Patent ID: 11898489
Assignee: ROLLS-ROYCE PLC
Field: Engines, pumps, turbines (Mechanical engineering)
Classification: CPC F  Y | IPC F

Claim 0:
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, wherein a compressor exit temperature is defined as an average temperature of airflow at the exit from the compressor at cruise conditions and a core entry temperature is defined as an average temperature of airflow entering the engine core at cruise conditions, and a core compressor temperature rise is defined as:, the
    ⁢
       
    compressor
    ⁢
       
    exit
    ⁢
       
    temperature
   
   
    the
    ⁢
       
    core
    ⁢
       
    entry
    ⁢
       
    temperature
   
  
  ;, the engine core further comprising an annular splitter at which the flow is divided between a core flow that flows through the engine core, and a bypass flow that flows along a bypass duct, wherein stagnation streamlines around the circumference of the engine, stagnating on a leading edge of the annular splitter, form a streamsurface forming a radially inner boundary of a streamtube that contains all of the bypass flow; and
a fan located upstream of the engine core, the fan comprising a plurality of fan blades extending from a hub, each fan blade having a leading edge and a trailing edge, each fan blade having a radially outer portion lying within the streamtube that contains the bypass flow, and wherein a fan rotor entry temperature is defined as an average temperature of airflow across the leading edge of each fan blade at cruise conditions and a fan tip rotor exit temperature is defined as an average temperature of airflow across the radially outer portion of each fan blade at the trailing edge at cruise conditions and a fan tip temperature rise is defined as:, the
    ⁢
       
    fan
    ⁢
       
    tip
    ⁢
       
    rotor
    ⁢
       
    exit
    ⁢
       
    temperature
   
   
    the
    ⁢
       
    fan
    ⁢
       
    rotor
    ⁢
       
    entry
    ⁢
       
    temperature
   
  
  ,
 

and wherein the gas turbine engine is configured to have a core compressor to fan tip temperature rise ratio of:, the
   ⁢
      
   core
   ⁢
      
   compressor
   ⁢
      
   temperature
   ⁢
      
   rise
  
  
   the
   ⁢
      
   fan
   ⁢
      
   tip
   ⁢
      
   temperature
   ⁢
      
   rise
  
 

in the range from 2.67 to 3.8,
the compressor exit temperature in the range of from 750 K to 1050 K, and
wherein the cruise conditions means the conditions at mid-cruise of the aircraft to which the gas turbine engine is attached, and the cruise conditions at mid-cruise means the conditions experienced by the aircraft and the gas turbine engine at a point between top of climb and start of descent at which 50% of the total fuel that is burned between the top of climb and the start of descent has been burned.