Granted Patent B2

US 12,196,104 · App. 18/496,251 · Granted Jan 14, 2025

Inventors: Michael Bartholomew (Oakham, GB); David Higginbottom (Leicester, GB)

Assignee: Rolls-Royce PLC

F01D5/147F01D5/30F02C7/32F05D2220/36

Abstract

Claims (34)

1. A fan blade for a gas turbine engine, the fan blade comprising:

an aerofoil portion comprising a leading edge extending from a root to a tip, a distance between the leading edge at the root and the leading edge at the tip defining a blade span, wherein:

a leading edge thickness is defined as a thickness of a cross-section at a given radius along the blade span at a location along a camber line that is 9% of a total length of the camber line from the leading edge;

for cross-sections through the aerofoil portion at radii between 50% and 70% of the blade span from a root radius, the leading edge thickness comprises a first maximum value;

for cross-sections through the aerofoil portion at radii greater than 70% of the blade span from the root radius, the leading edge thickness comprises a second maximum value; and

the second maximum value is between 105% and 125% of the first maximum value.

2. The fan blade of claim 1 , wherein:

for cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius, the leading edge thickness comprises a first average value;

for cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius, the leading edge thickness comprises a second average value; and

the second average value is between 105% and 125% of the first average value.

3. The fan blade of claim 1 , wherein:

for all cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius, the leading edge thickness comprises a first constant value;

for all cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius, the leading edge thickness comprises a second constant value; and

the second constant value is between 105% and 125% of the first constant value.

4. The fan blade of claim 1 , wherein:

for all cross-sections through the aerofoil portion at radii greater than 80% of the blade span from the root radius, the leading edge thickness is greater than 105% and less than 125% of the leading edge thickness for all cross-sections through the aerofoil portion at radii between 60% and 70% of the blade span from the root radius.

5. The fan blade of claim 1 , wherein:

for cross-sections through the aerofoil portion at radii between 70% and 80% of the blade span from the root radius, the leading edge thickness increases linearly with respect to the blade span.

6. The fan blade of claim 1 , further comprising:

a platform; and

a root portion;

wherein the root portion extends between the platform and the root of the aerofoil portion.

7. The fan blade of claim 6 , wherein a radial extent of the root portion is less than or equal to 7% of the blade span.

8. The fan blade of claim 1 , further comprising a tip portion that extends at least radially away from the tip of the aerofoil portion.

9. The fan blade of claim 8 , wherein a radial extent of the tip portion is less than or equal to 7% of the blade span.

10. A gas turbine engine for an aircraft, the gas turbine engine comprising:

an engine core comprising a turbine, a compressor, and a core shaft connecting the turbine to the compressor;

a fan located upstream of the engine core, the fan comprising a plurality of fan blades; and

a gearbox that receives an input from the core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the core shaft;

wherein each of the plurality of fan blades are the fan blade of claim 1 .

11. The gas turbine engine of claim 10 , wherein:

the turbine is a first turbine, the compressor is a first compressor, and the core shaft is a first core shaft;

the engine core further comprises a second turbine, a second compressor, and a second core shaft connecting the second turbine to the second compressor; and

the second turbine, the second compressor, and the second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.

Assignments (1)

Priority Claims (1)

Continuity (1)

References Cited (19)