A gas turbine engine article includes a substrate that has a pre-bond surface that includes a serpentine groove. A coating is disposed on the pre-bond surface and mechanically interlocks with the serpentine groove.

A method for manufacturing a structural surface heat exchanger of preset or left-hand final shape for an aircraft includes the steps of forming, shaping and assembling, by welding or brazing, a first corrugated skin and a second smooth skin in order to obtain channels. Each channel is delimited by a corrugation of the first skin and the second smooth skin so as to form a structural surface heat exchanger of preset or left-hand final shape, wherein a fluid is configured to circulate in the channels and air is configured to circulate in contact with the second smooth skin.

A turbine blade, a turbine, and a gas turbine which have enhanced cooling performance are provided. The turbine blade may include: an airfoil having a blade shape and including a suction side formed in a convex shape and a pressure side formed in a concave shape; a platform coupled to a lower portion of the airfoil; and a root member protruding downward from the platform and coupled to a rotor disk, wherein the platform may include: a plurality of inlets through which air is drawn into the platform; a plurality of outlets through which the air is discharged from the platform; and a cooling passage connecting the inlets to the outlets and including a plurality of dispersion spaces, a width in each dispersion spaces increasing from the inlet to the outlet, and a narrow space formed between the dispersion spaces and having a width less than the width of each dispersion spaces.

A fixed vane turbocharger includes: an impeller; a housing including, inside thereof, an impeller housing space which accommodates the impeller, a scroll flow passage formed on a radially outer side of the impeller, and a communication flow passage which brings the impeller housing space and the scroll flow passage into communication; and at least one fixed vane unit disposed in the communication flow passage and fixed to a portion of the housing on an inner side of the scroll flow passage with respect to a radial direction of the impeller. Each of the at least one fixed vane unit includes at least two vane portions and a coupling portion coupling the two vane portions, and is formed of a single sheet metal member.

A first-stage turbine vane supporting structure reduces chordal leakage between a first-stage turbine vane segment and its supporting ring. The structure includes a first-stage turbine vane segment including an outer platform, an inner platform, and a first-stage turbine vane disposed between the inner and outer platforms; an inner rail protruding from the inner platform in a longitudinal direction of the first-stage turbine vane; a supporting ring for supporting the first-stage turbine vane segment by engaging with the inner rail while facing one surface of the inner rail; a supporting member engaging with the supporting ring to define a U-shaped space bordered by three side surfaces surrounding the inner rail; and a flexible member interposed between the inner rail and one of the three side surfaces of the U-shaped space. A protrusion formed on the inner rail engages with a sealing surface of the supporting ring in an airtight manner.

An engine pod for a gas turbine engine which includes a pod wall having an inside and an outside. The pod wall includes a fixed downstream portion and a displaceable upstream portion which is displaceable in the axial direction between a first upstream position and a second downstream position. At its downstream end facing the fixed portion, the upstream portion forms a radially outer rear edge and axially spaced therefrom a radially inner rear edge, with a recess in between. It is provided that adjacent to the recess, an air-permeable structure is formed in the upstream portion which is intended and configured, in the first upstream position of the displaceable portion, to conduct air flowing in the region of the recess to the inside of the displaceable portion. According to a further aspect of the invention, the axial position of the radially inner rear edge varies in the circumferential direction.

The invention relates to a three-dimensional screen printing method for producing a green part from printing material for a powder metallurgical component, wherein the printing material contains a fraction of powder, more particularly metal powder or ceramic powder, and binder or consists of these materials, characterized in that a screen printing mask has a screen printing structure having openings for pressing the printing material through, the openings being partly undulate so that the green part at least partly has a three-dimensional undulate structure and/or undulate edges.

Airfoils for gas turbine engines are describe. The airfoils include a leading edge having an interior surface with an inflection point line extending radially between a root and a tip of the airfoil. The inflection point line is defined at a location of minimum radii that separates a pressure side and a suction side of the airfoil body. An interlaced trip strip array is arranged along the leading edge and includes a chevron trip strip having an apex and ligaments extending from the apex to form a chevron shape and a skew trip strip arranged proximate to the chevron trip strip with a leading end proximate the inflection point line. The skew trip strip is positioned adjacent to the chevron trip strip such that a gap is formed between the skew trip strip and one of the ligaments of the chevron trip strip.

Methods, apparatus, systems and articles of manufacture are disclosed. A shroud assembly of a gas turbine engine includes: a first shroud arm having a first end and a second end, the first end to couple to an outer wall and the second end to couple to a first shroud pad, and a second shroud arm having a first end and a second end, the first end to couple to the outer wall and the second end to couple to a second shroud pad, at least one of the first shroud pad or the second shroud pad to move radially outward toward the outer wall in response to a rotor blade contacting the at least one of the first shroud pad or the second shroud pad.

Provided is a method for sealing an annular gap in a turbine with a housing and a rotor, in which method a) an annular recess, which is arranged coaxially with respect to the rotor and opposite the tips of the rotor blades of a stage, is provided in the housing or in at least one element attached to the housing, b) a subdivided insert ring, which is formed such that it can be inserted into the recess in a form-fitting manner with play, is provided, c) the insert ring is inserted into the recess, more particularly in that the insert ring is pushed into the recess at least substantially in the axial direction, and d) the insert ring is fixed using securing elements in such a way that an at least substantially axially directed clamping force is applied by means of the securing elements. A turbine is also provided.