IPC분류정보
국가/구분 |
United States(US) Patent
등록
|
국제특허분류(IPC7판) |
|
출원번호 |
US-0847640
(2007-08-30)
|
등록번호 |
US-8336289
(2012-12-25)
|
발명자
/ 주소 |
|
출원인 / 주소 |
- United Technologies Corporation
|
인용정보 |
피인용 횟수 :
9 인용 특허 :
16 |
초록
▼
Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assem
Gas turbine engine systems and related methods involving multiple gas turbine cores are provided. In this regard, a representative gas turbine engine includes: an inlet; a blade assembly mounted to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, each of the multiple gas turbine cores being independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.
대표청구항
▼
1. A gas turbine engine comprising: an inlet;a blade assembly mounted to a main shaft to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, wherein each of the gas turbine cores is independently operative in a first state, in which rotational e
1. A gas turbine engine comprising: an inlet;a blade assembly mounted to a main shaft to receive intake air via the inlet; and multiple gas turbine cores located downstream of the blade assembly, wherein each of the gas turbine cores is independently operative in a first state, in which rotational energy is provided to rotate the blade assembly via the main shaft, and a second state, in which rotational energy is not provided to rotate the blade assembly, and wherein each of the gas turbine cores comprises a compressor that is mechanically connected to the blade assembly in the first state. 2. The gas turbine engine of claim 1, wherein each of the gas turbine cores further comprises a combustion section, a turbine and a shaft interconnecting the turbine and the compressor such that rotational energy of the turbine is applied via the shaft to the compressor. 3. The gas turbine engine of claim 1, wherein: the shaft of each of the gas turbine cores is oriented parallel to the main shaft. 4. The gas turbine engine of claim 1, further comprising a clutch operative to selectively apply and remove rotational energy, imparted by at least one of the gas turbine cores, to the blade assembly. 5. The gas turbine engine of claim 4, further comprising a gearbox operative to apply rotational energy, imparted by the clutch, to the blade assembly. 6. The gas turbine engine of claim 1, further comprising a gearbox operative to apply rotational energy, imparted by at least one of the gas turbine cores, to the blade assembly. 7. The gas turbine engine of claim 1, further comprising an electrical generator operative to convert rotational energy of at least one of the gas turbine cores to electricity. 8. The gas turbine engine of claim 7, wherein the electrical generator is operative to generate electricity despite rotational energy of the corresponding gas turbine core not being applied to the blade assembly. 9. The gas turbine engine of claim 1, wherein the multiple gas turbine cores are annularly positioned about the longitudinal axis of the gas turbine engine. 10. The gas turbine engine of claim 1, wherein the blade assembly is a variable pitch blade assembly. 11. The gas turbine engine of claim 1, wherein: the gas turbine engine is a turbofan; andthe blade assembly comprises a fan. 12. The gas turbine engine of claim 1, wherein each the multiple gas turbine cores is configured with a single spool. 13. The gas turbine engine of claim 1, wherein: each compressor is mechanically connected to the main shaft in the first state. 14. A gas turbine engine, comprising: an inlet;a rotatable blade assembly that is positioned to receive intake air from the inlet, which rotatable blade assembly is mounted on a blade assembly shaft; anda plurality of gas turbine cores, each comprising: a compressor section;a combustion section operative to receive compressed gas from the compressor section;a shaft;a turbine section operative to impart rotational energy to the compressor section via the shaft; anda drive segment coupled to the shaft and operative to provide rotational energy from the shaft to the blade assembly shaft, the drive segment being offset with respect to a centerline of the blade assembly. 15. The gas turbine engine of claim 14, further comprising: a plurality of clutches, each interconnected with a respective one of the drive segments; anda gearbox operative to impart rotational energy, provided via one or more of the clutches, to the blade assembly. 16. The gas turbine engine of claim 14, wherein each of the gas turbine cores is independently operative in a first state, in which rotational energy is provided to rotate the blade assembly, and a second state, in which rotational energy is not provided to rotate the blade assembly.
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