公开(公告)号：US20220099035A1

公开(公告)日：2022-03-31

申请号：US17466086

申请日：2021-09-03

Applicant: ROLLS-ROYCE PLC

Inventor： Craig W BEMMENT ,  Pascal DUNNING

IPC: F02C9/28 ,  F01D11/14 ,  F01D13/00 ,  F01D25/24

Abstract: A gas turbine engine for an aircraft includes an engine core including a first, lower pressure, turbine, a first compressor, and a first core shaft connecting the first turbine to the first compressor; and a second, higher pressure, turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor, and a fan located upstream of the engine core and including a plurality of fan blades extending from a hub. A turbine to fan tip temperature change ratio of a low pressure turbine temperature change to a fan tip temperature rise is in the range from 1.46 to 2.0.

公开(公告)号：US20210310408A1

公开(公告)日：2021-10-07

申请号：US17345328

申请日：2021-06-11

Applicant: ROLLS-ROYCE plc

Inventor： Pascal DUNNING ,  Michael J. WHITTLE ,  Roderick M. TOWNES

IPC: F02C3/073 ,  F02C7/36 ,  F02C3/107 ,  F01D5/28

Abstract: A highly efficient gas turbine engine includes the fan of the gas turbine engine driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

公开(公告)号：US20210310407A1

公开(公告)日：2021-10-07

申请号：US17345588

申请日：2021-06-11

Applicant: ROLLS-ROYCE PLC

Inventor： Craig W BEMMENT ,  Pascal DUNNING

IPC: F02C3/04 ,  F02C7/36

Abstract: A gas turbine engine for an aircraft comprises an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor; and a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across a radially outer portion of each fan blade at the trailing edge at cruise conditions. A core to fan tip temperature rise ratio is in the range from 2.845 to 3.8.

公开(公告)号：US20210222632A1

公开(公告)日：2021-07-22

申请号：US17143629

申请日：2021-01-07

Applicant: ROLLS-ROYCE PLC

Inventor： Craig W. BEMMENT ,  Pascal DUNNING

IPC: F02C9/00 ,  F02C3/04 ,  F02C7/36 ,  B64D27/10 ,  B64D31/00

Abstract: A gas turbine engine comprises a fan, a compressor, a low pressure turbine and a high pressure turbine. The fan diameter is greater than 250 cm and less than 381 cm; and the gas turbine engine has a first thrust at sea level static conditions and a second thrust at end of runway conditions; and a thrust take-off ratio greater than 1.32; wherein the thrust take-off ratio is the ratio of the first thrust to the second thrust.