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US 2026/0116550 A1Published application
Gas turbine engine compression system with core compressor pressure ratio
AssigneeROLLS-ROYC plc
InventorsGareth M. ARMSTRONG, Nicholas Howarth
Life of this patent
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Abstract
A gas turbine engine has a compression system radius ratio defined as the ratio of the radius of the tip of a fan blade to the radius of the tip of the most downstream compressor blade in the range of from 5 to 9. This results in an optimum balance between installation benefits, operability, maintenance requirements and engine efficiency when the gas turbine engine is installed on an aircraft.
Claims20
A gas turbine engine for an aircraft, the gas turbine engine comprising:
- an engine core comprising a turbine, a compressor, and a combustor;
- a fan comprising a plurality of fan blades;
- a bypass duct defined radially outside the core;
- and
- a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
- 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 to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 at cruise conditions;
- a fan root 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 to the mean total pressure of the flow at the fan inlet, is no greater than 1.24 at cruise conditions;
- and
- each fan blade is manufactured at least in part from carbon fibre.
6 more elements
A gas turbine engine for an aircraft, the gas turbine engine comprising:
- an engine core comprising a turbine, a compressor, and a combustor;
- a fan comprising a plurality of fan blades;
- a bypass duct defined radially outside the core;
- and
- a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
- 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 to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 at cruise conditions;
- a fan root 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 to the mean total pressure of the flow at the fan inlet, is no greater than 1.24 at cruise conditions;
- and
- the core compressor has exactly 12, 13 or 14 stages.
6 more elements
A gas turbine engine for an aircraft, the gas turbine engine comprising:
- an engine core comprising a turbine, a compressor, and a combustor;
- a fan comprising a plurality of fan blades;
- a bypass duct defined radially outside the core;
- and
- a gearbox that receives an input from a core shaft that is connected to at least a part of the turbine, the gearbox outputting drive to the fan so as to drive the fan at a lower rotational speed than the core shaft, wherein:
- 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 to the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct, is in a range 0.8 to 0.95 at cruise conditions;
- a fan root 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 to the mean total pressure of the flow at the fan inlet, is no greater than 1.24 at cruise conditions;
- and
- a bypass ratio is defined as the ratio of the mass flow rate of the flow through the bypass duct to the mass flow rate of the flow through the core is in a range of 10 to 11.5 at cruise conditions.
6 more elements
Dependencies are inferred from the claim text.