Geared turbofan engine with high compressor exit temperature
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
F02K-003/06
F02C-003/13
F02C-007/36
F02C-003/06
F02C-003/10
출원번호
US-0950393
(2015-11-24)
등록번호
US-10119466
(2018-11-06)
발명자
/ 주소
Schwarz, Frederick M.
Hasel, Karl L.
출원인 / 주소
United Technologies Corporation
대리인 / 주소
Carlson, Gaskey & Olds, P.C.
인용정보
피인용 횟수 :
0인용 특허 :
26
초록▼
A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit tempera
A gas turbine engine comprises a fan includes a plurality of fan blades rotatable about an axis. A compressor section includes at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between about 1000° F. and about 1500° F. A combustor is in fluid communication with the compressor section. A turbine section is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.
대표청구항▼
1. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit t
1. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor;a geared architecture driven by the turbine section for rotating the fan about the axis, wherein the geared architecture is positioned upstream of a low pressure turbine and downstream of a compressor rotor of the first compressor section driven by the low pressure turbine; andan inducer forming an additional compression section positioned upstream of the first and second compressor sections. 2. The gas turbine engine according to claim 1 wherein the average exit temperature is between 1100° F. and 1450° F. 3. The gas turbine engine according to claim 1 wherein the first compressor section comprises a low pressure compressor and the second compressor section comprises a high pressure compressor, and wherein the turbine section comprises a low pressure turbine that drives the low pressure compressor via a first shaft and a high pressure turbine that drives the high pressure compressor via a second shaft, and wherein the geared architecture couples the first shaft to the fan. 4. The gas turbine engine as set forth in claim 1 wherein the average exit temperature is defined at sea level, end of takeoff power and at a rated thrust for the gas turbine engine. 5. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section and a second compressor section, wherein the first compressor section comprises a low pressure compressor and the second compressor section comprises a high pressure compressor, and wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;wherein the high pressure compressor includes a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration, and wherein the first blade and disk configuration comprises a plurality of slots to receive the plurality of blades and including a plurality of rim cavities for honeycomb seals, and wherein the second blade and disk configuration comprises integrally formed blades such that there are no rim cavities or associated honeycomb seals;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor; anda geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the turbine section comprises a low pressure turbine that drives the low pressure compressor via a first shaft and a high pressure turbine that drives the high pressure compressor via a second shaft, and wherein the geared architecture couples the first shaft to the fan. 6. The gas turbine engine according to claim 5 wherein the first stage is positioned forward of the second stage. 7. The gas turbine engine as set forth in claim 5, wherein the geared architecture is positioned intermediate the fan and a compressor rotor driven by a low pressure turbine. 8. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section and a second compressor section, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, wherein the turbine section comprises three separate turbine sections including a fan drive turbine, an intermediate pressure turbine, and a secondary pressure turbine;the intermediate pressure turbine driving a first compressor rotor of the first compressor section;the secondary pressure turbine driving a second compressor rotor of the second compressor section; anda geared architecture driven by the fan drive turbine for rotating the fan about the axis. 9. The gas turbine engine according to claim 8 wherein the first compressor section comprises a low pressure compressor and the second compressor section comprises a high pressure compressor, and wherein the fan drive turbine drives the geared architecture via a shaft, and wherein the intermediate pressure turbine is configured to drive the first compressor rotor, and the secondary pressure turbine is configured to drive the second compressor rotor. 10. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;the second compressor section including a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration, and wherein the first blade and disk configuration comprises a plurality of slots to receive the plurality of blades and including a plurality of rim cavities for honeycomb seals, and wherein the second blade and disk configuration comprises integrally formed blades such that there are no rim cavities or associated honeycomb seals;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, wherein the turbine section comprises a first turbine section that drives the first compressor section via a first shaft at the first speed and a second turbine section that drives the second compressor section via a second shaft at the second speed; anda geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the geared architecture couples the first shaft to the fan. 11. The gas turbine engine according to claim 10 wherein the average exit temperature of the second compressor section is between 1100° F. and 1450° F. 12. The gas turbine engine according to claim 10 wherein the first stage is positioned forward of the second stage. 13. The gas turbine engine according to claim 10 wherein the average exit temperature is defined at Sea Level, end of takeoff power and at a rated thrust for the gas turbine engine. 14. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;the second compressor section including a plurality of stages with each stage comprising a disk with a plurality of blades extending radially outwardly from a rim of the disk, and wherein the plurality of stages includes at least a first stage having a first blade and disk configuration and a second stage having a second blade and disk configuration that is different than the first blade and disk configuration;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor, wherein the turbine section comprises a first turbine section that drives the first compressor section via a first shaft at the first speed and a second turbine section that drives the second compressor section via a second shaft at the second speed;a geared architecture driven by the turbine section for rotating the fan about the axis, and wherein the geared architecture couples the first shaft to the fan; andan inducer forming an additional compression section positioned in front of the first and second compressor sections. 15. The gas turbine engine according to claim 14 wherein the inducer is configured to rotate at a speed common with that of the fan. 16. A gas turbine engine comprising: a fan including a plurality of fan blades rotatable about an axis;a compressor section including at least a first compressor section rotating at a first speed and a second compressor section rotating at a second speed greater than the first speed, wherein components of the second compressor section are configured to operate at an average exit temperature that is between 1000° F. and 1500° F.;a combustor in fluid communication with the compressor section;a turbine section in fluid communication with the combustor;a geared architecture driven by the turbine section for rotating the fan about the axis;an inducer forming an additional compression section positioned in front of the first and second compressor sections, wherein the inducer is configured to rotate at a higher speed than the fan through an additional output of the geared architecture.
Curley Robert C. (White Marsh MD) Fisher Mark T. (Bel Air MD) Dileonardi James V. (Baltimore MD) DePinho ; Jr. A. Norton (Baldwin MD), Light weight fan blade containment system.
Allmon Barry L. (Maineville OH) Tongeman Kevin B. (Cincinnati OH), Low pressure drop radial inflow air-oil separating arrangement and separator employed therein.
Brodell Robert F. (Marlborough CT) Hovan Edward J. (Manchester CT) Selfors Steven T. (Somerville MA) Loffredo Constantino V. (Newington CT) Duesler Paul W. (Manchester CT), Nacelle and mounting arrangement for an aircraft engine.
Duesler Paul W. ; Loffredo Constantino V. ; Prosser ; Jr. Harold T. ; Jones Christopher W., Variable area fan exhaust nozzle having mechanically separate sleeve and thrust reverser actuation systems.
Rey Nancy M. ; Miller Robin M. ; Tillman Thomas G. ; Rukus Robert M. ; Kettle John L. ; Dunphy James R. ; Chaudhry Zaffir A. ; Pearson David D. ; Dreitlein Kenneth C. ; Loffredo Constantino V. ; Wyno, Variable area nozzle for gas turbine engines driven by shape memory alloy actuators.
※ AI-Helper는 부적절한 답변을 할 수 있습니다.