Externally mounted partially reusable launch vehicle system and its launch and recovery methods

Abstract

This invention discloses an externally mounted partially reusable launch vehicle system and its launch and recovery method, relating to the field of launch vehicle technology. It includes a disposable core stage and at least one externally mounted recovery module. The externally mounted recovery module is equipped with a main engine and flight control equipment, has a lifting body aerodynamic shape, and is detachably mounted to the outside of the disposable core stage via a mounting bracket and separation mechanism. A propellant tank is connected to the main engine via a propellant delivery pipeline equipped with a self-breaking connector to supply propellant. The self-breaking connector is used to disconnect the propellant delivery pipeline when the externally mounted recovery module separates from the disposable core stage. After separating from the disposable core stage, the externally mounted recovery module can return and land using a lifting body gliding method. This invention achieves the unity of complete recovery of high-value subsystems and disposal of low-value structures by integrating high-value subsystems into the externally mounted recovery module and retaining the low-value tank structure in the disposable core stage, reducing reentry heat protection costs and payload capacity loss.

Description

Technical Field

[0001] This invention relates to the field of launch vehicle technology, and in particular to an externally mounted partially reusable launch vehicle system and its launch and recovery method. Background Technology

[0002] The technology for reusing and recovering launch vehicles has been continuously developing since the late 20th century. Currently, the main technical routes that have been established or verified in engineering include the following three categories:

[0003] Single-use launch vehicles. All rocket structure, engines, and electronic equipment are discarded after a single flight, eliminating the problem of recovery. However, all hardware must be purchased again for each flight, resulting in high launch costs, which is the fundamental driving force behind the development of reusable technology.

[0004] Vertical stage recovery. Represented by thrust reverse vertical landing: the entire stage re-enters the atmosphere, relying on multiple engine ignitions for deceleration and landing buffer mechanisms for support, to achieve whole-stage recovery. This approach requires the entire stage tank to withstand re-entry aerodynamic heating or to be equipped with thermal protection for the re-entry phase, requires the engine to have deep throttling and multiple start-up capabilities, and must be equipped with additional mechanisms such as landing legs and grid rudders for the entire stage.

[0005] Engine nacelle recovery via parachute. Examples include parachute deceleration, helicopter aerial capture, or sea surface salvage: only the engine nacelle and storage tank are separated and recovered via parachute, while the rest is used once.

[0006] All of the above routes suffer from insurmountable engineering defects: The drawback of whole-stage vertical recovery is the huge cost of dead weight. The propellant tank is the component with the largest surface area in the rocket. Whole-stage recovery means that thermal protection or thermal management design must be carried out on the huge surface. In addition, the landing buffer mechanism, depth throttling device, and return propellant surplus have added mass, which makes the rocket's carrying capacity significantly reduced. At the same time, the workload of inspection and refurbishment after whole-stage reentry is large and the reuse preparation cycle is long.

[0007] The drawbacks of parachute recovery of engine nacelles are the incomplete recovery of the target and poor recovery accuracy. The recovered object lacks aerodynamic control capabilities, has a large dispersion of landing points, and is highly dependent on weather conditions and the acquisition platform. More importantly, high-value electronic payloads such as flight control computers, inertial measurement units, telemetry and safety control equipment are usually jettisoned with the rocket body and are not included in the recovery scope.

[0008] A two-dimensional analysis of the value and reentry difficulty of a launch vehicle stage reveals a severe inversion between the two, which is the core contradiction that current technology has failed to resolve: High-value subsystems (including the main engine, control system, telemetry / safety control system, and pressurization and delivery system) are compact in size and concentrated in mass, accounting for the highest proportion of value among all rocket hardware; their individual reentry and recovery requires a small thermal protection area and involves low engineering difficulty. Low-value subsystems (including propellant tanks, inter-tank sections, interstage sections, and fairings) are large in size and have a huge surface area, with extremely low value per unit area; if recovered as a whole, the reentry thermal protection area requirement is extremely high, and the engineering cost is enormous.

[0009] The full-stage recovery route essentially involves paying the highest thermal protection cost for low-value, large-surface-area structures; the nacelle recovery route only salvages some high-value hardware while abandoning the electronic payload. Neither approach offers a systematic solution to the aforementioned inverted contradiction. Summary of the Invention

[0010] In view of this, and addressing the contradiction between the inverted value density and reentry difficulty in existing technologies, the purpose of this invention is to provide an externally mounted partially reusable launch vehicle system and its launch and recovery method. This system externally mounts the high-value, compact power and electronic subsystems onto a disposable core stage, achieving controlled gliding return and horizontal landing through a lifting body aerodynamic design. The low-value, large-surface-area propellant tank structure is retained on the disposable core stage, without being recovered or heat-protected. This maximizes the preservation of recovery value while minimizing the need for reentry heat protection area and the loss of carrying capacity.

[0011] To achieve the above objectives, the present invention adopts the following technical solution: The core concept of this invention is to decouple the functions of the launch vehicle substages according to their value density and reentry difficulty, and to split them into two types of parts.

[0012] The first type is a single-use core stage, which retains only the functions of propellant storage and structural support. It includes propellant tanks, inter-tank sections, inter-stage sections, fairings, and a simple passivation system. It maintains a thin-walled and lightweight design, does not take any reentry thermal protection reinforcement, and is passivated and emptied after the mission and falls uncontrolled to the predetermined landing area.

[0013] The second type is the external recovery module, which integrates all high-value subsystems such as the main engine, flight control, telemetry and safety control, and pressurization and delivery into an independent aircraft with a lifting body aerodynamic shape, thermal protection and landing gear, and is externally mounted outside the core stage. During flight, the core stage tank supplies propellant to the engine of the external recovery module through a large-diameter cryogenic pipeline. After the substage shuts down and the first and second stages separate, the external recovery module glides with the core stage assembly to a near-vacuum environment of more than 80km, then separates from the core stage, glides back using its own lifting body shape and lands horizontally on the runway.

[0014] The essence of this architecture lies in the fact that it does not reclaim the entire sub-level, but rather extracts the high-value subsystems as a whole and discards the low-value structures. The reclaimed objects cover all high-value subsystems, while the thermal protection objects are limited to small-volume modules, thereby achieving a balance between maximizing the value of reclamation and minimizing the cost of thermal protection.

[0015] This invention provides, in one aspect, an externally mounted partially reusable launch vehicle system, comprising: A single-use core stage, comprising at least a propellant tank; and At least one external recovery module is provided with a main engine and flight control equipment, has a lifting body aerodynamic shape, and is detachably installed on the outside of the disposable core stage via a rack and separation mechanism; The propellant tank is connected to the main engine of the external recovery module via a propellant delivery pipeline equipped with a self-disconnecting connector to supply propellant; the self-disconnecting connector is used to disconnect the propellant delivery pipeline when the external recovery module separates from the disposable core stage; after separating from the disposable core stage, the external recovery module can return in a lifting body gliding manner and land horizontally.

[0016] This invention integrates high-value subsystems such as the main engine and flight control equipment into a separate external recovery module, which is then detachably installed on the outside of the disposable core stage. This achieves physical separation between high-value subsystems and low-value tank structures. The disposable core stage does not require reentry thermal protection and landing recovery mechanisms, avoiding the cost of thermal protection for large surface area tanks. After separation, the external recovery module returns via a lifting body gliding mode and lands horizontally with precise and controllable landing point. The recovered objects completely cover the engine and electronic payload, overcoming the dual defects of high dead weight cost of whole-stage recovery and abandonment of electronic payload during nacelle recovery in the prior art.

[0017] As a further improvement to the above technical solution, the self-breaking connector is disposed on the propellant delivery pipeline between the disposable core stage and the external recovery module; The self-disconnecting connector includes: The first self-sealing valve and the second self-sealing valve are respectively located at one end of the propellant delivery pipeline near the disposable core stage and the other end near the external recovery module; A discharge valve assembly is disposed on the propellant delivery pipeline between the first self-sealing valve and the second self-sealing valve, for discharging the propellant in the propellant delivery pipeline between the first self-sealing valve and the second self-sealing valve after the first self-sealing valve and the second self-sealing valve are closed; A cutting device is installed on the propellant delivery pipeline and located between the first self-sealing valve and the second self-sealing valve, for disconnecting the propellant delivery pipeline after the propellant has been emptied.

[0018] As a further improvement to the above technical solution, the self-disconnecting connector also includes a pipeline compartment door; the pipeline compartment door is disposed on the outer wall of the external recovery module and is used to seal the pipeline break after the propellant delivery pipeline is disconnected, so as to maintain the aerodynamic shape of the outer surface of the external recovery module.

[0019] As a further improvement to the above technical solution, the number of external recycling modules is even, and every two external recycling modules form a symmetrical group. The two external recycling modules in the same symmetrical group are arranged symmetrically on opposite sides of the disposable core level along the radial direction of the disposable core level. Each external recycling module is installed on the disposable core level through the bracket and separation mechanism.

[0020] As a further improvement to the above technical solution, the hanger and separation mechanism includes a through-load-bearing beam and a catapult pusher; The through-load-bearing beam is radially arranged along the disposable core level, and the two external recycling modules in each of the symmetrical groups are mounted on both ends of the same through-load-bearing beam; The ejector thruster is installed on the through-bracing beam and is used to simultaneously apply thrusting forces in opposite directions to the two external recovery modules on the same through-bracing beam, so that the two external recovery modules on the same through-bracing beam exit the disposable core stage synchronously. The through-bracing beam and the disposable core are connected by a support that transmits only shear force, so that the separation reaction force of the two external recovery modules in the same symmetrical group closes along the corresponding through-bracing beam within the group.

[0021] As a further improvement to the above technical solution, the disposable core stage is provided with an annular distribution manifold, which is arranged circumferentially along the disposable core stage, and the propellant tank is connected to the inlet of the annular distribution manifold. The propellant delivery pipeline consists of multiple equal-length pipes. The annular distribution main pipe is connected to the main engine of each of the external recovery modules through the multiple propellant delivery pipelines to form an equal-resistance grouped supply system for uniform propellant supply.

[0022] As a further improvement to the above technical solution, a detachment panel is fixed on the outer wall of each of the external recycling modules at the position between the first self-sealing valve and the second self-sealing valve. The release panel is equipped with an electrical release connector. One end of the electrical release connector is electrically connected to the external recycling module, and the other end is electrically connected to the disposable core via an electrical cable. The electrical release connector can be released and disconnected when the external recycling module is separated from the disposable core. The cutting device has a detonating cord, which is arranged around the propellant delivery pipeline and located between the pull-out panel and the first self-sealing valve, for cutting off the propellant delivery pipeline; The propellant delivery pipeline passes through the middle of the pull-out panel and is arranged in a centralized manner alongside the electrical cables.

[0023] As a further improvement to the above technical solution, the second self-sealing valve and the remaining section of the disconnected propellant delivery pipeline connected thereto constitute an integrated quick-change assembly that can be replaced as a whole after being recovered by the external recovery module.

[0024] Another aspect of the present invention provides a launch and recovery method based on an externally mounted partially reusable launch vehicle system, comprising the following steps: S1: The main engine of the external recovery module provides thrust to drive the launch vehicle to ascend, and the propellant tank supplies propellant to the main engine through the propellant delivery pipeline; S2: After the first stage of flight is completed, the main engine shuts down at an altitude of over 50km, and the first and second stages separate. The external recovery module glides along with the disposable core stage due to inertia. S3: When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the self-disconnecting connector disconnects the propellant delivery pipeline, the rack and separation mechanism are unlocked, and the external recovery module is separated from the disposable core stage. S4: The external recovery module returns in a gliding manner and lands horizontally on the runway, while the disposable core stage passesivates and falls after being emptied.

[0025] As a further improvement to the above technical solution, in step S2: the main engine is synchronously shut down at an altitude of over 50km; the first and second self-sealing valves of the self-disconnecting connector are closed, and the propellant is respectively sealed on the outside of the first and second self-sealing valves; after the first and second stage separation is implemented, the external recovery module slides along with the inertia of the disposable core stage; In step S3: when the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly of the self-breaking connector opens to drain the propellant in the pipe section between the first self-sealing valve and the second self-sealing valve; after the pressure in the pipe drops below a set threshold, the cutting device cuts off the propellant delivery pipeline; the hanger and separation mechanism unlock, and the ejector pusher pushes the external recovery module away from the disposable core stage; After the external recovery module separates from the disposable core stage, the pipeline compartment door is closed and locked, sealing the break in the propellant delivery pipeline; In step S4: After the external recovery module lands and is recovered, the integrated quick-change assembly consisting of the second self-sealing valve and the remaining pipe section of the propellant delivery pipeline connected to it is replaced as a whole and leak-checked for reuse.

[0026] As can be seen from the above technical solutions, compared with the prior art, the present invention discloses an externally mounted partially reusable launch vehicle system and its launch and recovery method, which has the following advantages and beneficial effects.

[0027] 1. Complete recovery of high-value subsystems with minimized heat protection costs. This invention integrates high-value subsystems such as the main engine and flight control equipment into a separate external recovery module, which is detachably installed outside the disposable core stage. Propellant tanks supply propellant to the main engine via propellant delivery pipelines equipped with self-disconnecting connectors, achieving physical separation of high-value power and electronic subsystems from low-value tank structures. The disposable core stage does not require reentry heat protection and landing recovery mechanisms; only the small-volume structure of the external recovery module needs heat protection. The reentry heat protection area is equivalent to 5%-10% of the total rocket surface area, fundamentally avoiding the engineering challenge of paying high heat protection costs for large-surface-area tanks in whole-stage recovery. At the same time, the recovered objects completely cover the engine and electronic payloads, overcoming the dual drawbacks of high dead weight costs in whole-stage recovery and abandoning electronic payloads in nacelle recovery.

[0028] 2. Self-severing connector enables controllable disconnection of the pipeline during flight, ensuring inherent safety during the separation process. This invention achieves a complete timing logic of "first sealing, then emptying, and then cutting off" for the propellant delivery pipeline by installing a self-severing connector on the propellant delivery pipeline, consisting of a first self-sealing valve, a second self-sealing valve, a discharge valve assembly, and a cutting device. The cutting off of the propellant delivery pipeline is carried out in a near-vacuum environment (altitude greater than 80km or dynamic pressure less than 200Pa). After the discharge valve assembly empties the propellant in the section between the two valves, the pressure inside the pipeline drops to a set threshold as a prerequisite for the cutting device to detonate. This eliminates the risks of high dynamic pressure flutter and oxygen-rich deflagration from both timing and environmental conditions, ensuring the inherent safety of the separation process. The matching pipeline compartment door automatically seals the pipeline break after the propellant delivery pipeline is disconnected, restoring the continuous aerodynamic shape of the external recovery module and ensuring the aerodynamic integrity of the gliding return.

[0029] 3. Symmetrical grouping layout achieves torque self-balancing, eliminating the need for local reinforcement in the load-bearing beam structure. This invention arranges the external recovery modules in an even number of radially symmetrical groups, enabling the thrust and aerodynamic torque of the two modules within the same symmetrical group to self-balance radially. This eliminates the rolling torque and transonic shock wave interference generated by a single-sided external mount without the need for counterweights. Two modules within the same symmetrical group are mounted at both ends of the same through-loaded load-bearing beam, with the ejector thruster installed on the through-loaded beam. The separation reaction force closes within the group along the through-loaded load-bearing beam without being transmitted to the tank's thin wall. The through-loaded load-bearing beam is connected to the disposable core stage via supports that only transmit shear force, eliminating the need for the traditional single-sided mounting reinforcement structure. The saved structural mass is directly converted into increased carrying capacity. Furthermore, the external recovery modules can be expanded in groups of 2, 4, or 6, and the same module product can be adapted to core stages of different sizes, reducing development and verification costs.

[0030] 4. A ring-shaped distribution manifold, in conjunction with equal-length pipelines, achieves equal flow resistance supply, eliminating the need for active balancing. This invention utilizes a circumferentially arranged ring-shaped distribution manifold on the disposable core stage. The propellant tank is connected to the inlet of the ring-shaped distribution manifold, which is then connected to the main engines of each external recovery module via multiple equal-length propellant delivery pipelines. This ensures that the inlet flow resistance of each module is essentially the same, the pressure deviation before the pumps of each module is controlled within 2%, and the thrust imbalance is less than 1.5%. This avoids thrust imbalance caused by differences in branch pipeline lengths in multi-module layouts, eliminating the need for active balancing using propellant.

[0031] 5. The release panel achieves integrated interfaces and clean separation, with synchronous disconnection of pipelines and electrical circuits. This invention fixes the release panel on the outer wall of the external recovery module, positioned between the first and second self-sealing valves. This allows the propellant delivery pipeline to pass through the release panel and be arranged parallel to the electrical cables, integrating the previously dispersed fluid and electrical interfaces onto the same panel. This avoids pipeline interference and cable entanglement caused by discrete arrangements. The detonating cord of the cutting device is positioned between the release panel and the first self-sealing valve. After the detonating cord detonates and cuts the propellant delivery pipeline, the cut position is located on the side of the release panel closest to the first self-sealing valve. The remaining pipeline section on the module side after cutting is carried away by the external recovery module, achieving reliable cutting of the propellant delivery pipeline at the release panel position. The electrical release connector on the release panel automatically releases and disconnects when the module separates from the core, achieving synchronous coordination between propellant delivery pipeline cutting and electrical circuit disconnection. This avoids cables tangling or pulling on the module after separation, ensuring a clean and efficient separation process.

[0032] 6. Integrated quick-change assembly enables rapid module reuse. This invention constructs an integrated quick-change assembly by combining the second self-sealing valve and the remaining section of the propellant delivery pipeline. This assembly is mechanically connected to the pipeline inside the module. After the module is recovered, it can be replaced as a whole and subjected to helium mass spectrometry leak detection. There is no need to enter the cabin for welding operations, which greatly reduces the reuse preparation time and reduces the time cost of re-mounting the module to perform missions.

[0033] 7. Near-vacuum separation ensures the module has defined initial conditions, enabling gliding landing and unpowered return. The launch and recovery method provided by this invention utilizes a complete timeline: "first-stage flight, shutdown above 50km, first and second-stage separation, module gliding with the core stage's inertia, self-disconnecting connectors to break pipelines at altitudes above 80km or dynamic pressures below 200Pa, unlocking and separating the hangar and separation mechanism, module gliding back to horizontal landing, and core stage descent." This utilizes the core stage's remaining kinetic energy to carry the module to a separation altitude above 80km without consuming its own propellant. The separation point is in a near-vacuum environment, with negligible aerodynamic interference, ensuring the module obtains defined initial separation conditions. The module returns unpowered via gliding, resulting in a lower impact load on horizontal landing than vertical landing, and fewer reuse and maintenance items. After the core stage is passivated and emptied, it falls, eliminating the need for thermal protection and a landing mechanism, fundamentally avoiding the engineering challenges of paying high thermal protection costs for large surface area structures during full-stage recovery.

[0034] 8. The method steps and system structure work in synergy, ensuring a closed-loop controllable process. This invention further elaborates on the specific execution steps of the self-disconnecting connector, including "valve closure, venting, pressure threshold confirmation, and detonating cord cutting," as well as operational details such as ejection thruster pushing away, sealing the pipeline compartment door at the break point, and overall replacement and leak detection of the quick-change components after recovery. This forms a fully closed-loop operation process from engine shutdown, propellant plugging, first and second stage separation, high-altitude taxiing, pipeline venting, pressure confirmation, disconnection and separation, aerodynamic shape restoration, to reuse preparation. The timing and condition constraints of each step are clearly defined, ensuring the reliable operation of the recovery system and its rapid reuse economy in complex flight environments. Attached Figure Description

[0035] To more clearly illustrate the technical solutions in the embodiments of the present invention 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 embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on the provided drawings without creative effort.

[0036] Figure 1 This is a schematic diagram of the externally mounted reusable launch vehicle system of the present invention.

[0037] Figure 2This is a schematic diagram of the structure of the externally mounted reusable launch vehicle system of the present invention, which includes two externally mounted recovery modules.

[0038] Figure 3 This is a schematic diagram of the structure of the externally mounted reusable launch vehicle system of the present invention, which includes four externally mounted recovery modules.

[0039] Figure 4 This is a schematic diagram of the structure of the externally mounted reusable launch vehicle system of the present invention, which includes six externally mounted recovery modules.

[0040] Figure 5 This is a schematic diagram of the external recovery module structure of the externally mounted partially reusable launch vehicle system of the present invention. It shows the integrated layout of the subsystems inside the module, the mechanical-fluid interface between the module and the core stage, and the relative arrangement of the discharge valve assembly of the self-breaking joint, the detonating cord cutting device, and the pipeline compartment door.

[0041] Figure 6 This is a flight mission profile and separation / recovery sequence diagram for a Level 1 application, showing the entire process from vertical takeoff, Level 1 shutdown, Level 1 and Level 2 separation, inertial taxiing of the combined module, high-altitude discharge and cutting separation to module runway landing and core stage passivation and descent.

[0042] Figure 7 The diagram shows a comparison of the functional decoupling principle, where (a) is the existing whole-stage recovery structure and (b) is the external recovery structure of the present invention, illustrating the essential difference between the two in terms of reentry heat protection area.

[0043] In the diagram: 1. Disposable core stage; 11. Propellant tank; 12. Fairing; 13. Tail section / interface section; 2. External recovery module; 21. Main engine; 22. Flight control equipment; 221. Flight control computer; 222. Inertial measurement unit; 223. Telemetry / safety control equipment; 23. Pressurized gas cylinder group; 24. Turbop pump and its associated valves and pipelines; 25. Lifting body wing; 26. Vertical tail; 27. Landing gear; 28. Thermal protective skin; 29. ​​Aerodynamic control surfaces; 3. Plugs and separation mechanism; 31. Through-load beam; 4. Self-breaking joint; 41. Emission valve assembly; 42. Cutting device; 43. Piping compartment door; 5. Annular distribution manifold; 6. Pull-out panel. Detailed Implementation

[0044] Embodiments of the present invention are described in detail below, examples of which are illustrated in the accompanying drawings, wherein the same or similar reference numerals denote the same or similar elements or elements having the same or similar functions throughout. The embodiments described below with reference to the accompanying drawings are exemplary and intended to explain the present invention, and should not be construed as limiting the present invention.

[0045] In the description of this invention, it should be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing this invention and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this invention.

[0046] Furthermore, 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 technical features indicated. Thus, a feature defined as "first" or "second" may explicitly or implicitly include one or more of that feature. In the description of this invention, "a plurality of" means two or more, unless otherwise explicitly specified.

[0047] In this invention, unless otherwise explicitly specified and limited, the terms "installation," "connection," "linking," and "fixing," etc., should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral part; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; they can refer to the internal communication of two components or the interaction between two components. Those skilled in the art can understand the specific meaning of the above terms in this invention according to the specific circumstances.

[0048] According to embodiments of the present invention, such as Figures 1 to 7 As shown, the externally mounted reusable launch vehicle system includes: a disposable core stage 1 and at least one externally mounted recovery module 2.

[0049] The disposable core stage 1 includes at least a propellant tank 11; at least one external recovery module 2 is equipped with a main engine 21 and flight control equipment 22, has a lifting body aerodynamic shape, and is detachably mounted on the outside of the disposable core stage 1 via a rack and separation mechanism 3.

[0050] The propellant tank 11 is connected to the main engine 21 of the external recovery module 2 through a propellant delivery pipeline equipped with a self-disconnecting connector 4 to supply propellant; the self-disconnecting connector 4 is used to disconnect the propellant delivery pipeline when the external recovery module 2 separates from the disposable core stage 1; after separating from the disposable core stage 1, the external recovery module 2 can return in a lifting body gliding manner and land horizontally.

[0051] Specifically, the disposable core stage 1 includes a propellant tank 11, a fairing 12, and a tail / interface section 13. The propellant tank 11 stores propellants such as liquid oxygen and kerosene (or liquid oxygen and methane); the fairing 12 accommodates the payload; and the tail / interface section 13 provides a simplified passivation system for the mechanical interfaces of external modules and propellant piping channels, used for automatic depressurization and evacuation after separation to ensure the safety of the core stage during uncontrolled descent. The core stage itself does not have a main engine or only a small amount of auxiliary power; its main thrust is provided by the external recovery module.

[0052] The external recovery module 2 is an independent aircraft with a lifting body shape. Its integrated subsystems include a built-in main engine 21, a flight control device 22 consisting of a flight control computer 221, an inertial measurement unit 222 and telemetry / safety control equipment 223, a pressurized gas cylinder group 23, a turbopump and its associated valves and pipelines 24, as well as an external lifting body wing 25, a vertical tail 26, a landing gear 27, a thermal protective skin 28, and aerodynamic control surfaces 29. The main engine 21 is one or more, providing all or most of the thrust for the core stage flight phase; the flight control computer 221 and inertial measurement unit 222 are responsible for the overall flight control and autonomous guidance during the module reentry phase; the telemetry / safety control equipment 223 is responsible for transmitting telemetry parameters and receiving and executing safety control commands for the entire rocket; the pressurized gas cylinder group 23 provides the gas source for pressurizing the tanks and actuating the valves; the turbopump and its associated valves and pipelines 24 are used for pressurizing and transporting propellant and regulating its flow; the lifting body wing 25 and vertical tail 26 provide the lift-to-drag ratio and directional stability for reentry gliding; the landing gear 27 is wheeled or skid-mounted, used to support runway level landing; the thermal protection skin 28 covers the exterior of the external recovery module 2, using high-temperature resistant alloys or thin-layer ablation materials to adapt to suborbital / low-heat reentry; the aerodynamic control surfaces 29 are responsible for attitude and glide path control during the gliding phase.

[0053] In some embodiments, the self-disconnecting connector 4 is disposed on the propellant delivery pipeline between the disposable core stage 1 and the external recovery module 2; Self-disconnecting connector 4 includes: The first self-sealing valve and the second self-sealing valve are respectively located at one end of the propellant delivery pipeline near the disposable core stage 1 and the other end near the external recovery module 2; The discharge valve assembly 41 is located on the propellant delivery pipeline between the first self-sealing valve and the second self-sealing valve, and is used to discharge the propellant in the propellant delivery pipeline between the two valves after the first self-sealing valve and the second self-sealing valve are closed. The cutting device 42 is installed on the propellant delivery pipeline and located between the first self-sealing valve and the second self-sealing valve, and is used to disconnect the propellant delivery pipeline after the propellant is emptied.

[0054] By installing a self-shutdown connector 4 consisting of a first self-sealing valve, a second self-sealing valve, a discharge valve assembly 41, and a cutting device 42 on the propellant delivery pipeline, controllable segmented disconnection of the propellant delivery pipeline during flight is achieved: after the self-sealing valves at both ends are closed, the middle section is emptied through the discharge valve assembly 41, and finally cut off by the cutting device 42, forming a complete timing logic of "sealing first, then discharging, and finally cutting". This physically eliminates the risk of propellant leakage and oxygen-rich deflagration during the cutting process, ensuring the inherent safety of the separation process.

[0055] In some embodiments, the self-disconnecting connector 4 further includes a pipeline compartment door 43; the pipeline compartment door 43 is disposed on the outer wall of the external recovery module 2 and is used to close the pipeline break after the propellant delivery pipeline is disconnected, so as to maintain the aerodynamic shape of the outer surface of the external recovery module 2.

[0056] By setting a pipeline compartment door 43 on the outer wall of the external recovery module 2, the pipeline break is automatically closed after the propellant delivery pipeline is disconnected, so that the outer surface of the external recovery module 2 can restore the continuous aerodynamic shape. This avoids the additional aerodynamic interference caused by the protrusion or depression of the break during the reentry gliding phase, and ensures the directional stability and landing accuracy of the lifting body during gliding return.

[0057] It should be noted that the first self-sealing valve, the second self-sealing valve, the discharge valve assembly 41, the cutting device 42, and the pipeline compartment door 43 together constitute a large-diameter cryogenic propellant quick-break connector (i.e., self-breaking connector 4). Specifically: the first and second self-sealing valves are respectively located at both ends of the propellant delivery pipeline, used to immediately close the pipeline after shutdown and seal the propellant on both sides of the valves; the discharge valve assembly 41 is located on the pipeline section between the two valves, used to empty the propellant in that section after the two valves are closed; the cutting device 42 is located between the two self-sealing valves and spaced apart from the discharge valve assembly 41, used to cut off the propellant delivery pipeline after the pipeline is emptied. The pipeline compartment door 43 is located on the outer wall of the external recovery module 2, used to seal the break on the module surface after the pipeline is cut. All four components work collaboratively in the sequence of "first sealing, then discharging, then cutting off, and finally sealing the break," jointly achieving the controllable disconnection and break closure of the large-diameter cryogenic propellant delivery pipeline during flight.

[0058] Specifically, the pipeline hatch 43 closes under the drive of a spring or actuator and is mechanically locked using a center dead point mechanism. After the locking is confirmed by dual sensors, the external recovery module 2 enters the gliding trim procedure. Since the separation of the module from the disposable core stage 1 is carried out in a near-vacuum environment (dynamic pressure less than 200 Pa), the aerodynamic load on the pipeline hatch 43 when it closes is small, and the closing action is reliable. The center dead point locking mechanism ensures that the pipeline hatch 43 remains closed after the dynamic pressure in the reentry stage rises, preventing the hatch from being accidentally opened due to changes in aerodynamic pressure.

[0059] In some embodiments, the pipeline compartment door 43 may also adopt a sliding cover structure that slides and closes in the direction of airflow.

[0060] In some embodiments, the number of external recycling modules 2 is even, and every two external recycling modules 2 form a symmetrical group. The two external recycling modules 2 in the same symmetrical group are arranged symmetrically on opposite sides of the disposable core stage 1 along the radial direction of the disposable core stage 1. Each external recycling module 2 is installed on the disposable core stage 1 through a bracket and a separation mechanism 3.

[0061] By setting the number of external recovery modules 2 to an even number and arranging them in pairs radially symmetrical groups, the thrust and aerodynamic torque of two external recovery modules 2 in the same symmetrical group are self-balanced radially. This eliminates the rolling torque and transonic shock wave interference generated by a single external attachment without the need for counterweights, simplifying the design of the entire rocket attitude control system. At the same time, the number of external recovery modules 2 can be expanded to two, four or six groups, and the same module product can be adapted to core stages of different sizes, reducing the development and verification costs.

[0062] In some embodiments, the hanger and separation mechanism 3 includes a through-load beam 31 and a catapult thruster; The through-bearing beam 31 is radially arranged along the disposable core stage 1, and the two external recycling modules 2 in each symmetrical group are mounted at both ends of the same through-bearing beam 31; The ejector is installed on the through-bearing beam 31 and is used to simultaneously apply thrusts in opposite directions to two external recovery modules 2 on the same through-bearing beam 31, so that the two external recovery modules 2 on the same through-bearing beam 31 exit the disposable core stage 1 synchronously. The through-bearing beam 31 is connected to the disposable core stage 1 by a support that only transmits shear force, so that the separation reaction force of the two external recovery modules 2 in the same symmetrical group is closed along the corresponding through-bearing beam 31 within the group.

[0063] By setting a through-load-bearing beam 31 that runs through the tail section of the disposable core stage 1, two external recovery modules 2 in the same symmetrical group are mounted on both ends of the same through-load-bearing beam 31, and the ejector pusher is installed on the through-load-bearing beam 31. This allows the pushing forces of the two external recovery modules 2 to cancel each other out within the group during separation, and the separation reaction force to close along the inside of the through-load-bearing beam 31 without being transmitted to the thin wall of the tank. Combined with the support connection that only transmits shear force, the hanging point reinforcement frame and center of gravity counterweight required by traditional single-sided mounting are eliminated, and the saved structural mass is directly converted into an increase in carrying capacity.

[0064] In some embodiments, the hanger and separation mechanism 3 further includes a locking mechanism, which is installed on the through-bearing beam 31 and is used to detachably lock the external recycling module 2 onto the through-bearing beam 31.

[0065] In some embodiments, a ring-shaped distribution manifold 5 is provided on the disposable core stage 1, the ring-shaped distribution manifold 5 is arranged circumferentially along the disposable core stage 1, and the propellant tank 11 is connected to the inlet of the ring-shaped distribution manifold 5. The propellant delivery pipeline consists of multiple equal-length pipelines. The annular distribution main pipe 5 is connected to the main engine 21 of each external recovery module 2 through the multiple propellant delivery pipelines to form an equal-resistance grouped supply system for uniform propellant supply.

[0066] By setting an annular distribution manifold 5 on the disposable core stage 1, and connecting each external recovery module 2 through multiple propellant delivery pipelines of equal length via the annular distribution manifold 5, the flow resistance before the pump of each external recovery module 2 is basically the same, avoiding the thrust imbalance caused by the difference in branch pipeline length under the multi-module layout, and eliminating the need to consume propellant for active balancing.

[0067] In some embodiments, each external recycling module 2 has a detachment panel 6 fixed on its outer wall at the position between the first self-sealing valve and the second self-sealing valve. The release panel 6 is equipped with an electrical release connector. One end of the electrical release connector is electrically connected to the external recycling module 2, and the other end is electrically connected to the disposable core 1 through an electrical cable. The electrical release connector can be released and disconnected when the external recycling module 2 is separated from the disposable core 1. The cutting device 42 is a detonating cord cutting device with a detonating cord. The cutting device 42 is located outside the disposable core stage 1 and the external recovery module 2 to ensure that the detonation products are discharged outward when the detonating cord is detonated, without causing damage to the internal structure of the module. The detonating cord of the cutting device 42 is arranged around the propellant delivery pipeline and is located between the release panel 6 and the first self-sealing valve, and is used to cut off the propellant delivery pipeline. The propellant delivery pipeline passes through the pull-out panel 6 in the middle and is arranged in a centralized manner alongside the electrical cables.

[0068] By arranging the propellant delivery pipeline and electrical cables side-by-side on the release panel 6, the previously dispersed pipeline and electrical interfaces are integrated onto the same panel, simplifying the interface layout between the module and the core stage and avoiding pipeline interference and cable entanglement caused by discrete arrangements. After the detonating cord of the cutting device 42 detonates and cuts the propellant delivery pipeline, the cutting position is located on the side of the release panel 6 near the first self-sealing valve, allowing the remaining pipeline section on the module side to be carried away with the external recovery module 2, achieving reliable cutting of the propellant delivery pipeline at the release panel 6. The electrical release connector on the release panel 6 automatically disconnects when the module separates from the core stage, achieving synchronous coordination between the cutting of the propellant delivery pipeline and the disconnection of the electrical circuit, preventing cables from tangling or pulling on the external recovery module 2 after separation, and ensuring a clean and efficient separation process.

[0069] In some embodiments, the self-shutdown connector 4 operates in the following sequence: After the first stage of flight is completed and the main engine 21 is shut down, the first and second self-sealing valves of the self-shutdown connector 4 are closed first, sealing the propellant in the propellant delivery pipeline to both sides of the valves; after the first and second stages separate, the external recovery module 2 glides along with the disposable core stage 1 assembly due to inertia. When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly 41 of the self-shutdown connector 4 opens, and the liquid oxygen trapped in the pipe section between the first and second self-sealing valves automatically flashes under the dual drive of its own saturated vapor pressure and the pipe section trapping pressure, and is discharged through the directional nozzle of the discharge valve assembly 41. The residual propellant on the fuel side is blown out and emptied by the trapping air pressure; the propellant tank 11 of the disposable core stage 1 is simultaneously passivated and emptied. Once the pressure inside the pipe drops below the set threshold, the cutting device 42 detonates, cutting off the entire emptied propellant delivery pipeline within milliseconds. After the hanger and separation mechanism 3 unlock, the external recovery module 2 separates from the disposable core stage 1, and the remaining pipe section on the module side after cutting is carried away with the external recovery module 2. After separation, the pipeline compartment door 43 closes and locks, sealing the break in the propellant delivery pipeline on the surface of the external recovery module 2, restoring the continuous aerodynamic shape of the outer surface of the external recovery module 2. The above timing sequence achieves dual redundant sealing under large-diameter cryogenic conditions, millisecond-level cutting, emptying of residual liquid before cutting, and no leakage at both ends of the break after separation, as well as aerodynamic sealing of the break.

[0070] Specifically, the separation sequence adopts the overall sequence of "shutdown, coasting, discharge, separation": After the main engine 21 shuts down at an altitude greater than 50km and the first and second self-sealing valves are closed, the first and second stages of the launch vehicle are separated first. The external recovery module 2 continues to coast with the disposable core stage 1 assembly, and the altitude of the assembly's coasting peak is greater than 80km. When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly 41 is opened to discharge air, the cutting device 42 is cut off, the mounting bracket and separation mechanism 3 are unlocked, and the external recovery module 2 is ejected by the ejection thruster. The external recovery module 2 is given a determined separation speed and initial attitude by the ejection thruster. At this time, about one to two minutes have passed since the main engine 21 shut down, and the aftereffect thrust has decayed to a negligible level. Moreover, the separation is carried out under near-vacuum ballistic conditions, the aerodynamic interference is negligible, and the separation dynamics are clean.

[0071] Specifically, the two external recovery modules 2 within the same symmetrical group are synchronously unlocked by the same cutting device 42 via an equal-length symmetrical detonation transmission network. The detonation transmission speed is approximately 7000 m / s, and the difference in the length of the detonation transmission path is controlled at the decimeter level. The mechanical motion dispersion of the two external recovery modules 2 is compressed from ±1 to 3 milliseconds for conventional pyrotechnics to the order of hundreds of microseconds. Physically, this topology does not have a single-sided detonation mode; electrical faults can only cause the entire group to mis-detonate. The pyrotechnic control circuit adopts layered redundancy: mis-detonation protection is placed at the energy level, and each pyrotechnic device has two built-in detonators, which detonate upon activation of either detonator; accidental detonation protection is placed at the authorization level, and the detonation output is controlled by two independent enable signals connected in series: a separate program command and an independent safety verification channel. A single short circuit, static electricity, or electromagnetic interference cannot constitute the detonation condition.

[0072] Specifically, when using a two- or three-group layout, all external recovery modules 2 are unlocked and separated simultaneously; or a grouped sequential separation is adopted: each symmetrical group separates sequentially according to a set time sequence, with only one symmetrical group separating at a time. The separation interval between adjacent groups can be tens to hundreds of milliseconds, preferably not less than 100 milliseconds. The ejection direction of each symmetrical group is along its respective axis of symmetry, and the groups are naturally 90° or 60° apart. The thrust velocity vector has outward and backward components, so that the flight path of the rear separation group is offset from that of the front separation group in three-dimensional space. This decomposes the complex dynamic problem of simultaneous separation of multiple bodies into several two-body symmetrical separations. Each separation maintains torque self-balance, and the attitude of the core stage 1 is not disturbed.

[0073] In some embodiments, the propellant tank 11 includes a liquid oxygen tank and a fuel tank, used to store liquid oxygen and fuel, respectively. Two annular distribution manifolds 5 are provided on the disposable core stage 1, namely a liquid oxygen annular distribution manifold and a fuel annular distribution manifold, both arranged circumferentially along the disposable core stage 1. The liquid oxygen tank of the propellant tank 11 is connected to the inlet of the liquid oxygen annular distribution manifold, and the fuel tank is connected to the inlet of the fuel annular distribution manifold.

[0074] The propellant delivery pipelines are independent of the annular distribution manifold 5 and include liquid oxygen pipelines and fuel pipelines. Multiple liquid oxygen pipelines of equal length extend from the annular liquid oxygen distribution manifold to each external recovery module 2, connecting to the liquid oxygen inlet of the main engine 21 of each external recovery module 2. Multiple fuel pipelines of equal length extend from the annular fuel distribution manifold to each external recovery module 2, connecting to the fuel inlet of the main engine 21 of each external recovery module 2. Each external recovery module 2 corresponds to two propellant delivery pipelines: one liquid oxygen pipeline and one fuel pipeline, used for delivering liquid oxygen and fuel respectively.

[0075] Both the liquid oxygen pipeline and the fuel pipeline are equipped with self-shutdown connectors 4. Each self-shutdown connector 4 includes a first self-sealing valve, a second self-sealing valve, a discharge valve assembly 41, and a cutting device 42.

[0076] On the liquid oxygen pipeline, the first self-sealing valve and the second self-sealing valve are respectively located at one end near the disposable core stage 1 and the other end near the external recovery module 2, and are used to immediately close the liquid oxygen pipeline after the engine is turned off, sealing the liquid oxygen on both sides of the valve; the discharge valve assembly 41 is located on the liquid oxygen pipeline between the first self-sealing valve and the second self-sealing valve, and is used to discharge the liquid oxygen in this section of the liquid oxygen pipeline after the two valves are closed.

[0077] On the fuel line, the first self-sealing valve and the second self-sealing valve are respectively located at one end near the disposable core stage 1 and the other end near the external recovery module 2, and are used to immediately close the fuel line after the engine is turned off, sealing the fuel on both sides of the valve; the exhaust valve assembly 41 is located on the fuel line between the first self-sealing valve and the second self-sealing valve, and is used to empty the fuel in this section of the fuel line after the two valves are closed.

[0078] The cutting device 42 is a frame-shaped structure surrounding all liquid oxygen and fuel pipelines. It is used to simultaneously cut off all liquid oxygen and fuel pipelines in a single detonation after the propellant in the liquid oxygen and fuel pipelines has been emptied. This frame-shaped structure replaces the single-pipe cutting ring, solving the problem of synchronous cutting when multiple pipelines are discretely arranged, ensuring that all propellant delivery pipelines within the same symmetrical group are disconnected simultaneously. The detonating cord of the cutting device 42 is externally constrained by a sheath to direct the detonation products outwards when the detonating cord is detonated. The detonating cord also has a sacrificial protective ring to absorb residual detonation energy and prevent damage to surrounding structures. A distance of at least five times the pipe diameter is maintained between the detonating cord cut and the sealing surfaces of the valve seats on both sides to ensure that the sealing surfaces are not affected by the detonation wave, guaranteeing the reliability of the seal after the self-sealing valve is closed.

[0079] In some embodiments, the second self-sealing valve and the remaining section of the disconnected propellant delivery pipeline connected thereto constitute an integrated quick-change assembly that can be replaced as a whole after being recovered by the external recovery module 2.

[0080] Specifically, the quick-change assembly and the propellant delivery pipeline within the external recovery module 2 are connected by a detachable mechanical connection using flanges with cryogenic metal seals or vacuum clamps. Welding, non-destructive testing, and pressure testing of the quick-change assembly are pre-completed in the workshop. After the external recovery module 2 is recovered, replacement is completed simply by pulling out the old quick-change assembly, inserting the new one, and performing a helium mass spectrometry leak test; no welding work is required inside the module compartment. The entire replacement process takes only hours, significantly reducing the reuse preparation time for the external recovery module 2 to be remounted and perform tasks, and lowering the complexity and labor costs of maintenance operations.

[0081] In some embodiments, the external payload recovery modules 2 are arranged in groups of two, radially symmetrically, around the core stage: each pair of external payload recovery modules 2 is located on both sides of the core stage along the same radial axis, forming a symmetrical group; depending on the core stage size and total thrust requirements, one group of two modules, two groups of four modules (two symmetrical axes are 90° apart), or three groups of six modules (adjacent symmetrical axes are 60° apart) can be set. The symmetrical grouping layout enables the thrust and aerodynamic torque on each radial axis to self-balance, eliminating the rolling torque and transonic shock wave interference generated by unilateral external payloads, ensuring attitude stability during the first stage flight phase; the even-number grouping expansion method allows the external payload thrust to be scaled up proportionally with the core stage size without changing the basic layout principles and module configuration. The propellant tank 11 simultaneously supplies propellant to the main engine 21 of each external payload recovery module 2 through the propellant delivery pipeline.

[0082] It is important to emphasize that the symmetrical group layout is not simply an adjustment of positional relationships. After the number of modules is expanded from a single module to 2, 4, or 6 symmetrical group configurations, the three sets of interface structures for supply, load bearing, and release are restructured accordingly, thereby bringing about quantifiable technological advancements: (1) Equal flow resistance grouped supply system. The single-module layout only requires one liquid oxygen pipeline and one fuel pipeline, and there is no distribution problem; under the multi-group layout, if multiple pipelines of unequal lengths are directly led out from the propellant tank 11, the flow resistance difference of each branch will cause the pressure deviation before the pump of each external recovery module 2 to be more than 5% and the thrust imbalance to reach 5% to 8%, which requires active adjustment and sacrificing propellant for balancing. To this end, a ring-shaped distribution manifold 5 is provided on the disposable core stage 1. The ring-shaped distribution manifold 5 is arranged around the periphery of the disposable core stage 1, and the propellant tank 11 is connected to the inlet of the ring-shaped distribution manifold 5. There are multiple propellant delivery pipelines of equal length. The ring-shaped distribution manifold 5 is connected to the main engine 21 of each external recovery module 2 through the multiple propellant delivery pipelines. A flow resistance matching component is provided at the inlet of each propellant delivery pipeline. Each symmetrical group of propellant delivery pipelines is equipped with a group independent shut-off valve to cooperate with the sealing of the unseparated group when the group is separated in sequence, so that the pump pressure deviation of each external recovery module 2 is less than 2% and the thrust imbalance is less than 1.5%, without the need for propellant balancing loss.

[0083] (2) Symmetrical groups share a common through-load beam 31. When mounted on one side, the separation reaction force of the catapult thruster is transmitted to the thin wall of the tank as a lateral concentrated load. The mounting area needs to be reinforced with a reinforcing frame and a stringer. In the single-module layout, a counterweight is also required to adjust the center of gravity of the entire rocket. In this embodiment, the two external recovery modules 2 of each symmetrical group are mounted on both ends of the same through-load beam 31 that radially penetrates the disposable core stage 1. The catapult thruster is installed on the through-load beam 31 to apply thrusting forces in opposite directions to the two external recovery modules 2 on the same through-load beam 31 at the same time. The through-load beam 31 and the disposable core stage 1 are connected by a support that only transmits shear force, so that the separation reaction force of the two external recovery modules 2 in the same symmetrical group is closed within the group along the corresponding through-load beam 31 and is not transmitted to the thin wall of the tank, and no counterweight is required.

[0084] (3) Grouped integrated fluid and electrical disconnect panel 6. A single module has only two fluid connectors, which can be arranged discretely; the layout of four or six modules increases the number of fluid connectors to 8 to 12 and includes cable bundles. The three major problems of synchronous distribution of discrete connector cutting, non-coplanar separation surfaces and multi-point residual liquid splashing make it physically infeasible. Therefore, each external recovery module 2 has a release panel 6 fixed on its outer wall at the position between the first and second self-sealing valves. All fluid connectors and electrical release connectors of each symmetrical group are integrated on a release panel 6 within the same separation surface. The cutting device 42 has a detonating cord, which is upgraded from a cutting ring around a single pipe to a cutting frame around all connectors of the release panel 6, and the entire group is cut off in one detonation. The propellant delivery pipeline passes through the release panel 6 in the middle and is arranged in parallel with the electrical cables. The release panel 6 is taken away with the external recovery module 2, and its remaining opening is completely sealed by the pipeline compartment door 43. Each symmetrical group is equipped with an independent discharge valve assembly 41, which has a directional nozzle. The directional nozzles of the discharge valve assemblies 41 of the two external recovery modules 2 in the same symmetrical group are arranged symmetrically, and the discharge reaction force is self-balanced.

[0085] (4) Fault-based degradation control. If any external payload recovery module 2 malfunctions during flight, the symmetrical modules in its symmetrical group are shut down, the overall thrust of the rocket remains balanced, and (N-2) / N thrust is retained for continued degradation flight, where N represents the number of external payload recovery modules 2. At the same time, the grouped sequential separation halves the impact energy released in each separation, and the peak value of the impact response spectrum at the inertial measurement unit mounting point is approximately 0.5 to 0.6 times that of all modules separating simultaneously.

[0086] The quantitative comparison between symmetrical grouped layouts and single-sided discrete layouts is shown in the table below.

[0087]

[0088] Note: The data in the table above are the design goals and magnitude estimates of this scheme. The specific values ​​are subject to the detailed design of the model.

[0089] In some embodiments, the inertial measurement unit 222 is installed with flexible vibration damping, and the impact response spectrum limit of the installation point is incorporated into the design requirements. Before separation, an attitude reference is set. After separation, navigation is maintained for 1 to 2 seconds using a combination of magnetometer, barometric altimeter, and satellite navigation. After zero-bias online re-estimation, the inertial measurement unit switches back to integrated inertial navigation. The inertial measurement unit is selected as a missile-grade product with an impact-resistant range. The impulses of each separation cancel each other out within the symmetrical group, minimizing the residual attitude disturbance of the core stage 1. Each external payload recovery module 2 obtains determined initial separation conditions. After separation, each module 2 glides independently along a predetermined route using its respective flight control equipment 22. Multiple external payload recovery modules 2 can enter the same landing site sequentially at predetermined intervals, achieving orderly recovery of multiple modules.

[0090] In some embodiments, when the external recovery module 2 is applied to a single stage, the separation velocity is in the suborbital range, with a Mach number of approximately 4 to 8 and a low peak heat flux density. The thermal protection skin can be made of high-temperature resistant alloy skin or a thin-layer ablation material, eliminating the need for ceramic tiles or complex insulation structures. The external recovery module 2 begins reentry gliding from a separation altitude of over 80 km. Based on a lift-to-drag ratio of 3 to 4, its gliding range is estimated at approximately 240 to 320 km, which can cover the distance required for returning to a domestic landing site from the separation point.

[0091] In some embodiments, the basic return mode of the external payload recovery module 2 is unpowered gliding, without a return engine, aviation fuel, or related systems, resulting in a simple structure and fewer reuse preparation items. The external payload recovery module 2 utilizes its own lifting body aerodynamic shape to glide and decelerate within the atmosphere, using aerodynamic control surfaces to control its heading and glide path. In the final stage, it deploys its landing gear and performs a horizontal landing on the runway. The impact load of a horizontal landing is far lower than that of a vertical thrust-reverse landing, and the module's shape is consistent with conventional aircraft, significantly simplifying the reuse preparation process for inspection, maintenance, and re-attachment.

[0092] In some embodiments, the external recovery module 2 is also equipped with an electric ducted fan as an auxiliary return power source. The electric ducted fan is powered by an onboard battery and is activated when gliding to the vicinity of the landing site to provide cruise power to extend the return distance or correct the glide path. The electric ducted fan does not require aviation fuel or an ignition starting device, starts quickly, is easy to maintain, and is not subjected to a high-temperature gas environment during reentry.

[0093] In some embodiments, when the external recovery module 2 is applied to the second stage, the separation speed is at the orbital level, and the reentry thermal environment is harsh. The external recovery module 2 adopts complete lifting body thermal protection, and disperses the peak heat flow over a longer time segment through high lift-to-drag ratio gliding. Since the heat protection object is only a small volume module, the heat protection area is reduced by about an order of magnitude compared to whole-stage recovery, reducing the thermal protection requirements of orbital-level recovery to an engineering-feasible range.

[0094] Another aspect of the present invention provides a launch and recovery method based on an externally mounted partially reusable launch vehicle system, comprising the following steps: S1: The main engine 21 of the external recovery module 2 provides thrust to drive the launch vehicle to ascend, and the propellant tank 11 supplies propellant to the main engine 21 through the propellant delivery pipeline; S2: After the first stage flight is completed, the main engine 21 shuts down at an altitude of over 50km, and the first and second stages are separated. The external recovery module 2 glides along with the disposable core stage 1. S3: When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the self-disconnecting connector 4 disconnects the propellant delivery pipeline, the rack and separation mechanism 3 are unlocked, and the external recovery module 2 is separated from the disposable core stage 1. S4: External recovery module 2 returns via a gliding motion and lands horizontally on the runway. The core stage 1 is passivated and emptied before falling.

[0095] The launch and recovery method of the externally mounted partially reusable launch vehicle system provided by this invention is based on the core concept of functional decoupling of the first stage of the launch vehicle according to value density. The high-value main engine 21 and flight control equipment 22 are integrated into the externally mounted recovery module 2, while the low-value propellant tank 11 is retained in the disposable core stage 1. Propellant is supplied from the core stage tank to the externally mounted module engine during flight, so that the power and the tank are physically separated but functionally coordinated, thereby recovering the highest hardware value with the lowest reentry heat protection cost. In step S2, after the main engine 21 shuts down at an altitude above 50km, the first and second stages separate. The module glides with the inertia of the disposable core stage 1, using the remaining kinetic energy of the core stage to carry the module to a separation altitude above 80km. The module can complete the ascent without consuming its own propellant, thus maximizing the energy margin required for the module's return. In step S3, when the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the self-disconnecting connector 4 disconnects the propellant delivery pipeline, using a near-vacuum environment to cut off the pipeline, thus eliminating the risks of high dynamic pressure flutter and oxygen-rich deflagration from both timing and environmental conditions. After the pylon and separation mechanism 3 are unlocked, the module separates from the core stage. The aerodynamic interference at the separation point is negligible, ensuring that the module obtains a definite initial separation attitude and velocity vector, establishing precise initial conditions for subsequent gliding return. In step S4, the module returns by gliding as a lifting body and lands horizontally on the runway, using aerodynamic lift to achieve a long-distance unpowered return, without the need to carry a return engine and aviation fuel. The module has a simple structure and requires fewer reusable maintenance items. The core stage 1 is passivated and emptied before falling, and the core stage does not take thermal protection measures throughout the process, fundamentally avoiding the engineering problem of paying high thermal protection costs for large surface area structures in whole-stage recovery.

[0096] In some embodiments, in step S2: the main engine 21 is synchronously shut down at an altitude above 50km; the first and second self-sealing valves of the self-disconnecting connector 4 are closed, and the propellant is sealed on the outside of the first and second self-sealing valves respectively; after the first and second stage separation is implemented, the external recovery module 2 slides along with the inertia of the disposable core stage 1; In step S3: when the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly of the self-severing connector 4 opens to drain the propellant in the pipe section between the first self-sealing valve and the second self-sealing valve; after the pressure in the pipe drops below the set threshold, the cutting device cuts off the propellant delivery pipeline; the hanger and separation mechanism 3 are unlocked, and the ejector pusher pushes the external recovery module 2 away from the disposable core stage 1; After the external recovery module 2 is separated from the disposable core stage 1, the pipeline compartment door 43 is closed and locked, sealing the break in the propellant delivery pipeline; In step S4: After the external recovery module 2 lands and is recovered, the integrated quick-change assembly consisting of the second self-sealing valve and the remaining pipe section of the propellant delivery pipeline connected to it is replaced as a whole and leak-checked for reuse.

[0097] By breaking down the overall "disconnection" function of the self-disconnecting connector 4 into three sequential sub-steps—valve closure, pipeline evacuation, and detonating cord cutting—and by limiting the first and second self-sealing valves to be simultaneously shut down above 50km in step S2, and sealing the propellant on the outside of the two valves respectively, the propellant is cut off on both sides of the core stage and module, avoiding disturbance to the separation process caused by continuous propellant flow or backflow before pipeline cutting. In step S3, the opening conditions for the discharge valve assembly 41 are limited to an altitude greater than 80km or a dynamic pressure less than 200Pa, the evacuation target being the propellant in the pipeline section between the first and second self-sealing valves, and the precondition for starting the cutting device 42 is that the pressure in the pipeline drops below a set threshold. This forms a complete safety logic of "sealing first, then evacuating, pressure confirmation, and then cutting off," using the pipeline pressure threshold as an objective criterion to ensure that the pipeline is emptied before cutting off, fundamentally eliminating the risk of deflagration caused by residual propellant in the pipeline coming into contact with high-temperature gas when the detonating cord detonates. At the same time, step S3 also limits the opening conditions for the discharge valve assembly 41 to be opened at an altitude greater than 80km or a dynamic pressure less than 200Pa, the evacuation target being the propellant in the pipeline section between the first and second self-sealing valves, and the precondition for starting the cutting device 42 is that the pressure in the pipeline drops below a set threshold. After the mounting bracket and separation mechanism 3 are unlocked, the ejection thruster pushes the external recovery module 2 away from the disposable core stage 1, enabling the module to obtain a defined separation velocity and initial attitude, ensuring the accuracy of the initial conditions for gliding return. After separation, the pipeline compartment door 43 closes and locks, sealing the propellant delivery pipeline break, so that the outer surface of the module restores its continuous aerodynamic shape before reentry, avoiding the adverse effects of airflow separation at the break on gliding stability. In step S4, it is further specified that after recovery, the integrated quick-change assembly consisting of the second self-sealing valve and the remaining section of the propellant delivery pipeline connected to it is replaced as a whole and leak-checked for reuse. This allows the pipeline interface to be reused and prepared without entering the compartment for welding work after the module is recovered. The steps form a complete closed loop from engine shutdown, propellant plugging, high-altitude gliding, venting, pressure confirmation, pipeline cutting, ejection separation, aerodynamic shape restoration to quick-change assembly replacement and leak checking. The timing and condition constraints are clear, ensuring the reliable operation and rapid reuse economy of the recovery system in complex flight environments.

[0098] In some embodiments, the case of setting up a set (2 modules) of external recovery module 2 for the first stage application of a two-stage liquid-fueled launch vehicle is illustrated as an example. Figure 1 and Figure 2As shown, two external recovery modules 2 are symmetrically installed on both sides of the tail section of the disposable core stage 1 via brackets and separation mechanisms 3. Each external recovery module 2 integrates one or more main engines 21 and all high-value subsystems. The propellant tank 11 of the disposable core stage 1 supplies propellant to the main engines 21 of the external recovery modules 2 on both sides through a propellant delivery pipeline equipped with a self-breaking connector 4. When two groups (4 modules) or three groups (6 modules) are used, each symmetrical group is distributed circumferentially along the disposable core stage 1. When two groups are set, the two axes of symmetry are 90° to each other, and when three groups are set, the adjacent axes of symmetry are 60° to each other. The mechanical connection, propellant supply and separation methods with the disposable core stage 1 are the same as in this embodiment.

[0099] Specifically, the two external recovery modules 2 share a single radially penetrating load-bearing beam 31 that runs through the disposable core stage 1. The ejection separation reaction force closes within the group along the load-bearing beam 31 without being transmitted to the thin wall of the propellant tank. The load-bearing beam 31 is connected to the disposable core stage 1 by a support that only transmits shear force. The disposable core stage 1 is equipped with an annular distribution manifold 5, which is arranged circumferentially along the disposable core stage 1. The propellant tank 11 is connected to the inlet of the annular distribution manifold 5. There are multiple propellant delivery pipelines of equal length. The annular distribution manifold 5 is connected to the main engine 21 of each external recovery module 2 through the multiple propellant delivery pipelines to form an equal flow resistance group supply system for uniform propellant supply. Each external recovery module 2 has a release panel 6 fixed on its outer wall at the position between the first self-sealing valve and the second self-sealing valve. All fluid interfaces and electrical interfaces are integrated on the release panel 6. The cutting device 42 has a detonating cord, which is arranged around the propellant delivery pipeline and located between the release panel 6 and the first self-sealing valve. The entire set is cut off and released in one detonation.

[0100] Specifically, the pipeline compartment door 43 is located on the outer wall of the external recovery module 2 to seal the pipeline break after the propellant delivery pipeline is disconnected, thereby maintaining the aerodynamic shape of the outer surface of the external recovery module 2. The discharge valve assembly 41 has a directional nozzle, and the directional nozzles of the discharge valve assemblies 41 of the two external recovery modules 2 in the same symmetrical group are symmetrically arranged, so that the discharge reaction force is self-balanced.

[0101] The flight proceeds in the following sequence: Step ①: The launch vehicle takes off vertically. All or most of the thrust of the first stage is provided by the main engine 21 of the external recovery module 2 on both sides. The propellant is continuously supplied from the propellant tank 11 of the disposable core stage 1 through the propellant delivery pipeline equipped with self-disconnecting connector 4. The symmetrical layout enables the thrust and aerodynamic torque to be self-balanced, and the rolling torque generated by the external module on one side can be eliminated without the need for counterweight.

[0102] Step ②: After the first stage of flight ends, the main engines 21 of the external recovery modules 2 on both sides shut down synchronously at an altitude greater than 50km, with a flight Mach number of about 4 to 8, which is in the suborbital range; the first and second self-sealing valves of the self-disconnecting connector 4 then close, sealing the propellant on the outside of the first and second self-sealing valves respectively.

[0103] Step 3: Implement first and second stage separation, and the second stage continues to fly; after the first and second stage separation, the external recovery modules 2 on both sides continue to glide along with the disposable core stage 1 assembly. The apex altitude of the assembly is greater than 80km. At this time, about one to two minutes have passed since the main engine 21 was shut down, and the aftereffect thrust has decayed to a negligible level.

[0104] Step 4: When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly 41 of the self-severing connector 4 opens. The liquid oxygen trapped in the pipe section between the first and second self-sealing valves automatically flashes under the dual drive of its own saturated vapor pressure and the pipe section trapping pressure, and is discharged through the directional nozzle of the discharge valve assembly 41. The residual propellant on the fuel side is blown out and emptied by the trapping air pressure; the propellant tank 11 of the disposable core stage 1 is simultaneously passivated and emptied. After the pressure inside the pipe drops below the set threshold, the cutting device 42 of the self-severing connector 4 detonates, cutting off the emptied propellant delivery pipeline as a whole within milliseconds; the hangar and separation mechanism 3 are unlocked, and the external recovery modules 2 on both sides are simultaneously pushed away from the disposable core stage 1 by the ejector thrusters installed on the through load-bearing beam 31 through the symmetrical detonation network. The separation point is in a near-vacuum environment with a dynamic pressure of less than 200Pa and negligible aerodynamic interference. The external recovery modules 2 obtain a definite separation speed and initial attitude. After the external recovery module 2 is separated from the disposable core stage 1, the pipeline compartment door 43 is closed and locked at the center dead point, sealing the break left on the surface of the propellant delivery pipeline, so that the outer surface of the external recovery module 2 can be restored to a continuous aerodynamic shape.

[0105] Step ③: The disposable core stage 1 is depressurized and emptied by the passivation system (the propellant tank 11 is emptied simultaneously with step ④), and falls uncontrolled along the trajectory to the predetermined sea or uninhabited landing area; the disposable core stage 1 is not thermally protected throughout the entire process.

[0106] Step 5: From a separation altitude of over 80km, the external recovery modules 2 on both sides re-enter and glide to decelerate using their lifting body aerodynamic shape. The flight control equipment 22 corrects the course and glide path via aerodynamic control surfaces based on data from the inertial measurement unit. The peak re-entry heat flux density is low and is absorbed by the thermal protection skin. The basic return method of the external recovery modules 2 is unpowered gliding, eliminating the need for a return engine and aviation fuel. Utilizing their own lifting body aerodynamic shape, they glide and decelerate within the atmosphere. Based on a lift-to-drag ratio of 3 to 4, the gliding range is estimated at approximately 240 to 320km, which can cover the distance required for returning to the domestic landing site from the separation point.

[0107] Step 6: External payload recovery module 2 flies to the landing site, deploys its landing gear, and performs a horizontal landing on the runway. The impact load of a horizontal landing is much lower than that of a vertical thrust-reverse landing, and the module's shape is consistent with that of a conventional aircraft, significantly simplifying the reuse preparation process for inspection, maintenance, and remounting. After landing, the second self-sealing valve and the remaining section of the propellant delivery pipeline connected to it form an integrated quick-change assembly, which is replaced and leak-checked as a whole, without the need for welding work inside the cabin. The reuse preparation time is measured in hours. After inspection and preparation, external payload recovery module 2 can be remounted to a new disposable core stage 1 to perform the next mission.

[0108] The above embodiments of the present invention do not constitute a limitation on the scope of protection, and those skilled in the art can make the following equivalent substitutions or modifications: (1) Number of modules: The external recovery modules 2 are arranged in groups of two, radially symmetrically. Preferably, there are 2 modules in one group, 4 modules in two groups (the two axes of symmetry are 90° to each other), or 6 modules in three groups (adjacent axes of symmetry are 60° to each other). Larger core stages can use more than three groups. The separation method can be that all external recovery modules 2 are separated at the same time, or separated in a symmetrical group sequence - each symmetrical group is separated in sequence according to the set time, and only one symmetrical group is separated at a time. The separation interval between adjacent groups is not less than 100 milliseconds. As a variation, a single module layout can also be used, but the center of mass and pressure center must be balanced to compensate for aerodynamic asymmetry.

[0109] (2) Propellant type: The propellant tank 11 of the disposable core stage 1 and the main engine 21 of the external recovery module 2 can be adapted to different propellant combinations such as liquid oxygen / kerosene and liquid oxygen / methane. The sealing material and structural parameters of the self-breaking joint 4 are adjusted accordingly.

[0110] (3) Aerodynamic configuration: The aerodynamic shape of the lifting body can be different forms such as blended wing-body, delta wing or winged column. The number and arrangement of aerodynamic control surfaces such as flaps and rudders can be adapted to the design.

[0111] (4) Return method: The basic plan is to return by gliding without power; or an electric ducted fan powered by the airborne battery can be set up as an auxiliary power source for the return, which is activated when gliding to the vicinity of the landing site, so as to extend the return distance, improve the glide path correction capability and weather adaptability.

[0112] (5) Landing method: The landing gear can be wheeled or skid-mounted; in the recovery area where there is no runway, arresting nets or offshore platforms can be used to assist in landing.

[0113] (6) Installation position: In addition to being installed at the tail section of the disposable core stage 1, the external recovery module 2 can also be adjusted along the axial direction of the disposable core stage 1 to match the distribution requirements of the entire rocket's center of mass and aerodynamic pressure center.

[0114] (7) Pipeline cutting method: Detonating cord cutting is the preferred option for the cutting device 42; alternatively, a cold cutting device such as a high-pressure gas-driven annular cutter or mechanical shearing can be used to adapt to occasions where the use of pyrotechnics is restricted.

[0115] (8) Emission and separation criteria: The primary criteria are an altitude greater than 80 km or a dynamic pressure less than 200 Pa, calculated from the inertial navigation altitude and dynamic pressure; alternatively, the internal pressure sensor or flight time can be used as backup criteria. The discharge valve assembly 41 can also serve as an interface for discharge and purging after ground refueling.

[0116] (9) Supply and load-bearing structure: The ring distribution main pipe 5 can be arranged on the outer wall of the tail section of the one-time core stage 1 or integrated into the internal structure of the tail section; the through load-bearing beam 31 can be a whole beam or a combination of segmented beam frames; the small diameter core stage can also use back-to-back hanging frames instead of the through load-bearing beam 31, but the transmission path of the separation reaction force must be demonstrated.

[0117] (10) Deformation of pipeline compartment door: The pipeline compartment door 43 can also adopt a sliding cover structure that slides and closes in the direction of airflow, and is closed and locked by a spring or actuator.

[0118] In the description of this specification, the references to terms such as "one embodiment," "some embodiments," "example," "specific example," or "some examples," etc., indicate that a specific feature, structure, material, or characteristic described in connection with that embodiment or example is included in at least one embodiment or example of the invention. In this specification, the illustrative expressions of the above terms do not necessarily refer to the same embodiment or example. Furthermore, the specific features, structures, materials, or characteristics described may be combined in any suitable manner in one or more embodiments or examples. In addition, those skilled in the art can combine and integrate the different embodiments or examples described in this specification.

[0119] Although embodiments of the present invention have been shown and described above, it is understood that the above embodiments are exemplary and should not be construed as limiting the present invention. Those skilled in the art can make changes, modifications, substitutions and variations to the above embodiments within the scope of the present invention.

Claims

1. An externally mounted, partially reusable launch vehicle system, characterized in that: include: A single-use core stage (1) includes at least a propellant tank (11); as well as At least one external recovery module (2) is provided with a main engine (21) and flight control equipment (22), has a lifting body aerodynamic shape, and is detachably installed on the outside of the disposable core stage (1) via a rack and separation mechanism (3); The propellant tank (11) is connected to the main engine (21) of the external recovery module (2) through a propellant delivery pipeline equipped with a self-disconnecting connector (4) to supply propellant; the self-disconnecting connector (4) is used to disconnect the propellant delivery pipeline when the external recovery module (2) separates from the disposable core stage (1); after the external recovery module (2) separates from the disposable core stage (1), it can return and land horizontally in a lifting body gliding manner.

2. The externally mounted partially reusable launch vehicle system according to claim 1, characterized in that, The self-disconnecting connector (4) is located on the propellant delivery pipeline between the disposable core stage (1) and the external recovery module (2); The self-breaking connector (4) includes: The first self-sealing valve and the second self-sealing valve are respectively located at one end of the propellant delivery pipeline near the disposable core stage (1) and the other end near the external recovery module (2); A discharge valve assembly (41) is disposed on the propellant delivery pipeline between the first self-sealing valve and the second self-sealing valve, for discharging the propellant in the propellant delivery pipeline between the first self-sealing valve and the second self-sealing valve after the first self-sealing valve and the second self-sealing valve are closed; A cutting device (42) is provided on the propellant delivery pipeline and located between the first self-sealing valve and the second self-sealing valve, for disconnecting the propellant delivery pipeline after the propellant is emptied.

3. The externally mounted partially reusable launch vehicle system according to claim 2, characterized in that, The self-disconnecting connector (4) also includes a pipeline compartment door (43); the pipeline compartment door (43) is disposed on the outer wall of the external recovery module (2) and is used to close the pipeline break after the propellant delivery pipeline is disconnected, so as to maintain the aerodynamic shape of the outer surface of the external recovery module (2).

4. The externally mounted partially reusable launch vehicle system according to claim 2, characterized in that, The number of external recycling modules (2) is even, and each pair of external recycling modules (2) forms a symmetrical group. The two external recycling modules (2) in the same symmetrical group are arranged symmetrically on opposite sides of the disposable core level (1) along the radial direction. Each external recycling module (2) is installed on the disposable core level (1) through the bracket and separation mechanism (3).

5. The externally mounted partially reusable launch vehicle system according to claim 4, characterized in that, The hanger and separation mechanism (3) includes a through load-bearing beam (31) and a catapult thruster; The through-bracing beam (31) is radially arranged through the disposable core stage (1), and the two external recycling modules (2) in each of the symmetrical groups are mounted on both ends of the same through-bracing beam (31); The ejector is installed on the through-bracing beam (31) and is used to simultaneously apply opposite thrusts to the two external recovery modules (2) on the same through-bracing beam (31) so that the two external recovery modules (2) on the same through-bracing beam (31) leave the disposable core stage (1) synchronously. The through-bracing beam (31) and the disposable core stage (1) are connected by a support that transmits only shear force, so that the separation reaction force of the two external recycling modules (2) in the same symmetrical group is closed along the corresponding through-bracing beam (31) within the group.

6. The externally mounted partially reusable launch vehicle system according to claim 5, characterized in that, The disposable core stage (1) is provided with an annular distribution manifold (5), which is arranged around the periphery of the disposable core stage (1), and the propellant tank (11) is connected to the inlet of the annular distribution manifold (5). The propellant delivery pipeline consists of multiple pipelines of equal length. The annular distribution main pipe (5) is connected to the main engine (21) of each of the external recovery modules (2) through the multiple propellant delivery pipelines to form an equal flow resistance grouping supply system for uniform propellant supply.

7. The externally mounted partially reusable launch vehicle system according to claim 6, characterized in that, Each of the external recycling modules (2) has a release panel (6) fixed on its outer wall at the position between the first self-sealing valve and the second self-sealing valve. The release panel (6) is provided with an electrical release connector. One end of the electrical release connector is electrically connected to the external recycling module (2), and the other end is electrically connected to the disposable core (1) through an electrical cable. The electrical release connector can be released and disconnected when the external recycling module (2) is separated from the disposable core (1). The cutting device (42) has a detonating cord that surrounds the propellant delivery pipeline and is located between the pull-out panel (6) and the first self-sealing valve, for cutting off the propellant delivery pipeline; The propellant delivery pipeline passes through the pull-out panel (6) in the middle and is arranged in a concentrated manner alongside the electrical cables.

8. The externally mounted partially reusable launch vehicle system according to claim 6, characterized in that, The second self-sealing valve and the remaining section of the disconnected propellant delivery pipeline connected thereto constitute an integrated quick-change assembly that can be replaced as a whole after being recovered by the external recovery module (2).

9. A method for launching and recovering an externally mounted partially reusable launch vehicle system according to any one of claims 1-8, characterized in that, Includes the following steps: S1: The main engine (21) of the external recovery module (2) provides thrust to drive the launch vehicle to ascend, and the propellant tank (11) supplies propellant to the main engine (21) through the propellant delivery pipeline; S2: After the first stage of flight is completed, the main engine (21) shuts down at an altitude of over 50km, and the first and second stages are separated. The external recovery module (2) glides along with the disposable core stage (1) by inertia. S3: When the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the self-disconnecting connector (4) disconnects the propellant delivery pipeline, the hanger and separation mechanism (3) unlocks, and the external recovery module (2) separates from the disposable core stage (1); S4: The external recovery module (2) returns in a gliding manner and lands horizontally on the runway, and the disposable core stage (1) is passivated and emptied before falling.

10. The launch and recovery method for an externally mounted partially reusable launch vehicle system according to claim 9, characterized in that, In step S2: the main engine (21) is shut down synchronously at an altitude of over 50km; the first and second self-sealing valves of the self-disconnecting connector (4) are closed, and the propellant is sealed on the outside of the first and second self-sealing valves respectively; after the first and second separation is implemented, the external recovery module (2) slides along with the disposable core stage (1) by inertia; In step S3: when the flight altitude is greater than 80km or the dynamic pressure is less than 200Pa, the discharge valve assembly of the self-breaking connector (4) is opened to drain the propellant in the pipe section between the first self-sealing valve and the second self-sealing valve; after the pressure in the pipe drops below the set threshold, the cutting device cuts off the propellant delivery pipeline; the hanger and separation mechanism (3) are unlocked, and the ejector pusher pushes the external recovery module (2) away from the disposable core stage (1). After the external recovery module (2) is separated from the disposable core stage (1), the pipeline compartment door (43) is closed and locked, sealing the break in the propellant delivery pipeline; In step S4: After the external recovery module (2) lands and is recovered, the integrated quick-change assembly consisting of the second self-sealing valve and the residual pipe section of the propellant delivery pipeline connected to it is replaced as a whole and leak checked for reuse.

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