公开(公告)号：US20250035037A1

公开(公告)日：2025-01-30

申请号：US18913201

申请日：2024-10-11

Applicant: ROLLS-ROYCE PLC

Inventor： Richard G. Stretton ,  Michael C. Willmot ,  Nicholas Grech

IPC: F02C3/06 ,  F01D17/16 ,  F02K1/00 ,  F02K3/06

Abstract: A gas turbine engine (10) for an aircraft comprises an engine core (11) comprising a turbine (19), a compressor (14), a core shaft (26), and a core exhaust nozzle (20), the core exhaust nozzle (20) having a core exhaust nozzle pressure ratio calculated using total pressure at the core nozzle exit (56); a fan (23) comprising a plurality of fan blades; and a nacelle (21) surrounding the fan (23) and the engine core (11) and defining a bypass duct (22), the bypass duct (22) comprising a bypass exhaust nozzle (18), the bypass exhaust nozzle (18) having a bypass exhaust nozzle pressure ratio calculated using total pressure at the bypass nozzle exit;
wherein a bypass to core ratio of: bypass ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio core ⁢ exhaust ⁢ nozzle ⁢ pressure ⁢ ratio is configured to be in the range from 1.1 to 2.0 under aircraft cruise conditions.

公开(公告)号：US11293346B2

公开(公告)日：2022-04-05

申请号：US16391626

申请日：2019-04-23

Applicant: ROLLS-ROYCE plc

Inventor： Michael I. Elliott ,  Peter Banister ,  Michael C. Willmot ,  Silvia Fernandez Arranz

IPC: F02C7/057 ,  F01D25/12 ,  F01D25/14 ,  F02C7/18 ,  F02C7/04 ,  F02C7/047

Abstract: There is provided an air intake system for providing air to a tip clearance control system. The air intake system comprises a ram-air intake having a scoop portion and a body portion. The body portion of the ram-air intake houses a heat exchanger.

公开(公告)号：US10882633B2

公开(公告)日：2021-01-05

申请号：US16741005

申请日：2020-01-13

Applicant: ROLLS-ROYCE plc

Inventor： Richard G. Stretton ,  Michael C. Willmot

IPC: F02C7/04 ,  B64D33/02 ,  B64D27/12 ,  F02K3/06

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio of: the ⁢ ⁢ turbine ⁢ ⁢ diameter ⁢ ⁢ at ⁢ ⁢ an ⁢ ⁢ axial location ⁢ ⁢ of ⁢ ⁢ the ⁢ ⁢ lowest ⁢ ⁢ pressure ⁢ ⁢ rotor ⁢ ⁢ stage a ⁢ ⁢ distance ⁢ ⁢ from ⁢ ⁢ a ⁢ ⁢ ground ⁢ ⁢ plane ⁢ ⁢ to ⁢ ⁢ the ⁢ ⁢ wing is in the range from 0.2 to 0.3.

公开(公告)号：US20200346779A1

公开(公告)日：2020-11-05

申请号：US16934767

申请日：2020-07-21

Applicant: ROLLS-ROYCE PLC

Inventor： Richard G. Stretton ,  Michael C. Willmot

IPC: B64D33/02 ,  B64D27/12 ,  F02C7/04 ,  F02K3/06

Abstract: A gas turbine engine of an aircraft includes: an engine core having a turbine including a lowest pressure rotor stage, a turbine diameter, a fan including a plurality of fan blades extending from a hub, an annular fan face at a leading edge of the fan; wherein a downstream blockage ratio is: the   turbine   diameter   at   an   axial   location of   the   lowest   pressure   rotor   stage ground   plane   to   wing   distance and a quasi-non-dimensional mass flow rate Q defined as: Q = W  T  0 P   0 · A flow where: W is mass flow rate through the fan in Kg/s; T0 is average stagnation temperature of the air at the fan face in Kelvin; P0 is average stagnation pressure of the air at the fan face in Pa; and Aflow is the flow area of the fan face in m2, and wherein a Q ratio of: the downstream blockage ratio×Q is in a range from 0.005 to 0.01.