Spacecraft unlocking shock measurement simulation test system

Abstract

A spacecraft unlocking impact measurement simulation test system belongs to the field of spacecraft ground test simulation technology. The present invention addresses the problems of high ground test cost and low efficiency in spacecraft pyrotechnic impact unlocking tests. It includes an on-board test structure, which is fixedly connected to the base using a simulated separation nut, and a wedge-shaped pad is provided on the upper end of the base to simulate the on-board clamping slider assembly; a simulated cabin plate is fixedly connected to the bottom end of the base; a simulated separation mechanism, which is fixedly connected to the upper end of a simulated bolt with a notch in the middle section using a bottom end face of a breakaway adapter frame, and the lower end of the simulated bolt is threadedly connected to the simulated separation nut; a pressure loading module, which uses a hydraulic press to pressurize the breakaway adapter frame to generate an upward thrust, and pulls the simulated bolt to generate a pre-tightening force; an acceleration impact sensor is provided on the base to monitor the impact load generated by the separation of the simulated separation mechanism and the on-board test structure. The present invention is used for spacecraft unlocking impact simulation testing.

Description

Technical Field

[0001] The invention relates to a spacecraft unlocking impact measurement simulation test system, and belongs to the technical field of spacecraft ground test simulation. Background Art

[0002] Reliable separation of the separation system is a core mission of spacecraft, and separation through explosive devices is the most mature and widely used separation method for spacecraft. The separation process generates large, high-frequency shock loads, which affect the accuracy, lifespan, and reliability of sensitive components and precision instruments and equipment, directly impacting the success of the entire spacecraft mission. Therefore, simulating separation scenarios on the ground to quantitatively acquire, analyze, and verify the separation shock loads and their responses plays a vital role in overall spacecraft performance.

[0003] Pyrotechnic separation technology mainly uses pyrotechnic cutters to cut off the clamping rod, detonate explosive bolts, and unlock pyrotechnic locks to quickly release the connection. However, directly conducting pyrotechnic impact unlocking tests has problems such as complex equipment, high cost, and non-repeatability, making ground testing difficult and inefficient.

[0004] Therefore, it is of great significance to design an equivalent ground simulation unlocking impact test system.

[0005] Existing ground simulation tests: Initially, simulations of pyrotechnic shock environments mostly employed standard pulse time-domain methods, employing typical time-domain signals such as sine and triangular waves. The impact energy generated by an explosion is concentrated in the high-frequency portion, which is a major factor in equipment failure. The impact effects produced by typical shock waves differ significantly from those produced by pyrotechnic explosions, often failing to characterize the high-frequency characteristics of blast shock. Since the 1970s, with in-depth research, shock response spectra have become a key engineering tool for characterizing high-frequency shock environments. Testing in pyrotechnic shock environments based on shock response spectrum theory can more accurately characterize the shock response characteristics of structures and reveal the relationship between shock loads and structural dynamic responses, resulting in test results that more closely resemble real-world conditions. Pyrotechnic shock testing techniques based on shock response spectra are primarily categorized as pyrotechnic explosion and mechanical impact. Pyrotechnic explosion methods suffer from high cost and non-repeatability, while mechanical impact methods face challenges such as long pre-test commissioning cycles and the inability of the generated shock waves to fully guarantee the frequency of the blast shock load. Summary of the Invention

[0006] Aiming at the problems of high ground testing cost and low efficiency in the impact unlocking test of spacecraft pyrotechnic devices, the present invention provides a spacecraft unlocking impact measurement simulation test system.

[0007] A spacecraft unlocking impact measurement simulation test system of the present invention comprises:

[0008] The onboard test structure uses a simulated separation nut to be fixedly connected to the base in the base cavity, and a wedge-shaped pad is set at the upper end of the base along the outer contour of the simulated separation nut to simulate the onboard compression slider assembly; a simulated cabin plate is used to simulate the onboard platform cabin mounting plate and is fixedly connected to the bottom end of the base;

[0009] The simulated separation mechanism adopts a fixed connection between the bottom end surface of the breakaway adapter frame and the upper end surface of the notch of the simulated bolt with a notch in the middle section, and the lower end of the simulated bolt is threadedly connected to the simulated separation nut; a support base is provided in the breakaway adapter frame, and the support base adopts a hollow design and is placed on the wedge-shaped pad by passing the support legs through the bottom end surface of the breakaway adapter frame;

[0010] The pressure loading module uses a hydraulic press to apply pressure to the breakaway adapter frame to generate an upward thrust, pulling the simulated bolt to generate a pre-tightening force; the hydraulic press is set on the support base inside the breakaway adapter frame; the hydraulic press monitors the force loading condition through a pressure gauge; a bolt pre-tightening force sensor is set at the bottom of the nut of the simulated bolt to monitor the pre-tightening force generated by the simulated bolt; an acceleration impact sensor is set on the base to monitor the impact load generated by the separation of the simulated separation mechanism and the on-board test structure.

[0011] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the hydraulic press is connected to the electric oil pump via a hydraulic oil pipe; a safety valve is provided on the hydraulic oil pipe.

[0012] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the pressure loading module also includes an inductive limit switch; the inductive limit switch is arranged on the buffer pad; the buffer pad is correspondingly arranged under the on-board test structure, for buffering the impact force generated by the falling of the on-board test structure; the on-board test structure falls onto the buffer pad, triggering the inductive limit switch, causing the hydraulic press to stop pressurizing.

[0013] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the hydraulic press is used to apply pressure, control the pre-tightening force of the simulated bolt to reach a preset pre-tightening force value, and then release it, so that the simulated separation mechanism is separated from the on-board test structure; the stress wave formed by the simulated bolt separation propagates along the base to the satellite.

[0014] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the simulation bolt is a T-shaped pull-off rod made of titanium alloy TC4, and the notch position in the middle section is the brittle fracture position. After the fracture, the simulation separation mechanism is separated from the on-board test structure.

[0015] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the simulated bolt notch radius r is:

[0016]

[0017] Where F is the preload force that causes the simulated bolt to fracture under actual working conditions, σ b To simulate the tensile stress of the bolt.

[0018] According to the spacecraft unlocking shock measurement simulation test system of the present invention, the acceleration shock sensor is arranged between the base and the simulation cabin panel.

[0019] The spacecraft unlocking impact measurement simulation test system according to the present invention uses a data acquisition card to collect monitoring data from a bolt preload sensor and an acceleration impact sensor. Based on the monitoring data transmitted by the data acquisition card, the controller controls the pressure loading module via a loading controller to ensure that the preload force of the simulated bolt reaches a preset preload force value. When the simulated separation mechanism separates from the onboard test structure, the controller controls the hydraulic press to stop pressurizing based on a trigger signal from an inductive limit switch.

[0020] The loading controller also controls the hydraulic press and safety valve according to the monitoring results of the pressure gauge to prevent overload and maintain pressure.

[0021] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the shape, structure and material of the base are the same as those of the corresponding base of the spacecraft to be simulated.

[0022] Beneficial effects of the present invention: The present invention is used for simulation testing of shock response during separation and unlocking of a spacecraft, and is suitable for shock testing in situations such as spacecraft payload release and separation, and interstage separation.

[0023] The test system described in the present invention can be applied to solve the equivalent measurement of impacts generated by various separation mechanisms, reduce the test cost of separation impact testing, and reduce manpower and material resources; simplify the impact test equipment, reduce the difficulty of ground testing; improve test efficiency, and can accurately, quickly and repeatedly perform buffer verification tests.

[0024] The equivalent separation mechanism of the present invention is simple and ingenious. A breakaway adapter converts the thrust provided by the hydraulic device into a pulling force (directed upward), while a hollow support base transmits the reaction force (directed downward), causing the bolt to break. Because the power source is a split, lightweight hydraulic press with small size and light weight, both the breakaway adapter and the support base in the simulated separation mechanism can be flexibly designed to adapt, making the device highly adaptable.

[0025] During the establishment of the system of the present invention, a finite element model of the preload release process can be established to numerically simulate the impact response generated by the separation process, and an equivalent scheme can be designed with reference to the simulation results. The accuracy of the separation impact test simulated by the present invention is sufficiently high. By using an electric hydraulic device as a driving source, the equivalent preload loading process is stable and reliable, and the obtained preload release effect is close to that on a real satellite. The amplitude change trend of the impact response spectrum is basically consistent, and the peak simulation deviation is less than 5%. Titanium alloy rods are used as test simulation bolts. Compared with explosive bolts, pyrotechnic nuts and other pyrotechnic products, the cost is lower, the safety is stronger, and the test form is flexible. Different test requirements can be achieved by simply replacing different breaking rods. The preload sensor makes the preload size intuitively measurable, thereby improving the test accuracy.

[0026] Experimental verification demonstrates that the proposed system can quickly and efficiently determine the impact suppression effects of different cushioning schemes, verifying spacecraft cushioning schemes with good repeatability and efficiency. Through pressure loading control and acquisition measurement, the system can accurately control the magnitude of the impact source released by different preload forces. It can also quantitatively analyze and verify the effectiveness of various impact reduction measures at low cost, with strong engineering feasibility. This system plays an important role in researching impact cushioning and verifying energy absorption methods. BRIEF DESCRIPTION OF THE DRAWINGS

[0027] Figure 1 This is a design flow chart of the spacecraft unlocking impact measurement simulation test system of the present invention;

[0028] Figure 2 It is a schematic diagram of a single set of inter-cabin compression and release mechanisms of the spacecraft to be simulated;

[0029] Figure 3 It is a schematic diagram of establishing a simulation model in the system design process of the present invention;

[0030] Figure 4 It is a schematic diagram of the acceleration response observation point location of the simulation model;

[0031] Figure 5 yes Figure 4 The acceleration response curve of measuring point 1 (point 1);

[0032] Figure 6 yes Figure 4 The acceleration response curve of measuring point 2 (point2);

[0033] Figure 7 yes Figure 4 The acceleration response curve of measuring point 3 (point 3);

[0034] Figure 8 yes Figure 4 The acceleration response curve of measuring point 4 (point 4);

[0035] Figure 9 It is the equivalent stress cloud diagram of the model section at 2ms during the simulation model establishment process;

[0036] Figure 10 It is the equivalent stress cloud diagram of the model section at 4ms during the simulation model establishment process;

[0037] Figure 11 This is the stress cloud diagram of the base section at 6ms during the simulation model establishment process;

[0038] Figure 12 This is the stress cloud diagram of the base section at 8ms during the simulation model establishment process;

[0039] Figure 13 Schematic diagram of the hardware structure of the spacecraft unlocking impact measurement simulation test system of the present invention;

[0040] Figure 14 yes Figure 13 The enlarged view of the part in the middle A frame;

[0041] Figure 15 yes Figure 13 Schematic diagram of the structural dimensions;

[0042] Figure 16 yes Figure 13 Schematic diagram of the side view structure size;

[0043] Figure 17 It is the control principle diagram of the system of the present invention;

[0044] Figure 18 It is a schematic diagram of the structure of the onboard test structure;

[0045] Figure 19 It is a structural diagram of the simulated separation mechanism;

[0046] Figure 20 yes Figure 19 Side view of;

[0047] Figure 21 This is a schematic diagram of the onboard test structure and simulated separation mechanism before the simulated bolt breaks;

[0048] Figure 22 This is a schematic diagram of the onboard test structure and simulated separation mechanism after the simulated bolt fracture;

[0049] Figure 23 It is a structural diagram of the simulated bolt;

[0050] Figure 24 This is a comparison chart of acceleration shock response curve test results obtained by using the system of the present invention to conduct a simulated unlocking shock test; g represents acceleration; Curve represents curve;

[0051] Figure 25 This is a schematic diagram of the targeted energy absorption improvement design of the base;

[0052] Figure 26 It is a schematic diagram of setting a buffer link at the lower end of the simulated separation nut. DETAILED DESCRIPTION

[0053] 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.

[0054] 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.

[0055] 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.

[0056] Specific implementation method 1. Combination Figures 1 to 23 As shown, the present invention provides a spacecraft unlocking impact measurement simulation test system, comprising:

[0057] First, the spacecraft's payload compression and release mechanisms must be analyzed to design a corresponding ground-based simulated unlocking impact test device. In this ground-based simulation, a lightweight hydraulic device applies a preload, which then snaps a metal rod, creating an ignition-like action for rapid unlocking. While retaining the basic configuration of the existing inter-cabin connection, the test device is simplified by creating an equivalent mechanical environment. Accelerations at each measurement point are collected to obtain the impact response at the corresponding key locations on the spacecraft. Repeated testing can be performed cost-effectively simply by replacing simulated bolts.

[0058] First, based on the actual physical process of the release mechanism, the causes of the impact load and the shock wave transmission path are analyzed, laying the foundation for the design of the ground simulation device.

[0059] by Figure 2Take the single set of clamping and releasing mechanism between the platform cabin and the payload cabin of the spacecraft shown as an example. The mechanism consists of a cover plate, a connecting tube, a clamping rod assembly, a clamping slider assembly, a base, a strut, an in-position switch, and a pyrotechnic separation nut. Among them: the bottom of the base is fixedly connected to the single-machine structural plate of the platform cabin; the clamping rod assembly provides a clamping load of 80kN when clamped, and the clamping rod assembly is released under the drive of the energetic material to achieve the detachment of the payload cabin. During the research and development stage of the spacecraft, it is necessary to test the impact environment of the onboard equipment to obtain the impact response of the key positions on the satellite, and then guide the next engineering practice. The traditional onboard test method is cumbersome, the excessive number of components inevitably makes the assembly complicated, and the pyrotechnics are expensive and non-reusable. These shortcomings make the test difficult and inefficient under the traditional experimental scheme, so it is of great significance to design an equivalent ground simulation unlocking impact test system.

[0060] At the moment of separation and unlocking, the sudden release of the compression rod is the main factor in the generation of shock waves. The huge preload release will cause stress waves to propagate on the satellite, which is the main reason for the structural resonance response. Therefore, it is necessary to establish a numerical model of the sudden release of the load of the separation mechanism under the preloaded state of the bolts, and analyze the impact magnitude and shock wave transmission law:

[0061] Can be used in LS-DYNA Figure 2 The mechanism is simulated and analyzed to understand the propagation mechanism of the impact load generated by the release of the preload, providing support for the design of the equivalent test plan. An equivalent simplified model is established, in which the platform cabin and the base are connected by studs and nuts, both of which are made of titanium alloy, to constrain the bottom of the base. The preload release process is simulated by gradually increasing the tension on the studs to a peak of 80kN and then suddenly reducing it to 0. The model is established as follows Figure 3 shown.

[0062] Set up four equally spaced acceleration response observation points on the side of the base from top to bottom, and the grid division is as follows: Figure 4 shown.

[0063] The acceleration response curve of the observation point is obtained by simulation, such as Figures 5 to 8 As shown in the figure. When the external load is released, the acceleration shock response of each point on the base increases instantly, with high amplitude and frequency. The peak values ​​of the four measuring points are 155180m / s respectively. 2 、110860m / s 2 、107330m / s 2 、84363m / s 2 , indicating that the shock wave generated by the separation propagates along the base toward the satellite, and the shock response will continue to decay with the increase of the transmission distance.

[0064] Figure 9 and Figure 10The equivalent stress cloud diagram of the entire cross section of the model shows that: 2ms ago, when the preload has not yet been released, the structure has a high stress level. 2ms later, when the external load is released instantaneously, the strain energy in the structure is released, and the stress wave in the stud begins to propagate outward. Figure 11 and Figure 12 The stress cloud diagram of the base section shows that the shock wave propagates to the surrounding structure, causing the entire mechanism to respond and vibrate.

[0065] From the above analysis, it can be seen that the stress wave propagation and structural resonance response caused by the release of the 80kN preload at the moment of separation of the clamping rod and the separation nut can produce a large impact response on the base, and the shock wave will propagate from the impact source along the base to the satellite. Therefore, the base can be used as a key position for shock response consideration. The impact acceleration measurement points of the original separation device and the equivalent experimental device can be mainly placed at the bottom edge of the base, that is, at one end connected to the spacecraft platform cabin.

[0066] Based on the numerical analysis results of preload release, an equivalent design of the separation mechanism is performed: the physical process of releasing the clamping assembly is simulated, and the mechanism design focuses on the loading and sudden release of the preload. The impact transmission path in the original structure is retained as an analysis object to examine whether the output results are valid:

[0067] Based on the above analysis, an equivalent test device was designed. The original base structure was used as the propagation carrier and collection object of the shock wave. The connection method between the pyrotechnic separation nut and the base was used to perform an equivalent design of the pyrotechnic separation mechanism. The final equivalent test system includes the on-board test structure, simulated separation mechanism, pressure loading module, and measurement and collection equipment. The overall configuration diagram is shown below. Figure 13 and Figure 14 shown.

[0068] The onboard test structure uses a simulated separation nut 110 to be fixedly connected to the base 120 within its inner cavity. A wedge-shaped pad 130 is provided at the upper end of the base 120 along the outer contour of the simulated separation nut 110 to simulate the onboard compression slider assembly. A simulated cabin plate 140 is used to simulate the onboard platform cabin mounting plate and is fixedly connected to the bottom end of the base 120.

[0069] The simulated separation mechanism adopts a method of fixedly connecting the bottom end surface of the breakaway adapter frame 210 with the upper end surface of the notch of the simulated bolt 220 with a notch in the middle section. The lower end of the simulated bolt 220 is threadedly connected to the simulated separation nut 110. A support base 230 is provided in the breakaway adapter frame 210. The support base 230 adopts a hollow design and is placed on the wedge-shaped pad 130 through the support legs passing through the bottom end surface of the breakaway adapter frame 210.

[0070] The pressure loading module uses a hydraulic press 310 to apply pressure to the break adapter frame 210 to generate an upward thrust, pulling the simulated bolt 220 to generate a pre-tightening force; the hydraulic press 310 is set on the support base 230 in the break adapter frame 210; the hydraulic press 310 monitors the force loading status through the pressure gauge 320; a bolt pre-tightening force sensor 330 is set at the bottom of the nut of the simulated bolt 220 to monitor the pre-tightening force generated by the simulated bolt 220; an acceleration impact sensor 340 is set on the base 120 to monitor the impact load generated by the separation of the simulated separation mechanism and the on-board test structure.

[0071] The pressure loading module can be composed of a separate electric hydraulic press, hydraulic oil pipes, and loading control equipment. The hydraulic press 310 is placed on a hollow support base and connected to an electric oil pump via hydraulic oil pipes to provide loading pressure for the simulated separation mechanism. A safety valve is installed on the hydraulic oil pipe. The loading control equipment includes a pressure gauge, inductive limit switch, electric hydraulic oil pump, and safety valve to achieve control of the loading pressure required for the experiment.

[0072] In this embodiment, the pressure loading module also includes an inductive limit switch 350; the inductive limit switch 350 is arranged on a buffer pad 400; the buffer pad 400 is correspondingly arranged below the on-board test structure to buffer the impact force generated by the falling of the on-board test structure; the on-board test structure falls onto the buffer pad 400, triggering the inductive limit switch 350, causing the hydraulic press 310 to stop pressurizing.

[0073] The hydraulic press 310 applies pressure to control the pre-tightening force of the simulated bolt 220 to reach a preset pre-tightening force value and then releases it, separating the simulated separation mechanism from the on-board test structure; the stress wave generated by the separation of the simulated bolt 220 propagates along the base 120 toward the satellite.

[0074] At the moment the mechanism completes separation, the impact load generated can be measured by the impact sensor attached to the onboard test structure. The onboard test structure then drops onto the cushion, triggering the inductive limit switch and stopping the pressure loading. The pressure gauge can monitor the loading status in real time to prevent overload risks. At the same time, the safety valve can maintain pressure and increase equipment safety. The system control schematic diagram is shown below. Figure 17 shown.

[0075] On-board test structure Figure 18 To ensure that the impact test and the actual impact transmission path on the satellite are consistent, the shape and material of the base are consistent with those on the satellite. The simulated cabin plate is used to simulate the platform cabin mounting plate on the satellite, and its impact response is measured. The simulated separation nut does not contain energetic material and is fixed to the base via eight bolts. The upper end will connect to the simulated bolts in the simulated separation mechanism. The clamping slider assembly on the satellite is simplified and replaced by a wedge-shaped spacer placed on the base.

[0076] Simulated separation mechanism such as Figure 19 and Figure 20 As shown in the figure, the simulated bolts are respectively passed through the breakaway adapter frame, the wedge-shaped pad, and tightened with the simulated separation nut. The support base adopts a hollow design and passes through the breakaway adapter frame through the notch and is placed on the wedge-shaped pad, on which a hydraulic press is placed.

[0077] For example, simulated bolt 220 is a T-shaped breakaway rod made of titanium alloy TC4. The mid-section notch represents a brittle fracture site, which, upon breaking, separates the simulated separation mechanism from the onboard test structure. A bolt preload sensor 330 is located at the base of the nut of simulated bolt 220 to facilitate real-time monitoring of the applied preload.

[0078] This embodiment also provides a protective net 500 for the safety of test personnel. The test personnel control the electric hydraulic device behind the protective net 500 to apply thrust to the breakaway adapter frame 210, causing the breakaway adapter frame 210 to pull the simulated bolt, causing the T-shaped breakaway rod to undergo brittle fracture, simulating a real separation situation. The onboard test structure falls off, triggering the inductive limit switch, and the pressure loading module stops loading. The internal structure before and after the bolt breaks is as shown in the figure. Figure 21 and Figure 22 shown. Figure 13 As shown, a steel cable 600 is further provided for suspending and breaking the adapter frame 210 .

[0079] According to the actual working conditions or the impact magnitude collected by actual measurement, the simulated bolts in the simulation experiment are designed. By selecting the breaking rods of different sizes and materials, various task requirements can be met:

[0080] By replacing the same simulated bolts, multiple sets of tests can be quickly repeated. Alternatively, the impact response caused by releasing different preloads can be simulated by changing the simulated bolt fracture depth or material. The hydraulic press is placed on the main onboard test structure via a support and counter-support. It applies a thrust to the breakaway adapter frame while simultaneously applying a reaction force to the onboard test structure, ultimately causing the simulated bolt to fracture brittlely. The fracture rod notch radius is then calculated based on the tensile strength of the selected material and the actual preload under operating conditions.

[0081] Furthermore, the radius r of the notch of the simulated bolt 220 is:

[0082]

[0083] Where F is the preload force that causes the simulated bolt to fracture under actual working conditions, σ b To simulate the tensile stress of the bolt.

[0084] In the experiment, the tensile strength of the broken rod is 950 MPa. When the preload force is 80 kN under the simulated actual working conditions, the bolt will fracture brittlely at the fracture under this stress. The radius r of the fracture is selected as:

[0085]

[0086] Simulate bolt structure such as Figure 23 shown.

[0087] As an example, the acceleration impact sensor 340 is disposed between the base 120 and the simulation deck 140 .

[0088] In actual use, measuring points can be designed as needed, and acceleration impact sensors 340 can be set at the measuring points. For example, two measuring points can be evenly set along the circumferential direction under the base 120, and two measuring points can be set in an array form on the outside of the base 120 on the simulation cabin plate 140.

[0089] Furthermore, a data acquisition card is used to collect monitoring data from the bolt preload sensor 330 and the acceleration impact sensor 340. The controller controls the pressure loading module through the loading controller based on the monitoring data transmitted by the data acquisition card, so that the preload force of the simulated bolt 220 reaches the preset preload force value. When the simulated separation mechanism separates from the onboard test structure, the hydraulic press 310 is controlled to stop pressurizing based on the trigger signal of the inductive limit switch 350.

[0090] The loading controller also controls the hydraulic press 310 and the safety valve according to the monitoring results of the pressure gauge 320 to prevent overloading and maintain pressure.

[0091] In this embodiment, the shape, structure and material of the base 120 are the same as those of the corresponding base of the spacecraft to be simulated.

[0092] In the simulation experiment, the SRS impact response data at the corresponding measuring points are obtained and compared to determine the effectiveness of the equivalent simulation experiment.

[0093] The spacecraft was subjected to a cabin unlocking impact test. The cabin mechanism pyrotechnics were successfully unlocked and the slider was ejected successfully. The impact acceleration response of the key position on the spacecraft during the unlocking process was obtained. The test results are as follows: Figure 24 As shown by the red dotted line in the middle, the peak of the shock response spectrum at the equipment installation location is basically concentrated in the frequency band near 3000 Hz, and the shock response peak is 9617g at 3758 Hz.

[0094] The acceleration shock response curve obtained by simulating the unlocking shock test is as follows: Figure 24 As shown by the black solid line in the middle, the trend of its amplitude change with frequency is basically consistent with the on-board pyrotechnic unlocking. The peak of the impact response spectrum is concentrated in the frequency band near 3000 Hz, and the peak of the impact response is 10054g at 3274 Hz. The relative error of the peak simulation is less than 5%. It can be determined that the equivalent test of the system of the present invention can well simulate the real impact effect caused by the actuation of the pyrotechnic device.

[0095] The system of the present invention is scalable and can perform comparative verification of buffering schemes.

[0096] The system of the present invention is low-cost, sophisticated in design, and highly scalable, and is suitable for simulating the collection of shock responses caused by the release of preload. When the shock response is too large and will damage sensitive components and equipment on the spacecraft, the original mechanism must be designed to reduce the shock. The following scheme can play a more important role in verifying and comparing various buffering schemes. The buffering scheme that can be verified and compared is as follows: Figure 25 and Figure 26 shown.

[0097] exist Figure 25 In the process, the base is changed and an energy-absorbing structure is added to it to play a buffering role. At the same time, the discontinuous structure can increase the reduction of the impact response amplitude; Figure 26 In the simulation, a foam metal buffer pad, a honeycomb aluminum buffer pad, etc. can be set at the bottom of the separation nut to absorb energy through material crushing.

[0098] The system of the present invention can be used to study buffering measures, verify buffering schemes, etc. The impact magnitude is adjustable, the device has strong scalability, and can carry out low-cost and high-efficiency comparative experiments.

[0099] 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 used in conjunction with other described embodiments.

Claims

1. A spacecraft unlocking impact measurement simulation test system, characterized in that: include: The on-board test structure adopts a simulated separation nut (110) to be fixedly connected to the base (120) in the inner cavity of the base (120), and a wedge-shaped pad (130) is provided at the upper end of the base (120) along the outer contour position corresponding to the simulated separation nut (110) to simulate the on-board pressing slider assembly; a simulated cabin plate (140) is adopted to simulate the on-board platform cabin mounting plate, and is fixedly connected to the bottom end of the base (120); The simulated separation mechanism adopts a fixed connection between the bottom end surface of the breakaway adapter frame (210) and the upper end surface of the notch of the simulated bolt (220) with a notch in the middle section, and the lower end of the simulated bolt (220) is correspondingly threadedly connected to the simulated separation nut (110); a support base (230) is provided in the breakaway adapter frame (210), and the support base (230) adopts a hollow design and is placed on the wedge-shaped pad (130) through the support legs passing through the bottom end surface of the breakaway adapter frame (210); the notch position in the middle section is a brittle fracture position, and after fracture, the simulated separation mechanism is separated from the on-board test structure; The pressure loading module uses a hydraulic press (310) to apply pressure to the breakaway adapter frame (210) to generate an upward thrust, thereby pulling the simulated bolt (220) to generate a pre-tightening force; the hydraulic press (310) is arranged on a support base (230) in the breakaway adapter frame (210); the hydraulic press (310) monitors the force loading condition through a pressure gauge (320); A bolt preload sensor (330) is provided at the bottom of the nut of the simulated bolt (220) to monitor the preload generated by the simulated bolt (220); an acceleration impact sensor (340) is provided on the base (120) to monitor the impact load generated by the separation of the simulated separation mechanism and the onboard test structure; The hydraulic press (310) applies pressure to control the pre-tightening force of the simulated bolt (220) to reach a preset pre-tightening force value and then releases the pre-tightening force, thereby separating the simulated separation mechanism from the on-board test structure; the stress wave generated by the separation of the simulated bolt (220) propagates toward the satellite along the base (120).

2. The spacecraft unlocking impact measurement simulation test system according to claim 1, characterized in that: The hydraulic press (310) is connected to the electric oil pump via a hydraulic oil pipe; a safety valve is provided on the hydraulic oil pipe.

3. The spacecraft unlocking impact measurement simulation test system according to claim 2, characterized in that: The pressure loading module further comprises an inductive limit switch (350); the inductive limit switch (350) is arranged on a buffer pad (400); the buffer pad (400) is correspondingly arranged below the onboard test structure and is used to buffer the impact force generated by the falling of the onboard test structure; when the onboard test structure falls onto the buffer pad (400), the inductive limit switch (350) is triggered, causing the hydraulic press (310) to stop pressurizing.

4. The spacecraft unlocking impact measurement simulation test system according to claim 3, characterized in that: The simulated bolt (220) is a T-shaped breaking rod made of titanium alloy TC4.

5. The spacecraft unlocking impact measurement simulation test system according to claim 4, characterized in that: The notch radius r of the simulated bolt (220) is: ， In the formula F The preload force that causes the simulated bolt to fracture brittlely in actual working conditions is To simulate the tensile stress of the bolt.

6. The spacecraft unlocking impact measurement simulation test system according to claim 5, characterized in that: The acceleration impact sensor (340) is arranged between the base (120) and the simulation cabin plate (140).

7. The spacecraft unlocking impact measurement simulation test system according to claim 6, characterized in that: A data acquisition card is used to collect monitoring data from a bolt preload sensor (330) and an acceleration impact sensor (340); a controller controls a pressure loading module via a loading controller based on the monitoring data transmitted by the data acquisition card, so that the preload force of the simulated bolt (220) reaches a preset preload force value, and when the simulated separation mechanism is separated from the onboard test structure, the hydraulic press (310) is controlled to stop pressurizing based on a trigger signal from an inductive limit switch (350); The loading controller also controls the hydraulic press (310) and the safety valve according to the monitoring result of the pressure gauge (320) to prevent overloading and achieve pressure maintenance.

8. The spacecraft unlocking impact measurement simulation test system according to claim 1, characterized in that: The shape, structure and material of the base (120) are the same as those of the corresponding base of the spacecraft to be simulated.

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