Abstract: A structure for assembling turbine blade seals, a gas turbine including the same, and a method of assembling turbine blade seals are provided. The structure for assembling turbine blade seals includes a turbine blade including an airfoil, a platform, and a root, a turbine rotor disk to which the root of the turbine blade is mounted, a seal plate mounted between the platform and one side of the turbine rotor disk to seal a cooling channel defined within the root and the platform, an insertion pin inserted through the turbine rotor disk and the seal plate to fix the seal plate to the turbine rotor disk, and a retainer configured to fix the insertion pin and to prevent the insertion pin from falling out.

Abstract: A tie rod assembly structure, a gas turbine having the same, and a method of assembling tie rod are provided. The tie rod assembly structure includes a tie rod on which a plurality of rotor disks are mounted, a bearing support shaft mounted to the tie rod to support the rotor disks and on which a bearing is mounted, a first nut mounted on the tie rod on one side of the bearing support shaft, and a second nut that is fastened to the first nut to tension the tie rod and then is in close contact with the bearing support shaft to support the bearing support shaft.

Abstract: An airfoil of either of a turbine blade or a turbine vane includes a cooling passage; at least one disk body disposed on an inner wall of the cooling passage and configured to reduce a flow cross-sectional area of the cooling passage to increase a fluid pressure of cooling fluid flowing through the cooling passage; and at least one through-hole formed in each of the at least one disk body such that the cooling fluid flows through the at least one through-hole and forms a vortex on a downstream side of the at least one through-hole. The cooling passage includes an inlet supplied with the cooling fluid and an end opposite to the inlet, and the at least one disk body is disposed at the inlet of the cooling passage and is configured to increase the fluid pressure of the cooling fluid flowing into the cooling passage.

Abstract: Impingement insert for an airfoil of a blade/vane of a gas turbine is provided. The impingement insert includes a double-walled section having an outer and an inner walls, that define—an inner channel at an inner surface of the inner wall, an outer channel at an outer surface of the outer wall and a middle channel between the outer and the inner walls. Impingement cooling holes are provided in the outer wall that use the cooling air of the middle channel to eject impingement jets into the outer channel. The impingement insert includes at least one extraction duct that extends between the outer and the inner walls across the middle channel, and has an inlet at the outer channel, and an outlet at the inner channel, for flowing the cooling air, after impingement, from the outer channel into the inner channel.

Abstract: An impingement cooling structure is provided. The impingement cooling structure includes a flow channel formed between a first wall and a second wall facing the first wall, a plurality of impingement cooling holes disposed in the first wall such that the plurality of impingement cooling holes are spaced apart from each other along the flow channel, and a flow diverter convexly protruding from a surface of the second wall in each space between injection axes of the plurality of impingement cooling holes.

Abstract: A reverse osmosis apparatus for a seawater desalination system is provided. The reverse osmosis apparatus includes a barrel in which a plurality of vessels receiving reverse osmosis membrane units are arranged, a feed tank provided in an intermediate portion of the barrel and connected to a seawater inlet, a first water tank provided inside a first end portion of the barrel and connected to a plurality of first vessels connected to a first side of the feed tank, and a second water tank provided inside a second end portion of the barrel and connected to a plurality of second vessels connected to a second side of the feed tank.

Abstract: A combustor includes a liner having an outlet end to pass combustion gas and a liner flange protruding outward from the outlet end; a transition piece to discharge combustion gas from the liner to a turbine, the transition piece having an inlet end for coupling to the outlet end of the liner and a transition piece flange protruding outward from the inlet end to face the liner flange; and a first elastic support installed on the liner flange to protrude toward the transition piece flange. A force applied from the transition piece elastically deforms an elastic arch of the first elastic support, which includes a movable support that is spaced apart from the liner flange if the force applied from the transition piece does not primarily deform the elastic arch. An auxiliary elastic support installed inside the first elastic support elastically deforms if the force secondarily deforms the elastic arch.

Abstract: A combustor nozzle and a combustor for a gas turbine including the same are provided. The combustor nozzle may include a plurality of vanes disposed radially on an outer circumferential surface thereof, each vane including an internal cavity to which fuel is supplied, wherein the vane may include an airfoil in cross-section, wherein at least one of a pressure surface and/or a suction surface of the airfoil is disposed in a height direction thereof with a plurality of nozzle holes communicating with the cavity, wherein the vane may include a V-shaped nozzle hole including a pair of outlets on both sides of the airfoil based on a leading edge of the airfoil, wherein the V-shaped nozzle hole may form an angle (?) diverging outward with respect to a horizontal plane across the leading edge, and an acute angle (?) radially outward with respect to a vertical plane perpendicular to the horizontal plane.

Abstract: A combustor includes an inner liner forming a combustion chamber; an outer liner surrounding the inner liner to form a cooling passage in which compressed air flows; and a plurality of cooling guides installed around an inner circumferential surface of the outer liner to surround the combustion chamber, each of the cooling guides protruding from the inner circumferential surface to create an impinging jet from the compressed air flowing in the cooling passage. The plurality of cooling guides surrounding the combustion chamber are installed at regular intervals in a flow direction of the compressed air, and are arranged in staggered axial rows. Each cooling guide includes an air guiding surface facing the flow of the compressed air to guide the compressed air toward the inner liner. Accordingly, liner cooling efficiency can be enhanced by more effectively guiding the impinging jet toward the inner liner.

Abstract: A gas turbine nozzle assembly of a gas turbine is provided. The turbine nozzle assembly may include a turbine nozzle extending from an inner diameter to an outer diameter and having an airfoil-shaped cross section having a leading edge and a trailing edge, and a pressure side and a suction side each of which extends from the leading edge to the trailing edge, wherein the turbine nozzle may include a hollow airfoil including a plurality of cavities positioned in the airfoil, an insert positioned in one or more of the plurality of cavities of the hollow airfoil, a plurality of cover plates, at least one of which is positioned at one of the inner diameter and at the outer diameter, and a plurality of impingement pans, at least one of which is positioned at one of the inner diameter and at the outer diameter.

Abstract: A compressor cleaning apparatus to clean a compressor of a gas turbine is provided. The compressor cleaning apparatus includes a nozzle configured to inject a cleaning fluid into an interior of a compressor, a fluid supply tube connected to the nozzle to supply the cleaning fluid to the nozzle, a first cleaning fluid supply connected to the fluid supply tube to supply a first cleaning fluid, and a second cleaning fluid supply connected to the fluid supply tube to supply a second cleaning fluid having a temperature higher than that of the first cleaning fluid.

Abstract: A liner cooling device for cooling a liner of a gas turbine is provided. The liner cooling device may include a support portion disposed between a liner and a transition piece of a gas turbine and configured to include a cooling flow passage through which cooling air moves to the transition piece. The support portion includes a support member disposed between the liner and the transition piece and an auxiliary support member disposed in the cooling flow passage and having a hole through which the cooling air passes.

Abstract: A method and apparatus for automatically generating periodic boiler combustion models and aperiodic boiler combustion models through automatic learning are provided. The method of automatically generating a boiler combustion model may include determining whether a specific event has occurred in association with a boiler, changing a training condition according to a result of the determining, generating a boiler combustion model trained on operation data measured in the boiler and stored in a database according to the training condition, and determining a precision of the generated boiler combustion model.

Abstract: Disclosed herein are a ring segment having an air pouch and a first cooling hole formed therein, and a turbine including the same. The air pouch and the first cooling hole are formed in a shield wall, thereby achieving an improvement in cooling performance as well as simplification of production process.

Abstract: A nozzle for a combustor in which a fuel containing hydrogen is burned is provided. The nozzle includes a first tube disposed in a center of the nozzle and having a first diameter, a plurality of second tubes circumferentially disposed around the first tube to be spaced apart from the first tube and each having a second diameter smaller than that of the first tube, and a plurality of third tubes disposed around the first tube and each having a diameter smaller than the second diameter, wherein the first tube, the plurality of second tubes, and the plurality of third tubes are arranged in parallel with each other.

Abstract: An impingement insert for an airfoil of a turbomachine component is provided. The insert includes first and second body parts, each having inner and outer surfaces; and first and second contact parts provided on the outer surfaces of the first and the second body parts. The insert includes a flexible mechanical seal part between the body parts. A flow channel for cooling air is defined by the seal part and the inner surfaces of the body parts. One or both of the body parts include impingement holes. The insert has an elastic part connected to the body parts. When the elastic part is subjected to deformation, the elastic part is configured to apply a force, on the first and/or the second body parts, in a direction of increasing a separation between the first contact part and the second contact part.

Abstract: Disclosed herein is an electrostatic precipitation method using an electrostatic precipitator including a collection module having a collection electrode and a discharge electrode, a housing having an internal partition wall formed therein, an inlet-side passage switching member, and an outlet-side passage switching member. The electrostatic precipitation method includes collecting dust by applying a voltage to the discharge electrode while gas flows, closing some of the flow spaces, divided by the internal partition wall using the passage switching members, and performing dust collection for one of the opened flow spaces by applying a voltage to the discharge electrode therein, and performing washing for at least one of the closed flow spaces by supplying washing water to the collection electrode therein.

Abstract: The present invention provides an airfoil 110 with the squealer tip cooling system 50 for a turbine blade 100 at the blade tip 113, wherein the squealer tip cooling system 50 comprises a cooling passage 170 arranged within a squealer tip 117, wherein the cooling passage 170 at least partly extends toward a terminal end 74 of the squealer tip 117, and a pocket 172 at a lateral surface 75, 76 of the squealer tip 117, open externally and extending inwardly at least partly across the cooling passage 170. The pocket 172 intersects the cooling passage 170 and the pocket 172 comprises an impingement surface 70 facing the cooling passage 170, on which a cooling medium expelled through the cooling passage 170 impinges before being discharged externally through the pocket 172.

Abstract: Disclosed are a hybrid power generation facility and a control method thereof.

Abstract: A rotor and a turbo machine including the same are provided. The rotor includes a disk including a disk slot, a blade including a root member inserted into the disk slot and an airfoil disposed outside the root member in a radial direction of the disk, a locking sheet disposed inside the root member in the radial direction, and a fixing portion disposed inside the root member in the radial direction and configured to fix the locking sheet to the root member.