A sealing device includes a sealing main body surrounding a rotor shaft rotating about an axis from an outer periphery side. The sealing main body includes an inside surface facing an outside surface of the rotor shaft with clearance and a plurality of hole parts formed recessed to an outer side in a radial direction from the inside surface and arrayed in a circumferential direction of the inside surface and in a direction of the axis. The hole parts have an inner peripheral wall in which a spiral groove extending twisting toward the outer side in a radial direction is formed.

A scoop inlet for a turbomachine. The turbomachine includes a core engine casing relative to a centerline extending the length of the turbomachine. Further, the core engine casing at least partially defines a bypass flow passage. The scoop inlet includes an inlet at the core engine casing in fluid communication with the bypass flow passage. The inlet includes a mouth fluidly coupling the bypass flow passage to a cooling system. The mouth receives bypass bleed air from the bypass flow passage and defines a width in a circumferential direction relative to the centerline. The scoop inlet also includes a bypass bleed duct defined through the core engine casing. The bypass bleed duct fluidly couples the mouth to the cooling system. In addition, the scoop inlet includes a plurality of droplets at the mouth of the inlet partially extending from the mouth into the bypass flow passage.

A strut for a gas turbine engine includes an airfoil section extending in a spanwise direction between a first platform and a second platform, extending in a chordwise direction between a leading edge and trailing edge to define a chord length, and extending in a thickness direction between a first side and a second side to define a chord width. Exterior surfaces of the airfoil section define a leading portion between the leading edge and a widest location of the airfoil section relative to the thickness direction, and a trailing portion between the widest location and the trailing edge. The exterior surfaces establish a respective exterior contour for each span position between a 0% span position and a 100% span position. The exterior surfaces define a plurality of dimples in the leading portion.

The present invention relates to a seal arrangement for a turbomachine, in particular a gas turbine, having a plurality of rows, arranged in succession in the axial direction (A), of shells (1-3) connected to one another in the circumferential direction (U), wherein shells adjacent in the axial direction have cross sections opened counter to a throughflow direction (A) and/or a thread axis inclined counter to the throughflow direction.

A gas turbine hot part includes: a body portion; a porous portion forming at least a part of the body portion or disposed on at least a part of the body portion and allowing a cooling gas to pass therethrough; and at least one filter disposed upstream of the porous portion in a flow direction of the cooling gas and capable of trapping foreign substances that cannot pass through the porous portion.

A gas turbine engine component that includes a structure having a surface which includes multiple cooling channels having a width of 20-30 m and a depth of 25-50 m.

An airfoil for a gas turbine engine includes a first side wall; a second side wall joined to the first side wall at a leading edge and a trailing edge; and an internal cooling system arranged within the first and second side walls configured to direct cooling air through and out of the airfoil. The internal cooling system has a first cooling circuit that includes an acceleration channel generally extending in a radial outward direction. A first section of the acceleration channel decreases in cross-sectional area along the radial outward direction such that the cooling air is accelerated through the first section of the acceleration channel. The first cooling circuit further includes a trailing edge chamber fluidly coupled to receive at least a portion of the cooling air from the acceleration channel and extending generally in a chordwise aft direction from the acceleration channel to the trailing edge.

A turbine nozzle disposed in a turbine includes a suction side extending between a leading edge of the nozzle and a trailing edge of the turbine nozzle in an axial direction and transverse to a longitudinal axis of the turbine nozzle, and extending a height of the nozzle in a radial direction along the longitudinal axis, a pressure side disposed opposite the suction side and extending between the leading edge of the turbine nozzle and the trailing edge of the turbine nozzle in the axial direction, and extending the height of the nozzle in the radial direction, and a bulge disposed on the suction side of the nozzle protruding relative to the other portion of the suction side in a direction transverse to a both the radial and axial directions.

A gas turbine engine component includes a passage and a geometrically segmented coating section adjacent the passage. The geometrically segmented coating section includes a wall that has a first side bordering the passage and a second side opposite the first side. The second side includes an array of cells, and there is a coating disposed over the array of cells. The coating defines an exterior side. A cooling passage extends through the wall and the coating. The cooling passage fluidly connects the passage and the exterior side.

A cylindrically-shaped hybrid rocket engine solid fuel grain defines an axial combustion port. A fuel grain material comprises a compounded blend of thermoplastic fuel and aluminum. The fuel grain comprises fused stack layers, each layer comprising a plurality of fused abutting concentric beaded structures arrayed to define the combustion port; the port exhibits a rifling pattern or rifling inducing geometry along the port wall. When an oxidizer is introduced into the combustion port combustion occurs along the exposed port wall. Each beaded structure defines a geometry that increases the combustion surface area while inducing a vortex flow of oxidizer and fuel gas. As each layer ablates, an abutting layer exhibiting a similar geometry, is revealed, undergoes a gas phase change, and ablates. This process repeats and persists until oxidizer flow is terminated or the fuel grain material is exhausted. The fuel grain may be manufactured by an additive manufacturing process.