A gas turbine engine for an aircraft includes a compressor, a combustion chamber, and a turbine having at least one stator, and at least one rotor. Each stator and rotor is formed by a plurality of blades, a fluid channel is formed between two consecutive blades, and each blade has two opposing surfaces. The compressor is in fluid communication with a first group of stator channels, and the combustion chamber is in fluid communication with a second group of stator channels, such that heat exchange can be performed through two opposing surfaces of at least one stator blade. The outer and the inner walls define a duct for the passage of the heated fluid through the rotor blades, and the outer wall is also arranged for directing the compressed air towards the combustion chamber.

A (micro) gas turbine arrangement includes a gas turbine device having a combustor system, a turbine driven by an exhaust gas stream of the combustor system, and a compressor for supplying the combustor system with a compressed oxidant stream, as well as a recuperator for transferring at least a portion of the thermal power of the exhaust gas stream of the turbine to the compressed oxidant stream. At least one bypass diverts at least a portion of the oxidant stream or the exhaust gas stream around at least one heat exchanger of the recuperator, and at least one control element for adjusting the flow through the at least one bypass, to be able to adapt the quantity of heat emitted by the gas turbine arrangement at the design point, and thus to be able to improve the efficiency of a power-heat cogeneration system having such a gas turbine arrangement.

A (micro) gas turbine arrangement includes a gas turbine device having a combustor system, a turbine driven by an exhaust gas stream of the combustor system, and a compressor for supplying the combustor system with a compressed oxidant stream, as well as a recuperator for transferring at least a portion of the thermal power of the exhaust gas stream of the turbine to the compressed oxidant stream. At least one bypass diverts at least a portion of the oxidant stream or the exhaust gas stream around at least one heat exchanger of the recuperator, and at least one control element for adjusting the flow through the at least one bypass, to be able to adapt the quantity of heat emitted by the gas turbine arrangement at the design point, and thus to be able to improve the efficiency of a power-heat cogeneration system having such a gas turbine arrangement.

A gas turbine engine includes a rotatable first shaft, a first disk connected to the first shaft, a second disk connected to the first shaft, a combustor radially outward from the first disk and the second disk, and a heat exchanger connected to the combustor aft of the second disk. The first disk includes a row of low pressure compressor blades and a row of high pressure turbine blades connected to a radially outer end of the row of low pressure compressor blades. The second disk includes a row of high pressure compressor blades and a row of low pressure turbine blades connected to a radially outer end of the row of high pressure compressor blades.

A gas turbine engine includes a rotatable first shaft, a first disk connected to the first shaft, a second disk connected to the first shaft, a combustor radially outward from the first disk and the second disk, and a heat exchanger connected to the combustor aft of the second disk. The first disk includes a row of low pressure compressor blades and a row of high pressure turbine blades connected to a radially outer end of the row of low pressure compressor blades. The second disk includes a row of high pressure compressor blades and a row of low pressure turbine blades connected to a radially outer end of the row of high pressure compressor blades.

A hybrid-electric propulsion system includes a gas turbine engine, an electric machine coupled to and rotatably driven by the gas turbine engine to produce AC electric power, an energy storage system, and a propulsion unit. The gas turbine engine includes a combustor and a recuperator that places an exhaust air flow that is downstream from the combustor in a heat exchange relationship with a compressed air flow that is upstream from the combustor to transfer thermal energy from the exhaust flow to the compressed flow. The propulsion unit includes a fan and an electric motor rotably coupled to the fan, the electric motor being driven by electric power from one of the electric machine or the energy storage system.

A hybrid-electric propulsion system includes a gas turbine engine, an electric machine coupled to and rotatably driven by the gas turbine engine to produce AC electric power, an energy storage system, and a propulsion unit. The gas turbine engine includes a combustor and a recuperator that places an exhaust air flow that is downstream from the combustor in a heat exchange relationship with a compressed air flow that is upstream from the combustor to transfer thermal energy from the exhaust flow to the compressed flow. The propulsion unit includes a fan and an electric motor rotably coupled to the fan, the electric motor being driven by electric power from one of the electric machine or the energy storage system.

A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, an inlet duct, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter D, and the heat exchanger module comprises a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into the fan assembly. In use, the first fluid has a maximum temperature of 80° C., and the heat exchanger module transfers at least 300 kW of heat energy from the first fluid to the airflow.

A turbofan gas turbine engine comprises, in axial flow sequence, a heat exchanger module, an inlet duct, a fan assembly, a compressor module, and a turbine module. The fan assembly comprises a plurality of fan blades defining a fan diameter D, and the heat exchanger module comprises a plurality of heat transfer elements for transfer of heat from a first fluid contained within the heat transfer elements to an airflow passing over a surface of the heat transfer elements prior to entry of the airflow into the fan assembly. In use, the first fluid has a maximum temperature of 80° C., and the heat exchanger module transfers at least 300 kW of heat energy from the first fluid to the airflow.

A container-type compressed air storage power generation device (2) comprises compressors (5a-5c); a tank (8); power generators (9a-9c); a control device (12); and a container (4). The compressors (5a-5c) compress air. The tank (8) is driven by air supplied from the compressors (5a-5c). The power generators (9a-9c) are driven by air supplied from the tank (8). The control device drives and controls the compressors (5a-5c) and the power generators (9a-9c). The container (4) houses the compressors (5a-5c) and the power generators (9a-9c), and the tank (8) is disposed outside the container (4). Therefore, the container-type compressed air storage power generation device (2) is easy to transport and construct on-site.