33results about How to "Reduce metal temperature" patented technology

An engine cylinder head has an exhaust manifold cast in the cylinder head. A turbocharger with a turbine body is adapted to mount directly to an exhaust outlet mounting face of the exhaust manifold. A dual level water jacket within the cylinder head has separate coolant feeds for upper and lower cooling jackets. The cooling jackets extend above and below the exhaust outlet and are connected inward of the exhaust mounting face to reduce metal temperatures of the mounting face below those of the turbocharger exhaust inlet flange. Separate cores for the upper and lower jackets are connected at coolant inlet and outlet locations at opposite ends. Intermediate core print connectors form controlled flow passages between the upper and lower jackets and exhaust ports in the cylinder head and integrated manifold. The improved cooling in these areas lowers operating temperatures in the cylinder head and obviates the need for a separate exhaust manifold.

An engine cylinder head has an exhaust manifold cast in the cylinder head. A turbocharger with a turbine body is adapted to mount directly to an exhaust outlet mounting face of the exhaust manifold. A dual level water jacket within the cylinder head has separate coolant feeds for upper and lower cooling jackets. The cooling jackets extend above and below the exhaust outlet and are connected inward of the exhaust mounting face to reduce metal temperatures of the mounting face below those of the turbocharger exhaust inlet flange. Separate cores for the upper and lower jackets are connected at coolant inlet and outlet locations at opposite ends. Intermediate core print connectors form controlled flow passages between the upper and lower jackets and exhaust ports in the cylinder head and integrated manifold. The improved cooling in these areas lowers operating temperatures in the cylinder head and obviates the need for a separate exhaust manifold.

A turbine blade having an internal cooling system with dual serpentine cooling channels in communication with tip cooling channels is disclosed. In at least one embodiment, the cooling system may include first and second tip cooling channels in communication with the first and second serpentine cooling channels, respectively. The first tip cooling channel may extend from the leading edge to the trailing edge and be formed from a first suction side tip cooling channel and a first pressure side tip cooling channel. The second tip cooling channel may extend from a midchord region toward the trailing edge and may be positioned between the pressure and suction sides such that the second tip cooling channel is positioned generally between the first suction side and pressure side tip cooling channels. The first and second tip cooling channels may exhaust cooling fluids through the trailing edge.

A localized directional impingement cooling is used to reduce the metal temperatures on highly stressed regions of the tip shroud.

An interface region between a combustion liner and a transition duct of a gas turbine combustor is disclosed having improved cooling such that component life is increased and metal temperatures are lowered. An aft end of a combustion liner is telescopically received within the transition duct such that a combustion liner seal is in contact with an inner wall of the transition duct inlet ring. Increasing the dedicated cooling air supply to the combustion liner aft end, coupled with a modified combustion liner aft end geometry, significantly reduces turbulence and flow re-circulation, thereby resulting in lower metal temperatures and increased component life. Multiple embodiments of the interface region are disclosed depending on the amount of cooling required.

A turbine airfoil having walls with radial extending near wall cooled vortex flowing channels to provide for a high cooling effectiveness with a low cooling flow requirement. A thermal skin is bonded over a main spar in which skewed ribs extend from both the spar and the thermal skin to form a vortex flowing path along the radial channels. The vortex flowing radial channels can be single pass radial channels or connected to form a 3-pass serpentine flow circuit with tip holes to discharge the cooling air from the radial channels and cool the blade tip periphery.

A turbine blade having an internal cooling system with dual serpentine cooling channels in communication with tip cooling channels is disclosed. In at least one embodiment, the cooling system may include first and second tip cooling channels in communication with the first and second serpentine cooling channels, respectively. The first tip cooling channel may extend from the leading edge to the trailing edge and be formed from a first suction side tip cooling channel and a first pressure side tip cooling channel. The second tip cooling channel may extend from a midchord region toward the trailing edge and may be positioned between the pressure and suction sides such that the second tip cooling channel is positioned generally between the first suction side and pressure side tip cooling channels. The first and second tip cooling channels may exhaust cooling fluids through the trailing edge.

A turbine engine component having improved cooling is provided. The turbine engine component includes an airfoil portion having a leading edge, a trailing edge, a pressure side, a suction side, a root, and a tip, and at least one cooling circuit in a wall of the airfoil portion. The at least one cooling circuit has at least one passageway extending between the root and the tip. The at least one passageway has an aspect ratio of no greater than 2:1, and preferably substantially unity.

A cooling system for a turbine airfoil of a turbine engine has a multiple suction side cooling slots extending from a front edge on the suction side to the center of the trailing edge or even to the pressure side of the center line and a pressure side cooling slot curving to a pressure side outlet forward of the trailing edge and having a front pressure side lip that is aligned with or forward of the front edge of the suction side cooling slots. The suction side cooling slots receive cooling flow from the pressure side cooling slots through a boundary layer bleed valve, which is also aligned with or rearward of the pressure side lip. The cooling system may also combine double impingement cooling with these features. The cooling system minimizes shear mixing, reduces hot spots and can reduce the trailing edge thickness, resulting in more efficient stage performance and extended operational life.

A cooling system for a turbine airfoil of a turbine engine has a multiple suction side cooling slots extending from a front edge on the suction side to the center of the trailing edge or even to the pressure side of the center line and a pressure side cooling slot curving to a pressure side outlet forward of the trailing edge and having a front pressure side lip that is aligned with or forward of the front edge of the suction side cooling slots. The suction side cooling slots receive cooling flow from the pressure side cooling slots through a boundary layer bleed valve, which is also aligned with or rearward of the pressure side lip. The cooling system may also combine double impingement cooling with these features. The cooling system minimizes shear mixing, reduces hot spots and can reduce the trailing edge thickness, resulting in more efficient stage performance and extended operational life.

A turbine engine component having improved cooling is provided. The turbine engine component includes an airfoil portion having a leading edge, a trailing edge, a pressure side, a suction side, a root, and a tip, and at least one cooling circuit in a wall of the airfoil portion. The at least one cooling circuit has at least one passageway extending between the root and the tip. The at least one passageway has an aspect ratio of no greater than 2:1, and preferably substantially unity.

Provided is a gas turbine plant that enables active clearance control for ensuring tip clearance of first-stage turbine rotor blades required during start-up and for achieving the minimum tip clearance during load operation. In a gas turbine plant including a cooler in an air system used for cooling second-stage turbine stator blades, a first-stage segmented ring and a second-stage segmented ring that oppose tips of first-stage turbine rotor blades and second-stage turbine rotor blades are supported by the same blade ring member, and a cooling-air for the second-stage turbine stator blades forms a cooling air flow cooling the blade ring, to control thermal expansion of the blade ring and to control the clearance with respect to the tips.

A method for hybrid wafer-to-wafer bonding, comprising: providing two silicon wafers with Cu pattern structures, a conventional Cu BEOL process is adopted on the silicon wafers to obtain the planarized surface with copper and dielectric; removing part of the Cu on the planarized surface of the Cu pattern structures by adopting an etching process to form a certain amount of Cu recesses; depositing a layer of bonding metal on the surface of the Cu by adopting a selective deposition process; performing surface activation on the bonding metal and the dielectric by adopting a surface activation process; aligning and pressing the two silicon wafers together to obtain the dielectric bonding; and obtaining the metal bonding through the annealing process. The sufficient metal bonding can be obtained at low annealing temperature according to the present invent, thereby the risk of dielectric delaminating caused by thermal expansion mismatch is reduced, which is conducive to reduce the difficulty of process integration, save process time and improve product yield.

The invention relates to a turbine rotor blade of a gas turbine. The turbine rotor blade of the gas turbine comprises a blade profile, a blade platform, a blade top and a blade root, wherein a coolingair channel loop is arranged in the blade, a cooling air channel and a cooling air channel inlet are arranged in the loop, and ribs are arranged in the cooling air channel; and the blade surface is divided into a blade leading edge and a blade trail edge by a pressure surface and a suction surface. The blade tail edge is provided with a tail edge exhaust slot, a tail edge channel is provided witha guide groove structure, and at least one turbulence column is arranged in the guide groove structure; and the blade root cooling air channel is divided into at least two channels by a partition plate, and the partition plate is gradually bent towards the front edge side of the blade in the airflow direction. According to the turbine rotor blade of the gas turbine, the heat exchange effect of aleading edge channel, the tail edge exhaust slot and cooling air in the blade can be improved, and the leading edge metal temperature close to the blade platform and the metal temperature of the tailedge, the blade platform close to the tail edge and the blade root can be reduced.

Method for assembling into turbine sets turbine components having internal cooling passages. The turbine components are classified according to flow capability through the internal cooling passages and turbine sets are assembled in which the turbine components have the same flow capability classification.

A method for hybrid wafer-to-wafer bonding, comprising: providing two silicon wafers with Cu pattern structures, a conventional Cu BEOL process is adopted on the silicon wafers to obtain the planarized surface with copper and dielectric; removing part of the Cu on the planarized surface of the Cu pattern structures by adopting an etching process to form a certain amount of Cu recesses; depositing a layer of bonding metal on the surface of the Cu by adopting a selective deposition process; performing surface activation on the bonding metal and the dielectric by adopting a surface activation process; aligning and pressing the two silicon wafers together to obtain the dielectric bonding; and obtaining the metal bonding through the annealing process. The sufficient metal bonding can be obtained at low annealing temperature according to the present invent, thereby the risk of dielectric delaminating caused by thermal expansion mismatch is reduced, which is conducive to reduce the difficulty of process integration, save process time and improve product yield.

A radiant superheater arranged to hang at the upper portion of a furnace of a boiler. The radiant superheater is substantially planar and includes a first vertical pass, a first connection pass, a second vertical pass, a third vertical pass, a second connection pass, and a fourth vertical pass. Each vertical pass includes an upper end and a lower end. The vertical passes are connected in series, so that steam to be superheated enters at the upper end of the first vertical pass and flows through the first vertical pass and from the lower end of the first vertical pass via the first connection pass to the lower end of the second vertical pass and through the second vertical pass and from the upper end of the second vertical pass to the upper end of the third vertical pass and through the third vertical pass and from the lower end of the third vertical pass via the second connection pass to the lower end of the fourth vertical pass and through the fourth vertical pass, to be discharged from the upper end of the fourth vertical pass. The first connection pass is arranged below the second connection pass so as to shield the second connection pass from radiation from the lower portion of the furnace.

An improved nozzle vane for a gas turbine engine, comprising a vane wall having inner and outer wall surfaces, the wall surfaces being spaced from one another to define a plurality of fluid passageways for a cooling medium; discreet cavities formed by interior wall members disposed between the inner and outer wall surfaces and within the fluid passageway for the cooling medium; a plurality of impingement cooling sleeves disposed in the discreet cavities defined by the inner and outer wall surfaces and by interior wall members; and a plurality of non-round, e.g., serrated, openings in each of the impingement cooling sleeves, with the openings being sufficient in size and number to accommodate the flow of a cooling media.

A method includes providing a furnace including a radiant heating zone having metal heating coils and burners, concurrently applying a combustion media, having a combustibility, and a diluent to the burners, the burners burning the combustion media producing flames heating the radiant heating zone, and the diluent reducing the combustibility of the combustion media for reducing heat generated by the flames for reducing heat degradation of the metal heating coils.

A process for the co-production of superheated steam and methane includes reacting a gas containing carbon monoxide and hydrogen in a series of methanation regions to produce a product gas containing methane The gas containing carbon monoxide and hydrogen flows, in passing through the series of methanation regions, successively through at least a first internally cooled methanation region and a second internally cooled methanation region, cooling the first and second internally cooled methanation region with water, which water is vaporized in the first internally cooled methanation region to produce steam and which steam is subsequently superheated in the second internally cooled methanation region to produce superheated steam The internally cooled methanation regions include tubes filled with a methanation catalyst and a space surrounding the tubes The gas containing carbon monoxide and hydrogen and / or product gas flows through the tubes and the water flows through the space surrounding the tubes.

A turbine rotor blade and a gas turbine including the same, wherein the turbine rotor blade includes a blade body, the blade body includes a cold air channel and an impact plate; the cold air channel is arranged inside the blade body; the impact plate is arranged inside the blade body; wherein the impact plate is provided with an impact hole, the impact hole is connected to the cold air channel, and is used to pass the cooling air in the cold air channel through the impact hole to perform impact-enhanced cooling on the top area of ​​the blade body. The turbine rotor blade of the present invention adopts a complex cold air channel structure inside, and impact-enhanced cooling is adopted in the cold air channel inside the blade body near the tip area, and the characteristics of high cooling air pressure margin and high heat transfer coefficient of impact cooling of the rotor blade are utilized to reduce the metal temperature in the tip area of ​​the rotor blade, thereby improving the service life of the turbine rotor blade.

The invention discloses a gas turbine combustor transition section with a cooling structure, and belongs to the technical field of a gas turbine. The outer surface of a transition section body is provided with a plurality of cooling components; each cooling component comprises a connecting piece and a heat dissipation part; the cooling components are distributed in an array, and every two adjacent cooling members are distributed in a staggered manner; each heat dissipation part is in the shape of a bowl or is spiral or dendritic or is in other shapes that not only can surface area be increased but also flow resistance is smaller. According to the gas turbine combustor transition section disclosed by the invention, the temperature of the transition section can be lowered and the service life of the transition section can be prolonged by increasing the heat exchange area of the outer surface of the transition section; meanwhile, the integrity of the transition section body cannot be broken, the reliability of a thermal barrier coating cannot be influenced and the thermal insulation effect of the can be helped to be enhanced by the cooling structure; pressure loss brought by the cooling structure is low, so that the efficiency of the gas turbine can be helped to be improved.

A receiver with a configuration of saturated and superheated steam solar modules in a tower solar concentration power plant. The configuration allows the incidence of radiation on both sides of the superheated steam module, providing significant benefits for its durability and global control of the power plant.

A method includes providing a furnace including a radiant heating zone having metal heating coils and burners, concurrently applying a combustion media, having a combustibility, and a diluent to the burners, the burners burning the combustion media producing flames heating the radiant heating zone, and the diluent reducing the combustibility of the combustion media for reducing heat generated by the flames for reducing heat degradation of the metal heating coils.

The invention relates to a turbine rotor blade of a gas turbine, which includes a blade profile, a blade platform, a blade top and a blade root. There is a cooling air passage loop inside the blade, and a cooling air passage and a cooling air passage inlet are installed in the cooling air passage. Ribbed; blade surface is divided into blade leading edge and blade trailing edge by pressure surface and suction surface. The trailing edge of the blade has a trailing edge exhaust slot, and the trailing edge channel has a diversion groove structure, and there is at least one spoiler column in the diversion groove structure; the cooling air passage of the blade root is divided into at least two passages by a partition plate, The partition plate is gradually bent toward the leading edge side of the blade along the airflow direction. The invention can improve the heat exchange effect between the internal leading edge channel, the trailing edge exhaust slot and the cooling air, and reduce the metal temperature of the leading edge near the blade platform and the metal temperature of the trailing edge and the platform and blade root near the trailing edge.

The present invention discloses a cooling structure for a turbine stator blade, including a turbine stator blade and a baffle; the turbine stator blade includes an upper edge plate, a lower edge plate and a blade profile; the baffle is a thin-walled sheet metal having the same shape as the front section of the turbine stator blade, and is wrapped around the upper edge plate, the lower edge plate and the front section of the blade profile to form a slit cavity. The upper edge plate, the lower edge plate and the blade profile are respectively provided with a plurality of ribs and air holes at certain intervals in their front sections. The present invention designs a thin-walled baffle structure to form a certain slit cavity with the turbine stator blade, and the cooling airflow fully exchanges heat inside the cavity, so that the metal temperature of the front section of the blade profile and the front section of the edge plate of the stator blade is reduced, and when the gas flows out through the slit, an efficient air film can be formed on the gas side wall surface of the stator blade, thereby protecting the gas side wall surface of the stator blade and improving the utilization efficiency of the cold air. At the same time, the front section of the stator blade profile, which is the most easily damaged in the traditional design, is transferred to the thin-walled baffle, thereby protecting the stator blade itself from damage.

A system and method for sealing a multilateral junction in an oil well is described. Molten material, such as a eutectic or non-eutectic alloy, flows into the junction area and forms a durable seal upon cooling. An expanding alloy may be utilized, which expands upon solidification and which has a melting temperature that is higher than the maximum anticipated well temperature. The alloy is placed within a cavity in the well and held at a temperature above the melting point of the alloy, after which the alloy is cooled down to the ambient well temperature and solidifies and expands within the cavity.