Provided is a combustor assembly. The combust assembly includes: a plurality of swirlers through which a first fluid that is a part of a fluid discharged from a compressor passes; a base portion, in which the plurality of swirlers are provided, comprising a first through hole formed between one swirler and another swirler from among the plurality of swirlers so that a second fluid that is another part of the fluid discharged from the compressor and different from the first fluid passes through the first through hole; and a deflector provided in the base portion so as to face the first through hole for changing a moving direction of the second fluid.

The disclosure describes a compressor housing for a turbocharger system. A scroll section of a collecting volute may be defined by a wall of the compressor housing. The scroll section may extend around the axis of the compressor from an inlet of the volute to a scroll discharge stage. A plate section of the compressor housing may extend from the wall in a radial direction outward from the axis and may extend in a circumferential direction around a substantial portion of the collecting volute. The plate section may be configured to shield heat and may advantageously be formed integral with the compressor housing.

A turbine casing for an exhaust-gas turbine is provided herein. The turbine casing has a connection flange for bearing-casing-side attachment to a bearing casing in which casing-connection holes that are mutually spaced in the circumferential direction are provided, wherein, between adjacent casing-connection holes, material recesses that are open radially inwards towards a central longitudinal axis of the turbine casing are provided in the connection flange. Further provided herein is an exhaust-gas turbine that is equipped with the turbine casing.

A method for upgrading a gas turbine, the method includes: a) removing all guide vanes of the first guide vane stage; b) replacing the removed guide vanes of the first guide vane stage with new or reconditioned guide vanes, wherein blade platforms of the new or reconditioned guide vanes are provided with cooling air bores which fluidically connect a cooling air supply duct to the annular gap and open into the annular gap, and wherein the cooling air bores are arranged in such a manner that more cooling air bores open into regions of an annular gap that are arranged radially inwards from leading edges of the guide vanes than in other regions of the annular gap.

A vane device for a gas turbine is provided. The vane device includes a first airfoil having a first suction side and a first pressure side, a second airfoil having a second suction side and a second pressure side, an inner shroud and an outer shroud. The first airfoil and the second airfoil are arranged between the inner shroud and the outer shroud, wherein the first airfoil and the second airfoil are at least partially coated with a MCrAlY coating. At least the first suction side has a first coated surface section which is coated with a thermal barrier coating and which represents at least a part of the total surface of the first suction side. At least the inner shroud or the outer shroud has a further coated surface section which is coated with a further thermal barrier coating. A corresponding method of manufacturing is also provided.

A vane arc segment includes an airfoil fairing that has a fairing platform and a hollow airfoil section that extends there from. The hollow airfoil section defines an airfoil profile and surrounds an internal cavity. The fairing platform defines a gaspath side and a non-gaspath side. The airfoil fairing is formed of a fiber-reinforced composite comprised of fiber plies. The fiber plies include at least one cavity fiber ply that is arranged as a tube that circumscribes the internal cavity. The at least one cavity fiber ply extends through the fairing platform and defines at least a portion of an upstanding collar on the non-gaspath side of the fairing platform. The upstanding collar defines a collar profile. The tube necks down through a neck portion such that at least a portion of the collar profile is narrower than the airfoil profile.

An apparatus and method of reducing operating temperatures of a gas turbine engine, exposed to a service tube assembly by utilizing a skirt and radially outer mount. The reduction in temperature exposure minimizes temperatures of the service tube during engine operation to reduce the incidence of oil coking or varnish.

A ceramic matrix composite article, method for forming the article, and perforated ply which may be incorporated therein are disclosed. The article includes at least one shell ply forming an exterior wall having first and second portions and defining a plenum. An annular brace formed of at least one structural support ply is disposed within the plenum, including a first integral portion integral with and part of the first portion of the exterior wall, a first curved portion extending from the first integral portion and curving across the article plenum to the second portion of the exterior wall, a second integral portion integral with and part of the second portion of the exterior wall, a second curved portion extending from the second integral portion and curving across the article plenum to the first curved portion, and an overlap in which the first and second curved portions are integral.

A heat shield panel for a gas turbine engine combustor is disclosed. The heat shield panel includes a hot side defining a first surface having an outer perimeter, a cold side defining a second surface spaced from the first surface and a plurality of holes, each hole including a central axis having vector components defined by a common vector.

An airfoil includes an airfoil section that has an airfoil wall that define leading and trailing ends and first and second sides that join the leading and trailing ends. The first and second sides span in a longitudinal direction between first and second ends. The airfoil wall circumscribes an internal core cavity. The airfoil section extends form a platform. The platform defines a shelf that extends forward from the first end at the leading end of the airfoil section to a platform leading edge, aft from the first end at the trailing end of the airfoil section to a platform trailing edge, laterally from the first end at the first side of the airfoil section to a first platform side edge, and laterally from the first end at the second side of the airfoil section to a second platform side edge. The internal core cavity extends from the airfoil section into the shelf.