31 results about "Rocket engine nozzle" patented technology

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

The 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 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.

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.

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.

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 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 utility model relates to the field of hoisting tools, in particular to a solid rocket engine nozzle diffusion section hoisting and transferring tool. Comprising a supporting ring, a supporting ring felt, a limiting block felt, a limiting block, a lifting ring screw and a lifting belt complete set rigging, the supporting ring felt is bonded to the surface of the supporting ring, a threaded hole used for fixing a lifting ring is formed in the supporting ring, the limiting block is fixed to the upper surface of the supporting ring in a welded mode, and the limiting block felt is bonded to the inner side of the limiting block 4. And the sling complete set of rigging is used for hoisting by utilizing a lifting hook of the sling complete set to penetrate through the lifting bolt. The lifting and transferring tool is not limited by the space of the small end of the diffusion section, the supporting ring felt and the limiting block felt can protect the large end of the diffusion section from being scratched, and the four lifting belts of the four-leg lifting belt complete set rigging can play a role in limiting the diffusion section and can also play a role in protecting the outer surface of the diffusion section. And the outer surface of the diffusion section cannot be damaged in the hoisting process.

The application discloses a kind of liquid rocket engine nozzle brazing defect laser shearing speckle interferometry detection methods, suitable for liquid rocket engine nozzle brazing surface microdeformation and internal defect detection. Controllable internal pressure is applied to nozzle inner cavity by water pressure or air pressure active loading mode, which is different from traditional thermal loading, vacuum loading, vibration loading. By uniform speed pressurization to preset detection pressure value and stabilize, reference phase shift diagram is collected by shearing speckle interferometry optical path, and reference phase of reference state is obtained by multi-step TPS method reconstruction. After stopping loading, target phase of target state is obtained by the same method in the reconstruction recovery process. Full-field phase difference is calculated by phase reconstruction device and phase difference image is generated. The abnormal stripe area presented in the image is the nozzle brazing defect position. The application can stimulate microdeformation of defect area by internal pressure loading, which can adapt to complex structure of nozzle, improve detection sensitivity and eliminate background interference.

This invention provides a fixing device for rocket engine nozzles, including a ring frame. Several connecting arms are evenly distributed around the center of the ring frame at its lower end. A support arm is provided at the upper end of the ring frame corresponding to each connecting arm. Each connecting arm includes an integrally formed docking portion and a connecting portion. The docking portion is vertically arranged, and the connecting portion extends outward from the ring frame, with a connecting corner piece installed at its free end. The docking portion is fixed to the ring frame by a connecting assembly. The support arm includes a fixing frame, with an arc-shaped support plate at one end extending outward from the ring frame. The support plate is inclined, gradually tilting towards the center of the ring frame from bottom to top, and has a heat insulation structure on its outer side. The ring frame includes a paper honeycomb core material, with a carbon fiber composite panel covering the outer wall of the paper honeycomb core material. This invention effectively achieves lightweight and high specific stiffness, integrating structural load-bearing and thermal protection, avoiding the weight increase caused by external heat insulation devices.

The invention discloses a visual large-expansion-ratio ground test jet pipe, and belongs to the field of solid rocket engine jet pipe ground testing. The device comprises a standard welding straight joint flange, a round-to-square air pipe, a square air pipe, a honeycomb device, a plane spray pipe main body, two side cover plates of the spray pipe, a visual window and a window cover plate, the standard welding straight joint flange is a connecting device of the spray pipe and a high-pressure cold air source; the round-to-square air pipe is a connecting part of the standard welding straight joint flange and the downstream square air pipe; the square air pipe is a connecting part of the spray pipe inlet and the round-to-square air pipe; the honeycomb device is placed into the mounting groove from one side and is connected, pressed and fixed through a flange; the honeycomb device is a rectifying component in the square spray pipe; regular hexagonal holes with large length-diameter ratios are densely distributed in the honeycomb and are used for dividing airflow and weakening vortex components in the airflow; cover plates on the two sides of the spray pipe are thick metal plates provided with window installation grooves. The pressure distribution condition of the expansion section of the nozzle can be accurately recorded, and the flow separation point position can be captured.

The invention relates to a liquid rocket engine nozzle atomization test device and a liquid rocket engine nozzle atomization test method, in particular to a liquid rocket engine nozzle atomization test device and a liquid rocket engine nozzle atomization test method which have high-frequency speed change pressure measurement capability in a back pressure environment, and solves the problem of supply of gas-liquid dual-path media when liquid rocket engine nozzle products are tested in a back pressure cabin. And the problems of high-frequency rapid change sensor sealing and sensor signal line leading-out in a back pressure environment are solved. According to the device, the split design idea is adopted, the test tool is divided into the upper cavity and the lower cavity, pre-tightening force is applied at a time through the fastener, three seals between the nozzle and the upper cavity and the lower cavity are completed at the same time, and the device is convenient to use and has the gas-liquid double-path supply capacity. According to the device, the sensor is sealed in the back pressure cabin through the sealing cavity, so that the pressure in the sealing cavity is always atmospheric pressure; a signal line of the sensor is led out from the interior of the cabin through the metal hose, and then sealing between the signal line and the back pressure cabin is completed through the switching section, so that no leakage of pressure in the cabin is ensured.

The present application relates to a kind of solid rocket engine nozzle shell strength hydraulic test tool, belong to the field of solid rocket engine;Hydraulic end is the shell structure of axial horizontal placement;Nozzle shell and hydraulic end coaxial butt joint;Snap ring is arranged at the butt joint of nozzle shell and hydraulic end, relative fixation limiting of nozzle shell and hydraulic end is realized;Nozzle shell axial middle inner wall is stepped structure;Hydraulic nipple is coaxially clamped in the step of nozzle shell inner cavity;Convergent ring is attached to the inner wall of nozzle shell axial end directed to hydraulic end;Hydraulic connecting rod is coaxially arranged at the axis of nozzle shell;The center of hydraulic nipple is provided with through hole;The axial one end of hydraulic connecting rod is fixed with hydraulic end, and the axial other end of hydraulic connecting rod is inserted into the through hole of hydraulic nipple;The present application solves the problem that nozzle shell is easily deformed after hydraulic test under the premise of overall examination of nozzle shell structure performance.

The present application relates to the technical field of solid rocket engine nozzle, in particular to a long-time small-thrust nozzle diffuser section. It comprises a nozzle shell, a back wall heat insulation layer, a front heat sink ring, a middle heat sink ring, a back heat sink ring, a baffle ring, a back heat insulation layer, a diffuser section and a throat liner. The back wall heat insulation layer is located inside the nozzle shell. The front heat sink ring, the middle heat sink ring and the back heat insulation layer are located adjacent to the back wall heat insulation layer. The throat liner is located inside the front heat sink ring and the middle heat sink ring. The diffuser section is located adjacent to the middle heat sink ring, the back heat sink ring and the throat liner along the flow direction of the gas. The back heat sink ring is located adjacent to the back heat insulation layer and the diffuser section. The baffle ring is located at the tail end of the nozzle. The present application is suitable for solid rocket engine with working time over 100 seconds, improves the specific impulse of the solid rocket engine, improves the ablation condition of the diffuser section and improves the overall reliability of the solid rocket engine.

The invention discloses a solid rocket engine corrugated expansion section flexible deformation thrust vectoring nozzle which comprises a base section and a corrugated flexible expansion section and further comprises a driving device, the base section is composed of a convergence section shell, a throat liner and an expansion section shell, the convergence section shell is fixedly connected with the expansion section shell, and the throat liner is fixedly connected with the expansion section shell. The throat liner is arranged on the inner wall of a shell of the convergence section, the corrugated flexible diffusion section comprises a metal ring and an elastic body, and the outline of the elastic body is designed in a horn shape. According to the corrugated expansion section flexible deformation thrust vectoring nozzle of the solid rocket engine, the driving device is installed on the outer wall face of the nozzle body, so that the nozzle body avoids the severe gas environment, the requirement for the material performance is remarkably reduced, meanwhile, the reliable thrust vectoring function can still be given to the nozzle body, precise control over thrust vectoring is easier to achieve, and the service life of the nozzle body is prolonged. The manufacturing cost is reduced; and the engineering practicability is improved.

The invention discloses an engineering estimation method for chemical ablation of a carbon / carbon composite nozzle, and belongs to the technical field of thermal protection of solid rocket engine nozzles. The implementation method comprises the following steps of: establishing a carbon / carbon spray pipe model, dividing grids and performing fluid-solid coupling calculation to obtain spray pipe pressure, speed and temperature field results; and flow parameters are decoupled based on the Arecitinib formula, and the contribution mechanism of the temperature, the pressure intensity and the component concentration of the combustion chamber to the ablation rate of the spray pipe is obtained. After combustion chamber parameters are changed, the relation between parameter deviations and ablation mass loss deviations is compared and analyzed, and a fitting relation between the ablation rate and the combustion chamber parameters is calibrated through experiment fitting, so that the nozzle ablation rate is efficiently and quantitatively measured, parameters influencing chemical ablation of the carbon / carbon composite material nozzle can be efficiently obtained, and the accuracy of chemical ablation of the nozzle is improved. Thermal protection design, service life evaluation and maintenance of the spray pipe are facilitated, and engineering accidents such as burnthrough are reduced.

The utility model relates to a solid rocket engine nozzle expansion section ablation examination structure. The expansion section is composed of an ablation examination sample block and a low-cost trepanning expansion section, the ablation performance of the woven expansion section material in two states can be examined at a time, the examination sample block adopts a stepped plunger type structure in order to avoid gas fire penetration at a butt joint part, and a combined body of the stepped plunger type ablation sample block and the trepanning expansion section is formed. The structure is arranged in a spray pipe shell to form a spray pipe, and the spray pipe is connected with a real solid rocket engine, so that the aim of examining the ablation performance of various woven expansion section materials through one-time solid rocket engine test run under the same working condition can be fulfilled. According to the utility model, the plurality of stepped plunger type examination sample blocks are embedded into the expansion section of the spray pipe, so that the examination of the ablation performance of the expansion section material under the actual working condition of a solid rocket engine is realized, and the examination authenticity and the examination efficiency of the expansion section material are improved.

The invention discloses a threaded foam blanking cap and a forming method thereof, and belongs to the technical field of solid rocket engine jet pipe blanking caps, the threaded foam blanking cap is mounted at the tail of a rocket engine jet pipe, and the threaded foam blanking cap comprises a blanking cap body, a first trapezoidal thread structure is arranged on the outer side of the blanking cap body; a second trapezoidal thread structure is arranged on the inner side of the check ring; the first trapezoidal thread structure and the second trapezoidal thread structure are matched with each other and connected in a screwing mode, and the matching faces of the first trapezoidal thread structure and the second trapezoidal thread structure are constructed to generate mechanical self-locking and radial pressing force during screwing fastening, so that sealing connection of the blanking cap body and the check ring is achieved. Therefore, the rocket engine jet pipe is sealed; the blanking cap body and the check ring are both formed by foaming expandable polystyrene in a one-time compression molding mode. Vibration loosening is effectively resisted, and the structural stability under lasting connection and working conditions is ensured.

The utility model relates to the technical field of sealing devices, in particular to a pressure test sealing device of a liquid rocket engine jet pipe, which comprises a base, a supporting cylinder, a sealing ring and a driving assembly. According to the utility model, through the arrangement of the sealing ring which deforms under pressure, after the spray pipe is placed, the driving assembly drives the upper end of the sealing ring to move downwards, so that the sealing ring deforms and bulges outwards, the diameter of the sealing ring is increased, the throat part of the spray pipe is blocked until the sealing ring cannot deform, and then a liquid pressing sealing performance test is carried out. Through the sealing mode, the sealing of the throat part is ensured, so that the problem that the sealing effect of the throat part of the spray pipe cannot be ensured in the prior art is solved, the time of the preparation work of the sealing performance test is shortened, and the test efficiency is improved. And the larger the downward moving distance of the upper end of the sealing ring is, the larger the deformation degree of the sealing ring is, and the larger the diameter of the sealing ring is, so that the device can adapt to spray pipes with different throat sizes, and the universality is improved.