82 results about "Shock cooling" patented technology

The invention discloses a reflooding cooling process analysis method and system. The method comprises the following steps: step 1, obtaining reflooding process data; 2, according to the re-submerging process data, obtaining a liquid core shear stress relation and a constitutive relation, and obtaining a disturbance wave height; 3, according to local shock cooling in the cooling process, a liquid film dynamic equilibrium relation is constructed, and the average vapor film thickness is obtained; 4, axial distribution of the minimum film state boiling temperature is obtained according to the disturbance wave height and the average steam film thickness; 5, shock cooling moments at different axial heights are obtained according to the axial distribution of the minimum film state boiling temperature, the relationship among the shock cooling speed, the shock cooling moments and the heights is obtained, the shock cooling speed of all the shock cooling moments is solved, and then analysis of the re-submerging cooling process is completed; according to the method, the problem of strong correlation of time-physical behaviors of two-phase evolution in the whole re-submerging process is solved, and accurate prediction of the whole strong transient re-submerging cooling process is realized.

A cold plate assembly includes a cold plate having first surface attachable to heat generating electronic component(s) and opposite second surface enclosed by an encapsulating lid that has inlet and exhaust ports. A distributed impingement and recovery manifold (DIRM) is positioned above the second surface of the cold plate. The DIRM includes an intake manifold presenting at least one sequence of nozzle openings providing direct impingement of cooling liquid on corresponding sections of the second surface and a return manifold which facilitates a return of exhaust cooling liquid to the exhaust port following at least one direct impingement of the portion of cooling liquid onto the corresponding sections to provide distributed, localized impingement cooling at corresponding sections of the cold plate.

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.

Impact insert (80) for a turbomachinery component, the impact insert (80) comprising: - a double-walled section (1, 2) having an inner wall (81) and an outer wall (82) defining an inner channel (501) formed on an inner surface (81a) of the inner wall (81), an outer channel (503) formed on an outer surface (82b) of the outer wall (82), and a central channel (502) formed between the inner surface (82a) of the outer wall (82) and the outer surface (81b) of the inner wall (81); and - a plurality of impact cooling holes (85) formed in the outer wall (82) and configured to eject impact jets (86) into the outer channel (503), wherein the impact jets (86) are formed from the cooling air (5) of the central channel (502); characterized in that the impact insert (80) further - at least one suction line (9) extending between the outer wall (82) and the inner wall (81) over the central channel (502) and comprising an inlet (9a) on the outer channel (503) and an outlet (9b) on the inner channel (501) for sucking cooling air (5) from the outer channel (503) into the inner channel (501), and wherein the suction line (9) is aerodynamically shaped with respect to a flow of the cooling air (5) flowing through the central channel (502).

The invention provides a high-pressure turbine guide vane cooling structure and an aero-engine, and belongs to the technical field of aero-engines, the high-pressure turbine guide vane cooling structure comprises a high-pressure turbine guide vane, the interior of the high-pressure turbine guide vane is provided with a cavity, the basin side and the back side of the high-pressure turbine guide vane are provided with film holes communicating with the cavity, and the tail of the high-pressure turbine guide vane is provided with a tail crack; the impact guide pipe is arranged in the cavity, an airflow channel is formed between the body part of the impact guide pipe and the cavity, the tail part of the impact guide pipe is provided with a convex hull, the convex hull is in close contact with the cavity so as to separate the airflow channel, and the body part of the impact guide pipe is respectively provided with impact cooling holes communicated with the airflow channels on the basin side and the back side; and a tail impact hole communicated with the tail crack is formed in the convex hull. According to the high-pressure turbine guide vane cooling structure, the parallel flow path design of the basin side, the back side and the tail edge is achieved, through refined flow path control, the problem that the outflow amount of the tail edge is insufficient under the condition of the low expansion ratio is solved, and the cold air consumption and the cold air quality of each part per unit area are 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 discloses a bypass exhaust dechlorination method and bypass exhaust dechlorination equipment used by the bypass exhaust dechlorination method. The bypass exhaust dechlorination method comprises the following steps: step 1, feeding kiln tail flue gas into shock cooling equipment (1) for cooling; 2, the cooled flue gas enters high-temperature dust removal equipment to be subjected to gas-solid separation; and 3, the waste gas obtained through separation in the step 2 is fed into a preheater (101), and the temperature of the waste gas fed into the preheater (101) in the step 3 is higher than 400 DEG C. According to the bypass ventilation dechlorination method, the smoke of the kiln tail smoke chamber is extracted through the pressure difference between the kiln tail smoke chamber and the preheater, after high-temperature dust removal treatment, waste gas returns to the preheater and is subjected to denitrification treatment together with system waste gas, and additional equipment does not need to be additionally arranged.

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 discloses a diesel injector cooling unit in the technical field of diesel engines, and the diesel injector cooling unit comprises a plurality of cooling parts uniformly distributed around an oil nozzle body in the circumferential direction, and each cooling part comprises a bag body, a liquid spraying pipe communicated with the bag body and an extrusion part used for extruding the bag body to perform pulse type impact on the oil nozzle body by cooling liquid; the oil nozzle body is obliquely arranged and rotationally installed on the air cylinder cover, a cooling groove is formed in the air cylinder cover, and the cooling component is arranged on the inner wall of the peripheral side of the cooling groove. A driving mechanism is further arranged in the cooling groove and used for driving the oil nozzle body to rotate so as to achieve multi-direction oil injection. According to the device, rotation control of the oil nozzle body and a cooling system are highly integrated, an integrated unit of rotary oil injection and pulse impact cooling is formed, the oil injection performance is improved, cooling liquid is periodically compressed and injected through the bag body, the surface of the oil nozzle body is impacted in a pulse mode, and strong local convection heat exchange is formed.

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.

A cold plate assembly includes a cold plate having first surface attachable to heat generating electronic component(s) and opposite second surface enclosed by an encapsulating lid that has inlet and exhaust ports. A distributed impingement and recovery manifold (DIRM) is positioned above the second surface of the cold plate. The DIRM includes an intake manifold presenting at least one sequence of nozzle openings providing direct impingement of cooling liquid on corresponding sections of the second surface and a return manifold which facilitates a return of exhaust cooling liquid to the exhaust port following at least one direct impingement of the portion of cooling liquid onto the corresponding sections to provide distributed, localized impingement cooling at corresponding sections of the cold plate.

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.

The present invention provides a cooling device for a telemetry module for measuring the dynamic stress of a low-pressure fan of a turbofan engine, comprising: an engine cowl, a through hole is opened at the center position of the front end thereof, and a tension bolt with a vent hole is arranged at the through hole; a gas collecting box is arranged inside, and the air inlet of the gas collecting box is communicated with the air outlet end of the tension bolt; the housing of the gas collecting box is sequentially connected to a first mounting seat and a second mounting seat to form a cooling gas collecting cavity; a plurality of impingement cooling holes are uniformly arranged along the circumference on the housing of the first mounting seat for enabling the cooling gas to form a jet impingement cooling on the transmitting device of the telemetry module through the impingement cooling holes; the housing of the second mounting seat is connected to a telemetry module receiving device adapter seat to form a telemetry cavity, and a plurality of cooling gas exhaust holes are uniformly opened along the circumference on the adapter seat; the cooled gas enters the telemetry cavity and then is discharged from the cooling gas exhaust holes. The present invention can ensure that the telemetry module does not exceed the limit temperature during the entire fan dynamic stress measurement test, so as to achieve the reliability and stability of the cooling scheme.

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.

The utility model discloses a magnetic stirrer with an automatic water replenishing structure, which comprises a base and an electromagnet, a control panel is arranged on the front surface of the base, the electromagnet is mounted on the bottom surface of the inner wall of the base in an embedded manner, a shell is solely mounted on the upper surface of the base, the shell is connected with the automatic water replenishing structure, and the automatic water replenishing structure is connected with the electromagnet. And the automatic water supplementing structure can uniformly inject water into the inner shell through the annular pipe, the automatic water supplementing structure comprises a clamping block, the clamping block is fixedly installed at the upper end of the outer shell, the lower surface of the clamping block is fixedly connected with the upper end of the inner shell, and the inner shell is located in the outer shell. According to the magnetic stirrer with the automatic water replenishing structure, the automatic water replenishing structure is arranged, water can be replenished to the inner shell through the annular pipe, water can be uniformly added during water injection of the annular pipe, shock cooling of the water temperature can be avoided, the magnetic stirrer is further provided with a temperature sensing structure, and an air cylinder in the connecting shell can push a temperature sensor into a material box through a connecting plate.

A turbine nozzle (112) or blade (114) includes an airfoil body (134) defined by a concave pressure-side outer wall (150) and a convex suction-side outer wall (152) connected along leading and trailing edges (154, 156) and defining a radially extending chamber (160) therebetween. The airfoil body (134) has a smooth inner surface (162) facing the radially extending chamber (160). An impingement cooling structure (170) is located within the radially extending chamber (160). The impingement cooling structure (170) includes: a wall (172) spaced from the inner surface (162) of the airfoil body (134); a plurality of holes (174) defined through the wall (172); and a plurality of elongated thermoflex elements (190) defined in the wall (172).Since the nozzle (112) or the blade (114) is manufactured by additive manufacturing, the airfoil body (134) and the impingement cooling structure (170) include a plurality of integrated 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 invention discloses a combined cooling structure with a partition plate for a turbine blade of a gas turbine, which is characterized in that film holes are formed in the wall surface of a blade body of the turbine blade, a cold air channel is arranged in the blade body of the turbine blade, an impact plate is arranged in the cold air channel, and the impact plate and the wall surface of the blade body are spaced to form an impact cavity; the impact plate is provided with impact holes, partition plates are arranged in the impact cavity and divide the impact cavity into a plurality of impact sub-cavities and hollow cavities in the length direction of the impact cavity, the impact sub-cavities are correspondingly communicated with the impact holes and the air film holes, and the hollow cavities are sealed cavities defined by the partition plates, the impact plate and the wall face of the blade body. According to the turbine blade, the partition plates are arranged in the impact cavity to construct the multiple impact sub-cavities and the hollow cavity, the impact sub-cavities can force impacted cold air to only flow out of the air film holes in the corresponding areas, and therefore the influence of transverse flow on downstream impact cooling is reduced, the hollow cavity is a sealed cavity, no cold air exists in the hollow cavity, and the impact cooling effect is good. The actual space of the impact cavity is reduced, so that the utilization rate of cold air can be improved.