Small celestial body autonomous rendezvous method based on sight line measurement and trajectory maneuvering integrated design

Abstract

The invention provides a small celestial body autonomous rendezvous method based on sight line measurement and trajectory maneuvering integrated design. The method comprises the following steps: acquiring a planned nominal maneuvering trajectory and orbital transfer information noted on the ground; keeping the direction of the detector to the direction of the small celestial body after each orbital transfer, and measuring the central sight line direction of the target small celestial body; calculating the initial relative position and initial speed of the detector relative to the small celestial body and the relative position and speed of the detector at the current moment according to the sight line measurement result and the measurement result of the orbital transfer speed increment; according to the relative position at the orbital transfer moment, the state transition matrix and the flight time, determining the speed increment required for reaching the next nominal orbital transfer position; controlling the detector to autonomously adjust the attitude to the orbit control speed increment direction, and controlling the detector to autonomously start the orbit control process after the nominal orbit control time is reached; after orbital transfer is completed, whether the detector arrives near the terminal position of the intersection section or not is autonomously judged.

Description

Technical Field

[0001] The present invention belongs to the field of navigation and guidance in deep space small celestial body detection, and in particular relates to a method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering. Background Art

[0002] For small celestial body exploration missions, considering the limited orbit determination accuracy of the target small celestial body and the influence of the orbit determination accuracy of the detector, the orbit determination error of the detector relative to the small celestial body can reach hundreds of kilometers. For safety reasons, the ground can generally only directly guide the detector to a distance of hundreds to thousands of kilometers from the small celestial body, and then the GNC subsystem on the satellite can generally carry out relative navigation based on the measurement results of the relative navigation sensor, and autonomously approach to a distance of tens of kilometers closer to the small celestial body. Generally, this process is called the process of rendezvous with the small celestial body.

[0003] Due to weight and power limitations, the range of the onboard ranging radar can generally only reach a measurement range of tens of kilometers at most. On the other hand, even if the ranging radar can reach a range of hundreds of kilometers, it is difficult to accurately point to the target small celestial body with a diameter of only hundreds of meters. Therefore, for long-distance ranges beyond tens of kilometers, the detector generally only has the means to use optical cameras to measure the central line of sight direction of the target small celestial body, and cannot directly obtain the distance information of the target.

[0004] According to the observability theory of relative navigation, it is impossible to obtain complete information of relative position only by line of sight measurement; in order to obtain complete observability of relative position, it is necessary to combine it with certain trajectory maneuvers. For trajectory maneuvers, the larger the maneuver amplitude, the stronger the navigation observability is generally obtained, but on the other hand, the more fuel is consumed, which requires comprehensive consideration of the optimization of fuel consumption and observability.

[0005] Therefore, it is necessary to provide a method for autonomous rendezvous with small celestial bodies to solve the problem of rendezvous with target small celestial bodies within a range of tens to thousands of kilometers relying only on line-of-sight measurement. Summary of the invention

[0006] In view of the conditions in small body detection missions where only the central line of sight of a target small body can be measured within a range of tens to thousands of kilometers, the present invention provides a method for autonomous rendezvous with small bodies based on an integrated design of line of sight measurement and trajectory maneuvering, thereby solving the problem of rendezvous with a target small body relying solely on line of sight measurement.

[0007] The technical solution provided by the present invention is as follows:

[0008] In the first aspect, a method for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuvering comprises:

[0009] Before conducting the rendezvous flight, obtain the planned nominal maneuver trajectory, speed increment for each maneuver and nominal change of trajectory time recorded on the ground;

[0010] After each orbit change, the probe is kept pointing toward the small celestial body, and the central line of sight direction of the target small celestial body is measured using an optical camera at set intervals;

[0011] After each measurement of the line of sight direction of the target small celestial body is completed, the initial relative position and initial velocity of the probe relative to the small celestial body are obtained by using the least square method for relative navigation solution based on the accumulated line of sight measurement results within the arc segments of the previous two orbit changes and the measurement results of the velocity increments of the previous two orbit changes, or when the number of orbit changes is less than two, the accumulated line of sight measurement results starting from the handover point and the measurement results of the velocity increments of each orbit change, and then the relative position and velocity of the probe at the current moment are obtained by using the state transfer matrix;

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

[0013] When reaching the second moment before the orbit change, the probe is controlled to autonomously adjust its attitude to the orbit control velocity increment direction, and after reaching the nominal orbit control time, the probe is controlled to autonomously start the orbit control process;

[0014] After completing the orbit change, it will autonomously determine whether the probe has reached the terminal position of the rendezvous segment. If it has, it will control the probe to brake autonomously and stop the rendezvous process. Otherwise, it will repeat the attitude adjustment to point to the target small celestial body and continue the subsequent rendezvous flight process.

[0015] Second, a control device for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuvering, located on the probe, includes:

[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 autonomous rendezvous with small celestial bodies method based on integrated design of line of sight measurement and trajectory maneuvering as described in the first aspect.

[0019] In a third aspect, a readable storage medium stores a computer program, which, when executed by a processor, implements the method for autonomous rendezvous with small celestial bodies based on the integrated design of line of sight measurement and trajectory maneuvering as described in the first aspect.

[0020] In a fourth aspect, a computer program product is provided, comprising: a computer program, which, when executed, executes the autonomous rendezvous with small celestial bodies method based on integrated design of line of sight measurement and trajectory maneuvering as described in the first aspect.

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

[0022] The present invention provides a method for autonomous rendezvous with a small celestial body based on an integrated design of line of sight measurement and trajectory maneuvering. Before carrying out the rendezvous flight, a planned nominal maneuvering trajectory, a speed increment for each maneuver, and a nominal orbit change time recorded on the ground are obtained; after each orbit change, the direction of the probe to the small celestial body is kept pointed, and the central line of sight direction of the target small celestial body is measured at set intervals using an optical camera; after each measurement of the central line of sight direction of the target small celestial body is completed, based on the accumulated line of sight measurement results within the arc segments of the previous two orbit changes and the measurement results of the speed increments of the previous two orbit changes, or when the number of orbit changes is less than two, based on the accumulated line of sight measurement results starting from the handover point and the measurement results of the speed increments of each orbit change, the least squares method is used to perform relative navigation solution to obtain the initial relative position and initial velocity of the probe relative to the small celestial body, and then the state transfer matrix is ​​used to obtain The relative position and speed of the probe at the current moment; at the first moment before each orbit change, the speed increment required to reach the next nominal orbit change position is determined according to the relative position at the orbit change moment, the state transfer matrix, and the flight time; when reaching 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, it is autonomously determined whether the probe has reached the vicinity of the terminal position of the rendezvous segment, and if so, the probe is controlled to autonomously brake and stop the rendezvous process, otherwise, the attitude is repeatedly adjusted to the pointing attitude of the target small celestial body and the subsequent rendezvous flight process is continued; the method of the present invention solves the problem of rendezvous with the target small celestial body within a range of tens to thousands of kilometers by relying only on line-of-sight measurement, and can be extended to the rendezvous of near-Earth spacecraft relying only on line of sight, greatly reducing the number and cost of hardware configurations of the rendezvous sensor. BRIEF DESCRIPTION OF THE DRAWINGS

[0023] Figure 1 This is a flow chart of a method for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuvering;

[0024] Figure 2 Define schematics for observability;

[0025] Figure 3 Schematic diagram for comparing trajectories before and after optimization. DETAILED DESCRIPTION

[0026] The following detailed description of the present invention will make the features and advantages of the present invention more clear and explicit.

[0027] The word “exemplary” is used exclusively herein to mean “serving as an example, example, or illustration.” Any embodiment described herein as “exemplary” is not necessarily to be construed as preferred or advantageous over other embodiments.

[0028] The present invention provides a method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering, such as Figure 1 As shown, the following steps are included:

[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 injected into the satellite from the ground.

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

[0031] For the maneuvering trajectory of the rendezvous process, firstly, the total number of orbit changes n and the time t of each orbit change in the rendezvous process are determined according to the relative position and speed of the handover point determined by the ground orbit measurement, the time constraint of the rendezvous, and the fuel consumption constraint. i (i=1,...n). The time t for each track change i The impact of factors such as the payload photography and observation needs, ground pointing and control needs, on-board relative navigation solution, and long-term relative orbit extrapolation model accuracy during the flight should be comprehensively considered, and it can generally be set to about 1 to 2 days.

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

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

[0034] Among them J O is the observability index, J f is the fuel consumption index, w is the weighting coefficient between (0,1), which needs to be repeatedly debugged according to the actual results, and can generally be taken in the range of 0.05 to 0.15.

[0035] Assuming the total number of orbit changes is n, the fuel consumption index J in the above formula is f The definition is as follows:

[0036]

[0037] Where X is the speed increment for each track change X = (ΔV1...ΔVn ) T , is the variable to be optimized; ΔV sum It is the maximum value of speed increment, which needs to be determined according to the available fuel consumption allocated for the actual task and can generally be set in the range of 5 to 20 m / s.

[0038] Observability Index J O It is defined as follows:

[0039]

[0040] Where cond(.) means to find the condition number of the corresponding matrix, M i is the observability matrix corresponding to the i-th solved arc segment, which is defined as:

[0041]

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

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

[0044]

[0045] Among them, Φ(t i ,t0) is the position and velocity of the target small celestial body relative to the small celestial body in the orbital system from t0 to t i The state transfer matrix at time should generally be determined according to the specific orbital characteristics of the target small celestial body: For the case where the target celestial body is a circular orbit, the CW equation Φ can be used CW (t i ,t0) for description; for the case where the target celestial body is in an elliptical orbit, the TH equation Φ can be used TH (t i ,t0) is described; the CW equation and TH equation are common knowledge in the field of orbital dynamics and will not be described here.

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

[0047]

[0048] n(t i ) is the i-th line of sight measurement result vector under the orbit system of the target small celestial body, and the formula is:

[0049]

[0050] Given the initial relative position r0 and initial velocity v0 of the probe in the orbital system of the target small celestial body, the optimization index described by formula (1) is used to call the fmincon function of Matlab for optimization, and the nominal velocity increment X = (ΔV1...ΔV n ) T .

[0051] The trajectories before and after optimization are as follows Figure 3 shown.

[0052] Step 2: After each orbit change, the probe will maintain its direction toward the small celestial body and use an optical camera to measure the central line of sight direction of the target small celestial body at regular intervals.

[0053] Step 3. After each measurement of the central line of sight direction of the target small celestial body is completed, the satellite performs relative navigation solution based on the least squares method according to the accumulated line of sight measurement results within the arc segments of the previous two orbit changes and the measurement results of the velocity increments of the previous two orbit changes, or when the number of orbit changes is less than two, the accumulated line of sight measurement results starting from the handover point and the measurement results of the velocity increments of each orbit change, to obtain the initial relative position r0 and initial velocity v0 of the probe relative to the small celestial body, and then uses the state transfer matrix to obtain the relative position and velocity r(t), v(t) of the probe at the current moment.

[0054] The specific solution process is as follows.

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

[0056] Assume that at time t i The direction vector of the target small celestial body measured by the optical camera is n i,mea , i=1,...,m; the relative position and velocity at the initial time t0 to be solved are The initial relative position and velocity priors are The relative navigation solution uses the following least squares optimization index:

[0057]

[0058] Where m is the number of line of sight measurements; σ i The measurement error of the target center line of sight direction vector should be considered in the setting of the value. The measurement error of the line of sight of the actual optical sensor should be taken into account. It can generally be taken as 0.001°~0.005°; is the a priori error diagonal matrix of the initial relative position and velocity, where R 0_err is the prior error of relative position, V 0_err is the prior error of relative velocity; n i,pre (x0) is the measurement time t obtained by extrapolation based on the initial value x0 i The corresponding sight line vector prediction result.

[0059] Before the second orbit change, the ground orbit measurement results are used to determine the prior information x 0,pre , at this time R 0_err It should be determined according to the actual orbit determination error; after the second orbit change, the solution of the on-board relative navigation is used as the prior information x 0,pre , at this time R 0_err It should be set in the form of proportional error of relative position, and can generally be set within 10% of the navigation result of the current position. 0_err , should generally be greater than the relative speed 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] For equation (8), find the optimal solution x0 that minimizes it. Assume that equation (8) can only be established after perturbation δx0 near the given guess value x0, we can get:

[0062]

[0063] Where Q is an orthogonal matrix, R is an upper triangular matrix, The 3×3 dimensional diagonal matrix representing the line of sight measurement error is defined as the QR decomposition form of the following matrix:

[0064]

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

[0066] According to formula (9), we can get the new estimated value of x0:

[0067] x0=x0+δx0 (11)

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

[0069] (c) Current state solution and line of sight prediction

[0070] After solving the initial state x0, the relative position and velocity at the current time t can be calculated through the state transfer matrix and the velocity increment of each track change.

[0071] Suppose the number of track changes in the time interval [t0, t] is m, and the time of each track change is t k , k = 1, ..., m, and the actual velocity increment of each orbit change in the target orbit system is ΔV k,orbctl , 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 formula (12), we can get t i The prediction result of the sight vector at the moment is:

[0076]

[0077] Among them C CB is the direction cosine matrix of the optical camera measurement system relative to the detector body, which is generally given by ground precision measurement before launch; C BI is the direction cosine matrix of the detector system relative to the inertial system, which is generally directly given by star-sensing measurements; C oi is the direction cosine matrix of the orbital system of the small celestial body relative to the inertial system, which is usually extrapolated from the initial position and velocity of the small celestial body in the inertial system given by the satellite based on the measured orbit to t i Calculated at the moment.

[0078] Step 4: At the first moment before each orbit change, such as 1 hour before, the probe will autonomously carry out guidance calculations, that is, determine the speed increment required to reach the next nominal orbit change position based on the relative position at the time of orbit change, the relative orbit description model, and the flight time.

[0079] At the first moment before each orbit change, such as 1 hour before, the velocity increment required for guidance is determined in the following form:

[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 is the target velocity increment calculated by the guidance law, t orbctl The time when the track change is about to be carried out, t f is the next track change moment, Φ r (t f ,t orbctl ) represents the time from the initial moment t orbctl At the end time t f The state transfer matrix Φ(t f ,t orbctl ) first 3 rows, r(t f ) represents the relative position at the next track change time. f、 ,t orbctl and r(t f ) are given by the trajectory optimization results before the intersection.

[0082] Variable ΔV cmd represents the velocity increment required for guidance in the orbital coordinate system of the target asteroid.

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

[0084] When the satellite reaches 30 minutes before the nominal orbit control time, it will autonomously adjust its attitude to the orbit control target direction and autonomously start the orbit control engine to complete the velocity increment ΔV calculated in step 4. cmd During the orbit control process, an accelerometer is used to record the actual velocity increment ΔV generated by the orbit control in the target orbit system. orbctl .

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

[0086] The present invention also provides a control device for autonomous rendezvous with a small celestial body 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 of each maneuver and the nominal trajectory change time recorded on the ground before conducting the rendezvous flight;

[0088] The second module is used to keep the probe pointing to the small celestial body after each orbit change, and to measure the central line of sight direction of the target small celestial body using an optical camera at set intervals;

[0089] The third module is used to perform relative navigation calculation using the least square method after completing the measurement of the line of sight direction of the center of the target small celestial body each time, based on the accumulated line of sight measurement results within the arc segments of the previous two orbit changes and the measurement results of the velocity increments of the previous two orbit changes, or when the number of orbit changes is less than two, based on the accumulated line of sight measurement results starting from the handover point and the measurement results of the velocity increments of each orbit change, to obtain the initial relative position and initial velocity of the probe relative to the small celestial body, and then use the state transfer matrix to obtain the relative position and velocity of the probe at the current moment;

[0090] The fourth module is used to determine the speed increment required to reach the next nominal track change position at the first moment before each track change according to the relative position at the track change moment, the state transfer matrix, and the flight time;

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

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

[0093] Second, a control device for autonomous rendezvous with a small celestial body based on integrated design of line-of-sight measurement and trajectory maneuvering, located on the probe, includes:

[0094] one or more processors;

[0095] a 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 above-mentioned method for autonomous rendezvous with small celestial bodies based on integrated design of line of sight measurement and trajectory maneuvering.

[0097] In a third aspect, a readable storage medium stores a computer program, which, when executed by a processor, implements the above-mentioned method for autonomous rendezvous with small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering.

[0098] In a fourth aspect, a computer program product is provided, comprising: a computer program (also referred to as code, or instruction), which, when executed, executes the above-mentioned method for autonomous rendezvous with small celestial bodies based on integrated design of line-of-sight measurement and trajectory maneuvering.

[0099] In the above embodiments, it can be implemented in whole or in part by software, hardware, firmware or any combination thereof. When implemented using software, it can be implemented in whole or in part in the form of 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, the process or function described in the embodiment of the present application is generated in whole or in part. 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 computer-readable storage medium, for example, the computer instructions can be transmitted from a website site, computer, server or data center by wired (e.g., coaxial cable, optical fiber, digital subscriber line (digital subscriber line, DSL)) or wireless (e.g., infrared, wireless, microwave, etc.) mode to another website site, computer, server or data center.

[0100] Those of ordinary skill in the art will appreciate that the units and algorithm steps of each example 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 performed in hardware or software depends on the specific application and design constraints of the technical solution. Professional and technical personnel can use different methods to implement the described functions for each specific application, but such implementation should not be considered to be beyond the scope of this application.

[0101] Those skilled in the art can clearly understand that, for the convenience and brevity of description, the specific working processes of the above-described equipment, devices and modules can refer to the corresponding processes in the aforementioned 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 device embodiments described above are only illustrative, for example, the division of the modules is only a logical function division, and there may be other division methods in actual implementation.

[0103] In addition, if the function is implemented in the form of a software functional unit and sold or used as an independent product, it can be stored in a computer-readable storage medium. Based on this understanding, the technical solution of the present application can be essentially or partly embodied in the form of a software product that contributes to the prior art. The computer software product is stored in a storage medium, including several instructions for a computer device (which can be a personal computer, a server, or a network device, etc.) to perform all or part of the steps of the method described in each embodiment of the present application. The aforementioned storage medium includes: U disk, mobile hard disk, read-only memory (ROM), random access memory (RAM), disk or optical disk, etc., and other media that can store program codes.

[0104] The present invention has been described in detail above in conjunction with specific implementations and exemplary examples, but these descriptions cannot be understood as limiting the present invention. Those skilled in the art understand that, without departing from the spirit and scope of the present invention, a variety of equivalent substitutions, modifications or improvements may be made to the technical solution of the present invention and its implementation methods, all of which fall within the scope of the present invention. The scope of protection of the present invention shall be subject to the attached claims.

[0105] The contents not described in detail in the specification of the present invention belong to the common knowledge of those skilled in the art.

Claims

1. A method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering, characterized in that: include: Before conducting the rendezvous flight, obtain the planned nominal maneuver trajectory, speed increment for each maneuver and nominal change of trajectory time recorded on the ground; After each orbit change, the probe is kept pointing toward the small celestial body, and the central line of sight direction of the target small celestial body is measured using an optical camera at set intervals; After each measurement of the line of sight direction of the target small celestial body is completed, the initial relative position and initial velocity of the probe relative to the small celestial body are obtained by using the least square method for relative navigation solution based on the accumulated line of sight measurement results within the arc segments of the previous two orbit changes and the measurement results of the velocity increments of the previous two orbit changes, or when the number of orbit changes is less than two, the accumulated line of sight measurement results starting from the handover point and the measurement results of the velocity increments of each orbit change, and then the relative position and velocity of the probe at the current moment are obtained by using the state transfer matrix; At the first moment before each orbit change, the speed increment required to reach the next nominal orbit change position is determined based on the relative position at the orbit change moment, the state transfer matrix, and the flight time; At the second moment before the orbit change, the probe is controlled to autonomously adjust its attitude to the orbit control velocity increment direction, and after reaching the nominal orbit control time, the probe is controlled to autonomously start the orbit control process; After completing the orbit change, it will autonomously determine whether the probe has reached the terminal position of the rendezvous segment. If it has, it will control the probe to brake autonomously and stop the rendezvous process. Otherwise, it will repeat the attitude adjustment to point to the target small celestial body and continue the subsequent rendezvous flight process.

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

3. The method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering according to claim 1, characterized in that: The speed increment of each maneuver is optimized as follows: Optimize the following optimization indicators: J w =wJ f +(1-w)J o (1) Among them J O is the observability index, J f is the fuel consumption index, w is the weighting coefficient between (0,1); Assuming the total number of orbit changes is n, the fuel consumption index J in the above formula is f The definition is as follows: Where X is the speed increment for each track change X = (ΔV1 ... ΔV n ) T , is the variable to be optimized; ΔV sum is the maximum value of speed increment; Observability Index J O It is defined as follows: Where cond(.) means to find the condition number of the corresponding matrix, M i is the observability matrix corresponding to the i-th solved arc segment, which is defined as: Assume that a line of sight measurement is performed before the i-th track change, and the measurement result is the line of sight vector n(t i ), then n(t i ) with respect to the partial derivative of the initial state variable H i In the following form: Among them, Φ(t i ,t0) is the position and velocity of the target small celestial body relative to the small celestial body in the orbital system from t0 to t i The state transfer matrix at time; r(t i ) is t i The position of the probe relative to the target small celestial body in the orbital system of the target small celestial body at the time is calculated by the initial relative position r0 and initial velocity v0 of the probe relative to the target small celestial body at the initial time recorded on the ground, and the state transfer matrix Φ(t i ,t0) and from the initial time to t i The speed increment ΔV of each track change at the time i , i=1,...,n is calculated, the formula is as follows: n(t i ) is the i-th line of sight measurement result vector under the orbit system of the target small celestial body, and the formula is: Given the initial relative position r0 and velocity v0 of the probe in the orbital system of the target small celestial body, the optimization index described by equation (1) is optimized to obtain the optimized nominal velocity increment X for each orbit change = (ΔV1...ΔV n ) T .

4. The method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering according to claim 3, characterized in that: The initial relative position and initial velocity of the probe relative to the small celestial body are obtained as follows: Establish an optimization index based on the least squares method: in, is the initial relative position r0 and velocity v0, m is the number of line of sight measurements; is the prior value of the initial relative position and velocity, σ i is the measurement error of the sight direction vector of the target center; is the a priori error diagonal matrix of the initial relative position and velocity, where R 0_err is the prior error of relative position, V 0_err is the prior error of relative velocity; n i,pre (x0) is the measurement time t obtained by extrapolation based on the initial value x0 i The corresponding sight vector prediction result; Solve the optimal solution x0 that minimizes the optimization index J(x0). Specifically, suppose that around the given guess value x0, perturbation δx0 is required to satisfy equation (8), and we get: Where Q is an orthogonal matrix, R is an upper triangular matrix, The 3×3 dimensional diagonal matrix representing the line of sight measurement error is defined as the QR decomposition form of the following matrix: According to formula (9), we get the new estimated value of x0: x0=x0+δx0 (10) Substitute the updated x0 into equation (8) again and iterate repeatedly until the modulus of δx0 is less than the set threshold.

5. The method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering according to claim 4, characterized in that: The relative position and speed of the detector at the current moment are obtained in the following way: Suppose the number of track changes in the time interval [t0, t] is m, and the time of each track change is t k , k = 1, ..., m, and the actual velocity increment of each orbit change in the target orbit system is ΔV k,orbctl , k = 1, ..., m, then the formulas for the relative position r(t) and velocity v(t) at time t are as follows: in:

6. The method for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering according to claim 1, characterized in that: The speed increment required to reach the next nominal track change position is obtained by: The velocity increment required for guidance is determined using the following form: ΔV cmd =Φ r -1 (t f ,t orbctl )[r(t f )-Φ r (t f ,t orbctl )Φ(t orbctl ,t0)x0] (14) The variable ΔV cmd represents the velocity increment required for guidance in the orbital coordinate system of the target asteroid; ΔV cmd is the target speed increment, t orbctl The time when the track change is about to be carried out, t f is the next track change moment, Φ r (t f ,t orbctl ) represents the time from the initial moment t orbctl At the end time t f The state transfer matrix Φ(t f ,t orbctl ) first 3 rows, r(t f ) indicates the relative position at the next track change time.

7. A control device for autonomous rendezvous with a small celestial body based on integrated design of line of sight measurement and trajectory maneuvering, characterized in that: Located on the detector, including: one or more processors; a 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 with small celestial bodies method based on integrated design of line of sight measurement and trajectory maneuvering as described in one of claims 1 to 6.

8. A readable storage medium, characterized in that: A computer program is stored thereon, and when the program is executed by a processor, the method for autonomous rendezvous with small celestial bodies based on an integrated design of line-of-sight measurement and trajectory maneuvering as described in one of claims 1 to 6 is implemented.

9. A computer program product, characterized in that The computer program product comprises: a computer program, which, when being executed, executes the autonomous rendezvous with small celestial bodies method based on integrated design of line-of-sight measurement and trajectory maneuvering as claimed in any one of claims 1 to 6.

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