A momentum analysis method for identifying instantaneous force impact on satellite space

Abstract

The present invention provides a momentum analysis method for identifying instantaneous force impact on satellite space, determines the attitude control thrust curve under current conditions, and determines the thrust size of the attitude control thruster under different jet duration conditions; sequentially calculates the impulse of each thruster in each control beat, the conversion matrix from the satellite system to the satellite orbit coordinate system, and the satellite semi-major axis change; uses the precise orbit determination data before and after the anomaly occurs to calculate the actual semi-major axis change before and after the anomaly. If the difference is greater than the set threshold, it is determined that it is caused by the impact of external force in space. The present invention can provide direct data support for satellite attitude anomaly positioning, influence domain analysis, source tracing analysis, and assessment of the mass and size of the impactor.

Description

Technical Field

[0001] The invention belongs to the field of aerospace measurement and control, and relates to a momentum analysis method of satellite forces. Background Art

[0002] During the operation of satellites in orbit, they face the threat of complex space environment. Space debris and micrometeoroids are running at high speed in space and have extremely high kinetic energy. If they collide with spacecraft, they will cause serious harm to the latter. my country has the ability to monitor larger space targets and avoid satellites, but due to technical limitations, tiny space debris and micrometeoroids still pose a great threat to my country's spacecraft in orbit. In the case of a small probability collision with tiny debris that cannot be monitored, the satellite is affected by the instantaneous impact and causes consequences. First, it causes damage to onboard components or fatal damage to the entire satellite, or even the entire satellite disintegrates; second, the satellite attitude may become unstable, causing abnormal jet control and causing changes in the satellite orbit; third, the instantaneous force impact will bring a speed increment to the entire satellite, which will also cause orbit changes.

[0003] For such anomalies, how to analyze whether the instantaneous force impact on the satellite is caused by internal factors or external impact is very important for the accurate positioning of the anomaly, standardized handling, and the country's correct judgment and external voice on such space events. At present, the force situation of the satellite is usually analyzed through changes in satellite telemetry data and post-orbit determination. It is difficult to accurately trace the cause through the telemetry data analysis method before and after the anomaly occurs, and it is impossible to directly prove whether the satellite anomaly is caused by external force impact or internal force; the post-orbit determination method can only evaluate the influence of known force models, and cannot evaluate the magnitude, force time and force direction of unknown forces. Summary of the invention

[0004] In order to overcome the shortcomings of the prior art, the present invention provides a momentum analysis method for identifying instantaneous force impacts on satellites in space. Based on the attitude and orbit control design characteristics of the satellite, the jet pulse width, jet duration, jet thrust, jet direction, and three-axis attitude of the orbit system are determined by analysis. The velocity increment generated by the jets of multiple attitude control engines after the abnormality occurs is calculated to cause the change of the satellite's semi-major axis. The method is compared with the precise orbit determination result to identify whether the abnormal jet causing the attitude control is caused by an external force impact. This method can provide direct data support for satellite attitude abnormality positioning, influence domain analysis, source tracing analysis, and assessment of the mass and size of the impactor.

[0005] The technical solution adopted by the present invention to solve the technical problem comprises the following steps:

[0006] (1) Determine the attitude control thrust curve under current conditions and determine the thrust of the attitude control thruster under different jet durations;

[0007] (2) Let the satellite thruster jet pulse width sequence be {tk}, the thrust sequence of the corresponding pulse width is {F tk}, the installation matrix of each thruster is M bt , calculate the impulse I of each thruster in each control cycle k =F tk t k ;

[0008] If the thrusters are not installed in the forward direction, the vector representation of the impulse vector of each thruster in each control beat in the satellite body coordinate system is calculated according to the installation direction of each thruster in the satellite body coordinate system. bk =M bt I k ;

[0009] (3) Calculate the transformation matrix M from the satellite local system to the satellite orbit coordinate system ob =M oi ·(M bi ) T , where M bi represents the transformation matrix from the inertial system to the local system, M oi Represents the transformation matrix from the 2000.0 inertial coordinate system to the orbital coordinate system;

[0010] (4) Based on the impulse generated by the jth thruster and the impulse of the satellite in the direction of motion, calculate the change in the satellite's semi-major axis Δa g ;

[0011] (5) Using the precise orbit determination data before and after the anomaly, calculate the actual semi-major axis change Δa before and after the anomaly to obtain the difference Δa s =Δa-Δa g If the difference is greater than the set threshold, it is determined to be caused by the impact of external force in space.

[0012] The step (3) comprises the following steps: taking the attitude measurement data of the satellite before the abnormality as the initial value, and setting the attitude in the J2000 inertial system to quaternion q GU0 , the gyro angular velocity vector ω bk After expanding to a 4-dimensional vector, the inertial system attitude quaternion q is calculated by integration GUk =[q0q1q2q3]; According to the inertial attitude quaternion, the transformation matrix M from the inertial system to the current system is calculated bi ; According to the definition of satellite orbital coordinate system, the transformation matrix M of space vector from 2000.0 inertial coordinate system to orbital coordinate system is oi ; Calculate the transformation matrix M from the satellite system to the satellite orbit coordinate system ob =M oi ·(M bi ) T .

[0013] The inertial system to the system conversion matrix

[0014]

[0015] The conversion matrix M from the 2000.0 inertial coordinate system to the orbital coordinate system oi for:

[0016]

[0017]

[0018] M oi (1,m)=M oi (2,p)×M oi (3,k)

[0019] Where i = 1, 2, 3 corresponds to the transformation matrix M oi The three components of each row vector in M oi (2,p) and M oi (3,k) is the matrix M oi The row vectors of the second and third rows in , is represented as a unit vector of the space vector r, Velocity vector of space vector r.

[0020] In step (4), the impulse generated by the j-th thruster in the orbital system is N represents the total number of N jet pulse widths generated by the j-th thruster during the jet time, I bji It represents the impulse generated by the ith pulse width of the jth thruster in this system.

[0021] In step (4), assuming that a satellite has M attitude control thrusters, the thrust in each direction is the sum of the components generated by the M attitude control thrusters in each direction. The impulse in the forward direction T in the orbital system is

[0022] In step (4), let the satellite mass be m, the satellite orbit velocity be n, the orbit eccentricity be e, and the true anomaly be f. Then the semi-major axis change caused by the satellite attitude control jet is I ox Represents the impulse of the satellite in the x direction.

[0023] In the step (5), the difference Δa s =Δa-Δa g If it is greater than 3σ of the calculation error, it is determined to be caused by the impact of external force in space.

[0024] The beneficial effects of the present invention are: a complete description of the use of abnormal jet impulse analysis of satellite attitude control thrusters to estimate the change in the semi-major axis of the satellite orbit, and the method of determining the data difference with the precise orbit to realize the analysis and judgment of the impact of external forces on the satellite. The present invention introduces the process and implementation method, and the present invention can be continuously improved and optimized in the future by increasing or reducing the calculation process links, changing the process method, optimizing and improving the calculation accuracy, etc. according to the changes in the situation. The first is to calibrate the thrust link of the attitude control thruster. When the thruster does not need to be calibrated or the thruster is equipped with a thrust measurement device, this process can be omitted or directly calculated using known values; the second is to determine the satellite attitude matrix link. This process can also establish different attitude determination methods by changing the calculation attitude conversion order, changing the calculation attitude intermediate matrix, and the attitude description method; the third is to improve the attitude estimation accuracy. The attitude estimation accuracy is improved by other attitude measurement methods, error correction methods, and optimization calculation methods. For example, the attitude accuracy is improved by correcting the gyro integral. In addition, in analyzing the possibility of spacecraft being subjected to external forces, in addition to estimating the magnitude of the forces acting on the spacecraft, the judgment criteria can be given according to the magnitude of the change in the semi-major axis in the force results, and the magnitude of the forces and the magnitude of the orbital changes can be combined to comprehensively and accurately judge the external forces on the spacecraft, so as to improve the accuracy of the application of the method.

[0025] The present invention breaks through the traditional method of estimating force results according to the known force conditions of the spacecraft, adopts a reverse calculation method, and designs an algorithm based on the instantaneous force impact momentum analysis model. Starting from the results and processes, the reverse tracing analysis provides a new idea for spacecraft abnormality analysis and judgment, provides a research and judgment plan for the result determination of in-orbit satellites after being impacted by external forces, provides a data source for analyzing the mass and size of the impactor, and improves the method of perceiving the on-orbit state of spacecraft. It not only has a major innovation in the analysis of external forces in the aerospace field, but the calculation results of the present invention also provide strong evidence for the positioning and characterization of such events. BRIEF DESCRIPTION OF THE DRAWINGS

[0026] Figure 1 It is the flow chart of the algorithm of the present invention.

[0027] Figure 2 It is a schematic diagram of the average pulse width of the attitude control jet according to the example of the present invention.

[0028] Figure 3 It is a schematic diagram of impulse generated by each jet beat under this system in the example of the present invention.

[0029] Figure 4 It is a schematic diagram of the attitude quaternion in the inertial system of the example of the present invention.

[0030] Figure 5 It is a schematic diagram of impulses generated at each beat in the example track system of the present invention. DETAILED DESCRIPTION

[0031] The present invention is further described below in conjunction with the accompanying drawings and embodiments. The present invention includes but is not limited to the following embodiments.

[0032] like Figure 1 As shown, the present invention calculates the impulse generated by the jet on the basis of calibrating the thrust of the attitude control thruster, combines the satellite jet impulse under this system with the orbital system attitude, calculates the impulse generated by the orbital system and the orbital change, and estimates the possibility of external force impact by the difference with the subsequent precise orbit determination result. The specific calculation steps are as follows:

[0033] 1. Calibrate the thrust of the attitude control thruster

[0034] The size of the satellite attitude control thruster is determined by the inlet pressure and the injection duration. Therefore, the attitude control thrust curve under the current conditions should be determined first, and the thrust size of the attitude control thruster under different injection durations should be determined.

[0035] 2. Calculate the abnormal jet impulse of attitude control

[0036] Let the satellite thruster jet pulse width sequence be {t k}, the thrust sequence of the corresponding pulse width is {F tk}, the installation matrix of each thruster is M bt .

[0037] Calculate the impulse I of each thruster in each control cycle k =F tk t k .

[0038] If the thrusters are not installed in the forward direction, the vector representation of the impulse vector of each thruster in each control beat in the satellite body coordinate system is calculated according to the installation direction of each thruster in the satellite body coordinate system. bk =M bt I k .

[0039] 3. Determine the satellite attitude matrix

[0040] During the period of abnormal satellite attitude ejection, the attitude angle information output by the star sensor is invalid. In order to obtain the real attitude information, it is necessary to use the angular velocity information output by the gyroscope to integrate and calculate the satellite attitude.

[0041] Calculate the satellite’s instantaneous attitude quaternion.

[0042] The satellite attitude measurement data before the abnormality is taken as the initial value, and the attitude in the J2000 inertial system is expressed by quaternion q GU0 =[q0 q1 q2 q3], the gyro angular velocity vector is expressed as ω bk =[ω x ω y ωz ] T , expanded to a 4-dimensional vector ω bk =[ω x ω y ω z 0] T . Integrate to calculate the inertial system attitude quaternion q GUk =[q0 q1 q2 q3].

[0043] Calculate the transformation matrix from the inertial system to the local system.

[0044] According to the inertial attitude quaternion, the transformation matrix M from the inertial system to the local system is calculated. bi .

[0045]

[0046] Calculate the transformation matrix from the inertial system to the orbital system.

[0047] According to the definition of satellite orbital coordinate system, the transformation matrix M of space vector from 2000.0 inertial coordinate system to orbital coordinate system is oi for:

[0048]

[0049]

[0050] M oi (1,m)=M oi (2,p)×M oi (3,k)

[0051] Where: i = 1, 2, 3 corresponds to the transformation matrix M oi The three components of each row vector in M. oi (2,p) and M oi (3,k) is the matrix M oi The row vectors for the second and third rows in . is represented as a unit vector of the space vector r, Velocity vector of space vector r.

[0052] Calculate the transformation matrix from the satellite local coordinate system to the satellite orbit coordinate system.

[0053] M ob =M oi ·(M bi ) T .

[0054] 4. Calculate the change in attitude control jet trajectory

[0055] Calculate the impulse generated by the jth thruster

[0056] Taking the j-th thruster as an example and the jet impulse calculation as an example, it is assumed that the j-th thruster generates a total of N jet pulse widths within the jet time.

[0057]

[0058] Calculate the satellite's impulse in the direction of motion

[0059] Assuming that there are M attitude control thrusters, the impulse I of the satellite in the x direction is ox for:

[0060] Calculate the change in the satellite's semi-major axis

[0061] Assume that the satellite mass is m, the satellite orbit speed is n, the orbit eccentricity is e, and the true anomaly angle is f. Then the change of the semi-major axis caused by the satellite attitude control jet is Δa: g for:

[0062]

[0063] 5. Analyze the possibility of satellite being impacted by external forces

[0064] Using the precise orbit determination data before and after the anomaly, the actual semi-major axis change Δa before and after the anomaly can be calculated to obtain the difference Δa s =Δa-Δa g If the difference is greater than 3σ of the calculation error, it can be determined that it is caused by the impact of external force in space.

[0065] The following is a specific example. Based on the instantaneous force impact momentum analysis model algorithm, the force impact of a certain orbiting satellite (hereinafter referred to as Y satellite) is calculated. By comparing the calculation results and combining them with the attitude control jet impulse analysis, it can be determined that it is impacted by external forces, and the change in the semi-major axis of the orbit caused by the external force impact can be estimated, which provides an effective analysis method for analyzing the impact of external force impact on Y satellite.

[0066] Assuming that satellite Y has 6 thrusters installed on the -X plane, the attitude control jet will produce a +X velocity increment on the satellite, raising the semi-major axis.

[0067] By establishing an attitude control jet model, the impact of the jet on the semi-major axis of the satellite can be calculated. The following issues need to be considered in this process:

[0068] (1) Since the control subsystem has different jet pulse widths during attitude control, and the thrust of the thruster under different thruster inlet pressures and different jet pulse widths is also different, the magnitude of the attitude control thrust under the current conditions should be determined first;

[0069] (2) Due to the limitation of telemetry parameter design and sampling period, there is no direct parameter to characterize the thruster jet duration and jet pulse width. Therefore, it is necessary to consider how to determine the thruster jet duration and jet pulse width under the current conditions.

[0070] (3) The thrusters used for attitude control jets have different installation angles;

[0071] (4) During the satellite jet, the star sensor output is invalid and there is no real attitude information on the satellite.

[0072] 1. Determination of thruster thrust

[0073] (1) Step 1: Under certain pressure conditions, measure the thrust of multiple thrusters of the same model at different working pulse widths, and use the average value as the benchmark.

[0074] (2) Step 2: Considering the actual thruster inlet pressure of the current satellite, the corresponding relationship between the current thrust of the satellite and the jet pulse width is calculated by interpolation. The results are shown in Table 1:

[0075] Table 1 The change of thruster thrust with jet pulse width

[0076]

[0077] 2. Abnormal jet impulse calculation

[0078] (1) Step 1: Based on the analysis of satellite telemetry data, it is assumed that the duration of the abnormal attitude control jet is 119 seconds. The working conditions of the thrusters involved in the jet are shown in Table 2.

[0079] Table 2 Statistical results of thruster jet

[0080]

[0081] (2) Step 2: Calculate the average pulse width based on the jet duration increment and jet number in each frame of telemetry data. The result is as follows: Figure 2 shown.

[0082] (3) The third step: According to the thrust calibration results, the actual thrust size can be obtained, and the impulse of each thruster in each control cycle can be obtained. According to the thruster installation layout, the impulse of the impulse vector of each thruster in each control cycle in the body coordinate system can be calculated. The calculation results are as follows: Figure 3 shown.

[0083] 3. Attitude matrix determination based on gyro data

[0084] Based on the gyro measurement data, the satellite attitude quaternion is calculated by integration. The result is as follows Figure 4 As shown in Figure 2, the transformation matrix from the satellite system to the orbital system can be obtained based on the satellite attitude quaternion.

[0085] 4. Estimation of orbital changes caused by jets

[0086] According to the attitude transformation matrix calculated in the previous section, the jet impulse of each beat of the satellite in the orbital coordinate system is calculated, such as Figure 5 shown.

[0087] The component I of the impulse vector generated by each thruster in each control cycle in the X-axis direction of the orbital coordinate system is x By accumulating, the impulse components along the X-axis direction of the orbital coordinate system generated by all working thrusters during the entire control process can be obtained.

[0088] After calculation, we can get: I x =-508.2Ns.

[0089] Since the impulse is negative, that is, the impulse direction is along the -X axis, the resulting orbital altitude change should be an increase. According to the current satellite mass and current orbit, the change in the semi-major axis caused by the jet is calculated to be 1034m.

[0090] 5. Estimation of the possibility of external impact

[0091] After the attitude control jet was completed, according to the orbit determination results, the actual orbit was raised by 972 meters, which was 62 meters lower than the calculated theoretical semi-major axis. Without considering the thruster thrust calibration error, other error analysis mainly includes:

[0092] (1) There are errors in satellite attitude determination. Onboard telemetry does not have true attitude information. The ground calculates the satellite's three-axis attitude through gyro integration. Affected by the gyro constant drift and sampling period, the attitude error is about 1 degree, and the impact on the orbit can be ignored.

[0093] (2) In the telemetry data, the attitude sampling period is 0.5 seconds, while there are two sampling periods for the jet duration, 2 seconds and 10 seconds. After averaging the data of the 10-second sampling period, there will still be deviations in the attitude information used during the jet. The attitude error is within about 3 degrees, which has little impact on the orbit.

[0094] (3) The influence of jet pulse width on thrust size was taken into consideration, the attitude control jet pulse width was calculated, and the whole satellite impulse calculation was introduced. Without considering the thrust model error, the thrust error is very small and can be ignored.

[0095] (4) In the early stage of jetting, the temperature of the catalyst bed is relatively low, which has a certain impact on the jetting efficiency. In the calculation, the thrust efficiency in the first 10 seconds is estimated to be 80%, and the theoretical orbit elevation value is reduced by about 16 meters.

[0096] In summary, the actual orbit is more than 62 meters lower than the theoretically calculated orbit, which exceeds 3σ of the theoretical error. It can be judged that the satellite is likely to be impacted by instantaneous external forces in space.

Claims

1. A momentum analysis method for identifying instantaneous force impact on satellite space, characterized in that: The following steps are involved: (1) Determine the attitude control thrust curve under current conditions and determine the thrust of the attitude control thruster under different jet durations; (2) Let the satellite thruster jet pulse width sequence be {t k }, the thrust sequence of the corresponding pulse width is {F tk }, the installation matrix of each thruster is M bt , calculate the impulse I of each thruster in each control cycle k =F tk t k ; If the thrusters are not installed in the forward direction, the vector representation of the impulse vector of each thruster in each control beat in the satellite body coordinate system is calculated according to the installation direction of each thruster in the satellite body coordinate system. bk =M bt I k ; (3) Calculate the transformation matrix M from the satellite local system to the satellite orbit coordinate system ob =M oi ·(M bi ) T , where M bi represents the transformation matrix from the inertial system to the local system, M oi Represents the transformation matrix from the 2000.0 inertial coordinate system to the orbital coordinate system; (4) Calculate the change in the satellite semi-major axis △a based on the impulse generated by the jth thruster and the impulse of the satellite in the direction of motion g ; In step (4), the impulse generated by the j-th thruster in the orbital system is N represents the total number of N jet pulse widths generated by the j-th thruster during the jet time, I bji represents the impulse generated by the ith pulse width of the jth thruster in this system; Assuming that a satellite has M attitude control thrusters, the thrust in each direction is the sum of the components generated by the M attitude control thrusters in each direction. The impulse in the forward direction T in the orbital system is Assuming the satellite mass is m, the satellite orbit speed is n, the orbit eccentricity is e, and the true anomaly angle is f, the change in the semi-major axis caused by the satellite attitude control jet is I ox represents the impulse of the satellite in the x direction; (5) Using the precise orbit determination data before and after the anomaly, calculate the actual semi-major axis change △a before and after the anomaly, and obtain the difference △a s =△a-△a g If the difference is greater than the set threshold, it is determined to be caused by the impact of external force in space.

2. The momentum analysis method for identifying instantaneous force impact on a satellite in space according to claim 1 is characterized in that: The step (3) comprises the following steps: taking the attitude measurement data of the satellite before the abnormality as the initial value, and setting the attitude in the J2000 inertial system to quaternion q GU0 , the gyro angular velocity vector ω bk After expanding to a 4-dimensional vector, the inertial system attitude quaternion q is calculated by integration GUk =[q0 q1 q2 q3]; According to the inertial attitude quaternion, the transformation matrix M from the inertial system to the current system is calculated bi ; According to the definition of satellite orbital coordinate system, the transformation matrix M of space vector from 2000.0 inertial coordinate system to orbital coordinate system is oi ; Calculate the transformation matrix M from the satellite system to the satellite orbit coordinate system ob =M oi ·(M bi ) T .

3. The momentum analysis method for identifying instantaneous force impact on a satellite in space according to claim 1 is characterized in that: The inertial system to the system conversion matrix 4. The momentum analysis method for identifying instantaneous force impact on a satellite in space according to claim 1 is characterized in that: The conversion matrix M from the 2000.0 inertial coordinate system to the orbital coordinate system oi for: M oi (1,m)=M oi (2,p)×M oi (3,k) Where i = 1, 2, 3 corresponds to the transformation matrix M oi The three components of each row vector in M oi (2,p) and M oi (3,k) is the matrix M oi The row vectors of the second and third rows in , is represented as a unit vector of the space vector r, Velocity vector of space vector r.

5. The momentum analysis method for identifying instantaneous force impact on satellite space according to claim 1 is characterized in that: In the step (5), the difference △a s =△a-△a g If it is greater than 3σ of the calculation error, it is determined to be caused by the impact of external force in space.

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