Patent Publication Number: US-2023160338-A1

Title: Geared turbofan engine with targeted modular efficiency

Description:
CROSS-REFERENCE TO RELATED APPLICATION 
     This application is a continuation of U.S. patent application Ser. No. 17/737,179 filed on May 5, 2022, which is a continuation of U.S. patent application Ser. No. 17/038,608 filed on Sep. 30, 2020, now U.S. Pat. No. 11,371,427 granted on Jun. 28, 2022, which is a continuation of U.S. patent application Ser. No. 16/531,704 filed on Aug. 5, 2019, now U.S. Pat. No. 10,823,052 granted on Nov. 3, 2020, which is a continuation of U.S. patent application Ser. No. 14/651,923 filed on Jun. 12, 2015, now U.S. Pat. No. 10,371,047 granted on Aug. 6, 2019, which is a National Phase application of International Application No. PCT/US2014/057127 filed on Sep. 24, 2014, which claims priority to U.S. Provisional Application No. 61/891,475 filed on Oct. 16, 2013. 
    
    
     BACKGROUND 
     A gas turbine engine typically includes a fan section, a compressor section, a combustor section and a turbine section. Air entering the compressor section is compressed and delivered into the combustion section where it is mixed with fuel and ignited to generate a high-speed exhaust gas flow. The high-speed exhaust gas flow expands through the turbine section to drive the compressor and the fan section. The compressor section typically includes low and high pressure compressors, and the turbine section includes low and high pressure turbines. 
     The high pressure turbine drives the high pressure compressor through an outer shaft to form a high spool, and the low pressure turbine drives the low pressure compressor through an inner shaft to form a low spool. The fan section may also be driven by the low inner shaft. A direct drive gas turbine engine includes a fan section driven by the low spool such that the low pressure compressor, low pressure turbine and fan section rotate at a common speed in a common direction. 
     A speed reduction device such as an epicyclical gear assembly may be utilized to drive the fan section such that the fan section may rotate at a speed different than the turbine section so as to increase the overall propulsive efficiency of the engine by allowing an increase in the fan diameter and a reduction in a fan pressure rise. In such engine architectures, a shaft driven by one of the turbine sections provides an input to the epicyclical gear assembly that drives the fan section at a reduced speed such that both the turbine section and the fan section can rotate at closer to their individual optimal speeds. 
     Although geared architectures have improved propulsive efficiency, turbine engine manufacturers continue to seek further improvements to engine performance including improvements to thermal, transfer and propulsive efficiencies. 
     SUMMARY 
     A turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan section including a fan blade having a leading edge and hub to tip ratio of less than about 0.34 and greater than about 0.020 measured at the leading edge and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge radial tip length that is greater than about 67% of the radial tip length of a leading edge of a first stage of the first compression section. A second compression section includes a last blade trailing edge radial tip length that is greater than about 57% of a radial tip length of a leading edge of a first stage of the first compression section. 
     In a further embodiment of the foregoing turbine engine, the fan section provides a low fan pressure ratio less than about 1.6. 
     In a further embodiment of any of the foregoing turbine engines, the fan section provides a low fan pressure ratio between about 1.45 and about 1.20. 
     In a further embodiment of any of the foregoing turbine engines, the fan section provides a bypass ratio greater than about 8. 
     In a further embodiment of any of the foregoing turbine engines, the fan section provides a bypass ratio greater than about 8. 
     In a further embodiment of any of the foregoing turbine engines, the fan section provides a bypass ratio greater than about 12. 
     In a further embodiment of any of the foregoing turbine engines, includes a turbine section that has a fan drive turbine and at least two turbine stages forward of a first turbine blade of the fan drive turbine. 
     In a further embodiment of any of the foregoing turbine engines, the fan drive turbine includes at least three stages. 
     In a further embodiment of any of the foregoing turbine engines, the fan drive turbine is coupled to the drive the first compression section. 
     In a further embodiment of any of the foregoing turbine engines, at least one of the at least two turbine stages is coupled to drive the second compression section. 
     In a further embodiment of any of the foregoing turbine engines, the at least two stages include a single turbine second forward of the fan drive turbine. 
     In a further embodiment of any of the foregoing turbine engines, the first compression section includes three stages and the second compression section includes eight stages. 
     A turbine engine according to an exemplary embodiment of this disclosure, among other possible things includes a fan section providing a bypass ratio greater than about 12, and a speed change mechanism with gear ratio greater than about 2.6 to 1. A first compression section includes a last blade trailing edge tip length that is greater than 67% of the radial length of a first stage leading edge of the first compression section. A second compression section includes a last blade with a trailing edge tip that includes a radial length that is greater than 57% of a radial length of the first stage leading edge of the first compression section. 
     In a further embodiment of any of the foregoing turbine engines, the fan section includes a plurality of fan blades supported on a hub. A leading edge of at least one of the fan blades includes a leading edge and a hub to tip ratio is less than about 0.34 and greater than about 0.020 measured at the leading edge. 
     In a further embodiment of any of the foregoing turbine engines, includes a turbine section that has a fan drive turbine and at least two turbine stages forward of a first turbine blade of the fan drive turbine. 
     In a further embodiment of any of the foregoing turbine engines, at least one of the at least two turbine stages is coupled to drive the second compression section. 
     In a further embodiment of any of the foregoing turbine engines, the at least two stages include a single turbine section forward of the fan drive turbine. 
     In a further embodiment of any of the foregoing turbine engines, the fan section provides a fan pressure ratio between about 1.45 and about 1.20. 
     In a further embodiment of any of the foregoing turbine engines, the first compression section includes three stages and the second compression section includes eight stages. 
     Although the different examples have the specific components shown in the illustrations, embodiments of this disclosure are not limited to those particular combinations. It is possible to use some of the components or features from one of the examples in combination with features or components from another one of the examples. 
     These and other features disclosed herein can be best understood from the following specification and drawings, the following of which is a brief description. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIG.  1    is a schematic view of an example turbine engine according to an embodiment. 
         FIG.  2    is a schematic view of a compressor section of the example turbine engine according to an embodiment. 
         FIG.  3    is a schematic view of another compressor section of the example turbine engine according to an embodiment. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    schematically illustrates an example gas turbine engine  20  that includes a fan section  22 , a compressor section  24 , a combustor section  26  and a turbine section  28 . The fan section  22  drives air along a bypass flow path B while the compressor section  24  draws air in along a core flow path C where air is compressed and communicated to a combustor section  26 . In the combustor section  26 , air is mixed with fuel and ignited to generate a high pressure exhaust gas stream that expands through the turbine section  28  where energy is extracted and utilized to drive the fan section  22  and the compressor section  24 . 
     Although the disclosed non-limiting embodiment depicts one gas turbine engine, it should be understood that the concepts and teachings described herein may be applied to other types of turbine engines; for example a turbine engine including a three-spool architecture in which three spools concentrically rotate about a common axis and where a low spool enables a low pressure turbine to drive a fan via a gearbox, an intermediate spool that enables an intermediate pressure turbine to drive a first compressor of the compressor section, and a high spool that enables a high pressure turbine to drive a high pressure compressor of the compressor section. 
     The example engine  20  generally includes a low speed spool  30  and a high speed spool  32  mounted for rotation about an engine central longitudinal axis A relative to an engine static structure  36  via several bearing systems  38 . It should be understood that various bearing systems  38  at various locations may alternatively or additionally be provided. 
     The low speed spool  30  generally includes an inner shaft  40  that connects the fan section  22  and a low pressure (or first) compressor section  44  to a low pressure (or first) turbine section  46 . The inner shaft  40  drives the fan section  22  through a speed change device, such as a geared architecture  48 , to drive the fan section  22  at a lower speed than the low speed spool  30 . The high-speed spool  32  includes an outer shaft  50  that interconnects a high pressure (or second) compressor section  52  and a high pressure (or second) turbine section  54 . The inner shaft  40  and the outer shaft  50  are concentric and rotate via the bearing systems  38  about the engine central longitudinal axis A. 
     A combustor  56  is arranged between the high pressure compressor  52  and the high pressure turbine  54 . In one example, the high pressure turbine  54  includes at least two stages to provide a double stage high pressure turbine  54 . As used herein, a “high pressure” compressor or turbine experiences a higher pressure than a corresponding “low pressure” compressor or turbine. 
     The example low pressure turbine  46  has a pressure ratio that is greater than about 5. The pressure ratio of the example low pressure turbine  46  is measured prior to an inlet of the low pressure turbine  46  as related to the pressure measured at the outlet of the low pressure turbine  46  prior to an exhaust nozzle. The low pressure turbine  46  is coupled to the fan section  22  through the geared architecture  48  and therefore is also referred to interchangeably in this disclosure as the fan drive turbine  46 . 
     A mid-turbine frame  58  of the engine static structure  36  is arranged generally between the high pressure turbine  54  and the low pressure turbine  46 . The mid-turbine frame  58  further supports bearing systems  38  in the turbine section  28  as well as setting airflow entering the fan drive turbine  46 . 
     Airflow through the core airflow path C is compressed by the low pressure compressor  44  then by the high pressure compressor  52  mixed with fuel and ignited in the combustor  56  to produce high speed exhaust gases that are then expanded through the high pressure turbine  54  and fan drive turbine  46 . The mid-turbine frame  58  includes vanes  60 , which are in the core airflow path and function as an inlet guide vane for the low pressure turbine  46 . Utilizing the vane  60  of the mid-turbine frame  58  as the inlet guide vane for low pressure turbine  46  decreases the length of the low pressure turbine  46  without increasing the axial length of the mid-turbine frame  58 . Choosing a high gearbox input to output ratio, reduces the number of vane rows in the fan drive turbine  46  and shortens the axial length of the turbine section  28 . Thus, the compactness of the gas turbine engine  20  is increased and a higher power density may be achieved. 
     The disclosed gas turbine engine  20  in one example is a high-bypass geared aircraft engine. In a further example, the gas turbine engine  20  includes a bypass ratio greater than about eight (8), with an example embodiment being greater than about twelve (12). The geared architecture  48  is an epicyclical gear train, such as a planetary gear system, star gear system or other known gear system, with a gear reduction ratio of greater than about 2.6. 
     In one disclosed embodiment, the gas turbine engine  20  includes a bypass ratio greater than about twelve (12:1) and a diameter of the fan blades  42  is significantly larger than an outer diameter of the low pressure compressor  44 . It should be understood, however, that the above parameters are only exemplary of one embodiment of a gas turbine engine including a geared architecture and that the present disclosure is applicable to other gas turbine engines. 
     A significant amount of thrust is provided by flow through the bypass flow path B due to the high bypass ratio. The fan section  22  of the engine  20  is designed for a particular flight condition—typically cruise at about 0.8 Mach and about 35,000 feet. The flight condition of 0.8 Mach and 35,000 ft., with the engine at its best fuel consumption—also known as “bucket cruise Thrust Specific Fuel Consumption (‘TSFC’)”—is the industry standard parameter of pound-mass (Ibm) of fuel per hour being burned divided by pound-force (Ibf) of thrust the engine produces at that minimum point. 
     “Low fan pressure ratio” is the pressure ratio across the fan blade alone, without a Fan Exit Guide Vane (“FEGV”) system. The low fan pressure ratio as disclosed herein according to one non-limiting embodiment is less than about 1.50. In another non-limiting embodiment, the low fan pressure ratio is between 1.45 and 1.20. 
     “Low corrected fan tip speed” is the actual fan tip speed in ft/sec divided by an industry standard temperature correction of [(Tram °R)/(518.7°R)] 0.5 . The “Low corrected fan tip speed”, as disclosed herein according to one non-limiting embodiment, is less than about 1150 ft/second. 
     The example gas turbine engine includes the fan section  22  that comprises in one non-limiting embodiment less than about 26 fan blades  42 . In another non-limiting embodiment, the fan section  22  includes less than about 20 fan blades  42 . Moreover, in one disclosed embodiment the low pressure turbine  46  includes no more than about 6 turbine rotor stages schematically indicated at  34 . In another non-limiting example embodiment, the low pressure turbine  46  includes about 3 turbine rotor states. A ratio between the number of fan blades  42  and the number of low pressure turbine rotors is between about 3.3 and about 8.6. The example low pressure turbine  46  provides the driving power to rotate the fan section  22  and therefore the relationship between the number of turbine rotor stages  34  in the low pressure turbine  46  and the number of blades  42  in the fan section  22  disclose an example gas turbine engine  20  with increased power transfer efficiency. 
     An example disclosed engine  20  provides a system-level combination of component (module) efficiencies and a system-level combination of features within these modules that are used to arrive at uniquely high engine efficiency (i.e. Thrust Specific Fuel Consumption) at takeoff and at bucket cruise. The disclosed combination of components provide benefit in a commercial engine with very high bypass ratio in achieving the stated, very low, thrust specific fuel consumption (see table  1 ) and is especially beneficial to a single aisle aircraft where the overall pressure ratio of the compressor is less than 50. 
     
       
         
           
               
               
               
               
             
               
                   
                 TABLE 1 
               
               
                   
                   
               
               
                   
                 Sea level takeoff, 
                 [2]Sea level takeoff, 
                   
               
               
                   
                 86 deg F., 0.0 Mn: 
                 86 deg F., 0.0 Mn: 
               
               
                   
                 Test Stand Operation: 
                 Test Stand Operation: 
               
               
                   
                 no power extraction, no 
                 no power extraction, no 
                 Bucket Cruise, 
               
               
                   
                 Environmental Control 
                 Environmental Control 
                 0.8 Mn, 35,000 ft, 
               
               
                   
                 System bleed 
                 System bleed 
                 Standard Day 
               
               
                   
                   
               
             
            
               
                   
               
            
           
           
               
               
               
               
            
               
                 Thrust Specific Fuel 
                   
                 0.2751 
                 0.53717 
               
               
                 Consumption [1] 
               
               
                 Speed change 
                 At least 
                 2.6 
                 2.6 
               
               
                 (Input/output) 
               
               
                 Component efficiency 
               
               
                 Fan OD 
                 at least 0.90 
                 0.9344 
                 0.96501 
               
               
                 Speed Change Mechanism 
                 at least 0.985 
                 0.9949 
                 0.99374 
               
               
                 First Compressor or LPC 
                 at least 0.84 
                 0.8695 
                 0.86622 
               
               
                 Second Compressor or HPC 
                 at least 0.82 
                 0.8495 
                 0.8356 
               
               
                 Turbine Section(s) 
                 at least 0.85 for the single 
                 0.87544 
                 0.8938 
               
               
                 excluding the 
                 HPT or combined 
               
               
                 fan drive turbine 
                 efficiency if two or more 
               
               
                   
                 turbines are used 
               
               
                 Fan Drive turbine 
                 at least 0.89 
                 0.9251 
                 0.9266 
               
               
                   
               
            
           
         
       
     
     The combination of module efficiency includes among other possible things, the fan section  22  with the fan blades  42  supported on a fan hub  64 . Each of the fan blades  42  includes a leading edge  62  that extends a radial distance  66  from the engine axis A. The fan hub  64  extends a radial distance  68  from the engine axis A. A low hub-tip ratio of fan hub radial radius  68  to the radius at the leading edge  62  of the fan blade  42  is less than 0.34 and greater than 0.020. The disclosed range of ratios is desirable in that the lower this value is, the smaller the outer fan section and inlet section has to be to accommodate a given amount of air, and maintaining this dimension within the desired ratio range enables a reduction in engine weight relative to an engine with a higher hub to tip ratio. In one example embodiment, the fan section  22  further provides a low fan pressure ratio that is between about 1.45 and about 1.20, and a bypass ratio greater than about 8.0. 
     The disclosed engine  20  includes the geared architecture  48  with a gear ratio greater than about 2.6 to 1. In this example, the speed change system is the geared architecture, which is an epicyclical gearbox and which includes planet gears or star gears interspersed by baffles for gathering and directing lubricant during operation. 
     The example turbine section  28  has at least two turbine stages forward of the first turbine stage  94  included in the fan drive turbine  46 . In this example, the high pressure or second turbine includes two turbine stages  96  forward of the fan drive turbine  46 . In this example, the two turbine stages  96  are part of a single high pressure turbine  54  with at least two turbine rotors  96 , however, it is within the contemplation of this disclosure that the at least two turbine rotors forward of the fan drive turbine  46  could be part of multiple turbines that rotate independent of each other, for example, two separate turbine sections with at least one turbine rotor each. 
     Referring to  FIG.  2   , with continued reference to  FIG.  1   , the first compression section  44 , which in one disclosed example is a low pressure compressor (LPC)  44 , includes three stages. The example LPC  44  includes first compressor blade  70  with a leading edge  72  and a last compressor blade  74  with trailing edge  76 . A tip of the leading edge  72  of the first blade  70  extends a radial distance  78  from the engine axis A. A tip of the trailing edge  76  of the last blade  74  extends a radial distance  80  from the engine axis A. The first compression section  44  is configured such that a ratio between the radial distance  80  at the trailing edge  76  is greater than 67% of the radial distance  78  of the leading edge  72  of the first blade  70 . The example configuration provided by the disclosed ratio enables improved airflow through the first compressor section  44  that provides improved efficiency. The disclosed relationship between the leading edge  72  and the trailing edge  76  enables a beneficial modest slope to the engine casing structures spanning the compressor section  24 . The modest slope provides for minimal effects to tip clearances of the compressor blades due to axial shifting of the compressor rotor due to overall aerodynamic loading. 
     Referring to  FIG.  3    with continued reference to  FIG.  1   , the second compression section  52 , which in one disclosed example is a high pressure compressor (HPC), includes at least eight stages. The example HPC  52  includes a first blade  82  with a leading edge  84  that extends a radial distance  86  from the engine axis A to a tip. The second compressor section  52  also includes a last blade  88  having a trailing edge  90  that extends a radial distance  92  from the engine axis A to the tip. A ratio between the leading edge  84  and the trailing edge  90  defines the configuration of the compressor  52  that provides the improved efficiency. In one disclosed example, the radial distance  92  of the trailing edge  90  of the last blade  88  is greater than about 57% of the radial distance  86  of the leading edge  84  of the first blade  82  of the second compressor section  52 . 
     A geared turbine arrangement for short range aircraft can uniquely exploit the particular aspects of an aircraft duty cycle that is characterized by an unusually low proportion of time in cruise operation versus the total time spent at takeoff and climb power (for a representative time span such as between engine overhauls). 
     A definition of a short range aircraft is one with a total flight length less than about 300 nautical miles. 
     
       
         
           
               
               
               
             
               
                 TABLE 2 
               
               
                   
               
               
                 Engine 
                 #1 
                 #2 
               
               
                   
               
             
            
               
                   
               
            
           
           
               
               
               
               
               
            
               
                 Max takeoff weight 
                 53,060 kg 
                 (117,000 lb) 
                 58,967 kg 
                 (130,000 lb) 
               
               
                 Max landing weight 
                 49,895 kg 
                 (110,000 lb) 
                 55,111 kg 
                 (121,500 lb) 
               
               
                 Maximum cargo 
                 3,629 kg 
                 (8,000 lb) 
                 4,853 kg 
                 (10,700 lb) 
               
               
                 payload 
               
               
                 Maximum payload 
                 13,676 kg 
                 (30,150 lb) 
                 16,284 kg 
                 (35,900 lb) 
               
               
                 (total) 
               
               
                 Max range 
                 2,778 km 
                 (1,500 nmi) 
                 2,778 km 
                 (1,500 nmi) 
               
               
                 Take off run 
                 1,219 m 
                 (3,999 ft) 
                 1,524 m 
                 (5,000 ft) 
               
               
                 at MTOW 
               
               
                 Landing field 
                 1,341 m 
                 (4,400 ft) 
                 1,448 m 
                 (4,751 ft) 
               
               
                 length at MLW 
               
               
                   
               
            
           
         
       
     
     As is shown in Table 2, a short range aircraft for purposes of this disclosure is defined as including a single aisle configuration with 2, 3 seating or 3, 3 seating. Conventionally, a short range aircraft has a capacity of about 200 passengers or less. Moreover, an example short range aircraft will have a maximum range of only about 1500 nautical miles. 
     Because of the extremely high utilization in terms of cumulative hours at relatively high power during take-off conditions, the disclosed geared turbine engine  20  arrangement is configured differently to achieve a beneficial balance of fuel burn and maintenance costs. The high power utilization is a result of frequent operation at high power conditions that generate high turbine temperatures, elevated turbine cooling air temperatures and elevated temperatures at the rear stage of the compressor. The result of such operation is that LPC pressures rise, temperature rise and efficiency may be lower than for a long range aircraft. In a long range aircraft that operates for longer periods and a greater portion of the cumulative operating hours, maximizing LPC efficiency is desired provides a significant benefit, and is a key difference when compared to short range aircraft. Pressure and temperature rise can be increased due to the less frequent use of takeoff power between overhaul periods which could be around 4000 hours for both the short range and long range commercial airliner. 
     Although an example embodiment has been disclosed, a worker of ordinary skill in this art would recognize that certain modifications would come within the scope of this disclosure. For that reason, the following claims should be studied to determine the scope and content of this disclosure.