A system and method for measuring average temperature of gas in an axial cross-section of a gas turbine engine gas path, involving diverting gas samples from different positions in the axial cross-section to a gas mixing chamber and measuring a temperature of the resulting mixed gas.

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 high-performance metal alloy is disclosed being suitable for additive manufacturing of machine components, in particular machine components which are subjected to high gas temperature stress. Exemplary machine components are statoric components of gas turbines, such as nozzles.

The aircraft-engine gas turbine includes an outer sealing ring for sealing an array of rotor blades that can be attached to a housing by a clamping mechanism (80) in a friction fit, and a plurality of ring segments (20.sub.i, 20.sub.1+1), wherein .[.a free axial path length (a.sub.f) of a sealing ring segment counter to the direction of through-flow is at least as large as an axial engagement (a.sub.1) of a rotation locking member (10) of the outer sealing ring (a.sub.fa.sub.1), which is free of form fit counter to the direction of through-flow, and/or an axial overhang (a.sub.2) of a radial mounting rail (23) of the outer sealing ring (a.sub.fa.sub.2), and/or an axial offset (a.sub.3, a.sub.4) of a sealing fin (31, 41); and/or.]. a quotient of a specific clearance sum of the outer sealing ring attached to the housing in a friction fit .Iadd.and pi is at least as large as a difference between a maximum outer diameter of the outer sealing ring and a minimum inner diameter of the flow channel inlet of the housing.Iaddend..

A turbine vane for a gas turbine engine having a plurality of cooling holes defined therein is provided. The plurality of cooling holes provide fluid communication to a surface of the turbine vane, the plurality of cooling holes including holes noted by the following coordinates: TVA, TVB, TVC, TVD and TVE of Table 1.

A gas turbine engine for an aircraft includes an engine core with a turbine, a compressor, and a core shaft connecting them. The engine includes a fan, with a plurality of fan blades, located upstream of the core and a gearbox receiving an input from the core shaft and outputting drive so the fan is at a lower rotational speed than the core shaft. The turbine includes a plurality of stages of axially spaced rotor blades mounted on a rotor, which are surrounded by a turbine casing. The turbine has an inlet defined at an upstream end of a first stage of blades and an outlet defined at a downstream end of a last stage of blades and a ratio of the area of the outlet to the inlet is at between 2.5 and 3.5. This increases the pressure ratio of and power extracted from the turbine and the engine.

A nozzle segment for a gas turbine engine with turbine airflow passing through the gas turbine engine. The nozzle segment including an upper shroud and an inner hub. The nozzle segment including a first airfoil and second airfoil extending from the upper shroud to the inner hub. The first airfoil and second airfoil including conduits for delivering secondary air to displace a portion of the turbine airflow passing through the gas turbine engine.

A turbine nozzle having an airfoil profile substantially in accordance with Cartesian coordinate values of X, Y, and Z set forth in Table 1, and within an envelope of approximately +/0.049 inches, where the X and Y values are in inches and the Z values are non-dimensional values from 0 to 1 and convertible to Z distances in inches by multiplying the Z values by the height of the airfoil in inches. The X and Y values are distances which, when connected by smooth continuing arcs, define airfoil profile sections at each distance Z. The profile sections at each distance Z are joined smoothly to one another to form the airfoil shape. The X and Y values may also be scaled as a function of a first constant and the Z values may be scaled as a function of a second constant.

A first stage vane array of a high pressure turbine that may be for a geared turbofan engine includes a plurality of airfoils circumferentially spaced from one-another and orientated about an engine axis. Each airfoil has a leading edge and a trailing edge with the trailing edge being circumferentially separated by the next adjacent trailing edge by a pitch distance. The leading a trailing edges of each one of the plurality of airfoils are axially separated by an axial chord length. A pitch-to-chord ratio of the pitch distance over the axial chord length is equal to or greater than 1.7.

A gas turbine engine includes a main compressor. A tap is fluidly connected downstream of the main compressor. A heat exchanger is fluidly connected downstream of the tap. An auxiliary compressor unit is fluidly connected downstream of the heat exchanger. The auxiliary compressor unit is configured to compress air cooled by the heat exchanger with an overall auxiliary compressor unit pressure ratio between 1.1 and 6.0. An intercooling system for a gas turbine engine is also disclosed.