844results about "Rocket engine plants" patented technology

The invention discloses a sealing device for a jet pipe of a solid rocket engine. The sealing device comprises a front supporting body, a rear supporting cylinder and a sealing plug cover. The front section of the front supporting body is a circular-truncated-cone-shaped positioning cylinder, the rear section of the front supporting body is a circular-ring-shaped connecting cylinder, and internal threads are machined on the rear section of the inner wall of the circular-ring-shaped connecting cylinder. An external thread is machined on the outer wall of the rear supporting cylinder, a positioning circular plate is vertically fixed to the rear end of the rear supporting cylinder, and the outer edge of the positioning circular plate is machined to be a circular truncated cone face. The front supporting body and the rear supporting cylinder are inserted inwards from the two ends of the spray pipe respectively and then are in threaded butt joint, the outer wall of the front supporting body is tightly attached to the inner wall of a convergence section of the spray pipe and the inner wall of a spray throat, and the outer edge of a positioning circular plate of the rear supporting cylinder is tightly attached to the inner wall of an expansion section of the spray pipe. And the sealing blanking cap is in threaded connection with the front end of the rear supporting cylinder. According to the spray pipe sealing device, ignition pressure can be reliably established, and residues are few during separation.

The invention provides a carrier rocket ground test system and method supporting visual process configuration. The system mainly comprises: (1) a visual process configuration module, which configures a test process file through a graphical interface; (2) a process analysis and scheduling module for analyzing a process file and converting the process file into a task sequence; the test execution engine module (3) is used for driving bottom hardware and executing a test instruction sequence; (4) the test resource library module is used for storing and managing various information of test elements and various information of test equipment; the execution monitoring and data management module (5) is used for displaying a test execution state in real time and recording whole-process data; according to the method, the test logic and the execution program are separated, and a user can quickly construct, modify and reuse an automatic test process in a graphical mode, so that the test flexibility is greatly improved, and the test adaptability and efficiency are remarkably improved.

The invention discloses a circumferential liquid film surface wave observation device and method based on a miniature high-speed camera. The circumferential liquid film surface wave observation device comprises a centrifugal nozzle, a laser, the high-speed miniature camera and a synchronizer, the centrifugal nozzle can spray a conical liquid film; the high-speed miniature camera is located at the central gas nucleus of the conical liquid film, and the optical axis coincides with the axis of the centrifugal nozzle; a narrow-band optical filter is arranged on the surface of an optical lens of the high-speed miniature camera; the laser can emit plane laser perpendicular to the axis of the centrifugal nozzle to the conical liquid film; wherein the wavelength of the planar laser is equal to the central wavelength of the narrow-band optical filter; and the synchronizer is respectively connected with the laser and the high-speed miniature camera. According to the method, a frozen instantaneous circumferential liquid film form without background interference can be obtained, a surface wave growth factor and a breaking critical criterion can be accurately quantified, and the development and evolution process of the circumferential surface wave of the liquid film is further obtained through high-speed shooting.

A reusable upper stage rocket or other atmospheric re-entry vehicle includes a nose, a base opposite the nose, and a propulsion engine toward the base. The propulsion engine includes a high pressure chamber and a nozzle configured to exhaust gas generated by the high pressure chamber. The nozzle includes an initial nozzle portion, a secondary nozzle portion downstream of the initial nozzle portion, and a nozzle exit at a downstream end of the secondary nozzle portion. The secondary nozzle portion includes an inner expansion surface, an outer expansion surface outboard of the inner expansion surface, and an expansion cavity defined between the inner expansion surface and the outer expansion surface.

A gas turbine engine including a fuel injector having an outer valve body with an inner passage, a wall at least partially enclosing the inner passage, the wall having an inner surface, an outer surface, and one or more channels passing through the wall between the outer surface of the wall and the inner surface of the wall, and a retractable valve body positioned within the inner passage of the outer valve body. The retractable valve body has one or more slots in fluid communication with the fluid spin chamber. The retractable valve body is positioned within the inner passage to move along a reciprocating axis at one or more valve positions including an advanced valve position, a retracted valve position, an open valve position, and a closed valve position.

The energy recycling heat control device provided by the embodiment of the present application comprises a liquid collecting chamber, an engine assembly and a spring, the engine assembly and the spring are arranged between the liquid collecting chamber and an injector, the pushing action of a piston by the target gas filling parameters in a sealed space of the engine assembly is combined with the spring mechanical parameters of the spring to control the distance between a heat plate and the outer wall of the liquid collecting chamber, so that the heat energy is recycled and reused, which is beneficial to improve the rapid response speed of the engine, reduce the power consumption, and thus reduce the risk of explosion after shutdown.

The invention provides a high-power two-stage pump axial force adjusting device with a complex leakage channel. The high-power two-stage pump axial force adjusting device comprises a first-stage pump and a second-stage pump. A primary pump impeller is installed on the main shaft and comprises a primary pump impeller front cover plate, the primary pump impeller front cover plate is coaxially sleeved with a primary pump front floating ring, and axial force adjustment is achieved by adjusting the flow of a primary pump front leakage channel through a radial sealing gap of the primary pump front floating ring. The first-stage pump impeller rear cover plate comprises a first-stage rear cover plate body, the axial rear end face of the first-stage rear cover plate body integrally extends backwards in the axial direction to form a first-stage throttling convex shoulder, and the first-stage throttling convex shoulder is coaxially sleeved with a first-stage pump rear floating ring. The first-stage throttling convex shoulder divides a first-stage pump rear leakage channel into a high-pressure area before floating ring leakage and a low-pressure area after floating ring leakage, area distribution of the high-pressure area and the low-pressure area is changed, and adjustment of axial force is achieved. According to the device, the axial force adjustment problem of the high-power two-stage pump with a complex leakage channel is solved, and the working reliability of an engine is improved.

Additively manufactured thrust chambers and thrust chambers with integral fluid manifolds, and hybrid additive manufacturing methods for their production, are provided. Hybrid additive manufacturing techniques may combine a variety of processes including, WAAM, PBF, cold spray and DED, for example, to produce objects with variant dimensional requirements, i.e., large overall size and small features. Hybrid additive manufacturing may be defined as provide various process layers within any manufactured object. These process layers in turn allow for the introduction of variable feature and size distribution throughout the manufactured object. Hybrid process layers according to aspects may also allow the use of a variety of materials or may use a single material across the various process layers.

A system and method are disclosed for improving fuel-air mixing in the combustion chamber of pulse combustors by deflecting the pressurized fuel injected from the one or more fuel injectors at a predetermined angle towards the centerline longitudinal axis of the combustion chamber.

The invention relates to a reusable propellant storage tank of a carrier rocket, and belongs to the technical field of liquid rocket propellant storage and management. The propellant storage box comprises a storage box body, and a lower-layer partition plate assembly, a middle-layer partition plate assembly and an upper-layer partition plate assembly which are arranged in a storage cavity of the storage box body, and the three layers of partition plate assemblies are of disc-shaped structures covering the cross section of the storage box. A first circulation area is arranged in the middle of the lower-layer partition plate assembly, a sealing plate and a second circulation area surrounding the sealing plate are arranged in the middle of the middle-layer partition plate assembly, a third circulation area is arranged in the middle of the upper-layer partition plate assembly, and the first circulation area, the second circulation area and the third circulation area are staggered in the radial direction, so that a nonlinear damping type liquid circulation path is formed in the storage tank. Therefore, violent shaking of the propellant in a weightlessness or variable acceleration environment is inhibited, and reliable restarting of the engine is guaranteed. According to the storage box, the structure is simple and reliable, the shaking restraining effect is remarkable, and the storage box is particularly suitable for propellant management of a reusable carrier rocket.

The invention provides a rocket engine short-time high-pressure high-speed jet flow simulation system and method.The simulation system comprises a high-pressure air source, a high-pressure storage tank, a Laval nozzle, a quick release mechanism, a test cabin and a measurement and control system, and the quick release mechanism is arranged in the throat of the Laval nozzle; the quick release mechanism is used for sealing and releasing the high-pressure gas stored in the high-pressure storage tank, the quick release mechanism is controlled by the measurement and control system to execute the release action, the high-pressure gas generates high-pressure high-speed jet flow through the Laval nozzle and acts on a to-be-tested piece in the test cabin, and high-pressure high-speed jet flow simulation of the rocket engine is achieved. Compared with a traditional device and method, cold airflow (air and / or nitrogen) is used for simulation, and the safety is remarkably improved; jet flows under different working conditions can be simulated only by adjusting parameters of the high-pressure storage tank, and reproducibility is good; core physical characteristics of high pressure, high speed, high noise and the like of real rocket jet flow are effectively reproduced, and key support is provided for structural bearing and acoustic vibration environment adaptability assessment.

Carbide-based fuel assembly includes outer structural member of ceramic matrix composite material, the interior surface of which is lined in higher temperature regions with an insulation layer of porous refractory ceramic material. A continuous insulation layer extends the length of the fuel assembly or separate insulation layer sections have a thickness increasing step-wise along the length of the fuel assembly from upper (inlet) section towards bottom (outlet) section. Fuel element positioned inward of the insulation layer and between support meshes has a fuel composition including HALEU and the form of a plurality of individual elongated fuel bodies or one or more fuel monolith bodies containing coolant flow channels. Fuel assemblies are distributively arranged in a moderator block, with upper end of the outer structural member attached to an inlet for propellant and lower end of the outer structural member operatively interfaced with a nozzle forming a nuclear thermal propulsion reactor.

The invention provides a hybrid power system for a dual-mode patrolling bomb and a use mode of the hybrid power system, and belongs to the technical field of aviation power. The problem that an existing patrolling bomb power system is difficult to meet the requirements for long-time cruising and instantaneous explosion acceleration at the same time is solved. The system comprises a hydrogen supply mechanism, a hydrogen peroxide supply mechanism, a fuel cell, a motor and a combustion chamber, the hydrogen supply mechanism is connected with an anode of the fuel cell, the fuel cell is connected with the motor, the motor is connected with a propeller, and the hydrogen peroxide supply mechanism is provided with two supply lines. The first supply line is connected with a cathode of the fuel cell through the hydrogen peroxide decomposition unit, the second supply line is connected with the combustion chamber, a first flow regulating valve is arranged on the first supply line, a second flow regulating valve is arranged on the second supply line, a grain is arranged in the combustion chamber, and the grain is spontaneous combustion fuel with a metal framework. The combustion chamber is connected with the exhaust nozzle. The hybrid power device is mainly used for providing hybrid power for the dual-mode patrol ammunition.

The utility model discloses a collection device for combustion products of an aero-engine fuel-rich gas generator, which adopts an asbestos fiber material, a zeolite molecular sieve, sandy soil and a filter screen for multiple filtration, and filters and collects solid-liquid components in outlet flue gas; and under the action of the water-cooling coil pipe, the residual outlet flue gas is cooled to a safe temperature, and the low-pressure gas tank is utilized to complete the collection of gas components. The collecting device disclosed by the utility model is stable in material performance, low in price, good in usability and service life, and capable of obtaining remarkable working and economic benefits.

The invention discloses a thrust rod elastic clamping structure for a rocket power device and a thrust rod. The thrust rod elastic clamping structure aims at solving the technical problems that according to traditional thread or flange connection, the machining precision requirement is high, installation is tedious, the negative mass is large, and storage deformation cannot be compensated. The structure comprises at least one axial clamping assembly and at least one circumferential clamping assembly which are arranged at the end of the thrust rod. The axial clamping assembly is matched with the first locked part through the first locking part and used for achieving axial connection of the structure and bearing main thrust. The circumferential clamping assembly is matched with the second locked part through the second locking part, so that circumferential locking of the structure is achieved, and torsion is resisted. The two assemblies are spatially arranged in parallel and functionally decoupled to jointly form a composite connection interface. By means of elastic clamping connection, rapid disassembly and assembly, automatic centering, vibration buffering and circumferential reliable locking between the thrust rod and the combustion chamber shell are achieved, and the position final assembly efficiency and the connection reliability are remarkably improved.

The invention relates to a zero-loss methane propellant supercooling filling system and method, and the system comprises a liquid nitrogen supply unit, a methane supply unit, a supercooling methane storage tank, a heat exchange pipeline unit and a temperature control unit; the heat exchange pipeline unit comprises a first heat exchanger, a second heat exchanger, a third heat exchanger and a separation and recovery module; the liquid nitrogen supply unit is connected with the cold end of the first heat exchanger. The cold ends of the first heat exchanger, the second heat exchanger and the third heat exchanger are sequentially connected in series. Methane output by the methane supply unit exchanges heat with the third heat exchanger and then is input into the supercooled methane storage tank; methane output by the exhaust end of the supercooled methane storage tank exchanges heat with the second heat exchanger and then is input into the separation and recovery module; and the separation and recovery module is connected with the supercooled methane storage tank and the methane supply unit and is used for converging the recovered methane into the supercooled methane storage tank and / or returning the recovered methane to the methane supply unit. According to the method, freezing in methane supercooling filling is effectively avoided, and lossless filling is achieved.

The present disclosure is directed to a supersonic oblique rotating detonation wave engine (SORDE) and systems and methods for generating a supersonic oblique rotating detonation wave. The SORDE is configured to produce and sustain a supersonic oblique rotating detonation wave through the injection of fuel at supersonic speeds into an inlet air flow between Mach 1 and Mach 7. The SORDE and method include injecting fuel into the inlet air in an amount to generate an equivalence ratio of 0.2 to 2.5. Some embodiments include a plurality of fuel injector ports each having a diameter of about 0.010 inches to about 0.040 inches; an annular wedge disposed in or upstream of the detonation chamber with an angle of about 5 degrees to about 40 degrees relative to a longitudinal axis of the engine; and / or a cylindrical center body disposed in or upstream of the detonation chamber.

A self-contained vertical takeoff and landing vehicle (VTOL) with aspect ratio of height to width no greater than 5.3:1 and with both rocket and airbreathing propulsion systems that can selectively operate one at a time or both at the same time and fuel tanks that have internal lattice structures formed of groups of interconnected multiplicities of fuel cells such as gyroids that line and are integrated with a vehicle skin to form a structural skeleton strengthened with said lattice structures that also serve to reduce fuel slosh compared with fuel tanks lacking such lattice structures and provide other benefits such as fuel temperature profile control.

The utility model discloses an ignition agent filling or discharging tool for an ignition agent storage box, and belongs to the technical field of ignition agent filling or discharging of ignition agent storage boxes. Comprising two fixing nuts, two threaded sleeves, two first check rings, two first sealing rings, two second check rings, two second sealing rings, two rotating shafts, two shells, two outer sleeve nuts and a storage box, the fixing nuts are in threaded connection with one ends of the rotating shafts, one ends of the fixing nuts are in lap joint with the threaded sleeves, the rotating shafts are sleeved with the threaded sleeves, and the storage box is connected with the outer sleeve nuts. The screw sleeve is in threaded connection outside the shell; according to the utility model, the ignition agent storage tank can be filled or discharged in an aerobic environment, the problem of high skill requirements on operators is solved, and meanwhile, compared with the traditional technology, the operation efficiency is greatly improved, the safety and reliability of the operation process are ensured, the operation is simple, and the practicability is high. And safety accidents caused by combustion of the ignition agent in contact with the air are avoided.

The high-pressure cantilever low-specific-speed centrifugal pump comprises a main shaft, the main shaft comprises a main shaft body, and the main shaft body is sequentially and coaxially sleeved with a floating ring, an auxiliary bearing and a centrifugal wheel from back to front in the axial direction. The floating ring is coaxially sleeved with an auxiliary bearing seat, and the axial front section of the auxiliary bearing seat is further coaxially sleeved with an auxiliary bearing. The outer side wall, close to the axial front end, of the main shaft body is further provided with a plurality of stop piece inserting grooves. The spindle body is coaxially sleeved with a nut locking gasket, and spindle external threads are arranged on the outer side wall of the axial front end of the spindle body. And the compression nut is arranged at the axial front end of the nut locking gasket and is connected with the main shaft body through the main shaft external thread. A plurality of locking piece installation grooves are further formed in the outer side wall of the compression nut, one end of the nut locking gasket is installed in the stop piece insertion groove, and the other end of the nut locking gasket is installed in the locking piece installation grooves. The two-stage booster pump solves the problems that the floating ring and the sealing convex shoulder are collided and abraded, and the strength of the compression nut is insufficient, and the reliability of the two-stage booster pump is improved.

The invention discloses an intermittent emission online control system of a low-temperature engine. The intermittent emission online control system comprises a control gas cylinder, an electric gas valve and a pre-cooling discharge valve which are sequentially connected through a pipeline, the precooling discharge valve is respectively communicated with the electric gas valve, the liquid oxygen main valve, the rocket liquid oxygen storage tank and a discharge pipeline through pipelines; the system further comprises a rocket-borne computer and a temperature sensor, and the temperature sensor is installed on a low-temperature pipeline connected with the engine. The temperature sensor is connected with the rocket-borne computer through a measuring cable; and the electric air valve is connected with the rocket-borne computer through a control cable. The invention further discloses a low-temperature engine intermittent emission online control method. According to the method, through real-time monitoring and dynamic optimization of the intermittent emission time sequence, the adaptability of an existing emission control method to different flight tasks is improved, the propellant emission consumption is reduced, and the success rate and economical efficiency of the rocket flight tasks are improved.

The invention relates to the technical field of rocket engines, in particular to a liquid-solid coupling type thermoelectric hybrid rocket engine which comprises a combustion starting structure and a butt joint tail spraying structure, the lower end of the combustion starting structure communicates with the butt joint tail spraying structure, and the combustion starting structure and the butt joint tail spraying structure are relatively fixed. A heat conduction pipe is arranged on a tail spraying pipe in the butt joint tail spraying structure in a sleeving mode, and the side end of the heat conduction pipe communicates with the butt joint sleeve. The combustion starting structure is used for fuel matching and combustion treatment, and the hydraulic control rod stretches out and draws back to control the lifting adjustment of the top tray and change the positions of the spring rod, the movable frame and the conical guide frame, so that the conical guide frame is adjusted on the exhaust nozzle, the gap position between the conical guide frame and the exhaust nozzle is changed, and the combustion wake flow is controlled; the combustion starting structure comprises a mixing treatment component, the lower end of the mixing treatment component is fixedly connected with a combustion chamber, and an igniter is installed on the combustion chamber. And through the arrangement of the structure, rocket launching treatment work can be conveniently carried out.

The application discloses a rocket conventional propellant filling training system and method, and relates to the related field of spaceflight equipment; the system comprises a simulation tank; liquid outlets of a plurality of storage tanks are respectively communicated with filling and discharging connectors on the simulation tank through filling pipelines; liquid inlets and liquid outlets of the plurality of storage tanks are respectively communicated with each other through reflux pipelines; a regulating and controlling device is arranged on the filling pipelines, can detect the temperature, pressure, flow, and flow rate in the filling pipelines, and regulate and control the flow and flow rate in the filling process; a gas distribution platform is connected with the simulation tank and the storage tanks, and can respectively introduce inert gas into the simulation tank and the storage tanks; and a filling control console is used for receiving detection data of the regulating and controlling device, and controlling the working processes of the regulating and controlling device and the gas distribution platform. The rocket conventional propellant filling training system and method can provide simulation filling training for post personnel and school students engaged in propellant filling in each launch site, fill the blank of teaching and training equipment, and improve the training effect and post ability.

The invention relates to an engine test bed device, in particular to a horizontal engine test bed fuel gas flow guiding device. In order to overcome the defect that in the prior art, an engine test bed fuel gas flow guide device cannot meet the requirements of ground test run and high modulus test at the same time, the horizontal engine test bed fuel gas flow guide device comprises a fuel gas flow guide cylinder, a cooling unit and an installation support arranged below the fuel gas flow guide cylinder. The fuel gas guide cylinder comprises a horizontal section and an inclined section which are communicated in sequence, and the horizontal section is detachably connected with the inclined section; the cooling unit conveys cooling water to a plurality of spraying pipes arranged on the inner side wall of the fuel gas guide cylinder, water is sprayed through the spraying pipes to form a protective water film, water is supplied to a cooling water tank at the bottom of the inclined section, and water is sprayed through the upper surface of the cooling water tank to cool fuel gas.

The application discloses a shaft-symmetrical rocket-based combined engine with a variable geometry throat, which comprises a combustion chamber and a square flow channel section in communication in an axial direction from front to back; the shell of the combustion chamber is surrounded by a bottom support bottom plate and a top plate, the bottom support bottom plate is horizontal, and the top plate is a semicircular arc shape which is arched upward; the square flow channel section is a cuboid cavity, and an upper variable structure geometry throat and a lower variable structure geometry throat which are consistent with the size of corresponding holes are embedded in the holes at the upper part and the lower part; the upper variable structure geometry throat and the lower variable structure geometry throat are both plate bodies in the shape of an inclined wedge, the tip of the inclined wedge faces the rear, and the inclined surface sides of the inclined wedge both face the cavity; the upper variable structure geometry throat and the lower variable structure geometry throat are used for changing the flow channel geometry throat area in the square flow channel section. The engine is matched with the flight state by reducing the throat area with the increase of the flight Mach number, the engine performance is improved, and the engine thrust vector adjustment can be realized.

The invention relates to the technical field of solid rocket engine combustion chamber composite material shells, and provides a large-aperture-ratio fiber winding composite material shell structure which comprises a front sealing head, a barrel section, a front skirt and a rear skirt assembly. The front sealing head is connected with one end of the cylinder section, the front skirt sleeve is arranged on the front sealing head and extends to the outside of one end of the cylinder section, and the other end of the cylinder section is connected with the rear skirt assembly. A traditional end socket is omitted, stable winding forming of the large-aperture-ratio composite material winding shell is achieved, and the requirement that the composite material shell bears the high internal pressure load effect is met. According to the structure, a traditional composite shell forming wire hanging tool and a combustion chamber shell skirt structure are combined, the structure and the function are integrated, the problem of fiber wire sliding during winding of a large-aperture-ratio structure is solved, the manufacturability of winding forming of the large-aperture-ratio composite shell is greatly improved, and the internal pressure bearing capacity of the large-opening composite shell is improved.

This invention relates to an on-orbit high-density storage and supply device and method for electric propulsion working propellant, belonging to the field of spacecraft deep space exploration and on-orbit servicing technology, which can effectively save space and mass resources of spacecraft. The cooling circuit consists of a storage tank, a refrigerator, a pump, and an injection pipe connected sequentially; the pressurization circuit consists of a storage tank, a heating evaporator, and a third solenoid valve connected in a loop; and the gaseous working propellant circuit consists of a storage tank, a vaporizer, and a buffer chamber connected sequentially. This invention simplifies the complexity of the propulsion system and reduces potential failure risks, effectively supporting the application of high-power, high-specific-impulse, and long-life new electric propulsion technologies, providing important support for deep space exploration and on-orbit servicing missions.

The invention relates to a hybrid engine (10) for projectiles in which a rocket engine with an annular nozzle (7) is arranged along with a base bleed unit with a circular nozzle (8) where the circular nozzle (8) of the base bleed unit is arranged centered in the annular nozzle (7) on the rocket engine and where the hybrid engine (10) is arranged with a propellant charge for the rocket engine (2) and a propellant charge for the base bleed unit (5). The invention further relates to a projectile (100) outfitted with a hybrid engine.

The invention discloses a solid rocket engine adiabatic pasting device, and relates to the technical field of rocket engine adiabatic pasting devices.The solid rocket engine adiabatic pasting device comprises a center shaft, a rotating frame is installed at one end of the center shaft, and a manipulator is installed at the other end of the center shaft; a gluing assembly, a guide wheel, a first rolling wheel, a cut-off assembly and a second rolling wheel are sequentially arranged on the rotating frame in a circumferential shape, a driving assembly for driving the rotating frame to rotate and the gluing assembly to conduct gluing is installed on the rotating frame, one end of the rotating frame extends into the engine through the center shaft, the driving assembly is started through the controller, and the cutting-off assembly conducts cutting-off on the rotating frame. The driving assembly can drive the gluing assembly to output a heat insulation sheet with glue, the heat insulation sheet is pasted on the inner wall of the engine after being guided by the heat insulation sheet guide wheel and rolled by the first rolling wheel, the heat insulation paste on the inner wall of the engine is pasted through rotation of the rotating frame, pasting is more uniform and smooth, operation is convenient, and the pasting efficiency and the pasting quality are improved.

This invention discloses a four-modal, wide-speed-range aerospace propulsion system, belonging to the field of aerospace propulsion technology. The system integrates pure rocket takeoff, air-assisted combustion acceleration, air-assisted combustion detonation cruise, and pure oxygen detonation orbit insertion. It draws upon the operating logic of multi-cylinder annular array aero-engines to adapt to aerospace rocket propulsion scenarios. Multiple Laval detonation tubes are arranged in an annular array. A top inlet dual-rotary gas distribution valve and a bottom outlet dual-rotary gas distribution valve enable independent hydrogen supply and exhaust gas discharge. Side-wall jet gas distribution valves provide time-separated hydrogen and oxygen supply, with detonation occurring only within the detonation tubes. The valve assembly employs a fixed-phase, one-time lock-in design, combined with hydrogen-oxygen regeneration cooling and an exhaust pressure stabilization chamber to achieve stable operation. It features a one-button emergency switchback to pure rocket operation under all conditions. The modular structure adapts to vacuum-atmosphere cross-domain operating conditions, solving the pain points of traditional aerospace propulsion such as backfire, premixed combustion, timing misalignment, and low arrangement efficiency. It possesses extremely high engineering feasibility and industrialization potential in the aerospace propulsion field.