617 results about "Rocket engine" patented technology

The invention discloses a solid rocket engine bearing structure lightweight determination method based on a carbon fiber composite material, and relates to the technical field of composite material structure design, and the method comprises the steps: obtaining the structure size and geometric shape of a fixed body; replacing a titanium alloy material with a carbon fiber composite material, and determining a carbon fiber composite material structure response model based on a finite element modeling method; determining an optimal layering scheme based on structural response sensitivity analysis and a multi-objective optimization method; then determining a thickness optimization scheme based on the structural size and geometric shape of the fixed body and load distribution; determining a reinforcement optimization scheme based on a parametric modeling method; determining a metal reinforcement configuration by adopting a structural rigidity-quality collaborative evaluation method; and finally, determining a lightweight structure of the fixed body based on a composite material layering process design method. The mass of the fixing body structure can be reduced, the mechanical property of the fixing body structure is improved, and the high-performance design requirement of a solid rocket engine is met.

The invention provides a carrier rocket engine stage-switching motor redundancy control system, which is combined with a multi-input instruction redundancy voting method, a voting path software and hardware arbitration method and the like to complete high-reliability carrier rocket engine stage-switching motor control. According to the system, two-out-of-three voting is carried out by using a triple-redundancy input control instruction and interactive heartbeat, software and hardware joint arbitration is carried out on voting results of three modules, and an instruction control channel is determined to complete phase sequence control of the stage conversion winding. The method has strong fault capability, has no single-point failure link, and can ensure the normal operation of the engine conversion motor in the flight process of the carrier rocket.

The present application relates to a kind of rocket engine injector thin-wall milling self-positioning fixture and assembly method, it mainly includes upper positioning tire body component, conical surface expansion ring, tire unloading bottom plate, pressing cover plate, tire unloading lifting ring, pressing bolt, expansion screw, lower positioning bottom plate component, locking bolt, long fixed bolt, lower positioning bottom plate component is fixed in machine tool workbench by locking bolt long-term, always not disassembled.Product is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion screw before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thin-wall;Injector thin-wall is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion screw before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thin-wall;Injector thin-wall is positioned and clamped and supported by conical surface expansion ring, tire unloading bottom plate, pressing cover plate, pressing bolt, expansion sleeve before milling injector thin-wall and positioning tire body component, limit the radial and axial movement of injector thinn

The invention discloses a digital twinborn body and a digital twinborn method for a hydraulic test system of a turbine pump of a liquid rocket engine, and solves the technical problem of high production and test resource consumption caused by research and development of engine products. The basic architecture of the digital twinning body comprises a digital twinning engine, a physical entity and a digital virtual body, and dynamic unification of digital virtual body prediction performance and object actual measurement performance is realized through virtual-real interaction and model training and optimization; finally, the display effect that a digital virtual body completely replaces a liquid rocket engine turbine pump hydraulic test system entity can be achieved, and the contradiction between increasing turbine pump hydraulic test tasks and limited physical test resources is greatly relieved.

The invention relates to the field of heat treatment processes, and discloses a deformation control method of a thin-wall special-shaped structure sheet metal part for a liquid rocket engine, the sheet metal part is made of a 30CrMnSiA trough plate, and the deformation control method comprises the following steps: vertically hanging the trough plate through steel wire binding during quenching; fixing the lower end of the trough plate in the vertical direction to prevent swinging in the quenching process; during quenching, the temperature of a quenching furnace is increased to the set preheating temperature below the quenching temperature, then the fixed trough plate part is put into the quenching furnace, the temperature is slowly increased to the quenching temperature along with the furnace, and the stress applied to a trough plate machine is pre-eliminated in the slow temperature increasing process; the temperature of quenching oil is increased to 40 DEG C or above, quenching deformation of the trough plate caused by a steam film generated by the oil is prevented, and a heat treatment production line is used for automatic feeding, turnover, quenching, cleaning and tempering heat treatment to complete the trough plate heat treatment procedure. According to the method, the hardness performance of the groove plate cannot be affected, warping and torsional deformation of the groove plate can be effectively controlled, the first-pass yield is improved, the reworking process and the rejection rate are reduced, and the production cost of the groove plate is saved.

The invention relates to the technical field of rocket launching, in particular to a steam protection rocket launching system which comprises a steam energy storage system, a steam conveying system, a steam releasing system and a launching rack structure. The steam energy storage system comprises a steam boiler, a steam pressure accumulator and a steam energy storage control unit and is used for preparing and storing high-pressure steam before rocket launching. The steam conveying system comprises a steam blending table and a steam supply pipeline and is used for conveying high-pressure steam in the steam pressure accumulator to the steam release system; the steam release system comprises a steam injection unit and a steam injection control system and is used for releasing steam when the rocket is ignited, launched or recovered, forming a steam protection air curtain and reducing thermal shock of high-temperature airflow on the structure. By establishing the high-pressure steam protection air curtain, active protection of structural parts, pipeline parts, hydraulic parts and other parts of the rocket launching system influenced by the gas tail flame area of the rocket engine is achieved, and the durability and reliability of the launching system are improved.

The application discloses a TPMS structure micro-channel heat exchanger and a manufacturing method, and belongs to the field of structural design and manufacturing of rocket engine heat exchangers. The heat exchanger core is a plurality of spiral flow channel structures adopting TPMS. The inlet side of a liquid oxygen inlet and outlet collector is connected with the inlet of the core spiral flow channel through a liquid oxygen partition plate, and the outlet side of the liquid oxygen inlet and outlet collector is connected with the outlet of the core spiral flow channel. A fuel gas inlet pipeline is connected with the inlet of the core spiral flow channel through a fuel gas partition plate, and a fuel gas outlet pipeline is connected with the outlet of the core spiral flow channel. The liquid oxygen partition plate and the fuel gas partition plate are each provided with a plurality of through holes, the through holes correspond to the spiral flow channels in the core one by one, and each through hole is connected with the inlet of one spiral flow channel in the core. The application adopts the TPMS spiral flow channel structure to increase the internal heat exchange area of the heat exchanger and improve the heating efficiency of the heat exchanger on fuel in the working process of the engine.

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

The invention provides a multi-channel temperature parameter pre-cooling process liquid rocket engine temperature field display method, and solves the problems that in the prior art, as a data acquisition system measures temperature parameters of a plurality of measuring points when a power system carries out preparatory work, the temperature parameters of the key temperature measuring points are measured on the basis of a temperature sensor; and the change condition of a structural system cannot be visually displayed after long-time numerical value display. According to the method, on the basis that simulation calculation is high in time complexity and cannot achieve real-time calculation, the recurrent neural network is adopted for feature analysis, model training is completed on the basis of gradient descent, interpolation calculation of the temperature field in the channel dimension is finally achieved, the resolution of channel dimension data can be effectively improved, then the resolution of a temperature field image is improved, and the method is suitable for large-scale popularization and application. And subsequent thrust chamber temperature field modeling and analysis based on the whole pre-cooling process are facilitated.

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

The invention relates to the field of spectrum external field testing, and discloses a rocket engine fuel gas measuring device which comprises a support and a hollow probe assembly installed on the support. In order to solve the problems of high-temperature and high-pressure impact and thermal ablation of the tail flame of the rocket, the support and the probe assembly are both designed to be of a wedge-shaped structure, internal circulating water paths are arranged on the support and the probe assembly for forced cooling, the probe assembly is of a U-shaped structure, and an optical measurement light path is arranged in the probe assembly. According to the optical measurement light path, gas parameters are measured through a light refraction principle by using a lighting tube, a prism and a collimating lens connector which are inserted into a probe, so that a measurement component is far away from a high-temperature area. According to the invention, the wedge-shaped structure, active water cooling and high-temperature-resistant materials are combined, so that the effects of good cooling performance, stable structure and ablation resistance are realized, and the measurement precision in a harsh environment can be ensured.

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

The invention discloses a sulfonated carbon nanotube-amino silicon dioxide double-phase reinforced epoxy adhesive, which is prepared from the following components in parts by mass: 0.5 to 2 parts of sulfonated carbon nanotube, 1 to 3 parts of amino silicon dioxide, 100 to 200 parts of epoxy resin matrix and 30 to 60 parts of curing agent. The sulfonated carbon nanotubes and the amino silicon dioxide form a uniformly distributed disperse network in an epoxy resin matrix, so that the load is effectively transmitted, the tensile shear strength of the adhesive is improved by 96% or above, the standard deviation of the tensile shear strength is reduced by about 59% compared with that of the unmodified adhesive, and large-scale application is facilitated. Due to the fact that the sulfonic acid base body is good in heat resistance, meanwhile, due to introduction of silicon dioxide, pinning points can be formed to inhibit high-temperature slippage of a resin chain, and the tensile shear strength retention rate of the adhesive at 200 DEG C / 30 min is about 84%; and the adhesive is beneficial to non-combustion chamber parts which have high requirements on the adhesive property at a certain temperature in tail pipes and diffusion sections in the field of solid rocket engines.

The invention provides a variable circulation hybrid electric propulsion system applied to an unmanned aerial vehicle and a propulsion method of the variable circulation hybrid electric propulsion system, and belongs to the technical field of aircraft propulsion. The problems of insufficient power and endurance of the unmanned aerial vehicle are solved. The four liquid rocket engines are arranged around the unmanned aerial vehicle body through the vehicle arms; each rotor wing is correspondingly arranged above the liquid rocket engine; a motor and a pump are arranged in the liquid-propellant rocket engine, and the combustion chamber is located below a shell of the liquid-propellant rocket engine and connected with the spray pipe. A fuel cell in the unmanned aerial vehicle body is located at the upper end of the unmanned aerial vehicle body. An energy storage cell is embedded into the lower middle portion of the fuel cell. The energy management system is embedded into the middle-upper part of the fuel cell; and the oxygen tank is positioned at the lower end of the fuselage. The fuel cell is adopted to charge the energy storage cell, the energy storage cell drives the pump to stably convey fuel to the combustion chamber, the problem that a traditional liquid rocket engine depends on complex external energy when started for multiple times is effectively solved, and rapid and reliable repeated starting of the engine is achieved.

The invention discloses an electrolytic preheating-electrolytic ignition method for a three-electrode hydroxylamine nitrate liquid propellant, and belongs to the field of liquid rocket engines and electrochemistry. Three electrolysis electrodes are arranged in the electrolysis chamber and comprise two positive electrodes and one negative electrode, so that two groups of electrolysis electrode pairs are formed. The method comprises the following steps: injecting a nitric acid hydroxylamine-based liquid propellant into an electrolysis chamber, and electrifying a group of electrolysis electrodes to start electrolysis preheating; after the preset switching time is up, closing the current working electrolysis electrode pair, and switching to another group of non-working electrolysis electrodes to continue electrolysis; and low-voltage and rapid electrolytic ignition is realized. According to the three-electrode electrolysis structure, the time for ignition and reaction to reach the highest temperature can be shortened, and the system response speed is increased; the preheating and ignition voltage can be changed, and the large voltage adjusting and optimizing space is achieved; meanwhile, the electrolytic ignition current is relatively stable, and the instantaneous power consumption is small; and electrolysis power consumption can be saved, electrolysis electrode loss is reduced, and reuse of the electrode is facilitated.

The invention belongs to the technical field of ignition, and discloses a three-component pintle rotational flow torch igniter. According to the three-component pintle rotational flow torch igniter, liquid fuel is jetted through the central pintle rotational flow nozzle, and a gas oxidizing agent is jetted through the peripheral rotational flow nozzle. The liquid fuel from the pintle swirl nozzle moves in the radial direction, and the gas oxidizing agent from the swirl nozzle rotates in the axial direction, so that the interaction between the liquid fuel and the gas oxidizing agent is enhanced, the atomization quality is improved, and the ignition success rate is further improved. Besides, the flame temperature of the torch igniter is increased through secondary oxygen or air jetted from the body part of the torch igniter, and meanwhile, the wide-range adjustment of the flame temperature of the torch igniter is realized. The device is suitable for large combustion devices such as liquid rocket engines, gas generators and combustion heaters.

The invention relates to a spray cooling system for a rocket engine test bed. The spray cooling system comprises a spray ring module and a diversion trench module, the spraying ring module is arranged at the wellhead position of a test bed, is coaxial with the center of a test frame, is spirally arranged by adopting a stainless steel seamless steel pipe, and is in threaded connection with a ground pre-embedded steel plate of the test bed; the diversion trench module is arranged at the bottom of the test bed and comprises a supporting frame and a diversion trench obliquely formed in the top of the supporting frame, the diversion trench is composed of a core area spraying structure and a non-core area spraying structure, and the supporting frame is welded and fixed to a pre-embedded steel plate on the ground of the test bed; and the spraying ring module and the diversion trench module are connected with an external water source through a water pump and a valve. According to the invention, the wellhead foundation position can be protected from ablation, gas can be quickly guided to a place far away from a test bed, and flames, smoke dust and disintegrating slag are prevented from being reflected back to impact the test bed and damage an engine and other corollary equipment. The spray cooling system can meet the cooling and noise reduction requirements in the test run process of rocket engines with different thrust.

The invention discloses a digital twin system for a liquid-propellant rocket engine, relates to the digital technical field of liquid-propellant rocket engines, and aims to solve the problem that a digital twin system in the prior art cannot meet data driving interaction between a test model of the liquid-propellant rocket engine and a physical space and real-time synchronous feedback of the state of each test part. The system at least comprises a physical space module, a digital space module, a training module, a model construction module, a service platform and a core support module, based on this, the system can establish a physical field of a plurality of target test parts according to test data of the plurality of target test parts of the liquid rocket engine, perform real-time data driving interaction between the physical field and a physical space, and perform real-time synchronous feedback on the state of each test part; data driving interaction between a test model and a physical space and real-time synchronous feedback of the state of each test part in a liquid rocket engine test are met.

The utility model relates to a kind of liquid rocket engine ground test drive's emergency shutdown control system and test device. Among them, liquid rocket engine ground test drive's emergency shutdown control system, including PXI device, SIS system, programmable logic controller, first engine working condition abnormality judging sensor, second engine working condition abnormality judging sensor, pressure sensor, flammable hydrocarbon gas detector, flame detector, control electric circuit and solenoid valve. By applying a variety of sensors, and using PXI device, SIS system and programmable logic controller are composed of redundant control system, reach the purpose of improving liquid rocket engine ground test emergency shutdown reliability, enhance the safety of liquid rocket engine ground test.

This invention discloses a design method and inducer for a liquid rocket engine turbopump. Based on the overall design requirements of the liquid rocket engine, the operating conditions of the inducer are determined. According to the design requirements and empirical formulas, the basic parameters of the inducer are initially determined. Based on the range of core parameters, different combinations of core parameters are used to establish multiple initial three-dimensional models of the inducer. Simulations are performed on each of the established initial three-dimensional models to obtain the cavitation performance of the inducer. Based on the flow analysis results, the inducer with the optimal blade profile among all parameter combinations is obtained. The selected inducer with the optimal blade profile is then optimized and adjusted until the liquid rocket engine turbopump inducer with the best cavitation performance is obtained. This invention can design a liquid rocket engine turbopump inducer with good cavitation performance.

The application relates to a high-temperature-resistant and erosion-resistant heat insulation layer structure and a laying process method, and belongs to the field of solid rocket engine charge design. The application comprises a debonding layer, a first surface layer, a cylinder heat insulation layer, a ceramic-carbon cloth layer, a buffer layer, an aerogel filling layer, a bottom layer, a V-shaped sheet and a tail second surface layer. The application adds the ceramic-carbon cloth layer and the aerosol layer in the heat insulation layer, improves the strength of the heat insulation layer and reduces the thermal conductivity, so that the anti-erosion and high-temperature-resistant performance of the heat insulation structure is improved; the V-shaped sheet is pasted at the position close to the tail joint of the combustion shell of the debonding layer and the surface layer, so that the excessive materials in the charge process and the high-temperature and high-pressure gas and high-temperature condensed phase particles in the engine working process are effectively prevented from entering the debonding gap, and then the damage caused by the abnormal ablation of the debonding layer and the surface layer is avoided.

A propulsion-based payload transporter (PPT) and methods of its operation are provided. The PPT may operate in a gravitational field such as that on the lunar surface. Via rocket engines that produce thrust, the PPT is configured to descend onto and over a payload to be moved. Once in position with respect to the payload, latches engage with the payload so as to secure the payload in a payload holding bay of the PPT. The rocket engines produce thrust with an exhaust plume that is angled away from the payload to avoid payload damage. The rocket engines, which are located at the sides of the payload holding bay, produce thrust to propel the PPT and the payload in a substantially horizontal trajectory with respect to gravity and toward a landing destination.

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 discloses a nano-structure thermal protection coating for hydrogen fuel gas turbine and a preparation method thereof. The nano-structure thermal protection coating takes high-temperature alloy as a substrate, and sequentially deposits a bonding bottom layer, a yttrium stabilized zirconium oxide or zirconate intermediate layer and a silicate top layer on the surface of the substrate. The application starts from the design of the coating structure, and aims at the problems of low service temperature and poor water-oxygen corrosion resistance of the existing thermal protection coating for gas turbine. The water-oxygen corrosion resistance function of the silicate material is superimposed on the thermal protection coating for gas turbine, and the coating structure is optimized to form a new thermal protection coating structure resistant to high-temperature corrosion. The new thermal protection coating structure can not only meet more extreme service environments, but also resist water-oxygen corrosion at high temperature. The new thermal protection coating structure is suitable for hydrogen fuel power generation gas turbine, and is also suitable for hydrogen fuel or hydrogen mixed marine engines, hydrogen fuel rocket engines and other hydrogen fuel gas turbines.

The invention discloses a welding butt-joint device of a solid rocket engine cable cover support, and relates to the technical field of welding auxiliary equipment.The welding butt-joint device comprises a positioning base and an engine shell, positioning plates are arranged on the two sides of the positioning base, the welding butt-joint device further comprises the cable cover support, an opening is formed in the positioning base, and the cable cover support is arranged in the opening. A plurality of sets of installation parts are formed on the two sides of the inner wall of the open hole, butt joint cylinders are arranged at the bottom ends of the installation parts, positioning pressing plates are arranged on the symmetrical installation parts, the positioning pressing plates position the cable cover support, inserting rods are arranged on the two sides of the upper end face of each positioning pressing plate, locking mechanisms are further arranged, and the locking mechanisms which are fixed after being inserted are arranged between the butt joint cylinders and the inserting rods. The installation efficiency of the positioning pressing plate is improved, in this way, a plurality of cable cover supports can be rapidly and accurately fixed at a time, the cable cover supports are positioned through the positioning mechanism, and the welding stability of the cable cover supports is further guaranteed.

The utility model provides a solid rocket engine part press fitting device which comprises a press machine, a pressing transmission assembly and a base, the base is located under a pressing head of the press machine, and the base is provided with a positioning part for directionally fixing the lower end of an engine part to be press-fitted. The upper end of an engine part to be pressed is connected with a pressing head of the press machine through the pressing transmission assembly. According to the utility model, a product is vertically placed, the centering property of an engine part is good, the engine part can be quickly centered and assembled, and meanwhile, the self gravity of the engine part and the pressure of a press are utilized, so that the engine part can be conveniently and stably pressed axially and fixed radially, and the product is prevented from toppling over; the assembly efficiency and safety are high, and the problems that a conventional horizontal press-fitting component is poor in centering performance and unstable in pressing force are effectively solved.

The invention discloses a coupling suitable for a high rotating speed, and relates to the field of aerospace liquid rocket engine tests, the coupling comprises an input shaft sleeve, an output shaft sleeve and a flexible connection assembly, and the flexible connection assembly is connected between the input shaft sleeve and the output shaft sleeve; wherein the input shaft sleeve is used for being connected with an output shaft of the speed increaser and transmitting torque output by the speed increaser; the output shaft sleeve is used for being connected with an input shaft of a tester and transmitting torque output by the tester; the flexible connecting assembly has the elastic deformation capacity and is used for converting hard connection between the input shaft sleeve and the output shaft sleeve into flexible connection. The input shaft sleeve and the output shaft sleeve are connected through the flexible connecting assembly, and the requirements for axial and radial position precision and coaxiality during installation are effectively lowered; meanwhile, the bearing stress conditions of the output shaft of the speed increaser and the input shaft of the tester are greatly improved, and the risks of bearing damage and shaft breakage are remarkably reduced.

The invention relates to a five-engine parallel rack of a liquid rocket engine. The five-engine parallel rack comprises a main supporting structure and a servo supporting structure, the main supporting structure is used for bearing thrust of an engine and comprises a rack upper end face, a rack lower end face, two pairs of inclined supporting rods and a plurality of pairs of main force bearing rods. After the support is in place, when the rocket is horizontally placed, the extension set I, the extension set IV and the extension set V are located below the axis of the rocket, and the extension set II and the extension set III are located above the axis of the rocket; except the V extension set, each of the other extension sets corresponds to one servo supporting structure, the upper end of each servo supporting structure is fixed on the upper end face of the rack, and the lower end of each servo supporting structure is fixed on the lower end face of the rack; the servo mechanisms are installed on the servo supporting structures and used for driving the corresponding extension sets, and each servo supporting structure is used for bearing servo acting force of the corresponding servo mechanism. The five-engine parallel-connection thrust transmission requirement is met, and meanwhile the five-engine parallel-connection thrust transmission device adapts to the non-cyclic symmetrical layout form of five engines.

The invention relates to a plugging device for a liquid rocket engine ignition chamber flow resistance characteristic test. The plugging device comprises a tray, a non-metal expansion sealing ring, a metal extrusion structure and a locking nut. The non-metal expansion sealing ring is mounted in the product pipeline, and a boss is designed at the bottom; a bottom boss of the non-metal expansion sealing ring is mounted in an annular groove of the tray; the upper end of the metal extrusion structure is located in the product pipeline, and the lower end of the metal extrusion structure penetrates through an inner hole of the non-metal expansion sealing ring and a tray via hole and then is fixedly connected with the locking nut. The tightening torque is applied to the locking nut, the metal extrusion structure moves downwards, the non-metal expansion sealing ring is extruded, circumferential extrusion expansion of the non-metal expansion sealing ring is achieved, and plugging is achieved. According to the invention, the problem of plugging the light pipe attached with the supporting structure is solved, the sealing ring is fixedly arranged on the inner wall of the pipeline through extrusion expansion, the plugging of the flow channel is realized, and the flow resistance characteristic test of the specific flow channel is further realized.