A compressor for an aircraft engine. A rotor includes blades rotatable about an axis. Blade tips extend between leading and trailing edges. A shroud surrounds the rotor, with an inner surface surrounding the tips. Grooves are defined in the shroud inner surface adjacent the tips. The grooves extend circumferentially about the shroud and radially from inlet openings to closed end surfaces. Groove sidewalls extend circumferentially about the axis. The grooves are axially spaced-apart, the most upstream inlet opening having an upstream end disposed upstream of the leading edges of the blades. The grooves have a swept angle from the inner surface, with a center of the inlet openings is axially offset of a center of the closed-end surfaces. The grooves span an overall axial distance corresponding to 30% or more of the blades’ chord length. The grooves have circumferential interruptions defined by baffles, and extend non-continuously around a shroud circumference.

Turbine and compressor casing abradable components for turbine engines include abradable surfaces with a zonal system of forward (zone A) and rear or aft sections (zone B) surface features. The zone A surface profile comprises an array pattern of non-directional depression dimples, or upwardly projecting dimples, or both, in the abradable surface. The dimpled forward zone A surface features reduce surface solidity in a controlled manner, to help increase abradability during blade tip rubbing incidents, yet they provide sufficient material to resist incoming hot working fluid erosion of the abradable surface. In addition, the dimples provide generic forward section aerodynamic profiling to the abradable surface, compatible with different blade airfoil-camber profiles. The aft zone B surface features comprise an array pattern of ridges and grooves.

A circumferential seal assembly suitable for forming a thin film between a rotatable runner and a sealing ring is presented. The assembly includes an annular seal housing, a rotatable runner, an annular seal ring, and a plurality of groove structures. Each groove structure includes a groove and an optional feed groove. The groove includes at least two adjoining steps defined by base walls arranged to decrease depthwise. Two adjoining base walls are disposed about a base shoulder. Each base shoulder locally redirects a longitudinal flow to form an outward radial flow in the direction of the annular seal ring. The base walls are bounded by and intersect a pair of side walls. A side wall includes side shoulders which narrows the groove widthwise and locally redirects the longitudinal flow to form a lateral flow in the direction of the other side wall. Outward and lateral flows separately or in combination enhance stiffness of a thin-film layer between the annular seal ring and the rotatable runner.

A compressor wheel for a compressor of a turbocharger has a hub (16) and a multiplicity of blades (12) on the hub (16). In intermediate spaces of the multiplicity of blades (12), a channel is in each case formed between a suction side (24) and a pressure side (26). The channel guides fluid that flows in axially in relation to a rotation axis (22) radially or radially-axially outward. The hub (16) in relation to the rotation axis (22) in at least one channel is contoured with a rotationally symmetrical portion (18′) and a non-rotationally symmetrical portion (18). The non-rotationally symmetrical portion (18) is formed by radii that are variable in the flow direction, and a transition region (28) between the hub (16) and the blade (12) adjoins the non-rotationally symmetrical portion (18).

A fan containment system for fitment around an array of radially extending fan blades mounted on a hub in an axial gas turbine engine, including: a fan case having an annular casing element for encircling an array of fan blades and an annular fan track liner lining a radially inner side of the casing element, including a first fan track liner panel positioned circumferentially adjacent a second fan track liner panel. An engagement arrangement engages the first fan track liner panel to the second, extending along at least a portion of adjacent sides of the liner panels, and is configured and the liner panels are shaped such that axial movement of the first fan track liner panel by a distance less than the length of the engagement arrangement, being measured in a direction parallel to the adjacent sides of the liner panels, disengages the first liner panel from the second.

A bearing compartment seal for a gas turbine engine includes a seal ring that defines an axis and has a radially inward facing sealing surface. A seal runner is configured to rotate relative to the seal ring. The seal runner has a runner surface facing the radially inward facing sealing surface. A plurality of grooves are spaced circumferentially along the runner surface. The plurality of grooves have a length in an axial direction that is at least 50% of an axial length of the runner surface.

A rotor disk is provided. The rotor disk may comprise a disk lug and a trench. The disk lug may be fixed to a distal surface of the rotor disk. The trench may be disposed on a surface of the disk lug. The trench may extend radially inwards from a distal surface of the disk lug. The trench may be configured to at least partially define a flow path by which cooling air may reach a distal surface of the disk lug in order to provide disk lug cooling.

An engine component with an abradable material and treatment is disclosed herein. An example engine component includes a substrate having a first thickness and defining at least a first substrate surface, the substrate extending along an axial direction, an abradable material on the first substrate surface, the abradable material defining a first abradable material surface, a second abradable material surface spaced radially apart from the first abradable material surface and defining a second thickness, and a treatment, the treatment extending from the first abradable material surface through at least a portion of the first thickness of the substrate.

Various embodiments of an airfoil and machines with airfoils are disclosed. The airfoils include a thicker leading airfoil portion and a thinner trailing airfoil portion. In one embodiment, the leading airfoil portion is formed by bending a body of the airfoil back toward itself. In another embodiment, the leading airfoil portion has a solid geometry and includes two elliptic surfaces. To prevent detachment of airflow, the leading airfoil portion includes at least two arc portions or surfaces that act to direct the airflow down to the trailing airfoil portion in a manner that stabilizes vortexes that may form in the region of changing thickness.

A turbomachine includes a compressor including variable-pitch stationary vanes each extending radially between a rotary hub and a stationary casing surrounding this rotary hub, each variable-pitch vane including a blade having a base spaced apart by a first radial gap from a stationary wall of the casing, and a tip spaced apart by a second radial gap from a rotary wall of the rotary hub. The stationary wall of the casing or the rotary wall of the rotary hub includes at the blade a shape treatment arranged to channel an air leak passing through the corresponding gap.