Spacecraft unlocking impact measurement simulation test system

Abstract

The invention discloses a spacecraft unlocking impact measurement simulation test system, and belongs to the technical field of spacecraft ground test simulation. The invention aims to solve the problems of high ground test cost and low efficiency in the impact unlocking test of the spacecraft initiating explosive device. Comprising an on-satellite test structure, a simulation separation nut is fixedly connected with a base, and a wedge-shaped cushion block is arranged at the upper end of the base to simulate an on-satellite compression sliding block assembly; a simulation cabin plate is fixedly connected with the bottom end of the base; the bottom end surface of a snapping transfer frame body is correspondingly and fixedly connected with the upper end surface of a notch of a simulation bolt with a notch in the middle section, and the lower end of the simulation bolt is correspondingly in threaded connection with a simulation separation nut; the pressure loading module adopts a hydraulic machine to pressurize the snapping transfer frame body to generate upward thrust, and pulls the simulation bolt to generate pre-tightening force; and an acceleration impact sensor is arranged on the base and is used for monitoring an impact load generated by separation of the simulation separation mechanism and the on-satellite test structure. The device is used for spacecraft unlocking impact simulation test.

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 one of the core tasks of spacecraft, and separation by pyrotechnics is the most mature and widely used separation method for spacecraft. During the separation process, a large high-frequency impact load will be generated, and the separation impact will affect the accuracy, life and reliability of sensitive components and precision instruments and equipment of the spacecraft, and directly affect the success of the entire spacecraft mission. Therefore, simulating the separation scene on the ground, quantitatively acquiring, analyzing, and verifying the separation impact load and its response are important for the overall performance of the spacecraft.

[0003] The pyrotechnic separation technology mainly uses pyrotechnic cutters to cut off the clamping rods, detonate explosive bolts, and unlock pyrotechnic locks to quickly release the connection. However, directly carrying out 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: The initial simulation of pyrotechnic shock environment mostly uses the standard pulse time domain method, which uses typical time domain signals such as sine waves and triangular waves. The impact energy generated by the explosion is concentrated in the high-frequency part, which is also the main factor causing equipment failure. There is a big difference between the impact effect generated by the typical shock wave and the impact effect of the pyrotechnic explosion, and it is often impossible to characterize the high-frequency characteristics of the explosion shock. Since the 1970s, with the deepening of research, the shock response spectrum has become the main engineering quantity for describing the high-frequency shock environment. Based on the shock response spectrum theory, the test of the pyrotechnic shock simulation environment can reasonably describe the shock response characteristics of the structure, and at the same time reflect the relationship between the shock load and the dynamic response of the structure, so that the test effect is closer to the real state. The pyrotechnic shock test technology based on the shock response spectrum is mainly divided into pyrotechnic explosion and mechanical impact. Among them, the pyrotechnic explosion has problems such as high cost and non-repeatability; the mechanical impact has problems such as long pre-test debugging cycle and the generated shock wave cannot fully guarantee the frequency of the explosion shock load. Summary of the invention

[0006] In view of the problems of high ground testing cost and low efficiency in the impact unlocking test of spacecraft pyrotechnics, 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 on-board test structure uses a simulated separation nut to be fixedly connected to the base in the inner cavity of the base, and a wedge-shaped pad is set at the upper end of the base along the outer contour position corresponding to the simulated separation nut to simulate the on-board compression slider assembly; a simulated cabin plate is used to simulate the on-board platform cabin mounting plate, which 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 simulated bolt with a notch in the middle section, and the lower end of the simulated bolt is threadedly connected with the simulated separation nut; a support base is arranged in the breakaway adapter frame, and the support base adopts a hollow design and is placed on a wedge-shaped pad by passing the support feet through the bottom end surface of the breakaway adapter frame;

[0010] The pressure loading module uses a hydraulic press to pressurize the break adapter frame to generate an upward thrust, pulling the simulated bolt to generate a pre-tightening force; the hydraulic press is arranged on a support base inside the break adapter frame; the hydraulic press monitors the force loading condition through a pressure gauge; a bolt pre-tightening force sensor is arranged 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 arranged 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 through a hydraulic oil pipe; a safety valve is arranged 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, so that the hydraulic press stops pressurizing.

[0013] According to the spacecraft unlocking impact measurement simulation test system of the present invention, the hydraulic press is used to apply pressure, and the preload force of the simulated bolt is controlled to reach a preset preload force value and then released, so that the simulated separation mechanism is separated from the on-board test structure; the stress wave formed by the separation of the simulated bolt is propagated 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 break 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] According to the spacecraft unlocking impact measurement simulation test system of the present invention, a data acquisition card is used to collect monitoring data of a bolt preload sensor and an acceleration impact sensor; the controller controls the pressure loading module through a loading controller according to the monitoring data transmitted by the data acquisition card, so that the preload force of the simulated bolt reaches a preset preload force value, and when the simulated separation mechanism is separated from the onboard test structure, the hydraulic press is controlled to stop pressurizing according to a trigger signal of 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 impact response during separation and unlocking of a spacecraft, and is suitable for impact testing in situations such as spacecraft load release separation and interstage separation.

[0023] The test system of 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 the investment in manpower and material resources; simplify the impact test equipment, reduce the difficulty of ground testing; improve the test efficiency, and can accurately, quickly, and repeatedly perform buffer verification tests.

[0024] The equivalent separation mechanism of the system of the present invention is simple and ingenious. The pull-off transfer structure converts the thrust provided by the hydraulic device into a pulling force (upward), and the hollow support and reverse base transmits the reaction force (downward), so that the bolt breaks. Since the power source is a separate lightweight hydraulic press with small size and light weight, the pull-off transfer structure and the support and reverse base in the simulated separation mechanism can be flexibly designed for adaptability, and the device has the characteristics of strong adaptability.

[0025] During the establishment of the system of the present invention, a finite element model of the preload release process can be established to perform numerical simulation on the impact response generated during the separation process, and an equivalent scheme can be designed with reference to the simulation results; the precision of the separation impact test simulated by the present invention is sufficiently high, and the equivalent preload loading process is stable and reliable by using an electric hydraulic device as a driving source, 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, which are lower in cost, safer, and more flexible in test form than explosive bolts, pyrotechnic nuts and other pyrotechnic products; different test requirements can be achieved by simply replacing different breaking rods, and the preload sensor makes the preload size intuitively measurable, thereby improving the test accuracy;

[0026] The experiment proves that the system of the present invention can quickly and efficiently obtain the impact suppression effect under different buffering schemes, verify the spacecraft buffering scheme, have good repeatability and high efficiency, accurately control the magnitude of the impact source released by different preload forces through pressure loading control and acquisition measurement, and can quantitatively analyze and verify the effects of various impact reduction measures at low cost, with strong engineering feasibility. It plays an important role in studying impact buffering and verifying energy absorption methods. BRIEF DESCRIPTION OF THE DRAWINGS

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

[0028] Figure 2 is a schematic diagram of a single set of inter-cabin compression and release mechanisms of a 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 locations of the simulation model;

[0031] Figure 5 yes Figure 4 The acceleration response curve of the measuring point 1 (point1);

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

[0033] Figure 7 yes Figure 4 The acceleration response curve of the measuring point 3 (point3);

[0034] Figure 8 yes Figure 4 The acceleration response curve of the measuring point 4 (point4);

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

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

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

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

[0039] Fig.13 It is a schematic diagram of the hardware structure of the spacecraft unlocking impact measurement simulation test system of the present invention;

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

[0041] Fig.15 yes Fig.13 Schematic diagram of structural dimensions;

[0042] Fig.16 yes Fig.13 A schematic diagram of the side view structure dimensions;

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

[0044] Fig.18 It is a schematic diagram of the structure of the on-board test structure;

[0045] Fig.19 It is a structural schematic diagram of the simulated separation mechanism;

[0046] Fig. 20 yes Fig.19 Side view of

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

[0048] Fig. 22 It is a schematic diagram of the on-board test structure and simulated separation mechanism after the simulated bolt breaks;

[0049] Fig.23 It is a structural schematic diagram of the simulated bolt;

[0050] Fig.24 It is a comparison diagram of acceleration impact response curve test results obtained by using the system of the present invention to conduct a simulated unlocking impact test; g represents acceleration; Curve represents curve;

[0051] Fig.25 It is a schematic diagram of the targeted energy absorption improvement design of the base;

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

[0053] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without creative work are within the scope of protection of the present invention.

[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 in conjunction with 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 load compression and release mechanism needs to be analyzed to design the corresponding equivalent ground simulation unlocking impact test device. In the ground simulation device, the preload is applied by a light and small hydraulic device, and then the metal rod is pulled off to achieve rapid unlocking with the equivalent pyrotechnic action; on the basis of retaining the original basic configuration of the cabin connection, the equivalent mechanical environment is simplified; the acceleration of each measuring point can be collected to obtain the impact response at the corresponding key position on the spacecraft; and low-cost multiple tests can be achieved by simply replacing the simulation bolts.

[0058] Firstly, based on the actual physical process of the release of the separation mechanism, the causes of the impact load and the transmission path of the shock wave 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, wherein: the lower part 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, too many components will inevitably complicate the assembly, and the pyrotechnics are expensive and cannot be reused. 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 clamping rod is the main factor in the generation of shock waves. The release of the huge preload 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 preload state of the bolts, and analyze the impact magnitude and shock wave transmission law:

[0061] 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, and to provide support for the design of the equivalent test scheme. An equivalent simplified model is established, in which the platform cabin and the base are connected by studs and nuts, and the materials are all titanium alloy, which constrains the bottom of the base. The preload release process is simulated by gradually increasing the tension of the studs to a peak value of 80kN and then suddenly reducing it to 0. The model is established as follows: Figure 3 shown.

[0062] Four equidistant acceleration response observation points are set 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. When the external load is released, the acceleration impact response of each point on the base increases instantly, with a higher amplitude and frequency. The peak values ​​of the four measuring points are 155180m / s 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] Fig. 9 and Fig.10The equivalent stress cloud diagram of the overall cross-section of the model shows that: 2ms ago, when the preload force has not yet been released, the structure has a high stress level. After 2ms, the external load is released instantly, and the strain energy in the structure is released, and the stress wave in the stud begins to propagate outward. Fig.11 and Fig.12 From the stress cloud diagram of the base section, it can be seen that the shock wave propagates to the surrounding structure, causing the overall mechanism to respond and vibrate.

[0065] From the above analysis, it can be seen that the stress wave propagation and structural resonance response formed 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 considerations. The impact acceleration measurement points of the original separation device and the equivalent experimental device can be mainly placed at the lower edge of the base, that is, connected to one end of the spacecraft platform cabin.

[0066] According to the numerical analysis results of the preload release, the separation mechanism is designed equivalently: the physical process of the release of the clamping component is simulated, and the loading and sudden release of the preload are considered in the mechanism design. The impact transmission path in the original structure is retained as the analysis object to examine whether the output result is valid:

[0067] Based on the above analysis, the equivalent test device is designed. The original base structure is used as the propagation carrier and collection object of the shock wave. The connection method between the pyrotechnic separation nut and the base is used to design the pyrotechnic separation mechanism equivalently. Finally, the overall 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 in the figure below. Fig.13 and Fig.14 shown.

[0068] 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 arranged 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;

[0069] The simulated separation mechanism adopts a fixed connection between the bottom end surface of the break 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 threadedly connected with the simulated separation nut 110; a support base 230 is arranged in the break adapter frame 210, and the support base 230 adopts a hollow design and is placed on the wedge-shaped pad 130 through the support feet passing through the bottom end surface of the break adapter frame 210;

[0070] The pressure loading module uses a hydraulic press 310 to pressurize the break adapter frame 210 to generate an upward thrust, and pulls the simulated bolt 220 to generate a pre-tightening force; the hydraulic press 310 is arranged on the support base 230 in the break adapter frame 210; the hydraulic press 310 monitors the force loading condition through the pressure gauge 320; a bolt pre-tightening force sensor 330 is arranged 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 arranged 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, a hydraulic oil pipe and a loading control device. The hydraulic press 310 is placed on a hollow support base and connected to an electric oil pump through a hydraulic oil pipe to provide loading pressure for the simulated separation mechanism; a safety valve is provided on the hydraulic oil pipe. The loading control device includes a pressure gauge, an inductive limit switch, an electric hydraulic oil pump and a safety valve, etc., which can realize the 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 release it, so that the simulated separation mechanism is separated from the on-board test structure; the stress wave formed by the separation of the simulated bolt 220 propagates toward the satellite along the base 120.

[0074] At the moment when the mechanism completes the separation action, the impact load generated can be measured by the impact sensor posted on the onboard test structure. Then the onboard test structure falls onto the cushion to trigger the inductive limit switch, and the pressure stops loading. The pressure gauge can monitor the loading status in real time to prevent overload risks. At the same time, the safety valve can realize the pressure maintenance function to increase the safety of the equipment. The system control schematic diagram is shown in the figure below. Fig.17 shown.

[0075] The on-board test structure is as follows Fig.18 As shown. In order to ensure that the impact test is the same as the actual impact transmission path on the satellite, the shape and material of the base are consistent with those on the satellite; the simulated cabin plate is used to simulate the installation plate of the platform cabin on the satellite and measure the impact response thereon; there is no energetic material in the simulated separation nut, which is fixed to the base by 8 bolts, and the upper end will be connected to the simulated bolts in the simulated separation mechanism; the clamping slider assembly on the satellite is simplified and replaced by a wedge-shaped pad placed on the base.

[0076] Simulate separation mechanism such as Fig.19 and Fig. 20 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] As an example, the simulated bolt 220 is a T-shaped breaking rod made of titanium alloy TC4, and the notch in the middle section is a brittle fracture position, which separates the simulated separation mechanism from the onboard test structure after fracture. A bolt preload sensor 330 is arranged at the bottom of the nut of the simulated bolt 220 to facilitate real-time observation 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, so that the breakaway adapter frame 210 pulls the simulated bolts, causing the T-shaped breakaway rod to undergo brittle fracture to simulate the actual separation situation. The on-board 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 FIG. Fig.21 and Fig. 22 shown. Fig.13 As shown, a steel cable 600 is also provided for suspending the break 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 task are designed, and various task requirements can be met by selecting breaking rods of different sizes and materials:

[0080] Multiple sets of repeated tests can be quickly carried out by replacing the same simulated bolts, or the impact response caused by the release of different preloads can be simulated by changing the fracture depth or material of the simulated bolts. The hydraulic press is placed on the main structure of the on-board test through the support and counter-support. While applying thrust to the break adapter frame, it also applies reaction force to the on-board test structure, eventually causing brittle fracture of the simulated bolts, and then calculating the break rod notch radius based on the tensile strength of the selected material and the actual preload size under working conditions.

[0081] Further, 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 950Mpa. When the preload force of the simulated actual working condition is 80kN, the bolt just breaks brittlely at the fracture under this stress. The radius r of the fracture is selected as:

[0085]

[0086] Simulate bolt structure such as Fig.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 on the outer side of the base 120 on the simulation cabin plate 140.

[0089] Furthermore, a data acquisition card is used to collect monitoring data of the bolt preload sensor 330 and the acceleration impact sensor 340; the controller controls the pressure loading module through the loading controller according to 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, and when the simulated separation mechanism is separated from the on-board test structure, the hydraulic press 310 is controlled to stop pressurizing according to 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 result 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 the effectiveness of the equivalent simulation experiment is determined by comparison.

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

[0094] The acceleration shock response curve obtained by simulating the unlocking shock test is as follows: Fig.24 As shown by the black solid line in the middle, the trend of its amplitude changing 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, the peak of the impact response is 10054g at 3274 Hz, and 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 pyrotechnics.

[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 to 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: Fig.25 and Fig.26 shown.

[0097] exist Fig.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; Fig.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 expandability, 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 examples of the principles and applications of the present invention. It should therefore be understood that many modifications may be made to the exemplary embodiments and that other arrangements may be devised without departing from the spirit and scope of the present invention as defined by the appended claims. It should be understood that the various dependent claims and features described herein may be combined in a manner different from that described in the original claims. It should also be understood that the features described in conjunction with a single embodiment may be used in 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 arranged 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; and adopts a simulated cabin plate (140) to simulate the on-board platform cabin mounting plate, which is fixedly connected to the bottom end of the base (120); The simulated separation mechanism adopts a bottom end surface of a breakaway adapter frame (210) and an upper end surface position of a simulated bolt (220) having a notch in the middle section to be fixedly connected correspondingly, and the lower end of the simulated bolt (220) is threadedly connected correspondingly to a simulated separation nut (110); a support base (230) is arranged in the breakaway adapter frame (210), and the support base (230) adopts a hollow design and is placed on a wedge-shaped pad (130) by passing a support foot through the bottom end surface of the breakaway adapter frame (210); The pressure loading module uses a hydraulic press (310) to pressurize a breakaway adapter frame (210) to generate an upward thrust, and pulls a 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 pre-tightening force sensor (330) is arranged at the bottom of the nut of the simulated bolt (220) to monitor the pre-tightening force generated by the simulated bolt (220); and an acceleration impact sensor (340) is arranged on the base (120) to monitor the impact load generated by the separation of the simulated separation mechanism and the on-board test structure.

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 on-board test structure and is used to buffer the impact force generated by the on-board test structure falling; the on-board test structure falls onto the buffer pad (400), triggering the inductive limit switch (350) and 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 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 release the pre-tightening force, so that the simulated separation mechanism is separated from the on-board test structure; the stress wave generated by the separation of the simulated bolt (220) propagates toward the on-board satellite along the base (120).

5. The spacecraft unlocking impact measurement simulation test system according to claim 4, characterized in that: The simulated bolt (220) is a T-shaped breaking rod made of titanium alloy TC4, wherein the notch in the middle section is a brittle fracture position, and after the fracture, the simulated separation mechanism is separated from the on-board test structure.

6. The spacecraft unlocking impact measurement simulation test system according to claim 5, characterized in that: The radius r of the notch of the simulated bolt (220) is: 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.

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

8. The spacecraft unlocking impact measurement simulation test system according to claim 7, characterized in that: A data acquisition card is used to collect monitoring data of a bolt preload force sensor (330) and an acceleration impact sensor (340); a controller controls a pressure loading module through a loading controller according to 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 according to a trigger signal of 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 maintain pressure.

9. 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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