Engineered groove features (EGFs) are formed within thermal barrier coatings (TBCs) of turbine engine components. The EGFs are advantageously aligned with likely stress zones within the TBC or randomly aligned in a convenient two-dimensional or polygonal planform pattern on the TBC surface and into the TBC layer. The EGFs localize thermal stress- or foreign object damage (FOD)-induced crack propagation within the TBC that might otherwise allow excessive TBC spallation and subsequent thermal exposure damage to the turbine component underlying substrate. Propagation of a crack is arrested when it reaches an EGF, so that it does not cross over the groove to otherwise undamaged zones of the TBC layer. In some embodiments, the EGFs are combined with engineered surface features (ESFs) that are formed in the component substrate or within intermediate layers applied between the substrate and the TBC.

An impeller for a turbine engine is disclosed, such as a turbojet engine or a turboprop engine of an aircraft, for rotating about an axis of rotation in a predetermined direction of rotation and for having a flow of gas passing therethrough, comprising a series of blades distributed circumferentially about the axis of rotation of the wheel, the radially external end of each blade comprising a heel supporting at least one radially external lip, extending circumferentially in an oblique direction relative to the radial plane perpendicular to the axis of rotation, the lip being angled circumferentially in the downstream direction in the aforementioned direction of rotation such as to enable a compression of the gases located upstream from the lip during the rotation of the impeller.

A geared turbofan engine includes a first rotor, a fan, a second rotor, a gear train, a fan casing, a nacelle and a plurality of discrete acoustic liner segments. The fan is connected to the first rotor and is capable of rotation at frequencies between 200 and 6000 Hz and has a fan pressure ratio of between 1.25 and 1.60. The gear train connects the first rotor to the second rotor. The fan casing and nacelle are arranged circumferentially about a centerline and define a bypass flow duct in which the fan is disposed. The plurality of discrete acoustic liner segments with varied geometric properties are disposed along the bypass flow duct.

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan section including a fan, the fan including a plurality of fan blades, a diameter of the fan having a dimension D that is based on a dimension of the fan blades, each fan blade having a leading edge, and a forward most portion on the leading edges of the fan blades in a first reference plane, a geared architecture, a turbine section including a high pressure turbine and a low pressure turbine, the low pressure turbine driving the fan through the geared architecture, a nacelle surrounding the fan, the nacelle including an inlet portion forward of the fan, a forward edge on the inlet portion in a second reference plane, and a length of the inlet portion having a dimension L measured along an engine axis between the first reference plane and the second reference plane. A dimensional relationship of L/D is no more than 0.45.

According to an example embodiment, a gas turbine engine assembly includes, among other things, a fan that has a plurality of fan blades. A diameter of the fan has a dimension D that is based on a dimension of the fan blades. Each fan blade has a leading edge. An inlet portion is situated forward of the fan. A length of the inlet portion has a dimension L between a location of the leading edge of at least some of the fan blades and a forward edge on the inlet portion. A dimensional relationship of L/D is between about 0.2 and 0.45.

A strut for a gas turbine engine includes a body defining a first side and an opposite second side. The first side is spaced from the second side along the circumferential direction of the gas turbine engine. Additionally, the body includes an inner section, a middle section, and an outer section. Each of the inner, middle, and outer sections are arranged in series order along a span of the strut and define a thickness between the first and second sides of the strut. The thickness of the middle section is greater than the thicknesses of the inner section and of the outer section to increase the strut's resistance to buckling in response to forces exerted thereon.

A turbine abradable component includes a support surface and a thermally sprayed ceramic/metallic abradable substrate coupled to the support surface for orientation proximal a rotating turbine blade tip circumferential swept path. An elongated pixelated major planform pattern (PMPP) of a plurality of discontinuous micro surface features (MSF) project from the substrate surface. The PMPP repeats radially along the swept path in the blade tip rotational direction, for selectively directing airflow between the blade tip and the substrate surface. Each MSF is defined by a pair of first opposed lateral walls defining a width, length and height that occupy a volume envelope of 1-12 cubic millimeters. The PMPP arrays of MSFs provide airflow control of hot gasses in the gap between the abradable surface and the blade tip with smaller potential rubbing surface area than solid projecting ribs with similar planform profiles.

A turbine engine casing flow-path segment that is locally diffusing, followed by a flow-path segment contracting in the vicinity of a fan blade. This contraction accelerates the fluid flow axially forward of the fan blade leading edge at the tip and converges with the linear flow-path aft of the fan blade leading edge but forward of the fan blade trailing edge. More diffused fluid flow results in increased flow capacity of the fan, and increased fan efficiency.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.

A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.