公开(公告)号：WO2014107202A3

公开(公告)日：2014-09-18

申请号：PCT/US2013063341

申请日：2013-10-04

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L

IPC: F02C7/057 ,  F02C3/04 ,  F02K3/04 ,  F02K3/075

CPC classification number: F02K3/075 ,  F02K3/06

Abstract: A gas turbine engine is typically comprised of a fan stage, multiple compressor stages, and multiple turbine stages. These stages are made up of alternating rotating blade rows and static vane rows. The total number of blades and vanes is the airfoil count. An overall pressure ratio is greater than 30. A bypass ratio is greater than 8. A stage ratio is the product of the bypass ratio and the overall pressure ratio divided by the number of stages. An airfoil ratio is that product divided by the airfoil count. The stage ratio is greater than or equal to 22 and/or the airfoil ratio is greater than or equal to.12.

Abstract translation: 燃气涡轮发动机通常包括风扇级，多个压缩机级和多个涡轮级。 这些阶段由交替旋转的叶片排和静叶片排组成。 叶片和叶片的总数是翼型数量。 总压比大于30.旁路比大于8.级比是旁路比和总压比的乘积除以级数。 翼型比是该产品除以翼型计数。 级比大于或等于22和/或翼型比大于或等于12。

公开(公告)号：CA2886267A1

公开(公告)日：2014-07-10

申请号：CA2886267

申请日：2013-10-04

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L

IPC: F02C7/057 ,  F02C3/04 ,  F02K3/04 ,  F02K3/075

Abstract: A gas turbine engine is typically comprised of a fan stage, multiple compressor stages, and multiple turbine stages. These stages are made up of alternating rotating blade rows and static vane rows. The total number of blades and vanes is the airfoil count. An overall pressure ratio is greater than 30. A bypass ratio is greater than 8. A stage ratio is the product of the bypass ratio and the overall pressure ratio divided by the number of stages. An airfoil ratio is that product divided by the airfoil count. The stage ratio is greater than or equal to 22 and/or the airfoil ratio is greater than or equal to.12.

公开(公告)号：WO2015050619A2

公开(公告)日：2015-04-09

申请号：PCT/US2014046781

申请日：2014-07-16

Applicant: UNITED TECHNOLOGIES CORP

Inventor： SCHWARZ FREDERICK M ,  HASEL KARL L ,  MERRY BRIAN D

IPC: F02C7/00 ,  F02C7/32

CPC classification number: F02C3/06 ,  F01D5/14 ,  F02C3/107 ,  F02C7/36 ,  F02K3/06 ,  F05D2220/32 ,  F05D2220/36 ,  F05D2240/30 ,  F05D2240/303 ,  F05D2240/307 ,  F05D2260/40311 ,  F16H1/28

Abstract: A ratio of an outer diameter of a fan hub at a leading edge of the blades to an outer tip diameter of the blades at the leading edge is greater than or equal to about 0.24 and less than or equal to about 0.38. The fan tip diameter is greater than or equal to about 84 inches (213.36 centimeters) and a fan tip speed is less than or equal to about 1050 ft / second (320.04 meters / second). A bypass ratio, a gear ratio and an AN2 value are also claimed. The fan drive turbine has between three and six stages.

Abstract translation: 在叶片的前缘处的风扇轮毂的外径与前缘处的叶片的外端直径的比率大于或等于约0.24且小于或等于约0.38。 风扇叶尖直径大于或等于约84英寸（213.36厘米），风扇叶尖速度小于或等于约1050英尺/秒（320.04米/秒）。 还要求旁路比，齿轮比和AN2值。 风扇驱动涡轮机有三到六级。

公开(公告)号：WO2013154639A3

公开(公告)日：2014-01-03

申请号：PCT/US2013021831

申请日：2013-01-17

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L ,  SMITH PETER G ,  OCHS STUART S

IPC: F02K3/02

CPC classification number: F01D17/162 ,  F04D29/563 ,  F05D2220/36 ,  F05D2240/121

Abstract: A turbofan engine includes a fan driven by a low pressure turbine through a gear reduction. The gear reduction has a gear ratio of greater than or equal to about 2.4. The low pressure turbine has an expansion ratio greater than or equal to about 5. The fan has a bypass ratio greater than or equal to about 8. In other features, a turbofan engine includes a variable geometry fan exit guide vane (FEGV) system having a multiple of circumferentially spaced radially extending fan exit guide vanes. Rotation of the fan exit guide vanes between a nominal position and a rotated position selectively changes a fan bypass flow path to permit efficient operation at various flight conditions.

Abstract translation: 涡轮风扇发动机包括由低压涡轮机通过齿轮减速器驱动的风扇。 齿轮减速器具有大于或等于约2.4的齿轮比。 低压涡轮机具有大于或等于约5的膨胀比。风扇具有大于或等于约8的旁通比。在其它特征中，涡轮风扇发动机包括可变几何风扇出口导叶（FEGV）系统，其具有 周向间隔开的径向延伸的风扇出口导叶的多个。 风扇出口导向叶片在标称位置和旋转位置之间的旋转选择性地改变风扇旁路流动路径以允许在各种飞行条件下的有效操作。

公开(公告)号：CA2862597C

公开(公告)日：2018-12-11

申请号：CA2862597

申请日：2014-09-10

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L

IPC: F02C7/36 ,  F02C7/00

Abstract: An example gas turbine engine includes, among other things, a geared architecture rotatably coupled to the fan drive shaft, and a high pressure compressor. The gas turbine engine is configured so that a core temperature at an exit of the high-pressure compressor is approximately in a range of about 1150 to about 1350 degrees Fahrenheit at take-off. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to a primary stream exhaust velocity, is approximately in a range of about 0.75 to about 0.90. A Bypass Ratio of the engine is greater than about 8Ø

公开(公告)号：CA2849372C

公开(公告)日：2015-12-08

申请号：CA2849372

申请日：2013-09-23

Applicant: UNITED TECHNOLOGIES CORP

Inventor： SHERIDAN WILLIAM G ,  HASEL KARL L

IPC: F02C7/36 ,  F16H1/32

Abstract: A gas turbine engine according to an exemplary aspect of the present disclosure includes, among other things, a fan section including a fan rotatable about an axis and a speed reduction device in communication with the fan. The speed reduction device includes a star drive gear system with a star gear ratio of at least 1.5. A fan blade tip speed of the fan is less than 1400 fps.

公开(公告)号：CA2862597A1

公开(公告)日：2015-04-03

申请号：CA2862597

申请日：2014-09-10

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L

IPC: F02C7/36 ,  F02C7/00

Abstract: An example gas turbine engine includes, among other things, a geared architecture rotatably coupled to the fan drive shaft, and a high pressure compressor. The gas turbine engine is configured so that a core temperature at an exit of the high-pressure compressor is approximately in a range of about 1150 to about 1350 degrees Fahrenheit at take-off. The gas turbine engine is configured so that an Exhaust Velocity Ratio, defined by a ratio of a fan stream exhaust velocity to a primary stream exhaust velocity, is approximately in a range of about 0.75 to about 0.90. A Bypass Ratio of the engine is greater than about 8Ø

公开(公告)号：RU2676150C1

公开(公告)日：2018-12-26

申请号：RU2017137647

申请日：2017-10-27

Applicant: UNITED TECHNOLOGIES CORP

Inventor： SHERIDAN WILLIAM G ,  HASEL KARL L

IPC: F02K3/068 ,  F02C7/36

Abstract: Газотурбинныйдвигательсодержит, средипрочего, вентиляторнуюсекцию, содержащуювентилятор, выполненныйс возможностьювращениявокругоси, иредуктор, взаимодействующийс указаннымвентилятором. Указанныйредукторсодержитэпициклическуюзубчатуюпередачуприводавентиляторас передаточнымотношениемпланетарнойпередачипоменьшеймере 2,5. Скоростьконцалопастивентиляторасоставляетменее 1400 футовв секунду. Турбиннаясекциянизкогодавлениявзаимодействуетс указаннымредуктором. Причемуказаннаятурбиннаясекциянизкогодавленияимееттрииличетыреступении степеньдвухконтурности, лежащуюв интервалеот 11,0 до 22,0. Достигаетсяповышениеэффективностиэксплуатации. 2 н. и 13 з.п. ф-лы, 3 ил.

公开(公告)号：CA2923326A1

公开(公告)日：2016-09-20

申请号：CA2923326

申请日：2016-03-08

Applicant: UNITED TECHNOLOGIES CORP

Inventor： SCHWARZ FREDERICK M ,  HASEL KARL L

IPC: F02C7/36

Abstract: A gas turbine engine includes a fan with a plurality of fan blades rotatable about an axis, and a compressor section that includes at least first and second compressor sections. An average exit temperature of the compressor section is between about 1000 °F and about 1500 °F. The engine also includes a combustor that is in fluid communication with the compressor section, and a turbine section that is in fluid communication with the combustor. A geared architecture is driven by the turbine section for rotating the fan about the axis.

公开(公告)号：CA2850225A1

公开(公告)日：2013-07-04

申请号：CA2850225

申请日：2012-12-26

Applicant: UNITED TECHNOLOGIES CORP

Inventor： HASEL KARL L ,  STAUBACH JOSEPH B ,  MERRY BRIAN D ,  SUCIU GABRIEL L ,  DYE CHRISTOPHER M

IPC: F01D1/04

Abstract: A gas turbine engine includes a fan section, a gear arrangement configured to drive the fan section, a compressor section and a turbine section. The compressor section includes a low pressure compressor section and a high pressure compressor section. The turbine section is configured to drive compressor section and the gear arrangement. An overall pressure ratio, which is provided by a combination of a pressure ratio across said low pressure compressor section and a pressure ratio across said high pressure compressor section, is greater than about 35. The pressure ratio across the low pressure compressor section is between about 3 and about 8 whereas the pressure ratio across the high pressure compressor section is between about 7 and about 15.