The method can include determining a mass flow rate W of working fluid circulating through the compressor stage, determining a control parameter value associated to the geometrical configuration of the variable geometry element based on the determined value of mass flow rate W; and changing the geometrical configuration of the variable geometry element in accordance with the determined control parameter value.

An aircraft gas turbine engine has an engine core with a turbine, compressor, and core shaft connecting the turbine to the compressor, a fan upstream of the engine core; and a gearbox. The engine core has three bearings, one forward, two rearward, to support the core shaft, a minor span being the axial distance between the two rearward bearings. A first blade to bearing ratio of the minor span divided by the product of the mass, radius at mid-height, and the square of the angular velocity at cruise for a blade of the lowest pressure set may have a value in the range from 2.0×10.sup.−6 to 7.5×10.sup.−6 kg.sup.−1.Math.rad.sup.−2.Math.s.sup.2. A second blade to bearing ratio of the minor span divided by the product of mass and radius at mid-height for a blade of the lowest pressure set may have a value in the range from 0.8 to 6.0 kg.sup.−1.

A gas turbine engine for an aircraft includes an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor, and a fan including a plurality of fan blades located upstream of the engine core. The fan has a fan diameter in the range from 240 cm to 280 cm. The turbine is the lowest pressure turbine of the engine and the compressor is the lowest pressure compressor of the engine. The turbine includes a total of three sets of turbine blades. The engine core further includes three bearings arranged to support the core shaft. The three bearings include a forward bearing and two rearward bearings located downstream of a leading edge of a lowest pressure turbine blade of the turbine at a root of the blade.

A method of optimizing probe placement in a turbomachine is disclosed which includes establishing a design matrix A of size m×(2N+1) utilized in developing flow properties around an annulus of a turbomachine, where m represents the number of datapoints at different circumferential locations around the annulus, and N represents dominant wavelets generated by upstream and downstream stators and blade row interactions formed around an annulus, wherein m is greater or equal to 2N+1, and optimizing probe positioning by iteratively modifying probe positions placed around the annulus and for each iteration determining a condition number of the design matrix A for each set of probe positions until a predetermined threshold is achieved for the condition number representing an optimal probe layout.

A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

An engine component for a gas turbine engine, the engine component comprising a cooling architecture comprising at least one unit cell having a set of walls with a thickness, the set of walls defining fluidly separate conduits having multiple openings, each of the multiple openings having a hydraulic diameter; wherein the thickness (t) and the hydraulic diameter (D.sub.H) relate to each other by an equation:

[00001]$\frac{{\left(D_H+2t\right)}^2}{{\left(\left(D_H+2t\right)/D_H\right)}^{1/3}}$

to define a performance area factor (PAF).

A gas turbine engine for an aircraft including: an engine core including a turbine, a compressor, and a core shaft connecting the turbine to the compressor; a fan located upstream of the engine core, the fan including a plurality of fan blades; a gearbox that can receive an input from the core shaft, and can output drive to a fan shaft via an output of the gearbox so as to drive the fan at a lower rotational speed than the core shaft; and a fan shaft mounting structure arranged to mount the fan shaft within the engine, the fan shaft mounting structure including at least two supporting bearings connected to the fan shaft. A fan-gearbox axial distance is defined as the axial distance between the output of the gearbox and the fan axial centreline, the fan-gearbox axial distance being greater than or equal to 0.35 m.

A turbomachinery includes a casing, a rotor shaft rotatably attached to the casing, a first blade row fixed to either one of the rotor shaft or the casing, and a second blade row fixed to either one of the rotor shaft or the casing and arranged adjacent to the upstream side or downstream side of the first blade row, wherein the turbomachinery sets the number of first blades and the number of second blades in a manner that the interblade phase angle difference of the second blade row is ±180°.

A gas turbine engine includes a core turbine engine and a fan mechanically coupled to the core turbine engine. The fan includes a plurality of fan blades, each fan blade defining a base and an inner end along a radial direction of the gas turbine engine. The fan also includes a hub covering the base of each of the plurality of fan blades. Further, the fan includes one or more bearings for supporting rotation of the plurality of fan blades. The one or more bearings define a fan bearing radius along a radial direction of the gas turbine engine. Similarly, the hub defines a hub radius along the radial direction of the gas turbine engine. The ratio of the hub radius to the fan bearing radius is less than about three, providing for desired packaging of the various components within the fan of the gas turbine engine.

The present invention belongs to the technical field of control of aero-engines, and proposes an adaptive boosting algorithm-based turbofan engine direct data-driven control method. First, a turbofan engine controller is designed based on the Least Squares Support Vector Machine (LSSVM) algorithm, and further, the weight of a training sample is changed by an adaptive boosting algorithm so as to construct a turbofan engine direct data-driven controller combining a plurality of basic learners into strong learners. Compared with the previous solution only adopting LS SVM, the present invention enhances the control precision, improves the generalization ability of the algorithm, and effectively solves the problem of sparsity of samples by the adaptive boosting method. By the adaptive boosting algorithm-based turbofan engine direct data-driven control method designed by the present invention.