1515results about "Rocket engine plants" patented technology

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.

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 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.

The embodiment of the utility model provides a storage tank bottom structure, a storage tank and a carrier rocket, the storage tank bottom structure comprises a shell and a top cover, the shell is of a hollow arc-shaped revolving body structure which is through in the front-back direction, the two ends of the shell are the small-diameter end and the large-diameter end respectively in the axis direction of the shell, and the small-diameter end of the shell is fixedly connected with the top cover; a propellant is arranged on the top cover; a plurality of protruding structures are evenly distributed on the outer side face of the shell in the circumferential direction in the axis direction of the shell. According to the technical scheme, the bottom structure of the thrust bearing type storage tank is simple in process, easy to manufacture in batches and beneficial to achieving light weight of a carrier rocket.

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 relates to the technical field of plasma ignition, in particular to a split type plasma ignition device and ignition system for passivating gunpowder, which comprises an ignition tube assembly and an oscillation control assembly, the ignition tube assembly comprises an ignition tube, gunpowder and a hollow resonance coil; the oscillation control assembly comprises a coil support, an excitation coil and an oscillator circuit module. The excitation coil on the coil support is used for being matched with an insulation tube in an ignition tube in an inserted mode. An ignition pulse excites an oscillator circuit module to generate a high-frequency excitation signal, the high-frequency excitation signal is released by an excitation coil and coupled to a hollow resonance coil in an electromagnetic induction mode to form LC resonance, the LC resonance enables a discharge part to generate a high-intensity electric field, the high-intensity electric field excites plasmas on the surface of gunpowder to ignite the gunpowder, and ignition is completed. The device breaks through multiple limitations of a traditional ignition mode, and intrinsic safety and high reliability of passivation agent ignition are achieved.

In a rocket motor in which burn rate enhancement (BRE) wires are used to accelerate the burn rate of the solid propellent, embedded charge assemblies (ECAs) are configured as support structures for the BRE wires. Each ECA includes an energetic material that is configured to burn along with the solid propellent to produce thrust and, upon detonation, to break up the solid propellent to terminate thrust. The detonation may also be initiated as a part of process to prevent a higher-order reaction, such as in reaction to heating from a fire or other cause. By being located inside the casing, the energetic material and ECAs do not adversely affect aerodynamics of the flight vehicle of which the rocket motor is a part, such as a missile.

The invention discloses a solid rocket engine in-situ heat preservation ignition test run frame which comprises a fixing base, a movable fixing plate, a supporting assembly, a force bearing pier, a first force transmission piece, a second force transmission piece, a force measuring piece, a heat preservation bottom plate and a heat preservation cover. A solid rocket engine to be tested is movably arranged on the movable fixing plate through the two supporting assemblies, and the movable fixing plate is connected to the fixing base through the multiple plate springs. The heat preservation bottom plate is arranged on the lower side of the movable fixing plate, the heat preservation cover is buckled on the heat preservation bottom plate, cold air is conveyed into the heat preservation cover through the cold air connector, the solid rocket engine to be tested can be cooled, and when the temperature of the solid rocket engine to be tested reaches the temperature required by testing conditions, the movable door is opened, and the solid rocket engine to be tested is ignited. And the force measuring piece can obtain thrust data of the solid rocket engine. The solid rocket engine in-situ heat preservation ignition test run rack can accurately simulate the thrust test of the solid rocket engine in an ultralow-temperature working environment.

The invention discloses a liquid-propellant rocket engine turbine pump inducer design method and an inducer, and the method comprises the steps: determining the working conditions of the inducer according to the overall design requirements of a liquid-propellant rocket engine; according to design requirements and an empirical formula, basic parameters of the inducer are preliminarily determined; different core parameters are combined according to the core parameter range of the inducer, and a plurality of initial three-dimensional models of the inducer are established; respectively carrying out analogue simulation on the plurality of established initial three-dimensional models of the inducer to obtain the cavitation performance of the inducer; according to a flow analysis result, obtaining an inducer of an optimal blade profile in all parameter combinations; and parameter optimization and adjustment are conducted on the inducer with the optimal blade profile till the inducer with the optimal cavitation performance for the turbine pump of the liquid rocket engine is obtained. By means of the method, the liquid rocket engine turbine pump inducer with the good cavitation performance can be designed.

The invention provides a high-pressure embedded blanking cap assembly for a solid rocket engine jet pipe, which is characterized in that a throat part of the solid rocket engine jet pipe and the front area of an expansion section at the downstream of the throat part are thinned from the inner wall to the outer side to form a mounting concave part of the blanking cap assembly; the blanking cap assembly comprises a blanking cap, a pressing ring and a throat liner which are sequentially installed in the gas ignition direction, the throat liner is installed on the throat portion of the spray pipe, the blanking cap is a thin-wall metal piece and adheres to the throat liner into a whole, the blanking cap is used by nesting of the metal pressing ring, and the blanking cap and the metal pressing ring are jointly adheres to the front area of the expansion section on the downstream portion of the throat portion to plug an outlet of the throat portion. The inner hole of the compression ring serves as a composition section of the inner molded surface of the spray pipe and is consistent with the inner molded surface of the front area of the expansion section on the downstream of the throat. The blanking cap structure has the advantages of being good in opening pressure intensity stability, high in reliability and high in consistency under the short-time ignition high-pressure opening requirement.

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.

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 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.

An air breathing engine system has an engine inlet configured to receive a flow of air. A combustor is in fluid communication with the engine inlet. A rocket ejector is positioned between the engine inlet and combustor. The rocket ejector is connected to the combustor with a mixing tube. The rocket ejector adjustably outputs high-temperature gas into air in the mixing tube. The high-temperature gas changes a pressure and thermal energy of the air within the mixing tube, which can be used to optimize combustor operation, such as based on a trajectory, flight condition, or speed of an aircraft propelled by the air breathing engine system.

The invention relates to a pilot-operated bacteria valve. The pilot-operated type fungus valve comprises a main valve mechanism, the main valve mechanism comprises a main valve shell and a main valve element, the main valve shell is provided with a main valve cavity, a feeding channel, a discharging channel and a pilot channel, and the main valve element is configured to cut off the feeding channel and the discharging channel and conduct the feeding channel and the pilot channel in a closed state and conduct the feeding channel and the discharging channel in an open state; the pilot valve mechanism is connected with the main valve shell, the pilot valve mechanism is provided with a pilot valve cavity, and the pilot valve cavity communicates with the pilot channel; the flow guide mechanism comprises a flow guide pipeline, and the flow guide pipeline communicates with the guide valve cavity and the discharging channel; the pilot valve mechanism is configured to be used for separating the pilot channel and the flow guide pipeline or communicating the pilot channel and the flow guide pipeline. The pilot-operated mushroom valve controls the opening and closing of the main valve core through the pilot valve mechanism, so that the driving force required by the action of the main valve core can be reduced, the structural size of the pilot-operated mushroom valve is reduced, and the weight is reduced.

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.

The invention provides a conventional and detonation dual-mode spontaneous combustion propellant space engine, and belongs to the technical field of spacecraft propulsion systems. The conventional and detonation dual-mode spontaneous combustion propellant space engine comprises a propellant control valve, an engine head, an engine body and an external excitation device. The propellant control valve, the engine head part and the engine body part are connected in sequence, and the external excitation device is embedded into the outer wall of the engine body part; according to the invention, by adopting an external excitation mode, the space engine in a detonation mode is realized; by means of a selective excitation mode, the space engine capable of selecting a conventional mode and a detonation mode is achieved. The detonation combustion is applied to the spontaneous combustion propellant space engine, the upper limit of the combustion efficiency of the space engine of a constant-pressure thermodynamic cycle is broken through, and the combustion efficiency is further improved on the original basis.

The invention discloses a multi-source information sensing system based on radio frequency sensing. The multi-source information sensing system comprises an RFID sensing system, reliability optimization deployment in a strong random environment and vibration signal recovery and reconstruction. The RFID sensing system can be used for collecting original signal data of an engine under the non-contact condition. Secondly, reliability optimization deployment is carried out in a strong random environment, and deployment of system tags, antennas and sensing parameters is optimized under the condition that influences of various interference factors in the working environment of the engine are comprehensively considered, so that the performance of the sensing system is improved. The method can be used for realizing accurate recovery of vibration characteristic signals, ensuring quick and accurate sensing of key parameters of an engine, and realizing quick detection of typical fault modes. The technology has wide application potential and can play an important role in the fields of engine vibration monitoring, engine fault analysis, engine health management and the like.

The invention provides a diagnosis method and device for unstable combustion oscillation characteristics of a rocket engine, and relates to the field of rocket engine testing, and the method comprises the steps: obtaining part parameters of all parts of a thrust chamber of a to-be-analyzed rocket engine; based on the nozzle parameters and the channel parameters, nozzle outlet sound admittance is determined, and mass flow pulsation of each nozzle channel outlet is determined; engine combustion chamber design parameters and configured flame parameters are processed, and heat release pulsation of a single flame is determined; and based on the configured modal cutoff value, mass flow pulsation and heat release pulsation, solving the configured acoustic characteristic network to obtain a plurality of actual main frequencies and modal numbers corresponding to the actual main frequencies, and determining modal categories of the actual main frequencies. According to the dynamic response characteristics of the injection pulsation of a single nozzle or multiple nozzles and the heat release pulsation of the combustion flame, the actual main frequency of the low-temperature engine during oscillation combustion under different working conditions can be predicted, and relevant parameter analysis is carried out.

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 rocket motor comprising a non-permeable outer shell, an intra-wall channel between the non-permeable outer shell and an annular section, and a center chamber formed at least in part by a first propellant is disclosed. The first propellant is made up of a multi-grained fuel grain, and the center chamber is further formed by an annular section. The intra-wall channel is configured to permit combustion gases to flux between the non-permeable outer shell and the annular section. The annular section configured to permit combustion gases to flux through pores embedded within the annular section.

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 discloses a liquid rocket engine combustion tail flame detection device and method, the device comprises an air inlet pipe, an exhaust pipe, a gas chamber packaging pipe and an intelligent controller, a plurality of detection devices such as a temperature sensor and a spectrum analyzer are arranged in the gas chamber packaging pipe along the airflow direction, and the detection devices are connected with the intelligent controller to form a data transmission channel; the detection method comprises the steps that the device is initialized, and all devices are calibrated and debugged; after the engine is started, tail flame parameters are detected in real time and transmitted to the intelligent controller; the controller performs preprocessing and deep analysis on the data; comparing an analysis result with a preset normal parameter range to judge the working state of the tail flame; and if normal, the data are backed up periodically, and the long-term trend of the performance of the tail flame can also be predicted. According to the method, the tail flame parameters can be comprehensively and accurately detected in real time, intelligent analysis and abnormal early warning are achieved, and the safety and reliability of an engine are improved.

The invention discloses a carrier rocket propellant management device which comprises a porous partition plate, a screen structure, a screen support, a storage box, a conveying pipe and a pressurizing pipe. The porous partition plate is installed on the upper portion of the storage box and used for restraining the propellant from shaking in the long-time small negative overload section, and the low-temperature liquid propellant is prevented from being mixed with the air pillow. The porous partition plate is a disc formed by splicing fan-shaped pore plates and is mounted at the upper part of the storage box; the screen structure is positioned at the lower part of the storage box; the screen structure is formed by splicing a plurality of fan-shaped components into a circle and comprises an outer ring, a middle ring and an inner ring; the screen support comprises a plurality of vertical partition plates and a framework and is used for supporting the screen structure and playing an anti-vortex role at the same time; the framework is fixed on the vertical partition plate; the storage box is communicated to a rocket engine through a conveying pipe; and the pressurizing pipe is used for conveying pressurizing gas to the storage box. The device has the functions of gas isolation and liquid storage, can meet the matching requirements of a sinking engine, is beneficial to secondary weight reduction, and improves the effective load.

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.

Systems and methods for a fully reusable upper stage for a multi-stage launch vehicle are provided. The reusable upper stage uses an aerospike engine for main propulsion and for vertical landing. A heat shield can include a plurality of scarfed nozzles embedded radially around a semi-spherical surface of the heat shield, wherein inboard surfaces of the plurality of scarfed nozzles collectively define an aerospike contour. The heat shield can be actively cooled to dissipate heat encountered during reentry of the upper stage.

The invention relates to a multi-branch diffusion box bottom force transmission structure, and belongs to the technical field of aerospace vehicle propellant storage box structures, the multi-branch diffusion box bottom force transmission structure comprises a thrust ring base plate, a thrust ring outer ring, a thrust ring butt joint plate and a thrust ring inner ring, the thrust ring base plate is of a flat plate structure with a hole in the center, the thrust ring outer ring is of a cylindrical structure, and the inner diameters of the upper end face and the lower end face are different; the thrust ring butt joint plate is arranged on the upper end face of the thrust ring outer ring and used for being connected with an engine rack; the thrust ring inner ring is of a cylindrical structure with the upper end face and the lower end face being the same in inner diameter and is arranged in a cavity of the thrust ring outer ring, one end of the thrust ring inner ring is connected with the thrust ring butt-joint plate, and the other end of the thrust ring inner ring is connected with the thrust ring base plate which is connected with the storage tank bottom. The thrust ring structure is additionally arranged on the outer side of the bottom of the storage box, the thrust ring is connected with the bottom of the storage box through multiple branches, thrust transmitted by an engine rack is diffused to a main structure of the bottom of the storage box through the thrust ring, and the effects of saving an engine force transmission section and improving the engine force transmission efficiency are achieved.