Testing device for monitoring separation and locking postures of host satellite and boarding load

Abstract

The invention discloses a test device for monitoring separation and attitude locking of a host satellite and a boarding load, belongs to the field of mechanical design and test, and aims to overcome the defects of insufficient simulation in a microgravity environment and inaccurate satellite attitude data acquisition in an existing boarding load and host satellite separation test. The air foot supporting system enables the simulation device to be in a free floating state; the simulation device simulates the separation of the host satellite and the boarding load; the monitoring device monitors dynamic parameters in the separation process of the host satellite and the hosting load; the host satellite and the hosting load are connected through a mechanical interface system, and the host satellite and the hosting load which are connected are driven to be locked or separated through locking or separating of the mechanical interface system; the push rod motor is used for adjusting the inclination angle of the mechanical interface system so as to adjust the separation angle of the host satellite and the boarding load. The method is used for simulating separation of the host satellite and the hosting load, and attitude data in the separation process is obtained for subsequent analysis.

Description

Technical Field

[0001] The invention relates to the field of mechanical design and testing, and belongs to a test for separation of a host satellite and a hosted payload. Background Art

[0002] In the field of space technology, as satellite missions become increasingly complex, the management and separation technology of host satellites and hosted payloads has gradually become a hot topic in research and aerospace practice. Hosted payloads are small devices or modules attached to host satellites, usually used to perform specific tasks such as scientific experiments, data collection or communication assistance. Ensuring that these hosted payloads can be quickly and safely separated from the host satellite when needed is one of the key factors for the success of satellite missions.

[0003] Traditional methods of separating hosted payloads mainly rely on robotic arms or explosive bolts. These methods face many challenges and limitations in practical applications. The robotic arm requires highly precise control and complex programming, and its size and weight impose a heavy burden on the overall design of the satellite, increasing the launch cost. Although explosive bolts can achieve rapid separation, their one-time use characteristics and the shock and vibration generated by the explosion may cause unpredictable damage to surrounding equipment. In addition, these methods have slow response speeds in emergency situations (such as satellite failures or collision threats) and cannot effectively protect hosted payloads.

[0004] With the increase in space mission requirements and the development of technology, there is an urgent need for a system that can achieve rapid and safe separation of hosted payloads from host satellites without relying on human intervention. In addition, in existing research and engineering practices, the experiments to achieve separation of hosted payloads from host satellites have the shortcomings of insufficient simulation in microgravity environments and inaccurate satellite attitude data, resulting in many uncertainties and risks in separation operations in actual missions.

[0005] Among them, the inaccurate satellite attitude data is mainly due to the fact that in the existing separation tests of the hosted payload and the host satellite, traditional monitoring technology is used to monitor the satellite attitude, that is, it mainly relies on a single sensor or a single type of acquisition equipment (such as image sensor or inertial sensor) to obtain satellite attitude data. These methods have some obvious shortcomings. First, it is difficult for a single sensor to obtain comprehensive attitude change information during the complex separation process, resulting in inaccurate data. In addition, due to the severe vibration and impact generated during separation, the accuracy and reliability of a single sensor are easily affected, which may lead to attitude data deviation. Although the existing image acquisition technology can provide visual data, when the attitude changes drastically or the lighting conditions change, the acquisition and processing of data will be limited, making it difficult to accurately reflect the motion state of the satellite in real time. The use of a single inertial sensor may increase the attitude calculation error because it cannot consider attitude changes in multiple directions at the same time. Therefore, the existing method fails to obtain accurate satellite attitude. Summary of the invention

[0006] The purpose of the present invention is to solve the problems that the existing separation test of hosted payload and host satellite has the shortcomings of insufficient simulation in microgravity environment and inaccurate satellite attitude data, which leads to many uncertainties and risks in the separation operation in actual missions. A test device for monitoring the separation and locking attitude of the host satellite and hosted payload is proposed.

[0007] A test device for monitoring the separation and locking attitude of a host satellite and a hosted payload, the test device comprising a simulation device, an air-foot support system and a monitoring device;

[0008] An air foot support system is installed at the bottom of the simulation device. By introducing gas into the top of the air pipe on the air foot support system, the gas is ejected from the bottom of the air pipe, so that the simulation device is in a free floating state, simulating the floating state in space.

[0009] The simulation device includes the host satellite, hosted payload, mechanical interface system and push rod motor;

[0010] The host satellite and the hosted payload are connected via a mechanical interface system, and the mechanical interface system is used to lock or separate, thereby driving the connected host satellite and hosted payload to lock or separate;

[0011] Push rod motor, used to adjust the tilt angle of the mechanical interface system, thereby adjusting the separation direction of the host satellite and the hosted payload;

[0012] Monitoring device, used to monitor dynamic parameters during the separation process of host satellite and hosted payload.

[0013] Preferably, the host satellite comprises a host satellite frame, a host satellite side plate and a host satellite bottom plate;

[0014] The host satellite frame is a cube frame, the host satellite side panels are rotatably connected to the side frames of the main satellite frame by rotating bolts, and the host satellite bottom plate is fixed to the bottom frame of the host satellite frame by bolts;

[0015] The boarding load includes a boarding load frame, a boarding load side plate and a boarding load bottom plate;

[0016] The host load frame is a cubic frame, the host load side plate is rotatably connected to the side frame of the host load frame by rotating bolts, and the host load bottom plate is fixed to the bottom frame of the host load frame by bolts; and the host load side plate is arranged opposite to the host satellite side plate;

[0017] The mechanical interface system includes a first separation component and a second separation component;

[0018] Separation piece No. 1 is arranged in the host payload frame, and separation piece No. 2 is arranged in the host satellite frame;

[0019] Through holes are provided at the same positions on the host satellite side plate and the host load side plate, and the host satellite side plate and the host load side plate are parallel and oppositely arranged;

[0020] The No. 1 separation piece passes through the through holes on the side plate of the host payload and the side plate of the host satellite in sequence, and is plugged into the No. 2 separation piece. The No. 1 separation piece and the No. 2 separation piece can be locked or separated.

[0021] Preferably, the first separation component includes an antenna satellite substrate, an ejection plate interface and a threaded cover;

[0022] The host payload, the antenna satellite base plate and the ejection plate interface are connected in sequence, and the threaded cover is put on the ejection plate interface;

[0023] Separation part No. 2 includes a power spring device, a locking steel ball, a spring plate, a base and a release motor;

[0024] The ejection plate interface passes through the through holes on the host payload side plate and the host satellite side plate in turn, and is plugged into the base through the compression power spring device; the No. 1 separation piece is restricted in the host payload frame through the threaded cover;

[0025] The base includes a hollow pillar and two circular support plates. The first circular support plate is fixed at the bottom of the hollow pillar, and the second circular support plate is sleeved and fixed on the hollow pillar. Through holes are evenly arranged on the side wall of the hollow pillar at the upper part of the second circular support plate. An annular groove is arranged on the outer circumferential surface of the ejection plate interface.

[0026] The spring plate is sleeved on the hollow pillar and supported on the second annular support plate. The tops of the two release motors are in contact with the bottom surface of the spring plate. When the two release motors are not started, the spring plate squeezes and fixes the locking steel ball between the through hole and the groove of the ejection plate interface to achieve the locking of the host satellite and the hosted payload; when the two release motors are started, the spring plate is driven to move upward along the base so that the annular groove on the inner circumference of the spring plate is docked with the through hole, and the locking steel ball enters from the annular groove on the ejection plate interface between the through hole and the annular groove on the spring plate to achieve the separation of the host satellite and the hosted payload;

[0027] Preferably, the second separation member further comprises a power regulating motor;

[0028] A power regulating motor is arranged between the base and the power spring device;

[0029] The power regulating motor is used to drive the power spring device to move upward and compress the power spring device.

[0030] Preferably, the second separation element further comprises a laser sensor;

[0031] Laser sensor is used to measure the relative displacement of the host satellite side plate and the hosted payload side plate.

[0032] Preferably, the monitoring device includes an acceleration sensor, an IMU inertial navigation measurement instrument and a dynamic capture mechanism;

[0033] Both the host satellite and the hosted payload are equipped with acceleration sensors and IMU inertial navigation measurement instruments to monitor the acceleration and angular velocity of the host satellite and the hosted payload in real time;

[0034] The dynamic capture mechanism is used to capture the relative displacement and respective acceleration of the host satellite and hosted payload in real time.

[0035] Preferably, the test device further comprises a host computer;

[0036] A processor is provided in the host computer, which is used to sequentially perform time alignment and space alignment on the relative displacement of the host satellite and the hosted payload captured by the dynamic capture mechanism in real time and the relative displacement output by the laser sensor in real time, and perform data fusion on the two relative displacements after spatial alignment to obtain the final relative displacement; at the same time, the acceleration of the host satellite output by the dynamic capture mechanism in real time and the acceleration of the host satellite output by the acceleration sensor located in the host satellite are sequentially performed time alignment and space alignment, and the two accelerations after spatial alignment are data fused to obtain the final host satellite acceleration, and the acceleration of the hosted payload output by the dynamic capture mechanism in real time and the acceleration of the hosted payload output by the acceleration sensor located in the hosted payload are sequentially performed time alignment and space alignment, and the two accelerations after spatial alignment are data fused to obtain the final hosted payload acceleration;

[0037] It is also used to receive the angular velocity of the host satellite and the hosted payload, and display the final hosted payload acceleration, the final host satellite acceleration and the final relative displacement of the host satellite and the hosted payload.

[0038] Preferably, the process of time registration is:

[0039] The acceleration of the host satellite captured by the dynamic capture mechanism in real time is fitted into a plurality of first host satellite acceleration curves, and the acceleration of the host satellite output by the acceleration sensor in real time is fitted into a plurality of second host satellite acceleration curves; the acceleration of the host payload captured by the dynamic capture mechanism in real time is fitted into a plurality of first host payload acceleration curves, and the acceleration of the host payload output by the acceleration sensor in real time is fitted into a plurality of second host payload acceleration curves; the relative displacement captured by the dynamic capture mechanism in real time is fitted into a plurality of first displacement curves, and the relative displacement output by the laser sensor in real time is fitted into a plurality of second displacement curves, and each fitted displacement curve and acceleration curve is expressed as:

[0040]

[0041] In the formula, a0, a1, a2 are the polynomial coefficients of the fitting curve, x i is the displacement information at the i-th moment measured by the laser sensor or dynamic capture mechanism, t i is the measurement time at the i-th moment, j represents the polynomial order of the fitting curve;

[0042] The error between each first host satellite acceleration curve and the acceleration acquisition value is calculated as the first host satellite acceleration error, the error between each second host satellite acceleration curve and the acceleration acquisition value is calculated as the second host satellite acceleration error, the error between each first host payload acceleration curve and the acceleration acquisition value is calculated as the first host payload acceleration error, the error between each second host payload acceleration curve and the acceleration acquisition value is calculated as the second host payload acceleration error, the error between each first displacement curve and the displacement acquisition value is calculated as the first displacement error, and the error between each second displacement curve and the displacement acquisition value is calculated as the second displacement error. The error δ i It is expressed as:

[0043]

[0044] In the formula, P(t i ) is the acceleration information at the i-th moment in the first host satellite acceleration curve, the second host satellite acceleration curve, the first hosted payload acceleration curve, or the second hosted payload acceleration curve, or the displacement information at the i-th moment in the first displacement curve, or the displacement information at the i-th moment in the second displacement curve;

[0045] A first host satellite acceleration curve corresponding to a minimum value is selected from multiple first host satellite acceleration errors as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value is selected from multiple second host satellite acceleration errors as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value is selected from multiple first host payload acceleration errors as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value is selected from multiple second host payload acceleration errors as the second host payload acceleration curve after screening, a first displacement curve corresponding to a minimum value is selected from multiple first displacement errors as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value is selected from multiple second displacement errors as the second displacement curve after screening.

[0046] Preferably, the process of spatial registration is:

[0047] The state coordinate systems of the first screened host satellite acceleration curve, the second screened host satellite acceleration curve, the first screened hosted payload acceleration curve, the second screened hosted payload acceleration curve, the first screened displacement curve and the second screened displacement curve are all converted into ground coordinate systems.

[0048] Preferably, the process of data fusion is:

[0049] According to the data fusion formula of formula 3, the first host satellite acceleration curve and the second host satellite acceleration curve after screening in the ground coordinate system are fused to obtain the final host satellite acceleration curve, the first host load acceleration curve and the second host load acceleration curve after screening in the ground coordinate system are fused to obtain the final host load acceleration curve, and the first displacement curve and the second displacement curve after screening in the ground coordinate system are fused to obtain the final displacement curve:

[0050]

[0051] In the formula, is the fused acceleration curve or fused displacement curve, W i is the weight of the i-th acceleration curve or the i-th displacement curve, i = 1, 2, X i is the displacement information or acceleration measured by the i-th sensor;

[0052] Among them, W i According to formula 4, we can obtain:

[0053]

[0054] In the formula, σ 2 is the total variance, E represents the expectation, and X represents the true value of the target information. is the error variance measured by the i-th sensor, is the minimum total variance of all sensor measurements.

[0055] The beneficial effects of the present invention are:

[0056] The present invention provides a host satellite and hosted payload separation and locking attitude monitoring test device, which can realize accurate simulation and testing of the separation process of the host satellite and hosted payload in a microgravity environment, thereby optimizing the design of the mechanical interface and improving the reliability and safety of the separation process.

[0057] Compared with these traditional methods, the present invention adopts a combination of multiple acquisition devices (dynamic capture camera system, acceleration sensor and laser sensor) to realize multi-dimensional and high-precision attitude data acquisition during the satellite separation process, and through data fusion processing technology, improves the accuracy of the data and the reliability of subsequent analysis.

[0058] The mechanical interface system of the present invention has automatic adaptability and a rapid response mechanism, and can quickly perform a separation operation in an emergency without human intervention, thereby improving the safety of load protection.

[0059] The device of the present invention has a compact structure and a high degree of integration, reduces the volume and weight of the equipment, reduces the launch cost, and improves the overall efficiency of the system.

[0060] The microgravity environment simulation system based on the air floating platform of the present invention has the advantages of providing a real microgravity environment, making the test results closer to the actual space conditions, and improving the accuracy and reliability of the test.

[0061] The present invention adopts high-precision acceleration sensor and IMU inertial navigation system to collect data, which has the advantages of accurately measuring the changes of acceleration and angular velocity during the separation process, providing high-precision data support, and improving the accuracy of experimental analysis.

[0062] The present invention adopts a dynamic capture camera system to record the attitude changes of the host satellite and the hosted payload in real time, providing visual data support, which is helpful for detailed analysis of the attitude changes during the separation process.

[0063] The present invention adopts a multi-sensor (dynamic capture camera system, acceleration sensor and laser sensor) information fusion mechanism to fuse the information collected by each sensor, reduce the cumulative error caused by the acceleration sensor, improve the precision, and enhance the accuracy and reliability of the test.

[0064] The present invention has an adjustable spring stiffness test mechanism, which can study the influence of spring stiffness on separation speed and direction, optimize the mechanical interface design, and ensure the stability and safety of the separation process.

[0065] The test device of the present invention adopts an air foot support system, an acceleration sensor, an IMU inertial navigation system and six dynamic capture camera technologies. The air foot support system is mainly used to carry out microgravity unloading of the host satellite and the hosted payload to simulate the floating state in space. The air foot further realizes the support in the microgravity environment by spraying air, so that the host satellite and the hosted payload can maintain a stable suspension state during the test. The acceleration sensor is used to measure the acceleration of the host satellite and the hosted payload, and the velocity change and displacement change are obtained by integration, so as to analyze the dynamic characteristics during the separation process. The IMU inertial navigation system is used to measure the angular velocity of the host satellite and the hosted payload, and the angle change is obtained by integration, providing data support for attitude control and adjustment. The dynamic capture camera technology is used to capture the attitude changes of the host satellite and the hosted payload, and the attitude change process is recorded through images to provide intuitive test data.

[0066] The present invention mainly studies the separation test of host satellite and hosted payload based on mechanical interface. In orbit, both host satellite and hosted payload are in floating state, so when conducting the test, it is necessary to use gas foot support to create microgravity phenomenon. This experiment mainly studies the influence of the stiffness of the spring in the mechanical interface on the separation speed and separation direction of the host satellite and hosted payload. In addition, the center of mass of the host satellite and the hosted payload may not be along the separation direction of the spring, so it is easy to have attitude changes during the separation process, especially attitude changes such as rotation. If the separation speed is slow, there is a risk of collision between the host satellite and the hosted payload, which is professionally called near-field collision. In addition, since there may be processing errors in the design of the mechanical interface, the mechanical interface itself is prone to the separation direction not along the axial direction of the spring, and thus there is also a risk of collision between the host satellite and the hosted payload.

[0067] This experiment mainly focuses on the following points: First, the influence of the stiffness of the spring in the mechanical interface on the separation speed and separation direction of the host satellite and the hosted payload; second, the influence of the separation speed and separation direction on the attitude of the host satellite and the hosted payload; third, the influence of the processing error based on the mechanical interface on the attitude of the host satellite and the hosted payload.

[0068] The present invention conducts an in-depth study on the influence of the machining error of the mechanical interface on the separation process through error analysis and optimization design, and proposes an optimization design scheme. This not only improves the manufacturing accuracy of the mechanical interface, but also reduces the errors and risks that may occur during the separation process. Through precise measurement and analysis, the present invention can optimize the design of the mechanical interface to ensure the rapid and safe separation of the host satellite and the hosted payload. This is of great significance for protecting the safety of the hosted payload in emergency situations (such as host satellite failure or collision threat) and ensuring the smooth execution of subsequent tasks. In addition, the test device of the present invention is compact in design and easy to operate, which reduces the manufacturing and operating costs of the test equipment. Its efficient testing capabilities and versatility make it have significant economic benefits and practical value in scientific research and engineering applications. In these ways, the present invention solves the limitations of traditional separation methods, improves the flexibility, adaptability and safety of the separation process, and provides reliable technical support for future space missions. BRIEF DESCRIPTION OF THE DRAWINGS

[0069] Figure 1 A schematic diagram of the structure of the test device for monitoring the separation and locking attitude of the host satellite and the hosted payload;

[0070] Figure 2 A schematic diagram of a structure in which the simulation device is made to float freely based on the air flotation method;

[0071] Figure 3 It is the three-view drawing of the simulation device, where: Figure 3 (a) is the front view of the simulation device. Figure 3 (a) is the left view of the simulation device. Figure 3 (c) is a top view of the simulation device;

[0072] Figure 4 Schematic diagram of the push rod motor changing the inclination angle of the host satellite side panel;

[0073] Figure 5 This is a schematic diagram of the overall structure of the host satellite;

[0074] Figure 6 This is a schematic diagram of the host satellite base plate structure;

[0075] Figure 7 This is a schematic diagram of the host satellite side panel structure;

[0076] Figure 8 It is a schematic diagram of the overall structure of the mechanical interface system;

[0077] Fig. 9 It is the exploded diagram of the mechanical interface system;

[0078] Fig.10 It is a state diagram of the mechanical interface system in the locked state;

[0079] Fig.11 for Fig.10 Middle AA section view;

[0080] Fig.12 It is a schematic diagram of the separation process of the mechanical interface system;

[0081] Fig.13 Schematic diagram of the separation dynamics of the host satellite and the hosted payload;

[0082] Fig.14 This is the overall test flow chart of the test device for monitoring the separation and locking attitude of the host satellite and the hosted payload. DETAILED DESCRIPTION

[0083] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[0084] It should be noted that, in the absence of conflict, the embodiments of the present invention and the features in the embodiments may be combined with each other.

[0085] The present invention will be further described below in conjunction with the accompanying drawings and specific embodiments, but they are not intended to limit the present invention.

[0086] Example:

[0087] Combination Figures 1 to 3 This embodiment is described, a test device for monitoring the separation and locking attitude of a host satellite and a hosted payload, the device comprising a simulation device 1, an air-foot support system 2 and a monitoring device;

[0088] An air foot support system 2 is installed at the bottom of the simulation device 1. By introducing gas into the top of the air pipe 2-1 on the air foot support system 2, the gas is ejected from the bottom of the air pipe 2-1, so that the simulation device 1 is in a free floating state, simulating the floating state in space; the simulation device 1 is used to simulate the separation of the host satellite 1-1 and the hosted payload 1-2; the monitoring device is used to monitor the dynamic parameters during the separation process of the host satellite 1-1 and the hosted payload 1-2;

[0089] The simulation device 1 includes a host satellite 1-1, a hosted payload 1-2, a mechanical interface system 1-3 and a push rod motor 1-4;

[0090] The host satellite 1-1 and the hosted payload 1-2 are connected via a mechanical interface system 1-3, and the mechanical interface system 1-3 is used to lock or separate, thereby driving the connected host satellite 1-1 and the hosted payload 1-2 to lock or separate;

[0091] The push rod motor 1-4 is used to adjust the tilt angle of the mechanical interface system 1-3, thereby adjusting the separation angle of the host satellite 1-1 and the hosted payload 1-2.

[0092] Specifically, Figure 1 The reference numeral 3 in the figure is an air-floating platform, and the air-floating platform 4 is located at the bottom of the air-foot support system 2. The air-foot support system 2 is mainly used to simulate the microgravity environment, so that the host satellite 5 and the hosted payload 6 can float freely during the test. After floating, the air-foot support system 2 is separated from the air-floating platform 4. The design of the air-foot support system 2 takes into account the uniformity and stability of the airflow to ensure that the host satellite and the hosted payload can maintain a stable suspension state during the test, thereby simulating the actual operating conditions in the space environment. The air-foot support system 4 is an important part of the air-floating platform. The air-foot support system 4 forms an air cushion by spraying air between the air foot and the ground, thereby reducing the contact friction between the object and the ground, so that the object can float on the air cushion, similar to the state in the microgravity environment. By adjusting the air pressure and flow of the air cushion, the gravity of the object is balanced by the air buoyancy, thereby achieving the effect of unloading gravity. In this way, the net force on the object on the air foot is close to zero, simulating the microgravity environment. The gas introduced into the air pipe 2-1 is nitrogen.

[0093] There are two push rod motors 1-4, which are respectively distributed on both sides of the host satellite and are mainly used to adjust the ejection direction of the mechanical interface system 1-3. Figure 3 As shown. The push rod motor precisely controls the ejection direction and force to ensure that the hosted payload can move along the predetermined trajectory during the separation process to avoid collision with the host satellite. The design of the push rod motor takes into account a variety of working modes and control algorithms, and can flexibly adjust the ejection parameters according to test requirements to optimize the effect of the separation process.

[0094] The structural composition of the host satellite 1-1, the hosted payload 1-2 and the mechanical interface system 1-3 is further defined below:

[0095] The host satellite 1-1 includes a host satellite frame 1-1-1, a host satellite side plate 1-1-2 and a host satellite bottom plate 1-1-3;

[0096] The host satellite side plate 1-1-2 is rotatably connected to the side frame of the main satellite frame 1-1-1 by rotating bolts 1-1-4, and the host satellite bottom plate 1-1-3 is fixed to the bottom frame of the host satellite frame 1-1-1 by bolts;

[0097] The boarding load 1-2 includes a boarding load frame 1-2-1, a boarding load side plate 1-2-2 and a boarding load bottom plate 1-2-3;

[0098] The host load side plate 1-2-2 is rotatably connected to the side frame of the host load frame 1-2-1 by rotating bolts 1-1-4, and the host load bottom plate 1-2-3 is fixed to the bottom frame of the host load frame 1-2-1 by bolts; and the host load side plate 1-2-2 is arranged opposite to the host satellite side plate 1-1-2;

[0099] The mechanical interface system 1-3 includes a separation component No. 1 and a separation component No. 2;

[0100] Separator No. 1 is arranged in the host payload frame 1-2-1, and separator No. 2 is arranged in the host satellite frame 1-1-1;

[0101] Through holes are provided at the same positions on the host satellite side plate 1-1-2 and the host load side plate 1-2-2, and the host satellite side plate 1-1-2 and the host load side plate 1-2-2 are parallel and oppositely arranged;

[0102] Separation piece No. 1 passes through the through holes on the host payload side plate 1-2-2 and the host satellite side plate 1-1-2 in sequence, and is plugged into separation piece No. 2, so that separation piece No. 1 and separation piece No. 2 can be locked or separated.

[0103] Specifically, the host satellite 1-1 is used as the main carrier to install the hosted payload 1-2 and the mechanical interface system 1-3, such as Figure 2 As shown. The structural design of the host satellite 5 takes into account multiple installation positions and interfaces, and can flexibly adapt to different types of hosted payloads and mechanical interface systems. Multiple sensors and control devices are integrated inside the host satellite 5, which can monitor the status of the host satellite 5 and various parameters during the separation process in real time. In addition, the host satellite 5 was simplified in the experiment. The host satellite 5 is mainly built with an aluminum profile frame. The purpose is to reduce the weight of the test piece, and the center of mass of the host satellite 5 can be adjusted quickly and easily by adding or deleting mass blocks, so as to verify the separation under eccentric center of mass conditions. The host satellite bottom plate 20 is connected to the aluminum profile frame by bolts 21. The host satellite side panel 1-1-2 is connected to the aluminum profile frame by rotating bolts 1-1-4.

[0104] The hosted payload 1-2 is installed on the host satellite 1-1 and connected to the host satellite 1-1 through a mechanical interface system 1-3. Figure 2As shown. Multiple sensors and control devices are integrated inside the boarding load 1-2, which can monitor the status of the boarding load 1-2 and various parameters during the separation process in real time. In addition, the boarding load 1-2 is simplified in the test. The boarding load 1-2 is mainly built with an aluminum profile frame. The purpose is to reduce the weight of the test piece, and the center of mass of the boarding load 1-2 can be adjusted quickly and easily by adding or removing mass blocks, so as to verify the separation under eccentric center of mass conditions. The boarding load side plate and the boarding load bottom plate are connected to the aluminum profile frame by bolts, and the boarding load side plate is fixed to the antenna satellite substrate by bolts. During the separation process, the boarding load 1-2 is quickly and safely separated and locked through the mechanical interface system 1-3, ensuring a rapid response in an emergency and protecting the safety of the boarding load.

[0105] The composition of the No. 1 separation component and the No. 2 separation component is further described below:

[0106] Separable part No. 1 includes antenna satellite substrate 1-3-1, ejection plate interface 1-3-2 and threaded cover 1-3-3;

[0107] The lodged payload 1-2, the antenna satellite substrate 1-3-1 and the ejection board interface 1-3-2 are connected in sequence, and the threaded cover 1-3-3 is sleeved on the ejection board interface 1-3-2;

[0108] The second separation part includes a power spring device 1-3-4, a locking steel ball 1-3-5, a spring plate 1-3-6, a base 1-3-7, a release motor 1-3-8, a power adjustment motor 1-3-9, a laser sensor 1-3-10 and a fixing bolt 1-3-11;

[0109] The ejection plate interface 1-3-2 passes through the through holes on the host payload side plate 1-2-2 and the host satellite side plate 1-1-2 in turn, and is plugged into the interior of the base 1-3-7 through the compression power spring device 1-3-4; the No. 1 separation piece is restricted in the host payload frame 1-2-1 through the threaded cover 1-3-3;

[0110] The base 1-3-7 includes a hollow pillar and two circular support plates, the two circular support plates are divided into a first circular support plate and a second circular support plate; the first circular support plate is fixed to the bottom of the pillar, the second circular support plate is sleeved on the pillar, the pillar and the two circular support plates are an integral structure, and through holes are evenly provided on the side wall of the pillar located at the upper part of the second circular support plate; an annular groove 1-3-2-1 is provided on the outer circumferential surface of the ejection plate interface 1-3-2;

[0111] The spring plate 1-3-6 is sleeved on the pillar and supported on the second circular support plate. The tops of the two release motors 1-3-8 are in contact with the bottom of the spring plate 1-3-6. When the two release motors 3-8 are not started, the spring plate 1-3-6 squeezes and fixes the locking steel ball 1-3-5 between the through hole and the groove of the ejection plate interface 1-3-2, thereby locking the host satellite 1-1 and the hosted payload 1-2; when the two release motors 1-3-8 are started, the spring plate 1-3-6 is driven to move upward along the base 1-3-7, so that the annular groove on the inner circumference of the spring plate 1-3-6 is docked with the through hole, and the locking steel ball 1-3-5 enters from the annular groove on the ejection plate interface 1-3-2 between the through hole and the annular groove on the spring plate 1-3-6, thereby separating the host satellite 1-1 and the hosted payload 1-2.

[0112] Specifically, the mechanical interface system 1-3 is located between the host satellite and the hosted payload to achieve rapid separation and locking of the two. Fig.10 The components of the mechanical interface system 7 are as shown. Fig. 9 As shown. The design of the mechanical interface system takes into account a variety of separation mechanisms and safety locking devices, which can be adjusted according to different test requirements. The mechanical interface system uses springs and locking devices to achieve rapid separation of the hosted load while ensuring a smooth and safe separation process. The automatic adaptability of the mechanical interface system enables it to respond quickly to separation commands without human intervention, which is particularly suitable for load protection and task execution in emergency situations.

[0113] In the mechanical interface system 1-3 system, the hosted payload 1-2, the antenna satellite substrate and the ejection plate interface are fixedly connected in the initial state, and are separated as a whole when separated. The lower surface of the ejection plate interface contacts the upper surface of the host satellite top plate, and the threaded cover is fixedly connected to the lower surface of the host satellite top plate. In addition, to ensure that the hosted payload can work normally after separation, the ejection plate interface and the host satellite should also be equipped with electrical interfaces, information transmission interfaces and thermal protection subsystems.

[0114] In terms of the design and separation method of the mechanical interface system, the base 1-3-7 is fixed on the host satellite, and the release motor 1-3-8 is installed on both sides of the base 1-3-7 and connected to the base through fixing bolts. The release motor is a push rod motor, the top of which is directly connected to the spring plate to drive the spring plate to move along the base.

[0115] The power adjustment motor 1-3-9 is located in the middle of the base 1-3-7 and is also connected by fixing bolts. This is also a push rod motor, and the top is connected to the power spring device 1-3-4. The ejection speed is controlled by adjusting the compression of the device. The laser sensor 1-3-10 is installed near the power adjustment motor 1-3-9 to measure the displacement of the power spring device 1-3-4.

[0116] The bottom plate of the power spring device 1-3-4 is connected to the top of the power adjustment motor 1-3-9, and its top plate contacts the bottom of the ejection board interface. The power spring device includes four symmetrically distributed springs, which generate an upward force when compressed. When unlocked, this force pushes the top plate to separate the ejection board interface 1-3-2 and its connected host load 1-2 and the antenna satellite substrate 1-3-1.

[0117] The locking steel ball 1-3-5 is placed in the middle of the base 8, and there is a groove on the side of the ejection plate interface 1-3-2. In the initial state, the spring plate 1-3-6 squeezes the locking steel ball 1-3-5 into the groove to keep the structure stable. Once the release motor 1-3-8 is started, the spring plate 1-3-6 moves along the base, so that the locking steel ball 1-3-5 enters the cavity groove between the spring plate 1-3-6 and the base, thereby unlocking the ejection plate interface 1-3-2. This design not only ensures stability and safety during the separation process, but also realizes the automatic separation and locking functions.

[0118] The composition of the second separation member is further defined below:

[0119] Separator No. 2 also includes a power regulating motor 1-3-9;

[0120] A power regulating motor 1-3-9 is provided between the base 1-3-7 and the power spring device 1-3-4;

[0121] The power regulating motor 1-3-9 is used to drive the power spring device 1-3-4 to move upward and compress the power spring device 1-3-4.

[0122] The composition of the second separation member is further defined below:

[0123] Separation part No. 2 also includes laser sensor 1-3-10;

[0124] Laser sensor 1-3-10 is used to measure the relative displacement between the host satellite side plate 1-1-2 and the hosted payload side plate 1-2-2.

[0125] The following is the composition of the monitoring device:

[0126] The monitoring device includes an acceleration sensor, an IMU inertial navigation measurement instrument and a dynamic capture mechanism 3;

[0127] The host satellite 1-1 and the hosted payload 1-2 are both equipped with acceleration sensors and IMU inertial navigation measurement instruments for real-time monitoring of the acceleration and angular velocity of the host satellite 1-1 and the hosted payload 1-2;

[0128] The dynamic capture mechanism is used to capture the relative displacement and respective acceleration of the host satellite 1-1 and the hosted payload 1-2 in real time.

[0129] The following further defines the composition of the device:

[0130] The device also includes a host computer 4;

[0131] A processor is provided in the host computer 4, which is used to sequentially perform time alignment and space alignment on the relative displacement of the host satellite 1-1 and the hosted payload 1-2 captured by the dynamic capture mechanism in real time and the relative displacement output by the laser sensor 1-3-10 in real time, and perform data fusion on the two relative displacements after spatial alignment to obtain the final relative displacement; at the same time, the acceleration of the host satellite 1-1 output by the dynamic capture mechanism in real time and the acceleration of the host satellite 1-1 output by the acceleration sensor located in the host satellite 1-1 are sequentially performed time alignment and space alignment, and the two accelerations after spatial alignment are data fused to obtain the final host satellite 1-1 acceleration, and the acceleration of the hosted payload 1-2 output by the dynamic capture mechanism in real time and the acceleration of the hosted payload 1-2 output by the acceleration sensor located in the hosted payload 1-2 are sequentially performed time alignment and space alignment, and the two accelerations after spatial alignment are data fused to obtain the final hosted payload 1-2 acceleration;

[0132] It is also used to receive the angular velocity of the host satellite 1-1 and the hosted payload 1-2, and display the final hosted payload 1-2 acceleration, the final host satellite 1-1 acceleration and the final relative displacement of the host satellite 1-1 and the hosted payload 1-2.

[0133] Specifically, the entire separation process is the preparation stage before unlocking, and the speed is 0; the second process is from unlocking the unlocking device to complete separation, that is, the compression spring is compressed from being fully opened, and the satellite suddenly changes from a speed of 0; the third process is the disappearance of the elastic force after separation, and the satellite completes the complete separation state, at which time the attitude and speed tend to be stable. This embodiment mainly considers the attitude feedback information and acceleration / speed feedback information of the satellite during the second separation process. Since multiple sensors are used and the measurement of the target is usually performed asynchronously, if asynchronous data is directly used for fusion, not only the measurement accuracy of the system cannot be improved, but the tracking effect of the separated satellite will be reduced. Therefore, the data from different measurement sensors such as the dynamic capture system, gyroscope, accelerometer and laser sensor are spatiotemporally aligned, that is, the observation data of different sensors are transformed into the same spatial coordinate system and aligned on the time axis to make them have spatiotemporal unity.

[0134] The process of temporal registration is described below:

[0135] The acceleration of the host satellite 1-1 captured by the dynamic capture mechanism in real time is fitted into a plurality of first host satellite acceleration curves, and the acceleration of the host satellite 1-1 output by the acceleration sensor in real time is fitted into a plurality of second host satellite acceleration curves; the acceleration of the host payload 1-2 captured by the dynamic capture mechanism in real time is fitted into a plurality of first host payload acceleration curves, and the acceleration of the host payload 1-2 output by the acceleration sensor in real time is fitted into a plurality of second host payload acceleration curves; the relative displacement captured by the dynamic capture mechanism in real time is fitted into a plurality of first displacement curves, and the relative displacement output by the laser sensor 1-3-10 in real time is fitted into a plurality of second displacement curves, and each fitted displacement curve and acceleration curve is expressed as:

[0136]

[0137] In the formula, a0, a1, a2 are the polynomial coefficients of the fitting curve, x i is the displacement information at the i-th moment measured by the laser sensor or dynamic capture mechanism, t i is the measurement time at the i-th moment, j represents the polynomial order of the fitting curve;

[0138] The error between each first host satellite acceleration curve and the acceleration acquisition value is calculated as the first host satellite acceleration error, the error between each second host satellite acceleration curve and the acceleration acquisition value is calculated as the second host satellite acceleration error, the error between each first host payload acceleration curve and the acceleration acquisition value is calculated as the first host payload acceleration error, the error between each second host payload acceleration curve and the acceleration acquisition value is calculated as the second host payload acceleration error, the error between each first displacement curve and the displacement acquisition value is calculated as the first displacement error, and the error between each second displacement curve and the displacement acquisition value is calculated as the second displacement error. The error δ i It is expressed as:

[0139]

[0140] In the formula, P(t i ) is the acceleration information at the i-th moment in the first host satellite acceleration curve, the second host satellite acceleration curve, the first hosted payload acceleration curve, or the second hosted payload acceleration curve, or the displacement information at the i-th moment in the first displacement curve, or the displacement information at the i-th moment in the second displacement curve;

[0141] A first host satellite acceleration curve corresponding to a minimum value is selected from multiple first host satellite acceleration errors as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value is selected from multiple second host satellite acceleration errors as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value is selected from multiple first host payload acceleration errors as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value is selected from multiple second host payload acceleration errors as the second host payload acceleration curve after screening, a first displacement curve corresponding to a minimum value is selected from multiple first displacement errors as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value is selected from multiple second displacement errors as the second displacement curve after screening.

[0142] Specifically, because the frequencies of the two accelerations or two displacements collected by two different elements are different, time alignment is required to fit the acceleration and displacement into a curve form. In this way, at the same moment, the point value taken on each curve is the data at the same time, and the time is unified.

[0143] The process of obtaining formula 1 is:

[0144] definition Use formula 2 to calculate the error and convert formula 2 into a polynomial function form, that is: Transforming formula 6, we get: Writing Equation 7 in variance form: Taking the partial derivative of Equation 8 and taking it to zero, we get: According to Formula 9, the set of equations that the polynomial coefficients should satisfy can be obtained as shown in Formula 1.

[0145] The process of spatial registration is described below:

[0146] The state coordinate systems of the first screened host satellite acceleration curve, the second screened host satellite acceleration curve, the first screened hosted payload acceleration curve, the second screened hosted payload acceleration curve, the first screened displacement curve and the second screened displacement curve are all converted into ground coordinate systems.

[0147] Specifically, spatial registration is to register the coordinate systems of each sensor so that they work in a unified measurement coordinate system. According to the installation distance of the sensor, the spatial registration algorithm can be divided into two cases: system-level registration and platform-level registration. In order to fuse the reference coordinate systems of each sensor, it is necessary to select a common coordinate system as the fusion standard. In this embodiment, considering the position of each sensor and the position of the sensor itself, first, the ground system is used as the fusion reference coordinate system, and then, according to the measurement value of the sensor and the geographic coordinates of the sensor itself, the state coordinates of the target are converted into the ground system. In this way, the spatial registration is completed, and the root mean square error (RMSE) is used as a quantitative indicator to evaluate the registration effect.

[0148]

[0149] In the formula, is the position information of the target measurement value after fusion at time k, X k is the true value of the target position at time k, and N is the number of sampling times.

[0150] The process of data fusion is described below:

[0151] According to the data fusion formula of formula 3, the first host satellite acceleration curve and the second host satellite acceleration curve after screening in the ground coordinate system are fused to obtain the fused host satellite acceleration curve, the first host load acceleration curve and the second host load acceleration curve after screening in the ground coordinate system are fused to obtain the fused host load acceleration curve, and the first displacement curve and the second displacement curve after screening in the ground coordinate system are fused to obtain the fused displacement curve:

[0152]

[0153] In the formula, is the fused acceleration curve or fused displacement curve, W i is the weight of the i-th acceleration curve or the i-th displacement curve, i = 1, 2, X i is the displacement information or acceleration measured by the i-th sensor;

[0154] Among them, W i According to formula 4, we can obtain:

[0155]

[0156] In the formula, σ 2 is the total variance, E represents the expectation, and X represents the true value of the target information. is the error variance measured by the i-th sensor, is the minimum total variance of all sensor measurements.

[0157] Specifically, after synchronizing the data collected by each sensor to the same spatiotemporal scale, an adaptive weighted fusion algorithm is used to assign weights with the minimum root mean square error (LMS) as the optimal criterion, which can better perform data fusion based on the measurement characteristics of different sensors.

[0158] Assume that at a certain moment, there are n sensors measuring the same target at the same location. The true value of the target location information is X, and the measurement values ​​of each sensor are X1, X2, …X n And they are independent of each other, and the weight of each sensor is W1, W2, …W n The fusion result is As in formula 3;

[0159] The error variances of the measured values ​​of each sensor are Calculate the total variance σ 2 like:

[0160]

[0161] In the formula, E is the expectation.

[0162] Because X1,X2,…X n If they are independent and X is an unbiased estimate, then:

[0163] E(XX i )(XX j )=0,

[0164] Then we can deduce:

[0165]

[0166] The weight corresponding to the minimum total variance is calculated as follows:

[0167] E(XX i )(XX j )=0,

[0168] At this time, the minimum value of the total variance is as follows:

[0169]

[0170] In this way, high-precision position and velocity information can be obtained during the satellite separation process.

[0171] like Fig.13 As shown, the spacecraft consists of two small satellites. Assume that the mass of the main satellite is M1 and the inertia is J1, and the mass of the auxiliary satellite is M2 and the inertia is J2. The following information can be calculated based on the fusion data.

[0172] is the unit vector of the spring force direction, and are the vector distances from the origin of the inertial coordinate system to the center of mass of the two satellites, and is the radius vector from the origin of the inertial system to the spring action point, l1 and l2 are the perpendicular distances from the center of mass to the direction of the spring force.

[0173] The equations of motion for the two spacecraft are as follows:

[0174]

[0175] Where F is the spring force, and its magnitude is |F|=k(l0-l t -r2·r F +r1·r F +b1+b2), the unit vector of the force direction is r F =[r Fx r Fy r Fz ], then F=|F|r F l0 is the original length of the spring, l t The shortest length of the spring when compressed.

[0176] b1 and b2 are the distances from the connection point to the centroid projection point. The specific expression is:

[0177]

[0178] The motion process of the two satellites can be divided into two stages depending on whether the spring is in a compressed state:

[0179] (1) When r2·r F -r1·r F -b1-b2<l0-l t When , the spring is in a compressed state, and there is a spring force F between the two satellites;

[0180] (2) When r2·r F -r1·r F -b1-b2≥l0-l t When , the spring is in its original length state and there is no interaction force between the two satellites.

[0181] Working principle:

[0182] The overall flow chart of the experiment is as follows: Fig.14As shown. In order to elaborate on the purpose and advantages of the present invention, a host satellite and hosted payload separation and locking attitude monitoring test device is described in detail in combination with the accompanying drawings and actual operation examples. The design of the test device mainly studies the influence of the stiffness of the spring in the mechanical interface on the separation speed and separation direction of the host satellite and the hosted payload, the influence of the separation speed and separation direction on the attitude of the host satellite and the hosted payload, and the influence of the machining error based on the mechanical interface on the attitude of the host satellite and the hosted payload.

[0183] Step 1: Equipment installation and initialization: Install the host satellite 1-1 and the hosted payload 1-2 on the air floating platform 4 and connect them through the mechanical interface system 1-3. The air support system 2 is adjusted to an appropriate height to ensure that the host satellite and the hosted payload can simulate the floating state in the microgravity environment during the test.

[0184] Step 2: Multiple cameras of the dynamic capture system 3 are arranged around the air floating platform to capture the dynamics of the host satellite and the hosted payload during the separation process in all directions.

[0185] Step 3: IMU inertial navigation measurement instruments and acceleration sensors are installed on the host satellite and hosted payload to monitor their attitude and acceleration changes in real time.

[0186] Step 4: The laser sensor is positioned next to the power spring assembly 1-3-4 to monitor the spring compression and release process.

[0187] Step 5: Control system activation: The host computer starts, integrates all sensor inputs, sets data acquisition frequency and recording parameters, such as separation speed, direction, attitude change, etc.

[0188] Step 6: Separation direction adjustment: Adjust the push rod motors 1-4 to change the separation direction of the host satellite side panels, that is, the corresponding formula The force direction unit vector r F =r Fx r Fy r Fz ].

[0189] Step 7: Initial preparation: Determine the pre-compression amount and spring stiffness of the power spring device 1-3-4 according to the mass, size and shape of the boarding load 1-2. Use Hooke's law to calculate the spring force F of the spring in the pre-compression state to ensure that the spring force is sufficient to successfully separate the boarding load 1-2.

[0190] Step 8: Initial Connection: Fig. 9As shown. The host payload 1-2 is fixedly connected to the ejection plate interface 10 through the antenna satellite substrate 9, and the whole structure is installed on the host satellite top plate. At this time, ensure that the locking steel ball 1-3-5 is squeezed by the spring plate 1-3-6 and fixed in the groove of the ejection plate interface 1-3-2 to ensure the stability of the system.

[0191] Step 9: Power adjustment: Fig. 9 Start the power regulating motor 1-3-9 to move the bottom plate of the power spring device 1-3-4 upward and compress the spring to a predetermined position, that is, the corresponding formula The shortest length l of the spring compression t The displacement of the power spring device is monitored by laser sensor 1-3-10 to ensure that the compression amount and spring force reach the predetermined calculated value, that is, the corresponding formula Medium spring force F.

[0192] Step 10: Pre-separation inspection: Check all components to ensure that the host load 1-2 can work normally after separation. System testing and data feedback confirm that all preparations have been completed and the separation system is in standby mode.

[0193] Step 11: Start the release motor: Fig. 9 Operate the release motor 1-3-8 to push the spring plate 1-3-6 to move to a specific position along the base 1-3-7, so that the locking steel ball 1-3-5 falls off, and the ejection plate interface 1-3-2 is unlocked.

[0194] Step 12: Separation process: Fig. 9 As shown. The power spring device 1-3-4 releases the spring force, pushing the ejection plate interface 1-3-2 upward and apart. The relationship between the compression amount of the power spring and the propulsion speed, the stress change of the locking spring and the maximum load capacity of the mechanism are analyzed to ensure the reliability of the propulsion and launch system.

[0195] Step 13: Data collection: During the experiment, the dynamic capture system 3, gyroscope, acceleration sensor and laser sensor synchronously record all dynamic parameters during the separation process. These data include the speed, acceleration, rotation angle and precise time point of the separation moment of the host satellite 1-1 and the hosted payload 1-2.

[0196] Step 14: Use the dedicated software in the host computer to perform preliminary processing on the collected data, such as filtering, denoising and data synchronization. This step ensures the accuracy of the analysis and the reliability of the data.

[0197] Step 15: Multi-condition analysis: 1) Change the rotation angle of the host satellite side panel to change the separation direction of the hosted load; 2) Adjust the spring stiffness k and pre-compression l t; 3) Explore the impact of machining errors based on mechanical interfaces on the attitude of the host satellite and hosted payload.

[0198] Step 16: Design multiple working conditions: According to the actual space environment and mission requirements, design working conditions including different spring stiffness, different pre-compression, and different center of mass offset. Each working condition attempts to simulate possible extreme or special situations.

[0199] Step 17: Working condition test: Conduct actual tests on each designed working condition and collect data on separation speed, direction, posture change, etc. In addition, introduce intentional processing errors, such as uneven spring stiffness and change in contact surface roughness, to evaluate the specific effects of these factors on separation performance.

[0200] Step 18: Multi-source information fusion and high-precision information solution: Although the accelerometer has high accuracy, there is a defect of error accumulation due to integration in the process of calculating velocity and displacement, and the dynamic capture system and laser sensor have low accuracy. Therefore, the multi-source information fusion technology is used to ensure high-precision solution of satellite position and velocity information and reduce measurement errors.

[0201] Step 19: Performance evaluation: Analyze the deviation of separation speed and direction from the preset target, and evaluate the spring stiffness k and pre-compression l t The influence on the separation effect. By establishing a mathematical model, that is, the formula Compare experimental data with theoretical predictions and identify possible causes of deviations.

[0202] Step 20: Error source identification: Identify the main error sources that affect separation performance by comparing the data differences under different operating conditions and ideal conditions. Use statistical analysis methods, such as analysis of variance (ANOVA), to determine the significance of the error sources.

[0203] Step 21: Systematic Error Correction: Correct the identified significant error sources. Depending on the nature of the error, adjust the design of the mechanical interface or propose new manufacturing and assembly standards. This step may include redesigning some components or adjusting the assembly process.

[0204] Step 22: Parameter optimization: Use an optimization algorithm, such as a genetic algorithm or gradient descent, to iteratively search for the optimal solution to ensure that the separation action achieves the highest efficiency and accuracy.

[0205] Step 23: Optimization verification test: The adjusted parameters are verified to be effective through another experiment. Repeat the experiment, collect new data, and perform data processing and performance evaluation again to ensure that the adjustments made have effectively improved system performance.

[0206] Step 24: Verification and feedback: After implementing the corrective measures, re-test to verify the effect. Based on the new test results, further data analysis is performed to evaluate the error.

[0207] Step 25: Through this series of tests and analyses, the mechanical interface system of the present invention can effectively achieve rapid and safe separation of the host satellite and the hosted payload. The test results will be used to further optimize the mechanical interface design and provide technical support for future space missions. Future research will focus on further reducing processing errors, improving separation efficiency, and exploring new materials and technologies to enhance the reliability and performance of the entire system.

[0208] Through this series of tests and analyses, the mechanical interface system of the present invention can effectively achieve the rapid and safe separation of the host satellite and the hosted payload. The purpose of this test is to consider the spring stiffness k, pre-compression amount l of the host satellite and the hosted payload under different working conditions. t and the separation direction r F , to meet the needs of space missions and avoid near-field interference effects, that is, the risk of close collision between the host satellite and the hosted payload. The test results will be used to further optimize the mechanical interface design and provide technical support for future space missions. Future research will focus on further reducing processing errors, improving separation efficiency, and exploring new materials and technologies to improve the reliability and performance of the entire system.

[0209] Although the present invention is described herein with reference to specific embodiments, it should be understood that these embodiments are merely examples of the principles and applications of the present invention. It should therefore be understood that many modifications may be made to the exemplary embodiments and that other arrangements may be devised without departing from the spirit and scope of the present invention as defined by the appended claims. It should be understood that the various dependent claims and features described herein may be combined in a manner different from that described in the original claims. It should also be understood that features described in conjunction with individual embodiments may be used in other described embodiments.

Claims

1. A test device for monitoring the separation and locking attitude of a host satellite and a hosted payload, characterized in that: The test device comprises a simulation device (1), an air foot support system (2) and a monitoring device; An air foot support system (2) is installed at the bottom of the simulation device (1). By introducing gas into the top end of an air pipe (2-1) on the air foot support system (2), the gas is ejected from the bottom end of the air pipe (2-1), so that the simulation device (1) is in a free floating state, simulating a floating state in space. The simulation device (1) includes a host satellite (1-1), a hosted payload (1-2), a mechanical interface system (1-3) and a push rod motor (1-4); The host satellite (1-1) and the hosted payload (1-2) are connected via a mechanical interface system (1-3), and the mechanical interface system (1-3) is used to lock or separate, thereby driving the connected host satellite (1-1) and the hosted payload (1-2) to lock or separate; A push rod motor (1-4) is used to adjust the tilt angle of the mechanical interface system (1-3), thereby adjusting the separation direction of the host satellite (1-1) and the hosted payload (1-2); A monitoring device is used to monitor dynamic parameters during the separation process of a host satellite (1-1) and a hosted payload (1-2).

2. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 1, characterized in that: The host satellite (1-1) comprises a host satellite frame (1-1-1), a host satellite side plate (1-1-2) and a host satellite bottom plate (1-1-3); The host satellite frame (1-1-1) is a cube frame, the host satellite side plate (1-1-2) is rotatably connected to the side frame of the main satellite frame (1-1-1) by rotating bolts (1-1-4), and the host satellite bottom plate (1-1-3) is fixed to the bottom surface frame of the host satellite frame (1-1-1) by bolts; The boarding load (1-2) includes a boarding load frame (1-2-1), a boarding load side plate (1-2-2) and a boarding load bottom plate (1-2-3); The host load frame (1-2-1) is a cubic frame, the host load side plate (1-2-2) is rotatably connected to the side frame of the host load frame (1-2-1) by rotating bolts (1-1-4), and the host load bottom plate (1-2-3) is fixed to the bottom surface frame of the host load frame (1-2-1) by bolts; and the host load side plate (1-2-2) and the host satellite side plate (1-1-2) are arranged opposite to each other; The mechanical interface system (1-3) includes a first separation component and a second separation component; The first separation component is arranged in the host payload frame (1-2-1), and the second separation component is arranged in the host satellite frame (1-1-1); Through holes are provided at the same positions on the host satellite side plate (1-1-2) and the host load side plate (1-2-2), and the host satellite side plate (1-1-2) and the host load side plate (1-2-2) are arranged in parallel and opposite to each other; The No. 1 separation piece passes through the through holes on the host payload side plate (1-2-2) and the host satellite side plate (1-1-2) in sequence and is plugged into the No. 2 separation piece, so that the No. 1 separation piece and the No. 2 separation piece can be locked or separated.

3. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 1, characterized in that: Separable part No. 1 includes an antenna satellite substrate (1-3-1), an ejection plate interface (1-3-2) and a threaded cover (1-3-3); The lodging payload (1-2), the antenna satellite substrate (1-3-1) and the ejection plate interface (1-3-2) are connected in sequence, and the threaded cover (1-3-3) is sleeved on the ejection plate interface (1-3-2); The second separation member includes a power spring device (1-3-4), a locking steel ball (1-3-5), a spring plate (1-3-6), a base (1-3-7) and a release motor (1-3-8); The ejection plate interface (1-3-2) passes through the through holes on the host load side plate (1-2-2) and the host satellite side plate (1-1-2) in sequence, and is plugged into the interior of the base (1-3-7) through the compression power spring device (1-3-4); the No. 1 separation component is restricted in the host load frame (1-2-1) through the threaded cover (1-3-3); The base (1-3-7) comprises a hollow pillar and two annular support plates, the one fixed at the bottom of the hollow pillar is named as the first annular support plate, the one sleeved and fixed on the hollow pillar is named as the second annular support plate, through holes are evenly arranged on the side wall of the hollow pillar at the upper part of the second annular support plate; an annular groove (1-3-2-1) is arranged on the outer circumferential surface of the ejection plate interface (1-3-2); The spring plate (1-3-6) is sleeved on the hollow pillar and supported on the second annular support plate. The tops of the two release motors (1-3-8) are in contact with the bottom surface of the spring plate (1-3-6). When the two release motors (3-8) are not started, the spring plate (1-3-6) squeezes and fixes the locking steel ball (1-3-5) between the through hole and the groove of the ejection plate interface (1-3-2), thereby achieving the locking of the host satellite (1-1) and the hosted payload (1-2). When both release motors (1-3-8) are started, the spring plate (1-3-6) is driven to move upward along the base (1-3-7), so that the annular groove on the inner circumference of the spring plate (1-3-6) is docked with the through hole, and the locking steel ball (1-3-5) enters between the through hole and the annular groove on the spring plate (1-3-6) from the annular groove on the ejection plate interface (1-3-2), thereby realizing the separation of the host satellite (1-1) and the hosted payload (1-2).

4. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 3, characterized in that: The second separation component also includes a power regulating motor (1-3-9); A power regulating motor (1-3-9) is arranged between the base (1-3-7) and the power spring device (1-3-4); The power regulating motor (1-3-9) is used to drive the power spring device (1-3-4) to move upward and compress the power spring device (1-3-4).

5. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 4, characterized in that: Separator No. 2 also includes a laser sensor (1-3-10); The laser sensor (1-3-10) is used to measure the relative displacement between the host satellite side plate (1-1-2) and the hosted payload side plate (1-2-2).

6. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 1, characterized in that: The monitoring device includes an acceleration sensor, an IMU inertial navigation measurement instrument and a dynamic capture mechanism (3); The host satellite (1-1) and the hosted payload (1-2) are both provided with acceleration sensors and IMU inertial navigation measurement instruments for real-time monitoring of the acceleration and angular velocity of the host satellite (1-1) and the hosted payload (1-2); The dynamic capture mechanism is used to capture the relative displacement and respective acceleration of the host satellite (1-1) and the hosted payload (1-2) in real time.

7. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 6, characterized in that: The test device also includes a host computer (4); A processor is provided in the upper computer (4) for sequentially performing time registration and space registration on the relative displacement of the host satellite (1-1) and the hosted payload (1-2) captured in real time by the dynamic capture mechanism and the relative displacement output in real time by the laser sensor (1-3-10), performing data fusion on the two relative displacements after spatial registration to obtain a final relative displacement; and simultaneously performing time registration and space registration on the acceleration of the host satellite (1-1) output in real time by the dynamic capture mechanism and the acceleration of the host satellite (1-1) output by the acceleration sensor located in the host satellite (1-1). -1) are sequentially time-aligned and spatially-aligned, the two accelerations after spatial alignment are data-fused to obtain the final acceleration of the host satellite (1-1), and the acceleration of the host payload (1-2) output in real time by the dynamic capture mechanism and the acceleration of the host payload (1-2) output by the acceleration sensor located in the host payload (1-2) are sequentially time-aligned and spatially-aligned, the two accelerations after spatial alignment are data-fused to obtain the final acceleration of the host payload (1-2); It is also used to receive the angular velocity of the host satellite (1-1) and the hosted payload (1-2), and display, at the same time, the final acceleration of the hosted payload (1-2), the final acceleration of the host satellite (1-1) and the final relative displacement of the host satellite (1-1) and the hosted payload (1-2).

8. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 7, characterized in that: The process of temporal registration is: The acceleration of the host satellite (1-1) captured in real time by the dynamic capture mechanism is fitted into a plurality of first host satellite acceleration curves, and the acceleration of the host satellite (1-1) output in real time by the acceleration sensor is fitted into a plurality of second host satellite acceleration curves; the acceleration of the host load (1-2) captured in real time by the dynamic capture mechanism is fitted into a plurality of first host load acceleration curves, and the acceleration of the host load (1-2) output in real time by the acceleration sensor is fitted into a plurality of second host load acceleration curves; the relative displacement captured in real time by the dynamic capture mechanism is fitted into a plurality of first displacement curves, and the relative displacement output in real time by the laser sensor (1-3-10) is fitted into a plurality of second displacement curves, and each fitted displacement curve and acceleration curve is expressed as: In the formula, a0, a1, a2 are the polynomial coefficients of the fitting curve, x i is the displacement information at the i-th moment measured by the laser sensor or dynamic capture mechanism, t i is the measurement time at the i-th moment, j represents the polynomial order of the fitting curve; The error between each first host satellite acceleration curve and the acceleration acquisition value is calculated as the first host satellite acceleration error, the error between each second host satellite acceleration curve and the acceleration acquisition value is calculated as the second host satellite acceleration error, the error between each first host payload acceleration curve and the acceleration acquisition value is calculated as the first host payload acceleration error, the error between each second host payload acceleration curve and the acceleration acquisition value is calculated as the second host payload acceleration error, the error between each first displacement curve and the displacement acquisition value is calculated as the first displacement error, and the error between each second displacement curve and the displacement acquisition value is calculated as the second displacement error. The error δ i It is expressed as: In the formula, P(t i ) is the acceleration information at the i-th moment in the first host satellite acceleration curve, the second host satellite acceleration curve, the first hosted payload acceleration curve, or the second hosted payload acceleration curve, or the displacement information at the i-th moment in the first displacement curve, or the displacement information at the i-th moment in the second displacement curve; A first host satellite acceleration curve corresponding to a minimum value is selected from multiple first host satellite acceleration errors as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value is selected from multiple second host satellite acceleration errors as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value is selected from multiple first host payload acceleration errors as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value is selected from multiple second host payload acceleration errors as the second host payload acceleration curve after screening, a first displacement curve corresponding to a minimum value is selected from multiple first displacement errors as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value is selected from multiple second displacement errors as the second displacement curve after screening.

9. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 8, characterized in that: The process of spatial registration is: The state coordinate systems of the first screened host satellite acceleration curve, the second screened host satellite acceleration curve, the first screened hosted payload acceleration curve, the second screened hosted payload acceleration curve, the first screened displacement curve and the second screened displacement curve are all converted into ground coordinate systems.

10. The test device for monitoring the separation and locking attitude of a host satellite and a hosted payload according to claim 9, characterized in that: The process of data fusion is: According to the data fusion formula of formula 3, the first host satellite acceleration curve and the second host satellite acceleration curve after screening in the ground coordinate system are fused to obtain the final host satellite acceleration curve, the first host load acceleration curve and the second host load acceleration curve after screening in the ground coordinate system are fused to obtain the final host load acceleration curve, and the first displacement curve and the second displacement curve after screening in the ground coordinate system are fused to obtain the final displacement curve: In the formula, is the fused acceleration curve or fused displacement curve, W i is the weight of the i-th acceleration curve or the i-th displacement curve, i = 1, 2, X i is the displacement information or acceleration measured by the i-th sensor; Among them, W i According to formula 4, we can obtain: In the formula, σ 2 is the total variance, E represents the expectation, X represents the true value of the target information, σ i 2 is the error variance measured by the i-th sensor, is the minimum total variance of all sensor measurements.

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