US 12,351,330 B2
Aircraft propulsion system with reduced rotational load of an engine shaft
Akinori Kita, Wako (JP); Takeshi Matsumoto, Wako (JP); and Daiko Tsutsumi, Wako (JP)
Assigned to HONDA MOTOR CO., LTD., Tokyo (JP)
Filed by HONDA MOTOR CO., LTD., Tokyo (JP)
Filed on Nov. 30, 2021, as Appl. No. 17/537,546.
Claims priority of application No. 2020-212826 (JP), filed on Dec. 22, 2020.
Prior Publication US 2022/0194577 A1, Jun. 23, 2022
Int. Cl. B64D 35/02 (2024.01); B64D 27/10 (2006.01); B64D 27/24 (2006.01); B64D 31/00 (2006.01); B64U 30/20 (2023.01); B64U 50/32 (2023.01); B64U 10/14 (2023.01); B64U 50/19 (2023.01)
CPC B64D 35/02 (2013.01) [B64D 27/10 (2013.01); B64D 27/24 (2013.01); B64D 31/00 (2013.01); B64U 30/20 (2023.01); B64U 50/32 (2023.01); B64U 10/14 (2023.01); B64U 50/19 (2023.01)] 5 Claims
OG exemplary drawing
 
1. An aircraft propulsion system, including:
a plurality of gas turbine engines attached to an airframe of an aircraft;
an electric generator, wherein the electric generator is connected to each engine shaft of each of the gas turbine engines;
a plurality of electric motors driven using electric power generated by the electric generator;
a plurality of rotors attached to the airframe of the aircraft and driven using driving force respectively output by the plurality of electric motors; and
a controller which controls an operation state of the plurality of gas turbine engines,
in which, when a flight state of the aircraft is a first state after the plurality of gas turbine engines operate and the aircraft takes off, the controller causes some of the plurality of gas turbine engines to operate while stopping the remaining engines,
wherein the plurality of gas turbine engines have a compressor that compresses intake air, a compression chamber disposed downstream of the compressor, and a combustor that is supplied with compressed air from the compression chamber,
wherein at least two gas turbine engines of the plurality of gas turbine engines includes:
a bleed air hole provided in the compressor;
a bleed air pipe extending from the bleed air hole of one gas turbine engine of the at least two gas turbine engines to the compression chamber of another gas turbine engine of the at least two gas turbine engines; and
a bleed air valve provided in the bleed air pipe,
when the another gas turbine engine which has stopped is caused to start up again, the controller opens the bleed air valve in the one gas turbine engine which is in operation and supplies the compressed air of the compressor in the one gas turbine engine which is in operation to the compression chamber of the another gas turbine engine which has stopped through the bleed air hole and the bleed air pipe,
wherein the compression chamber contains the combustor,
wherein a first gas turbine engine and a second gas turbine engine are provided as the plurality of gas turbine engines,
wherein the aircraft propulsion system further includes:
a first bleed air pipe of the first gas turbine engine, which connects the compressor of the first gas turbine engine and the compression chamber of the second gas turbine engine;
a second bleed air pipe of the second gas turbine engine, which connects the compressor of the second gas turbine engine and the compression chamber of the first gas turbine engine;
a first bleed air valve, which is only one provided in the first bleed air pipe and which switches communication and interruption between the compressor of the first gas turbine engine and the compression chamber of the second gas turbine engine; and
a second bleed air valve, which is only one provided in the second bleed air pipe and which switches communication and interruption between the compressor of the second gas turbine engine and the compression chamber of the first gas turbine engine.