阅读说明：本技术 气体涡轮引擎中的温度 (Temperature in gas turbine engines ) 是由 M.J.惠特尔 P.邓宁 R.M.汤斯 于 2019-07-31 设计创作，主要内容包括：本公开涉及气体涡轮引擎中的温度。提供了一种高效的气体涡轮引擎。该气体涡轮引擎的风扇经由齿轮箱从涡轮驱动,使得该风扇具有比该驱动涡轮更低的旋转速度,从而提供效率增益。该高效的风扇系统与具有低冷却流需求和/或高温性能并且对于给定功率可具有特别低质量的核心相配合。(The present disclosure relates to temperatures in gas turbine engines. An efficient gas turbine engine is provided. The fan of the gas turbine engine is driven from the turbine via a gearbox such that the fan has a lower rotational speed than the driven turbine, thereby providing an efficiency gain. The efficient fan system is compatible with cores that have low cooling flow requirements and/or high temperature performance and can have particularly low quality for a given power.)

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

an engine core (11) comprising:

a turbine (17, 19), a combustor (16) and a compressor (14, 15), the turbine comprising a first turbine (19) and a second turbine (17) and the compressor comprising a first compressor (14) and a second compressor (15);

a first spindle (26) connecting the first turbine to the first compressor;

a second spindle (27) connecting the second turbine to the second compressor, the second turbine, the second compressor and the second spindle being arranged to rotate at a higher rotational speed than the first spindle, the gas turbine engine further comprising:

a bypass duct located radially outside the engine core;

a fan (23) comprising a plurality of fan blades; and

a gearbox (30) receiving an input from the first spindle (26) and outputting a drive to the fan so as to drive the fan at a lower rotational speed than the first spindle, wherein:

a portion of the flow (C) entering the engine core bypasses the combustor and is used as a turbine cooling flow to cool the turbine;

the fan diameter is in the range of 225cm to 400 cm; and is

At cruise conditions, the cooling flow to bypass flow efficiency ratio is no greater than 0.02.

2. The gas turbine engine of claim 1, wherein the cooling flow to bypass flow efficiency ratio is in the range of 0.005 to 0.02, optionally 0.006 to 0.016, optionally 0.007 to 0.013.

3. The gas turbine engine of any preceding claim, wherein

The second turbine includes at least one ceramic matrix composite component.

4. The gas turbine engine of claim 5, wherein the mass of ceramic matrix composite in the second turbine is in the range of 2% to 15% of the total mass of the second turbine, and optionally in the range of 4% to 10% of the total mass of the second turbine.

5. The gas turbine engine (10) for an aircraft according to claim 3 or claim 4, wherein:

the first turbine includes at least one ceramic matrix composite component; and, optionally,

the mass of the ceramic matrix composite in the first turbine and the second turbine is in the range 1% to 15%, optionally 2% to 12% of the total mass of the first turbine and the second turbine.

6. The gas turbine engine (10) for an aircraft according to any one of the preceding claims, wherein:

the turbine comprises at least one row of stator blades (171); and is

The stator vanes (171) of the most axially upstream row are of a metal or ceramic matrix composite material.

7. The gas turbine engine (10) for an aircraft according to any one of the preceding claims, wherein:

the turbine includes at least one row of rotor blades (172); and is

The most axially upstream row of rotor blades (172) is a metal or ceramic matrix composite.

8. The gas turbine engine (10) of any one of the preceding claims, wherein:

the turbine includes at least one row of rotor blades (172), the most axially upstream row of rotor blades being radially surrounded by a seal segment (175); and is

The seal segment includes a ceramic matrix composite.

9. The gas turbine engine (10) of any one of the preceding claims, wherein:

the turbine comprises at least two rows of stator blades (171, 173); and is

The second most axially upstream row of stator vanes (173) comprises a ceramic matrix composite material.

10. The gas turbine engine (10) for an aircraft according to any one of the preceding claims, wherein:

the turbine includes at least two rows of rotor blades (174); and is

The second most axially upstream row of rotor blades (174) comprises a ceramic matrix composite material.

11. The gas turbine engine (10) for an aircraft of claim 10, wherein:

the second axially-most upstream row of rotor blades is radially surrounded by a ceramic matrix composite seal segment.

12. The gas turbine engine of any preceding claim, wherein the axially most upstream row of stator vanes (191) in the first turbine comprises a ceramic matrix composite.

13. The gas turbine engine of any preceding claim, wherein the axially most upstream row of rotor blades (192) in the first turbine comprises a ceramic matrix composite, the gas turbine engine optionally further comprising a ceramic matrix composite seal segment surrounding the axially most upstream row of rotor blades (192) in the first turbine.

14. The gas turbine engine according to any preceding claim, wherein a turbine inlet temperature is defined as the temperature at the inlet of the most axially upstream turbine rotor at maximum power conditions of the gas turbine engine, the turbine inlet temperature being at least 1800K, optionally at least 1850K, 1900K, 1950K or 2000K.

15. The gas turbine engine of any preceding claim, wherein:

the fan diameter is in the range of 250cm to 280cm or 325cm to 370 cm; and/or

The gear reduction ratio of the gearbox is in the range of 3.3 to 4; and/or

The maximum net thrust of the engine at sea level is in the range of 160kN to 550kN, optionally in the range of 160kN to 250kN or 300kN to 500 kN.