Mini-disks of gas turbine engines are provided having an axially extending portion extending axially with respect to an axis of the engine, the axially extending portion configured to engage with a hub arm of a compressor of the engine, a radially extending portion extends radially with respect to the axis, the radially extending portion configured to engage with an attachment of a turbine disk of the gas turbine engine, an intermediate portion extending between the axially extending portion and the radially extending portion, and at least one mini-disk connector configured to engage with a portion of the turbine disk of the gas turbine engine to prevent radial movement of the mini-disk during operation.

Provided is a radial diffuser that includes a housing, a plurality of diffuser vanes, and a plurality of deswirl vanes and at least one vane extension providing a service routing. Each of the vane extensions is disposed after a radial section and may extend into or through a transition and into the deswirl cascade. At least a portion of the vane extensions include a service passage extending therethrough. Each service passage is configured to allow a service conduit to extend therethrough without adversely crossing either a diffusion flow passage or a transition flow passage.

A method of polishing a metallic workpiece includes: mounting the workpiece in a hopper; loading the hopper with a polishing media comprising, by weight percent, more than 98% metallic chips, less than 2% liquid, and less than 0.05% abrasive; and oscillating the hopper for a run time, thereby generating a flow of the polishing media over the workpiece, until a predetermined surface finish is achieved on the workpiece.

A gas turbine engine includes a main engine compressor section. A booster compressor changing a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor operates at airflow conditions greater than a demand by the pneumatic system. A speed change system driving the booster compressor at speeds corresponding to a demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

A gas turbine engine includes a main engine compressor section. A booster compressor changes a pressure of airflow received from the main engine compressor section to a pressure desired for a pneumatic system. The booster compressor is configured to operate at airflow conditions greater than a demand of the pneumatic system. An exhaust valve controls airflow between an exhaust outlet and an outlet passage to the pneumatic system. The exhaust valve is operable to exhaust airflow from the booster compressor in excess of the demand of the pneumatic system. A bleed air system for a gas turbine engine and a method of controlling engine bleed airflow are also disclosed.

A gas turbine engine includes a compressor section including a first compressor, a turbine section including a first turbine and a second turbine, a first shaft and a second shaft, the first shaft interconnecting the first turbine and the second compressor, and a geared architecture. The first shaft is supported on a first bearing in an overhung manner. A performance ratio is between 0.5 and 1.5.

Systems and methods for cooling a core flow-path of a compressor section of a gas turbine engine are provided. In various embodiments, a cooling system for a gas turbine engine may comprise a valve system located radially outward from an engine case, the valve system being coupled between an air duct and the engine case, the valve system having an actuation device configured to at least one of open or close a valve in response to a command from an electronic engine controller, wherein cooling air from the air duct can pass through the valve system and enter the engine case into a high pressure compressor plenum when the valve system is in an open position.

A sub-assembly for a low-pressure compressor of an aircraft turbine engine includes a straightener provided with cantilevered vanes and a rotor hub having a cavity covered by an inner shroud opposite the vanes. Orifices are made in the inner shroud to allow an air flow to circulate in the downstream to upstream direction of the low-pressure compressor.

A coated substrate comprises: a metallic substrate; a bondcoat atop the substrate; and a ceramic barrier coat atop the bondcoat. The bondcoat has a combined content of one or more of molybdenum, chromium, and vanadium of at least 50 percent by weight.

A blisk fan is provided for a turbine engine propulsion system. The blisk fan includes a hub configured to rotate about a rotational axis at a maximum rotational speed, and a plurality of blades extending radially outward from the hub to define a fan leading edge tip diameter. Each of the blades has a first vibratory mode at a natural frequency, which is greater than a first fan order and less than a second fan order at the maximum rotational speed. The compression system preferably has a balance factor of the compression system between 1.9 and 3.2.