A gas turbine engine is provided that includes an engine core and a bypass duct. The engine core includes a compressor section, a combustor section, a turbine section and a core flowpath. The compressor section includes a radial flow compressor rotor. The core flowpath extends through the compressor section, the combustor section and the turbine section from a core inlet to a core exhaust. The bypass duct includes a bypass flowpath that extends outside of the engine core from a bypass inlet to a bypass exhaust. The bypass inlet is disposed along the compressor section and fluidly coupled with core flowpath.

An apparatus and method for a generator for generating power having a rotary union between a rotor and a stator, the rotary union comprising a first fluid passage having a converging nozzle and located within the stator having a first inlet and a first outlet, a second fluid passage having a diverging nozzle and located within the rotor and having a second inlet and a second outlet, the second inlet is aligned with the first outlet and at least one gap located between the rotor and stator.

A radial flow turbine heat engine includes a compressor, a recuperator, a combustor and a turbine. A compressor outlet manifold collects compressed gas from the compressor through a plurality of compressor outlets. A turbine inlet manifold supplies combustion gas to the turbine through a plurality of turbine inlets. The compressor outlet manifold comprises a plurality of compressor outlet manifold ducts and the turbine inlet manifold comprises a plurality of turbine inlet manifold ducts. These manifold ducts are circumferentially interdigitated with respect of each other around the shaft of the turbine to provide a flow path for compressed gas through the recuperator located radially inwardly with respect to the rotation axis of the shaft compared to the flow path for the combustion gas in the hot side portion of the heat engine. Furthermore, a flow path for the combustion gas is provided to the turbine inlets which are radially inward with respect to the rotation axis of the shaft compared to flow paths for the compressed gas within the compressor outlet manifold proximal to those turbine inlets.

A gas turbine engine for an aircraft such as a UAV includes a compressor connected to a turbine with a combustor to produce a hot gas stream. The rotor is supported by two radial foil bearings. An axial thrust bearing assembly is positioned between the compressor disk and the turbine disk and includes an axial thrust bearing radial disk extending from a hollow axial tube. Compressed air is bled off from the compressor and passed into an axial thrust balance chamber to provide the axial thrust balance for the rotor. The compressed air from the thrust bearing chamber then flows through both of the radial foil bearings for cooling, is collected in and around the hollow tube, and then discharged into the inlet of the turbine. An orifice can be adjusted to meter and control a pressure occurring in the thrust balance chamber.

A disclosed micro gas turbine system includes a micro gas turbine apparatus and an extracting cycle apparatus. The micro gas turbine apparatus includes a first compressor, a burner, and a first turbine. The first turbine expands a combustion gas generated by the burner. The extracting cycle apparatus includes a second compressor and a second turbine. The second compressor receives a flow of extracted air that is generated by extracting a part of a working fluid discharged from the first compressor. The second turbine expands the working fluid discharged from the second compressor. The working fluid discharged from the second turbine cools down the first turbine.

An apparatus and method for a generator for generating power having a rotary union between a rotor and a stator, the rotary union comprising a first fluid passage having a converging nozzle and located within the stator having a first inlet and a first outlet, a second fluid passage having a diverging nozzle and located within the rotor and having a second inlet and a second outlet, the second inlet is aligned with the first outlet and at least one gap located between the rotor and stator.

A core engine article includes a combustor liner defining a combustion chamber therein and a turbine nozzle. The combustor liner includes a plurality of injector ports, and the plurality of injector ports have a shape that tapers to a corner on a forward side of the injector ports. The turbine nozzle includes a plurality of airfoils. The combustor liner and turbine nozzle are integral with one another. A method of making a core engine article is also disclosed.

A thermodynamic apparatus (10) comprising a compressor module (100), a turbine module (200), and a regenerative heat exchanger (300) centred on a central axis (12). The compressor module (100), turbine module (200) and regenerative heat exchanger (300) are arranged in series along the central axis (12) such that the regenerative heat exchanger (300) is provided between the compressor module (100) and the turbine module (200).

The invention concerns land-based gas turbine plants with a multi-spool gas turbine arrangement for generating electrical power to supply a load (200). The invention comprises at least three spools (10a-10c). Each of the at least three spools (10a-10c) comprises a shaft (11a-11c), a compressor (C1-C3) and a turbine (T1-T3). Each one of the shafts (11a-11c) of the at least three spools (10a-10c) are independently rotatable with respect to each other. The invention further comprises electrical generators (G1-G3) mounted on each of the shafts (11a-11c) of the at least three spools (10a-10c), the output power of the generators being independently controllable and at least 60 percent of a total output power supplied to said load (200) in a form of electrical and rotational power is generated by the at least three generators (G1-G3) in the form of electrical energy.

A radial flow turbine heat engine includes a compressor, a recuperator, a combustor and a turbine. A compressor outlet manifold collects compressed gas from the compressor through a plurality of compressor outlets. A turbine inlet manifold supplies combustion gas to the turbine through a plurality of turbine inlets. The compressor outlet manifold comprises a plurality of compressor outlet manifold ducts and the turbine inlet manifold comprises a plurality of turbine inlet manifold ducts. These manifold ducts are circumferentially interdigitated with respect of each other around the shaft of the turbine to provide a flow path for compressed gas through the recuperator located radially inwardly with respect to the rotation axis of the shaft compared to the flow path for the combustion gas in the hot side portion of the heat engine. Furthermore, a flow path for the combustion gas is provided to the turbine inlets which are radially inward with respect to the rotation axis of the shaft compared to flow paths for the compressed gas within the compressor outlet manifold proximal to those turbine inlets.