A seal device includes a first sidewall, a second sidewall and a carbon seal element. The first sidewall is arranged with a first tool die which centers the first sidewall about an axis. The first sidewall axially contacts the first tool die along the axis. The carbon seal element is arranged with a centering element which centers the carbon seal element about the axis. The carbon seal element circumscribes and radially contacts the centering element. The second sidewall is arranged with a second tool die which centers the second sidewall about the axis. The second sidewall axially contacts the second tool die along the axis. The first tool die and the second tool die are moved axially together along the axis to press fit the second sidewall into a bore of the first sidewall and axially capture the carbon seal element between the first sidewall and the second sidewall.

A seal device includes a first sidewall, a second sidewall and a carbon seal element. The first sidewall is arranged with a first tool die which centers the first sidewall about an axis. The first sidewall axially contacts the first tool die along the axis. The carbon seal element is arranged with a centering element which centers the carbon seal element about the axis. The carbon seal element circumscribes and radially contacts the centering element. The second sidewall is arranged with a second tool die which centers the second sidewall about the axis. The second sidewall axially contacts the second tool die along the axis. The first tool die and the second tool die are moved axially together along the axis to press fit the second sidewall into a bore of the first sidewall and axially capture the carbon seal element between the first sidewall and the second sidewall.

A gas turbine engine according to an example of the present disclosure includes, among other things, a fan section, and a compressor section including at least a first compressor section and a second compressor section. A power ratio is provided by the combination of the first compressor section and the second compressor section. A method of design a gas turbine engine is also disclosed.

The present application concerns a test engine hood for a turbine engine, such as a double flow turbine. The test hood allows replacement of a flight engine hood during tests on a test bench on the ground where the temperature conditions could damage the flight hood. The test hood includes a tubular wall of carbon-fiber epoxy composite, and metal flanges upstream and downstream. To provide thermal protection, the test hood includes a layer of silicone with a majority of polysiloxane. The layer covers the entire inner surface of the wall to create a barrier. The present application also concerns a method for testing a turbine engine on a test bench, where the turbine engine is fitted with a test casing. The present application also concerns a use of silicone for thermal insulation of the inside of the test hood of the turbine engine on a test bench on the ground.

The present application relates to a stator for a low-pressure compressor of an axial turbine engine. The stator includes an annular row of stator vanes with connection vanes, and support vanes which are grouped into vaned casings. The casings include internal platforms and external platforms, and are integrally produced by means of additive production based on titanium powder. The stator has a circular internal shroud which is formed by the internal platforms of the vanes of the casings and internal shroud segments which are fixed to the connection vanes. The internal platforms include pairs of contact surfaces in order to interlock the shroud segments therein by moving into abutment and sliding against the contact surfaces. The contact surfaces are perpendicular to the rotation axis and form axial stops.

The invention deals with a splitter nose delimiting the inlet of a low-pressure compressor of an axial turbine engine. The splitter nose comprises a separation surface with an upstream circular edge suitable for separating a flow entering into the turbine engine into a primary flow and a secondary flow, and a plasma de-icing device. The device comprises two annular layers of dielectric material (42; 44) partially forming the separation surface, an electrode forming the upstream edge, an electrode forming an outer wall of the splitter nose, an electrode forming an outer shroud which supports blades, an electrode delimiting the primary flow. The device generates plasmas (46; 48; 50) opposing the presence of ice on the partitions of the splitter nose. The invention also deals with a turbine engine with a splitter nose that is provided with a de-icing system downstream of the fan.

A method of forming a circuit in a composite component includes providing a plurality of preform modules comprised of an organic matrix composite material, applying at least one electrical circuit on at least first and second preform modules of the plurality of preform modules, and arranging the first and second preform modules such that the electrical circuit of the first preform module is in contact with the electrical circuit of the second preform module. An additional step includes molding the first and second preform modules together to form a one-piece molded component such that the electrical circuits of the first and second preform modules form a complete circuit.

A gas turbine engine includes a fan section and a compressor section. The compressor section includes both a first compressor section and a second compressor section. A turbine section includes at least one turbine and driving the second compressor section and a fan drive turbine driving at least a gear arrangement to drive the fan section. A power ratio is provided by the combination of the first compressor section and the second compressor section, with the power ratio being provided by a first power input to the first compressor section and a second power input to the second compressor section, the power ratio being equal to, or greater than, about 1.0 and less than, or equal to, about 1.4.

A method may comprise: laying up a first plurality of plies of material comprising thermoplastic resin and fiber to form an inner skin preform, the inner skin preform being a continuous sheet including alternating peaks and valleys; laying up a second plurality of plies of material comprising thermoplastic resin and fiber to form an outer skin preform; and joining the inner skin preform to the outer skin preform.

A low-pressure compressor for a turbine engine, such as an aircraft turbojet engine includes a rotor with two rows of rotor blades between which two annular ribs are positioned; and one annular row of stator blades between the rotor blades. An internal shroud is connected to the stator blades. The internal shroud includes abradable material collaborating with the annular ribs, and annular teeth made of an abradable material and which extend radially towards the rotor, so as to provide sealing. The system may be used in a method for manufacturing a bypass turbojet engine compressor.