Efficient aircraft engine

Abstract

A highly efficient gas turbine engine is a system wherein the fan of the gas turbine engine is driven from a turbine via a gearbox, such that the fan has a lower rotational speed than the driving turbine, thereby providing efficiency gains. The efficient fan system is mated to a core that has low cooling flow requirements and/or high temperature capability, and which may have particularly low mass for a given power.

1 . A gas turbine engine for an aircraft comprising:
 an engine core comprising:   a turbine, a combustor, and a compressor, the turbine comprising a first turbine and a second turbine and the compressor comprising a first compressor and a second compressor;   a first core shaft connecting the first turbine to the first compressor;   a second core shaft connecting the second turbine to the second compressor, the second turbine, second compressor, and second core shaft being arranged to rotate at a higher rotational speed than the first core shaft, the gas turbine engine further comprising:
 a bypass duct radially outside the engine core; 
 a fan comprising a plurality of fan blades; and 
 a gearbox that receives an input from the first core shaft and outputs drive to the fan so as to drive the fan at a lower rotational speed than the first core shaft, wherein: 
 part (C) of the flow (B) that enters the engine core bypasses the combustor and is used as turbine cooling flow to cool the turbine; 
 the turbine entry temperature, defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, is in the range of from 1800K to 2100K; and 
 at cruise conditions, the cooling to bypass flow efficiency ratio is no greater than 0.02. 
   
     
     
         2 . A gas turbine engine according to  claim 1 , wherein the cooling to bypass efficiency ratio is in the range of from between 0.005 and 0.02. 
     
     
         3 . A gas turbine engine according to  claim 1 , wherein the cooling to bypass efficiency ratio is in the range of from between 0.006 and 0.016. 
     
     
         4 . A gas turbine engine according to  claim 1 , wherein the cooling to bypass efficiency ratio is in the range of from between 0.007 and 0.013. 
     
     
         5 . A gas turbine engine according to  claim 1 , wherein:
 the second turbine comprises at least one ceramic matrix composite component.   
     
     
         6 . A gas turbine engine according to  claim 5 , wherein the mass of ceramic matrix composite in the second turbine is in the range of from 2% to 15% of the total mass of the second turbine, and optionally in the range of from 4% to 10% of the total mass of the second turbine. 
     
     
         7 . A gas turbine engine for an aircraft according to  claim 5 , wherein:
 the first turbine comprises at least one ceramic matrix composite component; and, optionally,   the mass of ceramic matrix composite in the first and second turbines is in the range of from 1% to 15%, optionally 2% to 12%, of the total mass of the first and second turbines.   
     
     
         8 . A gas turbine engine for an aircraft according to  claim 1 , wherein:
 the turbine comprises at least one row of stator vanes; and   the most axially upstream row of stator vanes are metallic or ceramic matrix composite.   
     
     
         9 . A gas turbine engine for an aircraft according to  claim 1 , wherein:
 the turbine comprises at least one row of rotor blades; and   the most axially upstream row of rotor blades are metallic or ceramic matrix composite.   
     
     
         10 . A gas turbine engine according to  claim 1 , wherein:
 the turbine comprises at least one row of rotor blades, the most axially upstream row of rotor blades being radially surrounded by seal segments; and   the seal segments comprise a ceramic matrix composite.   
     
     
         11 . A gas turbine engine according to  claim 1 , wherein:
 the turbine comprises at least two rows of stator vanes; and   the second most axially upstream row of stator vanes comprise a ceramic matrix composite.   
     
     
         12 . A gas turbine engine for an aircraft according to  claim 1 , wherein:
 the turbine comprises at least two rows of rotor blades; and   the second most axially upstream row of rotor blades comprise a ceramic matrix composite.   
     
     
         13 . A gas turbine engine for an aircraft according to  claim 12 , wherein:
 the second most axially upstream row of rotor blades is radially surrounded by ceramic matrix composite seal segments.   
     
     
         14 . A gas turbine engine according to  claim 1 , wherein the axially most upstream row of stator vanes in the first turbine comprise a ceramic matrix composite. 
     
     
         15 . A gas turbine engine according to  claim 1 , wherein the axially most upstream row of rotor blades in the first turbine comprise a ceramic matrix composite, the gas turbine engine optionally further comprising ceramic matrix composite seal segments surrounding the axially most upstream row of rotor blades in the first turbine. 
     
     
         16 . A gas turbine engine according to  claim 1 , wherein the turbine entry temperature, defined as the temperature at the inlet to the most axially upstream turbine rotor at a maximum power condition of the gas turbine engine, at least 1850K, 1900K, 1950K or 2000K. 
     
     
         17 . A gas turbine engine according to  claim 1 , wherein the fan diameter is in the range of from 225 cm to 400 cm, optionally in the range of from 250 cm to 280 cm, or 325 to 370 cm. 
     
     
         18 . A gas turbine engine according to  claim 1 , wherein the gear reduction ratio of the gearbox is in the range of from 3.3 to 4. 
     
     
         19 . A gas turbine engine according to  claim 1 , wherein the maximum net thrust of the engine at sea level is in the range of from 160 kN to 550 kN, optionally in the range of from 160 kN to 250 kN or 300 kN to 500 kN.