A method of operating a gas turbine engine. The method includes determining a first compressor non-dimensional rotational speed set point, determining a current first compressor rotational speed and first compressor inlet temperature, and transferring power between the first and second shafts such that the first shaft non-dimensional rotational speed matches the set point.

The invention concerns land-based gas turbine plants with a two-spool gas turbine arrangement for generating electrical power. The invention comprises two spools. The first spool comprises a first shaft, a first compressor and a first turbine. The second spool comprises at least a second shaft and a second turbine. The shafts of the first and the second spools are independently rotatable of each other. The invention further comprises the first generator and the second generator having nominally substantially equal power ratings, and the rotating parts of the first generator and the second generator have nominally substantially equal rotational speed ratings, and at least 60 percent of a total output power supplied to said load in a form of electrical and rotational power is generated by the two electrical generators.

A burner for a combustion chamber of a gas turbine with a mixing and injection device, which includes a limiting wall that defines a gas-flow channel and at least two streamlined bodies. Each streamlined body extends in a first transverse direction into the gas-flow channel, and has two lateral surfaces that are arranged essentially parallel to the main-flow direction. The lateral surfaces are joined to one another at their upstream and downstream sides to form leading and trailing edges of the body, respectively. At least one of the streamlined bodies includes a mixing structure and at least one fuel nozzle at its trailing edge for introducing at least one fuel essentially parallel to the main-flow direction into the flow channel. At least two of the streamlined bodies have different lengths along the first transverse direction such that they may be used for a can combustor.

Systems and methods are provided that use compressed gas to power a turbine, which in turn powers a generator, where an expansion of the compressed gas provides cooling for an electrical load that is powered by the generator.

A gas turbine engine includes: a compressor section including a compressor mean radius; a combustor section fluidly coupled downstream of the compressor section and include a combustor mean radius; and a turbine section fluidly coupled downstream of the combustor section and a turbine mid-span radius. The combustor mean radius is greater than each of the compressor mean radius and the turbine mid-span radius.

The present invention relates to a supercritical CO.sub.2 power generation system of a series recuperative type. According to an embodiment of the present invention, an inlet temperature of a turbine can be increased to increase a work of the turbine, thereby realizing a cycle design having improved turbine efficiency. Further, the number and diameter of pipes connected to a heat exchanger using an external heat source can be reduced to reduce the plumbing related costs, thereby improving economical efficiency.

A turbine engine assembly includes a turbine section including at plurality of turbine stages through which the gas flow expands to generate a mechanical power output. An inter-turbine burner between at least two of the plurality of turbine stages reheats the gas flow. A condenser extracts water from the gas flow exhausted from the turbine section, and an evaporator heats the water extracted by the condenser to generate a steam flow with the steam flow communicated to the inter-turbine burner and added to the gas flow expanded through the turbine section.

A gas turbine engine assembly including a tap that is at a location up stream of the combustor section for drawing a bleed airflow. An exhaust heat exchanger is configured to transfer thermal energy from the exhaust gas flow into the bleed airflow and communicate the heated bleed airflow into the turbine section where it is expanded to drive the turbine section.

Hybrid-electric powertrains for aircraft are disclosed herein. An example hybrid-electric powertrain includes a gas turbine propulsion engine including a first propulsor and a gas turbine engine to drive the first propulsor to produce thrust, a generator operably coupled to a drive shaft of the gas turbine engine, and an electric propulsion unit including a second propulsor and an electric motor to drive the second propulsor to produce thrust. During a first mode of operation, the gas turbine propulsion engine and the electric propulsion unit are activated to produce thrust, and the generator is driven by the gas turbine engine to produce electrical power to power the electric propulsion unit. During a second mode of operation, the gas turbine propulsion engine is activated to produce thrust and the electric propulsion unit is deactivated.

A system includes a gas turbine having a turbine stage disposed in a combustion gas path, wherein the turbine stage includes turbine vanes disposed upstream from turbine blades. The system includes an isothermal expansion system coupled to the turbine stage. The isothermal expansion system includes multi-fluid injectors configured to vary axial positions of combustion within a turbine stage expansion of the turbine stage to reduce temperature variations over the turbine stage expansion, wherein at least one of the multi-fluid injectors is coupled to each of the turbine vanes. Each of the multi-fluid injectors includes a fuel port configured to inject a fuel, an oxidant port configured to inject an oxidant, and a barrier fluid port configured to inject a barrier fluid between the fuel and the oxidant, wherein the barrier fluid is configured to delay mixing between the fuel and the oxidant.