376 results about "Supersonic speed" patented technology

The invention belongs to the technical field of gas turbine maintenance, and particularly relates to a solid-state repair method for a hot-end component of a gas turbine, which comprises the following steps: S1, acquiring damage information of the surface of a to-be-repaired matrix; s2, roughening treatment is conducted on a damaged area on the surface of the to-be-repaired base body, and electrochemical activation treatment is conducted on the roughened damaged area; s3, supersonic speed cold spraying treatment is conducted on the treated damaged area, and gradient materials are deposited on the damaged area layer by layer till a compact repairing layer is deposited on the surface of the base body; s4, selective laser remelting treatment is conducted on the compact repairing layer of the damaged area; s5, carrying out laser shock on the dense repair layer subjected to the remelting treatment; and S6, the base body subjected to laser shock treatment is machined till the standard geometric dimension of the base body is recovered. The method can solve the problems of large heat input, matrix damage, insufficient bonding strength and the like in the process of repairing the damaged area of the matrix in the traditional hot working technology.

The invention discloses a supersonic aircraft head thermal protection structure, and belongs to the technical field of hot end component cooling, the supersonic aircraft head thermal protection structure comprises a cooling system, a preheating system, a CBC power generation system and a secondary cooler, the cooling system comprises a coolant liquid storage tank, an electromagnetic adjustable pump and a cooling channel arranged in the wall surface of the aircraft head, a cooling medium flowing in a direction opposite to that of external high-temperature gas is arranged in the cooling channel; the preheating system comprises a fuel tank, a fuel adjustable pump and a fuel channel; the secondary cooler is used for transferring heat in the cooling channel into the fuel channel; the CBC power generation system comprises a heater, a primary cooler and a circulating working medium, the heater is used for transferring heat in the cooling channel to the circulating working medium, and the primary cooler is used for transferring heat in the circulating working medium into the fuel channel. Efficient thermal protection of the head of the supersonic aircraft can be achieved, the risk of structural failure caused by high temperature is reduced, and the safety of the aircraft is improved.

The invention discloses an optimization design method for a control surface of a wide-speed-range large-attack-angle reusable carrier, and relates to optimization design of aerodynamic configuration of an aircraft. The method comprises the following steps: S1, setting a control surface aerodynamic configuration design variable range, and generating an initial sample library by adopting Latin hypercube sampling; s2, aerodynamic parameters are obtained through CFD simulation, and a Kriging proxy model is trained; s3, evaluating the precision of the proxy model by taking a U learning function as a criterion, and stopping adding points when the minimum U function value is smaller than a threshold value; s4, constructing an optimization model which takes maximization of the lift-drag ratio and the static stability margin under the hypersonic speed as a target and takes the condition that the aerodynamic parameters of the supersonic speed / subsonic speed are not lower than a reference value and the hinge moment as constraints; and S5, performing iterative optimization by adopting an improved multi-target particle swarm algorithm based on genetic algorithm crossover mutation operation, updating the proxy model after the optimal solution of each generation is subjected to CFD verification, and outputting an optimal solution set. The problem of wide-speed-range aerodynamic configuration contradictions is solved, and the comprehensive performance of the carrier is remarkably improved.

The partial or full dehydration of organic or inorganic matter containing water by induction of matter into a vacuum and processing matter through a specially designed acceleration channel is disclosed. The inducted matter accelerates in air from zero speed to sub-sonic speed to reach supersonic speed. As the material transitions the sound barrier, it is subject to acoustic shock waves and an instant negative pressure drop occurs. The sound waves disintegrate, disinfect the material and extract part or all moisture from any organic or inorganic material.

The invention discloses a liquid metal-based radar antenna array surface non-uniform heat flux density heat dissipation structure, which comprises an antenna array surface, a radiation unit is mounted on the front surface of the antenna array surface, a liquid cooling plate is mounted on the back surface of the antenna array surface, and a T / R assembly on the antenna array surface is connected with the liquid cooling plate; a gradient porous structure formed based on a three-period minimal curved surface is welded at the bottom of the liquid cooling plate, and the liquid cooling plate is sealed with the liquid cooling shell through a sealing piece; the liquid cooling shell comprises a liquid inlet channel, a liquid outlet channel and a liquid cooling cavity, and liquid metal as cooling liquid exchanges heat with the liquid cooling plate through the liquid inlet channel and the liquid cooling cavity and then flows out through the liquid outlet channel. According to the invention, the liquid metal with high thermal conductivity is adopted as the cooling liquid and is combined with the gradient porous structure formed based on the three-period minimal curved surface, so that the problem of non-uniform heat flux density of the seeker detection radar in the flight process of a hypersonic aircraft is effectively solved, and compared with conventional gas-liquid two-phase flow heat dissipation, the stability is more excellent; and while heat exchange is enhanced, metal consumables can be saved, and the consumption of coolants can be reduced.

The invention provides low-temperature supersonic airflow collision ultramicro equipment, which belongs to the field of material crushing and processing, and comprises a feeding device, an airflow collision device, a closed-loop type circulation grading device, a material receiving device, a gas cooling device and a gas inlet device, and the gas inlet device is used for providing a high-pressure gas source for the equipment; the gas cooling device cools the high-pressure gas to a preset low temperature, and then the high-pressure gas is fed into the gas flow collision device after being subjected to pressure regulation through the gas feeding pipeline; the airflow collision device comprises a collision chamber and at least one pair of opposite nozzles, the horizontal distance between the paired nozzles can be adjusted, the air supply pipeline and the feeding pipeline are connected with the nozzles after being intersected, and high-pressure air is sprayed out through the nozzles to form supersonic jet flow to drive materials to collide with each other in the collision chamber. And the collided materials are subjected to closed-loop circulation classification and then enter a material receiving device, and large particles are subjected to circulation collision crushing again. The ultramicro equipment disclosed by the invention can be used for crushing various plant raw materials into 1-10 microns, and has a good application prospect.

The invention relates to the technical field of bag type dust removal, in particular to a supersonic speed injection pulse injection pressure control method and system, which are used for evaluating the effectiveness of single ash removal operation and optimizing the subsequent injection pressure, and the method comprises the following steps: when at least one injection unit of a bag type dust removal system executes injection operation, the injection pressure of at least one injection unit of the bag type dust removal system is controlled; a pressure monitoring device arranged on a blowing gas path of the blowing unit is used for collecting a pressure signal in the blowing gas path to obtain pressure waveform data, and the pressure waveform data represents the pressure change at the moment of blowing; extracting at least one preset dynamic characteristic parameter from the pressure waveform data, wherein the dynamic characteristic parameter reflects the blowing airflow state or the air permeability change of the filter bag; the problems that in the prior art, dust removal control is extensive, and fine evaluation and adjustment cannot be achieved are solved by collecting a pressure signal of a blowing gas path, extracting dynamic characteristic parameters, evaluating the effectiveness grade of dust removal, adjusting the blowing pressure accordingly and updating the mapping relation online at the same time.

The invention relates to the technical field of metal wire processing, and discloses a high-performance metal wire plastic coating process for aerospace engineering, which comprises the following steps: carrying out base material type selection on a metal wire, and carrying out surface treatment of precise cleaning, etching and depth gradient activation; a Ti-Cr alloy and graphene-copper composite transition layer is deposited through a magnetron sputtering technology, and a multi-layer coating structure is formed through electrostatic spraying of a base coating, a fluidized bed dipping function layer and high velocity oxy-fuel spraying of a nano composite coating in sequence; the molybdenum-gold composite sealing layer is prepared by adopting an electric arc spraying technology, and is subjected to vacuum ultraviolet curing enhancement, high-temperature gradient annealing, gradient sintering and vacuum curing treatment; and heat radiation on the surface of the coating is enhanced through laser-induced micro-nano structure regulation and silicon nitride film sputtering. Through collaborative optimization of base material treatment, coating construction, function strengthening and quality control, the interface bonding strength, the comprehensive performance and the environmental adaptability of the spaceflight metal wire coating are remarkably improved, and the service life of the spaceflight metal wire coating is remarkably prolonged.

The invention discloses an aerodynamic configuration structure of a precision guidance assembly and a design method, and belongs to the technical field of precision guidance, an assembly precursor adopts continuous three-section curve design: a supersonic speed section adopts a parabola profile, is provided with ribs on the surface, and is used for optimizing shock wave control under a supersonic speed working condition; the transonic speed section adopts a hyper-ellipsoid curved surface, ribs are arranged on the surface of the transonic speed section, and metamaterials are attached to the ribs and used for restraining transonic speed separation; the subsonic speed section adopts arc transition, and ribs are arranged on the surface and used for improving the lift force efficiency of the control surface; wherein the curvatures at the joints of the three sections of curves are continuous. According to the method, supersonic velocity-transonic velocity-subsonic velocity three-section collaborative design is realized, and meanwhile, a boundary layer is improved from a microscopic level by combining a collaborative drag reduction design of a microscopic rib and metamaterial technology, and the performance of a guidance assembly is optimized and improved.

The invention belongs to the technical field of turbine engine design, and particularly relates to a turbine interstage combustion chamber supersonic cruise engine and a control method thereof.The turbine interstage combustion chamber supersonic cruise engine comprises a fan adopting a FLADE fan; the gas compressor is arranged behind the fan; the main combustion chamber is arranged behind the gas compressor; the high-pressure turbine is arranged behind the main combustion chamber and is connected with the gas compressor through a high-pressure shaft; the mixing chamber is arranged behind the high-pressure turbine and is communicated with the fan through an outer duct; the interstage combustion chamber is arranged behind the mixing chamber; the low-pressure turbine is arranged behind the combustion chamber and is connected with the fan through a low-pressure shaft; the spraying pipe is arranged on the low-pressure turbine; and after the injection valve is arranged on the high-pressure turbine, the injection valve deflects to the inner runner surface of the outer duct when the injection valve is opened, and deflects to the outer runner surface of the outer duct when the injection valve is closed.

An airfoil for an aircraft having an airfoil geometry. The airfoil geometry includes an upper surface and a lower surface, the upper and lower surfaces are in accordance with Cartesian coordinate values of X and Y. The Cartesian coordinate values of X and Y are non-dimensional values from 0 to 1 convertible to dimensional values by multiplying the Cartesian coordinate values of X and Y by a chord length of the airfoil. The upper surface and the lower surface form an enclosed airfoil shape. The aircraft is a supersonic aircraft, and the airfoil has a thickness-to-chord ratio of about 10%.

The invention relates to the technical field of rocket engine simulation tests, in particular to a vertical high-altitude simulation test bed compatible with a ground test and a test method. Comprising a vacuum chamber which is provided with a ventilation assembly capable of being opened and closed; and the diffuser is communicated with the vacuum cabin and is vertically arranged at the tail end of the engine, and engine gas enters the diffuser. When a high-altitude simulation test is carried out, the long-nozzle engine is arranged in the vacuum cabin, and the openable and closable ventilation assembly is closed, so that the inner side of the vacuum cabin is isolated from the outside. Engine fuel gas enters the diffuser and is expanded in the diffuser to form supersonic airflow to suck the vacuum chamber, the fuel gas is injected to the injection expansion section under injection of the injection working medium, the fuel gas is mixed with the injection working medium airflow to continue to be pressurized in the supersonic speed state, and finally the fuel gas is exhausted from the injection expansion section when the pressure is higher than the atmospheric pressure. The air pressure in the vacuum chamber is lower than the atmospheric pressure, and high-altitude simulation test conditions are met.

The invention provides a wing-body fusion wide-speed-range aircraft layout design method based on consideration of an afterbody effect. The method comprises the following steps: firstly, generating a cone-guided waverider precursor of a wide-speed-range aircraft for supersonic flight; then, based on the basic airfoil profile, performing pneumatic optimization on the basic wing profile and the strake wing profile of the wide-speed-domain aircraft to obtain a multi-section wide-speed-domain waverider airfoil profile; then calculating the length of a waverider afterbody and designing a symmetry plane of the wide-speed-range aircraft; and finally, based on an osculating cone method and transonic speed strake wing constraints, obtaining three-dimensional layout parameters by using the waverider precursor parameters and the waverider afterbody length, and assembling the basic wing optimized airfoil profile, the strake wing optimized airfoil profile and the symmetry plane to generate the layout of the wide-speed-domain aircraft. According to the method, the cone-guided waverider generation method, the wide-speed-range airfoil optimization method and the osculating cone double-sweepback waverider generation method are coupled, and the aerodynamic design of the afterbody is considered, so that the aerodynamic performance of the wide-speed-range aircraft under the subsonic / transonic / supersonic speed is improved.

The invention belongs to the field of target drone fuel oil systems, and particularly relates to a self-adaptive supersonic speed target drone fuel oil system and a control method thereof. The system comprises a machine body oil tank, a first electromagnetic valve, an overflow valve, a pressurization oil tank, a fuel pump, an engine pump, a monitoring module and a control module. The body fuel tank comprises a main fuel tank and an auxiliary fuel tank which are arranged in a fuel tank in the middle of the target drone body and located in front of and behind the theoretical mass center of the target drone respectively. The pressurization oil tank is additionally arranged in the fuel oil cabin, fuel oil in the engine body oil tank enters the pressurization oil tank for pressure maintaining after being pressurized by the fuel oil pump, and then oil is supplied to the engine from the pressurization oil tank, so that stable oil supply to the engine is achieved by establishing a local pressure environment instead of a large pressure environment of the whole fuel oil cabin, and therefore when a fuel oil system works, the fuel oil can be stably supplied to the engine. The oil supply flow and pressure are controlled in a closed-loop mode in real time through the control system according to the engine state.

The invention relates to the technical field of metallurgical furnace charge conveying and control, and discloses a top-blown furnace multi-channel supersonic jet type feeding control method and system.The method comprises the steps that flow distribution in a nozzle is achieved through a piezoelectric ceramic high-speed valve installed at a pipeline inlet, and a nozzle servo motor is controlled at the top of a spray gun to adjust the spray angle of the spray gun; the feeding speed, pressure and temperature are monitored in real time through a pressure sensor and a flow sensor arranged in an inlet area of the pipeline and an infrared thermometer mounted at the top of the furnace; the carrier gas flow, the nozzle opening degree and the spraying angle are dynamically adjusted through control of the PLC executing mechanism, and collaborative optimization of the material flow vector, the flow field in the hearth and the heat and mass transfer process is achieved. The system comprises a spray gun, a piezoelectric ceramic high-speed valve, a flow sensor, an infrared thermometer, a nozzle control servo motor, a temperature measurement control small servo motor and a control room. The comprehensive operation efficiency of energy density distribution optimization, reaction efficiency improvement and energy consumption index reduction of the molten pool reactor is finally realized.

The invention relates to the technical field of hydrogen permeation membranes, in particular to a porous metal tube loaded high-entropy alloy hydrogen permeation membrane as well as a preparation method and application thereof. The preparation method comprises the steps that under the inert atmosphere, molybdenum and copper are subjected to mixed ball milling, then cobalt, chromium, iron and nickel are added for mixed ball milling, molybdenum and copper doped cobalt, chromium, iron and nickel high-entropy alloy powder is obtained, then the molybdenum and copper doped cobalt, chromium, iron and nickel high-entropy alloy powder is sprayed to the surface of a porous metal pipe through a supersonic speed hot spraying film making method, and the porous metal pipe is obtained. And a film layer is formed, and the porous metal tube loaded high-entropy alloy hydrogen permeation film is prepared. The high-entropy alloy hydrogen permeation membrane is high in corrosion resistance and suitable for extracting hydrogen from coke oven gas containing much sulfur / carbon impurity gas, and the problems that a conventional compact metal hydrogen permeation membrane is poor in sulfur / carbon resistance and thermal stability are solved.

A device and related method that provides, but is not limited thereto, a two-phase heat transfer device with unique combination of enhanced evaporation and increased cooling capacity. An advantage associated with the device and method includes, but is not limited thereto, increased cooling capacity per unit area, controlled and optimized evaporation, prevention of boiling, and prevention of drying of the evaporator. An aspect associated with an approach may include, but is not limited thereto, using anon-wetting coating or structure to keep working fluid away from the spaces between elongated members of an evaporator and using a wetting coating or structure to form thin films of working fluid around the distal region of the elongated members. For example it can be used to cool a computer chip, a skin of a hypersonic flying object, parabolic solar collector, turbine or engine blade, or any other heat source that requires high heat flux.

The invention relates to the technical field of thermal barrier coatings for gas turbines or aero-engines, and particularly discloses an ultra-thick and strong-bonding double-ceramic-layer thermal barrier coating and a preparation method thereof.The thermal barrier coating is composed of an MCrAlY bonding layer, a YSZ ceramic layer bottom layer and a high-entropy rare earth zirconate ceramic layer surface layer; the MCrAlY bonding layer is prepared through atmospheric plasma or hypersonic flame spraying, YSZ and high-entropy rare earth zirconate spherical powder serve as spraying raw materials, and the thermal barrier coating is prepared through an axial powder feeding high-energy plasma spraying system. The thermal barrier coating has the characteristics of super thickness (the coating thickness is 600-1200 microns), high bonding strength (gt; 19 MPa) and the like.

The invention discloses a preparation method for spraying a nickel-chromium-chromium carbide coating on the surface of a gas turbine part through high velocity oxy-fuel, and belongs to the technical field of preparation of the nickel-chromium-chromium carbide coating through thermal spraying. The preparation method comprises the following steps that preparation before spraying is conducted, and nickel-chromium-chromium carbide powder for thermal spraying is placed in a drying oven to be dried; cleaning is conducted, specifically, the spraying area and the adjacent area of the part are wiped and cleaned, and impurities such as oil stains are removed; protection before sand blasting is conducted, specifically, the non-sand-blasting surface and a small part of the spraying area of the part are protected through a high-temperature-resistant adhesive tape; a sand blowing machine is provided with a 0-0.6 MPa adjustable air pressure air source and is provided with an oil-water separator; spraying is conducted, specifically, a nickel-chromium-chromium carbide coating is sprayed through hypersonic flame; and cleaning is conducted, specifically, after spraying is completed, the spraying protection clamp and the adhesive tape are removed, and floating ash on the surface of the part is cleaned. The wear-resistant coating is prepared on the surface of the aero-engine and gas turbine part by adopting the process, a nickel-aluminum bottom layer does not need to be sprayed in advance, the quality is stable, and the production cost is reduced.

The invention discloses an anti-oxidation and anti-scouring heat prevention and insulation composite material, a preparation method and application, and belongs to the technical field of high-temperature thermal protection materials. According to the method, a two-step in-situ modification method is adopted, hydrolyzed organosilane reacts with phenolic resin to prepare silicon modified phenolic resin, then the silicon modified phenolic resin is bonded with an organoboron compound to obtain boron-silicon co-modified phenolic resin, a curing agent is added, then the boron-silicon co-modified phenolic resin is injected into a fiber reinforcement mold, and a finished product is obtained through a sol-gel reaction and normal-pressure drying. According to the invention, boron and silicon elements are embedded into phenolic aldehyde molecular chains through covalent bonds to form a boron-silicon-carbon hybrid network, a nano ceramic phase and a compact glassy protective layer are generated at high temperature, and the coating has excellent anti-scouring property, thermal stability and heat insulation property. The preparation process is simple and controllable, is suitable for large-scale production, can accurately regulate and control the performance, and is suitable for thermal protection parts of supersonic aircrafts.

The invention relates to the technical field of natural gas extraction, in particular to a supersonic shock wave hole atomization drainage gas production device, and aims at solving the problems that in the prior art, a throttler is poor in atomization efficiency and low in drainage gas production efficiency. The device comprises a connector and further comprises an atomization device. The atomization device comprises a central pipe and a nozzle; two ends of the central pipe are communicated to form a diversion channel; the flow guide channel is provided with an inlet end and an outlet end; the connector sleeves the inlet end; the nozzle is fixedly connected with the outlet end through the nozzle connector; a groove is formed in the inner wall, close to the inlet end, of the flow guide channel, when airflow flows into the groove, shock wave turbulent flow can be generated, so that a gas-liquid mixture in natural gas can be preliminarily sheared and broken in the groove and then is sprayed out of the nozzle, the supersonic shock wave hole atomization drainage gas production device can have higher atomization efficiency, and the atomization effect is better. And the natural gas drainage gas recovery efficiency is further improved.

The invention discloses an unmanned aerial vehicle supersonic airflow anti-recoil fire extinguishing device which comprises a fire extinguishing water spray pipe, an anti-recoil gas spray pipe, a high-low holder, a direction turntable, a camera, a direction control mechanism, a high-low control mechanism, a fire extinguishing water generation device, a gas generation device, a controller and an unmanned aerial vehicle. The high-low holder is installed on the direction rotary table and can rotate in the horizontal direction under the control of the direction control mechanism, the unmanned aerial vehicle is fixedly installed on the direction rotary table, the camera is fixedly installed on the high-low holder, and the fire extinguishing water spraying pipe is installed on the high-low holder and can rotate in the high-low direction under the control of the high-low control mechanism. The reverse recoil gas spraying pipe and the fire-extinguishing water spraying pipe are opposite in direction and are fixed together in an axis coincident mode, the fire-extinguishing water generating device comprises a fire-extinguishing water storage pond, a fire-extinguishing water pump and a grounding water pipe, the fire-extinguishing water storage pond is communicated with the fire-extinguishing water pump, and the fire-extinguishing water pump is communicated with the grounding water pipe; the gas generating device comprises a gas storage chamber, a gas compressor and a grounding gas pipe, the gas storage chamber is communicated with the gas compressor, and the gas compressor is communicated with the grounding gas pipe.

The utility model belongs to a spraying device, and provides a supersonic spraying device, a non-invasive spraying device and a spraying instrument.The problems that spraying is not uniform and operation is inconvenient still exist in an existing instrument, and the supersonic spraying device, the non-invasive spraying device and the spraying instrument.The supersonic spraying device comprises a liquid section, a gas inlet section, a contraction section, a throat part and an expansion section, and the gas inlet section, the contraction section, the throat part and the expansion section are coaxially and sequentially connected; liquid enters from the liquid section, gas is accelerated through the contraction section, reaches the throat part and is continuously accelerated at the expansion section, the liquid entering from the liquid section is atomized under the action of the gas at the throat part, is brought into the expansion section and is finally sprayed out from the spraying opening along with continuous acceleration, and the uniform spraying effect is achieved. According to the supersonic jet device, the ratio of the axial length L1 to the axial length L2 is designed, the structure of the supersonic jet device is further optimized, and actual verification shows that under the optimized ratio, the jet force of the supersonic jet device is obviously improved, the particle size and the jet range of jetted liquid drops are better, and the flow speed of the liquid drops is more stable.

A device and related method that provides, but is not limited thereto, a two-phase heat transfer device with unique combination of enhanced evaporation and increased cooling capacity. An advantage associated with the device and method includes, but is not limited thereto, increased cooling capacity per unit area, controlled and optimized evaporation, prevention of boiling, and prevention of drying of the evaporator. An aspect associated with an approach may include, but is not limited thereto, using a non-wetting coating or structure to keep working fluid away from the spaces between elongated members of an evaporator and using a wetting coating or structure to form thin films of working fluid around the distal region of the elongated members. For example it can be used to cool a computer chip, a skin of a hypersonic flying object, parabolic solar collector, turbine or engine blade, or any other heat source that requires high heat flux.

The invention discloses a large-intake-angle supersonic ultrathin cascade flow loss value extraction method based on infinitesimal flow averaging, and belongs to the field of fluid dynamics. According to the method, on the basis of ultrasonic probe vibration data obtained in the real testing process, the maximum displacement of the probe in the testing direction and the incoming flow direction is determined in combination with testing position information, and therefore the infinitesimal size in flow loss analysis is further defined. On the basis, a numerical model for simulating the flow state of the working medium in the cascade is constructed according to the physical structure of the test bed, and is simplified into a single-channel model. In order to obtain an accurate simulation result, a plurality of turbulence models are selected, and the optimal turbulence model is selected by means of a plurality of evaluation indexes for subsequent analysis. Setting infinitesimal elements in the simulated flow field, wherein the size and the distribution of initial positions of the infinitesimal elements are defined according to position data in an actual test. According to the method, the flow loss characteristic of the supersonic cascade can be effectively evaluated, and the cascade design efficiency and reliability are improved.

The invention provides a continuous supersonic speed airflow jet type soot blower which comprises a mixer, an ignition assembly, a shock wave tank and a jet assembly, the mixer is provided with an air inlet and a fuel gas inlet, the air inlet is used for introducing air into the mixer, the fuel gas inlet is used for introducing combustible gas into the mixer, and the jet assembly is used for jetting the fuel gas into the mixer. The ignition assembly comprises an ignition chamber and an igniter, the ignition chamber is connected with the mixer, the igniter is arranged in the ignition chamber and used for igniting mixed gas in the ignition chamber, the shock wave tank is connected with the ignition chamber and used for storing gas exhausted by the ignition chamber, the injection assembly comprises an injection pipe, one end of the injection pipe is connected with the shock wave tank, and the other end of the injection pipe is connected with the shock wave tank. And the other end of the injection pipe is arranged in the air pre-heater. The continuous supersonic speed airflow jet type soot blower has the advantages of being low in energy consumption and good in cleaning effect.

The invention discloses a thermal barrier coating with a double-bonding-layer structure and a preparation method thereof, and belongs to the field of thermal barrier coatings. According to the thermal barrier coating, nickel-based high-temperature alloy serves as a substrate, the thermal barrier coating sequentially comprises an MCrAlY lower bonding layer, a Cr2AlC upper bonding layer and an 8YSZ ceramic layer, the MCrAlY lower bonding layer and the Cr2AlC upper bonding layer are prepared through the high velocity oxy-fuel spraying technology, and the 8YSZ ceramic layer is prepared through the atmosphere plasma spraying technology. According to the thermal barrier coating system, through the synergistic effect of the oxidation resistance of the Cr2AlC upper bonding layer and the mutual diffusion inhibition characteristic of the MCrAlY lower bonding layer, the comprehensive performance of the thermal barrier coating is optimized, meanwhile, the anti-stripping capacity of the thermal barrier coating system is remarkably improved, and the thermal barrier coating system is suitable for various high-temperature alloy substrates, is not limited by alloy component differences, can further adapt to workpieces in complex geometrical shapes and is suitable for industrial production. The method can be used for protecting hot-end components such as aero-engine turbine blades and the like.

The invention provides a low-slag efficient smelting method and system based on immersed supersonic gas-solid injection, and relates to the field of metallurgy. The method comprises a raw material adding stage, a melting stage, an oxidation heating stage and a tapping stage. The low-slag efficient smelting system based on immersed supersonic speed gas-solid injection comprises an electric arc furnace, a powder injection tank, an immersed supersonic speed gas-solid oxygen lance and a control device, the electric arc furnace is used for smelting a target steel grade, the powder spraying tank is used for conveying a slagging material to the immersed supersonic gas-solid oxygen lance, and the immersed supersonic gas-solid oxygen lance is used for conveying the slagging material, methane, oxygen and shielding gas into a steel slag layer of the electric arc furnace; and the control system is used for controlling the immersed supersonic gas-solid oxygen lance. According to the low-slag efficient smelting method and system, the dephosphorization efficiency is improved by about 2.5-5%, the slag amount is reduced by 5-20 kg / t, the consumption of slag making materials is reduced by 5-10%, and the metal yield is improved by 0.5% or above.

The invention provides a high-load supersonic cascade and a gas compressor, and belongs to the technical field of gas turbine gas compressors. The blade grid comprises blades and vortex generators, the vortex generators are installed in the front edge areas of the blades, and the vortex generators and the suction surfaces of the blades are located on the surfaces of the same sides of the blades. In the airflow direction corresponding to the blades, the height of the vortex generator is gradually reduced, the width of the vortex generator is gradually increased, and the width of the vortex generator is the size of the vortex generator in the spreading direction of the blades. The vortex generators in specific shapes are arranged on the surfaces of the blades, so that after airflow passes through the vortex generators, two reverse vortex pairs spaced by a certain distance can be generated, and the two reverse vortex pairs can entrain the airflow from a mainstream area with more gas and supplement the airflow to a shock wave / boundary layer interference separation area lacking gas; therefore, flow separation of the rear half section of the suction surface is inhibited, flow loss is reduced, and aerodynamic performance of the cascade is improved.

The invention relates to the technical field of aircraft design, in particular to a sectional type hybrid power supersonic commercial passenger plane which comprises a supersonic main aircraft and a subsonic babysitter machine. In the take-off and landing stages, the subsonic babysitter carries the supersonic main aircraft to form a combined body, an engine of the supersonic main aircraft supplies power to a ducted fan of the subsonic babysitter and charges an energy storage device of the ducted fan, and the main aircraft and the subsonic babysitter cooperate to generate thrust to jointly fly; in the cruising stage, the supersonic main aircraft is separated from the subsonic babysitter, the supersonic main aircraft flies independently, and the subsonic babysitter returns. Through the innovative sectional type structural design, combination and configuration of the propelling system and power distribution and switching integration in the flight stage, efficient, safe and environment-friendly flight of the supersonic commercial airliner is achieved.