Return type sub-orbital load cabin

Abstract

The invention provides a return type sub-orbital load cabin. The load cabin comprises an end cap, a structural conical section, a pneumatic arc bottom, an effective load system, a measuring system and a recovery system. The end cap is arranged at the top of the structural conical section; the pneumatic arc bottom is arranged at the bottom of the structural conical section; the effective load system is arranged in the end cap or the structural cone section; the measuring system is arranged in the structural cone section; and the recovery system is arranged in the pneumatic arc bottom. According to the invention, the test cost is reduced, safe recovery can be realized, and long-time microgravity experiment and in-situ detection are supported. The return type sub-orbital load cabin serves as a flying platform capable of providing a microgravity experiment environment and realizing safe recovery.

Description

Technical Field

[0001] This application relates to the field of aerospace technology, and in particular to a reusable suborbital payload capsule. Background Technology

[0002] Suborbital flight refers to a spacecraft entering space at an altitude exceeding the 100km KAMAN line, but at a speed below the first cosmic velocity, thus failing to achieve orbital flight around the Earth.

[0003] Currently, the main experimental platforms capable of creating microgravity environments include the following: 1) Sounding rocket payload compartment: High level of microgravity ( The duration is approximately 180–200 seconds, and the test space is large, but traditional payload chambers are usually non-recoverable, resulting in high test costs. 2) Zero-gravity aircraft: low microgravity level ( The duration of a single microgravity event is approximately 20 seconds, which is easily affected by airflow and results in high experimental costs. 3) Orbiting vehicles (such as space stations and satellites): have a better microgravity environment ( (It has a long duration, but is significantly affected by factors such as attitude control, has a complex system, and has extremely high operating and recovery costs.) 4) Microgravity drop tower: Microgravity level is high ( The time is extremely short (usually <10 seconds), and it is only suitable for small-scale model experiments.

[0004] Currently, in terms of suborbital recovery technology, Blue Origin's New Shepard spacecraft uses a combination of retro-rocket engines and parachutes for recovery, but its system is complex, costly, and lacks in-situ near-space detection capabilities; MIT's microgravity experimental platform achieves lightweight design, but its microgravity duration is short and its functionality is limited; Japan's JAXA's BOV-7A payload bay relies on a balloon-borne aerostat, which is too low to achieve active recovery and suborbital microgravity experiments.

[0005] Furthermore, traditional reentry and recovery technologies are not suitable for suborbital low-speed parachute deployment conditions. For example, cluster parachute systems (such as NASA Orion) or controllable wing parachutes (such as the European IXV) suffer from low reliability, system complexity, and large landing point dispersion under suborbital low-speed parachute deployment conditions.

[0006] Therefore, there is a lack of a low-cost, reusable suborbital payload capsule that can support long-term microgravity experiments and in-situ exploration in the existing technology. Summary of the Invention

[0007] The purpose of this application is to provide a recoverable suborbital payload capsule that reduces experimental costs, enables safe recovery, and supports long-term microgravity experiments and in-situ exploration. As a flight platform capable of providing a microgravity experimental environment and achieving safe recovery, the recoverable suborbital payload capsule has significant value in space science research.

[0008] To achieve the above objectives, this application provides a reusable suborbital payload capsule, which includes: a nose cone, a structural cone section, an aerodynamic arc bottom, a payload system, a measurement system, and a recovery system; the nose cone is disposed at the top of the structural cone section; the aerodynamic arc bottom is disposed at the bottom of the structural cone section; the payload system is disposed within the nose cone or the structural cone section; the measurement system is disposed within the structural cone section; and the recovery system is disposed within the aerodynamic arc bottom.

[0009] The reentry suborbital payload capsule described above further includes: a hatch, an operating port cover, a heat shield, and an impact damping device; the hatch, the heat shield, and the impact damping device are disposed on the structural cone section; the operating port cover is disposed on the end cap and the impact damping device.

[0010] As described above, in the reentrant suborbital payload capsule, the inner wall of the structural cone section is connected to a payload mounting disk; the payload mounting disk is an aluminum alloy skeleton with an aluminum honeycomb core and a carbon fiber panel.

[0011] As described above, in the reusable suborbital payload bay, the measurement system includes: a power module, a radio frequency front-end, a central program controller, a GNSS telemetry antenna, a cable network system, and / or an external measurement system; the power module provides power to each device; the radio frequency front-end and the GNSS telemetry antenna provide data transmission functionality for the payload bay; the central program controller acts as the brain of the payload bay, controlling the orderly operation of each device and also functions as a payload manager; the cable network system connects the various devices.

[0012] The reentry suborbital payload capsule described above, wherein the external measurement system includes: an antenna, a camera device, multiple sensors, and multiple transducers.

[0013] As described above, the reentry-type suborbital payload capsule includes multiple sensors: an in-situ measurement sensor, a wall temperature sensor, a noise sensor, a triaxial impact sensor, a triaxial acceleration sensor, a low-frequency vibration sensor, and a high-frequency vibration sensor. The in-situ measurement sensor is used for in-situ measurements. The wall temperature sensor is used to measure the temperature of the inner wall of the payload capsule. The noise sensor is used to measure noise within the capsule. The triaxial impact sensor is used to measure impacts in three axes within the payload capsule. The triaxial acceleration sensor is used to measure acceleration within the capsule. The low-frequency vibration sensor and the high-frequency vibration sensor are used to measure vibrations on the payload mounting surface within the capsule.

[0014] The reentry suborbital payload capsule described above, wherein the payload system includes: a first payload, a second payload, and multiple standard unit modules.

[0015] The reusable suborbital payload capsule described above, wherein the recovery system includes: a parachute recovery system component and a parachute recovery controller; the parachute recovery controller integrates sensing devices.

[0016] The reentry suborbital payload capsule described above, wherein the parachute recovery system components include: a parachute capsule assembly, a deceleration parachute, a main parachute, an ejector, and a pin puller.

[0017] As described above, the reentrant suborbital payload capsule has an arc-shaped aerodynamic base and is made of polyimide-based glass fiber composite material.

[0018] The beneficial effects achieved by this application are as follows: (1) This application enables the safe recovery and partial reuse of the reusable suborbital payload capsule, significantly reducing the cost of a single mission, achieving high landing accuracy, eliminating the need for large-scale search and rescue, rapidly recovering experimental data, supporting long-term microgravity experiments and in-situ exploration, and providing a low-cost, high-frequency experimental platform for microgravity science, space exploration and technology verification.

[0019] (2) This application integrates the parachute compartment recovery system components into the aerodynamic base to improve space utilization.

[0020] (3) The structural cone section of this application adopts a reusable design and uses a heat protection system to achieve effective heat insulation, and the impact environment is controlled below 1000g. Attached Figure Description

[0021] To more clearly illustrate the technical solutions in the embodiments of this application or the prior art, the drawings used in the description of the embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are only some embodiments recorded in this application. For those skilled in the art, other drawings can be obtained based on these drawings.

[0022] Figure 1 This is a front view of a reentry suborbital payload capsule according to an embodiment of this application.

[0023] Figure 2 for Figure 1 Schematic diagram of the cross section at point CC.

[0024] Figure 3 for Figure 2 Sectional view at point BB.

[0025] Figure 4 for Figure 2 Sectional view at point AA.

[0026] Reference numerals: 1-End cap; 2-Structural cone section; 3-Aerodynamic arc bottom; 4-Parachute recovery system component; 5-Parachute recovery controller; 6-Impact damping device; 11-Operating hatch cover; 21-Hatch door; 22-Heat shield; 23-Separation and disengagement assembly; 24-First load mounting disk; 25-Second load mounting disk; 26-Reinforcing tie rod; 27-Standard unit module; 41-Power module; 42-RF front end; 43-Central program controller; 51-Triaxial impact sensor; 52-Triaxial impact sensor Transducer; 53-Triaxial accelerometer; 54-High-frequency vibration sensor; 55-Low-frequency vibration sensor; 56-Noise transducer; 57-Noise sensor; 58-Air temperature transducer; 59-Air temperature sensor; 60-Wall temperature transducer; 61-First payload; 62-Second payload; 71-First pressure sensor; 72-Second pressure sensor; 73-GNSS telemetry antenna; 74-Cable network system test interface; 75-External measurement system; 79-Camera device. Detailed Implementation

[0027] The technical solutions of the embodiments of this application will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. All other embodiments obtained by those skilled in the art based on the embodiments of this application without creative effort are within the scope of protection of this application.

[0028] like Figure 1-4As shown, this application provides a reusable suborbital payload capsule, which includes: a nose cone 1, a structural cone section 2, an aerodynamic arc base 3, a payload system, a measurement system, and a recovery system. The nose cone 1 is located at the top of the structural cone section 2 and is used to adjust the axial center of gravity of the entire capsule, bearing the main impact load during landing. The structural cone section 2 serves as the main structure of the payload capsule, providing not only the necessary internal space to accommodate the payload and other systems but also helping to stabilize the flight attitude through its conical profile. The aerodynamic arc base 3 is located at the bottom of the structural cone section 2. The payload system is located within the nose cone 1 or within the structural cone section 2. The payload system can withstand various mechanical and thermal environments during launch, on-orbit operation, and reentry. The measurement system is located within the structural cone section 2. The recovery system is located within the aerodynamic arc base 3. The measurement system monitors the payload capsule's flight parameters (such as position, velocity, acceleration, temperature, etc.) in real time, providing critical data to the ground control center. The recovery system ensures that the payload capsule can be effectively located and recovered after returning to Earth.

[0029] like Figure 1 As shown, the payload compartment also includes: equipment mounting brackets, a hatch 21, an operating port cover 11, a heat shield 22, and an impact damping device 6; the hatch 21, heat shield 22, and impact damping device 6 are mounted on the structural cone section 2; the operating port cover 11 is mounted on the end cap 1 and the impact damping device 6. The impact damping device 6 uses a combination of aluminum alloy, aluminum alloy honeycomb core, stainless steel, and polytetrafluoroethylene to reduce the impact during the separation of the compartment and the rocket.

[0030] This application discloses a reusable suborbital payload capsule with a classic blunt-nosed cone configuration, located at the foremost (top) end of the booster launch vehicle, with its lower end connected to the transfer section. It lacks an active attitude control system and utilizes a spin-stabilized parachute landing for recovery. Preferably, the payload capsule has a theoretical maximum diameter of Φ1016mm and a height of 1106mm. According to overall requirements, the payload capsule and transfer section are connected at six points (non-electro-coupled explosive bolts). The main load-bearing structure is a thin-walled, mesh-reinforced cone section integrally machined from aluminum alloy. Subsequent product finalization will utilize carbon fiber composite materials for further weight reduction and enhanced strength, enabling reusability.

[0031] Preferably, the end cap 1 is made of stainless steel and is mainly used to adjust the axial center of gravity of the entire cabin and resist impact during return and landing (with an internal sandwich structure), protect the internal payload, and isolate aerodynamic heat conduction and radiation to the internal payload (with an internal polyimide aluminum-coated film).

[0032] like Figure 2As shown, the inner bulkhead of structural cone segment 2 is connected to a load mounting disc; the load mounting disc is an aluminum alloy frame with an aluminum honeycomb core and a carbon fiber panel. The load mounting disc includes a first load mounting disc 24 and a second load mounting disc 25. The first load mounting disc 24 and the second load mounting disc 25 are parallel and spaced apart, distributed vertically. One load mounting disc (the upper load mounting disc) is connected to the crossbeam of the inner bulkhead of structural cone segment 2; the other load mounting disc (the lower load mounting disc) is connected to the lower end frame of structural cone segment 2 using corner joints. The first load mounting disc 24 and the second load mounting disc 25 are the main load-bearing structures for the effective load and measurement system equipment. Reinforcing tie rods 26 are provided between the first load mounting disc 24 and the bulkhead of structural cone segment 2, and between the second load mounting disc 25 and the bulkhead of structural cone segment 2. The reinforcing tie rods 26 are used to strengthen the stability and reliability of the connection between the first load mounting disc 24 and the bulkhead of structural cone segment 2, and between the second load mounting disc 25 and the bulkhead of structural cone segment 2, thereby enhancing the strength and stiffness of the entire structure.

[0033] like Figure 2 and 3 As shown, the reinforcing tie rods 26 comprise three sets, spaced apart along the circumference of the structural conical segment 2. The included angle between adjacent sets of reinforcing tie rods 26 is 120 degrees. The three sets of reinforcing tie rods 26 are positioned on the outer periphery of the umbrella-shaped structure, between the first load mounting disk 24 and the second load mounting disk 25. One end of each set of reinforcing tie rods 26 is connected to the inner wall of the structural conical segment 2, and the other end is connected to the first load mounting disk 24 and the second load mounting disk 25, forming a high-rigidity load-bearing frame. The three sets of reinforcing tie rods 26 ensure a stable and reliable connection between the first load mounting disk 24 and the second load mounting disk 25 and the structural conical segment 2. The three sets of reinforcing tie rods 26 are used to strengthen and distribute the load-bearing capacity, thereby enhancing the strength and rigidity of the entire structure.

[0034] like Figure 1 and 4 As shown, a separation and disconnection assembly 23 is installed on the lower surface of the first load mounting disk 24. When the load compartment is separated, the separation and disconnection assembly 23 reliably pulls out the power supply / signal / high frequency connector in one go with a constant pulling force, ensuring that the electrical connection is broken first, no rebound occurs, and the cable is not damaged. At the same time, the pull-out load is 100% borne by itself and is not transmitted to the pneumatic arc base 3.

[0035] In a specific embodiment of the present invention, the structural cone segment 2 is conical in shape, and in addition to the heat-resistant coating, the outer wall of the structural cone segment 2 also has a protruding heat shield 22. The heat shield 22 is also made of polyimide-based glass fiber composite material, which provides heat insulation while meeting the transmittance requirements of the GNSS antenna. Furthermore, the bottom is made of alumina transparent ceramic, which meets the requirements of high temperature resistance, heat insulation, and the visible light transmittance of the camera device 79 is above 80%.

[0036] As a specific embodiment of the present invention, the aerodynamic arc bottom 3 has an arc-shaped aerodynamic shape and is made of polyimide-based glass fiber composite material. Located at the bottom of the payload compartment, the aerodynamic arc bottom 3's arc design helps to provide lift during reentry, control the descent trajectory, and achieve stable aerodynamic deceleration. The aerodynamic arc bottom 3 is used to protect equipment inside the compartment and reduce aerodynamic drag.

[0037] As a specific embodiment of the present invention, the pneumatic arc bottom 3 is provided with a heat-resistant coating, which can effectively insulate heat and protect the equipment inside the cabin.

[0038] like Figure 3 As shown, the measurement system includes: a power module 41, an RF front-end 42, a central programmable controller 43, a GNSS telemetry antenna 73, a cable network system, and / or an external measurement system 75. The power module 41 provides power to all devices; the RF front-end 42 and the GNSS telemetry antenna 73 provide data transmission functionality to the payload bay; the GNSS telemetry antenna 73 also serves as an S-band telemetry transmitting antenna; the central programmable controller 43 acts as the brain of the payload bay, controlling the orderly operation of each device and also functions as a payload manager; the cable network system connects all devices, transmitting energy, data, and information; the external measurement system 75 collects reentry aerodynamic parameters such as surface heat flux density, boundary layer pressure, and spectral radiation intensity. The external measurement system 75 acquires in-situ data of the hypersonic aerodynamic / thermal / fluid fields during the reentry-descent phase.

[0039] like Figure 4 As shown, the cable network system is connected to the upper surface of the first load mounting disk 24 via the cable network system test interface 74. The cable network system test interface 74 is fixedly connected to the first load mounting disk 24.

[0040] As a specific embodiment of the present invention, the external measurement system 75 includes: an antenna, a camera device 79, multiple sensors, and multiple transducers. The antenna is used to receive a GNSS time reference (for synchronous sampling).

[0041] like Figure 3As shown, multiple sensors include: in-situ measurement sensors (high-precision air pressure sensor, temperature sensor, air pressure sensor, ozone concentration sensor), wall temperature sensor, noise sensor 57, triaxial impact sensor 51, triaxial acceleration sensor 53, low-frequency vibration sensor 55, high-frequency vibration sensor 54, air temperature sensor 59, first pressure sensor 71, and second pressure sensor 72; the in-situ measurement sensors are used for in-situ measurements; the wall temperature sensor is used to measure the temperature of the inner wall of the payload compartment; the noise sensor 57 is used to measure the noise inside the compartment; the triaxial impact sensor 51 is used to measure the impact in three directions inside the payload compartment; the triaxial acceleration sensor 53 is used to measure the acceleration inside the compartment; and the low-frequency vibration sensor 55 and the high-frequency vibration sensor 54 are used to measure the vibration of the payload mounting surface inside the compartment.

[0042] like Figure 3 As shown, the transducer includes: an air temperature transducer 58, a wall temperature transducer 60, a noise transducer 56, and a triaxial impact transducer 52. The wall temperature sensor and wall temperature transducer 60 are used to measure the temperature of the inner wall of the payload compartment; the noise sensor 57 and noise transducer 56 are used to measure the noise level inside the compartment; the triaxial impact sensor 51 and triaxial impact transducer 52 are used to measure the impact in the three axes inside the payload compartment.

[0043] like Figure 1-3 As shown, the payload system includes: a first payload 61, a second payload 62, and multiple standard unit modules 27. The first payload 61 has an envelope size of Φ650mm and a height of 450mm, supporting a payload weight of 138Kg. Depending on specific needs, it can be used as a whole or divided into multiple modules for combined utilization. The second payload 62 has an envelope size of 150mm×124mm×70mm and a weight of 3Kg. The standard unit modules 27 have an envelope size of 110mm×110mm×85mm and a weight of 3Kg, and can be used as payload resources, such as key equipment for near-space in-situ measurements. The output data of the first payload 61 is typically written into the top-level mission success criteria. The second payload 62 utilizes the "remaining resources from the main experiment" for flight, reducing costs and improving cabin utilization. The data from the second payload 62 serves as a "mission bonus," not affecting the success or failure of the main mission, but it can verify new technologies.

[0044] As a preferred embodiment of the present invention, the total weight of the payload compartment does not exceed 300 kg, the weight of the supported payload is not less than 150 kg, the envelope size of the first supported payload 61 is Φ650 mm × 450.6 mm (height) (150 L volume); the size of the second supported payload 62 and the three standard unit modules 27 can be customized (approximately 50 L volume).

[0045] like Figure 2 and 4 As shown, the recovery system includes: a parachute recovery system component 4 and a parachute recovery controller 5; the parachute recovery controller 5 integrates sensing devices, including an accelerometer.

[0046] As a specific embodiment of the present invention, the parachute recovery system component 4 includes: a parachute compartment assembly, a deceleration chute, a main parachute, an ejector, and a pin puller. The upper end of the parachute compartment assembly is connected to the reentry suborbital payload compartment structure via bolts. The lower inner wall of the parachute compartment assembly is provided with guide rails and limit blocks to provide an exit channel for the deceleration chute / main parachute pack and prevent swaying. The parachute compartment assembly provides an installation base and space chamber for the entire parachute system. After being ejected, the parachute compartment cover of the parachute compartment assembly forms a 180° outward flip, becoming the first "pull point" to pull out the deceleration chute pack. The deceleration chute is connected to the inner center ring of the parachute compartment cover via guide lines. The guide lines at the top of the main parachute are connected to the outer shell of the deceleration chute release device; once the deceleration chute detaches, the main parachute pack top cover is pulled out in the opposite direction by the deceleration chute straps. The main parachute is used for secondary deceleration, reducing the speed of the reentry capsule. The ejector is fixed to the inner wall of the parachute compartment via flanges, and the ejector ensures that the compartment cover leaves the compartment first, and then the parachute is pulled out. After the pin is pulled out by the ejector, the parachute canopy can only be flipped outward by the thrust of the catapult.

[0047] This application employs spin-stabilized uncontrolled reentry technology, combined with MBSE for aerodynamic and mass balancing. MBSE is used to lock the three variables of "aerodynamics, mass, and moment of inertia" into the same parameter model simultaneously. In one iteration, it can be seen whether the attitude drift exceeds the limit, ensuring the stability of the reentry attitude.

[0048] This application features a modular payload interface that supports various experimental configurations, with a payload weight ratio of up to 50% and a payload volume of up to 200L. It is also recyclable.

[0049] This application also includes a heat protection system comprising a multi-layered thermal insulation assembly, an alumina transparent ceramic window, a heat shield, and / or a thermal control coating; the multi-layered thermal insulation assembly is disposed within the end cap 1; the alumina transparent ceramic window and the heat shield are disposed within the structural cone section 2; the thermal control coating is applied to the outer surface of the end cap 1, the outer surface of the structural cone section 2, the outer surface of the aerodynamic arc bottom 3, and / or the inner wall of the outer surface of the parachute canopy. By implementing the heat protection system, the internal temperature during reentry is achieved to be ≤50°C.

[0050] This application provides a recoverable suborbital payload capsule, which serves as a recoverable suborbital spacecraft for microgravity science experiments and in-situ measurements in near-space (atmosphere gas composition, temperature, density, pressure, wind, and disturbances, etc.).

[0051] This application discloses a reusable suborbital payload capsule that can achieve the following: 1) Provides a microgravity environment and supports highly in-situ exploration; 2) Enables safe recovery of the payload compartment and reuse of some structures, significantly reducing testing costs; 3) Supports in-situ measurement of atmospheric parameters in near space (up to 200 km); 4) The system employs uncontrolled spin-stabilized reentry and parachute recovery, resulting in a simple and highly reliable system; 5) Improve the effective load-bearing efficiency, with its weight accounting for more than 50%.

[0052] The beneficial effects achieved by this application are as follows: (1) This application enables the safe recovery and partial reuse of the reusable suborbital payload capsule, significantly reducing the cost of a single mission, achieving high landing accuracy, eliminating the need for large-scale search and rescue, rapidly recovering experimental data, supporting long-term microgravity experiments and in-situ exploration, and providing a low-cost, high-frequency experimental platform for microgravity science, space exploration and technology verification.

[0053] (2) This application integrates the parachute compartment recovery system components into the aerodynamic base to improve space utilization.

[0054] (3) The structural cone section of this application adopts a reusable design and uses a heat protection system to achieve effective heat insulation, and the impact environment is controlled below 1000g.

[0055] In the description of this application, the terms "first" and "second" are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of indicated technical features. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of the stated features. In the description of this application, "multiple" means two or more, unless otherwise explicitly specified.

[0056] In the description of this application, the word "for example" is used to mean "used as an example, illustration, or description." Any embodiment described as "for example" in this application is not necessarily to be construed as being more preferred or advantageous than other embodiments. The following description is provided to enable any person skilled in the art to make and use the invention. Details are set forth in the following description for purposes of explanation. It should be understood that those skilled in the art will recognize that the invention can be made without using these specific details. In other instances, well-known structures and processes will not be described in detail to avoid obscuring the description of the invention with unnecessary detail. Therefore, the invention is not intended to be limited to the embodiments shown, but is consistent with the broadest scope of the principles and features disclosed in this application.

[0057] The above description is merely an embodiment of the present invention and is not intended to limit the invention. Various modifications and variations can be made to the present invention by those skilled in the art. Any modifications, equivalent substitutions, improvements, etc., made within the spirit and principle of the present invention should be included within the scope of the claims of the present invention.

Claims

1. A reusable suborbital payload capsule, characterized in that, The payload compartment includes: an end cap, a structural cone section, an aerodynamic arc bottom, a payload system, a measurement system, and a recovery system; The end cap is positioned at the top of the structural cone section; The pneumatic arc bottom is located at the bottom of the structural cone section; The payload system is located inside the end cap or the structural cone section; The measurement system is installed within the conical section of the structure; The recovery system is located inside the pneumatic arc base.

2. The reusable suborbital payload capsule according to claim 1, characterized in that, The payload compartment also includes: a hatch, an operating port cover, a heat shield, and an impact damping device; The hatch, the heat shield, and the impact damping device are mounted on the conical section of the structure; the operating port cover is mounted on the end cap and the impact damping device.

3. The reusable suborbital payload capsule according to claim 1, characterized in that, The inner wall of the conical section of the structure is connected to a load mounting disc. The load mounting disk is made of aluminum alloy frame, aluminum honeycomb core, and carbon fiber panel.

4. The reusable suborbital payload capsule according to claim 1, characterized in that, The measurement system includes: a power module, a radio frequency front end, a central program controller, a GNSS telemetry antenna, a cable network system, and / or an external measurement system; The power module provides energy to each device; The radio frequency front end and the GNSS telemetry antenna provide data transmission functionality for the payload compartment; The central program controller acts as the brain of the payload compartment, controlling the orderly operation of various devices and also functions as a payload manager. The cable network system connects the various devices.

5. The reusable suborbital payload capsule according to claim 4, characterized in that, The extravehicular measurement system includes an antenna, a camera device, multiple sensors, and multiple transducers.

6. The reusable suborbital payload capsule according to claim 5, characterized in that, Multiple sensors include: in-situ measurement sensors, wall temperature sensors, noise sensors, triaxial impact sensors, triaxial acceleration sensors, low-frequency vibration sensors, and high-frequency vibration sensors; The in-situ measurement sensor is used for in-situ measurement; The wall temperature sensor is used to measure the temperature of the inner wall of the payload compartment; The noise sensor is used to measure the noise inside the cabin; The triaxial impact sensor is used to measure the impact in the three axes inside the load chamber; The triaxial accelerometer is used to measure the acceleration inside the cabin; The low-frequency vibration sensor and the high-frequency vibration sensor are used to measure the vibration of the payload mounting surface inside the cabin.

7. The reusable suborbital payload capsule according to claim 1, characterized in that, The payload system includes: a first payload, a second payload, and multiple standard unit modules.

8. The reusable suborbital payload capsule according to claim 1, characterized in that, The recovery system includes: a parachute recovery system component and a parachute recovery controller; The parachute recovery controller integrates sensing devices.

9. The reusable suborbital payload capsule according to claim 8, characterized in that, The parachute recovery system components include: a parachute compartment assembly, a deceleration parachute, a main parachute, an ejector, and a pin puller.

10. The reusable suborbital payload capsule according to claim 1, characterized in that, The aerodynamic arc base has an arc-shaped aerodynamic shape and is made of polyimide-based glass fiber composite material.

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