on the handover point status, and uploads the planned nominal maneuver trajectory, the speed increment for each maneuver, and the nominal orbit change time to the satellite.

[0031] For the maneuver trajectory during the rendezvous process, firstly, based on the relative position and speed of the handover point determined by ground-based orbit determination, the time constraints and fuel consumption constraints of the rendezvous process, the total number of orbit changes n and the time t of each orbit change are determined. i (i = 1, ..., n). The time t for each orbital change. i The time should take into account factors such as the payload's photography and observation needs during flight, the ground pointing and telemetry needs, the on-board relative navigation calculations, and the accuracy of the long-term relative orbit extrapolation model. It can generally be set to about 1 to 2 days.

[0032] After completing the segmented planning of the trajectory, the following optimization metrics are used for optimization:

[0033] J w =wJ f +(1-w)J o (1)

[0034] J O As an observability indicator, J f For fuel consumption indicators, w is the weighting coefficient between (0,1), which needs to be repeatedly adjusted based on actual results. It can generally be taken in the range of 0.05 to 0.15.

[0035] Let the total number of track changes be n, then the fuel consumption index J in the above formula is... f Defined in the following form:

[0036]

[0037] Where X is the velocity increment for each orbital change, X = (ΔV1...ΔV)n ) T , where is the variable to be optimized; ΔV sum The maximum speed increment needs to be determined based on the available fuel consumption allocated for the actual task, and can generally be set to the range of 5 to 20 m / s.

[0038] Observability index J O Defined in the following form:

[0039]

[0040] Where cond(.) denotes the condition number of the corresponding matrix, M i Let be the observability matrix corresponding to the i-th arc segment, defined as:

[0041]

[0042] Taking a total of 8 orbital changes as an example, the observability of the entire flight trajectory is defined as follows: Figure 2 As shown.

[0043] M i The calculation is as follows: Assume that a line-of-sight measurement was performed before the i-th orbital change, and the measurement result is the line-of-sight vector n(t). i ), then n(t) i The partial derivative H with respect to the initial state variables i It is in the following form:

[0044]

[0045] Wherein, Φ(t) i (t0) represents the position and velocity of the target small celestial body relative to the small celestial body in the target small celestial body's orbital system, from t0 to t... i The state transition matrix at any given time should generally be determined based on the specific orbital characteristics of the target celestial body: for the case where the target celestial body has a circular orbit, the CW equation Φ can be used. CW (t i The description is based on t0); for the case where the target celestial body has an elliptical orbit, the TH equation Φ can be used. TH (t i The description is as follows: ,t0); The CW equation and TH equation are well-known in the field of orbital dynamics and will not be elaborated here.

[0046] r(t i ) for t i The position of the probe relative to the target celestial body in the orbital system at any given moment is determined by the initial relative position r0 and velocity v0 of the probe relative to the target celestial body at the initial moment noted on the ground, and the state transition matrix Φ(t). i ,t0) and from the initial time to t iThe velocity increment ΔV at each orbital change time i The formula for calculating i = 1, ..., n is as follows:

[0047]

[0048] n(t i Let be the vector representing the result of the i-th line-of-sight measurement under the target small celestial body's orbital system, and the formula is:

[0049]

[0050] Given the initial relative position r0 and initial velocity v0 of the probe in the target small celestial body orbit system, the optimization index described by Equation (1) is used. The optimization is performed by calling the fmincon function in Matlab, and the nominal velocity increment X = (ΔV1...ΔV0) for each orbit change is obtained. n ) T .

[0051] Trajectories before and after optimization are as follows Figure 3 As shown.

[0052] Step 2: After each orbit change, the probe will maintain its orientation toward the small celestial body and periodically measure the direction of the line of sight to the center of the target small celestial body using an optical camera.

[0053] Step 3: After each measurement of the central line of sight of the target celestial body, the satellite performs relative navigation calculations based on the accumulated line of sight measurements within the previous two orbit change arcs and the measured velocity increments of the previous two orbit change operations. Alternatively, if the number of orbit changes is less than two, the satellite performs relative navigation calculations based on the accumulated line of sight measurements from the handover point and the measured velocity increments of each orbit change operation. The initial relative position r0 and initial velocity v0 of the probe relative to the celestial body are obtained by solving the calculations. Then, the relative position and velocity r(t) and v(t) of the probe at the current moment are obtained using the state transition matrix.

[0054] The specific solution process is as follows.

[0055] (a) Establishing an optimization index based on the least squares method

[0056] Let at time t i The direction vector of the target small celestial body, measured by the optical camera, is n. i,mea Let i = 1, ..., m; the relative position and velocity at the initial time t0 to be solved are: The prior values ​​of the initial relative position and velocity are The relative navigation solution then uses the least squares method to optimize the index in the following form:

[0057]

[0058] Where m represents the number of line-of-sight measurements; σ i This represents the measurement error of the line-of-sight vector at the target center. The setting of this value should take into account the measurement error of the actual optical sensor on the line of sight, and it can generally be taken as 0.001° to 0.005°. Let R be the prior error diagonal matrix of the initial relative position and velocity, where R is the prior error matrix of the initial relative position and velocity. 0_err V represents the prior error of the relative position. 0_err n represents the prior error of relative velocity; i,pre (x0) represents the measurement time t obtained by extrapolation based on the initial value x0. i The corresponding line-of-sight vector prediction results.

[0059] Prior to the second orbit change, ground-based orbit determination results were used to determine the prior information x. 0,pre At this time R 0_err The determination should be based on the actual orbit measurement error; after the second orbit change, the solution results of on-board relative navigation are used as prior information x. 0,pre At this time R 0_err It should be set according to the proportional error of the relative position, and is generally set to within 10% of the current position navigation result. For V 0_err It should generally be greater than the relative velocity measurement track error, and can generally be set in the range of 0.1 to 0.3 m / s.

[0060] (b) Initial state solution based on least squares method

[0061] Find the optimal solution x0 that minimizes equation (8). Suppose that equation (8) only holds after perturbation δx0 around a given guessed value x0, then we can obtain:

[0062]

[0063] Where Q is an orthogonal matrix and R is an upper triangular matrix. To represent the line-of-sight measurement error, a 3×3 diagonal matrix is ​​defined as follows: (QR decomposition form of the matrix is ​​not provided in the original text.)

[0064]

[0065] For equation (9), considering that R is an upper triangular matrix, we can calculate from the last element back to obtain all the element values ​​of δx0, thereby reducing the amount of calculation.

[0066] According to equation (9), a new estimate of x0 can be obtained:

[0067] x0=x0+δx0 (11)

[0068] Considering the influence of linearization error when calculating the partial derivative matrix using formula (9), the updated x0 is substituted back into formula (9) and iterated several times until the magnitude of δx0 is less than the set threshold.

[0069] (c) Solving the current state and predicting the line of sight

[0070] After solving for the initial state x0, the relative position and velocity at the current time t can be calculated using the state transition matrix and the velocity increments of each orbital change.

[0071] Suppose that the orbital change occurs m times within the time interval [t0, t], and the times of each orbital change are t. k k = 1,...,m, and the actual velocity increment for each orbital change in the target orbital system is ΔV. k,orbctl If k = 1, ..., m, then the formulas for the relative position r(t) and velocity v(t) at time t are as follows:

[0072]

[0073] in:

[0074]

[0075] Based on equation (12), we can obtain t i The predicted result of the gaze vector at time t is:

[0076]

[0077] Where C CB C is the direction cosine matrix of the optical camera measurement system relative to the detector body, which is generally given by ground precision measurements before launch; BI The direction cosine matrix of the detector system relative to the inertial frame is generally given directly by star-sensor measurements; C oi This is the direction cosine matrix of the small celestial body's orbital frame relative to the inertial frame. It is generally extrapolated from the initial position and velocity of the small celestial body in the inertial frame given by the satellite's orbit determination to t. i The time was calculated.

[0078] Step 4: One hour before each orbit change, the probe will autonomously conduct guidance calculations, that is, based on the relative position at the time of the orbit change, the relative orbit description model, and the flight time, determine the speed increment required to reach the next nominal orbit change position.

[0079] The velocity increment required for guidance is determined one hour prior to each orbital change, using the following method:

[0080] ΔV cmd =Φ r -1 (tf ,t orbctl )[r(t f )-Φ r (t f ,t orbctl )Φ(t orbctl ,t0)x0] (14)

[0081] Where ΔV cmd The target velocity increment calculated by the guidance law, t orbctl The time when the orbital change is about to take place is t. f For the next orbital change moment, Φ r (t f ,t orbctl ) indicates from the initial time t orbctl At the final time t f The state transition matrix Φ(t) f ,t orbctl The first 3 lines of r(t) f () indicates the relative position at the next orbital change moment. t f、 ,t orbctl and r(t) f All of these are given from the trajectory optimization results before the intersection.

[0082] variable ΔV cmd This represents the velocity increment required for guidance in the target asteroid's orbital coordinate system.

[0083] Step 5: When the second moment before the orbit change is reached, such as 30 minutes before the orbit change (closer to the orbit change time), the satellite begins to autonomously adjust its attitude to the direction of the orbit control velocity increment. After reaching the nominal orbit control time, the satellite autonomously begins the orbit control process, and at the same time, the accelerometer is used to measure the actual velocity increment of the orbit change.

[0084] 30 minutes before the nominal orbit control time, the satellite autonomously adjusts its attitude to the target orbit control direction and autonomously activates the orbit control engine to complete the velocity increment ΔV calculated in step 4 of the orbit control process. cmd During orbit control, accelerometers are used to record the actual velocity increment ΔV generated by orbit control under the target orbital system. orbctl .

[0085] Step 6: After completing the orbit change, the satellite autonomously determines whether it has reached the vicinity of the rendezvous segment's end position. If it has, the satellite autonomously brakes and stops the rendezvous process; otherwise, it repeats the attitude adjustment for the target small celestial body and continues the subsequent rendezvous flight process.

[0086] The present invention also provides a control device for autonomous rendezvous small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering, comprising:

[0087] The first module is used to obtain the planned nominal maneuver trajectory, the speed increment for each maneuver, and the nominal orbit change time from the ground before conducting rendezvous flights;

[0088] The second module is used to maintain the detector's orientation toward the small celestial body after each orbit change, and to use an optical camera to measure the direction of the center line of sight of the target small celestial body at set intervals.

[0089] The third module is used to perform relative navigation calculations using the least squares method after each measurement of the line of sight direction of the target celestial body center. This is based on the accumulated line of sight measurement results within the first two orbit change arcs and the measurement results of the velocity increments of the first two orbit change operations. Alternatively, when the number of orbit changes is less than two, the module is based on the accumulated line of sight measurement results from the handover point and the measurement results of the velocity increments of each orbit change operation. This yields the initial relative position and initial velocity of the detector relative to the celestial body. Then, the state transition matrix is ​​used to obtain the detector's current relative position and velocity.

[0090] The fourth module is used to determine the speed increment required to reach the next nominal orbit change position at the first moment before each orbit change, based on the relative position at the time of the orbit change, the state transition matrix, and the flight time.

[0091] The fifth module is used to control the detector to autonomously adjust its attitude to the direction of the orbit control speed increment at the second moment before the orbit change, and to control the detector to autonomously start the orbit control process after the nominal orbit control time is reached.

[0092] The sixth module is used to autonomously determine whether the probe has reached the vicinity of the rendezvous segment's end position after completing the orbit change. If it has, it controls the probe to brake autonomously and stop the rendezvous process; otherwise, it repeats the attitude adjustment for pointing at the target small celestial body and continues the subsequent rendezvous flight process.

[0093] Secondly, a control device for an autonomously rendezvous small celestial body, based on an integrated design of line-of-sight measurement and trajectory maneuvering, is located on the detector and includes:

[0094] One or more processors;

[0095] Storage device for storing one or more programs.

[0096] When the one or more programs are executed by the one or more processors, the one or more processors implement the autonomous rendezvous small celestial body method based on the integrated design of line-of-sight measurement and trajectory maneuvering described above.

[0097] Thirdly, a readable storage medium storing a computer program that, when executed by a processor, implements the aforementioned autonomous rendezvous method for small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering.

[0098] Fourthly, a computer program product comprising: a computer program (also referred to as code or instructions) that, when executed, performs the autonomous rendezvous method for small celestial bodies based on the integrated design of line-of-sight measurement and trajectory maneuvering described above.

[0099] In the above embodiments, implementation can be achieved, in whole or in part, through software, hardware, firmware, or any combination thereof. When implemented in software, it can be implemented, in whole or in part, as a computer program product. The computer program product includes one or more computer instructions. When the computer instructions are loaded and executed on a computer, all or part of the processes or functions described in the embodiments of this application are generated. The computer can be a general-purpose computer, a special-purpose computer, a computer network, or other programmable device. The computer instructions can be stored in a computer-readable storage medium or transmitted from one computer-readable storage medium to another. For example, the computer instructions can be transmitted from one website, computer, server, or data center to another via wired (e.g., coaxial cable, fiber optic, digital subscriber line (DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) means.

[0100] Those skilled in the art will recognize that the units and algorithm steps of the various examples described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, or a combination of computer software and electronic hardware. Whether these functions are implemented in hardware or software depends on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the described functions for each specific application, but such implementation should not be considered beyond the scope of this application.

[0101] Those skilled in the art will clearly understand that, for the sake of convenience and brevity, the specific working processes of the devices, apparatuses, and modules described above can be referred to the corresponding processes in the foregoing method embodiments, and will not be repeated here.

[0102] In the several embodiments provided in this application, it should be understood that the disclosed devices, apparatuses, and methods can be implemented in other ways. For example, the apparatus embodiments described above are merely illustrative; for instance, the division of modules is only a logical functional division, and other division methods may be used in actual implementation.

[0103] Furthermore, if the aforementioned functions are implemented as software functional units and sold or used as independent products, they can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of this application, in essence, or the part that contributes to the prior art, or a portion of the technical solution, can be embodied in the form of a software product. This computer software product is stored in a storage medium and includes several instructions to cause a computer device (which may be a personal computer, server, or network device, etc.) to execute all or part of the steps of the methods described in the various embodiments of this application. The aforementioned storage medium includes various media capable of storing program code, such as USB flash drives, portable hard drives, read-only memory (ROM), random access memory (RAM), magnetic disks, or optical disks.

[0104] The present invention has been described in detail above with reference to specific embodiments and exemplary examples; however, these descriptions should not be construed as limiting the present invention. Those skilled in the art will understand that various equivalent substitutions, modifications, or improvements can be made to the technical solutions and embodiments of the present invention without departing from the spirit and scope of the invention, and all such modifications and improvements fall within the scope of the present invention. The scope of protection of the present invention is defined by the appended claims.

[0105] The contents not described in detail in this specification are common knowledge to those skilled in the art.

Claims

1. A method for autonomous rendezvous with small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering, characterized in that, include: Before conducting a rendezvous flight, obtain the planned nominal maneuver trajectory, the speed increment for each maneuver, and the nominal orbit change time from the ground. After each orbit change, the detector maintains its orientation toward the small celestial body, and at set intervals, the optical camera measures the direction of the center line of sight of the target small celestial body. After each measurement of the line of sight to the center of the target celestial body, the detector's initial relative position and initial velocity relative to the celestial body are calculated using the least squares method, based on the accumulated line of sight measurements within the first two orbit change arcs and the measured velocity increments of the first two orbit change operations. Alternatively, if there are fewer than two orbit change operations, the detector's initial relative position and initial velocity are calculated based on the accumulated line of sight measurements from the handover point and the measured velocity increments of each orbit change operation. Then, the detector's current relative position and velocity are obtained using the state transition matrix. At the moment before each orbit change, the velocity increment required to reach the next nominal orbit change position is determined based on the relative position at the moment of orbit change, the state transition matrix, and the flight time. At the second moment before the orbit change, the probe is controlled to autonomously adjust its attitude to the direction of the orbit control speed increment, and after reaching the nominal orbit control time, the probe is controlled to autonomously start the orbit control process. After completing the orbit change, the probe autonomously determines whether it has reached the vicinity of the end position of the rendezvous segment. If it has, it controls the probe to brake autonomously and stop the rendezvous process. Otherwise, it repeats the attitude adjustment for pointing to the target small celestial body and continues the subsequent rendezvous flight process. The speed increment for each maneuver is obtained through optimization in the following manner: The following optimization metrics will be optimized: （1） in J O As an observability indicator, J f As a fuel consumption indicator, w The weighting coefficients are those between (0,1); Let the total number of orbit changes be... n Then the fuel consumption index in the above formula J f Defined in the following form: （2） in X Speed ​​increment for each orbit change , where is the variable to be optimized; This represents the maximum speed increment. Observability indicators J O Defined in the following form: （3） in cond (.) indicates that the condition number of the corresponding matrix is ​​calculated. M i For the first i The observability matrix corresponding to the arc segment is defined as follows: i =1,…, n -2（4） Assuming in the first i A line-of-sight measurement was performed before the next trajectory change, and the result was the line-of-sight vector. n ( t i ),but n ( t i Partial derivatives with respect to the initial state variables H i It is in the following form: （5） in, The position and velocity of the target small celestial body relative to the small celestial body in the target small celestial body orbital system are from t 0 to t i The state transition matrix at time t; for t i The position of the probe relative to the target celestial body in the orbital system at any given moment is determined by the initial relative position of the probe to the target celestial body as indicated on the ground. r 0 and initial velocity v 0. State transition matrix and from the initial moment to t i Speed ​​increments at each orbital change time , i =1,..., n The calculation yields the following formula: （6） n ( t i ) for the target small celestial body orbit system i The vector result of the secondary line of sight measurement is given by the following formula: （7） Given the initial relative position of the probe in the target small celestial body's orbital system r 0 and speed v 0. The optimization index described by equation (1) is optimized to obtain the nominal speed increment for each track change.

2. The method for autonomous rendezvous of small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering design according to claim 1, characterized in that, The nominal maneuver trajectory includes the position of the probe relative to the small celestial body in the target small celestial body's orbital system during each orbital change during the rendezvous process.

3. The method for autonomous rendezvous of small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering design according to claim 1, characterized in that, The initial relative position and initial velocity of the detector with respect to the small celestial body were obtained in the following manner: Establish an optimization index based on the least squares method: （8） in, Initial relative position r 0 and speed v 0, m Number of times to measure line of sight; These are the prior values ​​for the initial relative position and velocity. The measurement error is the line-of-sight vector at the target center. Let be the prior error diagonal matrix of the initial relative position and velocity, where R 0_err This represents the prior error of the relative position. V 0_err This represents the prior error of the relative velocity; Based on the initial value x Measurement time obtained by extrapolation of 0 t i The corresponding line-of-sight vector prediction results; n i,mea For a moment t i The orientation vector of the target small celestial body obtained by the optical camera; Solve for the optimization index Minimize the optimal solution x 0, specifically: assuming a given guess value x Perturbation is needed near 0. Only when the latter equation (8) is true can we obtain: （9） in Q It is an orthogonal matrix. R It is an upper triangular matrix. To represent the line-of-sight measurement error, a 3×3 diagonal matrix is ​​defined as follows: (QR decomposition form of the matrix is ​​not provided in the original text.) According to equation (9), we get x A new estimate of 0: （10） The updated x Substitute 0 back into equation (8) and iterate repeatedly until... The modulus is less than the set threshold.

4. The method for autonomous rendezvous of small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering design according to claim 3, characterized in that, The relative position and velocity of the detector at the current moment are obtained in the following way: Set in the time interval Number of internal orbit changes m The times of each orbital change are as follows: t k , k =1,..., m And the actual velocity increment for each orbital change in the target orbital system is , k =1,..., m ,but t Relative position of time and speed The formula is as follows: （12） in: 。 5. The method for autonomous rendezvous of small celestial bodies based on integrated line-of-sight measurement and trajectory maneuvering design according to claim 4, characterized in that, The speed increment required to reach the next nominal track change position is obtained as follows: The required velocity increment for guidance is determined using the following method: （14） Among the variables This represents the velocity increment required for guidance in the target asteroid's orbital coordinate system; For the target speed increment, t orbctl The time for the orbital change is about to begin. t f For the next orbital change moment, Indicates from the initial time t orbctl To the end time t f State transition matrix The first 3 lines, This indicates the relative position at the next orbital change moment.

6. A control device for autonomous rendezvous small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering, characterized in that, Located on the detector, including: One or more processors; Storage device for storing one or more programs. When the one or more programs are executed by the one or more processors, the one or more processors implement the autonomous rendezvous small celestial body method based on the integrated design of line-of-sight measurement and trajectory maneuvering as described in any one of claims 1 to 5.

7. A readable storage medium, characterized in that, It stores a computer program that, when executed by a processor, implements the autonomous rendezvous method for small celestial bodies based on the integrated design of line-of-sight measurement and trajectory maneuvering as described in any one of claims 1 to 5.

8. A computer program product, characterized in that, The computer program product includes: a computer program that, when the computer program is run, executes the autonomous rendezvous method for small celestial bodies based on the integrated design of line-of-sight measurement and trajectory maneuvering as described in any one of claims 1 to 5.

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