A method for producing a tubular body with reduced internal stress uses 3D printing. The tubular body has an outer wall with a stiffening structure extending along at least part of the outer wall. The method sets a printing plane for 3D printing with a 3D printer, and prints a tubular body layer in the printing plane with the 3D printer. The tubular body layer has an outer wall layer and a stiffening structure layer extending in the printing plane along a periphery of the outer wall layer. The stiffening structure layer has at least two portions spaced apart from one another. The method produces an outer wall with a stiffening structure for a tubular body with reduced internal stress.

A turbomachinery stator apparatus includes: a compressor casing including a casing wall defining an arcuate flowpath surface and an opposed backside surface, the flowpath surface defining at least two spaced-apart rotor lands, a stator vane row of stator vanes disposed inside the compressor casing, wherein the casing wall includes a heat shield positioned outboard of the rotor lands immediately upstream or downstream of the stator vane row, and wherein a) the casing wall includes the heat shield and b) the stator vanes form a single monolithic whole.

A turbomachine includes a shroud and a hub spaced apart from the shroud to channel an airflow along a direction. The turbomachine includes a plurality of airfoils coupled between the shroud and the hub. At least one airfoil of the plurality of airfoils includes a leading edge spaced apart from a trailing edge in the direction of the airflow and a pressure side opposite a suction side. The turbomachine includes at least one local vortex generator array defined on the suction side so as to extend onto the hub or the shroud. The at least one local vortex generator array is defined downstream of the leading edge.

A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

Methods for creating an aircraft turbomachine vane using additive manufacturing include additively manufacturing a vane on a bed of powder using selective laser melting, the additive manufacturing being performed on a support plate so that first or second circumferential edges are manufactured first directly on the support plate, at least one temporary support member being produced simultaneously with the first or second edges. The methods also include removing the temporary support member by breaking its connection with the leading or trailing edge with a tool that is engaged in at least one recess thereof.

There is provided a cobalt-based alloy product comprising: in mass %, 0.08-0.25% C; 0.1% or less B; 10-30% Cr; 5% or less Fe and 30% or less Ni, the total amount of Fe and Ni being 30% or less; W and/or Mo, the total amount of W and Mo being 5-12%; at least one of Ti, Zr, Hf, V, Nb and Ta, the total amount of Ti, Zr, Hf, V, Nb and Ta being 0.5-2%; 0.5% or less Si; 0.5% or less Mn; 0.003-0.04% N; and the balance being Co and impurities. The product is a polycrystalline body of matrix phase crystal grains. In the matrix phase crystal grains, post-segregation cells with an average size of 0.13-2 μm are formed, wherein components constituting an MC type carbide phase comprising Ti, Zr, Hf, V, Nb and/or Ta are segregated along boundary regions of the post-segregation cells.

An assembly is provided for a gas turbine engine. This gas turbine engine assembly includes a split case structure. The split case structure includes a first wall, a second wall, a first case segment and a second case segment. The first wall extends axially along and circumferentially about an axial centerline. The second wall extends axially along and circumferentially about the axial centerline. The second wall is radially outboard of and axially overlaps the first wall. The first case segment is configured to form a first portion of the first wall and a first portion of the second wall. The second case segment is configured to form a second portion of the first wall and a second portion of the second wall. The second case segment is circumferentially adjacent and attached to the first case segment at a joint.

A compression ring for a shrouded compressor including a radially inner surface having one or more areas configured to mate flush with one or more portions of a radially outward surface of a shroud of the shrouded compressor, a radially outer surface located opposite the radially inner surface, a labyrinth seal located on the radially outer surface, a groove located within the radially inner surface, and a load ring located within the groove.

A component with a region to be cooled having a cooling channel which is arranged and designed so as to cool the region of the component during operation by a fluid flow, wherein the cooling channel is defined by a first channel side facing the region and by a second channel side facing away from the region. The first channel side forms a larger contact surface for the cooling channel than the second channel side. An additive manufacture process can produce the component.

A method for forming a component for a gas turbine engine may include forming a first portion of the component that includes a cast metal or metal alloy, forming a second portion of the component that includes presintered preform defining at least one support structure, positioning the second portion on the first portion to define an assembly such that the first portion and the second portion define at least one cooling channel therebetween, and heating the assembly to join the first portion and the second portion and form the component.