An aircraft thrust reverser inner wall and method of manufacturing the same. The aircraft thrust reverser inner wall may include a face sheet, a perforated back sheet, and a core sandwiched between the face sheet and the perforated back sheet. The face sheet may have an inner face sheet surface and an outer face sheet surface, and the core may have an inner core surface, an outer core surface, and a plurality of cell walls extending therebetween. An electro-depositable material may be applied, via electrodeposition, in a substantially continuous layer over the outer core surface, the cell walls, and the outer face sheet surface, thus bonding the face sheet and core together. The perforated back sheet may be attached to the core at the outer core surface, and a conductive coating may be applied to the inner face sheet surface.

A solid rocket propellant grain having rocket propellant and a membrane in contact with the rocket propellant. The membrane includes a highly heat conductive pattern which affects the propellant burning rate through localized conductive heat transfer from the combustion zone and into the uncombusted propellant. Different geometries for the thermally conductive pattern produce different combustion results.

A high-efficiency compressor section (10) for a gas turbine engine is disclosed. The compressor section includes a vane carrier (12) adapted to hold ring segment assemblies (16) that provide optimized blade tip gaps (28,29) during a variety of operating conditions. The ring segment assemblies include backing elements (30) and tip-facing elements (32) urged into a preferred orientation by biasing elements (40) that maintain contact along engagement surfaces (44,46). The backing and tip-facing elements have thermal properties sufficiently different to allow relative growth that strategically forms an interface gap (42,43) therebetween, resulting in blade tip gaps that are dynamically adjusted operation.

A high-efficiency compressor section (10) for a gas turbine engine is disclosed. The compressor section includes a vane carrier (12) adapted to hold ring segment assemblies (16) that provide optimized blade tip gaps (28,29) during a variety of operating conditions. The ring segment assemblies include backing elements (30) and tip-facing elements (32) urged into a preferred orientation by biasing elements (40) that maintain contact along engagement surfaces (44,46). The backing and tip-facing elements have thermal properties sufficiently different to allow relative growth and geometric properties strategically selected to strategically form an interface gap therebetween (42,43), resulting in blade tip gaps that are dynamically adjusted operation.

A method for joining engine components includes positioning a first plurality of thermal protection structures across a thermal protection space between a first thermal protection surface and a second thermal protection surface. The first and second engine components are locally joined by forming a first plurality of transient liquid phase (TLP) or partial transient liquid phase (PTLP) bonds along corresponding ones of the first plurality of thermal protection structures between the first thermal protection surface and the second thermal protection surface. The second thermal protection surface is formed from a second surface material different from a first surface material of the first thermal protection surface.

A turbine engine module, in particular an aircraft turbine engine, including an annular casing having an internal wall forming a channel wall; and a nozzle surrounded by the casing and including an annular external platform and an annular internal platform between which stator blades extend, the external platform having an external face that faces the internal wall of the casing and includes an annular groove oriented towards the outside and housing a sealing device, the sealing device coming into cylindrical contact with a track of the internal wall of the casing, wherein the internal wall of the casing includes a thermal barrier made of ceramic material directly above the track, the track being arranged between the thermal barrier and the sealing device.

The CPAP device includes a blower. The blower includes a motor and a blower fan rotatable by the motor. The motor includes a stator assembly including a plurality of coils arranged in a circumferential direction, and a rotor assembly including a permanent magnet and disposed in an inner side portion of the stator assembly in a radial direction. An output shaft is inserted into the rotor assembly. The output shaft rotates together with the rotor assembly. The blower fan is fixed to a tip end of one side of the output shaft. Part of the output shaft is inserted into an insertion hole provided at a center of the base. A plain bearing is interposed between an inner peripheral surface of the insertion hole and an outer peripheral surface of the output shaft.

A stator structure includes a silicon steel sheet, a first plastic material and a second plastic material. The silicon steel sheet has a plurality of pole faces. The first plastic material covers on the silicon steel sheet. The second plastic material forming on the first plastic material. The first plastic material covered on the plurality of the pole faces is not covered by the second plastic material. The thermal conductivity of the first plastic material is higher than that of the second plastic material.

A protective coating for a surface exposed to hot gas flow comprises a thermal layer, a conducting layer and an abrasive layer. The thermal layer comprises alumina having a sufficient amount of impurities to lower the thermal layer thermal conductivity. The layer is formed from a powder having a thermal conductivity no more than 10 BTU in/hr ft.sup.2 F., and overlies the surface. The conducting layer overlies the thermal layer. The abrasive layer comprises abrasive particles bonded in a metal matrix that is electroplated onto the conducting layer.

A turbocharger assembly can include a center housing; a turbine housing; and an insert disposed between the center housing and the turbine housing where a surface of the turbine housing and a first surface of the insert define a volute and where a surface of the center housing and a second surface of the insert define a chamber.