3192results about "Rocket engine plants" patented technology

The disclosed methods, systems, and kits provide the ability to deliver entire clean sheet designs from concept to first hot fire in under six weeks with instant specific impulses above 330 seconds in some of our engines. In examples, thrusters can be delivered that are at less than half of the mass budget allowable for them and they can be delivered in weeks.

The embodiment of the invention provides a liquid nitrogen heating, vaporizing and pressurizing system and a liquid rocket. The liquid nitrogen heating, vaporizing and pressurizing system comprises an oxygen tank, a fuel tank, a liquid nitrogen tank, a liquid nitrogen pump, a cavitation pipe, a liquid nitrogen heater and a liquid nitrogen tank pressurizing pipe. An oxygen tank pressurizing pipe is arranged at the top of the oxygen tank; a fuel tank pressurizing pipe is arranged at the top of the fuel tank; the liquid nitrogen box is arranged in the oxygen box, and the bottom of the liquid nitrogen box is connected with a liquid nitrogen conveying pipe; the liquid nitrogen pump is arranged on the liquid nitrogen conveying pipe; the cavitation pipe is used for stabilizing the liquid nitrogen conveying flow; the liquid nitrogen heater is used for heating liquid nitrogen conveyed by the liquid nitrogen pump into high-temperature nitrogen, the heated nitrogen is used for pressurizing the oxygen tank, the fuel tank and the liquid nitrogen tank at the same time, and the liquid nitrogen heater is connected with the oxygen tank pressurizing pipe and the fuel tank pressurizing pipe. According to the liquid nitrogen heating, vaporizing and pressurizing system, the integrated pressurizing technology is adopted, the pressurizing system is simplified through the technology, meanwhile, the cost and weight of the pressurizing system are reduced through liquid nitrogen heating, vaporizing and pressurizing, and the safety and stability of the pressurizing system are improved.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The invention relates to the technical field of space propellant fluid, in particular to a satellite low-temperature propellant liquid injection and storage combined device which comprises a radial injector, parallel channels, a liquid collector and a channel distributor, the radial injector is of a thin-wall cylinder structure, and the middle line of the radial injector coincides with the inner axial direction of a low-temperature storage box; the liquid collector is arranged at the lower end of the radial ejector, is welded and fixed with the bottom end of the radial ejector, and is communicated with a liquid opening of the low-temperature storage tank; the channel distributor is arranged at the upper end of the radial ejector, is welded and fixed with the top end of the radial ejector, and is communicated with an air port of the low-temperature storage tank; the top end of each parallel channel is fixedly welded to the channel distributor, and the bottom end of each parallel channel is fixedly welded to the liquid collector. The fluid storage and mixing efficiency under the space microgravity condition is achieved to the maximum extent, thermal stratification in the low-temperature storage tank is effectively eliminated through sufficient convection mixing, and the temperature uniformity in the storage tank is improved.

The invention relates to an attitude and orbit control engine high-altitude environment simulation ignition test system and a verification method, in particular to a high-altitude simulation test bed active injection verification system and a verification method, and solves the technical problems of high cost, high risk and low verification data validity during verification of a high-altitude simulation test bed in the prior art. The active injection verification system for the high-altitude simulation test bench comprises a vacuum chamber, a diffuser, a water spray cooler, a first corrugated pipe, a first gas collection chamber, a second corrugated pipe, n groups of coolers, a third corrugated pipe, a second gas collection chamber, a steam-jet pump, a pipeline and a data acquisition module, according to the verification method of the active ejection verification system of the high-altitude simulation test stand, ejection capacity verification is carried out in two stages, and the verification effectiveness is greatly improved on the premise that the verification cost is greatly reduced.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The invention discloses a diversion cooling device and a rocket engine ignition test system, and belongs to the technical field of rocket launching tests. The flow guide cooling device comprises a spraying piece, a flow guide mechanism and a water supply mechanism. The spraying piece is arranged at the initial section of the tail flame of the rocket engine, a water curtain barrier can be arranged through spraying from top to bottom, cooling and noise reduction are achieved, fuel gas of the tail flame is prevented from burning other devices, and the temperature of the cooled tail flame cannot be too high when the cooled tail flame makes contact with the flow guide mechanism, and the flow guide mechanism cannot be damaged; in the flow guide mechanism, a flow guide groove water spraying panel which is in an arc shape and protrudes downwards in an inclined mode and flow guide groove bodies located on the two sides of the flow guide groove water spraying panel are used for conducting comprehensive spraying and flow guide on the high-impact-force and high-temperature tail flame, cooling and noise reduction are achieved, and water spraying areas with different spraying densities can conduct targeted spraying according to the temperature of the inner flame and the outer flame of the tail flame. And the arc plate and the second nozzle are detachable, so that the arc plate and the second nozzle are convenient to replace, assemble and disassemble after a certain part is damaged by high temperature, the cost is reduced, and the use economy is improved.

Transpiration-cooled systems having permeable and non-permeable portions are provided. In one aspect, a nozzle for a rocket engine includes a non-permeable outer layer, a plurality of walls, a permeable layer, and a plurality of cooling channels. The non-permeable outer layer forms a predetermined shape of the nozzle and defines an interior space. The plurality of walls extend generally orthogonal to the non-permeable outer layer and into the interior space. The permeable layer is disposed between each set of adjacent walls of the plurality of walls. The permeable layer is spaced a distance from the non-permeable outer layer. The plurality of cooling channels are formed between the permeable layer, the non-permeable outer layer, and a set of adjacent walls of the plurality of walls. The plurality of cooling channels are configured to transport a fluid.

The utility model provides a storage box bottom structure with a rack, which comprises a box bottom formed by welding a circular top cover and a plurality of scalloped segments; the rack small end frame is arranged on the outer side of the center of the top cover and used for being connected with an engine; and the plurality of rack force transmission plates are uniformly connected to the circumferential position of the rack small end frame and respectively extend to the scalloped segments from the top cover. The box bottom structure has the advantages of being simple in structure and capable of improving carrying capacity.

The invention relates to an engine tail flame anomaly detection method and system based on spectrum contrast, and the method comprises the steps: simulating an ideal combustion process of an engine based on a radiation model, generating an ideal tail flame spectrum for reference, and extracting the reference spectrum intensity of a target substance; the actual tail flame spectrum of the tail flame of the engine to be detected is collected, and the actual spectrum intensity of the target substance is obtained; based on the reference spectral intensity and the actual spectral intensity, obtaining dynamic characteristics used for judging whether the to-be-detected engine fails or not; constructing a dynamic threshold range for judging whether the dynamic characteristics are abnormal or not; whether the dynamic features are abnormal or not is judged based on the dynamic threshold value range, if the dynamic features exceed the dynamic threshold value range, the dynamic features are abnormal, and if the dynamic features are within the dynamic threshold value range, the dynamic features are not abnormal. According to the invention, organic combination between simulation and actual monitoring is realized, and high-precision and real-time fault diagnosis can be flexibly realized.

The invention includes engines adapted for using apparatuses and methods for instantiating chemical reactants in a nanoporous carbon powder, and further includes uses for such engines and vehicles incorporating such engines.

The invention discloses a pilot-operated type self-driving fluid release valve, and belongs to the field of control valves. Comprising a main valve assembly, a pilot valve assembly and a one-way valve assembly. Valve bodies of the pilot valve assembly and the one-way valve assembly are integrated on a main valve body of the main valve assembly; the main valve assembly comprises a main valve body, a main valve piston, a main valve element and a main valve spring. The pilot valve assembly comprises a pilot valve air control opening, a pilot valve inlet, a pilot valve outlet channel and a pilot valve pressure relief opening. The one-way valve assembly is arranged at an outlet of the main valve and used for blocking medium backflow. Through mechanical configuration reconstruction and system integration innovation, the technical contradiction of the liquid rocket engine high-pressure gas control valve in weight, response speed and reliability is thoroughly solved.

The invention relates to an engine fault detection method based on multiple sensors and spectral data fusion, and the method comprises the steps: collecting the first spectral data, reference temperature and reference pressure of the tail flame of a normal engine, and extracting the ideal spectral characteristics of a target substance based on the first spectral data; acquiring second spectral data of engine tail flame of the to-be-detected engine, and acquiring actual spectral characteristics of the target substance based on the second spectral data; the actual tail flame temperature and the actual tail flame pressure of the engine tail flame of the engine to be detected are collected, and the actual spectrum characteristics are corrected to obtain actual spectrum correction characteristics; based on the ideal spectral feature and the actual spectral correction feature, obtaining a dynamic feature for judging whether the to-be-detected engine has a fault; constructing a dynamic threshold range; and judging whether the dynamic characteristics are abnormal based on the dynamic threshold range. The method can be flexibly and accurately used for engine fault prediction and timely alarm, and accidents can be fully reduced.

A fuel grain assembly for a hybrid rocket engine includes a hybrid fuel grain including fuel grain material, wherein an outermost portion of the fuel grain material defines an outer surface of the fuel grain, and wherein the fuel grain material includes a polymer based rocket fuel material; and a solid propellant material disposed in contact with the fuel grain material of the hybrid fuel grain, wherein the solid propellant material includes an oxidizer and a binder material, wherein a hollow combustion port extends through the fuel grain assembly, wherein at least a portion of a wall of the hollow combustion port is defined by the solid propellant material.

The invention describes an air plasma propulsion system and method comprising: an air turbine plant adapted to receive supply air from an external environment, and a plasma propulsion unit adapted to increase a speed of said air after a corresponding ionization and passage into air plasma phase. The plant comprises a turbine compressing air which transits in the plasma propulsion unit which ionizes it strongly and accelerates it through a frequency or microwaves in magnetic field produced by a laser beam.

The invention discloses a water-air bimodal self-adaptive rocket engine system and a control method, and belongs to the technical field of rocket control. The system comprises an environment sensing recognition module, a thrust output regulation and control module, a fuel combination switching module, a heat exchange regulation module, an attitude propulsion control module and a modal coordination control module, and precise recognition and rapid classification of a rocket operation environment are realized through combined judgment of a multi-point distributed sensing unit, density and flow characteristics; by constructing a multi-level thrust adjusting system, the dynamic change of the geometrical shape of the spray pipe is achieved, combustion supply ratio adjustment and multi-parameter linkage control are fused, the self-adaptability and response sensitivity of thrust output are remarkably improved, rapid conversion between high-pressure short-range propulsion and efficient long-range flight is achieved, and the high-pressure short-range propulsion and high-efficiency long-range flight are achieved. Regulation and control commands are preloaded and distributed in parallel according to the current environment, so that all functional modules can respond cooperatively, the operation mode in the new environment can be quickly adapted, and the system can capture the execution state of the regulation and control commands in real time.

A vacuum insulated blast tube includes an insulating layer of a burn resistant material such as phenolic resin formed on an interior surface of the blast tube to provide the necessary erosion and thermal insulation properties to protect the blast tube and a vacuum insulated layer formed in the walls of the blast tube itself to provide the additional thermal insulation properties to protect any non-propulsive sub-systems positioned in the void space around the blast tube. A void space in the walls of the blast tube is held under vacuum with a pressure of less than 25 Torr and a thermal conductivity Tcond_vac of less than one-third of the thermal conductivity of air Tcond_air to form the vacuum insulation layer.

The invention relates to a multifunctional pump rear swing liquid rocket engine single-machine rack which is characterized in that two main force bearing beams and two auxiliary force bearing beams are welded to form a lower force transmission frame; the four upper connectors and the four upper connecting beams are welded to form an upper force transmission frame. The four main force bearing rods are welded to the upper force transmission frame and the lower force transmission frame respectively. The two servo supports are arranged to serve as upper fulcrums of the servo mechanism, and the bidirectional swing requirement of the liquid rocket engine can be met. And a pair of turbine pump supporting seats are symmetrically arranged with one of the servo supports. And a pair of turbine pump auxiliary support lugs is arranged on the same side of the turbine pump support. The turbine pump support and the turbine pump auxiliary supporting lug are jointly responsible for being connected with the turbine pump, and stable fixed supporting of the turbine pump is achieved. The main force bearing beam is connected with the lower end force transmission assembly, the upper connector is connected with a rocket body or a multi-machine rack, and the thrust transmission function of the engine is achieved.

The invention discloses a solid rocket engine composite material skirt connecting structure and a forming method thereof.The connecting structure comprises a composite material skirt, a skirt hole reinforcing ring, a connected cabin section and a connecting piece, the composite material skirt and the skirt hole reinforcing ring are integrally formed, and the rear end of the composite material skirt and the rear end of the skirt hole reinforcing ring are connected with the connected cabin section through the connecting piece; the connected cabin section is of an L-shaped nested inner lap joint structure and has the butt joint limiting and bearing functions, and the thickness transition section at the front end of the composite skirt is connected with a fiber winding structure layer of a combustion chamber shell. Compared with a traditional metal skirt connecting structure, the weight reduction effect is obvious, compared with a traditional composite skirt hinge block connecting structure, the connecting mode is simpler, and the fiber continuity of a composite skirt connecting area is guaranteed; the problems that in-plane shear rigidity and strength of the composite material are insufficient, and orifice positions are prone to damage are solved. The structure and the forming process are simple, the maturity is high, and the cost is low.

An improved liquid-propellant engine utilizing a combination of a gas turbine, a generator, an electric pump fed engine, and a power management system. When in use, the gas turbine and generator are configured for providing power to an electric motor for operating a fuel pump and an oxidizer pump to pump fuel and oxidizer from individual tanks into a combustion chamber for mixture and combustion. The combustion of the fuel and oxidizer is configured for providing thrust from the base of a rocket to propel it. The generator of the present invention provides continuous electrical power to the electric pump(s) to continue the supply of fuel and oxidizer to the combustion chamber. The power management system of the present invention is electrically connected to the generator, the electric motor(s) configured for operating the fuel and oxidizer pumps, a flight computer, and fuel and oxidizer valves.

The invention provides an online fault detection method for a reusable liquid rocket engine, and aims to realize real-time online adjustment of a fault threshold according to the change of the working condition of the engine under the condition of large-range variable thrust of the reusable engine so as to form a red line shutdown strategy based on an online dynamic threshold. Faults occurring in the working process of the engine are pre-warned and judged, so that effective measures are taken in time, and the safety of the carrier rocket and the load is protected.

The invention relates to a working platform in the bottom of a rocket storage tank and a mounting method of the working platform. The bottom of the rocket storage tank at least comprises a scalloped segment and a manhole formed in the bottom of the rocket storage tank. The working platform comprises a supporting mechanism, a mounting mechanism and a plurality of standing mechanisms, one end of the supporting mechanism is used for extending into the bottom of the rocket storage tank through a manhole, and when one end of the supporting mechanism is located in the bottom of the rocket storage tank, the mounting mechanism is detachably mounted at the end, located in the bottom of the rocket storage tank, of the supporting mechanism; the multiple standing mechanisms are detachably mounted on the mounting mechanism in the circumferential direction of the mounting mechanism in sequence, buffer layers are mounted at the ends, away from the mounting mechanism, of the standing mechanisms, and the tail ends of the buffer layers are used for abutting against the side walls of the ends, located in the bottom of the rocket storage tank, of the scalloped segments. The occupied area of the rocket storage tank can be reduced, the product quality of the rocket storage tank is guaranteed, the working efficiency is improved, and danger is reduced.

A tailor-welding, spinning and variable-wall-thickness forming method for a large-sized extension section of a liquid propellant rocket engine, which belongs to the technical field of metal forming. The method comprises: performing laser numerical control machining on a GH3044 superalloy to obtain two chord-tangent circles; performing electron beam welding to obtain a complete circle; performing a vacuum solid solution treatment; performing pre-forming to obtain a shallow dish shape; performing a vacuum solid solution treatment; assembling a variable-wall-thickness flat blank and a spinning mandrel of a first spinning die, and then performing spinning on a spinning machine to form a curved generatrix variable-wall-thickness blank part; performing a vacuum solid solution treatment; and assembling the curved generatrix variable-wall-thickness blank part and a spinning mandrel of a second spinning die, and then performing spinning on the spinning machine to form a curved generatrix variable-wall-thickness part. The present invention is mainly used for the manufacturing of an ultra-large-sized GH3044 extension section of the liquid propellant rocket engine, which has a diameter of 1200 mm and a length of 1500 mm, and the product precision reaches a wall thickness tolerance of + / -0.05 mm for each point and a surface profile of ≤3 mm.

The invention provides a rocket power anchor. The rocket power anchor comprises a bullet of a spiral structure. The anchoring structure is detachably connected with the warhead, and the anchoring structure can stretch; the second-stage engine is detachably connected with the anchoring structure, a second-stage delay igniter capable of driving the anchoring structure to stretch is arranged at the joint of the second-stage engine and the anchoring structure, and the second-stage engine is filled with a second-stage propelling grain; the first-stage engine is detachably connected with the second-stage engine, a middle partition bin is arranged at the joint of the second-stage engine and the first-stage engine, a first-stage delay igniter is arranged on the middle partition bin, and the first-stage engine is filled with a first-stage propelling grain; the exhaust nozzle can be obliquely arranged at a certain angle, so that the rocket power anchor can rotate during movement; the lifting ring is rotationally connected to the first-stage engine; the total mass of the first-stage propelling grain is smaller than that of the second-stage propelling grain; and a pressure-bearing diaphragm is arranged on the middle partition bin.

The invention relates to a fluid supply system and method, in particular to a cryogenic propellant automatic supply system and method for a liquid rocket engine. The invention aims to solve the problems of tedious process, low efficiency and low safety of a method for pre-cooling and filling a liquid rocket engine by manual operation. Comprising a main container used for storing a low-temperature propellant, a supply pipeline connected to the output end of the main container and the input end of a low-temperature propellant cavity of the liquid rocket engine and used for supplying the low-temperature propellant, an exhaust pipeline connected with the supply pipeline, and a valve and a monitoring assembly arranged on the supply pipeline. The valve and monitoring assembly comprises a discharge valve, a first exhaust valve, a second exhaust valve, a third exhaust valve, a fourth exhaust valve, a first supply valve, a first bypass valve, a second supply valve, a second bypass valve, a blowdown valve, a flow meter assembly, a filter, a plurality of temperature monitoring sensors and a plurality of pressure monitoring sensors.

The invention discloses a free filling type metal wire embedded solid propellant forming device and process, and belongs to the field of solid propellants, the free filling type metal wire embedded solid propellant forming device comprises a bottom plate, a base, a mold cylinder and a top cover are sequentially arranged on the bottom plate, and a plurality of pressing structures are arranged on the upper portion of the top cover; a solid propellant coating barrel is arranged on the inner wall of the mold barrel; the lower end of the metal wire is fixed on the bottom plate, the upper end of the metal wire is fixed on the first fastener of the fastener mounting block on the top cover, and the first fastener is screwed into the threaded hole, so that the metal wire can be tensioned, and the internal stress of the metal wire can be adjusted by adjusting the expansion amount of the metal wire; according to the invention, the collimation of the metal wire in the propellant is ensured, the proper stress of the metal wire in the slurry casting process is also ensured, the problem that the metal wire is broken due to overlarge stress is avoided, and the technical problem that the yield of the solid propellant is low due to the fact that the metal wire is easy to break in the propellant in the prior art is solved.

The invention discloses a liquid rocket engine fault simulation system, method, medium and equipment, the system comprises a fault simulation module, an environment interference simulation module and an image acquisition and fault identification module, the fault simulation module is used for simulating various fault states of a liquid rocket engine in a test run process; the environment infection simulation module is used for simulating various environment interferences occurring in the test run process of the liquid rocket engine; and the image acquisition and fault identification module is used for acquiring video image data of the liquid rocket engine in the fault simulation and environment interference simulation process so as to carry out real-time identification and diagnosis on the fault of the liquid rocket engine. According to the invention, the fault of the liquid rocket engine in the test run process can be accurately identified.

The invention provides a liquid oxygen methane engine test bed composed of a multifunctional test system. The test bed comprises a gas path system, a liquid path system and a test station. Wherein the gas path system and the liquid path system are connected through a pipeline for gas supply, and the liquid path system is connected with the test run station and used for providing liquid fuel needed by testing for the test run station. The liquid path system comprises a low-pressure liquid path module and a high-pressure liquid path module. And the low-pressure liquid path module and the high-pressure liquid path module are respectively connected with a test run station. The test bed is composed of a set of gas path system, two sets of liquid path systems and three test stations. The two sets of liquid path systems jointly adopt one set of gas path system for gas supply, so that medium conveying can be carried out on the three test run stations according to requirements, the test run platform can meet the gas distribution requirements of engines of different models, meanwhile, the resource sharing of test equipment is improved, the total cost of equipment construction and maintenance is greatly reduced, and the working efficiency is improved. And the engine testing efficiency is improved.

The invention provides a thrust chamber, a power system and an aircraft. The thrust chamber comprises a combustion body which is cylindrical and is provided with a throat part, and a combustion chamber is defined by a first part, located on the upstream of the throat part, of the combustion body; the first annular side wall is connected with the combustion body and located at the end, away from the throat part, of the combustion chamber. Wherein the combustion body is provided with a first cooling channel, and the first cooling channel extends from the lower end of the combustion body to the upper end of the combustion body; the first annular side wall is provided with a plurality of first cooling holes communicating with the first cooling channel, the multiple first cooling holes are distributed in the circumferential direction of the combustion body and extend in the axial direction of the combustion body, and the first cooling holes communicate with the combustion chamber and are close to the inner wall of the combustion chamber. Coolant enters the combustion chamber from the upstream through the first cooling channel and the first cooling hole in sequence, the coolant is dispersed along the inner wall of the combustion chamber, the temperature near the inner wall of the combustion chamber can be reduced, and the wall face of the combustion chamber is protected.

The invention discloses a liquid engine ignition device. The liquid engine ignition device comprises a storage tank used for storing an ignition agent required by preset ignition times and an ignition tube used for storing the ignition agent required by single ignition; the ignition tube is communicated with the storage tank through a communicating tube; a one-way valve for preventing an ignition agent in the ignition pipe from flowing back into the storage tank is arranged on the communicating pipe; the ignition tube is configured to be capable of extruding an ignition agent which is stored in the ignition tube and is used for single ignition into an engine combustion chamber under the extrusion effect; the storage tank is configured to be capable of automatically filling the ignition tube with the discharged ignition agent with the ignition agent stored in the storage tank under the extrusion effect. According to the method for sharing the storage tank and the ignition tube, the reliability of the ignition tube and the multi-time ignition characteristic of the storage tank are combined, automatic quantitative ignition agent filling is achieved, and an efficient ignition scheme is provided for multi-time ignition of a rocket engine.