An engine case is provided having a first solid wall region and a second solid wall region with an internal region between the first and second sold wall regions. The internal region defines at least one cavity. One or more lattice structures are provided within the cavity that controls the flow of coolant air through the cavity. The cavity may be divided into two or more distinct cooling regions for allowing particular coolant flow paths to be provided to different parts of the engine case.

A seal for a gas turbine engine includes a plurality of seal arc segments. Each of the seal arc segments includes radially inner and outer sides and sloped first and second circumferential sides. The seal arc segments are circumferentially arranged about an axis such that the sloped first and second circumferential sides define gaps circumferentially between adjacent ones of the seal arc segments. Each of the gaps extends from the radially inner sides along a respective central gap axis that slopes with respect to a radial direction from the axis.

A gas turbine engine includes a first case, a second case, and a fastener. The first case has a first flange defining a first opening. The second case has a second flange defining a second opening. The second flange abuts the first flange. The fastener extends through the first opening and the second opening. The fastener defines a first bore that extends through a fastener head along an axis towards an outlet port defined by a fastener shank.

A compressor system (10) includes a motor (3) including a rotor (31) configured to rotate about an axis and a stator (32) disposed on an outer circumference side of the rotor (31), a compressor (2) including an impeller (22) configured to compress a working fluid (31) by rotating together with the rotor and a housing (23) covering the impeller (22) from an outer circumference side, and a heat exchange flow path (600) through which a fluid flowing inside the stator (32) or in a gap between the stator (32) and the rotor (31) is flowed and heat is exchangeable between the fluid and the housing (23).

A compressor system (10) includes a motor (3) including a rotor (31) configured to rotate about an axis and a stator (32) disposed on an outer circumference side of the rotor (31), a compressor (2) including an impeller (22) configured to compress a working fluid (31) by rotating together with the rotor and a housing (23) covering the impeller (22) from an outer circumference side, and a heat exchange flow path (600) through which a fluid flowing inside the stator (32) or in a gap between the stator (32) and the rotor (31) is flowed and heat is exchangeable between the fluid and the housing (23).

In a gas turbine for aeronautic engines, a stator body delimited by an outer lateral surface is cooled by an air cooling device having a plurality of circumferential tubes for distributing air on the outer lateral surface; each circumferential tube having a plurality of outlets for guiding respective cooling airflows towards the outer lateral surface and into a respective circumferential channel obtained between two groups of circumferential channels adjacent to each other and lapped by the flow of air leaving the circumferential channel.

In a gas turbine for aeronautic engines, a stator body delimited by an outer lateral surface is cooled by an air cooling device having a plurality of circumferential tubes for distributing air on the outer lateral surface; each circumferential tube having a plurality of outlets for guiding respective cooling airflows towards the outer lateral surface and into a respective circumferential channel obtained between two groups of circumferential channels adjacent to each other and lapped by the flow of air leaving the circumferential channel.

Aspects of the disclosure are directed to systems and methods for receiving operating state parameters associated with an operative state of an aircraft, determining a clearance value between a first structure of the engine and a second structure of the engine, where the clearance value is determined based on the operating state parameters and a passive clearance model that includes a specification of an uncertainty in the clearance value, determining that the clearance value deviates from a clearance target in an amount that is greater than a threshold, and engaging an active clearance control (ACC) mechanism based on the deviation.

Thermal lifting members for blade outer air seal supports of gas turbine engines include a hollow body defining a thermal cavity therein, at least one inlet fluid connector fluidly connected to the thermal cavity configured to supply hot fluid to the thermal cavity from a fluid source, at least one outlet fluid connector fluidly connected to the thermal cavity configured to allow the hot fluid to exit the thermal cavity, and at least one lifting hook configured to engage with a blade outer air seal support, wherein the thermal lifting member is configured to thermally expand outward when hot fluid is passed through the thermal cavity such that during thermal expansion the at least one lifting hook forces the blade outer air seal support to move outward.

Thermal lifting members for blade outer air seal supports of gas turbine engines include a hollow body defining a thermal cavity therein, at least one inlet fluid connector fluidly connected to the thermal cavity configured to supply hot fluid to the thermal cavity from a fluid source, at least one outlet fluid connector fluidly connected to the thermal cavity configured to allow the hot fluid to exit the thermal cavity, and at least one lifting hook configured to engage with a blade outer air seal support, wherein the thermal lifting member is configured to thermally expand outward when hot fluid is passed through the thermal cavity such that during thermal expansion the at least one lifting hook forces the blade outer air seal support to move outward.