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 control system for limiting a power turbine torque of a gas turbine engine is disclosed. In various embodiments, the control system includes an engine control module configured to output an effector command signal to a gas generator of the gas turbine engine; a power turbine governor module configured to output to the engine control module a power turbine torque request signal; and a power turbine torque limiter module configured to output to the power turbine governor module a power turbine speed rate signal to limit a power turbine speed overshoot of the gas turbine engine.

A gas turbine engine defining a centerline and a circumferential direction, the gas turbine engine including: a turbomachine comprising a compressor section, a combustion section, and a turbine section arranged in serial flow order, the turbomachine defining a working gas flowpath and a fan duct flowpath; a primary fan driven by the turbomachine defining a primary fan tip radius R.sub.1 and a primary fan hub radius R.sub.2; a secondary fan located downstream of the primary fan and driven by the turbomachine, at least a portion of an airflow from the primary fan configured to bypass the secondary fan, the secondary fan defining a secondary fan tip radius R.sub.3 and a secondary fan hub radius R.sub.4, wherein the secondary fan is configured to provide a fan duct airflow through the fan duct flowpath during operation to generate a fan duct thrust, wherein the fan duct thrust is equal to % Fn.sub.3S of a total engine thrust during operation of the gas turbine engine at a rated speed during standard day operating conditions; wherein a ratio of R.sub.1 to R.sub.3 equals $\sqrt{\left(\mathrm{EFP}\right)\frac{\left(1-Rq{R}_{\mathrm{Sec}.-{\mathrm{Fan}}^2}\right)}{\left(1-Rq{R}_{\mathrm{Prim}.-\mathrm{Fa}{n}^2}\right)}\left(\frac{1}{\%{\mathrm{Fn}}_{3S}}-1\right)};$

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 comprising a plurality of fan blades extending from a hub. First and second turbine entrance and exit temperatures are defined as average temperature of airflow at the entrance or exit to the respective turbine at cruise conditions. A low pressure turbine temperature change is defined as:

[00001]$\frac{\mathrm{the}.Math..Math.\mathrm{first}.Math..Math.\mathrm{turbine}.Math..Math.\mathrm{entrance}.Math..Math.\mathrm{temperature}}{\mathrm{the}.Math..Math.\mathrm{first}.Math..Math.\mathrm{turbine}.Math..Math.\mathrm{exit}.Math..Math.\mathrm{temperature}}.$

A high pressure turbine temperature change is defined as:

[00002]

A gas turbine is provided, the gas turbine engine including a turbomachine having an inlet splitter defining in part an inlet to a working gas flowpath and a fan duct splitter defining in part an inlet to a fan duct flowpath. The gas turbine engine also includes a primary fan driven by the turbomachine defining a primary fan tip radius R1, a primary fan hub radius R2, and a primary fan specific thrust rating TP; and a secondary fan downstream of the primary fan and driven by the turbomachine, the secondary fan defining a secondary fan tip radius R3, a secondary fan hub radius R4, and a secondary fan specific thrust rating TS; wherein the gas turbine engine defines an Effective Bypass Area, and wherein a ratio of R1 to R3 equals $\frac{R_1}{R_3}=\sqrt{\left(\mathrm{EFP}\right)\frac{\left(1-{\mathrm{RqR}}_{\left(\mathrm{Sec}.-\mathrm{Fan}\right)}^2\right)}{\left(1-{\mathrm{RqR}}_{\mathrm{Prim}.-\mathrm{Fan}}^2\right)}\left(\frac{T_P}{T_S}\right)\left(\mathrm{EBA}\right)}.$

A variable displacement exhaust turbocharger according to at least one embodiment comprises: a plurality of nozzle vanes; an annular drive ring; and a lever plate fitted at one end in a groove portion disposed in the drive ring via a connecting pin portion and connected at another end to each nozzle vane. The value of dimensionless slippage S between the groove portion and the connecting pin portion is equal to or less than 0.0016. The dimensionless slippage S is represented by a certain equation.

A gas turbine engine is provided. The gas turbine engine includes: a turbomachine having a compressor section, a combustion section, and a turbine section arranged in serial flow order, the compressor section having a high pressure compressor defining a high pressure compressor exit area (A.sub.HPCExit) in square inches; wherein the gas turbine engine defines a redline exhaust gas temperature (EGT) in degrees Celsius, a total sea level static thrust output (Fn.sub.Total) in pounds, and a corrected specific thrust, wherein the corrected specific thrust is greater than or equal to 42 and less than or equal to 90, the corrected specific determined as follows: Fn.sub.TotalEGT/(A.sub.HPCExit.sup.21000).

A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.

A turbomachine engine includes an engine core including a high-pressure compressor, a high-pressure turbine, and a combustion chamber in flow communication with the high-pressure compressor and the high-pressure turbine. The engine core has a length (L.sub.CORE), and the high-pressure compressor has an exit stage diameter (D.sub.CORE). A high-pressure shaft is coupled to the high-pressure compressor and the high-pressure turbine. The high-pressure shaft is characterized by a high-speed shaft rating (HSR) from 1.5 to 6.2, and a ratio of L.sub.CORE/D.sub.CORE is from 2.1 to 4.3.