54 results about "Rocket engine nozzle" patented technology

The invention provides a spacecraft quartz powder heat insulation tile which is prepared by adopting high-purity quartz powder (SiO2) as a main raw material and combining a high-temperature-resistant ceramic fiber reinforcement and a nano aerogel heat insulation layer through a high-temperature sintering process. The heat insulation tile has the following advantages: 1, the heat insulation tile has excellent high temperature resistance, and can tolerate 1600 DEG C high temperature for a long time and tolerate 1800 DEG C instantaneous thermal shock for a short time; 2, repeated use: the structure is stable, and high strength is still kept after multiple thermal cycles; light weight: the density is low (less than or equal to 0.5 g / cm < 3 >), and the load of a spacecraft is reduced; and 4, the thermal conductivity is low (less than or equal to 0.05 W / m.K), and heat transfer is effectively isolated. The method is suitable for thermal protection systems (TPS) of spacecrafts, such as thermal insulation of reentry capsules and rocket engine nozzles.

The invention belongs to the technical field of computer aided design, and provides a C / C composite material thermochemical response and pore scale structure evolution prediction method, which comprises the following steps: constructing a geometric model of the pore scale of a rocket engine jet pipe; establishing a fluid flow model in a D2Q9 form; establishing a multi-component mass transport model in a D2Q5 form; establishing a heat transfer model in a D2Q5 form; establishing a heterogeneous chemical reaction model of the fluid-solid boundary; establishing an updating model of the composite material based on a volume pixel method; and based on a geometric model, a fluid flow model, a multi-component mass transport model, a heat transfer model, a heterogeneous chemical reaction model and an update model, predicting the thermochemical response and pore scale structure evolution of the C / C composite material in the rocket engine nozzle. The method is provided for studying the ablation process of the C / C composite material, and theoretical basis and guidance direction are provided for thermal performance study and fine design of the spray pipe.

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 present application relates to the technical field of liquid rocket engine nozzle, and discloses a manufacturing method of an extension section of a large-scale liquid rocket second-stage engine nozzle, which is characterized in that the extension section is divided into n cylinder sections according to the size of the plate and the elongation rate of the plate, and each cylinder section is divided into m melon pieces to be prepared by means of tailor welding; the number n of the cylinder sections ensures that the number of the girth welds of the extension section is the least, and the welding of the girth welds is maximally avoided; the cylinder sections of the extension section are obtained by stress relaxation heat correction treatment of preformed cylinder sections, and the uncontrollable welding deformation problem is overcome. The present application solves the problem that the uncontrollable deformation caused by the springback of the sheet metal forming in the block welding leads to the difficulty in guaranteeing the profile tolerance of the inner surface of the nozzle extension section, improves the welding quality, reduces the rejection rate and production cost of the finished product, improves the production efficiency, has a wide application prospect, and any large-diameter thin-walled product can be processed according to the method of the present application.

The invention relates to a positioning and pressing device for assembling a liquid rocket engine jet pipe blanking cap, which comprises a circumferential positioning assembly for fixing the relative position of the device and a liquid rocket engine jet pipe, and a supporting assembly, an adjusting assembly, a guide shell and a rotating assembly which are uniformly distributed; the axial pressing assembly is used for achieving centering and reliable pressing after the blanking cap is pasted and comprises a telescopic assembly and a pressing part. The guide shell is of a hollow structure and wraps the rotating assembly. One end of the supporting assembly is rotationally connected with the rotating assembly, and the other end extends out of the guide shell and can axially and telescopically move along the guide section of the guide shell, so that the end face of the supporting assembly is attached to the inner wall molded surface of the liquid rocket engine jet pipe; the adjusting assembly is fixed in the guide shell and movably connected with the rotating assembly, the rotating assembly is driven to rotate through telescopic movement of the adjusting assembly, and then the supporting assembly is driven to stretch out and draw back in the axial direction of the guide section of the guide shell. The problem of poor centering performance is solved, and the blanking cap assembling efficiency is improved.

The invention provides a rocket engine exhaust nozzle blanking cap which is characterized in that a pressure-bearing surface of the blanking cap is a spherical curved surface which is convex inwards and concave outwards, and reinforcing ribs which are radially and uniformly distributed are arranged on the concave surface; the thickness distribution of the pressure-bearing surface is in gradient change: the circle center area is thinnest and is gradually thickened towards the edge; the blanking cap is integrally formed through injection molding; the injection molding material is short carbon fiber reinforced polyether-ether-ketone and contains 10-30% by mass of short carbon fibers. Under the action of internal and external pressure difference of 0.2 + / -0.02 MPa, a bonding interface is free of debonding, the blanking cap is free of fracture or excessive deformation, excellent airtight and liquid-tight performance is shown, the blanking cap can be reliably fractured within the pressure range of 0.20-1.5 MPa, the number of generated fragments is 8-10, and the size is small.

The invention discloses a forming device and method for a cloth clamping rubber blanking cap of a solid rocket engine jet pipe, and relates to the technical field of rocket accessories. The extension pipe is mounted on the manipulator, a plurality of through grooves are formed in the peripheral wall of the extension pipe, supporting pipes are arranged in the through grooves, and the supporting pipes can slide and rotate along the through grooves; the supporting pipe is provided with a connecting groove; the extrusion block can penetrate through the connecting groove and is fixedly connected with the extension pipe; the adjusting rod is in sliding connection with the supporting pipe, and the adjusting rod abuts against the extrusion block; the transverse plate is rotationally connected with the supporting pipe. A first vertical plate and a second vertical plate are fixed to the two ends of the transverse plate. The supporting plate penetrates through the transverse plate, is in sliding connection with the transverse plate and abuts against the adjusting rod, and the contact block is fixedly connected with the supporting plate; the connecting rod abuts against the contact block and is in sliding connection with the first vertical plate. The bearing plate is fixedly connected with the connecting rod and forms a clamping gap with the second vertical plate. The device is reasonable in design structure, and the coaxiality of the rubber layer and the clamping cloth can be guaranteed.

The utility model discloses a liquid-propellant rocket engine jet pipe sealing device, which belongs to the technical field of aerospace and comprises a connecting rod. The inner connecting block is rotationally mounted at one end of the connecting rod and is attached to the contraction section of the thrust chamber; the sealing assembly is arranged at the other end of the connecting rod in a sleeving mode and attached to the expansion section of the thrust chamber. According to the sealing device provided by the utility model, the effect of sealing the thrust chamber is achieved through the matching of the inner connecting block and the sealing assembly. And through the arrangement of the inner connecting block and the blocking cover, the sealing device is suitable for thrust chambers of different sizes, the applicability of the sealing device is improved, and operation is easy and convenient.

The application provides a solid rocket engine nozzle thrust contrast test device, which comprises a front head, a combustion chamber, a rear head, a nozzle A, a nozzle B, an upper clamp, a lower clamp, a six-component force measuring platform, a hexagonal bolt, a hexagonal nut, an internal hexagonal cylindrical head screw and a stud. The two nozzles A and B with different aerodynamic profiles are fixed on the end face of the rear head through the internal hexagonal cylindrical head screw; the combustion chamber is fixed by limiting with the upper clamp, the lower clamp, the hexagonal bolt and the hexagonal nut; the base of the lower clamp is fixed on the working plane of the six-component force measuring platform through the hexagonal bolt, the hexagonal nut and the stud; during the test, the high-temperature and high-pressure gas generated by the combustion chamber is simultaneously sprayed out from the nozzles A and B, and the generated thrust is transmitted to the six-component force measuring platform through the upper clamp, the lower clamp and the connecting piece. The application can realize the accurate test and contrast of the thrust performance of the two nozzles with different aerodynamic profiles under the same engine working condition.

The present application relates to the field of rocket engine nozzle testing, and particularly relates to a flexible rocket engine nozzle cold swing test simulation test system and a test method thereof, the system comprising a pressure vessel, a servo mechanism, an inclinometer, a linear displacement sensor and a real-time simulator, the pressure vessel provides a cold pressurized environment, the servo mechanism drives the nozzle to swing, the inclinometer calibrates the swing angle, and the linear displacement sensor detects the swing and outputs an analog signal. During testing, the swing angle-voltage relationship is calibrated through the inclinometer and the linear displacement sensor; then the real-time simulator sends a sine instruction to drive the nozzle to swing, and synchronously collects the displacement signal, analyzes the dynamic characteristics of the servo mechanism and the coordination with the nozzle. Through the swing angle-voltage equation obtained by calibrating the swing, combined with continuous attitude simulation controlled by the simulator, the nozzle swing angle data is calculated, and the dynamic response of the servo mechanism under the pressurized condition is evaluated, the present application can truly reflect the continuous swing working condition of the nozzle, and improve the reliability and authenticity of the test data.

The present application relates to a rocket engine nozzle, in particular to a rocket engine nozzle with a semi-extended nozzle and a design method thereof, which solves the technical problem that the use of an axisymmetric nozzle with a fixed area ratio will cause additional performance loss in the under-expanded state, and the use of a large area ratio nozzle will cause structural interference, thereby limiting the improvement of the comprehensive performance of the rocket. The present application is based on the original parallel nozzle, and the outer nozzle is designed to be semi-extended, thereby achieving the purpose of increasing the area ratio of the outer nozzle and improving the comprehensive performance of the rocket, increasing the effective wind area of the rocket, providing additional lateral moment when the engine swings, being more conducive to the flight control of the rocket, and having the advantages of simple structure, no actuator, high reliability and the like. The profile of the semi-extended nozzle is designed by using the characteristic line method, and the wall cooling scheme adopts the combination of ablation cooling and radiation cooling, which has high compatibility with the traditional bell-shaped nozzle technology and is more realizable.

The application provides a liquid rocket engine model use method and system suitable for one-dimensional system simulation, and comprises the following steps: a liquid rocket engine model establishment step: establishing a liquid path orifice model and a gas path orifice model; making the liquid path orifice model equivalent to the flow characteristics of a liquid rocket engine valve and a head part, and making the gas path orifice model equivalent to the flow characteristics in a liquid rocket engine nozzle; a liquid rocket engine model running step: running the established liquid rocket engine model to perform dynamic calculation of the liquid rocket engine. The application uses simple liquid path orifice models and gas path orifice models to respectively equivalent to the flow characteristics of an engine valve and a head part and the flow characteristics in a nozzle, thereby greatly reducing the difficulty of one-dimensional modeling of the engine.

The application discloses a method for improving hydrogen embrittlement resistance of iron-based superalloy by grain boundary regulation and application thereof. The application uses iron-based superalloy powder containing trace elements as raw material, and regulates hydrogen embrittlement resistance of the iron-based superalloy by laser powder bed melting process combined with vacuum sealing tube treatment and five-step heat treatment process. The first three steps are high-temperature heat treatment, and the last two steps are low-temperature heat treatment. The hydrogen embrittlement resistant iron-based superalloy obtained by the process has high strength and plasticity matching at room temperature and 600 DEG C, and has small elongation loss after hydrogen charging, and the hydrogen embrittlement resistance is significantly better than that of Inconel 718 nickel-based alloy. The process regulates the grain boundary type and element distribution in the alloy through multi-stage synergistic regulation, not only significantly improves the comprehensive mechanical properties of the material, but also endows the material with excellent hydrogen embrittlement resistance, can effectively meet the strict requirements of liquid hydrogen rocket engine nozzle and key parts of aviation system on high reliability, long service life and extreme environment adaptability, and has wide application prospect.

The utility model discloses a rocket engine jet pipe blanking cap bonding tool which comprises a cover plate with a threaded hole in the center, a pressing rod connected to the threaded hole in the center of the cover plate in a threaded mode and a pressing block fixed to the front end of the pressing rod through a screw. A positioning face matched with the rear end face of the blanking cap is machined on the front end face of the pressing block. An annular connecting part with threads is forwards machined at the outer edge of the cover plate, and the internal threads of the annular connecting part of the cover plate are matched with the external threads at the outer edge of the rear end of the spray pipe; the cover plate is fixed to the rear end of the spray pipe through the annular connecting part, the pressing block at the front end of the pressing rod is made to forwards abut against the rear end face of the blocking cover by rotating the pressing rod, and the blocking cover is made to be pressed and fixed to the inner wall of the expansion section of the spray pipe. The bonding tool ensures that the bonding tool is completely attached to the matching surface of the blanking cap, and the problem that the blanking cap is difficult to bond is solved.

The invention discloses a testing device for the flow resistance performance of a liquid rocket engine nozzle, and relates to the field of mechanical structure design, and the testing device comprises a tray which is provided with a first through hole; the shell is provided with an internal cavity, and is provided with a second through hole and a third through hole which are communicated with the internal cavity; the annular part of the product is pressed between the tray and the shell; the product is inserted into the internal cavity through the second through hole; the sealing screw is in threaded connection with the third through hole of the shell, and one end of the sealing screw enters the inner cavity through the third through hole and compresses the product; and the gland is in threaded connection with the shell and covers the outer side of the sealing screw. The nozzle is fixed, and the flow resistance characteristic of runners uniformly distributed on the side surface of the nozzle is measured.

A composite blanking cap structure of an underwater solid rocket engine jet pipe comprises a jet pipe shell, a jet pipe lining, a blanking cap, a blanking cap pressing plate, a fireproof plate and a sealing ring, the jet pipe shell and the jet pipe lining are bonded through an adhesive, the blanking cap, the blanking cap pressing plate and the fireproof plate are sequentially assembled on the jet pipe shell, axial sealing is achieved between the blanking cap and the jet pipe shell through the sealing ring, and the sealing ring is arranged between the blanking cap and the jet pipe shell. An annular square groove is formed in the side, making contact with the spray pipe shell, of the plug cover, the diameter of the center of the annular square groove is the same as the diameter of an inner hole of the plug cover pressing plate and the diameter of an inner hole of the fireproof plate, and the spray pipe shell, the plug cover and the plug cover pressing plate are radially positioned through hole-shaft fit. By means of the structural design, the underwater solid rocket engine nozzle blanking cap has the functional characteristics of being good in sealing performance, weak in internal pressure bearing capacity and high in external pressure bearing capacity, meanwhile, it can be guaranteed that the coaxiality among the nozzle shell, the blanking cap body and the blanking cap pressing plate is good, high thermal protection capacity is achieved, and the reliability of the overall structure of the nozzle can be guaranteed.

A solid rocket engine underwater thrust prediction method comprises the steps of solid rocket engine one-dimensional inner trajectory pre-calculation, solid rocket engine charge combustion three-dimensional structure topology evolution simulation and solid rocket engine nozzle-gas bubble-water field coupling-modeling. Wherein the one-dimensional trajectory pre-calculation of the solid rocket engine is based on an engine ignition working process simulation model, and a time-varying curve of each physical parameter in a combustion chamber in an engine ignition transient process under each water depth condition is calculated; solid rocket engine charge combustion three-dimensional structure topological evolution simulation is based on a propellant non-parallel layer combustion hypothesis, and a charge combustion process topological structure change rule when a variable water depth engine works is obtained through simulation; according to the solid rocket engine nozzle-gas bubble-water field coupling modeling, the combined action of a flow field and wake flow in a combustion chamber is comprehensively considered, and the underwater solid rocket engine thrust prediction method of the wake flow of the solid rocket engine nozzle under the high-back-pressure water medium effect is established.

The invention relates to a thrust follow-control solid rocket engine pintle cooling structure which comprises a pintle, a hemispherical shell is arranged at one end of the pintle, a cavity is formed in one section in the pintle in the axial direction of the pintle, and one end of the cavity extends into the hemispherical shell; a cavity tube is arranged on the other section in the throat bolt in the axial direction, one end of the cavity tube is communicated with the other end of the cavity, and the other end of the cavity tube extends to the other end of the throat bolt; a plurality of rows of air film holes are formed in the wall surface, from one section of the cavity to the hemispherical shell, of the pintle, the plurality of rows of air film holes are distributed in the wall surface in the axial direction of the pintle, and a plurality of air film holes are uniformly formed in each row of air film holes in the circumferential direction of the pintle; one end of the air film hole is communicated with the cavity, and the other end of the air film hole is communicated with the outer side of the throat bolt. According to the throat bolt cooling structure, air film cooling and sweating cooling are integrated, the cooling efficiency is high, the cooling area is wide, and the cooling efficiency is greatly improved through combination of the two cooling modes.

The invention provides a rocket engine jet pipe machining method and tool, and relates to the technical field of engine jet pipe machining. According to the rocket engine jet pipe machining method and tool, a fixed base is included, two vertical mounting bases are fixedly connected to the upper surface of the fixed base, and a horizontal moving base capable of horizontally moving is connected between the two vertical mounting bases. According to the paint spraying device, an output shaft of a third servo motor drives a second bevel gear to rotate, the second bevel gear drives four first bevel gears to rotate, and the first bevel gears drive second threaded rotating rods to rotate, so that a threaded mounting block drives a movable sliding strip to walk, and four driving long-strip rods drive four paint spraying pipes to get close to one another; the driving air cylinder is started according to needs, the angle between the driving long strip rod and the coating spraying pipe is changed through the driving air cylinder, the coating spraying pipe is parallel to the inner wall of the rocket engine spraying pipe, the distances between the multiple spraying holes and the inner wall are close, and the spraying uniformity is greatly improved.

The application provides a rocket engine tail nozzle closure, the pressure bearing surface of the closure is a spherical surface with an inner convex and outer concave shape, and reinforcing ribs are uniformly distributed in a radial manner on the concave surface; the thickness of the pressure bearing surface changes in a gradient manner, the thinnest part is in the center of the circle, and the thickness gradually increases towards the edge; the closure is integrally formed by injection molding; the injection molding material is short carbon fiber reinforced polyether ether ketone, which contains 10-30% short carbon fibers in terms of mass fraction. The application can take into account both low pressure sealing and high pressure reliable rupture: under the action of an internal and external pressure difference of 0.2±0.02 MPa, the adhesive interface is not detached, the closure is not ruptured or excessively deformed, and excellent air and liquid tightness is exhibited, while reliable rupture can occur in a pressure range of 0.20-1.5 MPa, and the number of fragments generated is 8-10 pieces and the size is small.

The invention belongs to the technical field of engine parameter calculation, and particularly relates to an air suction cruise hypersonic aircraft nozzle parameter equivalent estimation method, which comprises the following steps of: enabling an air inlet channel, an isolation section and a combustion chamber of a scramjet engine to be equivalent to a combustion chamber of a rocket engine; the contraction ratio of the jet pipe of the scramjet engine is equivalent to the expansion ratio of the jet pipe of the rocket engine; the method comprises the following steps: enabling air flow and fuel injection flow entering a combustion chamber of the scramjet engine to be equivalent to premixed gas which is uniformly mixed according to a preset proportion, and determining formula parameters of an equivalent propellant based on the premixed gas; converting the flow field parameters at the air inlet of the scramjet engine into the total temperature and the total pressure of a combustion chamber of an equivalent rocket engine; calculating nozzle parameters of the scramjet engine by adopting a thermodynamic calculation method based on the parameters obtained in the above steps; the problems that an existing scramjet engine nozzle parameter is large in solving difficulty and high in calculation cost are solved.

Seals for gimbaling and / or fixed rocket engine nozzles, and associated systems and methods are disclosed. A representative rocket propulsion system includes a rocket engine having an exhaust nozzle, a seal plate carried by the exhaust nozzle, and a seal engaged with the seal plate. The seal includes at least one support, multiple pivotable first flaps, carried by the at least one support and positioned to contact the seal plate, and multiple pivotable second flaps, with an individual second flap positioned to shield a corresponding individual first flap. At least one forcing element is operatively coupled to at least one of the individual first flap or the individual second flap, to apply a pivoting force to the at least one of the individual first flap or the individual second flap.

The invention discloses a novel pressure hole structure for a solid rocket engine jet pipe. One side of the spray pipe shell is provided with a pressure-measuring filler neck and a sealing ring mounting hole, and the other side of the spray pipe shell is bonded and assembled with a convergence ring; an air guide sleeve is arranged on the convergence ring; the whole air guide sleeve is a rotary body, a blind hole with a large diameter is formed in the center of the air guide sleeve and serves as a main air guide hole, and a plurality of auxiliary air guide holes with small diameters are formed in the circumferential face of the top of the air guide sleeve and communicated with the main air guide hole. When the engine works, the auxiliary gas guide hole makes direct contact with gas, and gas pressure is collected to the main gas guide hole through the auxiliary gas guide hole and then measured through the pressure measuring filler neck. The novel pressure measuring hole structure has the advantages that the opening direction of the auxiliary gas guide holes in the gas guide sleeve is along the gas direction and is basically parallel to the gas flowing direction, and the gas guide holes are not prone to being blocked by particle impurities in gas; even if an individual auxiliary gas guide hole in the auxiliary gas guide holes is accidentally blocked, stable pressure measurement can still be achieved through other auxiliary gas guide holes, and the stability and reliability of pressure measurement of the solid rocket engine are improved.

This invention relates to the field of rocket engine nozzle testing technology, specifically disclosing a pneumatic control loop for nozzle load simulation. The loop includes a servo control loop, a manual control loop, an electronically controlled pressure regulating valve control loop, a master control valve, and a master controller. The servo control loop includes a first pressure regulating valve, and the electronically controlled pressure regulating valve control loop includes a second pressure regulating valve and an electronic pressure controller. The master controller provides working air pressure to the electronic pressure controller by controlling the second pressure regulating valve, and the electronic pressure controller controls the opening of the first pressure regulating valve. This invention uses high-pressure gas from a gas source, which enters the servo control loop or manual control loop through the master control valve before being input into the actuation chamber. This simulates the frictional torque experienced by the rocket engine nozzle, enabling a semi-physical simulation experiment of the nozzle. The electronic pressure controller acts as a pilot valve, controlling the gap between the valve core and valve seat in the first pressure regulating valve, thereby achieving precise control of the valve opening and a wider pressure regulation range.

The invention relates to a simulation method for manufacturing a rocket engine jet pipe through an inner and outer wall integrated forming scheme based on ABAQUS. The method specifically comprises the steps that ABAQUS simulation pretreatment is conducted; aBAQUS simulation solution is carried out; and ABAQUS simulation post-processing is carried out. According to the ABAQUS-based simulation method for manufacturing the rocket engine jet pipe through the inner wall and outer wall integrated forming scheme, the relation between geometric parameters, material parameters and manufacturing process parameters of the jet pipe and the performance of the jet pipe is effectively established; the optimized parameter combination is obtained by adjusting different numerical values of all parameters, the problem of cost and process limitation in the actual manufacturing process is solved, the method is a new idea for obtaining optimized parameters, and the method has important significance on selection of actual parameters and is suitable for popularization and application. The relation between geometric parameters, material parameters and manufacturing process parameters of the spray pipe and the performance of the spray pipe is effectively established, and the structural strength and the overall performance of the spray pipe can be remarkably improved by accurately controlling and optimizing the parameters.

The invention belongs to the technical field of engine exhaust nozzles, and discloses a small-length-width-ratio air turbine rocket engine exhaust nozzle design method based on flow matching, a circle-to-square transition shrinkage pipe section is designed, and the circle-to-square transition shrinkage pipe section comprises a circular end used for being connected with an air turbine rocket engine combustion chamber outlet and a square end far away from the circular end; a two-dimensional spray pipe section is designed and comprises straight panels fixedly connected to the end faces of two side plates of the square end, an adjusting plate rotationally connected to the end face of a bottom plate of the square end and a convergence and expansion plate fixedly connected to the end face of a top plate of the square end; the convergence expansion plate comprises a convergence section in smooth transition connection with the square end, a straight throat section in smooth transition connection with the tail end of the convergence section, and an expansion section in smooth transition connection with the straight throat section. By the adoption of the method, the structure of an axisymmetric convergence-expansion exhaust nozzle adjusting unit used by an existing air turbine rocket engine is simplified, the reliability of the axisymmetric convergence-expansion exhaust nozzle adjusting unit is improved, and integrated design of the axisymmetric convergence-expansion exhaust nozzle adjusting unit and other exhaust nozzles is achieved.

The application provides a rocket engine nozzle temperature measuring method and system, comprising the following steps: step S1: setting a standard temperature Te under a non-shielding condition; step S2: preparing a test piece, forming a base, fixing a platinum-rhodium thermocouple in contact with the base, and configuring a plurality of ceramic protection tubes on a thermocouple measuring end; collecting platinum-rhodium thermocouple temperature data and simultaneously measuring temperature by using a two-color infrared thermometer; step S3: taking the temperature measured by the two-color infrared thermometer in step S2 as a reference, and establishing a platinum-rhodium thermocouple calibration curve; step S4: removing the outer surface of a throat part of a rocket engine nozzle by machining to form a base; fixing a platinum-rhodium thermocouple in contact with the base, and connecting the thermocouple to a temperature data collection system; and step S5: obtaining a throat temperature Tt by interpolation calculation according to the platinum-rhodium thermocouple calibration curve. The application solves the problem that the throat wall temperature of a key parameter of a nozzle cannot be measured by an infrared thermometer.

The invention discloses a permeable nozzle fluid-solid coupling effect cold flow test bed and a test method, and is applied to the technical field of rocket engine nozzle tests. The strain pipe with high yield strength and low elastic modulus is adopted as a measuring element, the fluid-solid coupling effect of the spray pipe is amplified through material characteristics, high-precision strain gauges are arranged at key positions of the tail of the strain pipe and the tail of the spray pipe respectively, and synchronous and accurate measurement of the overall structural response and the local dynamic response of the spray pipe is achieved. And meanwhile, the opening degree of an electromagnetic valve of the air supply system is dynamically adjusted by monitoring a pressure intensity signal of the steady flow section in real time, and when the pressure intensity of the steady flow section reaches a preset threshold value, the electromagnetic valve is automatically closed, so that the pressure ratio of the spray pipe is in a change trend of first rising and then falling, and the working process of the spray pipe executing a lift-off and landing recovery task is simulated. In addition, the system can also simulate the lift-off and landing processes of the spray pipe in different working time periods by adjusting the combination of air storage tanks connected in series in the flow channel.

The invention relates to a manufacturing method of a rocket engine jet pipe with a buffer cavity, and belongs to the technical field of aerospace engine manufacturing. The manufacturing method comprises the steps that a three-dimensional model of the spray pipe is sectioned in the circumferential direction and segmented in the axial direction, and a plurality of sub-components are formed; a selective laser melting process is adopted, the sub-component is printed on equipment with the breadth of 650 mm, and a buffer cavity which is communicated with an internal cooling channel and extends in the circumferential direction is pre-formed in the axial butt joint face of the sub-component; the upper sub-section and the lower sub-section are connected into fan blade units through transverse welding, and then the multiple fan blade units are spliced into the complete spray pipe through longitudinal welding. During welding, protective gas is guided into the inner channel through the protective gas hood at the top end and flows out of the groove to form inner side protection. The buffer cavity extending in the circumferential direction can release thermal stress and enhance cooling turbulence during working. According to the method, low-cost and high-precision manufacturing of the nearly 3-meter spray pipe is realized by using conventional 650-mm-level equipment.

The present invention relates to a liquid rocket engine nozzle back-pressure low-temperature spray test system. A low-temperature helium medium is supplied via a helium cylinder, a helium delivery pipeline, a heat exchanger, and a low-temperature helium delivery pipeline. Cryogenic cooling of the helium is achieved via a nitrogen cylinder, a nitrogen booster delivery pipeline, a liquid nitrogen container, and a liquid nitrogen cooling delivery pipeline. A back-pressure chamber back-pressure environment is achieved via a nitrogen cylinder and a nitrogen back-pressure delivery pipeline. Liquid nitrogen test medium is supplied via a nitrogen cylinder, a nitrogen booster delivery pipeline, a liquid nitrogen container, and a liquid nitrogen test delivery pipeline. The back-pressure chamber is used to secure the product and create a low-temperature back-pressure spray environment, allowing for visual measurement of the spray field. This allows for measurement of the atomization characteristics of the nozzle in a low-temperature back-pressure environment.