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US 11,698,030 B2Granted
Geared gas turbine engine
AssigneeROLLS-ROYCE plc
InventorsNicholas Howarth, Gareth M Armstrong
Life of this patent
Estimating the term…
Abstract
A gas turbine engine has a compression system blade ratio defined as the ratio of the height of a fan blade to the height of the most downstream compressor blade in the range of from 45 to 95. 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 comprising:
- an engine core comprising a turbine, a compressor, and a combustor;
- a fan comprising a plurality of fan blades, each fan blade having a fan blade height defined as the radius of the leading edge at the tip of the blade minus the radius of the point where the leading edge intersects the radially inner gas-washed hub;
- 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:
- the blade height of the most downstream compressor blade is defined as the radius of the leading edge at the tip of the blade minus the radius of the point where the leading edge intersects the radially inner gas-washed surface;
- a compression system blade ratio defined as the ratio of the fan blade height to the height of the most downstream compressor blade is in the range of from 45 to 70;
- and
- the temperature measured at the exit to the combustor is in the range of from 1950 k to 2000 k at a maximum take-off condition.
5 more elements
A gas turbine engine for an aircraft comprising:
- an engine core comprising a turbine, a compressor, and a combustor;
- a fan comprising a plurality of fan blades, each fan blade having a fan blade height defined as the radius of the leading edge at the tip of the blade minus the radius of the point where the leading edge intersects the radially inner gas-washed hub;
- 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:
- the blade height of the most downstream compressor blade is defined as the radius of the leading edge at the tip of the blade minus the radius of the point where the leading edge intersects the radially inner gas-washed surface;
- a compression system blade ratio defined as the ratio of the fan blade height to the height of the most downstream compressor blade is in the range of from 45 to 70;
- and
- the temperature measured at the exit to the combustor is in the range of from 1600 k to 1650 k at cruise conditions.
5 more elements
Dependencies are inferred from the claim text.