The present invention includes an apparatus for protecting an aerodynamic surface from erosion including a screen capable of being applied to a leading edge of the aerodynamic surface; and an erosion protection coating applied to the screen before or after the screen is applied to the leading edge, wherein the erosion protection coating protects the aerodynamic surface from erosion.

A nozzle assembly is provided and includes an outer cowl, an inner nozzle sleeve disposable within the outer cowl, nested support rings by which the outer cowl and the inner nozzle sleeve are coupled, sets of rivets to respectively connect the nested support rings to the inner nozzle sleeve, the nested support rings together and the nested support rings to the outer cowl and an elongate washer configured to reinforce three or more rivets of one or more of the sets of rivets.

A vane stage includes an arcuate platform defining an axial centerline axis having a pair of flanges that extend radially inward from the platform. The flanges are axially spaced from one another and from respective forward and aft ends of the platform. The vane stage includes a vane extending radially outward from the platform and a seal carrier mounted to the flanges of the platform. A method for constructing a vane stage includes sliding a seal carrier between flanges of an arcuate platform. Each flange includes at least a pair of through holes and interfaces with a respective axial side of the seal carrier. The method includes drilling through holes in each axial side of the seal carrier by using the through holes of each flange as guides.

An airfoil for a gas turbine engine according to an example of the present disclosure includes, among other things, an airfoil section extending from a root section. The airfoil section includes a sheath that receives a core. The core includes first and second ligaments received in respective internal channels defined by the sheath, and each one of the first and second ligaments includes at least one interface portion in the root section dimensioned to receive a retention pin.

A vane guide assembly for a gas turbine engine, the vane guide assembly including: an airfoil having an end bonded to an opening of a platform by an adhesive; and a pull tab partially located in the adhesive and having a portion extending from a bondline formed by the adhesive.

A fan for a gas turbine engine includes: an annular casing; a disk disposed inside the casing and mounted for rotation about an axial centerline, the disk including a row of fan blades extending radially outwardly therefrom; each of the fan blades including an airfoil having circumferentially opposite pressure and suction sides extending radially in span from a root to a tip, and extending axially in chord between spaced-apart leading and trailing edges, with the airfoils defining corresponding flow passages therebetween for pressurizing air; the row including no more than 21 and no less than 13 of the fan blades; and wherein each of the fan blades has a solidity defined by a ratio of the airfoil chord over a circumferential pitch of the fan blades, measured at 60% of a radial distance from the root to the tip, of less than about 1.6.

An acoustic liner that may be for a turbofan engine may comprise a first sidewall, a second sidewall, wherein the first sidewall and the second sidewall define a plurality of cells disposed therebetween, and a pre-cut septum ribbon comprising a plurality of permeable septum coupled together via a plurality of connecting members, with each one of the plurality of permeable septum being generally in a respective one of the plurality of cells.

An airfoil that includes an airfoil body having a root and a tip, and convex and concave sides that extend between a leading edge and a trailing edge. The airfoil also includes at least a first cladding element that is attached to the airfoil body. The first cladding element includes a first portion and a second portion. The second portion is configured to separate from the first portion when the first cladding element encounters a force of at least a predetermined amount.

A turbomachine includes a shroud and a rotor, which includes first and second blades. A first blade tip and a second blade tip respectively include a base and a first layer. The second blade tip also includes an abrasive second layer layered over the respective first layer. The first layer has a lower material hardness than the shroud. The second layer has a lower thermal stability than the shroud and the first layer. The rotor performs a grind operation and, subsequently, a post-grind operation. The second layer, in the grind operation, contacts and removes material from the shroud, and wears away, thereby revealing the first layer of the second blade tip for the post-grind operation. The first layer of the first blade tip is spaced apart with at least some radial clearance from the shroud in the grind and post-grind operations.

A turbomachine airfoil element includes an airfoil that has pressure and suction sides spaced apart from one another in a thickness direction and joined to one another at leading and trailing edges. The airfoil extends in a radial direction a span that is in a range of 4.47-4.77 inch (113.5-121.2 mm). A chord length extends in a chordwise direction from the leading edge to the trailing edge at 50% span and is in a range of 2.57-2.87 inch (65.3-72.9 mm). The airfoil element includes at least two of a first mode with a frequency of 297±10% Hz, a second mode with a frequency of 1035±10% Hz, a third mode with a frequency of 1488±10% Hz, a fourth mode with a frequency of 1524±10% Hz, a fifth mode with a frequency of 2855±10% Hz and a sixth mode with a frequency of 4462±10% Hz.