A turboengine as disclosed includes an outer wall structure and an inner wall structure, wherein the inner wall structure is provided at a radially inner position with respect to the outer wall structure, and each of the wall structures has a surface, the surfaces being arranged facing each other in the radial direction. At least one guide vane member includes at least one airfoil, a radially inner end and a radially outer end. The inner wall structure and the outer wall structure are jointly provided as a vane carrier unit, wherein the inner wall structure and the outer wall structure are fixedly connected to each other by at least one bridging member extending between the inner wall structure and the outer wall structure.

A cooling device for cooling platforms of a guide vane ring of a gas turbine is arranged downstream inside a main flow channel of a combustion chamber. Cooling air passages are arranged in a wall of the platforms or of an intermediate piece that is connected therewith to guide cooling air for film cooling the surfaces of the platforms. At least in certain areas, the wall is configured with at least two layers having—as viewed from the main flow channel—an outer wall and a spaced apart inner wall forming a hollow space, wherein the hollow space can be impinged by cooling air through at least one cooling air blow-in opening inside the outer wall, and at least one cooling air blow-out opening is arranged inside the inner wall extending in the downstream direction to the surfaces of the platforms.

The invention relates to a measuring device for an aircraft engine, characterized by at least one probe device for measuring a physical and/or chemical state in at least one measuring space within the aircraft engine, wherein the at least one measuring space is fluidically connected to a cavity, and at least one air-conducting device, which is fluidically coupled to the cavity in such a manner that a fluid flow, in particular a gas flow, can be removed from the at least one cavity to a pressure sink. The invention also relates to an aircraft engine and to a measuring method.

A turbine nozzle has an airfoil that includes a pressure side portion of a nominal airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z of a pressure side as set forth in Table I. The Cartesian coordinate values of X, Y, and Z are non-dimensional values from 0% to 100% convertible to dimensional distances by multiplying the Cartesian coordinate values of X, Y and Z by a height of the airfoil defined along the Z axis. The X and Y values of the pressure side are coordinate values that, when connected by smooth continuing arcs, define pressure side sections of the pressure side portion of the nominal airfoil profile at each Z coordinate value. The pressure side sections may be joined smoothly with one another to form the pressure side portion.

A highly efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

A turbine article includes a substrate with a geometric surface having a multiple of divots recessed into the substrate, and a ceramic topcoat disposed over the geometric surface, the topcoat including at least a first layer having a first hardness and a second layer having a second hardness, the first hardness different than the second hardness.

An integrated combustor nozzle includes an inner liner segment; an outer liner segment; and a panel extending radially between the inner and outer liner segments. The panel includes a forward end, an aft end, and a side walls extending axially from the forward end to the aft end. The aft end defines a turbine nozzle having a trailing edge circumferentially offset from the forward end. The inner liner segment has a pair of sealing surfaces, each of which defines a first continuous curve in the circumferential direction. The outer liner segment has a pair of sealing surfaces, each of which defines a second continuous curve in the circumferential direction. In some instances, the curves are monotonic in the circumferential direction. A segmented annular combustor including an array of such integrated combustor nozzles is also provided.

A first-stage turbine vane supporting structure reduces chordal leakage between a first-stage turbine vane segment and its supporting ring. The structure includes a first-stage turbine vane segment including an outer platform, an inner platform, and a first-stage turbine vane disposed between the inner and outer platforms; an inner rail protruding from the inner platform in a longitudinal direction of the first-stage turbine vane; a supporting ring for supporting the first-stage turbine vane segment by engaging with the inner rail while facing one surface of the inner rail; a supporting member engaging with the supporting ring to define a U-shaped space bordered by three side surfaces surrounding the inner rail; and a flexible member interposed between the inner rail and one of the three side surfaces of the U-shaped space. A protrusion formed on the inner rail engages with a sealing surface of the supporting ring in an airtight manner.

A gas turbine engine comprises a main compressor section having a high pressure compressor with a downstream most end, and more upstream locations. A turbine section has a high pressure turbine. A first tap taps air from at least one of the more upstream locations in the main compressor section, passes the tapped air through a heat exchanger and then to a cooling compressor. The cooling compressor compresses air downstream of the heat exchanger. A second tap taps air from a location closer to the downstream most end than the location(s) of the first tap. The first and second tap mix together and are delivered into the high pressure turbine. An intercooling system for a gas turbine engine is also disclosed.

An outlet guide vane assembly in a gas turbine engine is presented. The outlet guide vane assembly includes an inner shroud and an outlet guide vane having an inner platform. The inner shroud has a flange arranged at aft side that is bolted to a flange of the inner platform arranged at forward side. The inner shroud flange has a protrusion that engages a recess of the inner platform flange forming a form fit connection interface between the inner shroud and the outlet guide vane. The inner platform has shiplaps arranged at two circumferential sides that overlap shiplaps of an adjacent inner platform forming a form fit connection interface between adjacent outlet guide vanes. The outlet guide vane assembly includes a plurality of segments circumferentially arranged. Each segment includes a plurality of outlet guide vanes assembled to an inner shroud.