30results about How to "Reduce thrust loss" patented technology

The invention relates to a pneumatic shock response spectrum tester, which comprises a base, a pneumatic hammerhead device, a resonant panel supporting device and a functional mechanism. The pneumatic hammerhead device comprises a hammerhead, guide rails and an internal cylinder, wherein one end of the internal cylinder is in sliding sealing with an open hole at one end of the hammerhead; grooves on the two sides of the hammerhead are clamped in the two symmetrically arranged guide rails; the guide rails are fixed on the base axially; a screw shaft at an end cover of the internal cylinder is connected with a joint bearing which is connected with a lug of an internal cylinder connecting block by a pin; and the bottom surface of the internal cylinder connecting block is fixed on the base by a thrust cylinder connecting seat. The resonant panel supporting device comprises a resonant panel, limiting clamping blocks, round rolling bars and a cushioning seat, wherein V-shaped grooves on the two sides of the resonant panel are clamped in a plurality of symmetrically arranged limiting clamping blocks; the resonant panel is supported by the round rolling bars fixed on a round bar mounting seat; and the limiting clamping blocks and the round bar mounting seat are fixed on the base. The pneumatic shock response spectrum tester is simply and compactly structured, simple and convenient to operate, short in stroke, high in excitation energy, convenient to arrange and maintain, safe and environment-friendly.

The invention relates to a nozzle insulation structure for thrust steering jet engines. The structure is suitable for field of medium and small thrust aero-engines having thrust steering demands. Thestructure consists of 1 # heat shield (9), 2 # heat shield (8), 3 # heat shield (7), 4 # heat shield (6), elastic connecting sleeve members (13) and air entraining holes (2). The heat shields forms anoverall structure member; a friction material is a high temperature alloy, and the friction material is simple in structure, light in weight and good in maintainability, and is capable of repeated assembly and disassembly; gaps are designed between the heat shields so that no additional load can be achieved during a nozzle steering process; a heat bearing part is separated with a bearing part, and therefore, high temperature resistance demands can be met, long service life can be achieved, and reliability of the structure can be enhanced; and through the controlling of the designed gaps, cooling air entraining amount can be decreased, and engine thrust losses can be reduced.

The invention belongs to the technical field of aircraft air intake ducts and discloses a combined subsonic air intake device of an aircraft. The device comprises an embedded air intake duct, an externally-hung air intake structure and an automatically-separated connection mechanism. The embedded air intake duct comprises a fixed air intake passage and a fixed flange plate. The externally-hung air intake structure comprises a jettisonable air intake passage, a front fairing, a rear fairing and a jettisonable flange plate. The automatically-separated connection mechanism comprises an explosive bolt and a connecting sheet. The jettisonable air intake passage, the front fairing and the rear fairing are fixed to the jettisonable flange plate through connecting pieces, the jettisonable flange plate is fixed to an inlet of the embedded air intake duct on the aircraft through the explosive bolt and the connecting sheet, and an airflow outlet of the externally-hung air intake structure is matched with an airflow inlet of the embedded air intake duct. According to the device, flow distortion is good, engine thrust loss is reduced, and the flying ability of the aircraft is guaranteed; the externally-hung air intake structure is separated from the aircraft at the tail section of the aircraft, and therefore the stealth defense penetration ability of the aircraft is guaranteed.

The invention discloses an aero-engine rectification cover hood anti-icing device adopting a loop type heat pipe and belongs to the field of protection of aero-engines. The aero-engine rectification cover hood anti-icing device comprises a loop type axial rotary heat pipe, wherein the axial rotary heat pipe sequentially comprises a condensation section, a heat insulation section and an evaporation section from front to back; the front end of the condensation section is fitted to an aero-engine rectification cover hood; the inner diameter of the front part of the condensation section is greater than that of the back part of the condensation section; the inner diameter from the evaporation section to the heat insulation section is gradually reduced; the inner diameter of the heat insulation section is greater than the maximum inner diameter of the condensation section; a liquid reflowing channel is communicated between the condensation section and the heat insulation section; an inlet of the liquid reflowing channel is located at the condensation section and an outlet of the liquid reflowing channel is located at the heat insulation section; a heating device is mounted on the evaporation section. The loop type axial rotary heat pipe is integrated on the inner surface of the aero-engine rectification cover hood so that generated heat is efficiently and stably transmitted to the surface of the aero-engine rectification cover hood, and anti-icing of the aero-engine rectification cover hood is realized.

The invention relates to a propeller, specifically to a combined pump-jet propeller. The combined pump-jet propeller is characterized by comprising an external guide pipe; the external guide pipe is of a contracted shape; two ends of the external guide pipe is respectively a water inlet and a water outlet; a pump-jet propeller and a water jet pump are arranged within the combined pump-jet propeller; the pump-jet propeller comprises a pump-jet rotor, a pump-jet stator and a pump-jet motor; the pump-jet motor is fixed within the external guide pipe through four fixed rotors; the rear part of thepump-jet motor is connected with the pump-jet rotor; the pump-jet stator is arranged behind the pump-jet rotor to be used for rectification; the pump-jet stator is connected with the external guide pipe; a pump-jet impeller motor is arranged at the center of the pump-jet stator; and the pump-jet impeller motor is connected with the water jet pump. The combined pump-jet propeller has the advantages that defects of the pump-jet propeller and the water jet pump complement each other on the premise that the advantages of both the pump-jet propeller and the water jet pump are kept; the combined propeller has the advantages of low radiation noise, less thrust loss, higher propelling efficiency and high critical speed; and the combined pump-jet propeller is suitable for submarines and vessels.

The invention discloses a protection device for natural gas pipeline laying. The protection device for natural gas pipeline laying structurally comprises protection device bodies, a first outer box, screws, a supporting bottom plate, a second outer box, auxiliary plates, a cover plate and rotary knobs. According to the protection device for natural gas pipeline laying, a natural gas pipeline is put on semicircular pipeline plates, under the action of pipeline plate auxiliary springs, the semicircular pipeline plates are stressed to press the natural gas pipeline, and when the height of the pipeline needs to be increased, a pedal is stamped with foot (the pedal pushes an oil cavity extrusion cylinder, at that time, a rubber ball of an upper end oil conveying structure is opened, oil is conveyed to a hydraulic cavity cylinder, and a lifting rod drives the second outer box to ascend); and when the height of the pipeline needs to be reduced, a pressure relieving valve is rotated, the oil in the hydraulic cavity cylinder reflows into a hydraulic oil tank through an oil pipe, and the lifting rod drives the second outer box to descend. The protection device is small in size, the pipelinecan be clamped and protected only by putting the natural gas pipeline into the semicircular pipeline plates, the height can be adjusted in a self-service manner after the pipeline is installed, and installation of the natural gas pipeline is facilitated.

The invention discloses a double-S-bend convergent-divergent spray pipe with an infrared suppression measure, and belongs to the field of aero-engines. The double-S-bend convergent-divergent spray pipe comprises an engine outer duct, an engine mixer, an engine inner duct, a cooling channel, a spray pipe contraction section, a spray pipe expansion section and a spray pipe throat part, wherein the cooling channel is located on the periphery of a leeward area of the S-bend spray pipe and is composed of a cooling channel wall covering the leeward area of the S-bend spray pipe and an S-bend spray pipe wall; an opening is formed in the starting end of the cooling channel, and the other sides are connected with the S-bend spray pipe wall through the edge of the cooling channel wall to form a closed structure and communicate with the interior of the S-bend spray pipe through bypass air inlet holes and cooling outlets; and the plurality of bypass air inlet holes are evenly distributed in the starting end of the cooling channel in the circumferential direction, and the cooling outlets are located in the tail end of the cooling channel and communicate with the S-bend spray pipe through a plurality of cooling air outlet holes. The temperature of the wall surface of the spray pipe is reduced by airflow in the cooling channel, so that the infrared radiation characteristic of the spray pipe is reduced; and flow separation of the leeward area of the double-S-bend spray pipe is improved through airflow exhausted by cooling holes, so that the thrust of the spray pipe is improved.

The invention relates to the technical field of rocket engines, particularly to a flexible extension effuser with self-adaption height, which includes a basic bell-shaped effuser part, a multi-stage bell-shaped extension part, a variable constrained point part and a support structure part, wherein part of the basic bell-shaped effuser part is a fixed effuser, one end of the fixed effuser is connected with a rocket engine, and one end is connected with the multi-grade bell-shaped extension part; the multi-grade bell-shaped extension part is a flexible extension effuser, is in a rolling state when in the launching initial state, and is unfolded along with the increase of flying height; the variable constrained point part is positioned on the flexible extension effuser, and is fixed to a setposition of the support structure part; the support structure part is positioned inside the flexible extension effuser, and is unfolded along with the increase of flying height. According to the flexible extension effuser, the current problem of low working efficiency of rockets with wide operation range, such as single-stage orbit rockets is solved, the air flow is kept fully expanded all the time, so that the rocket is in the optimal operation state at any height.

The present invention discloses one kind of retch equipment for water-jetting propelled ship. The retch equipment includes one boost pump jet with front edge connected to the boost pump and one rudder shell fitting to the boost pump jet through one connecting piece. The front edge of the rudder shell is fitted to the back edge of the boost pump jet. The present invention can guide the jetted water flow from the boost pump jet, and the water flow is turned inside the rudder shell to reach the ship retching aim. The present invention has the advantages of homogeneous distribution of stress, great thrust efficiency, minimized loss, etc.

The invention discloses a high-thrust double three-phase permanent magnet linear synchronous motor for a non-rope elevator hoisting system. The high-thrust double three-phase permanent magnet linear synchronous motor comprises a short secondary and a long primary; the long primary adopts a segmented structure, and each primary segment comprises an armature core and two armature windings arranged on armature teeth in a sleeving mode, wherein the two armature windings have a certain electrical angle difference in space, are mutually independent in a circuit and are sequentially switched during segment switching, and the switching point is the middle point of the primary segment; the long primary is installed on the wall of an elevator shaft, and a certain distance exists between every two adjacent segments; and the short secondary comprises back irons and magnetic poles, the magnetic poles adopt a Halbach structure and are in rigid connection with a car through the back ions, the short secondary adopts a segmented structure, and the segments are continuously installed. According to the motor, not only is the advantage of the good error tolerance performance of a multi-phase motor combined, but also a ubiquitous three-phase inverter power supply can be directly used for supplying power, and the application occasions and the development prospect f the multi-phase motor are widened.

The invention discloses a fire damper for a secondary detonation-breathing pulse detonation engine. A body of the fire damper is circular. The front end of the body is planar, and the rear end of the body is spherical. A plurality of circles of same conical holes are arranged on the circle of the body. The conical holes are equal in peripheral spacing. The outer diameter of the body is equal to the inner diameter of a straight barrel of an ejector. The inner diameter of the body is equal to the outer diameter of a pre-detonation tube. The body is mounted at one end of the pre-detonation tube, located at a lip of the ejector. An air channel is formed between the body and the lip of the ejector. Air flowing in from the front end of the fire damper passes maximum sectional areas of the conical holes and decreases along with the sectional areas of the conical holes. Air flowing in from the rear end of the fire damper passes minimum sectional areas of the conical holes first and increases along with the sectional areas of the conical holes, and forward flow resistance is smaller than reverse flow resistance. Forward transmission of high-temperature combustion products in the pulse detonation engine is improved effectively, and operational reliability of the pulse detonation engine is improved.

The invention provides a pump-jet propeller. The pump-jet propeller includes a conduit, stators, a rotor, a wheel hub, a rotating shaft and a fluid director, wherein the wheel hub, the rotor, the stators and part of the rotating shaft are located in the conduit; the wheel hub and rotor are successively sleeves the rotating shaft, and the rotating shaft can drive the rotor to rotate movably; the stators are evenly distributed between the wheel hub and the conduit at intervals, one ends of the stators are connected with the wheel hub, and the other ends of the stators are connected with the conduit; and the fluid director is located on the side, away from the wheel hub, of the rotor, and the fluid director is connected with one end of the rotating shaft and is coaxial with the rotor, the wheel hub and the rotating shaft. According to the pump-jet propeller, on the basis of retaining the original advantages, the fluid director is added to interact with trailing vortexes generated by the rotor and stators, the thrust loss is decreased, and greater thrust are achieved.

The invention discloses a high-pressure extrusion dehydrator and a using method thereof, and relates to the field of garbage treatment. The high-pressure extrusion dehydrator comprises an extrusion equipment frame, a feeding push head and an extrusion push head; a feeding part and an extrusion part are arranged on the two sides of the extrusion equipment frame; the feeding push head is arranged on the feeding part and the extrusion push head is arranged on the extrusion part; the feeding push head and the extrusion push head are both provided with a driving mechanism; a feeding opening is arranged at the upper portion of the feeding part; a discharging opening is arranged at one side of the extrusion part; a water outlet is arranged at the lower portion of the extrusion equipment frame; the extrusion part is provided with a sealing gate; the driving mechanisms is arranged on the gate; and side face drainage holes are arranged at the two sides of the extrusion push head; the bottom surface of the extrusion push head is provided with bottom surface drainage holes; and the two sides of the extrusion equipment frame are provided with frame side drainage grooves. According to the high-pressure extrusion dehydrator provided by the invention, multi-direction drainage in the extrusion process is achieved, the overall drainage efficiency is improved, and dry and wet component separation in the extrusion process is achieved.

The invention relates to the technical field of rocket engines, in particular to a flexible extension nozzle with adaptive height. It includes the bell-shaped nozzle part of the basic section, the multi-stage bell-shaped extension section, the variable constraint point part and the supporting structure part; among them, the bell-shaped nozzle part of the basic section is a fixed nozzle, one end of which is connected to the rocket engine, and one end It is connected to the multi-stage bell-shaped extension section; the multi-stage bell-shaped extension section is a flexible extension nozzle, which is rolled up when it is in the initial state of launch, and then unfolded as the flight height increases; the variable restraint point part is located in the flexible extension On the nozzle, the support structure part is fixed to reach the predetermined position; the support structure part is located inside the flexible extension nozzle and expands as the flight height rises. The invention can solve the problem of low working efficiency of rockets with a wide working range, such as single-stage orbiting rockets, so that the airflow can always keep fully expanded and be in the best working state at any height.

The invention provides a multi-tube pulse knocking engine with a stable exhaust device. The engine comprises a knocking combustion chamber, a pressure stabilizing section and an exhaust nozzle. According to the engine, fuel gas is exhausted from knocking pipes and then enters the exhaust nozzle through the pressure stabilizing section, the pulsation characteristic of the fuel gas is weakened by the buffering effect of a pressure stabilizing cavity in the pressure stabilizing section, and thrust loss caused by exhaust over-expansion or under-expansion when a fixed geometric nozzle is adopted is reduced. The pressure stabilizing cavity has the specific effects that firstly, the section area of a flow channel of the pressure stabilizing cavity is larger than that of the knocking pipes, the fuel gas expands and decelerates after entering the pressure stabilizing cavity, then the pressure gradient in the fuel gas is reduced, in addition, when a shock wave structure generated by knocking combustion enters the pressure stabilizing cavity, diffraction occurs at the sudden expansion position of the section of the flow channel, the strength of shock waves is greatly weakened, and therefore pressure oscillation caused by the shock waves is abated; and secondly, the pressure stabilizing cavity is the confluence of the flow directions of the fuel gas in all the knocking pipes, and under the condition that all the knocking pipes work cooperatively, the pressure in the cavity can be always kept near the peak value. On the basis of the two points, the purposes of stabilizing the incoming flow pressure of the nozzle and improving the working efficiency of the nozzle are achieved. The engine can be applied to the fields of knocking propulsion and the like.

An aircraft includes a first battery, provisions for transport that are powered by a second battery, and a management system for transferring energy between the first battery and the second battery.

The invention discloses a corrugated turbulent flow plane cooling device and application, and belongs to the field of aero-engines. Comprising a cold side corrugated plate, a hot side flat plate and hollow turbulent flow columns, the cold side corrugated plate is located on one side of a cold air channel, and the hot side flat plate is located on one side of a high-temperature fuel gas channel; a hill-shaped middle channel with the sectional area changing periodically is formed between the cold side corrugated plate and the hot side flat plate. A plurality of hollow turbulent flow columns are arranged between the cold side corrugated plate and the hot side flat plate to form a middle channel containing the hollow turbulent flow columns, and the cold air channel is communicated with the high-temperature fuel gas channel through hollow holes in the hollow turbulent flow columns; the hollow turbulent flow columns are connected with wave troughs of the hot side flat plate and the cold side corrugated plate; and a gas film protection layer is formed on the hot-side flat plate by utilizing outflow of a trough high-pressure area of the cold-side corrugated plate. According to the invention, the amount of cold air near the wall surface of the spray pipe is reduced, and the thrust loss of the engine caused by cold air mixing is reduced under the condition of achieving higher cooling efficiency.

The invention relates to a flange structure suitable for an underwater electromagnetic launch device and relates to the technical field of launch. According to the flange structure, special launch cylinder mounting holes and water replenishment holes are designed according to the seawater flowing direction and the motor movement direction, before an aircraft is launched, the interior of the launchdevice communicates with sea to achieve pressure equalizing of the interior of the launch device and seawater, the effect that thrust of a launch motor has nothing to do with the water depth is achieved, and large-depth and variable-depth launch is achieved. When the aircraft is launched, the seawater is replenished for a negative-pressure area formed after the motor moves at a high speed, and under the premise that the structural rigidity is met, the thrust loss of the motor is reduced. According to the scheme, the structure is simple, work is reliable, and mounting and fixing connectors areprovided for a launching cylinder.

The invention relates to a plasma acceleration propulsion device, in particular to a plasma generation device and method based on a multi-stage magnetic field and a multi-section spray pipe. The technical problem that the specific impulse and thrust are not ideal due to the fact that a plasma generating device with a single-section spray pipe is not favorable for optimizing the configuration of a magnetic field and the trajectory of plasma jet is difficult to control is solved. The device comprises a working medium supply and distribution unit and a plasma generation acceleration unit, wherein the working medium supply and distribution unit comprises a working medium distributor, a second insulator, a transition electrode and a third insulator which are coaxially arranged from inside to outside; the plasma generation acceleration unit comprises n sections of spray pipes, n-1 sections of first insulators, accelerator electric field power supplies corresponding to the spray pipes, electromagnetic coils corresponding to the spray pipes and electromagnetic field power supplies corresponding to the electromagnetic coils; and the n sections of spray pipes and the n-1 sections of first insulators form a spray pipe assembly. The invention further provides a plasma generation method based on the multi-stage magnetic field and the multi-section spray pipe.

The invention relates to a flange structure suitable for an underwater electromagnetic launch device and relates to the technical field of launch. According to the flange structure, special launch cylinder mounting holes and water replenishment holes are designed according to the seawater flowing direction and the motor movement direction, before an aircraft is launched, the interior of the launchdevice communicates with sea to achieve pressure equalizing of the interior of the launch device and seawater, the effect that thrust of a launch motor has nothing to do with the water depth is achieved, and large-depth and variable-depth launch is achieved. When the aircraft is launched, the seawater is replenished for a negative-pressure area formed after the motor moves at a high speed, and under the premise that the structural rigidity is met, the thrust loss of the motor is reduced. According to the scheme, the structure is simple, work is reliable, and mounting and fixing connectors areprovided for a launching cylinder.

The invention discloses a balloon catheter for coronary artery balloon dilatation and a preparation method of the balloon catheter. The balloon catheter and the preparation method aim to achieve that proximal push force is transferred to the distal end, so that balloon inflating and contracting time are shortened. An inflatable balloon is connected with a liquid flowing chamber, the distal end of a proximal catheter body of the liquid flowing chamber is connected with a distal catheter body, the strength of the proximal catheter body is larger than that of the distal catheter body, the distal end of the proximal catheter body is provided with a U-shaped notch, an inner tube is arranged in the inflatable balloon and the distal catheter body, the distal end of the inner tube is connected with the distal end of the inflatable balloon to form a guide wire inlet, and an opening of the proximal end connected on the side wall of the proximal end of the distal catheter body forms a guide wire outlet. The preparation method includes the steps of manufacturing of all parts and assembly welding. Compared with the prior art, the balloon catheter and the preparation method have the advantages that inflating and deflating time of the balloon are shorter as liquid or gas flow is large; and since the strength of a distal metal tube is smaller than that of a proximal metal tube, and elasticity of the distal metal tube is larger than that of the proximal metal tube, distal conveyance of the catheter is facilitated when the catheter extends into a bend position of a blood vessel.

In recent years, unmanned aerial vehicle development is booming worldwide, and an unmanned aerial vehicle has many characteristics that a manned aerial vehicle does not have so that the unmanned aerial vehicle can be gradually concerned by people. The unmanned aerial vehicle has many advantages such as being low in construction cost and using cost and free of life danger that the manned aerial vehicle cannot match, and the unmanned aerial vehicle can excellently complete tasks that the manned aerial vehicle cannot do. On many occasions where the manned aerial vehicle cannot work, the unmanned aerial vehicle plays an enormous role. But many defects still exist in recovering of the unmanned aerial vehicle and therefore limit application of the unmanned aerial vehicle on many occasions. Accurate recovering and low-overload landing are the future developmental tendency of the unmanned aerial vehicle. The thrust turbine jet vector vertical take-off and landing unmanned aerial vehicle is one of the thrust vectoring unmanned aerial vehicles, has advantages that aircrafts of other types cannot match in the short take off and landing and super-maneuverable property aspects and has wide developmental prospects. Many foreign military powers have conducted in-depth studies on the aircraft, and verification aircrafts of various types are designed and manufactured to conduct experimental study.

The invention discloses a fire damper for a secondary detonation-breathing pulse detonation engine. A body of the fire damper is circular. The front end of the body is planar, and the rear end of the body is spherical. A plurality of circles of same conical holes are arranged on the circle of the body. The conical holes are equal in peripheral spacing. The outer diameter of the body is equal to the inner diameter of a straight barrel of an ejector. The inner diameter of the body is equal to the outer diameter of a pre-detonation tube. The body is mounted at one end of the pre-detonation tube, located at a lip of the ejector. An air channel is formed between the body and the lip of the ejector. Air flowing in from the front end of the fire damper passes maximum sectional areas of the conical holes and decreases along with the sectional areas of the conical holes. Air flowing in from the rear end of the fire damper passes minimum sectional areas of the conical holes first and increases along with the sectional areas of the conical holes, and forward flow resistance is smaller than reverse flow resistance. Forward transmission of high-temperature combustion products in the pulse detonation engine is improved effectively, and operational reliability of the pulse detonation engine is improved.

The invention relates to a pneumatic shock response spectrum tester, which comprises a base, a pneumatic hammerhead device, a resonant panel supporting device and a functional mechanism. The pneumatic hammerhead device comprises a hammerhead, guide rails and an internal cylinder, wherein one end of the internal cylinder is in sliding sealing with an open hole at one end of the hammerhead; grooveson the two sides of the hammerhead are clamped in the two symmetrically arranged guide rails; the guide rails are fixed on the base axially; a screw shaft at an end cover of the internal cylinder is connected with a joint bearing which is connected with a lug of an internal cylinder connecting block by a pin; and the bottom surface of the internal cylinder connecting block is fixed on the base bya thrust cylinder connecting seat. The resonant panel supporting device comprises a resonant panel, limiting clamping blocks, round rolling bars and a cushioning seat, wherein V-shaped grooves on thetwo sides of the resonant panel are clamped in a plurality of symmetrically arranged limiting clamping blocks; the resonant panel is supported by the round rolling bars fixed on a round bar mounting seat; and the limiting clamping blocks and the round bar mounting seat are fixed on the base. The pneumatic shock response spectrum tester is simply and compactly structured, simple and convenient to operate, short in stroke, high in excitation energy, convenient to arrange and maintain, safe and environment-friendly.

The invention discloses a pulse detonation engine with air entering from the side portion. The pulse detonation engine is at least composed of a head cone, a thrust wall, a detonation chamber, a blasting chamber, lateral air-inlet valves, main fuel spray nozzles, blasting fuel spray nozzles, igniters and a tail spray pipe. The thrust wall is arranged behind the head cone, and the blasting chamber is arranged behind the detonation chamber and shares the same space with the tail spray pipe; the blasting chamber comprises an energy collection cone and a movable baffle plate; the lateral air inlet valves are distributed on the pipe walls of the detonation chamber and the blasting chamber in the axial direction; the main fuel spray nozzles and the blasting fuel spray nozzles are distributed on the inner walls of the detonation chamber and the blasting chamber in the axial direction respectively; when the engine works, the multiple main fuel spray nozzles spray atomized fuel into the detonation chamber, the blasting fuel spray nozzles spray gasified blasting fuel into the blasting chamber, the igniters ignite to detonate a fuel mixture in the blasting chamber and then a mist-shaped mixture in the detonation chamber is detonated; and in the detonation cycle of the engine, the multiple lateral air inlet valves can achieve inlet air filling and mixing with the fuel rapidly, and detonation frequency and thrust of the engine are remarkably improved.

The invention discloses a baffling cooling device and application, and belongs to the field of aero-engines. Comprising a hot side flat plate, a baffle plate, an outer side cover plate, hollow turbulent flow columns, a baffle front plate and a baffle tail plate, a cold air channel is formed between the baffle plate and the outer side cover plate, and the hot side flat plate is located on one side of a high-temperature fuel gas channel; a plurality of hollow turbulent flow columns are arranged between the hot side flat plate and the baffle plate to form a middle channel containing the hollow turbulent flow columns; the baffle front plate is packaged at the front ends of the baffle plate and the outer side cover plate, and the baffle tail plate is packaged at the tail ends of the hot side flat plate and the outer side cover plate to form a semi-closed space; cooling air enters the middle channel to form cooling air flow, passes through the hollow turbulent flow columns and then is folded back at the baffling tail plate to form reverse baffling to enter the cold air channel, and finally flows into the high-temperature fuel gas channel through hollow holes in the hollow turbulent flow columns, and a cooling air film attached to the wall face is formed on the inner wall of the hot side flat plate. Due to baffling of cold air, the temperature of the whole cooling structure is uniform, and internal thermal stress is reduced.