Thermal management assembly for a hybrid-electric aircraft propulsion system

A method of thermally managing an electrical assembly of a hybrid-electric aircraft propulsion system comprising a gas turbine engine and a nacelle, the gas turbine engine having an axis, the gas turbine engine comprising a first rotational assembly and an engine static structure, the first rotational assembly comprising a first shaft, a bladed first compressor rotor, and a bladed first turbine rotor, the first shaft interconnecting the bladed first compressor rotor and the bladed first turbine rotor, the engine static structure housing the first rotational assembly, the nacelle forming an exterior aerodynamic housing of the hybrid-electric aircraft propulsion system and circumscribing the gas turbine engine, the nacelle forming an annular bypass duct radially between the engine static structure and the nacelle, the method comprising:

• circulating a coolant through a control unit cooling circuit in fluid communication with a control unit of the electrical assembly, the control unit being electrically connected to an electric machine, the electric machine being mechanically coupled with the first rotational assembly, the control unit and the control unit cooling circuit being disposed within a bifurcation extending radially through the annular bypass duct and connecting the nacelle and the engine static structure;
• directing a bypass air flow from a propulsor to the bifurcation;
• and
• transferring heat from the coolant to the bypass air flow at the bifurcation;
• wherein the bifurcation comprises a bifurcation body housing the control unit and a thermal management assembly including the control unit cooling circuit, the bifurcation body forming an outer side surface of the bifurcation, the control unit comprising a plurality of plate fins extending from the control unit through the bifurcation body to the outer side surface, and the plurality of plate fins disposed at the control unit cooling circuit within the control unit.

A hybrid-electric aircraft propulsion system comprising:

• a propulsor;
• an engine coupled with the propulsor, the engine comprising an engine static structure;
• a nacelle forming an exterior aerodynamic housing the engine;
• a bifurcation extending between the engine static structure and the nacelle, the bifurcation including a bifurcation body extending between a leading edge and a trailing edge, the bifurcation body comprising a first circumferential side wall and a second circumferential side wall extending from the leading edge toward the trailing edge;
• an electrical assembly comprising a first electric machine and a first control unit, the first electric machine being coupled with the propulsor, the first control unit being electrically connected to the first electric machine, the first control unit disposed within the bifurcation between the first circumferential side wall and the second circumferential side wall;
• and
• a thermal management assembly comprising a control unit cooling circuit, a coolant pump, and a coolant tank, the control unit cooling circuit connected in fluid communication with and formed by the first control unit, the control unit cooling circuit disposed within the bifurcation between the first circumferential side wall and the second circumferential side wall, the coolant pump and the coolant tank connected in fluid communication with and further forming the control unit cooling circuit, the coolant pump and the coolant tank disposed within the bifurcation.