A method for digitally designing a support with the shape of a fibrous blank obtained by three-dimensional weaving intended to form a fibrous preform of a turbine engine blade or propeller after shaping and compaction in a mold, includes providing a set of points representative of a face of the fibrous blank, the face being intended to form the root of the blade or the propeller and a portion of an aerodynamic profile of the blade or the propeller, generating a web connecting the points of the set of points, and digitally designing the support including at least an imprint of the fibrous blank having the shape of the generated web.

A fiber-reinforced component for use in a gas turbine engine includes a first braided fiber sleeve forming a cooling channel and a plurality of fiber plies enclosing the first braided fiber sleeve, with the plurality of fiber plies forming first and second walls separated by the first braided fiber sleeve. The fiber-reinforced component further includes a matrix material between fibers of the braided fiber sleeve and the plurality of fiber plies.

A combustion chamber suitable in particular for use in a rocket engine comprises a combustion space, a first wall enclosing the combustion space and cooling duct fins, which extend from a surface of the first wall and separate adjacent cooling ducts from one another. At least one of the cooling duct fins has at its end facing away from the surface of the first wall a bent section, which at least partially covers a cooling duct adjacent to the cooling duct fin.

A method for manufacturing a blade in composite material having an added metal leading edge for gas turbine aeroengine, includes producing a blade body in composite material including in longitudinal direction, a blade root part, a shank part and an airfoil body part; manufacturing, via additive manufacturing, a leading edge extending in longitudinal direction between a lower end present at the shank part in composite material and an upper end present at the tip; bonding the manufactured leading edge onto the foremost edge portion of the airfoil body of the blade body in composite material.

A centrifugal fan frame structure includes a base plate and a plurality of side wall sections. The side wall sections are discontinuously formed along an outer periphery of the base plate through injection molding, such that a clearance is formed between any two adjacent side wall sections, and portions of the outer periphery of the base plate located corresponding to the clearances are not covered by any injection molded side wall section. And, the side wall sections and the base plate together define a receiving space between them. With the injection molded side wall sections being spaced along the outer periphery of the base plate, a wrapping force from the injection molded side wall sections is effectively reduced to prevent the base plate from structural deformation under the wrapping force.

Methods and masks for manufacturing component of gas turbine engines are described. The methods include applying a mask to a protected surface of the component, the component having a designated surface to be treated by a shot peen operation. The mask includes a full masking portion configured to prevent a shot peen media from impacting the protected surface. A masking control region is arranged around the designated surface. The masking control region is configured to control an amount of force imparted to the component by shot peen media during the shot peen operation, wherein the masking control region extends from the full masking portion to the designated surface. The designated surface is shot peened with shot peen media to form a compressive stress region within the component proximate the designated surface and a tapering transition of compressive forces within the component proximate the masking control region.

A fan includes a hub configured to be driven by motor to rotate about a fan rotational axis, blades that protrude radially from the hub and a band that surrounds the rotational axis and connects the tips of the blades. The band includes features that structurally reinforce the knit lines. In some embodiments, the features are thickened regions that protrude from outer surface of the band in a direction away from the hub. A thickened region is disposed between respective tips of each pair of adjacent blades. In some embodiments, the features are structurally—reinforcing ribs that protrude from the hub-facing surface of the band. A rib is disposed between respective tips of each pair of adjacent blades. In some embodiments, the features are both ribs and thickened regions.

A method of manufacturing a component for a gas turbine engine includes applying a thermoplastic polymer sheet over a composite body for the component; applying a shield over part of the composite body, the shield terminating at an end which overlies the thermoplastic polymer sheet and defines an interface between shielded and unshielded regions of the component; and pressing the shield into the thermoplastic polymer sheet so that the thermoplastic polymer sheet deforms around the end of the shield, such that the exterior profile of the component at the interface between the shielded and unshielded regions is flush.

A method for manufacturing a blade with a first portion and a second portion, the method includes forming the first portion that includes forming a model of the first portion from removable material, forming a first shell mould from the model of the first portion, and forming the single-crystal or columnar first portion m a first metal alloy in the first shell mould from a single-crystal seed, and forming the second portion in which the second portion is formed on the first portion, and in which the first portion and the second portion are made from different materials, the second portion being polycrystalline and formed from a second metal alloy. The blade includes a single-crystal or columnar first portion made from a first metal alloy and a polycrystalline second portion made from the second metal alloy different from the first metal alloy.

A turbine blade for a gas turbine, having a blade root and an aerodynamically curved blade airfoil arranged above the blade root. The blade airfoil has a pressure-side and a suction-side blade wall, together extending from a leading edge, that can receive a flow of working medium, to a trailing edge. A multiplicity of cooling air outlet openings are formed on the pressure-side blade wall, which extend upstream from the trailing edge with respect to the flow direction, and through these openings cooling air that is conveyed through the interior of the blade airfoil can exit. At least one of the cooling air outlet openings has a substantially rectangular or trapezoidal shape with rounded corners. At least the lower corner, pointing towards the leading edge, of the cooling air outlet opening forms a relief notch, which projects outwardly from the rectangular shape, with a rounded notch bottom.