55 results about "Shock cooling" patented technology

A jet impingement cooling apparatus includes a target having a flat surface and a plurality of recesses recessed from the flat surface, and a jet member having a plurality of orifices aligned with the plurality of recesses, respectively, wherein a cooling fluid is ejected from each orifice to the plurality of recesses.

The invention provides a spindle water cooling machine control method and system for a high-precision numerical control machine tool and a medium. The method comprises the steps that the working state of a main shaft is determined and obtained, the shunting flow of a three-way diverter valve is controlled according to the working state, a control strategy of the three-way diverter valve is obtained, and the three-way diverter valve is controlled according to the control strategy to conduct cooling water control. By arranging the three-way diverter valve, switching of cooling water between the main shaft branch and the bypass loop is achieved, dynamic adjustment of the variable frequency water pump and the refrigerating device is combined with the control unit, and the problem of shock cooling shrinkage caused by continuous cooling when the main shaft temporarily stops is solved; meanwhile, the problems that the efficiency is low and the service life of equipment is shortened due to frequent start and stop or continuous full-load operation of the water cooling machine are solved, and intelligent control over the spindle water cooling machine of the high-precision numerical control machine tool is achieved.

This invention provides an impact cooling assembly, turbine blade, turbine retaining ring, and gas turbine. The impact cooling assembly includes an impact plate and multiple impact holes arranged in an array. Each impact hole has at least two laterally extending protrusions. In the impact hole with the smallest center distance from the other impact hole, the protrusion on the side of one impact hole facing the other is offset from the protrusion on the side of the other impact hole facing the first impact hole. The impact cooling assembly of this invention can generate a strong cooling effect and provide strong heat transfer uniformity to the impact target surface, thereby reducing the thermal stress of the component where the target surface is located.

A guide vane (200, 300, 400, 500, 700) in a gas turbine engine (100) includes an inner platform (202), an outer platform (204), and an airfoil (206) extending therebetween. Side surfaces (302, 304) of the inner and outer platforms between the guide vane (300) and an adjacent guide vane define a first seal slot (310, 318), a second seal slot (312, 320), and a third seal slot (314, 322) forming a closed loop with three corners (316, 324). At least one of the corners is rounded. The guide vane (500, 600) includes a spoiler rib (504) and a peg fin (506, 604) disposed in an interior portion (406) of the airfoil. The peg fin (604) is disposed in a region of a trailing edge (212). The inner platform defines a film cooling hole (410) disposed at an outer surface (408) facing the airfoil. The film cooling hole is arranged in a scalloped shape. An inner surface (804) of the inner platform and an outer surface (704) of the outer platform include impingement cooling ribs (802, 702). The inner platform projects further upstream in a flow direction than the outer platform.

The application relates to a double-wall impact cooling device and application, belonging to the field of aero-engines, which comprises a cold-side flat plate, a hot-side flat plate and hollow turbulence columns, the cold-side flat plate is located on one side of a cold gas channel and is provided with a plurality of impact holes, the hot-side flat plate is located on one side of a high-temperature gas channel, a plurality of hollow turbulence columns are arranged between the cold-side flat plate and the hot-side flat plate, an intermediate channel containing the hollow turbulence columns is formed, and the cold gas channel and the high-temperature gas channel are communicated through hollow holes in the hollow turbulence columns. The application fully utilizes the advantages of suction-strengthened impact convection cooling and the hollow holes easy to flow to form a gas film layer, reduces the temperature distribution and temperature gradient of the heat shield, and achieves the purpose of high-efficiency cooling.

The utility model discloses a bypass ventilation system with multiple stages of cyclone cylinders connected in series and in parallel. The bypass ventilation system with the multiple stages of cyclone cylinders connected in series and in parallel comprises an air taking opening, shock cooling equipment, a primary cyclone cylinder, an efficient cyclone cylinder and an exhaust fan which are sequentially connected. The insertion depth L of the inner barrel of the efficient cyclone barrel is set to be 1200-5200 mm, and the air speed of an air inlet of the efficient cyclone barrel is set to be 20-30 m / s. According to the bypass ventilation system with the multi-stage cyclones connected in series and in parallel, the efficient cyclones are used for replacing a bag type dust collector, secondary cooling is not carried out on smoke any more, the temperature of a waste gas outlet is increased, and the working quality of follow-up waste gas treatment equipment is guaranteed.

The invention relates to the technical field of new energy automobile detection, and discloses a bearing rotation testing machine for a new energy automobile, which comprises a box body, an observation window is rotatably arranged on the box body, a vertical plate is fixedly arranged in the box body, a driving wheel is rotatably arranged on the vertical plate, the driving wheel is driven by a driving source to rotate, and the observation window is fixedly arranged on the box body. Water is injected into the hollow shaft through the water supply assembly to absorb heat of a high-temperature part, then the waste heat utilization assembly is matched with the heat exchange assembly, warm water obtained after temperature rise is conveyed to the water soaking basin, and the temperature of water used for water soaking is close to the high temperature of a synchronous belt and a bearing; the shock cooling impact of high-temperature parts when meeting cold water is avoided, the problems that in traditional equipment, a belt is subjected to shock cooling shrinkage failure, and a bearing shafting generates internal stress are solved, the service life of the synchronous belt and the equipment parts is prolonged, and the consistency and stability of detection data are guaranteed.

The invention belongs to the field of aero-engine combustion chamber outlet parameter measurement, and particularly relates to an aero-engine combustion chamber outlet sample gas sampler, which adopts a local impact cooling heat exchange structure, improves the temperature resistance of the windward side of the sampler, and can prolong the service life of the sampler in a combustion chamber outlet complex parameter measurement process. The flow of cooling water is greatly increased, and the cooling capacity of the whole structure is improved; and the mounting seat cavity is designed to cool the mounting seat, so that the possibility that the mounting seat is ablated in a high-temperature state is reduced.

The application discloses a double-wall blade for a gas turbine and the gas turbine, and the blade comprises an outer shell and a plurality of inner walls, the outer shell defines a cavity, a plurality of air outlet slots extending along the length direction of the blade are arranged on the outer shell, the openings of the air outlet slots are directed to the trailing edge of the outer shell, the blade is provided with an air inlet communicating with the cavity, the plurality of inner walls are arranged in the cavity in the circumferential direction and connected with the outer shell, the outer wall surface of the plurality of inner walls and the inner wall surface of the outer shell form a plurality of cooling channels, the inner wall is provided with at least one through hole communicating the cooling channels with the cavity, and the end of the plurality of cooling channels close to the trailing edge of the outer shell is communicated with the plurality of air outlet slots one by one. The double-wall blade disclosed by the application uses the cooling mode combining impingement cooling and air film cooling, greatly improves the cooling effect of the cooling air on the high-temperature blade, and enables the blade to have strong heat exchange capacity.

The application belongs to the technical field of S-bend nozzle design of an aero-engine, and particularly relates to an aero-engine infrared radiation signal control S-bend nozzle, which comprises a round-to-square section, an inlet end of which is used to be connected to the rear end of a low-pressure turbine; an S-bend left side wall, an S-bend right side wall, an S-bend lower side wall and an S-bend upper side wall are connected to the outlet end of the round-to-square section, are spliced with each other, and the S-bend upper side wall has a plurality of impact cooling holes on the rearward visible part thereof, and the outer wall has an annular support edge surrounding the rearward visible part; a cover plate is connected to the annular support edge and has a low-temperature medium inlet hole thereon; a flow divider plate is arranged in the annular support edge, is connected to the cover plate, forms a flow dividing cavity with the cover plate, has a plurality of flow dividing holes thereon, and forms an impact cooling cavity with the S-bend upper side wall; a heat shield is connected to the inner side of the S-bend upper side wall, covers the rearward visible part, forms a heat insulation cavity with the S-bend upper side wall, and has a plurality of air film holes thereon.

The invention belongs to the technical field of aero-engines, and discloses a turbine casing impact heat exchange test method and system, and the method comprises the steps: determining impact hole parameter research variables, and the structure parameters and number of aperture replacement parts, and adjusting the diameters of impact holes through the aperture replacement parts; the number and arrangement of the impact holes are adjusted by installing hole diameter replacement parts and a plurality of plugs on the impact hole plates; the impact target distance is adjusted through different numbers of target distance replacement parts; and performing an impact cooling heat exchange test on each combined cartridge receiver impact heat exchange test piece to obtain heat exchange coefficients of impact target surfaces of different research variables. The hole diameter replacement part is arranged to change the hole diameter and hole arrangement, the target distance replacement part is arranged to change the impact distance, the number of test pieces can be reduced, the machining cost is reduced, the machining period is shortened, meanwhile, disassembly and assembly are easy, disassembly and assembly time is short, test consistency is guaranteed, effective measurement of the impact heat exchange test temperature is achieved, and therefore the corresponding impact heat exchange coefficient is obtained.

The invention belongs to the technical field of jet impingement cooling performance optimization, and particularly relates to a multi-objective optimization method and system for a variable-spacing multi-jet impingement cooling system.The method comprises the steps that the basic structure of the variable-spacing multi-jet impingement cooling system is determined, and a three-dimensional model is established; based on the three-dimensional model, utilizing Workbench software to establish a target simulation model; key parameters of the variable-spacing multi-jet impingement cooling system are determined through parameter sensitivity analysis, the key parameters serve as to-be-optimized design variables, sample points are obtained based on an experimental design method, to-be-optimized target values corresponding to the sample points are calculated through parameterized simulation, and the to-be-optimized target values are obtained. Establishing an agent model of the variable-spacing multi-jet impingement cooling system based on the target value to be optimized; and constructing a multi-objective optimization model based on the established agent model, and optimizing parameters of the variable-spacing multi-jet impingement cooling system through a multi-objective optimization algorithm. According to the invention, the heat dissipation performance of the server chip is improved.

The invention relates to the technical field of aero-engine low-pressure turbine blades, in particular to an internal cooling structure of an aero-engine low-pressure turbine blade, which comprises a blade body, and a front edge impingement cooling cavity, a convection heat exchange cavity and a top confluence cavity are arranged in the blade body; a top edge plate is arranged at the upper end of the blade body, and an air outlet is formed in the top edge plate; a longitudinal partition plate is arranged between the front edge impingement cooling cavity and the convection heat exchange cavity; a transverse partition plate is arranged between the convection heat exchange cavity and the top confluence cavity. The upper end of the longitudinal clapboard is connected with the front edge end of the transverse clapboard; a plurality of impingement cooling holes are formed in the longitudinal partition plates and the transverse partition plates; convex ribs are arranged on the inner wall of the front edge side of the front edge impingement cooling cavity; a first rib is arranged on the inner wall of the suction surface side of the convection heat exchange cavity; second ribs are arranged on the inner wall of the pressure face side of the convection heat exchange cavity.

The application discloses a blade tip structure and blade with improved cooling efficiency, comprising a groove formed in a tip clearance leakage flow area, the inner wall of the groove is inclined to the mean camber line of the blade, the outlet of the cooling hole of the blade is located on the inner wall of the groove, the flow direction of the cooling working medium output by the cooling hole is the same as the flow direction of the leakage flow, the dovetail groove and the cooling hole are combined, the dovetail groove suppresses the formation of the pressure side corner vortex and reduces the heat exchange coefficient on the near pressure surface side; the cooling hole directly blows out the cooling gas to the blade tip high heat exchange area for impact cooling, the cooling gas flows to the suction surface side under the action of the pressure difference after the impact cooling, the blade tip temperature distribution is improved, the cooling gas utilization rate is improved, and the blade tip temperature distribution is uniform; on the other hand, the high-speed cooling jet forms an approximate circumferential rib structure, plays a role of aerodynamic sealing, suppresses the development of the leakage flow to the downstream, reduces the blade tip leakage amount, improves the gas-thermal characteristics of the blade tip area, prolongs the service life of the high-pressure turbine blade, and guarantees the safe and effective operation of the gas turbine.

A turbine nozzle or blade includes an airfoil body defined by a concave pressure side outer wall and a convex suction side outer wall that connect along leading and trailing edges and, therebetween, form a radially extending chamber. The airfoil body has an inner surface facing the radially extending chamber. An impingement cooling structure is within the radially extending chamber. The impingement cooling structure includes: a wall spaced from the inner surface of the airfoil body; a plurality of holes defined through the wall; and a plurality of elongated thermal flex elements defined in the wall. Because the nozzle or blade is made by additive manufacturing, the airfoil body and the impingement cooling structure include a plurality of integral material layers.

An impingement plate (86, 186, 286) for providing cooling to a gas turbine engine component includes a first surface (88) with a second surface (90) opposite the first surface (88). The first surface (88) and the second surface (90) are surrounded by a perimeter (112) defining an edge (112A, 112B). A first plurality of holes (92) extend through the impingement plate (86, 186, 286). A plurality of protrusions (96) extend outward from one of the first surface (88) or the second surface (90) and have a protrusion surface transverse to one of the first surface (88) or the second surface (90). At least one second impingement hole (98) extends through the protrusion surface.

The utility model relates to the rolling technical field, and one embodiment of the utility model provides an elliptical material's control cooling and rolling device, it includes roller group, the outer side of roller group is provided with rolling mechanism, the rolling mechanism includes the side plate of placement in the outer side of roller group, and the inner side between side plate is rotatably installed with roll no. The utility model discloses through making water flow cover more fit elliptical material middle area, avoid edge portion and cool too much while middle part cooling deficiency, compared with traditional elliptical material's control cooling and rolling device, the elliptical material's control cooling and rolling device of this kind through flat ellipse spout setting flat ellipse spout, increase middle part water flow density, improve middle part water flow speed, strengthen the impact cooling effect, solve the uneven cooling problem of elliptical material because of cross section shape, avoid edge portion and cool too much fragile or middle part residual defect, through above technical scheme, solved the technical problem of the poor cooling effect of elliptical material in relevant technology.

An arrangement (1) for the rear impact cooling of an impact plate (2) thermally stressed on its front side (3) by a hot medium (4) comprises the impact plate (2), a perforated plate (6) arranged at a distance from the rear side (5) of the impact plate (2), cooling fluid passage holes in the perforated plate (6) directed towards the rear side (5) of the impact plate (2), a cooling fluid outflow channel (9) formed between the rear side (5) of the impact plate (2) and the perforated plate (6), open in a transverse flow direction (10) at a cooling fluid outlet opening, and a cooling fluid supply channel (13) to which the cooling fluid passage holes are fluidically connected. The cooling fluid supply channel (13), the cooling fluid passage holes, and the cooling fluid outflow channel (9) form a cooling fluid passage through the arrangement (1) that is closed up to the cooling fluid outlet opening.To increase heat transfer from the baffle plate (2) to a cooling fluid (15) flowing through the cooling fluid passage, fluidic oscillators (19) are arranged in the cooling fluid passage.

The invention discloses an efficient cooling low-emission combustor for a gas turbine, and belongs to the technical field of gas turbines. The combustor comprises a nozzle shell, a flame tube wall and a premixer, a cavity is formed in the flame tube wall and is divided into an outer ring cavity and an inner ring cavity by a partition plate, impact holes are formed in the partition plate, and cooling ring ribs are arranged in the inner ring cavity; the outer ring cavity is communicated with a fuel outlet of the nozzle shell through a first-stage fuel pipeline, and the inner ring cavity is communicated with the premixer through a second-stage fuel pipeline. Fuel enters from the outer ring cavity, impacts and cools the inner ring cavity through the impact holes, and then is subjected to heat exchange enhancement through the ring ribs, so that impact-ring rib combined cooling of the fuel on the flame tube wall is realized. The structure can improve the ignition and combustion performance, reduce the consumption of cooling air and increase the proportion of lean premixed combustion participating in air. Compared with a traditional air cooling structure, the cooling efficiency and the combustion stability can be remarkably improved, emission of NOx and CO is reduced, and the design of the gas turbine combustion chamber which is efficient, energy-saving and low in emission is achieved.

The utility model discloses a waterproof roll film cooling device, which relates to the technical field of waterproof roll processing, and comprises a cooling pool and a spraying pre-cooling mechanism, two opposite side walls of the cooling pool are correspondingly provided with threading grooves, the inner wall of the upper part of the cooling pool is fixedly provided with a water receiving groove, the bottom wall of the water receiving groove is provided with a downward pressing guide mechanism, and the spraying pre-cooling mechanism is provided with a spraying pre-cooling mechanism. The downward pressing guide mechanism is used for pressing the waterproof coiled material downwards into the cooling pond from the penetrating guide groove, a guiding-out groove is formed in one side wall of the cooling pond, a guiding-out roller is rotationally arranged in the guiding-out groove, and the spraying pre-cooling mechanism is arranged on the cooling pond and located above the water receiving groove. The utility model particularly provides waterproof coiled material film coating cooling equipment which is used for pre-cooling a waterproof coiled material, so that the shock cooling quality of the waterproof coiled material is prevented from being damaged, and the waterproof coiled material can be conveniently threaded in a cooling pool.

The application discloses a fin type cooling layer plate applied to a hot end part of a gas turbine and relates to the technical field of turbine blades. The fin type cooling layer plate is integrally formed by adopting a 3D printing technology and comprises an inner layer wall, a connecting rib and an outer layer wall. The outer layer wall is in a square tube shape, the connecting rib is in a fin type parallel layout, the connecting rib connects the inner layer wall and the outer layer wall and divides a cooling space into a T shape. The application constructs an impingement cooling microchannel by adopting the integral layer plate cooling structure in which the fin type connecting rib is connected with the inner layer wall and the outer layer wall, improves the effective cooling area and the cooling uniformity of cooling air, reduces the flow resistance, improves the comprehensive cooling efficiency, and the comprehensive cooling efficiency of the application is improved by 39% compared to the prior art, the effective cooling area is improved by 18%, and the total pressure loss coefficient is reduced by 16%. The application can significantly enhance the cooling effect and the cooling air utilization rate of the turbine blade layer plate cooling structure.

The turbine nozzle or blade includes an airfoil body defined by concave pressure side and convex suction side outer walls that connect along leading and trailing edges and form a radially extending chamber therebetween. The airfoil body has an inner surface facing the radially extending chamber. An impingement cooling structure is within the radially extending chamber. The impingement cooling structure includes a wall spaced from the inner surface of the airfoil body, a plurality of holes defined through the wall, and a plurality of elongated thermal flex elements defined in the wall. Because the nozzle or blade is fabricated by additive manufacturing, the airfoil body and impingement cooling structure include multiple monolithic layers of material.

The application provides a combustion chamber flame tube cooling wall surface structure and a flame tube, which are integrated by impingement cooling and mixing, and overcome the problem of uneven temperature distribution of the existing flame tube wall surface. The combustion chamber flame tube cooling wall surface structure comprises an impingement hole wall and an inner layer wall, an annular air passage is arranged between the impingement hole wall and the inner layer wall, the impingement hole wall and the inner layer wall are connected through an air film ring, an air film cooling ring slot groove is formed at the tail of the flame tube, the air film ring is provided with air film cooling holes, vertical first diameter holes are arranged on the impingement hole wall to form vertical impingement jets, the diameter of the first diameter holes is less than a first preset value, and main combustion holes and mixing holes are arranged on the inner layer wall. The application improves the traditional flame tube cooling structure, makes the air of the cooling wall surface participate in the mixing organization of the combustion chamber, and greatly reduces the amount of cooling air under the condition of ensuring the cooling effect.

The invention discloses a permanent magnet-based magnetic nanofluid jet impingement cooling device which mainly structurally comprises a water inlet pipe, a water outlet pipe, a cover plate, an upper partition plate, a jet pipe, a lower partition plate and a cavity which are sequentially matched from top to bottom, the cavity is divided into three parts by the upper partition plate and the lower partition plate, the upper part is a jet distribution layer, and the middle part is a backwater gathering layer; and the lower part is a jet flow layer which is connected with the jet flow distribution layer and the jet flow layer through jet flow pipes. Jet flow impact and nanofluid are used for strengthening heat transfer, meanwhile, the uniformity of jet flow impact heat transfer is improved through the combined action of rotational flow and a magnetic field, an ascending rotational flow guide piece is adopted on the bottom face to guide fluid to generate ascending rotational flow so as to prevent particles from precipitating on a heat exchange face, and meanwhile certain heat transfer can be strengthened as a fin. The permanent magnet magnetic field effect and the nano magnetic fluid flow are combined and applied to the jet flow impact process, the heat transfer performance is improved, the heat transfer uniformity is improved, and support is provided for cooling of heat exchange equipment with high heat flux density.

An integrated combustor nozzle (100) includes a combustion liner (110) extending radially between an inner liner section (106) and an outer liner section (108), the combustion liner (110) including a forward end portion (112), an aft end portion, a first sidewall (116), and a second sidewall (118). The aft end portion (114) of the combustion liner (110) defines a turbine nozzle (120). The combustion liner (110) defines a cavity (126) forward of the turbine nozzle (120). The cavity (126) extends between the first sidewall (116) and the second sidewall (118). An impingement cooling apparatus is positioned within the cavity (126). The impingement cooling apparatus includes a flange (310, 311). The impingement cooling apparatus further includes a plurality of impingement members (302) configured to direct a coolant to impinge on the first sidewall (116) and the second sidewall (118). Each impingement member (302) extends from a respective inlet (313) defined within the flange (310, 311) to a respective closed end (312). A plurality of impingement holes (304) are defined on each impingement member (302).

The invention discloses an automobile part machining mold, and relates to the technical field of automobile part machining, and the automobile part machining mold is technically characterized in that the mold comprises a fixed mold body fixed to a supporting table, a cooling cavity is formed in the fixed mold body, a first storage cavity and a second storage cavity are formed below the supporting table, and a first piston plate and a second piston plate are arranged in the first storage cavity and the second storage cavity in a sliding mode respectively; the first connecting pipe and the second connecting pipe at the two ends of the first storage cavity are both connected with a connector on the fixed mold, and the connector is connected with the second storage cavity through a third connecting pipe; the second storage cavity is communicated with the first storage cavity through a first communicating pipe, a cooling box on the supporting table is connected with the two ends of the second storage cavity through a connecting pipe set, a stepped cooling mode is adopted, liquid with waste heat is used for primary cooling, then cooling liquid is used for final cooling, and the situation that a mold is damaged due to shock cooling is avoided; meanwhile, molten plastic is prevented from being solidified in advance to influence filling, and product quality and mold service life are improved.

The invention provides a flow guide device and a gas turbine, and relates to the technical field of gas turbine equipment. The flow guide device comprises a flow guide partition plate and a cover plate; the cover plate and the bearing seat form an annular channel for cold air circulation; an annular airflow channel for adjusting the width and angle of a cold air outlet of the annular channel is arranged on the flow guide partition plate. Cold air at the air compressor end is guided to the position between the bearing seat and the rotor disc through the annular channel, the width and angle of the cold air outlet of the annular channel are adjusted through the annular air flow channel, the outflow speed and the outflow angle of the cold air are controlled, and impact cooling can be directionally formed on the surface of the rotor disc when the cold air flows out. On the premise that the cold air amount is given, the utilization efficiency of cooling air can be effectively improved, the key area on the rotor disc is efficiently cooled, the temperature field of the rotor disc is improved, the reliability of the rotor disc is improved, the service life of the rotor disc is prolonged, and then the overall efficiency of the gas turbine is guaranteed.