A turbomachine engine can include a fan assembly, a vane assembly, a core engine, a gearbox, and an overall engine efficiency rating. The fan assembly can include a plurality of fan blades. The vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include a low-pressure turbine. The gearbox includes an input and an output. The input of the gearbox is coupled to the low-pressure turbine of the core engine and comprises a first rotational speed, the output of the gearbox is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 3.2-4.0. The overall engine efficiency rating is within a range of 0.57-8.0.

A turbomachine engine can include a fan assembly, a pitch change mechanism, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The pitch change mechanism can be coupled to the fan assembly. The vane assembly can include a plurality of vanes. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

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.

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 gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

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 method of optimizing probe placement in a turbomachine is disclosed which includes determining wavenumber (Wn) of N dominant wavelets generated by upstream and downstream stators and blade row interactions formed around an annulus, establishing a design matrix A utilized in developing flow properties around the annulus having a dimension of m×(2N+1), iteratively modifying probe positions placed around the annulus and 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 optimal probe position, wherein the condition number is defined as norm A.Math.norm A+, wherein A+ represents inverse of A for a square matrix and a Moore-Penrose pseudoinverse of A for a rectangular matrix.

A method for reconstructing nonuniform circumferential flow in a turbomachine is disclosed which includes receiving one or more wavenumbers of interest, receiving positional information for a plurality of circumferential positions of a plurality of instrumentation probes, receiving signals from the plurality of instrumentation probes to generate a spatially under-sampled data, and from the spatially under-sampled data determining a multi-wavelet approximation reconstructing circumferential flow field.

A turbomachine engine can include an unducted fan assembly, an unducted vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The unducted fan assembly can include a plurality of fan blades. The unducted vane assembly can include a plurality of vanes, and the vanes can, in some instances, be disposed aft of the fan blades. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the unducted fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.

A turbomachine engine can include a fan assembly, a pitch change mechanism, a vane assembly, a core engine, a gearbox, and a gearbox efficiency rating. The fan assembly can include a plurality of fan blades. The pitch change mechanism can be coupled to the fan assembly. The vane assembly can include a plurality of vanes. The core engine can include one or more compressor sections and one or more turbine sections. The gearbox includes an input and an output. The input is coupled to the one or more turbine sections of the core engine and comprises a first rotational speed, the output is coupled to the fan assembly and has a second rotational speed, and a gear ratio of the first rotational speed to the second rotational speed is within a range of 4.1-14.0. The gearbox efficiency rating is 0.10-1.8.