公开(公告)号：US20180231021A1

公开(公告)日：2018-08-16

申请号：US15894276

申请日：2018-02-12

Applicant: ROLLS-ROYCE plc

Inventor： Mark J. WILSON ,  Gabriel GONZALEZ-GUTIERREZ ,  Marco BARALE ,  Benedict PHELPS ,  Kashmir S. JOHAL ,  Nigel HS SMITH

IPC: F04D29/38 ,  F02K3/06

Abstract: A fan blade is provided with an aerofoil portion for which, for cross-sections through the aerofoil portion at radii between 15% and 25% of the blade span from the root radius, the average leading edge thickness is greater than the leading edge thickness at the tip. The geometry of the fan blade may result in a lower susceptibility to flutter.

公开(公告)号：US20180231018A1

公开(公告)日：2018-08-16

申请号：US15894240

申请日：2018-02-12

Applicant: ROLLS-ROYCE plc

Inventor： Nigel HS SMITH ,  Mark J. WILSON ,  Gabriel GONZALEZ-GUTIERREZ ,  Marco BARALE ,  Benedict PHELPS ,  Kashmir S. JOHAL

IPC: F04D29/38 ,  F02K3/06

CPC classification number: F04D29/384 ,  F01D5/141 ,  F01D5/26 ,  F02K3/06 ,  F04D29/388 ,  F05D2220/32 ,  F05D2220/36 ,  F05D2240/301 ,  F05D2240/303 ,  Y02T50/672 ,  Y02T50/673

Abstract: A fan blade is provided with an aerofoil portion for which, at radii between 20% and 40% of the blade span, the location of the position of maximum thickness along the camber line is at less than a defined percentage of the total length of the camber line. For all cross-sections through the aerofoil portion at radii greater than 70% of the blade span, the location of the position of maximum thickness along the camber line is at more than a defined percentage of the total length of the camber line. The geometry of the fan blade may result in a lower susceptibility to flutter.

公开(公告)号：US20200011273A1

公开(公告)日：2020-01-09

申请号：US16411294

申请日：2019-05-14

Applicant: ROLLS-ROYCE plc

Inventor： Stephane M. M. BARALON ,  Mark J. WILSON ,  Benedict R. PHELPS

IPC: F02K3/06 ,  F02K3/02

Abstract: A gas turbine engine system has an engine core and a bypass duct. A fan drives the flow through the bypass duct. A bypass efficiency is defined as the efficiency of the fan compression of the bypass flow. The bypass efficiency is a function of the bypass flow rate at a given set of conditions. The fan bypass inlet mass flow rate at the reference operating point is appreciably higher than the mass flow rate through the bypass duct at the peak bypass efficiency at a given fan reference rotational speed and cruise conditions. This results in increased design flexibility and improved overall engine performance.

公开(公告)号：US20190162071A1

公开(公告)日：2019-05-30

申请号：US16170707

申请日：2018-10-25

Applicant: ROLLS-ROYCE plc

Inventor： Mark J. WILSON ,  Stephane M. M. BARALON ,  Benedict PHELPS

IPC: F01D5/16 ,  F01D5/14

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades 139 in which the camber distribution relative to covered passage of the fan 13 allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

公开(公告)号：US20190063368A1

公开(公告)日：2019-02-28

申请号：US16106813

申请日：2018-08-21

Applicant: ROLLS-ROYCE plc

Inventor： Benedict R. PHELPS ,  Mark J. WILSON ,  Gabriel GONZALEZ-GUTIERREZ ,  Nigel HS SMITH ,  Marco BARALE ,  Kashmir S. JOHAL ,  Stephane MM BARALON ,  Craig W. BEMMENT

IPC: F02K3/06 ,  F02C3/06 ,  F01D5/14

Abstract: A gas turbine engine 10 is provided in a fan root to tip pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the engine core (P102) to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct (P104), is no greater than a certain value. The gas turbine engine 10 may provide improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

公开(公告)号：US20220074308A1

公开(公告)日：2022-03-10

申请号：US17412759

申请日：2021-08-26

Applicant: ROLLS-ROYCE plc

Inventor： Benjamin MOHANKUMAR ,  Mark J. WILSON ,  Cesare A. HALL

IPC: F01D5/14 ,  B64D27/10 ,  B64D27/24

Abstract: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric S, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, and L/D is between 0.2 and 0.45.

公开(公告)号：US20190063369A1

公开(公告)日：2019-02-28

申请号：US16106853

申请日：2018-08-21

Applicant: ROLLS-ROYCE plc

Inventor： Benedict R. PHELPS ,  Mark J. WILSON ,  Gabriel GONZALEZ-GUTIERREZ ,  Nigel HS SMITH ,  Marco BARALE ,  Kashmir S. JOHAL ,  Stephane MM BARALON ,  Craig W. BEMMENT

IPC: F02K3/06 ,  F02C3/06 ,  F01D5/14

Abstract: A gas turbine engine 10 is provided in which a fan having fan blades in which the camber distribution along the span allows the gas turbine engine to operate with improved efficiency when compared with conventional engines, whilst retaining an acceptable flutter margin.

公开(公告)号：US20180231020A1

公开(公告)日：2018-08-16

申请号：US15894249

申请日：2018-02-12

Applicant: ROLLS-ROYCE plc

Inventor： Mark J. WILSON ,  Gabriel GONZALEZ-GUTIERREZ ,  Marco BARALE ,  Benedict PHELPS ,  Kashmir S. JOHAL ,  Nigel HS SMITH

IPC: F04D29/38 ,  F02K3/06

CPC classification number: F04D29/386 ,  F01D5/141 ,  F01D5/147 ,  F02K3/06 ,  F04D29/324 ,  F04D29/388 ,  F05D2220/32 ,  F05D2230/239 ,  F05D2240/303 ,  Y02T50/672 ,  Y02T50/673

Abstract: A fan blade for a gas turbine engine is arranged such that for any two points on its leading edge that are in the radially outer 40% of the blade span and are radially separated by at least 5% of the blade span, the radially outer of the two points is axially forward of the radially inner point. The radius of the leading edge of a given fan blade at the hub divided by the radius of the leading edge of the fan blade at the tip is less than or equal to 0.3. Such an arrangement may result in an improved operability range.

公开(公告)号：US20220162957A1

公开(公告)日：2022-05-26

申请号：US17453501

申请日：2021-11-04

Applicant: ROLLS-ROYCE plc

Inventor： Benjamin MOHANKUMAR ,  Mark J. WILSON ,  Cesare A. HALL

IPC: F01D15/12 ,  F01D25/24

Abstract: An aircraft engine comprising a fan, the fan having a diameter D and including a plurality of fan blades, the fan blades having a sweep metric Stip, each fan blade having a leading edge, and a forward-most portion on the leading edge of each fan blade being in a first reference plane. The aircraft engine further comprises a nacelle, comprising an intake portion forward of the fan, a forward edge on the intake portion being in a second reference plane, wherein the intake portion has a length L measured along an axis of the aircraft engine between the first reference plane and the second reference plane, the aircraft engine having a cruise design point condition Mrel, wherein Mrel is between 0.4 and 0.93, L/D is between 0.2 and 0.45 and Stip is from −1 to 0.1.

公开(公告)号：US20210404342A1

公开(公告)日：2021-12-30

申请号：US17352468

申请日：2021-06-21

Applicant: ROLLS-ROYCE plc

Inventor： Benedict R. PHELPS ,  Stephane M M BARALON ,  Mark J. WILSON

IPC: F01D5/14 ,  F02K3/06

Abstract: A fan blade for a gas turbine engine has a covered passage. A cross section through the fan blade at a point along the blade span is defined as having particular change in angle (α3−α1) of the camber line between the leading edge and the trailing edge and/or between the leading edge and the point on the camber line that corresponds to the start of the covered passage.