A self-contained battery pack automatic ejection device for suborbital vehicles

Abstract

An independent battery pack automatic ejection device for suborbital aircraft belongs to the technical field of suborbital aircraft safety and solves the safety problem of battery thermal runaway in the suborbital high-speed flight environment. The device comprises a catastrophe perception and decision unit, an energy and initiation management unit and a vector ejection power mechanism. The catastrophe perception and decision unit is used for perceiving the transient pressure, characteristic gas concentration and temperature of the battery module and performing hardware logic judgment on whether thermal runaway occurs. The energy and initiation management unit provides independent thermal battery power supply, and removes the physical isolation of the initiation circuit through the mechanical and electrical double insurance safety initiator to detonate the explosion bolt. The vector ejection power mechanism is used for providing ejection power and controllable trajectory avoidance. The core purpose of the present application is to integrate the three functions of intelligent catastrophe diagnosis, independent emergency power and controllable trajectory avoidance into one, realize the safe and controllable separation of the danger source and the aircraft body, and ensure the reliable ejection action and the safe separation trajectory in the extreme environment.

Description

Technical Field

[0001] This invention relates to an automatic battery pack ejection device for suborbital vehicles, belonging to the field of suborbital vehicle safety technology. Background Technology

[0002] Suborbital space tourism vehicles typically use high-energy-density battery packs as their power source. These battery packs are susceptible to thermal runaway under extreme conditions, potentially leading to explosions or violent combustion. The development speed of these runaway events can range from milliseconds to seconds, posing a threat to the vehicle's structure and crew safety. However, existing mechanical ejection or explosive separation methods are suitable for low-speed, atmospheric-pressure environments on Earth and cannot cope with the extreme aerodynamic loads, near-vacuum environments, and severe thermal shocks encountered during high-speed suborbital flight. Furthermore, in the high-speed suborbital environment, the projectile's trajectory is subject to complex aerodynamic disturbances, making a collision with the vehicle highly likely and potentially causing secondary disasters. There is a lack of rapid and reliable diagnostic mechanisms for the rapid catastrophic events of battery explosions and combustion, making errors such as accidental or delayed ejection prone to occur. Summary of the Invention

[0003] The technical problem solved by this invention is to overcome the shortcomings of the prior art and provide an independent automatic battery pack ejection device for suborbital spacecraft. In extreme emergency situations such as battery pack explosion or fire in suborbital spacecraft, it can achieve safe, rapid, and automatic ejection, thus solving the safety problem of battery thermal runaway in suborbital high-speed flight environment.

[0004] The technical solution of this invention is: An independent battery pack automatic ejection device for suborbital vehicles includes: an embedded disaster perception and decision-making unit, an independent emergency energy and detonation management unit, a vector ejection power mechanism, and a thermo-mechanical integrated ejection cabin 100. The entire launch device is mechanically connected to the main structure of the aircraft 200 via multiple explosive bolts 301 through the mounting flange around the thermo-mechanical integrated launch chamber 100. The embedded disaster perception and decision-making unit, the independent emergency energy and detonation management unit, the vector ejection power mechanism, and the battery module are all installed inside the thermo-mechanical integrated ejection chamber. The embedded disaster perception and decision-making unit is used to sense the transient pressure, characteristic gas concentration, and temperature of the battery module, and to perform hardware logic to determine whether thermal runaway has occurred. If thermal runaway occurs, it outputs an ejection trigger signal to the independent emergency energy and detonation management unit and the vector ejection power mechanism. The independent emergency energy and detonation management unit includes an armor-protected thermal battery and an electromechanical dual-safety detonator; the thermal battery supplies power to the entire projectile device through two circuits, main and backup; upon receiving the projectile trigger signal, the electromechanical dual-safety detonator releases the physical isolation of the detonation circuit, detonates the multiple explosive bolts 301, and the mechanical connection between the entire projectile device and the aircraft is severed. The vector launch power mechanism includes a main launch group engine, an attitude control group engine, and a timing controller. After receiving the launch trigger signal, the timing controller controls the main launch group engine and the attitude control group engine to work according to a predetermined timing sequence to launch the entire launch device. The main launch group engine is used to provide the main separation impulse, and the attitude control group engine is used to impart pitch angular velocity to the launch device through thrust vector control.

[0005] Furthermore, the integrated thermal-powered projectile chamber 100 adopts a multi-layer structure design, including a carbon fiber reinforced carbonized composite material outer cover 101, a ceramic fiber heat insulation layer 102, a battery support frame 103, and a mechanically weak link 104. The outermost layer is a carbon fiber reinforced carbonized composite material outer cover 101, which is coated with an anti-oxidation coating and has a smooth transition with the aircraft's heat protection layer; the inner side of the carbon fiber reinforced carbonized composite material outer cover 101 is a ceramic fiber heat insulation layer 102 for heat insulation. The battery support frame 103 is located inside the ceramic fiber heat insulation layer 102 and has a grid-reinforced structure. It is fixedly connected to the carbon fiber reinforced carbonized composite material outer cover 101 through the mechanical weak link 104; the battery module is installed on the battery support frame 103. The inner side of the wall of the carbon fiber reinforced carbonized composite material outer cover 101 is also provided with a mechanical weak point 104; the mechanical weak point 104 automatically disintegrates when the re-entry dynamic pressure reaches the design value.

[0006] Furthermore, the carbon fiber reinforced carbonized composite material outer cover is 8mm thick and has a temperature resistance of ≥2000℃; the ceramic fiber insulation layer is 15mm thick, with a density of 0.3g / cm³ and a thermal conductivity of <0.05W / (m·K); the battery support frame is made of T700 carbon fiber composite material; and the shear strength of the mechanical weak points is designed to be 50kPa.

[0007] Furthermore, the embedded disaster perception and decision-making unit 400 includes a transient pressure sensor 401, a characteristic gas sensor 402, a distributed thermocouple cluster 403, and a hard-wired logic decision circuit 404. A transient pressure sensor 401 is positioned in the air gap between the battery modules to measure the transient pressure of the battery modules and outputs a pressure signal to the hard-wired logic decision circuit 404. A characteristic gas sensor 402 is also positioned in the air gap between the battery modules to detect the characteristic gas generated by battery thermal runaway and outputs a gas concentration signal to the hard-wired logic decision circuit 404. A distributed thermocouple cluster 403 is arranged with several measuring points between the battery modules to form a temperature monitoring network and outputs a temperature signal to the hard-wired logic decision circuit 404. The hard-wired logic decision circuit 404 compares the received pressure signal, gas concentration signal, and temperature signal with the preset pressure threshold, gas concentration threshold, and temperature threshold, respectively. When all three signals meet the decision thresholds, it outputs the ejection trigger signal.

[0008] Furthermore, the transient pressure sensor 401 adopts a piezoresistive high-frequency response sensor with a range of 0~2MPa and a response time of <1ms; The characteristic gas sensor 402 is a VOC sensor based on electrochemical principle, with a detection range of 0-2000ppm and a response time of ≤1s. The characteristic gas generated by battery thermal runaway is CO. The distributed thermocouple cluster 403 uses K-type armored thermocouples and arranges 8-12 measuring points between battery modules to form a temperature monitoring network. The spacing between measuring points is no more than 50mm, and the temperature measurement range is -50~800℃.

[0009] Furthermore, the hard-wired logic decision circuit 404 uses an FPGA chip to implement hardware logic, avoiding the delays and instabilities that may be caused by software programs; the pressure threshold is set to be no less than 1.5 times the initial pressure value, the gas concentration threshold is set to CO≥500ppm, and the temperature gradient threshold is set to the temperature difference between adjacent measuring points≥50℃ / s. When the three signals simultaneously meet the decision thresholds, the ejection trigger signal is output within ≤5ms.

[0010] Furthermore, in the independent emergency power and detonation management unit, the thermal battery uses a lithium-thionyl chloride thermal battery, which is independent of the aircraft's main power network. The electromechanical dual-safety detonator includes an electronic password verification module and a mechanical slider isolation mechanism; the electronic password verification module is used to verify the projectile trigger signal, and the electronic password verification module adopts the AES-128 encryption algorithm, a 32-bit digital password, and the verification time is <1ms. After the electronic password verification is successful, the mechanical slider isolation mechanism is activated. This mechanical slider isolation mechanism is driven by a shape memory alloy and has an action time of ≤2ms. The mechanical slider moves away, releasing the physical isolation of the detonation circuit, making the detonation circuit conductive, detonating multiple explosive bolts, and disconnecting the mechanical connection between the entire projectile device and the aircraft.

[0011] Furthermore, in the vectoring propulsion mechanism, there are four main launch group engines, evenly distributed circumferentially, which ignite simultaneously to provide the main separation impulse. The nozzles of the main launch group engines are exposed on the surface of the thermo-mechanical integrated launch chamber 100. There are two attitude control group engines, asymmetrically arranged, which ignite 50ms later than the main launch group engines and impart pitch velocity to the launch chamber through thrust vector control. The ignition sequence is precisely controlled by a timing controller with a timing control accuracy of ±1ms.

[0012] Furthermore, the specific timing of the operation of the projectile device is as follows: T=0ms: Thermal runaway occurs in the battery module, and abnormal signals are detected by the transient pressure sensor, characteristic gas sensor, and distributed thermocouple cluster. T=5ms: The hard-wired logic decision circuit confirms that the three conditions are met simultaneously and outputs a projectile trigger signal; T=6ms: The electromechanical dual-safety detonator completes electronic password verification; T=8ms: The mechanical slider isolation mechanism works, the mechanical slider moves away, and the detonation circuit is connected; T=10ms: The explosive bolt detonates, and the mechanical connection is released; T=11ms: Main projectile engine ignition; T=60ms: Attitude control group engine ignition; T=500ms: The thermo-mechanical integrated ejection chamber completely leaves the influence zone of the aircraft; T=120s: The thermo-mechanical integrated projectile chamber reaches the dynamic pressure for disintegration, the weak mechanical links are destroyed, and it is passively disintegrated.

[0013] On the other hand, the present invention also proposes an automatic ejection method for an independent battery pack of a suborbital spacecraft, comprising the following steps: S1: Disaster perception and decision-making; The battery module is installed in the thermo-mechanical integrated ejection chamber. The status of the battery module is monitored in real time through transient pressure sensors, characteristic gas sensors, and distributed thermocouple clusters. When the three collected signals of transient pressure, characteristic gas concentration, and temperature all exceed the set threshold at the same time, it is determined that the battery module has thermal runaway and an ejection trigger signal is output. S2: Safety initiation and unlocking; Electronic password verification of the ejection trigger signal to confirm the signal validity. After the electronic password verification is passed, the physical isolation of the initiation circuit is released, making the initiation circuit fully conductive, and detonating the explosive bolts connecting the thermo-mechanical integrated ejection cabin and the main body of the aircraft to achieve mechanical connection separation. S3: Vector Launch and Evasion; After the mechanical connection is separated, the vector launch power mechanism works according to the preset program: the main launch group engine is first synchronously ignited to provide the main separation impulse, and the attitude control group engine is ignited after a predetermined delay to apply pitch angular velocity to the thermo-mechanical integrated launch pod. The separation trajectory of the thermo-mechanical integrated launch pod changes from a straight separation to a curved motion with a lateral component, ensuring rapid deviation from the aircraft's trajectory; S4: Safety procedures after ejection; After the thermo-integrated launcher is ejected, it re-enters and falls independently. When the dynamic pressure of the thermo-integrated launcher reaches the design value, the mechanical weak points designed in the thermo-integrated launcher break, the thermo-integrated launcher disintegrates and is fully ablated during the fall, avoiding damage to the ground.

[0014] The advantages of this invention compared to the prior art are: (1) The device of the present invention is specifically designed to solve the safety problem of battery thermal runaway in suborbital high-speed flight environment. In extreme emergency situations such as battery explosion or fire, it can achieve safe, fast and automatic ejection; it integrates intelligent diagnosis, independent emergency power and controllable ballistic avoidance functions. (2) This invention achieves millisecond-level high-reliability diagnosis of battery explosion and fire faults through embedded pressure, gas, and temperature three-source signal sensing and hard-wired logic "AND" judgment. Independent emergency power and electromechanical dual-insurance detonation management ensure the absolute independence and safety of the projectile system under extreme conditions. Safe, collision-free separation of the battery compartment and the aircraft is achieved through timing and direction control of the vector projectile power mechanism. The controllable disintegration design of the thermo-mechanical integrated projectile compartment solves the ground safety risks associated with projectile reentry.

[0015] (3) The embedded disaster sensing and decision-making unit of this invention is a multi-source sensor array composed of transient pressure sensors, characteristic gas sensors, and distributed thermocouple clusters. The pressure surge signal, the signal of excessive concentration of volatile gas in specific electrolyte, and the signal of abnormal temperature gradient generated by the sensors are transmitted to the logic decision circuit. Through the "three AND gate" hardware logic, the ejection trigger command is output within milliseconds. The "three AND gate" logic avoids false actions caused by sensor failure or interference, and realizes reliable diagnosis of the fire disaster mode of battery explosion and ignition.

[0016] (4) The safety detonator of the present invention adopts a dual insurance of mechanical slider isolation and electronic password verification. It can only be deactivated and the detonation circuit can be connected when the physical detonation signal is received and the electronic password verification is passed at the same time, thus ensuring the absolute safety of the system in storage, transportation and flight.

[0017] (5) The thermo-mechanical integrated ejection chamber of this invention is the structural body that carries the battery pack and all the above-mentioned units. From the outside to the inside, it consists of a carbon fiber reinforced carbonized composite material outer cover, a ceramic fiber heat insulation layer, and a battery support frame. The connection between each layer adopts a mechanical weak point design with a predetermined shear strength. The thermo-mechanical integrated design enables the ejection chamber to withstand extreme aerodynamic and thermal loads and remain intact in the early stage of separation. However, when the predetermined dynamic pressure is reached during reentry, the aerodynamic load exceeds the strength of the weak point and it passively disintegrates, eliminating the risk of large debris impacting the ground. Attached Figure Description

[0018] Figure 1 This is a cross-sectional schematic diagram of an independent battery pack automatic ejection device for a suborbital spacecraft integrated into a suborbital spacecraft according to the present invention. Figure 2 This is a block diagram of the embedded disaster sensing and decision-making unit of an independent battery pack automatic ejection device for suborbital spacecraft according to the present invention. Figure 3 This is a schematic diagram of the working timing and trajectory of the vector launch power mechanism of an independent battery pack automatic launch device for a suborbital spacecraft according to the present invention; Figure 4 This is an exploded view of the gradient protection and controllable disassembly structure of the thermo-mechanical integrated ejection chamber of an independent battery pack automatic ejection device for suborbital spacecraft according to the present invention. Detailed Implementation

[0019] The specific embodiments of the present invention will now be described in further detail with reference to the accompanying drawings.

[0020] This invention provides an independent battery pack automatic ejection device for suborbital vehicles. Its core purpose is to integrate three major functions: intelligent disaster diagnosis, independent emergency power, and controllable trajectory avoidance, so as to achieve safe and controllable separation of the hazard source from the vehicle body and ensure reliable ejection action and safe separation trajectory in extreme environments.

[0021] like Figure 1 As shown, the present invention proposes an independent automatic battery pack ejection device for suborbital spacecraft, comprising: an embedded disaster perception and decision-making unit, an independent emergency energy and detonation management unit, a vector ejection power mechanism, and a thermo-mechanical integrated ejection cabin 100. The entire launch device is mechanically connected to the main structure of the aircraft 200 via multiple explosive bolts 301 through the mounting flange around the thermo-mechanical integrated launch chamber 100. The embedded disaster perception and decision-making unit, the independent emergency energy and detonation management unit, the vector ejection power mechanism, and the battery module are all installed inside the thermo-mechanical integrated ejection chamber. The embedded disaster perception and decision-making unit is used to sense the transient pressure, characteristic gas concentration, and temperature of the battery module, and to perform hardware logic to determine whether thermal runaway has occurred. If thermal runaway occurs, it outputs an ejection trigger signal to the independent emergency energy and detonation management unit and the vector ejection power mechanism. The independent emergency energy and detonation management unit includes an armor-protected thermal battery and an electromechanical dual-safety detonator; the thermal battery supplies power to the entire projectile device through two circuits, main and backup; upon receiving the projectile trigger signal, the electromechanical dual-safety detonator releases the physical isolation of the detonation circuit, detonates the multiple explosive bolts 301, and the mechanical connection between the entire projectile device and the aircraft is severed. The vector launch power mechanism includes a main launch group engine, an attitude control group engine, and a timing controller. After receiving the launch trigger signal, the timing controller controls the main launch group engine and the attitude control group engine to work according to a predetermined timing sequence to launch the entire launch device. The main launch group engine is used to provide the main separation impulse, and the attitude control group engine is used to impart pitch angular velocity to the launch device through thrust vector control.

[0022] like Figure 4 As shown, the thermo-mechanical integrated projectile chamber 100 adopts a multi-layer structure design, including a carbon fiber reinforced carbonized composite material outer cover 101, a ceramic fiber heat insulation layer 102, a battery support frame 103, and a mechanical weak point 104. The outermost layer is a carbon fiber reinforced carbonized composite material outer cover 101, which is coated with an anti-oxidation coating and has a smooth transition with the aircraft's heat protection layer; the inner side of the carbon fiber reinforced carbonized composite material outer cover 101 is a ceramic fiber heat insulation layer 102 for heat insulation. The battery support frame 103 is located inside the ceramic fiber heat insulation layer 102 and has a grid-reinforced structure. It is fixedly connected to the carbon fiber reinforced carbonized composite material outer cover 101 through the mechanical weak link 104; the battery module is installed on the battery support frame 103. The inner side of the wall of the carbon fiber reinforced carbonized composite material outer cover 101 is also provided with a mechanical weak point 104; the mechanical weak point 104 automatically disintegrates when the re-entry dynamic pressure reaches the design value.

[0023] Preferably, the carbon fiber reinforced carbonized composite material outer cover is 8mm thick and has a temperature resistance of ≥2000℃; the ceramic fiber insulation layer is 15mm thick, with a density of 0.3g / cm³ and a thermal conductivity of <0.05W / (m·K); the battery support frame is made of T700 carbon fiber composite material; and the shear strength of the mechanical weak point is designed to be 50kPa.

[0024] like Figure 2As shown, the embedded disaster perception and decision-making unit 400 includes a transient pressure sensor 401, a characteristic gas sensor 402, a distributed thermocouple cluster 403, and a hard-wired logic decision circuit 404. A transient pressure sensor 401 is positioned in the air gap between the battery modules to measure the transient pressure of the battery modules and outputs a pressure signal to the hard-wired logic decision circuit 404. A characteristic gas sensor 402 is also positioned in the air gap between the battery modules to detect the characteristic gas generated by battery thermal runaway and outputs a gas concentration signal to the hard-wired logic decision circuit 404. A distributed thermocouple cluster 403 is arranged with several measuring points between the battery modules to form a temperature monitoring network and outputs a temperature signal to the hard-wired logic decision circuit 404. The hard-wired logic decision circuit 404 compares the received pressure signal, gas concentration signal, and temperature signal with the preset pressure threshold, gas concentration threshold, and temperature threshold, respectively. When all three signals meet the decision thresholds, it outputs the ejection trigger signal.

[0025] Preferably, the transient pressure sensor 401 is a piezoresistive high-frequency response sensor with a range of 0~2MPa and a response time of <1ms; Preferably, the characteristic gas sensor 402 is a VOC sensor based on electrochemical principle, with a detection range of 0-2000ppm, a response time of ≤1s, and the characteristic gas generated by battery thermal runaway is CO; Preferably, the distributed thermocouple cluster 403 uses K-type armored thermocouples, with 8-12 measuring points arranged between battery modules to form a temperature monitoring network. The spacing between the measuring points is no more than 50mm, and the temperature measurement range is -50~800℃.

[0026] In this invention, the hard-wired logic decision circuit 404 uses an FPGA chip to implement hardware logic, avoiding the delay and instability that may be caused by software programs; the pressure threshold is set to be no less than 1.5 times the initial pressure value, the gas concentration threshold is set to CO≥500ppm, and the temperature gradient threshold is set to the temperature difference between adjacent measuring points≥50℃ / s. When the three signals simultaneously meet the decision thresholds, the ejection trigger signal is output within ≤5ms.

[0027] In this invention, the thermal battery in the independent emergency energy and detonation management unit is a lithium-thionyl chloride thermal battery, which is independent of the aircraft's main power network. The electromechanical dual-safety detonator includes an electronic password verification module and a mechanical slider isolation mechanism; the electronic password verification module is used to verify the projectile trigger signal, and the electronic password verification module adopts the AES-128 encryption algorithm, a 32-bit digital password, and the verification time is <1ms. Specifically, the electronic password verification module employs a dynamic voting verification strategy based on flight state adaptation. The module receives a launch trigger signal from the logic decision circuit 404, which encapsulates a dynamic checksum generated in real-time based on the current flight timing code and the device's unique identification code. The module uses clock synchronization information to decrypt and match the dynamic checksum. Verification succeeds when the signal generation timing is logically consistent with the physical state changes of the battery compartment. This design aims to resist signal interference or malicious injection in the complex electromagnetic environment of suborbital flight, ensuring the uniqueness and logical rationality of emergency commands.

[0028] After the electronic password verification is successful, the mechanical slider isolation mechanism is activated. This mechanical slider isolation mechanism is driven by a shape memory alloy and has an action time of ≤2ms. The mechanical slider moves away, releasing the physical isolation of the detonation circuit, making the detonation circuit conductive, detonating multiple explosive bolts, and disconnecting the mechanical connection between the entire projectile device and the aircraft.

[0029] In this invention, the vector launch power mechanism consists of four main launch group engines, evenly distributed circumferentially, which ignite simultaneously to provide the main separation impulse. The nozzles of the main launch group engines are exposed on the surface of the thermo-mechanical integrated launch chamber 100. The attitude control group consists of two engines, asymmetrically arranged, which ignite 50ms later than the main launch group engines and impart pitch velocity to the launch chamber through thrust vector control. The ignition sequence is precisely controlled by a timing controller with a timing control accuracy of ±1ms.

[0030] like Figure 3 As shown, the specific timing of the projectile's operation is as follows: T=0ms: Thermal runaway occurs in the battery module, and abnormal signals are detected by the transient pressure sensor, characteristic gas sensor, and distributed thermocouple cluster. T=5ms: The hard-wired logic decision circuit confirms that the three conditions are met simultaneously and outputs a projectile trigger signal; T=6ms: The electromechanical dual-safety detonator completes electronic password verification; T=8ms: The mechanical slider isolation mechanism works, the mechanical slider moves away, and the detonation circuit is connected; T=10ms: The explosive bolt detonates, and the mechanical connection is released; T=11ms: Main projectile engine ignition; T=60ms: Attitude control group engine ignition; T=500ms: The thermo-mechanical integrated ejection chamber completely leaves the influence zone of the aircraft; T=120s: The thermo-mechanical integrated projectile chamber reaches the dynamic pressure for disintegration, the weak mechanical links are destroyed, and it is passively disintegrated.

[0031] Example 1: This invention discloses a stand-alone automatic battery pack ejection device for suborbital vehicles, installed as a pre-integrated modular system in the battery compartment of the suborbital vehicle. The entire device is mechanically connected to the main structure 200 of the vehicle via mounting flanges around the thermo-mechanically integrated ejection chamber 100 and multiple explosive bolts 301. Figure 1 As shown. Necessary signal monitoring lines are maintained between the launch device and the main body of the aircraft, and an isolation design is employed. The outer surface of the launch device smoothly transitions into the aircraft's heat shield, maintaining the integrity of its aerodynamic shape.

[0032] The automatic launch device includes: an embedded disaster perception and decision-making unit, an independent emergency energy and detonation management unit, a vector launch power mechanism, and a thermo-mechanical integrated launch chamber 100.

[0033] like Figure 2 As shown, the embedded disaster perception and decision-making unit 400 includes a transient pressure sensor 401, a characteristic gas sensor 402, a distributed thermocouple cluster 403, and a hard-wired logic decision circuit 404.

[0034] The transient pressure sensor 401 is a piezoresistive high-frequency response sensor with a range of 0~2MPa and a response time of <1ms. It is placed in the critical air gap between battery modules.

[0035] The characteristic gas sensor 402 is a VOC sensor based on electrochemical principle, with a detection range of 0-2000ppm and a response time of ≤1s. It detects characteristic gases (such as CO, HF, CH4, etc.) generated by battery thermal runaway. This embodiment focuses on CO, the most representative characteristic gas.

[0036] The distributed thermocouple cluster 403 uses K-type armored thermocouples and arranges 8-12 measuring points between battery modules to form a temperature monitoring network. The spacing between measuring points is no more than 50mm, and the temperature measurement range is -50~800℃.

[0037] The hard-wired logic decision circuit 404 uses an FPGA chip to implement the hardware logic, avoiding the latency and instability that may be caused by software programs. The pressure threshold is set to no less than 150 kPa (approximately 1.5 times the initial pressure value), the gas concentration threshold is set to CO ≥ 500 ppm, and the temperature gradient threshold is set to the temperature difference between adjacent measuring points ≥ 50℃ / s. When the three signals simultaneously meet the decision thresholds, the FPGA outputs a projectile trigger signal within ≤ 5 ms.

[0038] The embedded disaster sensing and decision-making unit of this invention is a multi-source sensor array composed of transient pressure sensors, characteristic gas sensors, and distributed thermocouple clusters. The sensors generate signals such as sudden pressure increases, excessive concentrations of specific electrolyte volatile gases, and abnormal temperatures, which are transmitted to the logic decision circuit. Through a three-AND gate hardware logic, a launch trigger command is output within milliseconds. The three-AND gate logic avoids malfunctions caused by sensor failures or interference, enabling reliable diagnosis of battery explosion and fire disaster modes.

[0039] The independent emergency power and detonation management unit is designed to be completely independent, including an armored thermal battery and an electromechanical dual-safety detonator.

[0040] The armor-protected thermal battery is a lithium-thionyl chloride thermal battery with a nominal voltage of 28V and a capacity of 20Ah. The battery is externally protected by a 3mm thick titanium alloy shell, internally filled with silica aerogel insulation material. The entire projectile system is powered by two sets of power management circuits, one main and one backup.

[0041] The electromechanical dual-safety initiator includes an electronic password verification module and a mechanical slider isolation mechanism. The electronic password verification module uses the AES-128 encryption algorithm, a 32-bit digital password, and a verification time of <1ms. The mechanical slider isolation mechanism is driven by a shape memory alloy, with an actuation time of ≤2ms, providing physical isolation of the initiation circuit under normal conditions. The initiation circuit includes two sets of initiation lines: a main line and a backup line.

[0042] The mechanical slider isolation mechanism consists of a slider, a groove, isolation contacts, and a shape memory alloy drive unit. Under normal conditions, the slider is in the "safe" position under the action of the shape memory alloy drive unit. The isolation contacts on the slider physically isolate the critical path of the detonation circuit. After successful electronic password verification, the controller sends a command to the shape memory alloy drive unit, energizing it. The drive unit then actuates, generating linear displacement that pulls or pushes the slider, moving it along the groove from the "safe" position to the "ready to detonate" position. Once the slider reaches the "ready to detonate" position, the isolation contacts on the slider precisely engage with the corresponding contacts fixed to the mechanism, physically connecting the detonation circuit and awaiting the detonation command.

[0043] The vectoring propulsion mechanism consists of several small solid rocket motors arranged asymmetrically in the circumferential direction. These motors are divided into a main launch group and an attitude control group. First, the main launch group ignites synchronously, providing the primary impulse needed to propel the battery compartment away from the spacecraft. Then, following a preset program and a delay of several milliseconds, the attitude control group ignites, using its thrust vectoring to impart a precise pitch or yaw rate to the launch compartment. This ensures that after separation from the spacecraft, it follows an evasive trajectory with the maximum deviation from the spacecraft's orbit, guaranteeing its rapid departure from the spacecraft.

[0044] Specifically, the vectoring propulsion mechanism includes the main launch group engine, the attitude control group engine, and the timing controller.

[0045] The main projectile launcher group has four engines, evenly distributed circumferentially, each with a thrust of 5 kN and an operating time of 0.5 s. The attitude control group has two engines, asymmetrically arranged, each with a thrust of 2 kN and an operating time of 0.3 s.

[0046] The main ejector group engines ignite simultaneously, providing the main separation impulse.

[0047] The attitude control group's engine ignites after a 50ms delay, and the projectile is given pitch angular velocity through thrust vector control.

[0048] The ignition timing is precisely controlled by a dedicated timing controller with a timing control accuracy of ±1ms.

[0049] In this invention, the thermo-mechanical integrated ejection chamber is the structural body that carries the battery pack and all the aforementioned units. From the outside to the inside, it consists of a carbon fiber reinforced carbonized composite material outer casing, a ceramic fiber thermal insulation layer, and a battery support frame. The connections between each layer employ a mechanically weak point design with a predetermined shear strength. The thermo-mechanical integrated design enables the ejection chamber to withstand extreme aerodynamic and thermal loads and remain intact during the initial separation phase. However, during reentry, when the predetermined dynamic pressure is reached, the aerodynamic load exceeds the strength of the weak points, causing passive disintegration and eliminating the risk of large debris impacting the ground.

[0050] Specifically, the thermal-power integrated projectile chamber 100 adopts a multi-layer structure design, including a carbon fiber reinforced carbonized composite material outer cover 101, a ceramic fiber heat insulation layer 102, a battery support frame 103, and a mechanical weak point 104.

[0051] The carbon fiber reinforced carbonized composite material outer cover is 8mm thick, coated with an anti-oxidation coating, and has a temperature resistance of ≥2000℃, with a smooth transition to the aircraft's heat protection layer.

[0052] The ceramic fiber insulation layer 102 has a thickness of 15mm, a density of 0.3g / cm³, and a thermal conductivity of <0.05W / (m·K) (@1000℃).

[0053] The battery support frame 103 is made of T700 carbon fiber composite material with a grid-reinforced structure and is connected to the external structure through the mechanical weak link 104.

[0054] The shear strength of the mechanically weak link 104 is designed to be 50 kPa, and it will automatically disintegrate when the re-entry dynamic pressure reaches the design value.

[0055] Example 2: This invention also proposes an automatic ejection method for an independent battery pack of a suborbital spacecraft, comprising the following steps: S1: Disaster perception and decision-making; The battery module is installed in the thermo-mechanical integrated ejection chamber. The status of the battery module is monitored in real time through transient pressure sensors, characteristic gas sensors, and distributed thermocouple clusters. When the three collected signals of transient pressure, characteristic gas concentration, and temperature all exceed the set threshold at the same time, it is determined that the battery module has thermal runaway and an ejection trigger signal is output. S2: Safety initiation and unlocking; Electronic password verification of the ejection trigger signal to confirm the signal validity. After the electronic password verification is passed, the physical isolation of the initiation circuit is released, making the initiation circuit fully conductive, and detonating the explosive bolts connecting the thermo-mechanical integrated ejection cabin and the main body of the aircraft to achieve mechanical connection separation. S3: Vector Launch and Evasion; After the mechanical connection is separated, the vector launch power mechanism works according to the preset program: the main launch group engine is first synchronously ignited to provide the main separation impulse, and the attitude control group engine is ignited after a predetermined delay to apply pitch angular velocity to the thermo-mechanical integrated launch pod. The separation trajectory of the thermo-mechanical integrated launch pod changes from a straight separation to a curved motion with a lateral component, ensuring rapid deviation from the aircraft's trajectory; S4: Safety procedures after ejection; After the thermo-integrated launcher is ejected, it re-enters and falls independently. When the dynamic pressure of the thermo-integrated launcher reaches the design value, the mechanical weak points designed in the thermo-integrated launcher break, the thermo-integrated launcher disintegrates and is fully ablated during the fall, avoiding damage to the ground.

[0056] The parts of this invention not described in detail are common knowledge to those skilled in the art.

Claims

1. An independent automatic battery pack ejection device for suborbital vehicles, characterized in that, include: Embedded disaster perception and decision-making unit, independent emergency energy and detonation management unit, vector ejection power mechanism and thermo-mechanical integrated ejection chamber (100). The entire launch device is mechanically connected to the main structure of the aircraft (200) through multiple explosive bolts (301) via the mounting flange around the thermo-mechanical integrated launch chamber (100). The embedded disaster perception and decision-making unit, the independent emergency energy and detonation management unit, the vector ejection power mechanism, and the battery module are all installed inside the thermo-mechanical integrated ejection chamber. The embedded disaster perception and decision-making unit is used to sense the transient pressure, characteristic gas concentration, and temperature of the battery module, and to perform hardware logic to determine whether thermal runaway has occurred. If thermal runaway occurs, it outputs an ejection trigger signal to the independent emergency energy and detonation management unit and the vector ejection power mechanism. The independent emergency energy and detonation management unit includes an armor-protected thermal battery and an electromechanical dual-safety detonator. The electromechanical dual-safety detonator includes an electronic password verification module and a mechanical slider isolation mechanism. After the electronic password verification is passed, the mechanical slider isolation mechanism is activated. The thermal battery supplies power to the entire projectile device through two circuits: a main circuit and a backup circuit. After receiving the projectile trigger signal, the electromechanical dual-safety detonator releases the physical isolation of the detonation circuit and detonates the multiple explosive bolts (301), thus disconnecting the mechanical connection between the entire projectile device and the aircraft. The vector launcher power mechanism includes a main launcher engine, an attitude control engine, and a timing controller. Upon receiving the launch trigger signal, the timing controller controls the main launcher engine and the attitude control engine to operate according to a predetermined timing sequence, launching the entire launcher. The main launcher engine provides the primary separation impulse, while the attitude control engine imparts a pitch velocity to the launcher via thrust vector control. The attitude control engine ignites with a delay relative to the main launcher engine, imparting a pitch velocity to the launcher via thrust vector control. The ignition timing is precisely controlled by the timing controller. The thermo-mechanical integrated projectile chamber (100) adopts a multi-layer structure design, including a carbon fiber reinforced carbonized composite material outer cover (101), a ceramic fiber heat insulation layer (102), a battery support frame (103), and a mechanical weak point (104). The outermost layer is a carbon fiber reinforced carbonized composite material outer cover (101), which is coated with an anti-oxidation coating and has a smooth transition with the aircraft's heat protection layer; the inner side of the carbon fiber reinforced carbonized composite material outer cover (101) is a ceramic fiber heat insulation layer (102) for heat insulation. The battery support frame (103) is located inside the ceramic fiber heat insulation layer (102) and is a grid-reinforced structure. It is fixedly connected to the carbon fiber reinforced carbonized composite material outer cover (101) through the mechanical weak link (104); the battery module is installed on the battery support frame (103). The inner side of the wall of the carbon fiber reinforced carbonized composite material outer cover (101) is also provided with a mechanical weak link (104); the mechanical weak link (104) automatically disintegrates when the re-entry dynamic pressure reaches the design value.

2. The independent automatic battery pack ejection device for suborbital spacecraft according to claim 1, characterized in that: The carbon fiber reinforced carbonized composite material outer cover is 8mm thick and has a temperature resistance of ≥2000℃; the ceramic fiber insulation layer is 15mm thick, with a density of 0.3g / cm³ and a thermal conductivity of <0.05W / (m·K); the battery support frame is made of T700 carbon fiber composite material; the shear strength of the mechanical weak points is designed to be 50kPa.

3. The independent automatic battery pack ejection device for suborbital spacecraft according to claim 1, characterized in that: The embedded disaster perception and decision unit (400) includes a transient pressure sensor (401), a characteristic gas sensor (402), a distributed thermocouple cluster (403), and a hard-wired logic decision circuit (404). A transient pressure sensor (401) is arranged in the air gap between the battery modules to measure the transient pressure of the battery modules and outputs a pressure signal to the hard-wired logic decision circuit (404); a characteristic gas sensor (402) is also arranged in the air gap between the battery modules to detect the characteristic gas generated by the thermal runaway of the battery and outputs a gas concentration signal to the hard-wired logic decision circuit (404); a distributed thermocouple cluster (403) is arranged with several measuring points between the battery modules to form a temperature monitoring network and outputs a temperature signal to the hard-wired logic decision circuit (404). The hard-wired logic decision circuit (404) compares the received pressure signal, gas concentration signal, and temperature signal with the preset pressure threshold, gas concentration threshold, and temperature threshold, respectively. When the three signals simultaneously meet the decision threshold, it outputs the ejection trigger signal. The three signals refer to the pressure signal, gas concentration signal, and temperature signal.

4. The independent automatic battery pack ejection device for suborbital spacecraft according to claim 3, characterized in that: The transient pressure sensor (401) is a piezoresistive high-frequency response sensor with a range of 0~2MPa and a response time of <1ms; The characteristic gas sensor (402) is an electrochemical VOC sensor with a detection range of 0-2000ppm and a response time of ≤1s. The characteristic gas generated by the thermal runaway of the battery is CO. The distributed thermocouple cluster (403) uses K-type armored thermocouples and arranges 8-12 measuring points between battery modules to form a temperature monitoring network. The distance between measuring points is no more than 50mm, and the temperature measurement range is -50~800℃.

5. The independent automatic battery pack ejection device for a suborbital spacecraft according to claim 4, characterized in that: The hard-wired logic decision circuit (404) uses an FPGA chip to implement hardware logic, avoiding the delay and instability that may be caused by software programs; the pressure threshold is set to be no less than 1.5 times the initial pressure value, the gas concentration threshold is set to CO≥500ppm, and the temperature gradient threshold is set to the temperature difference between adjacent measuring points≥50℃ / s. When the three signals simultaneously meet the decision threshold, the ejection trigger signal is output within ≤5ms.

6. The independent automatic battery pack ejection device for suborbital spacecraft according to claim 1, characterized in that: In the independent emergency power and detonation management unit, the thermal battery is a lithium-thionyl chloride thermal battery, which is independent of the aircraft's main power network. The electronic password verification module in the electromechanical dual-safety detonator is used to verify the projectile trigger signal. The electronic password verification module adopts the AES-128 encryption algorithm, a 32-bit digital password, and the verification time is <1ms. The mechanical slider isolation mechanism is driven by shape memory alloy with an action time of ≤2ms. When the mechanical slider moves away, it releases the physical isolation of the detonation circuit, making the detonation circuit conductive and detonating multiple explosive bolts. The mechanical connection between the entire projectile device and the aircraft is then severed.

7. The independent automatic battery pack ejection device for a suborbital spacecraft according to claim 1, characterized in that: In the vector launcher power mechanism, there are 4 main launcher engines, which are evenly distributed circumferentially and ignite simultaneously to provide the main separation impulse. The nozzles of the main launcher engines are exposed on the surface of the thermo-mechanical integrated launcher body (100). There are 2 attitude control engines, which are asymmetrically arranged and ignite 50ms later than the main launcher engines. The timing control accuracy of the ignition sequence is ±1ms.

8. The independent automatic battery pack ejection device for a suborbital spacecraft according to claim 7, characterized in that: The specific timing sequence of the operation of the projectile device is as follows: T=0ms: Thermal runaway occurs in the battery module, and abnormal signals are detected by the transient pressure sensor, characteristic gas sensor, and distributed thermocouple cluster. T=5ms: The hard-wired logic decision circuit confirms that the three conditions are met simultaneously and outputs a projectile trigger signal; T=6ms: The electromechanical dual-safety detonator completes electronic password verification; T=8ms: The mechanical slider isolation mechanism works, the mechanical slider moves away, and the detonation circuit is connected; T=10ms: The explosive bolt detonates, and the mechanical connection is released; T=11ms: Main projectile engine ignition; T=60ms: Attitude control group engine ignition; T=500ms: The thermo-mechanical integrated ejection chamber completely leaves the influence zone of the aircraft; T=120s: The thermo-mechanical integrated projectile chamber reaches the dynamic pressure for disintegration, the weak mechanical links are destroyed, and it is passively disintegrated.

9. A method for automatically launching a suborbital battery pack based on the automatic battery pack launching device for suborbital vehicles as described in claim 1, characterized in that, include: S1: Disaster perception and decision-making; The battery module is installed in the thermo-mechanical integrated ejection chamber. The status of the battery module is monitored in real time through transient pressure sensors, characteristic gas sensors, and distributed thermocouple clusters. When the three collected signals of transient pressure, characteristic gas concentration, and temperature all exceed the set threshold at the same time, it is determined that the battery module has thermal runaway and an ejection trigger signal is output. S2: Safety initiation and unlocking; Electronic password verification of the ejection trigger signal to confirm the signal validity. After the electronic password verification is passed, the physical isolation of the initiation circuit is released, making the initiation circuit fully conductive, and detonating the explosive bolts connecting the thermo-mechanical integrated ejection cabin and the main body of the aircraft to achieve mechanical connection separation. S3: Vector Launch and Evasion; After the mechanical connection is separated, the vector launch power mechanism works according to the preset program: the main launch group engine is first synchronously ignited to provide the main separation impulse, and the attitude control group engine is ignited after a predetermined delay to apply pitch angular velocity to the thermo-mechanical integrated launch pod. The separation trajectory of the thermo-mechanical integrated launch pod changes from a straight separation to a curved motion with a lateral component, ensuring rapid deviation from the aircraft's trajectory; S4: Safety procedures after disembarkation; After being ejected, the thermo-integrated projectile hull re-enters and falls independently. When the dynamic pressure of the thermo-integrated projectile hull reaches the design value, the mechanical weak points designed in the thermo-integrated projectile hull break, the thermo-integrated projectile hull disintegrates and is fully ablated during the fall, avoiding damage to the ground.

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