Gas turbine engine having optimized fan

Patent metadata

Field	Value
Publication number	US-11268386-B2
Application number	US-201916554713-A
Country	US
Kind code	B2
Filing date	Aug 29, 2019
Priority date	Sep 4, 2018
Publication date	Mar 8, 2022
Grant date	Mar 8, 2022

Abstract

A gas turbine engine comprises carbon fibre fan blades. At cruise conditions, the fan tip air angle θ in the range: 64 degrees≤θ≤67 degrees. Additionally or alternatively, the fan blade tip angle β is in the range of from 62 to 69 degrees. Arrangements in accordance with the present disclosure provide advantages which may include improved bird-strike performance. This may allow advantages associated with carbon fibre fan blades to be better exploited.

First claim

We claim:

1. A gas turbine engine for an aircraft 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, wherein:
the fan blades comprise a carbon fibre composite material; and
at engine cruise conditions, a fan tip air angle θ is in the range: 64 degrees≤θ≤67 degrees, the fan tip air angle θ being defined as:

θ
=

tan

-
1

⁡

(

V
ThetaBladeTip

Vx
air

)

where
⁢

:

V
ThetaBladeTip

=



ω
·

D
2



;

ω is a fan rotational speed in radians/second;
D is a diameter of the fan in metres at a leading edge thereof; and
Vx air is a mean axial velocity of a flow into the fan over the leading edge.

2. The gas turbine engine according to claim 1 , wherein the fan is directly coupled to at least one turbine stage by a rigid shaft so as to rotate at the same rotational speed as the at least one turbine stage to which it is connected.

3. The gas turbine engine according to claim 1 , wherein:
a fan blade tip angle β is defined as an angle between a tangent to a leading edge of a camber line in a cross-section through a fan blade at 90% of a blade span from a root and a projection of an axial direction onto the cross-section, and the fan blade tip angle β is within 2 degrees of the fan tip air angle θ.

4. The gas turbine engine according to claim 1 , wherein:
a fan blade tip angle β is defined as an angle between a tangent to a leading edge of a camber line in a cross-section through a fan blade at a tip thereof and an axial direction, the fan blade tip angle β being in the range of from 62 to 69 degrees.

5. The gas turbine engine according to claim 4 , wherein the blade tip angle β is in the range of from 63 to 68 degrees.

6. The gas turbine engine according to claim 1 , wherein the fan blades comprise a main body attached to a leading edge sheath, a material of the main body and a material of the leading edge sheath being different .

7. The gas turbine engine according to claim 6 , wherein the material of the leading edge sheath has better impact resistance than the material of the main body material.

8. The gas turbine engine according to claim 6 , wherein the material of the leading edge sheath comprises titanium.

9. The gas turbine engine according to claim 6 , wherein the material of the main body comprises the carbon fibre composite material.

10. The gas turbine engine according to claim 1 , wherein a specific thrust is defined as net engine thrust divided by mass flow rate through the engine, and at the engine cruise conditions, the specific thrust is in the range of from 70 Nkg −1 s to 100 Nkg −1 s.

11. The gas turbine engine according to claim 1 , wherein a quasi-non-dimensional mass flow rate Q is defined as:

Q
=

W
⁢

T
⁢

⁢
0

P
⁢

⁢

0
·

A
fan

.

where:
W is mass flow rate through the fan in Kg/s;
T 0 is average stagnation temperature of air at a fan face in Kelvin;
P 0 is average stagnation pressure of the air at the fan face in Pa;
A fan is an area of the fan face in m 2 , and
at the engine cruise conditions:
0.029 Kgs −1 N −l K 1/2 ≤Q≤0.036 Kgs −1 N −1 K 1/2 ,

12. The gas turbine engine according to claim 11 , wherein the fan rotational speed at the engine cruise conditions is less than 2500 rpm.

13. The gas turbine engine according to claim 1 , wherein a fan tip loading is defined as dH/Utip 2 , where dH is an enthalpy rise across the fan and Utip is a translational velocity of the fan blades at a tip of the leading edge, and at the engine cruise conditions, 0 . 28 <dH/Utip 2 <0.35.

14. The gas turbine engine according to claim 1 , wherein:
a fan pressure ratio, defined as the ratio of a mean total pressure of a flow at a fan exit to a mean total pressure of a flow at a fan inlet, is no greater than 1.5 at the engine cruise conditions; and/or
a fan root pressure ratio, defined as the ratio of a mean total pressure of a 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 1.25 at the engine cruise conditions, wherein, optionally, the ratio between the fan root pressure ratio to a fan tip pressure ratio at the engine cruise conditions is no greater than 0.95, where the fan tip pressure ratio is defined as the ratio of a mean total pressure of a flow at the fan exit that subsequently flows through a bypass duct to the mean total pressure of the flow at the fan inlet.

15. The gas turbine engine according to claim 1 , 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 lower rotational speed than the first core shaft.

16. The gas turbine engine according to claim 1 , wherein a forward speed of the gas turbine engine at the cruise conditions is in the range of from a Mach number of 0.75 to a Mach number of 0.85.

17. The gas turbine engine according to claim 1 , wherein a forward speed of the gas turbine engine at the cruise conditions is a Mach number of 0.8 and the cruise conditions corr

Assignees

• Rolls Royce Plc

Inventors

• Benson Christopher
• Baralon Stephane M M

Classifications

• F01D5/282

  Selecting composite materials, e.g. blades with reinforcing filaments · CPC title

• F05D2240/307

  related to the tip of a rotor blade · CPC title

• F05D2250/711

  convex · CPC title

• F01D5/20

  Specially-shaped blade tips to seal space between tips and stator {(F01D5/225 takes precedence)} · CPC title

• F05D2220/36

  specially adapted for the fan of turbofan engines · CPC title