High efficiency gas turbine engine

Abstract

A gas turbine engine has a quasi-non-dimensional mass flow rate in a defined range and a specific thrust in a defined range to achieve improved over all performance, taking into account fan operability and/or bird strike requirements as well as engine efficiency. The defined ranges of quasi-non-dimensional mass flow rate and specific thrust may be particularly beneficial for gas turbine engines in which the fan is driven by a turbine through a gearbox.

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, an annular fan face being defined at a leading edge of the fan; 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:   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; 
         T0 is average stagnation temperature of the air at the fan face in Kelvin; 
         P0 is average stagnation pressure of the air at the fan face in Pa; 
         A fan  is the area of the fan face in m 2 ; 
         a specific thrust is defined as net engine thrust divided by mass flow rate through the engine; and 
         at engine cruise conditions: 
         0.029 Kgs −1 N −1 K 1/2 ≤Q<0.036 Kgs −1 N −1 K 1/2 ; and 
         70 Nkg −1 s ≤specific thrust ≤110 
       
     
     
         2 . A gas turbine engine according to  claim 1 , wherein at cruise conditions, 0.03 Kgs −1 N −1 K 1/2 ≤Q≤0.035 Kgs −1 N −1 K 1/2 . 
     
     
         3 . A gas turbine engine according to  claim 1 , wherein at cruise conditions, 031 Kgs −1 N −1 K 1/2 ≤Q≤0.034 Kgs −1 N −1 K 1/2 . 
     
     
         4 . A gas turbine engine according to  claim 1 , wherein at cruise conditions, the specific thrust is less than 100 Nkg −1 s. 
     
     
         5 . A gas turbine engine according to  claim 1 , wherein a fan tip loading is defined as dH/Utip 2 , where dH is the enthalpy rise across the fan and Utip is the translational velocity of the fan blades at the tip of the leading edge, and at cruise conditions, 0.28 Jkg −1 K −1 /(ms −1 ) 2 <dH/Utip 2 <0.36 Jkg −1 K −1 /(ms −1 ) 2 . 
     
     
         6 . A gas turbine engine according to  claim 1 , wherein a fan pressure ratio, defined as the ratio of the mean total pressure of the flow at the fan exit to the mean total pressure of the flow at the fan inlet, is no greater than 1.5 at cruise conditions. 
     
     
         7 . A gas turbine engine according to any  claim 1 , 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:
 a fan root pressure ratio, defined as the ratio of the mean total pressure of the 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 cruise conditions.   
     
     
         8 . A gas turbine engine according to  claim 7 , wherein:
 a fan tip pressure ratio is defined as the ratio of the mean total pressure of the flow at the fan exit that subsequently flows through the bypass duct to the mean total pressure of the flow at the fan inlet; and   the ratio between the fan root pressure ratio to the fan tip pressure ratio at cruise conditions is less than 0.95.   
     
     
         9 . A gas turbine engine according to  claim 1 , wherein the ratio of the radius of fan blade at its hub to the radius of the fan blade at its tip is less than 0.33. 
     
     
         10 . A gas turbine engine according to  claim 1 , wherein the fan blades comprise a main body attached to a leading edge sheath, the main body and the leading edge sheath being formed using different materials. 
     
     
         11 . A gas turbine engine according to  claim 10 , wherein the leading edge sheath material comprises titanium and/or the main body material comprises carbon fibre or an aluminium alloy. 
     
     
         12 . A gas turbine engine according to  claim 1 , further comprising an intake that extends upstream of the fan blades, wherein:
 an intake length L is defined as the axial distance between the leading edge of the intake and the leading edge of the tip of the fan blades;   the fan diameter D is the diameter of the fan at the leading edge of the tips of the fan blades; and the ratio L/D is in the range of from 0.2 to 0.45.   
     
     
         13 . A gas turbine engine according to  claim 1 , wherein the gearbox has a reduction ratio in the range of from 3.1 to 3.7. 
     
     
         14 . A gas turbine engine according to  claim 1 , wherein the forward speed of the gas turbine engine at the cruise conditions is in the range of from Mn 0.75 to Mn 0.85. 
     
     
         15 . A gas turbine engine according to  claim 1 , wherein the forward speed of the gas turbine engine at the cruise conditions is Mn 0.8. 
     
     
         16 . A gas turbine engine according to  claim 1 , wherein the cruise conditions correspond to atmospheric conditions at an altitude that is in the range of from 10500 m to 11600 m. 
     
     
         17 . A gas turbine engine according to  claim 1 , wherein the cruise conditions correspond to atmospheric conditions at an altitude of 11000 m. 
     
     
         18 . A gas turbine engine according to  claim 1 , wherein the cruise conditions correspond to a forward Mach number of 0.8;
 a pressure of 23000 Pa; and   a temperature of −55 deg C.   
     
     
         19 . 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, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft.