A compressor rotor for a gas turbine engine has blades circumferentially distributed around and extending a span length from a central hub. The blades include alternating first and second blades having airfoils with corresponding geometric profiles. The airfoil of the first blade has a coating varying in thickness relative to the second blade to provide natural vibration frequencies different between the first and the second blades.

A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved overall performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

A rotor for a gas turbine engine includes a hub having an outer periphery, a first plurality of blades arranged on the outer periphery of the hub, and a second plurality of blades arranged on the outer periphery of the hub. The first plurality of blades is made of a first material and has a first elastic modulus over density ratio. The first plurality of blades has a first natural frequency. The second plurality of blades is made of a second material and has a second elastic modulus over density ratio. The second elastic modulus over density ratio is different from the first elastic modulus over density ratio. The second plurality of blades has a second natural frequency different from the first natural frequency. A method of fabricating a bladed rotor is also presented.

The invention relates to a blade made of composite material, provided with a reinforcement (17) made of a stronger material with regard to a leading edge (5) likely to be struck by solid objects, which comprises, under the aerodynamic portion of the blade, extensions (18, 19) attached to the blade root. According to the invention, the extensions (18, 19) are dissymmetric, the extension (18) located on the suction side being generally further back from the leading edge (5) than the extension (19) on the pressure side.

A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

A blade containment structure surrounding a fan in a turbofan engine is disclosed. The blade containment structure includes a cellular material to absorb energy and contain fragments of a blade thrown outward; an inner shell; a ductile back sheet spaced radially outward from the inner shell, the ductile back sheet and inner shell cooperating to define a nesting area for the cellular material, wherein the cellular material is bound at its radially inner surface by the inner shell and at its outer surface by the ductile back sheet; and a containment blanket overlaid on the ductile back sheet, the containment blanket being of the type effective to contain fragments of the blade that penetrate through the ductile back sheet.

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.

A locking spacer, which is fitted in a dovetail slot provided on an outer circumferential surface of a disk put on a rotor shaft, includes: a pair of first blocks each provided with a dovetail joint and a stepped seating surface with a first bolt hole, and configured to have a size occupying a portion of an internal space of the dovetail slot; a second block having a size to be inserted into a remaining portion of the internal space of the dovetail slot, and having a height corresponding to the seating surfaces; a fixing plate seated on both the seating surfaces of the first blocks and an upper surface of the second block, and provided with second bolt holes corresponding to the first bolt holes; and a bolt screwed into the first bolt hole through an associated second bolt hole.

Provided are a composite coating layer having improved erosion resistance and a turbine component including the same. The composite coating layer may include a TiN layer; and a TiAlN layer, wherein the composite coating layer is formed by alternately stacking the TiN layer and the TiAlN layer, and a total number of layers including the TiN layer and the TiAlN layer is 6 to 18, whereby the composite coating layer is capable of exhibiting high erosion resistance, high hardness, superior high-cycle fatigue characteristics and low surface roughness. Moreover, the turbine component including the composite coating layer is also capable of manifesting improved properties, such as high erosion resistance, high hardness, superior high-cycle fatigue characteristics and low roughness, thus remarkably increasing lifespan characteristics.

The present disclosure relates to a ring for controlling discharge gates for an aircraft turbine engine. The control ring extends around a casing of the turbine engine and includes a connector for connecting to the gates, characterized in that it comprises metal sectors and composite material sectors. The metal sectors may be connected to one another by the composite material sectors and the connector may be carried by the metal sectors.