A maneuvering guidance method for high orbit satellite close-in approach

Abstract

The application discloses a kind of high-orbit satellite close approach maneuvering guidance methods.Probe the position coordinate information of non-cooperative target, calculate the relative distance Δx of current target by own position coordinate information;Suppose the minimum approach distance index of this approach task is δ, judge the size of Δx and δ, if it does not satisfy the condition of Δx < δ, then the current position does not satisfy the approach task index, enter the cycle maneuvering guidance;Along the line-of-sight direction, apply thrust and continue for a certain time, the thrust size is determined by satellite thrust system performance, the thrust duration is determined by the computing unit of satellite or the computing power of ground control station;After a certain time of thrust, satellite reaches new position, detects target position coordinate again and makes relative distance calculation;Repeat the above cycle process until Δx < δ condition is satisfied.The whole process of the application is that engine is continuously started, which can reduce the number of engine start and stop, and can make the thrust utilization efficiency highest, which can effectively reduce the time required for approach task.

Description

TECHNICAL FIELD

[0001] The application belongs to the technical field of aerospace engineering, and particularly relates to a high-orbit satellite close-in maneuvering guidance method. BACKGROUND

[0002] With the rapid development of aerospace technology, various spacecrafts are more and more widely used in the military and civil technology of various space powers. As an important platform for space environment utilization research, the spacecraft has become an important research object in space science, and the spacecraft orbit maneuvering problem related thereto has become a research hotspot in the aerospace industry. Among them, the close-in problem of two spacecrafts is of great significance, for example, the service satellite rescues the client satellite, the space debris cleaning, and the emergency detection service, etc. all need to use the spacecraft close-in technology.

[0003] At present, the traditional close-in maneuvering method (referring to the relative distance between two spacecrafts within one hundred meters) usually performs orbit transfer at special points, and most of the time is spent in waiting for the transfer window, so it needs a long time; the requirement for the satellite's ability is high, which is reflected in the high precision requirement for the satellite's thrust system and detection system; during the maneuvering process, especially in the last period before the close-in ends, the engine starts and stops frequently. SUMMARY

[0004] The purpose of the application is to provide a high-orbit satellite close-in maneuvering guidance method.

[0005] The technical solution for realizing the purpose of the application is as follows: a high-orbit satellite close-in maneuvering guidance method, according to the working condition at the beginning of the task and the close-in requirement index, a task sequence is generated, the task sequence includes target position information R tar , thrust vector T, and thrust duration t cha , specifically including the following steps:

[0006] Step (1): line-of-sight vector calculation

[0007] In the earth inertial coordinate system, according to the detected target position coordinates R tar and the position coordinates R cha of itself, the relative position vector R between the target and itself is solved, and the unit vector is the line-of-sight vector e, so that the target information sequence related to the position vector and the line-of-sight direction can be obtained, and the above variable is expressed by the following formula:

[0008] R = R tar -R cha Formula (1)

[0009]

[0010] Step (2): Calculate the relative distance

[0011] The relative distance Δx between the satellite and the target at the current moment is calculated as follows:

[0012] △x=||R|| Equation (3)

[0013] Given that the target for the approach mission is δ, that is, the relative distance between the satellite and the non-cooperative target is less than δ when the mission ends, if Δx does not satisfy equation (4), then proceed to step (3); if Δx satisfies equation (4), then the approach mission ends and the mission is completed.

[0014] △x<δ Equation (4)

[0015] Step (3): Calculation of maximum thrust

[0016] Applying maximum thrust according to the line-of-sight vector direction is described based on the Earth's inertial coordinate system. The line-of-sight direction e obtained from equation (2) is transformed into the line-of-sight direction e based on the satellite coordinate system. v The engine is in accordance with e v The direction provides the satellite with maximum thrust T;

[0017] Step (4): Satellite motion calculation

[0018] During the thrust duration t cha Within the orbit, the satellite maintains motion at maximum thrust T, describing the satellite's motion at time t. cha The differential equation for motion within a time interval is:

[0019]

[0020] In the formula, r and v are the position vector and velocity vector of the satellite in the Earth's inertial coordinate system, respectively; v1 is the initial velocity of the satellite at that point; and μ is the Earth's gravitational constant, taken as μ = 398600 km. 3 / s 2 T is the thrust magnitude, m is the satellite mass, and e is the line-of-sight vector; after t cha After a period of time, the satellite reaches a new position, and the target also reaches a new position, updating the satellite's position vector R. cha and target position vector R tar Information; return to step (1) to perform a loop calculation.

[0021] Furthermore, step (3) specifically involves:

[0022] The parameters describing the satellite's attitude are yaw angle. Pitch angle θ, roll angle γ, and coordinate transformation matrices from Earth inertial coordinate system to satellite coordinate system. Represented as:

[0023]

[0024] Then, the conversion relationship between the line-of-sight direction e v based on the earth inertial coordinate system is expressed as:

[0025]

[0026] The satellite provides the maximum thrust T v in the direction of e v , that is, the maximum thrust T v based on the satellite coordinate system is expressed as:

[0027] T v = T v e v Equation (7)

[0028] Then, the maximum thrust T based on the earth inertial coordinate system is expressed as:

[0029]

[0030] Compared with the prior art, the present application has the following advantages:

[0031] The present application determines the thrust vector direction by the line-of-sight vector direction, continuously applies the thrust for a certain period of time until the next change of the thrust vector, and the engine is continuously turned on during the whole process, which can reduce the number of engine start-stop and maximize the utilization efficiency of the thrust; in theory, the satellite can be maneuvered at any point on the orbit, without moving to a special point (usually the perigee or apogee) for orbit change for a long time, which can effectively reduce the time required for approaching the target. BRIEF DESCRIPTION OF DRAWINGS

[0032] Figure 1 The flow chart of the orbit maneuver guidance calculation method of the high-orbit satellite for close-in approach to a non-cooperative target according to the present application.

[0033] Figure 2 The schematic diagram of thrust application of a high-orbit satellite.

[0034] Figure 3 The relative distance change diagram of a high-orbit satellite for close-in approach to a non-cooperative target.

[0035] 1-high-orbit satellite, 2-non-cooperative target. DETAILED DESCRIPTION

[0036] The present application will be further described in detail below with reference to the accompanying drawings.

[0037] The application provides a maneuvering guidance method for a high-orbit satellite to approach a target closely, comprising the following steps:

[0038] According to specific working conditions and approaching requirements at the beginning of a task, a task sequence is generated, including target position information R tar , thrust vector T, and thrust duration t cha .

[0039] Step (1), line-of-sight vector calculation

[0040] In the earth inertial coordinate system, according to the detected target position coordinates R tar and the position coordinates R cha of the satellite, the relative position vector R between the target and the satellite is solved, and the unit vector of the relative position vector R is the line-of-sight vector e, so that the target information sequence related to the position vector and the line-of-sight direction is obtained, and the above variables can be expressed by the following formula:

[0041] R = R tar - R cha Formula (1)

[0042]

[0043] Step (2), relative distance calculation

[0044] The relative distance △x between the satellite and the non-cooperative target at this time is calculated, and the calculation method is as follows:

[0045] △x = ||R|| Formula (3)

[0046] Given that the approaching task index is δ (i.e., the relative distance between the satellite and the non-cooperative target at the end of the task is less than δ), if △x does not satisfy formula (4), step (3) is entered; if △x satisfies formula (4), it indicates that the approaching task is completed, and the task is completed.

[0047] △x < δ Formula (4)

[0048] Step (3), maximum thrust calculation

[0049] The maximum thrust is applied according to the line-of-sight vector direction, and no matter the configuration of the thrust system on the satellite, the basic principle is to provide different thrusts by multiple engines to cooperate with each other, so that the total thrust direction is the same as the line-of-sight direction. It is worth noting that the line-of-sight direction e obtained by formula (2) is described based on the earth inertial coordinate system, and the thrust direction of the satellite is also related to the attitude of the satellite, so the line-of-sight direction e based on the earth inertial coordinate system cannot be directly used as a reference for the thrust direction of the satellite, and it still needs to be converted into the line-of-sight direction e v based on the satellite coordinate system, and the engine provides the maximum thrust for the satellite according to the direction of e v .

[0050] wherein the parameters describing the satellite attitude are yaw angle pitch angle θ and roll angle γ, then the coordinate transformation matrix from the earth inertial coordinate system to the satellite coordinate system may be expressed as:

[0051]

[0052] Then, the conversion relationship between the line-of-sight direction e v based on the satellite coordinate system and the line-of-sight direction e based on the earth inertial coordinate system can be expressed as:

[0053]

[0054] The satellite provides the maximum thrust in the direction of e v , which is denoted as T v , i.e. the maximum thrust T v based on the satellite coordinate system can be expressed as:

[0055] T v = T v e v Equation (7)

[0056] Then the maximum thrust T based on the earth inertial coordinate system can be expressed as:

[0057]

[0058] Step (four), satellite motion calculation

[0059] In the duration of the thrust t cha , the satellite maintains the maximum thrust T motion, and the differential equation describing the satellite motion in t cha time is:

[0060]

[0061] In the equation, r and v are the position vector and velocity vector of the satellite in the earth inertial coordinate system, v1 is the initial velocity of the satellite at the point, μ is the earth gravity constant, μ = 398600 km 3 / s 2 , T is the thrust size, m is the satellite mass, and e is the line-of-sight vector. After t cha time of motion, the satellite reaches a new position, and the target also reaches a new position, and the satellite position vector R cha and the target position vector R tar information are updated. Return to step (one) for loop calculation.

[0062] Embodiment

[0063] AsFigure 1 As shown: After the satellite reaches a position where it can detect non-cooperative targets, maneuvering guidance begins. The onboard GNC detection system detects the position coordinates of the non-cooperative target and calculates the current relative distance Δx between itself and the target using its own position coordinates. Assuming the minimum approach distance index for this approach mission is δ, the satellite compares Δx with δ. If the condition Δx < δ is not met, it means the current position does not meet the approach mission index, and the satellite enters cyclic maneuvering guidance. Thrust is applied along the line of sight and maintained for a certain period of time. The magnitude of the thrust is determined by the performance of the satellite's thrust system, and the duration of the thrust is determined by the computing power of the onboard computing unit or the ground control station. The shorter the duration, the more control operations are performed, the higher the control accuracy, and the higher the computing power requirements. After the thrust is maintained for a certain period of time, the satellite reaches a new position, detects the target's position coordinates again, and calculates the relative distance. This cyclic process is repeated until the condition Δx < δ is met, making the distance between the satellite and the non-cooperative target less than the approach distance mission index, thus laying the foundation for subsequent rendezvous, capture, or interception maneuvers.

[0064] like Figure 2 As shown: Relative position vectors of the satellite and non-cooperative targets and thrust As collinear vectors, the thrust remains constant in each cycle of maneuver guidance until a change is made in the next cycle of maneuver guidance, and the engine does not shut down throughout the entire process.

[0065] like Figure 3 As shown: Taking a satellite in GEO orbit as an example, a close approach mission with an accuracy of 10m was achieved against a non-cooperative target with irregular speed changes at a relative distance of 36km. The satellite has a thrust error and a detection error of 10%. According to the maneuvering method proposed in this invention, the close approach mission against the non-cooperative target was achieved after 4940s, and the final approach distance was 4.6m, which met the mission requirements.

Claims

1. A method of close-proximity maneuvering guidance for a high orbit satellite, characterized by, According to the working condition at the beginning of the task and the approaching requirement index, a task sequence is generated, and the task sequence includes target position information R tar , thrust vector T, thrust duration t cha , and specifically includes the following steps: Step (1): Line-of-sight vector calculation In the inertial coordinate system of the earth, according to the detected position coordinates R tar of the target cha , the relative position vector R between the target and the self is solved, and the unit vector e is the line-of-sight vector, so that the target information sequence related to the position vector and the line-of-sight direction is obtained, and the above variable is expressed by the following formula: R = R tar - R cha Formula (1) Step (2): Relative distance calculation The relative distance between the satellite and the target at the current time is calculated as follows: △x = ||R|| Equation (3) Given the approaching task index δ, i.e., the relative distance between the satellite and the non-cooperative target at the end of the task is less than δ, if △x does not satisfy Equation (4), proceed to Step (3); if △x satisfies Equation (4), it indicates that the approaching task is completed, and the task is completed. △x < δ Equation (4) Step (3): Maximum thrust calculation Applying the maximum thrust in the direction of the line-of-sight vector is described in the Earth inertial coordinate system, transforming the line-of-sight direction e found from equation (2) into the line-of-sight direction e in the satellite coordinate system v , the engine providing the maximum thrust T to the satellite in the direction of e v . Step (4): Satellite motion calculation During the thrust duration t cha , the satellite moves with maximum thrust T, describing a trajectory in the time t cha , the differential equation of the satellite motion is wherein r, v are the position vector and velocity vector of the satellite in the earth inertial coordinate system, v1 is the initial velocity of the satellite at the point, μ is the earth gravity constant, μ = 398600 km 3 / s 2 , T is the thrust size, m is the satellite mass, e is the line-of-sight vector; after t cha time of movement, the satellite reaches a new position, the target also reaches a new position, the satellite position vector R cha and the target position vector R tar information are updated; return to step (1) for loop calculation; Step (3) is specifically: The parameters describing the satellite attitude are yaw angle pitch angle θ, roll angle γ, coordinate transformation matrix from the earth inertial coordinate system to the satellite coordinate system is expressed as: Then, the conversion relationship between the line-of-sight direction e based on the satellite coordinate system and the line-of-sight direction e based on the earth inertial coordinate system is expressed as: v Then, the conversion relationship between the line-of-sight direction e based on the satellite coordinate system and the line-of-sight direction e based on the earth inertial coordinate system is expressed as: Satellite according to e v The maximum thrust provided by the direction is denoted as T. v That is, the maximum thrust T based on the satellite coordinate system v Represented as: T v = T v e v Equation (7) Then the maximum thrust T based on the Earth inertial coordinate system is represented as:

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