Efficient jet

Abstract

A gas turbine engine includes: an engine core, compressor system, and core shaft. A compressor exit pressure is defined as an average airflow pressure at the exit of the highest pressure compressor at cruise conditions. The core has an annular splitter and bypass flow. Stagnation streamlines around the engine circumference form a streamsurface. A fan is upstream the core with blades having leading and trailing edges, and a radially inner portion within the streamtube. A fan root entry pressure is an average airflow pressure across the radially inner portion leading edge of each fan blade at cruise conditions. An overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure. A bypass jet velocity is defined as the jet velocity of air flow exiting the bypass exhaust nozzle at cruise conditions. A jet velocity is in a range between 4.7 m/s and 7.7 m/s.

1 - 17 . (canceled) 
     
     
         18 . A gas turbine engine for an aircraft comprising:
 an engine core comprising a turbine system comprising one or more turbines, a compressor system comprising one or more compressors, and a core shaft connecting the turbine system to the compressor system, wherein a compressor exit temperature is defined as a temperature of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions and a compressor exit pressure is defined as a pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core further comprising an annular splitter at which 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 a circumference of the engine, stagnating on a leading edge of the annular splitter, form a streamsurface forming a radially outer boundary of a streamtube that contains all of the core flow;   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 inner portion lying within the streamtube that contains the core flow, and wherein a fan root entry temperature is defined as an temperature of airflow across the leading edge of the radially inner portion of each fan blade at cruise conditions and wherein a fan root entry pressure is defined as an pressure of airflow across the leading edge of the radially inner portion of each fan blade at cruise conditions; and   a nacelle surrounding the engine core, the nacelle defining the bypass duct and a bypass exhaust nozzle, wherein:   an overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure,   a bypass nozzle pressure ratio is defined as the nozzle pressure ratio of the bypass exhaust nozzle at cruise conditions,   a core temperature rise is defined as the compressor exit temperature in Kelvin divided by the fan root entry temperature in Kelvin,   a temperature-pressure ratio defined as:   
       
         
           
             
               
                 the 
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                  
                 core 
                  
                 
                     
                 
                  
                 temperature 
                  
                 
                     
                 
                  
                 rise 
               
               
                 the 
                  
                 
                     
                 
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                 bypass 
                  
                 
                     
                 
                  
                 nozzle 
                  
                 
                     
                 
                  
                 pressure 
                  
                 
                     
                 
                  
                 ratio 
               
             
           
         
         is in a range between 1.52 and 1.8, and 
         the overall pressure ratio is in a range between 42.5 and 70, and 
         wherein an annular fan face is defined at a leading edge of the fan, and 
         a quasi-non-dimensional mass flow rate Q is defined as: 
       
       
         
           
             
               Q 
               = 
               
                 W 
                  
                 
                   
                     
                       T 
                       0 
                     
                   
                   
                     
                       P 
                       0 
                     
                     · 
                     
                       A 
                       
                         f 
                          
                         a 
                          
                         n 
                       
                     
                   
                 
               
             
           
         
         where: 
         W is mass flow rate through the fan in Kg/s; 
         T 0  is average stagnation temperature of the air at the fan face in Kelvin; 
         P 0  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 ; 
         wherein at cruise conditions, Q has a value in the range between 0.031 Kgs −1 N −1 K 1/2  and 0.036 Kgs −1 N −1 K 1/2 . 
       
     
     
         19 . A gas turbine engine for an aircraft comprising:
 an engine core comprising a turbine system comprising one or more turbines, a compressor system comprising one or more compressors, and a core shaft connecting the turbine system to the compressor system, wherein a compressor exit temperature is defined as a temperature of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions and a compressor exit pressure is defined as a pressure of airflow at the exit of the highest pressure compressor of the compressor system at cruise conditions, the engine core having a core radius defined between the centreline of the engine and a forwardmost tip of the engine core;   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, wherein a radially inner portion of each fan blade comprises the portion of each fan blade at a distance from the centreline of the engine less than the core radius, and wherein a fan root entry temperature is defined as a temperature of airflow across the leading edge of the radially inner portion of each fan blade at cruise conditions and wherein a fan root entry pressure is defined as a pressure of airflow across the leading edge of the radially inner portion of each fan blade at cruise conditions; and   a nacelle surrounding the engine core, the nacelle defining the bypass duct and a bypass exhaust nozzle, wherein:   an overall pressure ratio is defined as the compressor exit pressure divided by the fan root entry pressure,   a bypass nozzle pressure ratio is defined as the nozzle pressure ratio of the bypass exhaust nozzle at cruise conditions,   a core temperature rise is defined as the compressor exit temperature in Kelvin divided by the fan root entry temperature in Kelvin,   a temperature-pressure ratio defined as:   
       
         
           
             
               
                 the 
                  
                 
                     
                 
                  
                 core 
                  
                 
                     
                 
                  
                 temperature 
                  
                 
                     
                 
                  
                 rise 
               
               
                 the 
                  
                 
                     
                 
                  
                 bypass 
                  
                 
                     
                 
                  
                 nozzle 
                  
                 
                     
                 
                  
                 pressure 
                  
                 
                     
                 
                  
                 ratio 
               
             
           
         
         is in a range between 1.52 and 1.8, and 
         the overall pressure ratio is in a range between 42.5 and 70, and 
         wherein an annular fan face is defined at a leading edge of the fan, and 
         a quasi-non-dimensional mass flow rate Q is defined as: 
       
       
         
           
             
               Q 
               = 
               
                 W 
                  
                 
                   
                     
                       T 
                       0 
                     
                   
                   
                     
                       P 
                       0 
                     
                     · 
                     
                       A 
                       
                         f 
                          
                         a 
                          
                         n 
                       
                     
                   
                 
               
             
           
         
         where: 
         W is mass flow rate through the fan in Kg/s; 
         T 0  is average stagnation temperature of the air at the fan face in Kelvin; 
         P 0  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 ; 
         wherein at cruise conditions, Q has a value in the range between 0.031 Kgs −1 N −1 K 1/2  and 0.036 Kgs −1 N −1 K 1/2 . 
       
     
     
         20 . A gas turbine engine according to  claim 18 , wherein the overall pressure ratio is in a range between 50 and 70. 
     
     
         21 . A gas turbine engine according to  claim 18 , wherein the bypass nozzle pressure ratio is in a range between 2.02 and 2.25. 
     
     
         22 . A gas turbine engine according to  claim 18 , wherein the core temperature rise is in a range between 3.10 and 3.50. 
     
     
         23 . A gas turbine engine according to  claim 18 , wherein:
 the compressor system comprises a first compressor and a second compressor,   the turbine system comprises a first turbine and a second turbine,   the core shaft is a first core shaft connecting the first compressor and the first turbine,   the engine core further comprises a second core shaft connecting the second turbine to the second compressor,   the second turbine, second compressor, and second core shaft are arranged to rotate at a higher rotational speed than the first core shaft, wherein the first compressor comprises 3 compression stages and the second compressor comprises 10 compression stages or greater,   and optionally wherein the first compressor comprises 3 compression stages and the second compressor comprises 9 compression stages or greater.   
     
     
         24 . A gas turbine engine according to  claim 18 , wherein a stage pressure rise is generated across each compression stage of the compressor system, and the stage pressure rise of compressor stages provided in the compressor system is in the range between 1.3 and 1.4. 
     
     
         25 . A gas turbine engine according to  claim 18 , wherein, at cruise conditions, the specific thrust of the gas turbine engine is any one of:
 a) in a range between 50 Nkg −1 s and 100 Nkg −1 s;   b) in a range between 70 Nkg −1 s and 90 Nkg −1 s; or   c) equal to or below 90 Nkg −1 s   
     
     
         26 . The gas turbine engine according to  claim 18 , wherein Q takes a value less than or equal to 0.035 Kgs −1 N −1 K 1/2  at cruise conditions. 
     
     
         27 . The gas turbine engine according to  claim 18 , wherein a fan tip loading at cruise conditions 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 of the fan, and wherein the fan tip loading is in a range from 0.25 to 0.4. 
     
     
         28 . The gas turbine engine according to  claim 27 , wherein the fan tip loading takes a value in the range from 0.29 to 0.31 at cruise conditions. 
     
     
         29 . The gas turbine engine according to  claim 18 , wherein a fan tip pressure ratio is defined as the ratio of the mean total pressure of the air flow at the exit of the fan that subsequently flows through the bypass duct to the mean total pressure of the air flow at the inlet of the fan, and wherein, at cruise conditions:
 the fan tip pressure ratio is in a range between 1.35 and 1.44.   
     
     
         30 . The gas turbine engine according to  claim 18 , wherein a fan root pressure ratio is defined as the ratio of the mean total pressure of the air flow at the exit of the fan that subsequently flows through the engine core to the mean total pressure of the air flow at the inlet of the fan, and wherein, at cruise conditions:
 the fan root pressure ratio is in a range between 1.18 and 1.30.   
     
     
         31 . The gas turbine engine according to  claim 18 , wherein a fan pressure ratio is defined as the ratio of the mean total pressure of the air flow at the exit of the fan to the mean total pressure of the air flow at the inlet of the fan, and wherein, at cruise conditions:
 the fan pressure ratio is in a range between 1.35 and 1.43.   
     
     
         32 . The gas turbine engine according to  claim 18 , wherein cruise conditions means the conditions at mid-cruise of an aircraft to which the engine is attached, and optionally means the conditions experienced by the aircraft and engine at the midpoint between top of climb and start of decent. 
     
     
         33 . The gas turbine engine according to  claim 18 , wherein any one or more of:
 a) 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;   b) the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 11582 m and a forward Mach Number of 0.8;   c) the cruise conditions correspond to atmospheric conditions defined by the International Standard Atmosphere at an altitude of 10668 m and a forward Mach Number of 0.85; and/or   d) the cruise conditions correspond to atmospheric conditions at an altitude that is in a range of from 10500 m to 11600 m.   
     
     
         34 . The gas turbine engine according to  claim 18 , further comprising 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, and optionally wherein the gear box has a gear ratio in a range from 3.2 to 3.7.