Test device for monitoring the separation and locking attitude of the host satellite and hosted payload

Abstract

A test device for monitoring the separation and locking attitude of a host satellite and a hosted payload belongs to the field of mechanical design and testing. Existing tests for separation of a hosted payload and a host satellite have the disadvantages of insufficient simulation in a microgravity environment and inaccurate satellite attitude data. An air-support system enables the simulation device to be in a free-floating state; the simulation device simulates the separation of the host satellite and the hosted payload; the monitoring device monitors the dynamic parameters during the separation process of the host satellite and the hosted payload; the host satellite and the hosted payload are connected through a mechanical interface system, which is used to lock or separate, driving the connected host satellite and hosted payload to lock or separate; the push rod motor is used to adjust the tilt angle of the mechanical interface system, thereby adjusting the separation angle of the host satellite and the hosted payload. The present invention is used to simulate the separation of a host satellite and a hosted payload, and obtains attitude data during the separation process for subsequent analysis.

Description

Technical Field

[0001] The present 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, with the increasing complexity of satellite missions, the management and separation of host satellites and hosted payloads has become a hot topic in research and spaceflight practice. Hosted payloads are small devices or modules attached to host satellites, typically used to perform specific tasks such as scientific experiments, data collection, or communications assistance. Ensuring that these hosted payloads can be quickly and safely separated from their host satellites when needed is a key factor in satellite mission success.

[0003] Traditional methods for 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 place a heavy burden on the overall design of the satellite, increasing launch costs. Although explosive bolts can achieve rapid separation, their one-time use nature 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 failure or collision threats) and cannot effectively protect hosted payloads.

[0004] With the increasing demand for space missions and the advancement of technology, there is an urgent need for a system that can quickly and safely separate hosted payloads from their host satellites without relying on human intervention. However, existing research and engineering practices have limited simulation capabilities in microgravity environments and inaccurate satellite attitude data. This leads to numerous uncertainties and risks in actual separation operations.

[0005] The main reason for inaccurate satellite attitude data is that existing tests of separating hosted payloads from host satellites rely on traditional monitoring technologies. These technologies rely primarily on a single sensor or a single type of acquisition device (such as an image sensor or inertial sensor) to obtain satellite attitude data. These methods have several significant shortcomings. First, a single sensor struggles to obtain comprehensive information on attitude changes during the complex separation process, resulting in inaccurate data. Furthermore, the intense vibration and impact generated during separation can easily affect the accuracy and reliability of a single sensor, potentially leading to deviations in attitude data. While existing image acquisition technologies can provide visual data, data acquisition and processing are limited when attitude changes are dramatic or when lighting conditions vary, making it difficult to accurately reflect the satellite's motion in real time. Furthermore, the use of a single inertial sensor can't simultaneously account for attitude changes in multiple directions, leading to increased errors in attitude calculation. Consequently, existing methods fail to accurately capture satellite attitude. Summary of the Invention

[0006] The purpose of the present invention is to solve the problems that the existing separation test of the host payload and the host satellite has the shortcomings of insufficient simulation in the microgravity environment and inaccurate satellite attitude data, which leads to the separation operation in actual missions facing many uncertainties and risks. A test device for monitoring the separation and locking attitude of the host satellite and the 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 tube on the air foot support system, the gas is ejected from the bottom of the air tube, making the simulation device float freely, 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, which is used to lock or separate the connected host satellite and hosted payload.

[0011] Pusher 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 the 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 payload frame is a cube frame, the host payload side panels are rotatably connected to the side frames of the host payload frame by rotating bolts, and the host payload bottom plate is fixed to the bottom frame of the host payload frame by bolts; and the host payload side panels are arranged opposite to the host satellite side panels;

[0017] The mechanical interface system includes a No. 1 separation component and a No. 2 separation component;

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

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

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

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

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

[0023] Separator 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 component is confined within the host payload frame through a threaded cover;

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

[0026] The spring plate is sleeved on the hollow pillar and supported on the second circular 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, thereby locking 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 docks 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, thereby separating the host satellite and the hosted payload.

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

[0028] A power regulating motor is provided 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 separating element further comprises a laser sensor;

[0031] Laser sensor, 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 includes a host computer;

[0036] A processor is provided in the host computer, which is used to perform time alignment and spatial alignment on the relative displacement of the host satellite and the hosted payload captured in real time by the dynamic capture mechanism and the relative displacement output in real time by the laser sensor, 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 in real time by the dynamic capture mechanism and the acceleration of the host satellite output by the acceleration sensor located in the host satellite are respectively performed time alignment and spatial 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 in real time by the dynamic capture mechanism and the acceleration of the hosted payload output by the acceleration sensor located in the hosted payload are respectively performed time alignment and spatial 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 in real time by the dynamic capture mechanism is fitted into multiple first-type host satellite acceleration curves, and the acceleration of the host satellite output by the acceleration sensor in real time is fitted into multiple second-type host satellite acceleration curves; the acceleration of the hosted payload captured in real time by the dynamic capture mechanism is fitted into multiple first-type hosted payload acceleration curves, and the acceleration of the hosted payload output by the acceleration sensor in real time is fitted into multiple second-type hosted payload acceleration curves; the relative displacement captured in real time by the dynamic capture mechanism is fitted into multiple first-type displacement curves, and the relative displacement output by the laser sensor in real time is fitted into multiple second-type displacement curves. Each fitted displacement curve and acceleration curve is expressed as:

[0040]

[0041] In the formula, a0, a1, and a2 are the polynomial coefficients of the fitting curve, and 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 of the i-th moment, j represents the polynomial order of the fitting curve;

[0042] Calculate the error between each first host satellite acceleration curve and the acceleration acquisition value as the first host satellite acceleration error, calculate the error between each second host satellite acceleration curve and the acceleration acquisition value as the second host satellite acceleration error, calculate the error between each first hosted payload acceleration curve and the acceleration acquisition value as the first hosted payload acceleration error, calculate the error between each second hosted payload acceleration curve and the acceleration acquisition value as the second hosted payload acceleration error, calculate the error between each first displacement curve and the displacement acquisition value as the first displacement error, and calculate the error between each second displacement curve and the displacement acquisition value as the second displacement error. The error δ i Expressed as:

[0043]

[0044] Where, 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 from multiple first host satellite acceleration errors is selected as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value from multiple second host satellite acceleration errors is selected as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value from multiple first host payload acceleration errors is selected as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value from multiple second host payload acceleration errors is selected as the second host payload acceleration curve after screening; a first displacement curve corresponding to a minimum value from multiple first displacement errors is selected as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value from multiple second displacement errors is selected 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 the ground coordinate system.

[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 payload acceleration curve and the second host payload acceleration curve after screening in the ground coordinate system are fused to obtain the final host payload acceleration curve. At the same time, 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] Where, 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] Where, 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, 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 the 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 collection 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 high 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 flotation 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 in acceleration and angular velocity during the separation process, providing high-precision data support, and improving the accuracy of experimental analysis.

[0062] The present invention uses 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 and facilitating 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 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 support in a microgravity environment by ejecting air, so that the host satellite and the hosted payload can maintain a stable suspended 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 record the attitude change process through images to provide intuitive test data.

[0066] The present invention mainly studies the separation test of the host satellite and the hosted payload based on the mechanical interface. In orbit, the host satellite and the hosted payload are both in a floating state, so when conducting the test, it is necessary to use air-foot support to create the 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 the 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 a near-field collision. In addition, since there may be processing errors in the design of the mechanical interface, it is easy for the mechanical interface itself to have a separation direction that is 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 missions. 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 Schematic diagram of the experimental device for monitoring the separation and locking attitude of the host satellite and hosted payload;

[0070] Figure 2 This is a schematic diagram of the structure of the simulation device that is in a free-floating state 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 main 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] Figure 9 This is the exploded diagram of the mechanical interface system;

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

[0079] Figure 11 for Figure 10 Middle AA section;

[0080] Figure 12 Schematic diagram of the separation process of the mechanical interface system;

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

[0082] Figure 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 clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. 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 making any creative efforts shall fall 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 with reference to the accompanying drawings and specific embodiments, but they are not intended to limit the present invention.

[0086] Example:

[0087] Combine Figures 1 to 3 This embodiment describes 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, making the simulation device 1 float freely, 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. The mechanical interface system 1-3 is used to lock or separate the connected host satellite 1-1 and the hosted payload 1-2.

[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 between the host satellite 1-1 and the hosted payload 1-2.

[0092] Specifically, Figure 1 Reference numeral 3 denotes an air flotation platform. Air flotation platform 4 is located at the bottom of air foot support system 2. Air foot support system 2 is primarily used to simulate a microgravity environment, allowing the host satellite 5 and hosted payload 6 to float freely during testing. Once floating, air foot support system 2 detaches from air flotation platform 4. The design of air foot support system 2 takes into account the uniformity and stability of the airflow to ensure that the host satellite and hosted payload maintain a stable suspension during testing, thereby simulating actual operating conditions in a space environment. Air foot support system 4 is a key component of the air flotation platform. Air foot support system 4 sprays air between the air foot and the ground, creating an air cushion. This reduces contact friction between objects and the ground, allowing the objects to float on the air cushion, similar to the conditions in a microgravity environment. By adjusting the air pressure and flow rate of the air cushion, the weight of the object is balanced by the buoyancy force, achieving the effect of unloading gravity. This ensures that the net force on the object on the air foot is close to zero, simulating a microgravity environment. Nitrogen is introduced into air pipe 2-1.

[0093] There are two push rod motors 1-4, which are 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 in the figure, the pusher motor precisely controls the ejection direction and force, ensuring that the hosted payload follows the predetermined trajectory during separation, avoiding collision with the host satellite. The pusher motor design incorporates multiple operating modes and control algorithms, allowing for flexible adjustment of ejection parameters based on test requirements to optimize the separation process.

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

[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 payload side plate 1-2-2 is rotatably connected to the side frame of the host payload frame 1-2-1 by rotating bolts 1-1-4, and the host payload bottom plate 1-2-3 is fixed to the bottom frame of the host payload frame 1-2-1 by bolts; and the host payload 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 first separation component and a second separation component;

[0100] Separator No. 1 is set in the host payload frame 1-2-1, and separator No. 2 is set 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 payload side plate 1-2-2. The host satellite side plate 1-1-2 and the host payload side plate 1-2-2 are parallel and oppositely arranged.

[0102] 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. The No. 1 separation piece and the No. 2 separation piece can be locked or separated.

[0103] Specifically, the host satellite 1-1 serves as the main carrier for installing 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 during the test. 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 removing mass blocks, so as to verify the separation situation 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 host payload 1-2, which can monitor the status of the host payload 1-2 and various parameters during the separation process in real time. In addition, the host payload 1-2 is simplified in the test. The host payload 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 host payload 1-2 can be adjusted quickly and easily by adding or removing mass blocks, so as to verify the separation situation under eccentric center of mass conditions. The host payload side plate and the host payload bottom plate are connected to the aluminum profile frame by bolts, and the host payload side plate is fixed to the antenna satellite substrate by bolts. During the separation process, the host payload 1-2 is quickly and safely separated and locked through the mechanical interface system 1-3 to ensure a rapid response in an emergency and protect the safety of the host payload.

[0105] The following further introduces the composition of the No. 1 separation component and the No. 2 separation component:

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

[0107] The hosted 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 put on the ejection plate interface 1-3-2;

[0108] Separation component No. 2 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 sequence, and is inserted into the interior of the base 1-3-7 through the compression power spring device 1-3-4; the No. 1 separation component is confined within 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 named as the first circular support plate and the second circular support plate; the first circular support plate is fixed to the bottom of the pillar, and 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 above 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 docks 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, such as Figure 10 As shown. The components of the mechanical interface system 7 are as follows Figure 9 As shown in the figure, the mechanical interface system is designed with multiple separation mechanisms and safety locking devices, allowing for adjustments to suit diverse testing requirements. Springs and locking devices enable rapid separation of the hosted payload while ensuring a smooth and safe separation process. The system's adaptive capabilities allow it to quickly respond to separation commands without manual intervention, making it particularly suitable for payload protection and mission execution in emergency situations.

[0113] In the mechanical interface systems 1-3, hosted payloads 1-2, the antenna satellite baseplate, and the ejection plate interface are initially fixedly connected and separate as a single unit during separation. The lower surface of the ejection plate interface contacts the upper surface of the host satellite's top plate, while the threaded cap is fixedly connected to the lower surface of the host satellite's top plate. Furthermore, to ensure the hosted payload can function properly after separation, the ejection plate interface and the host satellite are equipped with electrical interfaces, information transmission interfaces, and thermal protection subsystems.

[0114] Regarding the design and separation of the mechanical interface system, base 1-3-7 is fixed to the host satellite, and release motors 1-3-8 are mounted on either side of base 1-3-7 and connected to the base via bolts. The release motors are pushrod motors, with their tops directly connected to spring plates, which drive the spring plates along the base.

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

[0116] The base of the power spring assembly 1-3-4 is connected to the top of the power adjustment motor 1-3-9, while its top plate contacts the bottom of the ejection plate interface. The power spring assembly comprises four symmetrically distributed springs that generate an upward force when compressed. When unlocked, this force pushes the top plate, separating the ejection plate interface 1-3-2, its attached hosted payload 1-2, and the antenna satellite baseplate 1-3-1.

[0117] The locking steel ball 1-3-5 is placed in the center of the base 8, and a groove is formed 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, maintaining structural stability. Once the release motor 1-3-8 is activated, the spring plate 1-3-6 moves along the base, allowing the locking steel ball 1-3-5 to enter 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 automated separation and locking functions.

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

[0119] Separator No. 2 also includes 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 piece is further defined below:

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

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

[0125] The following describes the components 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 this device:

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

[0131] A processor is provided in the host 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 by the dynamic capture mechanism in real time and the relative displacement output by the laser sensor 1-3-10 in real time, and performing data fusion on the two relative displacements after spatial registration to obtain a 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 registration and space registration, and the two accelerations after spatial registration are data fused to obtain a final host satellite 1-1 acceleration, as well as 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 registration and space registration, and the two accelerations after spatial registration are data fused to obtain a 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 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.

[0133] Specifically, the entire separation process is the preparatory phase before unlocking, with a speed of zero. The second phase is from unlocking the unlocking device to complete separation, where the compression spring goes from compressed to fully expanded, causing the satellite to undergo a sudden change from zero speed. The third phase is the post-separation phase, where the spring force disappears and the satellite completes its complete separation state, at which point its attitude and velocity stabilize. This embodiment primarily considers the satellite's attitude feedback and acceleration / velocity feedback during the second separation process. Because multiple sensors are used, and target measurements are typically performed asynchronously, directly fusion of asynchronous data would not only fail to improve the system's measurement accuracy, but would also reduce the tracking effectiveness of the separated satellite. Therefore, data from various measurement sensors, such as the motion capture system, gyroscopes, accelerometers, and laser sensors, undergo spatiotemporal registration. This involves transforming the observation data from these sensors into the same spatial coordinate system and aligning them on the time axis to ensure spatiotemporal consistency.

[0134] The process of temporal registration is as follows:

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

[0136]

[0137] In the formula, a0, a1, and a2 are the polynomial coefficients of the fitting curve, and 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 of the i-th moment, j represents the polynomial order of the fitting curve;

[0138] Calculate the error between each first host satellite acceleration curve and the acceleration acquisition value as the first host satellite acceleration error, calculate the error between each second host satellite acceleration curve and the acceleration acquisition value as the second host satellite acceleration error, calculate the error between each first hosted payload acceleration curve and the acceleration acquisition value as the first hosted payload acceleration error, calculate the error between each second hosted payload acceleration curve and the acceleration acquisition value as the second hosted payload acceleration error, calculate the error between each first displacement curve and the displacement acquisition value as the first displacement error, and calculate the error between each second displacement curve and the displacement acquisition value as the second displacement error. The error δ i Expressed as:

[0139]

[0140] Where, 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 from multiple first host satellite acceleration errors is selected as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value from multiple second host satellite acceleration errors is selected as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value from multiple first host payload acceleration errors is selected as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value from multiple second host payload acceleration errors is selected as the second host payload acceleration curve after screening; a first displacement curve corresponding to a minimum value from multiple first displacement errors is selected as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value from multiple second displacement errors is selected as the second displacement curve after screening.

[0142] Specifically, because the frequencies of the two accelerations or two displacements collected by two different components are different, time alignment is required to fit the accelerations and displacements into curve form. In this way, at the same moment, the point values ​​taken on each curve are 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 it into a polynomial function form, that is: Transforming formula 6, we get: Write 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 the ground coordinate system.

[0147] Specifically, spatial registration is to align 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, taking into account 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 the state coordinates of the target are converted into the ground system based on the measurement value of the sensor and the geographic coordinates of the sensor itself. 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] Where, is the position information after the target measurement value is fused 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 data fusion process is as follows:

[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 payload acceleration curve and the second host payload acceleration curve after screening in the ground coordinate system are fused to obtain the fused host payload acceleration curve. At the same time, 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] Where, 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] Where, 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, 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 position 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 The relationship 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] Where 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 that:

[0165]

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

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

[0168] The minimum value of the total variance at this time is as follows:

[0169]

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

[0171] like Figure 13 As shown in the figure, 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 radius vectors 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 point of action of the spring, 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, the magnitude of which 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 expressions are:

[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 experimental flow chart is as follows Figure 14To fully illustrate the objectives and advantages of the present invention, a test device for monitoring the separation and locked attitude of a host satellite and a hosted payload is described in detail, combined with the accompanying drawings and actual operation examples. The design of this test device primarily studies the effect of the stiffness of the spring in the mechanical interface on the separation speed and direction of the host satellite and hosted payload, the effect of the separation speed and direction on the attitude of the host satellite and hosted payload, and the effect of machining errors based on the mechanical interface on the attitude of the host satellite and hosted payload.

[0183] Step 1: Equipment Installation and Initialization: Install the host satellite 1-1 and hosted payload 1-2 onto the air flotation platform 4 and connect them via the mechanical interface system 1-3. Adjust the air support system 2 to an appropriate height to ensure that the host satellite and hosted payload can float in a simulated microgravity environment during the experiment.

[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: Position the laser sensor 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, and sets the data acquisition frequency and recording parameters, such as separation speed, direction, and attitude changes.

[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 and spring stiffness of the power spring assembly 1-3-4 based on the mass, size, and shape of the load 1-2. Calculate the spring force F in the pre-compressed state using Hooke's law to ensure that this force is sufficient to successfully separate the load 1-2.

[0190] Step 8: Initial Connection: Figure 9As shown, the hosted payload 1-2 is fixedly connected to the ejection plate interface 10 via the antenna satellite base plate 9, and the entire structure is mounted on the host satellite roof. 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 system stability.

[0191] Step 9: Power Adjustment: Figure 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 the predetermined position, which corresponds to the 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 Middle spring force F.

[0192] Step 10: Pre-Separation Inspection: Inspect all components to ensure that the host payloads 1-2 can function normally after separation. System testing and data feedback confirm that all preparations are complete and the separation system is in standby mode.

[0193] Step 11: Start the release motor: Figure 9 Operate the release motor 1-3-8 to push the spring plate 1-3-6 to move along the base 1-3-7 to a specific position, 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: Figure 9 As shown, the power spring assembly 1-3-4 releases its spring force, pushing the ejection plate interface 1-3-2 upward and apart. Analyze the relationship between the power spring's compression and propulsion speed, the stress changes in the locking spring, and the maximum load capacity of the mechanism to ensure the reliability of the propulsion and launch system.

[0195] Step 13: Data Collection: During the experiment, the dynamic capture system 3, gyroscopes, accelerometers, and laser sensors synchronously record all dynamic parameters during the separation process. This data includes the velocity, acceleration, rotation angle of the host satellite 1-1 and the hosted payload 1-2, as well as the precise moment of separation.

[0196] Step 14: Use dedicated software on 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 t3) 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 operating conditions: Based on the actual space environment and mission requirements, design operating conditions with different spring stiffness, pre-compression, and center of mass offset. Each operating condition attempts to simulate possible extreme or special conditions.

[0199] Step 17: Operating Condition Test: Conduct actual tests for each designed operating condition, collecting data on separation speed, direction, and posture changes. In addition, introduce intentional machining errors, such as uneven spring stiffness and variations in contact surface roughness, to evaluate their specific impact 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 during the calculation of velocity and displacement. The dynamic capture system and laser sensor have low accuracy. Therefore, the use of multi-source information fusion technology can 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 under different operating conditions with the ideal operating 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 any significant error sources identified. 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 the separation action achieves the highest efficiency and accuracy.

[0205] Step 23: Optimization Verification Test: Validate the effectiveness of the adjusted parameters through re-experimentation. Repeat the experiment, collect new data, and perform data processing and performance evaluation again to ensure that the adjustments have effectively improved system performance.

[0206] Step 24: Verification and Feedback: After implementing the corrective actions, rerun the test to verify the results. Based on the new test results, conduct further data analysis to assess the error.

[0207] Step 25: Through this series of tests and analyses, the mechanical interface system of this invention can effectively achieve rapid and safe separation of the host satellite and its 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 machining 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 , meeting mission requirements while avoiding near-field interference (NSI), a risk of close collision between the host satellite and its 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 machining errors, improving separation efficiency, and exploring new materials and technologies to enhance 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 illustrative of the principles and applications of the invention. It should be understood that many modifications may be made to the illustrative embodiments, and that other arrangements may be devised, without departing from the spirit and scope of the invention as defined by the appended claims. It should be understood that the various dependent claims and features described herein may be combined in ways other than those described in the original claims. It should also be understood that features described in conjunction with individual embodiments may be employed in conjunction with other described embodiments.

Claims

1. A test device for monitoring the separation and locking attitude of a host satellite and a hosted payload, characterized by: The test device comprises a simulation device (1), a gas 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 for monitoring dynamic parameters during the separation process of the host satellite (1-1) and the hosted payload (1-2); The test device also includes a host computer (4); A processor is provided in the host 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 data fusion 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, and the two accelerations after spatial alignment are data-fused to obtain a final host satellite (1-1) acceleration; and the acceleration of the hosted payload (1-2) output in real time by the dynamic capture mechanism and the acceleration of the hosted payload (1-2) output by the acceleration sensor located in the hosted payload (1-2) are sequentially time-aligned and spatially-aligned, and the two accelerations after spatial alignment are data-fused to obtain a final hosted payload (1-2) acceleration; It is also used to receive the angular velocity of the host satellite (1-1) and the hosted payload (1-2), and display it, and simultaneously display 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).

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) 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); 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 payload frame (1-2-1) is a cube frame, the host payload side plate (1-2-2) is rotatably connected to the side frame of the host payload frame (1-2-1) by rotating bolts (1-1-4), and the host payload bottom plate (1-2-3) is fixed to the bottom frame of the host payload frame (1-2-1) by bolts; and the host payload side plate (1-2-2) is arranged opposite to the host satellite side plate (1-1-2); The mechanical interface system (1-3) includes a first separation component and a second separation component; Separator No. 1 is set in the host payload frame (1-2-1), and separator No. 2 is set 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 payload side plate (1-2-2), and the host satellite side plate (1-1-2) and the host payload side plate (1-2-2) are arranged parallel and opposite to each other; The No. 1 separating 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 separating piece. The No. 1 separating piece and the No. 2 separating 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 antenna satellite base plate (1-3-1), ejection plate interface (1-3-2) and threaded cover (1-3-3); The host 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 separating component 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 payload 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 confined within the host payload frame (1-2-1) through the threaded cover (1-3-3); The base (1-3-7) includes a hollow pillar and two circular support plates. The first circular support plate is fixed to the bottom of the hollow pillar, and the second circular support plate is sleeved and fixed on the hollow pillar. Through holes are evenly provided on the side wall of the hollow pillar at the upper portion 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). The spring plate (1-3-6) is sleeved on the hollow 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 (1-3-8) 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: Separator No. 2 also includes a power regulating motor (1-3-9); A power regulating motor (1-3-9) is provided between the base (1-3-7) and the power spring device (1-3-4); The power regulating motor (1-3-9) is used for driving the power spring device (1-3-4) to move upwards 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 of 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 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); 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 1, 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-type 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-type host satellite acceleration curves; the acceleration of the hosted payload (1-2) captured in real time by the dynamic capture mechanism is fitted into a plurality of first-type hosted payload acceleration curves, and the acceleration of the hosted payload (1-2) output in real time by the acceleration sensor is fitted into a plurality of second-type hosted payload acceleration curves; the relative displacement captured in real time by the dynamic capture mechanism is fitted into a plurality of first-type displacement curves, and the relative displacement output in real time by the laser sensor (1-3-10) is fitted into a plurality of second-type displacement curves. Each fitted displacement curve and acceleration curve is expressed as: In the formula, a0, a1, and a2 are the polynomial coefficients of the fitting curve, and 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 of the i-th moment, j represents the polynomial order of the fitting curve; Calculate the error between each first host satellite acceleration curve and the acceleration acquisition value as the first host satellite acceleration error, calculate the error between each second host satellite acceleration curve and the acceleration acquisition value as the second host satellite acceleration error, calculate the error between each first hosted payload acceleration curve and the acceleration acquisition value as the first hosted payload acceleration error, calculate the error between each second hosted payload acceleration curve and the acceleration acquisition value as the second hosted payload acceleration error, calculate the error between each first displacement curve and the displacement acquisition value as the first displacement error, and calculate the error between each second displacement curve and the displacement acquisition value as the second displacement error. The error δ i Expressed as: Where, 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 from multiple first host satellite acceleration errors is selected as the first host satellite acceleration curve after screening, and a second host satellite acceleration curve corresponding to a minimum value from multiple second host satellite acceleration errors is selected as the second host satellite acceleration curve after screening; a first host payload acceleration curve corresponding to a minimum value from multiple first host payload acceleration errors is selected as the first host payload acceleration curve after screening, and a second host payload acceleration curve corresponding to a minimum value from multiple second host payload acceleration errors is selected as the second host payload acceleration curve after screening; a first displacement curve corresponding to a minimum value from multiple first displacement errors is selected as the first displacement curve after screening, and a second displacement curve corresponding to a minimum value from multiple second displacement errors is selected as the second displacement curve after screening.

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 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 the ground coordinate system.

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 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 payload acceleration curve and the second host payload acceleration curve after screening in the ground coordinate system are fused to obtain the final host payload acceleration curve. At the same time, the first displacement curve and the second displacement curve after screening in the ground coordinate system are fused to obtain the final displacement curve: Where, 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: Where, 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, The minimum total variance of all sensor measurements.

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