High-precision linear motion direct current driving system for spaceflight and control method of high-precision linear motion direct current driving system

Abstract

The invention discloses a high-precision linear motion direct-current driving system for spaceflight and a control method of the high-precision linear motion direct-current driving system, and relates to the technical field of spacecraft attitude driving control. The high-precision linear motion direct current driving control method for spaceflight comprises the steps of linear direct current motor response aging judgment, spacecraft driving load evaluation and spacecraft driving signal optimization. According to the method, the driving load influence evaluation requirement is determined through command signal response timeliness judgment, if the driving load influence evaluation requirement does not exist, the spacecraft attitude adjustment process continues to be monitored, otherwise, the driving load optimization requirement continues to be determined, and finally if the driving load optimization requirement exists, the oscillation signal influence quantification requirement continues to be determined. Otherwise, the oscillation signal optimization requirement is directly determined after the oscillation signal influences quantification, the spacecraft linear motion direct current driving stability is improved, and the problem that in the prior art, spacecraft linear motion direct current driving is not stable due to the spacecraft attitude adjustment load requirement difference is solved.

Description

Technical Field

[0001] The present invention relates to the technical field of spacecraft attitude drive control, and in particular to a high-precision linear motion DC drive system for aerospace and a control method thereof. Background Art

[0002] Spacecraft carry a variety of key equipment to accomplish their missions, encompassing scientific exploration, communications, navigation, Earth monitoring, weather forecasting, military reconnaissance, and other areas. A spacecraft's design and the types of equipment it carries directly determine its mission capability and efficiency. Typical spacecraft equipment includes attitude control systems, propulsion systems, energy systems, communications systems, remote sensing instruments, payload equipment, life support systems, and data storage and processing systems. The attitude control system includes reaction wheels (which control the spacecraft's attitude by varying its linear velocity), control thrust gyros (which precisely control the spacecraft's attitude using angular momentum), magnetometers, gyroscopes, solar sails, and control surfaces. Each device is driven by an independent linear DC motor. The magnetometers and gyroscopes provide information about the spacecraft's attitude; the solar sails and control surfaces control the spacecraft's direction.

[0003] For example, a reaction wheel generates angular momentum (the rotational momentum of an object about a specific axis) through rotation. When the linear velocity of the reaction wheel changes, the spacecraft's attitude also changes according to the law of conservation of angular momentum. For example, if the linear velocity of the reaction wheel increases or decreases, the spacecraft will rotate accordingly, changing its attitude. A linear DC motor can serve as the drive motor for the reaction wheel, which is connected to the DC motor's rotor via a shaft. The DC motor controls the wheel's rotational speed, and the reaction force acts on the spacecraft through the shaft, thereby changing the spacecraft's angular momentum and adjusting its attitude.

[0004] The attitude control system in the space shuttle is driven by a linear DC motor. This motor primarily generates linear thrust through the interaction of magnetic fields and currents, achieving linear motion. The primary output of a linear DC motor is thrust applied along the motor's direction of motion (usually a linear track). This thrust is generated by electromagnetic forces acting between the stator and mover, acting on a load along the linear direction. Closed-loop control during spacecraft attitude adjustment involves real-time monitoring of the spacecraft's attitude, angular velocity, acceleration, and other data to provide precise feedback to the attitude control system, enabling high-precision attitude adjustment. Real-time monitoring of sensor output signals allows for effective adjustment of the linear DC motor control process. Because signal transmission delays or phase shifts between sensors can affect the precise control of the overall system, algorithms (such as Kalman filtering and cross-correlation analysis) are often used in multi-sensor systems to correct for signal synchronization and phase shifts. During spacecraft attitude adjustment, the load on the linear DC motor continuously fluctuates, affecting its output thrust. To ensure precise attitude control, real-time monitoring of load fluctuations and timely adjustment of the linear DC motor's output thrust are essential. In high-precision linear motion systems, load fluctuations and motor thrust changes have a significant impact on posture control. The motor's output thrust is adjusted by monitoring load changes. Load change monitoring is usually achieved by thrust sensors, which can detect tiny deformations (strains) of the shaft caused by thrust.

[0005] During spacecraft attitude adjustments, oscillation signals are often generated by feedback from motors, sensors, or other control systems. The presence of oscillation signals can affect attitude control accuracy, especially in space missions requiring high-precision control. Linear DC motor control technology primarily uses a proportional-integral-derivative control (PID) controller to regulate motor speed and thrust for precise motion control. The PID control algorithm precisely tracks the target state by adjusting the proportional, integral, and derivative gains. Thermal management techniques typically rely on simple temperature sensor monitoring systems, activating cooling systems such as heat sinks or fans when the temperature exceeds a preset threshold.

[0006] For example, a Chinese patent application with publication number CN112947514B discloses a spacecraft attitude control method based on an event-driven superhelix algorithm, which includes: first, using a multivariable generalized superhelix algorithm to construct a control input for a spacecraft attitude control system, and then constructing an event-driven trigger condition, in which additional terms are introduced to obtain stronger robustness and faster convergence speed, and obtain a trigger moment. The spacecraft actuator samples the state signal at the trigger moment to complete the control of the spacecraft attitude.

[0007] For example, a Chinese patent application with publication number CN108803647B discloses a model-free data-driven control method for spacecraft attitude control, which includes: Step 1: applying typical excitations to a space assembly spacecraft multiple times through an actuator carried by a service spacecraft; Step 2: collecting response information of the assembly spacecraft through a sensor carried by the service spacecraft; Step 3: establishing a mapping relationship between the excitation and response of the assembly spacecraft; Step 4: designing an assembly spacecraft controller by dynamically linearizing discrete data; Step 5: fitting the assembly spacecraft controller designed in Step 4 according to the classical dynamic model of the spacecraft.

[0008] The above technology has at least the following technical problems: In existing technologies, during spacecraft operation, the load requirements of the drive system may change rapidly due to factors such as the potential for changes in load when performing different attitude adjustment tasks. This can lead to system response delays and accuracy, affecting the accuracy and real-time performance of attitude adjustment, increasing position deviations, and thus affecting the stability of the spacecraft's attitude adjustment. It can also cause heat accumulation in the motor and drive, further affecting the stability of the drive system.

[0009] Furthermore, load variations can trigger heat accumulation in the drive system, affecting the operational stability of the linear DC motor and driver. If the drive system fails to adjust and compensate for load variations in a timely manner, the linear DC motor's output can become unstable, impacting the accuracy and stability of attitude adjustments. Large load fluctuations can cause a lag in the system's dynamic response, and the motor may be unable to quickly track control commands, leading to control distortion or reduced accuracy. This can lead to instability in the spacecraft's linear motion DC drive due to variations in the load requirements for attitude adjustments. Summary of the Invention

[0010] To address the technical problem of unstable DC drives for linear motion of spacecraft due to differences in load requirements for attitude adjustment in existing technologies, the present invention provides a high-precision DC linear motion drive system for aerospace and a control method thereof. The technical solution is as follows: On the one hand, a high-precision linear motion DC drive control method for aerospace is provided, which includes: S1, before the spacecraft attitude is adjusted, the command signal response timeliness of the linear DC motor output thrust is judged, and the drive load impact assessment requirement is determined based on the command signal response timeliness judgment; S2, if there is no drive load impact assessment requirement, continue to monitor the spacecraft attitude adjustment process, otherwise, perform drive load impact assessment based on the parameters in the spacecraft attitude adjustment process, and determine the drive load optimization requirement, if there is a drive load optimization requirement, determine the oscillation signal impact quantification requirement after the drive load optimization, otherwise directly perform oscillation signal impact quantification based on the parameters in the drive control process; S3, if there is an oscillation signal optimization requirement, determine whether to provide feedback to the preset personnel based on the result of the command signal response timeliness judgment after the oscillation signal optimization, otherwise continue to monitor the spacecraft attitude adjustment process.

[0011] On the other hand, a high-precision linear motion DC drive system for aerospace use is provided, which includes: an attitude adjustment requirement setting module, an attitude adjustment drive activation module and an attitude feedback and adjustment module; wherein, the attitude adjustment requirement setting module is used to detect the real-time attitude of the spacecraft through an attitude sensor when the spacecraft performs attitude adjustment, and generate an attitude adjustment signal based on the current spacecraft attitude; the attitude adjustment drive activation module is used to adjust the spacecraft attitude through a linear DC motor according to the attitude adjustment signal; the attitude feedback and adjustment module is used to determine whether to perform interference suppression based on the adjusted spacecraft attitude, and if interference suppression is performed, readjust the spacecraft attitude after interference suppression, otherwise continue to monitor the spacecraft attitude adjustment process, and interference suppression includes drive load optimization and oscillation signal optimization.

[0012] The beneficial effects brought about by the technical solution provided by the embodiment of the present invention include at least: 1. By determining the timeliness of the command signal response of the linear DC motor's output thrust and determining the drive load impact assessment requirements based on the command signal response timeliness determination, it is beneficial to timely discover possible response problems of the linear DC motor and provide a basis for subsequent spacecraft attitude adjustment. Then, based on the parameters in the spacecraft attitude adjustment process, the drive load impact is assessed and the drive load optimization requirements are determined. This can avoid damage to the linear DC motor caused by unreasonable load. Finally, based on the parameters in the drive control process, the oscillation signal impact is quantified and the oscillation signal optimization requirements are determined. This can prevent the spacecraft attitude adjustment from getting out of control due to oscillation, thereby improving the stability of the spacecraft attitude adjustment.

[0013] 2. By obtaining sensor data from various locations on the spacecraft, it is determined whether to perform oscillation signal optimization. If oscillation signal optimization is to be performed, data from the sensors installed at various locations on the spacecraft are obtained. If the fluctuation of the sensor output signal at each location is within the preset fluctuation range, oscillation signal optimization is not performed. Otherwise, oscillation signal optimization is performed based on the set lag time. The operating status of the DC drive system for spacecraft attitude adjustment is comprehensively evaluated, thereby reducing the impact of oscillation signals on spacecraft attitude adjustment and the operation of the linear DC motor.

[0014] 3. By correcting the output thrust of the linear DC motor, the linear DC motor is ensured to operate stably when the load increases, which helps to prevent overload of the linear DC motor. The load is then distributed in parallel to multiple motors to disperse the load pressure, improve the efficiency of spacecraft attitude adjustment, and avoid overloading of a single motor. Then, by adjusting the output power correction factor, the linear DC motor used for spacecraft attitude adjustment can be ensured to maintain high efficiency. Finally, the heat dissipation device is conditionally activated based on the temperature judgment result, realizing the rational utilization of heat dissipation resources, thereby ensuring that the linear DC motor operates within a safe temperature range and avoiding degradation of spacecraft attitude adjustment performance due to overheating. BRIEF DESCRIPTION OF THE DRAWINGS

[0015] In order to more clearly illustrate the technical solutions in the embodiments of the present invention, the following briefly introduces the drawings required for use in the description of the embodiments. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on these drawings without creative work.

[0016] Figure 1 This is a flow chart of spacecraft attitude adjustment drive optimization provided by an embodiment of the present invention; Figure 2 This is a flow chart of a high-precision linear motion DC drive control method for aerospace use provided by an embodiment of the present invention; Figure 3 is a flow chart of drive load optimization provided by an embodiment of the present invention; Figure 4 This is an interface diagram of a motor performance diagnosis interface provided by an embodiment of the present invention; Figure 5 This is an interface diagram of a motor signal diagnostic interface provided by an embodiment of the present invention; Figure 6 The diagram is a schematic structural diagram of a high-precision linear motion DC drive system for aerospace use provided by an embodiment of the present invention. DETAILED DESCRIPTION

[0017] In order to make the technical problems, technical solutions and advantages to be solved by the present invention clearer, a detailed description will be given below with reference to the accompanying drawings and specific embodiments.

[0018] like Figure 1 The flowchart of the spacecraft attitude adjustment drive optimization shown in the figure, wherein the connection points are explained as follows: 1 indicates quantifying the impact of the oscillation signal based on the parameters in the drive control process; 2 indicates continuing to monitor the spacecraft attitude adjustment process; the specific logic is: if the output thrust adjustment time is longer than the preset output thrust adjustment time, the drive load impact is evaluated based on the parameters in the spacecraft attitude adjustment process, otherwise the spacecraft attitude adjustment process is continued to be monitored; if the drive load impact evaluation result is greater than the preset load fluctuation threshold, there is a demand for drive load optimization, otherwise there is no demand for drive load optimization, and the subsequent oscillation signal impact quantification is performed directly; if there is no demand for drive load optimization, the oscillation signal impact is quantified directly based on the parameters in the drive control process; if there is a demand for drive load optimization, then after the drive load optimization, if the drive load impact evaluation result after the drive load optimization is greater than the preset load fluctuation threshold, then the drive load optimization is required, otherwise, there is no demand for drive load optimization, and the subsequent oscillation signal impact quantification is performed directly; if there is no demand for drive load optimization, the oscillation signal impact is quantified directly based on the parameters in the drive control process; if there is a demand for drive load optimization, then after the drive load optimization, if the drive load impact evaluation result after the drive load optimization is greater than the preset load fluctuation threshold, then the drive load optimization is required. If the load fluctuation threshold is greater than the preset value, the oscillation signal influence is quantified based on the parameters in the drive control process; otherwise, if the output thrust adjustment time after the drive load optimization is greater than the preset output thrust adjustment time, the oscillation signal influence is quantified based on the parameters in the drive control process; otherwise, the spacecraft attitude adjustment process is continued to be monitored; if the oscillation signal influence index is greater than the preset oscillation signal influence quantification threshold, there is an oscillation signal optimization requirement; otherwise, the spacecraft attitude adjustment process is continued to be monitored; if there is an oscillation signal optimization requirement, then if the output signal fluctuation of the sensor at each position is within the preset fluctuation range, the oscillation signal optimization is not performed; otherwise, the oscillation signal is optimized according to the set lag time; if the output thrust adjustment time after the oscillation signal optimization is greater than the preset output thrust adjustment time, feedback is given to the preset personnel; otherwise, the spacecraft attitude adjustment process is continued to be monitored.

[0019] As an embodiment of the first aspect, the present invention provides a high-precision linear motion DC drive control method for aerospace. Figure 2 The flowchart of a high-precision linear motion DC drive method for aerospace is shown, and the method includes the following steps: S1: Linear DC Motor Response Timeliness Assessment: Before adjusting the spacecraft's attitude, the linear DC motor's output thrust command signal response timeliness assessment is performed. Based on this timeliness assessment, the drive load impact assessment requirements are determined. Before adjusting the spacecraft's attitude, the linear DC motor response timeliness assessment module assesses the linear DC motor's command signal response timeliness to verify that the linear DC motor adjusts its output thrust according to the given command within the predetermined timeframe, thereby ensuring the stability and accuracy of attitude adjustments.

[0020] S2: Spacecraft drive load assessment: If there is no need for drive load impact assessment, continue to monitor the spacecraft attitude adjustment process; otherwise, perform drive load impact assessment based on the parameters in the spacecraft attitude adjustment process, and determine the drive load optimization requirements; if there is a drive load optimization requirement, determine the oscillation signal impact quantification requirement after the drive load optimization; otherwise, directly quantify the oscillation signal impact based on the parameters in the drive control process; if the drive load impact assessment result is not greater than the preset load fluctuation threshold, continue to monitor the spacecraft attitude adjustment process without optimization; otherwise, optimize the drive load to reduce the impact of load fluctuations on attitude adjustment.

[0021] S3: Spacecraft drive signal optimization: If there is a need for oscillation signal optimization, determine whether to provide feedback to the preset personnel based on the result of the command signal response timeliness judgment after the oscillation signal optimization; otherwise, continue to monitor the spacecraft attitude adjustment process; after evaluating and optimizing the linear DC motor drive load, further optimize the drive signal to ensure that no unnecessary oscillation signals or control errors are generated during the spacecraft attitude adjustment process to meet high-precision control requirements.

[0022] In this embodiment, mechanical components (such as antennas, solar panels, and instruments) within a spacecraft (e.g., a space shuttle) generate linear thrust during motion. Frictional thrust between these components can vary with the state of motion. Furthermore, the nonlinear characteristics of the linear DC motor and driver (such as motor torque fluctuation and driver output noise) can all contribute to load fluctuations. A sudden increase in load increases the power output of the linear DC motor, leading to increased heat. If the cooling system fails to dissipate heat in a timely manner, the linear DC motor may overheat. During orbital operation, temperature fluctuations can cause thermal expansion and contraction of mechanical components, impacting load characteristics. Load demands may vary as the spacecraft performs different attitude adjustment tasks. Furthermore, orbital adjustment tasks can cause changes in spacecraft acceleration, potentially leading to load fluctuations.

[0023] Furthermore, flexible attachments (such as large solar panels and antennas) on spacecraft (such as the space shuttle) are prone to vibration during attitude adjustments, resulting in oscillations in the output signal. When a spacecraft performs rapid attitude maneuvers, nonlinear switching control can easily excite sustained vibrations in the flexible attachments, compromising the smoothness and accuracy of the maneuvers. The flexible attachments are coupled to the central rigid body, generating flexible vibrations that can reduce the spacecraft's attitude pointing accuracy. Flexible vibration refers to the bending vibration of a structure under load and typically occurs in flexible or elastic objects or structures, particularly long and slender objects such as spacecraft attachments, antennas, and solar panels. Oscillations can lead to unstable output thrust from the drive system, compromising the accuracy and stability of attitude adjustments. Load fluctuations and the presence of oscillations can combine to affect control system stability.

[0024] By continuously monitoring and optimizing the linear DC motor's response time, load influence, and oscillation signal, it is possible to precisely control the spacecraft's attitude adjustment, avoid attitude errors and oscillations, and ensure that the linear DC motor is always in the best working state, thereby improving the stability of the spacecraft's (such as the space shuttle) attitude adjustment.

[0025] Furthermore, the specific method for determining the timeliness of the command signal response of the linear DC motor output thrust is as follows: The thrust output of the linear DC motor corresponding to the controlled device during spacecraft attitude adjustment is monitored in real time, and the duration of the output thrust adjustment is determined. This duration is determined by measuring the time from when the linear DC motor corresponding to the controlled device receives a command signal until the output thrust reaches a steady-state value. The thrust output of the linear DC motor is represented by a thrust sensor (such as a force sensor, which directly measures the motor's thrust output) installed at the motor's output terminal or on a load connected to the motor, and the sensor's output signal is read in real time. This step ensures that the thrust output of the linear DC motor meets design requirements during spacecraft attitude adjustment, ensuring the accuracy of spacecraft attitude adjustment. Obtaining the output thrust adjustment duration provides critical data for determining the timely response of the drive, helping to promptly identify potential problems with the drive response.

[0026] If the output thrust adjustment time is longer than the preset output thrust adjustment time obtained from the preset database, it means that there is a response delay, and the drive load impact assessment is performed based on the parameters during the spacecraft attitude adjustment process, otherwise the monitoring of the spacecraft attitude adjustment process continues; the preset output thrust adjustment time is set by the preset personnel; by comparing with the preset output thrust adjustment time, it is possible to quickly and accurately determine whether the linear DC motor has a response delay problem, providing a clear basis for subsequent processing decisions; when there is a response delay, performing a drive load impact assessment can deeply analyze the cause of the problem, provide targeted directions for solving the response delay problem, and help improve the reliability and stability of the linear motion DC drive system.

[0027] In this embodiment, through the above steps, the real-time determination and evaluation of the timeliness of the linear DC motor's output thrust response can promptly discover and resolve the linear DC motor's response delay problem, ensuring that the linear DC motor's output thrust can accurately and promptly respond to the command signal for spacecraft attitude adjustment, thereby enabling the spacecraft to be precisely adjusted according to a predetermined trajectory and attitude, thereby improving the accuracy and stability of the attitude adjustment of a spacecraft (such as a space shuttle), and meeting the requirements of space missions for high-precision attitude control.

[0028] Furthermore, the driving load impact assessment is performed based on the parameters during the spacecraft attitude adjustment process. The specific method is as follows: The attitude change comparison coefficient is obtained by comparing the average attitude change of the spacecraft during the attitude adjustment process with the preset attitude change obtained from the preset database, that is, ; Where F1 represents the average attitude change of the spacecraft during the attitude adjustment process, F 1,0 It represents a preset attitude change; the average attitude change represents the average value of the spacecraft attitude change within a preset time period, and the spacecraft attitude change is represented by the absolute value of the result of deviation processing of the attitude angle at the current preset time point and the attitude angle at the previous preset time point obtained by the inertial measurement unit; wherein, the preset time period and the preset time point are set by the preset personnel, and the preset attitude change, the preset linear DC motor heating rate, and the preset output thrust stability coefficient are set by the preset personnel; the comparison processing in this application represents the ratio operation between the actual value and the reference value, and the deviation processing represents the difference operation.

[0029] It's understandable that larger attitude changes may require higher thrust from the linear DC motor when performing attitude adjustments, resulting in a greater impact on the load. Larger attitude changes may also increase the workload of the linear DC motor, affecting its heating rate. By comparing the actual attitude change with the preset attitude change, we can quantitatively assess whether the spacecraft's attitude adjustments meet expectations.

[0030] According to the motor heating rate of the linear DC motor corresponding to the device to be controlled during the spacecraft attitude adjustment process and the preset motor heating rate obtained from the preset database, the heating rate comparison coefficient is obtained, that is, Where F2 represents the motor heating rate of the linear DC motor of the controlled device during the spacecraft attitude adjustment process, and F 2,0 Indicates the preset motor heating rate. The heating rate of the linear DC motor corresponding to the controlled device is expressed as the average value of the deviation between the temperature at the current preset time point and the temperature at the previous preset time point, which are collected by the temperature sensor. A larger posture change and a more unstable output thrust will lead to an increase in the load on the linear DC motor in a short period of time, thereby accelerating the heating rate. By comparing the actual linear DC motor heating rate with the preset heating rate, a quantitative assessment is made as to whether the motor temperature change is normal.

[0031] According to the output thrust stability coefficient of the linear DC motor corresponding to the device to be controlled during the spacecraft attitude adjustment process and the preset output thrust stability coefficient obtained from the preset database, an inverse comparison process is performed to obtain the output thrust stability inverse comparison coefficient, that is, Where F3 represents the output thrust stability coefficient of the linear DC motor during the spacecraft attitude adjustment process, F 30 Represents a preset output thrust stability coefficient; the output thrust stability coefficient is obtained by comparing the average value and the corresponding standard deviation of the output thrust of the linear DC motor corresponding to the controlled device collected within a preset time period; the inverse comparison processing in this application represents a ratio operation between the reference value and the actual value; by comparing the stability of the actual output thrust with the preset value, the stability of the motor output thrust is evaluated.

[0032] The drive load impact evaluation balance parameter is introduced, and the attitude change contrast coefficient, heating rate contrast coefficient and output thrust stability inverse contrast coefficient are weighted coupled to obtain the drive load impact evaluation result. The drive load impact evaluation result is used to quantitatively evaluate the influence of the load fluctuation of the linear DC motor on the spacecraft attitude adjustment process.

[0033] The specific limiting expression for the impact of the driving load on the evaluation results is:

[0034] Where F represents the evaluation result of the driving load impact during the spacecraft attitude adjustment process, α1 represents the balance parameter of the first driving load impact evaluation, α2 represents the balance parameter of the second driving load impact evaluation, and α3 represents the balance parameter of the third driving load impact evaluation.

[0035] It should be added that the drive load impact evaluation balance parameters are obtained from a preset database, specifically including a first drive load impact evaluation balance parameter, a second drive load impact evaluation balance parameter, and a third drive load impact evaluation balance parameter, which respectively represent the degree of influence of the parameters in the corresponding spacecraft attitude adjustment process on the oscillation signal influence index; the parameters in the spacecraft attitude adjustment process include the average attitude change of the spacecraft, the motor heating rate and the output thrust stability coefficient of the linear DC motor corresponding to the controlled device; the sum of the three is 1, for example, the parameters in the spacecraft attitude adjustment process are respectively compared with the corresponding preset drive load impact evaluation balance parameters to form a corresponding drive load impact evaluation mapping set, and the parameters in the real-time spacecraft attitude adjustment process are respectively input into the corresponding drive load impact evaluation mapping set to obtain the corresponding drive load impact evaluation balance parameters; the drive load impact evaluation mapping set includes a first drive load impact evaluation mapping set, a second drive load impact evaluation mapping set, and a third drive load impact evaluation mapping set.

[0036] In this embodiment, through multi-dimensional parameter comparison and comprehensive evaluation, the impact of the load fluctuation of the linear DC motor on the attitude adjustment of a spacecraft (such as a space shuttle) can be accurately quantified, which is conducive to timely detection of abnormal conditions during the spacecraft attitude adjustment process, provides a basis for subsequent drive load optimization, reduces the spacecraft attitude deviation caused by motor load fluctuations, and improves the reliability of the entire high-precision linear motion DC drive control method for aerospace.

[0037] like Figure 3 The flowchart of the drive load optimization shown has the following specific logic: based on the load of the linear DC motor corresponding to the device to be controlled, it is determined whether to correct the output thrust of the linear DC motor according to the thrust correction factor; then, based on the output thrust of the linear DC motor corresponding to the device to be controlled, it is determined whether to distribute the load of the linear DC motor in parallel; then, based on the output thrust of the linear DC motor corresponding to the device to be controlled, it is determined whether to correct the output power of the linear DC motor according to the output power correction factor; finally, based on the temperature of the linear DC motor corresponding to the device to be controlled, it is determined whether to activate the heat dissipation device according to the duration of the heat dissipation device being on.

[0038] like Figure 4 The motor performance diagnostics interface shown in the figure shows the selected attitude control device, displaying the corresponding motor and motor information. Selecting reaction wheel No. 1 displays the following information: device name, installation location, associated tasks, and information about each linear DC motor. Linear DC motor information includes linear speed, thrust, motor temperature, cumulative operating time, and abnormality duration. Expanding to view abnormal motor temperature information displays the following information: motor winding temperature, gearbox temperature, heat dissipation status, and abnormality information.

[0039] To further illustrate, the specific process of determining the drive load optimization requirements is as follows: If the drive load impact assessment result exceeds a preset load fluctuation threshold value obtained from a preset database, then drive load optimization is required and step A2 is executed. Otherwise, no drive load optimization is required and subsequent oscillation signal impact quantification proceeds directly. The preset load fluctuation threshold value, the preset linear DC motor load, the first preset linear DC motor output thrust, and the second preset linear DC motor output thrust are set by the preset personnel.

[0040] A2. If the load of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor load obtained from the preset database, the output thrust of the linear DC motor corresponding to the device to be controlled is corrected according to the thrust correction factor, and A3 is executed. Otherwise, A3 is directly executed. The thrust correction factor represents the result obtained by inputting the load deviation coefficient and the drive load impact evaluation result deviation coefficient into the thrust correction mapping set obtained from the preset database to adjust the output thrust of the linear DC motor corresponding to the device to be controlled; the correction in this application represents a multiplication operation, for example, multiplying the thrust correction factor and the linear DC motor thrust; pulse width modulation is used to control the current, thereby controlling the thrust output by the motor; by correcting the output thrust of the linear DC motor, it is ensured that the linear DC motor can operate stably when the load is larger, which helps to avoid overload and possible damage of the linear DC motor, thereby improving the accuracy and stability of spacecraft attitude control.

[0041] A3. If the output thrust of the linear DC motor corresponding to the device to be controlled is greater than the first preset linear DC motor output thrust obtained from the preset database, the load of the linear DC motor is distributed in parallel and A4 is executed, otherwise A4 is executed directly; the load can be distributed in parallel according to the equal distribution method, equal distribution: the total load is evenly distributed to multiple motors, for example, if there are two motors working in parallel and the total load is F, then each motor carries a load of F / 2; by distributing the load in parallel to multiple motors, the load pressure is dispersed, and the stability of the aerospace high-precision linear motion DC drive system is maintained; if the load is reasonably distributed, the carrying capacity of the aerospace high-precision linear motion DC drive system can be improved to avoid overload of a single motor.

[0042] A4. If the output thrust of the linear DC motor corresponding to the device to be controlled is greater than the second preset linear DC motor output thrust obtained from the preset database, the output power of the linear DC motor corresponding to the device to be controlled is corrected according to the output power correction factor. Otherwise, feedback is provided to the preset personnel. The output power is adjusted by adjusting the linear speed. By adjusting the output power correction factor, the linear DC motor can be ensured to maintain high efficiency under increasing load. The feedback can prompt the preset personnel to adjust the drive control strategy in a timely manner to avoid performance degradation or failure of the linear motion DC drive system for spacecraft attitude adjustment.

[0043] The load deviation coefficient is obtained by performing relative deviation processing on the load of the linear DC motor corresponding to the controlled device and the preset linear DC motor load, and describes the degree of difference between the actual linear DC motor load and the preset load. The drive load impact evaluation result deviation coefficient is obtained by performing relative deviation processing on the drive load impact evaluation result and the preset load fluctuation threshold, and measures the relative deviation between the load assessment and the preset load fluctuation threshold.

[0044] The output power correction factor represents the result obtained by inputting the output thrust deviation coefficient and the drive load impact evaluation result deviation coefficient into the output power correction mapping set obtained from the preset database to adjust the output power of the linear DC motor corresponding to the controlled device. The output thrust deviation coefficient is obtained by performing relative deviation processing on the output thrust of the linear DC motor corresponding to the controlled device and the output thrust of a second preset linear DC motor, reflecting the difference between the actual output thrust of the linear DC motor corresponding to the controlled device and the second preset output thrust.

[0045] It should be added that drive load optimization also includes: Determine whether the temperature of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor temperature obtained from the preset database: If the temperature of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor temperature obtained from the preset database, the heat dissipation device is activated according to the duration of the heat dissipation device's activation. Otherwise, the heat dissipation device (such as a fan, liquid cooling system, or heat pipe heat dissipation) is not activated. A temperature sensor installed on the linear DC motor collects the motor's actual temperature data in real time; the preset linear DC motor temperature is set by a preset personnel. By comparing the temperature of the linear DC motor corresponding to the device to be controlled with the preset linear DC motor temperature in real time, it is possible to promptly and accurately detect whether the linear DC motor is overheating, providing a basis for deciding whether to subsequently activate the heat dissipation device, thereby ensuring that the linear DC motor operates within a safe temperature range and avoiding degradation of the spacecraft's attitude adjustment performance due to overheating. Conditionally activating the heat dissipation device based on the temperature judgment result ensures the rational use of heat dissipation resources.

[0046] The cooling device is turned on for a period of time and the current linear DC motor temperature deviation coefficient is assigned a value based on the temperature reduction coefficient. The temperature reduction coefficient is obtained by comparing the cooling time in the historical time period with the linear DC motor temperature deviation coefficient. The linear DC motor temperature deviation coefficient is obtained by performing relative deviation processing on the temperature of the linear DC motor corresponding to the controlled device and the preset linear DC motor temperature. The assignment in this application represents a multiplication operation.

[0047] It's understandable that calculating the cooling device's on-time takes into account both the current temperature deviation of the linear DC motor corresponding to the controlled device and historical cooling data, making the determination of the cooling device's on-time more scientific and reasonable. The cooling coefficient reflects the patterns of historical cooling processes. Applying this to the current temperature deviation coefficient allows us to predict the time required for the motor temperature to drop to a safe range, thereby more precisely controlling the cooling device's on-time and improving the accuracy and effectiveness of temperature control during spacecraft attitude adjustments.

[0048] In this embodiment, through thrust correction, output thrust inspection and parallel distribution, the output power and load distribution of the linear DC motor corresponding to the controlled device are dynamically adjusted according to the actual load conditions, which can flexibly respond to different load conditions and improve the attitude adjustment efficiency of the spacecraft; enable the high-precision linear motion drive system to maintain efficient and stable operation during the spacecraft attitude adjustment process, while minimizing load fluctuations and overload risks.

[0049] By real-time monitoring and precise control of the temperature of the linear DC motor corresponding to the controlled device, dynamic decisions can be made based on historical data and real-time monitoring information, which can adapt to different working environments and changes in working conditions. By calculating parameters such as the temperature deviation coefficient and the temperature reduction coefficient, the heat dissipation strategy can be automatically optimized according to actual conditions, thereby improving the control capability of spacecraft attitude adjustment in complex aerospace environments, reducing the thermal stress caused by overheating of the linear DC motor, and thus extending the service life of the linear DC motor and the entire drive system.

[0050] Furthermore, the specific process for determining the quantitative requirements of the oscillation signal is as follows: B1: If the driving load impact evaluation result after driving load optimization is greater than the preset load fluctuation threshold obtained from the preset database, the oscillation signal impact is quantified based on the parameters in the driving control process, otherwise execute B2; by comparing the driving load impact evaluation result with the threshold, quickly judge whether the load fluctuation exceeds the normal range. If the driving load impact evaluation result after driving load optimization is still greater than the preset load fluctuation threshold, there may be an oscillation signal, and it is necessary to quantify the oscillation signal impact based on the parameters in the driving control process to avoid greater impact on the spacecraft attitude adjustment.

[0051] B2. If the output thrust adjustment time after the drive load is optimized is longer than the preset output thrust adjustment time obtained from the preset database, it indicates that there is a response delay. The oscillation signal impact is quantified based on the parameters in the drive control process. Otherwise, the monitoring of the spacecraft attitude adjustment process continues. By comparing the output thrust adjustment time with the preset value, it is determined whether the motor response is delayed. If the output thrust adjustment time after the drive load is optimized is longer than the preset output thrust adjustment time, a vibration signal may exist.

[0052] In this embodiment, through the dual judgment of two key indicators (drive load impact evaluation results and output thrust adjustment time), it is possible to accurately identify whether there is an oscillation signal impact, avoid misjudgment, provide a basis for subsequent oscillation signal optimization and adjustment, and intervene in time to prevent the oscillation signal from causing serious impact on the spacecraft attitude adjustment.

[0053] Furthermore, the oscillation signal influence is quantified based on the parameters in the drive control process. The specific method is as follows: According to the output thrust adjustment frequency of the linear DC motor corresponding to the device to be controlled in the drive control process and the preset output thrust adjustment frequency obtained from the preset database, the output thrust adjustment comparison coefficient is obtained, that is, Where Z1 represents the output thrust adjustment frequency of the linear DC motor corresponding to the controlled device during the drive control process, and Z 1,0 represents a preset output thrust adjustment frequency; the output thrust adjustment frequency is obtained by analyzing the output thrust change of the linear DC motor corresponding to the device to be controlled within a preset time period through fast Fourier transform; the output thrust change is obtained by calculating the difference between the output thrust at the current preset time point and the output thrust at the previous preset time point within the preset time period; wherein, the preset output thrust adjustment frequency, the preset sensor output signal frequency, and the preset phase deviation are set by the preset personnel; by comparing the output thrust adjustment frequency of the linear DC motor corresponding to the device to be controlled with the preset output thrust adjustment frequency, it is reflected whether the frequency characteristics of the output thrust of the linear DC motor corresponding to the device to be controlled deviate from the predetermined stable state.

[0054] According to the average sensor output signal frequency corresponding to the device to be controlled during the drive control process and the preset sensor output signal frequency obtained from the preset database, the extreme value deviation coefficient of the signal frequency is obtained, that is, Where Z2 represents the average sensor output signal frequency of the device to be controlled during the drive control process, and Z 2,MAX Indicates the preset maximum sensor output signal frequency, Z 2,MINRepresents the preset minimum sensor output signal frequency; the preset sensor output signal frequency includes the preset maximum sensor output signal frequency and the preset minimum sensor output signal frequency. The sensor output signal frequency is obtained by analyzing the sensor output signal within a preset time period through fast Fourier transform. The average sensor output signal frequency represents the average value of the sensor output signal frequencies of the dual sensors. By comparing the sensor output signal frequency with the preset sensor output signal frequency, the difference between the actual sensor signal frequency and the preset sensor output signal frequency is reflected.

[0055] According to the average phase deviation between the dual sensor output signals corresponding to the device to be controlled during the drive control process and the preset phase deviation obtained from the preset database, the phase deviation comparison coefficient is obtained, that is, Where Z3 represents the average phase deviation between the output signals of the two sensors during the drive control process, Z 3,0 Represents a preset phase deviation; the average phase deviation between the dual sensor output signals is obtained by comparing the time delay and signal period of the two sensor signals; by comparing the average phase deviation between the dual sensor output signals and the preset phase deviation, the phase deviation between the dual sensor output signals is reflected.

[0056] The oscillation signal influence quantitative balance parameter is introduced, and the output thrust adjustment contrast coefficient, signal frequency extreme value deviation coefficient and phase deviation contrast coefficient are weighted coupled to obtain the oscillation signal influence quantitative index, which is used to quantitatively evaluate the impact of the oscillation signal on the spacecraft attitude adjustment.

[0057] Among them, the specific limiting expression of the oscillation signal impact index is:

[0058] Where Z represents the oscillation signal influence index in the driving control process, β1 represents the quantitative balance parameter affected by the first oscillation signal, β2 represents the quantitative balance parameter affected by the second oscillation signal, and β3 represents the quantitative balance parameter affected by the third oscillation signal.

[0059] It should be added that the oscillation signal influence quantitative balance parameters are obtained from a preset database, specifically including a first oscillation signal influence quantitative balance parameter, a second oscillation signal influence quantitative balance parameter, and a third oscillation signal influence quantitative balance parameter, which respectively represent the degree of influence of the corresponding parameters based on the drive control process on the oscillation signal influence index; the parameters based on the drive control process include the output thrust adjustment frequency of the linear DC motor corresponding to the controlled device, the average sensor output signal frequency corresponding to the controlled device, and the average phase deviation between the dual sensor output signals; the sum of the three is 1. For example, the parameters based on the drive control process are respectively combined with the corresponding preset oscillation signal influence quantitative balance parameters to form a corresponding oscillation signal influence quantitative mapping set, and the real-time parameters based on the drive control process are respectively input into the corresponding oscillation signal influence quantitative mapping set to obtain the corresponding oscillation signal influence quantitative balance parameters; the oscillation signal influence quantitative mapping set includes a first oscillation signal influence quantitative mapping set, a second oscillation signal influence quantitative mapping set, and a third oscillation signal influence quantitative mapping set.

[0060] In this embodiment, through the above steps, the influence of the vibration signal is comprehensively quantified, and the stability of the spacecraft attitude adjustment can be optimized, thereby improving the accuracy of the spacecraft attitude control, so that the high-precision linear motion DC drive control method can still work stably in complex environments.

[0061] like Figure 5 The interface diagram of the motor signal diagnosis interface shown has the following specific contents: according to the selected posture control device, the corresponding motor and signal information are displayed. The signal information includes time domain characteristics, frequency domain characteristics, waveform display, current signal detection information, gain adjustment suggestions, and oscillation abnormality records; time domain characteristics include average current and peak value; frequency domain characteristics refer to abnormal frequency; waveform display includes real-time current waveform diagram and spectrum analysis diagram; current signal detection information includes oscillation signal amplitude, oscillation frequency, abnormal duration, and cause; maintenance application can be initiated according to gain adjustment suggestions; oscillation abnormality records include time, abnormality type, description, system measures, and status.

[0062] To further illustrate, the specific process of oscillation signal optimization is: In C1, if the oscillation signal impact index is greater than the preset oscillation signal impact quantification threshold obtained from the preset database, it indicates that oscillation signal optimization is required and C2 is executed. Otherwise, oscillation signal optimization is not required and monitoring during spacecraft attitude adjustment continues. The preset oscillation signal impact quantification threshold and preset vibration amplitude are set by the design personnel. By comparing the oscillation signal impact index with the threshold, it is quickly determined whether there is an oscillation signal that needs optimization.

[0063] C2, obtains data from sensors installed at various locations on the spacecraft; the installed sensors include but are not limited to inertial measurement units, temperature sensors, and acceleration sensors; if the sensor output signal fluctuations at various locations are within a preset fluctuation range, oscillation signal optimization is not performed, otherwise oscillation signal optimization is performed according to the set lag time; the sensor output signal fluctuations are represented by the variance of the recorded real-time output signals of the sensors; by obtaining sensor data from various locations on the spacecraft, the operating status of the DC drive system for spacecraft attitude adjustment is comprehensively evaluated, and a reasonable lag time is calculated through the lag time mapping set to avoid optimization that is too early or too late.

[0064] The lag time refers to the time interval after the oscillation signal is detected, and the optimization is delayed for a certain period of time. It is the result of the lag time mapping set obtained by inputting the average sensor signal change rate and the oscillation signal impact quantitative index into the preset database. The preset fluctuation range is the result of mapping the fluctuation range mapping set obtained by inputting the sensor position, oscillation signal impact index, vibration amplitude and vibration frequency of the sensor output signal into the preset database. The average sensor signal change rate is obtained by calculating the difference between consecutive data points (the current data point and the previous data point) to obtain the signal change rate, and then taking the average value within a certain time period to represent it. The vibration amplitude is often described by the peak-to-peak value, that is, the difference between the maximum and minimum values ​​of the signal. The amplitude of each frequency component in the frequency domain is obtained through fast Fourier transform, and the frequency corresponding to the maximum amplitude is the vibration frequency.

[0065] Oscillation signal optimization refers to the adaptive adjustment of the PID control algorithm. The gains in the PID control algorithm include integral gain, proportional gain, and differential gain. The adjustment of PID parameters is completed through fuzzy logic reasoning. Fuzzy control can derive appropriate PID gain values ​​based on the current signal state (such as error and error change rate). The current signal state includes but is not limited to the vibration amplitude and vibration frequency of the sensor output signal.

[0066] In this embodiment, the oscillation signal that needs to be optimized is accurately identified through dual judgment of the oscillation signal impact index and the vibration amplitude of the sensor output signal; the PID control algorithm is optimized to reduce the influence of the oscillation signal on the spacecraft attitude adjustment and the operation of the linear DC motor, thereby improving the stability and reliability of the DC drive system for spacecraft attitude adjustment; the lag time optimized by the PID control algorithm is set by presetting the fluctuation range and the lag time mapping set, thereby avoiding frequent optimization that may interfere with the high-precision linear motion DC drive control method for aerospace.

[0067] As an embodiment of the second aspect, an embodiment of the present invention provides a high-precision linear motion DC drive system for aerospace, such as Figure 6The schematic diagram of the structure of a high-precision linear motion DC drive system for aerospace is shown, comprising: an attitude adjustment demand setting module, an attitude adjustment drive activation module, and an attitude feedback and adjustment module; The attitude adjustment requirement setting module is used to detect the spacecraft's real-time attitude through attitude sensors during attitude adjustment and generate an attitude adjustment signal based on the current spacecraft attitude. Attitude sensors include but are not limited to inertial measurement units, star sensors, and gyroscopes. The spacecraft's real-time attitude includes but is not limited to the spacecraft's pitch, yaw, and roll angles. Based on the attitude error, a control algorithm (such as a PID control algorithm) is used to generate a corresponding attitude adjustment signal, which is used to drive the motor to perform attitude adjustment. The attitude error represents the difference between the target attitude and the current attitude, usually expressed as an angle error. The target attitude is set by a preset person based on mission requirements. By calculating the error between the current and target attitudes in real time, an accurate adjustment signal is generated to ensure that the spacecraft always approaches the target attitude during the attitude adjustment process, reducing attitude error.

[0068] The attitude adjustment drive activation module is used to adjust the spacecraft's attitude according to attitude adjustment signals using linear DC motors. These motors convert their rotational motion into linear or angular attitude adjustments. These rotational motion is converted into the desired linear motion or angular adjustment through a transmission mechanism (such as a ball screw or linear motor). The linear DC motor drive system precisely controls the position of various spacecraft components (such as antennas and solar panels) based on attitude adjustment signals, ensuring they are aligned to the target angle.

[0069] The attitude feedback and adjustment module determines whether to perform interference suppression based on the adjusted spacecraft attitude. If interference suppression is performed, the spacecraft attitude is readjusted after interference suppression is performed. Otherwise, the spacecraft attitude adjustment process continues to be monitored. Interference suppression includes drive load optimization and oscillation signal optimization. The adjusted spacecraft attitude is then interpolated with the target attitude to obtain the adjusted attitude error. If the adjusted attitude error is greater than the preset attitude error (set by the preset personnel), interference suppression is performed; otherwise, the spacecraft attitude adjustment process continues to be monitored. By continuously monitoring attitude changes and performing interference suppression when necessary, the spacecraft's attitude is ensured to be stable during mission execution, avoiding attitude deviations caused by external disturbances or system errors.

[0070] In this embodiment, through precise attitude adjustment requirement setting and drive system, it is possible to ensure that the spacecraft is adjusted to the target attitude with a small error. Especially in the case of microgravity and external disturbances in the aerospace environment, the high-precision linear motion DC drive system for aerospace can effectively respond, reduce the impact of external disturbances on the spacecraft attitude, and improve the accuracy of spacecraft attitude adjustment.

[0071] To summarize, the embodiment of the present application determines the timeliness of the command signal response of the linear DC motor output thrust, and determines the drive load impact assessment requirements based on the command signal response timeliness determination, which is conducive to timely discovering possible response problems of the linear DC motor and providing a basis for subsequent spacecraft attitude adjustment. Then, the drive load impact assessment is performed based on the parameters in the spacecraft attitude adjustment process, and the drive load optimization requirements are determined, which can avoid damage to the linear DC motor caused by unreasonable load. Finally, the oscillation signal impact is quantified based on the parameters in the drive control process, and the oscillation signal optimization requirements are determined, which can prevent the spacecraft attitude adjustment from getting out of control due to oscillation, thereby improving the stability of the spacecraft attitude adjustment.

[0072] Those skilled in the art will appreciate that embodiments of the present invention may be provided as methods, systems, or computer program products. Thus, the present invention may take the form of an entirely hardware embodiment, an entirely software embodiment, or an embodiment combining software and hardware aspects. Furthermore, the present invention may take the form of a computer program product implemented on one or more computer-usable storage media (including but not limited to magnetic disk storage, CD-ROM, optical storage, etc.) containing computer-usable program code.

[0073] The present invention is described with reference to flowcharts and / or block diagrams of systems, devices (systems), and computer program products according to embodiments of the present invention. It should be understood that each process and / or block in the flowcharts and / or block diagrams, as well as combinations of processes and / or blocks in the flowcharts and / or block diagrams, can be implemented by computer program instructions. These computer program instructions can be provided to a processor of a general-purpose computer, a special-purpose computer, an embedded processor, or other programmable data processing device to produce a machine, so that the instructions executed by the processor of the computer or other programmable data processing device generate instructions for implementing the processes in the flowcharts and / or block diagrams. Figure 1 a process or multiple processes and / or boxes Figure 1 A device that provides the functions specified in a block or multiple blocks.

[0074] These computer program instructions may also be stored in a computer readable memory that can direct a computer or other programmable data processing device to work in a specific manner, so that the instructions stored in the computer readable memory produce an article of manufacture comprising an instruction device, which implements the process Figure 1 a process or multiple processes and / or boxes Figure 1 The function specified in one or more boxes.

[0075] These computer program instructions can also be loaded onto a computer or other programmable data processing device so that a series of operational steps are executed on the computer or other programmable device to produce a computer-implemented process, thereby providing the instructions executed on the computer or other programmable device for implementing the process. Figure 1 a process or multiple processes and / or boxes Figure 1 A step that specifies a function in one or more boxes.

[0076] Although the preferred embodiments of the present invention have been described, those skilled in the art may make additional changes and modifications to these embodiments once they have learned the basic creative concept. Therefore, the appended claims are intended to be interpreted as including the preferred embodiments and all changes and modifications that fall within the scope of the present invention.

[0077] Obviously, those skilled in the art may make various changes and modifications to the present invention without departing from the spirit and scope of the present invention. Thus, if such changes and modifications fall within the scope of the claims and their equivalents, the present invention is intended to include such changes and modifications.

[0078] It should be understood that in various embodiments of the present invention, the size of the serial numbers of the above-mentioned processes does not mean the order of execution. The execution order of each process should be determined by its function and internal logic, and should not constitute any limitation on the implementation process of the embodiments of the present invention.

[0079] Those skilled 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. Professionals and technicians can use different methods to implement the described functions for each specific application, but such implementation should not be considered beyond the scope of the present invention.

[0080] Those skilled in the art will clearly understand that, for the convenience and brevity of description, the specific working processes of the above-described equipment, devices and units can refer to the corresponding processes in the aforementioned method embodiments and will not be repeated here.

[0081] The above description is merely a specific embodiment of the present invention, but the scope of protection of the present invention is not limited thereto. Any modifications or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in the present invention should be included within the scope of protection of the present invention. Therefore, the scope of protection of the present invention should be based on the scope of protection of the claims.

Claims

1. A high-precision linear motion DC drive control method for aerospace, characterized in that: The following steps are involved: S1, before adjusting the spacecraft attitude, performing a timeliness assessment of the command signal response of the linear DC motor output thrust, and determining the drive load impact assessment requirement based on the timeliness assessment of the command signal response; S2: If there is no need for drive load impact assessment, continue to monitor the spacecraft attitude adjustment process; otherwise, perform drive load impact assessment based on the parameters during the spacecraft attitude adjustment process and determine the drive load optimization requirement; if there is a drive load optimization requirement, determine the oscillation signal impact quantification requirement after the drive load optimization; otherwise, directly perform oscillation signal impact quantification based on the parameters during the drive control process; S3: If there is a need to optimize the oscillation signal, it is determined whether to provide feedback to the preset personnel based on the result of the timeliness judgment of the command signal response after the oscillation signal optimization; otherwise, the spacecraft attitude adjustment process continues to be monitored.

2. The high-precision linear motion DC drive control method for aerospace according to claim 1, characterized in that: The specific method for determining the timeliness of the command signal response of the linear DC motor output thrust is as follows: Monitor the output thrust of the linear DC motor corresponding to the controlled device in real time during spacecraft attitude adjustment, and obtain the output thrust adjustment duration; If the output thrust adjustment time is greater than the preset output thrust adjustment time obtained from the preset database, the drive load impact assessment is performed based on the parameters during the spacecraft attitude adjustment process, otherwise the monitoring of the spacecraft attitude adjustment process continues.

3. The high-precision linear motion DC drive control method for aerospace according to claim 1, characterized in that: The specific method for evaluating the driving load impact based on the parameters during the spacecraft attitude adjustment process is as follows: Comparing an average attitude change of the spacecraft during the spacecraft attitude adjustment process with a preset attitude change obtained from a preset database to obtain an attitude change comparison coefficient, wherein the average attitude change represents an average value of the spacecraft attitude change within a preset time period, and the spacecraft attitude change is represented by the absolute value of a deviation result obtained by performing deviation processing on an attitude angle at a current preset time point obtained by an inertial measurement unit and an attitude angle at a previous preset time point; A heating rate comparison coefficient is obtained by comparing a motor heating rate of a linear DC motor corresponding to a device to be controlled during spacecraft attitude adjustment with a preset motor heating rate obtained from a preset database, wherein the motor heating rate is represented by an average value of deviations between the temperature at a current preset time point and the temperature at a previous preset time point, which are acquired by a temperature sensor; performing an inverse comparison process on an output thrust stability coefficient of a linear DC motor corresponding to a device to be controlled during spacecraft attitude adjustment and a preset output thrust stability coefficient obtained from a preset database to obtain an output thrust stability inverse comparison coefficient, wherein the output thrust stability coefficient is obtained by comparing a standard deviation and an average value of the output thrust of the linear DC motor collected within a preset time period; A drive load impact assessment balance parameter is introduced, and a weighted coupling process is performed on the attitude change contrast coefficient, the heating rate contrast coefficient, and the output thrust stability inverse contrast coefficient to obtain the drive load impact assessment result. The drive load impact assessment result is used to quantitatively evaluate the influence of the load fluctuation of the linear DC motor on the spacecraft attitude adjustment process.

4. The high-precision linear motion DC drive control method for aerospace use according to claim 3, characterized in that: The specific process of determining the drive load optimization requirement is as follows: A1: If the driving load impact evaluation result is greater than a preset load fluctuation threshold obtained from a preset database, there is a need for driving load optimization, and A2 is executed; otherwise, there is no need for driving load optimization, and the subsequent oscillation signal impact quantification is directly performed; A2: If the load of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor load obtained from the preset database, the first drive load is optimized and A3 is executed; otherwise, A3 is directly executed; A3: If the output thrust of the linear DC motor corresponding to the device to be controlled is greater than the first preset linear DC motor output thrust obtained from the preset database, the loads of the linear DC motors corresponding to the device to be controlled are distributed in parallel and A4 is executed; otherwise, A4 is directly executed; A4: If the output thrust of the linear DC motor corresponding to the device to be controlled is greater than the second preset linear DC motor output thrust obtained from the preset database, the second driving load is optimized; otherwise, feedback is provided to the preset personnel.

5. The high-precision linear motion DC drive control method for aerospace use according to claim 4, characterized in that: The first drive load optimization means correcting the output thrust of the linear DC motor corresponding to the controlled device according to the thrust correction factor; The thrust correction factor represents the result obtained by inputting the load deviation coefficient and the drive load impact evaluation result deviation coefficient into the thrust correction mapping set obtained in the preset database; The second drive load optimization means correcting the output power of the linear DC motor corresponding to the device to be controlled according to the output power correction factor; The load deviation coefficient is obtained by performing relative deviation processing on the load of the linear DC motor corresponding to the controlled device and a preset linear DC motor load, and the drive load impact evaluation result deviation coefficient is obtained by performing relative deviation processing on the drive load impact evaluation result and a preset load fluctuation threshold; The output power correction factor represents a result obtained by inputting an output thrust deviation coefficient and a drive load impact evaluation result deviation coefficient into an output power correction mapping set obtained from a preset database. The output thrust deviation coefficient is obtained by performing relative deviation processing on the output thrust of the linear DC motor corresponding to the controlled device and the output thrust of a second preset linear DC motor.

6. The high-precision linear motion DC drive control method for aerospace use according to claim 5, characterized in that: The drive load optimization further includes: Determine whether the temperature of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor temperature obtained from the preset database: If the temperature of the linear DC motor corresponding to the device to be controlled is greater than the preset linear DC motor temperature obtained from the preset database, the heat dissipation device is activated according to the heat dissipation device activation time, otherwise the heat dissipation device is not activated; The duration of the heat dissipation device being on is obtained by assigning a current linear DC motor temperature deviation coefficient according to a temperature reduction coefficient. The temperature reduction coefficient is obtained by comparing the temperature reduction duration in a historical time period with the linear DC motor temperature deviation coefficient. The linear DC motor temperature deviation coefficient is obtained by performing relative deviation processing on the temperature of the linear DC motor corresponding to the device to be controlled and a preset linear DC motor temperature.

7. The high-precision linear motion DC drive control method for aerospace use according to claim 6, characterized in that: The specific process of determining the quantization requirement of the oscillation signal impact is as follows: B1: If the driving load impact evaluation result after driving load optimization is greater than the preset load fluctuation threshold obtained from the preset database, the oscillation signal impact is quantified based on the parameters in the driving control process, otherwise, B2 is executed; B2. If the output thrust adjustment duration after the drive load optimization is greater than the preset output thrust adjustment duration obtained from the preset database, the oscillation signal impact is quantified based on the parameters in the drive control process. Otherwise, the monitoring of the spacecraft attitude adjustment process is continued.

8. The high-precision linear motion DC drive control method for aerospace use according to claim 7, characterized in that: The quantification of the oscillation signal influence based on the parameters in the drive control process is performed in the following specific method: An output thrust adjustment comparison coefficient is obtained by comparing an output thrust adjustment frequency of the linear DC motor corresponding to the device to be controlled during the drive control process with a preset output thrust adjustment frequency obtained from a preset database, wherein the output thrust adjustment frequency is obtained by analyzing a change in the output thrust of the linear DC motor corresponding to the device to be controlled within a preset time period through fast Fourier transform; performing extreme value deviation processing on an average sensor output signal frequency corresponding to the device to be controlled during the drive control process and a preset sensor output signal frequency obtained from a preset database to obtain a signal frequency extreme value deviation coefficient, wherein the preset sensor output signal frequency includes a preset maximum sensor output signal frequency and a preset minimum sensor output signal frequency, and the sensor output signal frequency is obtained by analyzing the sensor output signal within a preset time period through fast Fourier transform; Comparing the average phase deviation between the output signals of the dual sensors corresponding to the device to be controlled during the drive control process with the preset phase deviation obtained from the preset database, a phase deviation comparison coefficient is obtained; The oscillation signal influence quantitative balance parameter is introduced, and the output thrust adjustment contrast coefficient, signal frequency extreme value deviation coefficient and phase deviation contrast coefficient are weighted coupled to obtain the oscillation signal influence quantitative index, which is used to quantitatively evaluate the influence of the oscillation signal on the spacecraft attitude adjustment.

9. The high-precision linear motion DC drive control method for aerospace use according to claim 8, characterized in that: The specific process of the oscillation signal optimization is as follows: C1: If the oscillation signal impact index is greater than the preset oscillation signal impact quantization threshold obtained from the preset database, it indicates that there is a need for oscillation signal optimization and C2 is executed. Otherwise, it indicates that there is no need for oscillation signal optimization and the monitoring of the spacecraft attitude adjustment process continues; C2, acquires data from sensors installed at various locations on the spacecraft. If the fluctuation of the sensor output signal at each location is within a preset fluctuation range, no oscillation signal optimization is performed. Otherwise, oscillation signal optimization is performed based on a preset lag time. The lag time represents the result obtained by inputting the average sensor signal change rate and the oscillation signal impact quantitative index into the preset database to obtain the lag time mapping set; the preset fluctuation range represents the result obtained by inputting the position of each sensor, the oscillation signal impact index, the vibration amplitude and the vibration frequency of the sensor output signal into the preset database to obtain the fluctuation range mapping set.

10. A high-precision linear motion DC drive system for aerospace, applying the high-precision linear motion DC drive control method for aerospace according to any one of claims 1 to 9, characterized in that: The system includes: a posture adjustment requirement setting module, a posture adjustment drive activation module and a posture feedback and adjustment module; The attitude adjustment requirement setting module is used to detect the real-time attitude of the spacecraft through the attitude sensor when the spacecraft performs attitude adjustment, and generate an attitude adjustment signal based on the current spacecraft attitude; The attitude adjustment drive activation module is used to adjust the attitude of the spacecraft through the linear DC motor according to the attitude adjustment signal; The attitude feedback and adjustment module is used to determine whether to perform interference suppression based on the adjusted spacecraft attitude. If interference suppression is performed, the spacecraft attitude is readjusted after the interference suppression. Otherwise, the spacecraft attitude adjustment process continues to be monitored. The interference suppression includes drive load optimization and oscillation signal optimization.

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