Patent Publication Number: US-2023151772-A1

Title: Reverse-flow gas turbine engine

Description:
This application is a continuation of U.S. patent application Ser. No. 17/368,317 filed Jul. 6, 2021, which is a continuation of U.S. patent application Ser. No. 15/266,321 filed Sep. 15, 2016, which are hereby incorporated herein by reference in their entireties. 
    
    
     TECHNICAL FIELD 
     The application relates generally to gas turbine engines and, more particularly, to gas turbine engines with a reverse flow core. 
     BACKGROUND OF THE ART 
     Reverse-flow gas turbine engines draw air into a central core of the engine near a rear portion of the engine, and exhaust combustion gases from a front portion of the engine. Gases therefore flow through the core from the rear to the front of the engine. 
     In some conventional reverse-flow engines, air is drawn into the core and compressed with a single compressor stage driven by a first turbine stage. A second turbine stage, separate from the first turbine stage and rotating a separate shaft, provides the rotational output of the engine. The first turbine stage is therefore performing all of the work to compress the air, which may affect the overall efficiency of the engine. 
     SUMMARY 
     In one aspect, there is provided a gas turbine engine, comprising: a first spool having a low pressure compressor section disposed forward of an air inlet along a direction of travel of the engine and in fluid communication with the air inlet, and a low pressure turbine section disposed forward of the low pressure compressor section and drivingly engaged thereto; a second spool having a high pressure compressor section disposed forward of the low pressure compressor section and in fluid communication therewith to receive pressurized air therefrom, and a high pressure turbine section disposed forward of the high pressure compressor section and drivingly engaged thereto, the high pressure turbine section disposed aft of the low pressure turbine section and in fluid communication therewith; and an output drive shaft drivingly engaged to the low pressure turbine section and extending forwardly therefrom, the drive shaft configurable to drivingly engage a rotatable load disposed forward of the low pressure turbine section. 
     In another aspect, there is provided a method of operating a gas turbine engine, comprising: drawing air into a core of the engine through a low pressure compressor section and then through a high pressure compressor section along a forward direction with respect to a direction of travel of the engine to provide pressurized air; igniting a mixture of the pressurized air and fuel to generate combustion gases; circulating the combustion gases along the forward direction through a high pressure turbine section and then through a low pressure turbine section to drive the high pressure and low pressure turbine sections with the combustion gases; driving the high pressure compressor section with the high pressure turbine section; and driving the low pressure compressor section and a rotatable load with the low pressure turbine section, the low pressure turbine section and the high pressure turbine section rotating independently from one another. 
     In yet another aspect, there is provided a gas turbine engine, comprising: an output drive shaft having a front end configurable to drivingly engage a rotatable load; a low pressure turbine section and a low pressure compressor section drivingly engaged to the drive shaft, the low pressure turbine section disposed forward of the low pressure compressor section; a high pressure shaft rotatable independently of the drive shaft; a high pressure turbine section and a high pressure compressor section drivingly engaged to the high pressure shaft, the high pressure compressor section disposed forward of the low pressure compressor section and in fluid communication therewith, and the high pressure turbine section disposed aft of the low pressure turbine section and in fluid communication therewith; wherein during operation of the engine, air flows toward the front end of the drive shaft through the low pressure and high pressure compressor sections, and combustion gases flow toward the front end of the drive shaft through the high pressure and low pressure turbine sections. 
    
    
     
       DESCRIPTION OF THE DRAWINGS 
       Reference is now made to the accompanying figures in which: 
         FIG.  1    is a schematic cross-sectional view of a gas turbine engine, according to an embodiment of the present disclosure; and 
         FIG.  2    is a schematic cross-sectional view of a gas turbine engine, according to another embodiment of the present disclosure. 
     
    
    
     DETAILED DESCRIPTION 
       FIG.  1    illustrates a gas turbine engine  10  of a type preferably provided for use in subsonic flight, generally comprising in serial flow communication an air inlet  11 , a compressor section  12  for pressurizing the air from the air inlet  11 , a combustor  13  in which the compressed air is mixed with fuel and ignited for generating an annular stream of hot combustion gases, a turbine section  14  for extracting energy from the combustion gases, an exhaust outlet  15  through which the combustion gases exit the gas turbine engine  10 . The engine  10  includes a propeller  16  which provides thrust for flight and taxiing. The gas turbine engine  10  has a longitudinal center axis  17 . 
     The gas turbine engine  10  (sometimes referred to herein simply as “engine  10 ”) has a central core  18  through which gases flow and which includes some of the turbomachinery of the engine  10 . The engine  10  is a “reverse-flow” engine  10  because gases flow through the core  18  from the air inlet  11  at a rear portion thereof, to the exhaust outlet  15  at a front portion thereof. This is in contrast to “through-flow” gas turbine engines in which gases flow through the core of the engine from a front portion to a rear portion. The direction of the flow of gases through the core  18  of the engine  10  disclosed herein can be better appreciated by considering that the gases flow through the core  18  in the same direction D as the one along which the engine  10  travels during flight. Stated differently, gases flow through the engine  10  from a rear end thereof towards the propeller  16 . 
     It will thus be appreciated that the expressions “forward” and “aft” used herein refer to the relative disposition of components of the engine  10 , in correspondence to the “forward” and “aft” directions of the engine  10  and aircraft including the engine  10  as defined with respect to the direction of travel. In the embodiment shown, a component of the engine  10  that is “forward” of another component is arranged within the engine  10  such that it is located closer to the propeller  16 . Similarly, a component of the engine  10  that is “aft” of another component is arranged within the engine  10  such that it is further away from the propeller  16 . 
     Still referring to  FIG.  1   , the engine  10  has multiple spools which perform compression to pressurize the air received through the air inlet  11 , and which extract energy from the combustion gases before they exit the core  18  via the exhaust outlet  15 . 
     A first spool  20  includes at least one component to compress the air that is part of the compressor section  12 , and at least one component to extract energy from the combustion gases that is part of the turbine section  14 . More particularly, the first spool  20  has a low pressure turbine section  21  which extracts energy from the combustion gases, and which is drivingly engaged (e.g. directly connected) to a low pressure compressor section  22  for pressurizing the air. The low pressure turbine section  21  (sometimes referred to herein simply as “LP turbine section  21 ”) drives the low pressure compressor section  22  (sometimes referred to herein simply as “LPC section  22 ”) thereby causing the LPC section  22  to pressurize the air. Both the LP turbine section  21  and the LPC section  22  are disposed along the center axis  17 . In the depicted embodiment, both the LP turbine section  21  and the LPC section  22  are axial rotatable components having an axis of rotation that is coaxial with the center axis  17 . They can each include one or more stages of rotors and stators, depending upon the desired engine thermodynamic cycle, for example. 
     In the depicted embodiment, the first spool  20  has a power shaft  23  which mechanically couples the LP turbine section  21  and the LPC section  22 , and extends between them. The power shaft  23  is coaxial with the center axis  17  of the engine  10 . The power shaft  23  allows the LP turbine section  21  to drive the LPC section  22  during operation of the engine  10 . The power shaft  23  is not limited to the configuration depicted in  FIG.  1   , and can also mechanically couple the LP turbine section  21  and the LPC section  22  in any other suitable way provided that it transmits a rotational drive from the LP turbine section  21  to the LPC section  22 . For example, the power shaft  23  can be combined with a geared LPC section  22  to allow the LPC section  22  to run at a different rotational speed from the LP turbine section  21 . This can provide more flexibility in the selection of design points for the LPC section  22 . 
     The LP turbine section  21  is forward of the LPC section  22 . The LP turbine section  21  is also aft of the exhaust outlet  15 . The LPC section  22  is forward of the air inlet  11 . This arrangement of the LP turbine section  21  and the LPC section  22  provides for a reverse-flow engine  10  that has one or more low pressure compressors located at the rear of the engine  10  which are driven by one or more low pressure turbines located at the front of the engine  10 . 
     Still referring to  FIG.  1   , the engine  10  includes an output drive shaft  24 . The drive shaft  24  extends forwardly from the LP turbine section  21  and is drivingly engaged thereto. The drive shaft  24  is distinct from the power shaft  23  and mechanically coupled thereto to be driven by the LP turbine section  21 . In the depicted embodiment, the drive shaft  24  and the power shaft  23  are coaxial and interconnected.  FIG.  1    shows that the power and drive shafts  23 , 24  are interconnected with a spline  25 . The spline  25 , which can include ridges or teeth on the drive shaft  24  that mesh with grooves in the power shaft  23  (or vice versa), allows for the transfer of torque between the drive shaft  24  and the power shaft  23 . In the depicted embodiment, the power shaft  23  lies at least partially within the drive shaft  24 , such that the spline  25  transfers the rotational drive or torque generated by the LP turbine section  21  from the drive shaft  24  to the power shaft  23 . The spline  25  can operate so that the power shaft  23  and the drive shaft  24  rotate at the same rotational speed. Other mechanical techniques can also be used to interconnect the power and drive shafts  23 , 24 . For example, the power and drive shafts  23 , 24  can be interconnected by curvic coupling, pins, and interference fits. Other configurations of the drive shaft  24  and the power shaft  23  are also possible. For example, the drive shaft  24  and the power shaft  23  can be a single shaft driven by the LP turbine section  21 . The drive shaft  24  therefore transfers the rotational output of the LP turbine section  21  in a forward direction to drive another component. 
     A rotatable load, which in the embodiment shown includes the propeller  16 , is mountable to the engine  10 , and when mounted, is drivingly engaged (e.g. directly connected) to the LP turbine section  21 , and is located forward of the LP turbine section  21 . In such a configuration, during operation of the engine  10 , the LP turbine section  21  drives the rotatable load such that a rotational drive produced by the LP turbine section  21  is transferred to the rotatable load. The rotatable load can therefore be any suitable component, or any combination of suitable components, that is capable of receiving the rotational drive from the LP turbine section  21 , as now described. 
     In the embodiment shown, a reduction gearbox  31  (sometimes referred to herein simply as “RGB  31 ”) is mechanically coupled to a front end of the drive shaft  24 , which extends between the RGB  31  and the LP turbine section  21 . The RGB  31  processes and outputs the rotational drive transferred thereto from the LP turbine section  21  via the drive shaft  24  through known gear reduction techniques. The RGB  31  allows for the propeller  16  to be driven at its optimal rotational speed, which is different from the rotational speed of the LP turbine section  21 . 
     The propeller  16  is mechanically coupled to the output of the RGB  31  via a propeller shaft  35 . The propeller shaft  35  allows the rotational drive outputted by the RGB  31  during operation of the engine  10  to be transferred to the propeller  16  to provide propulsion during flight. In an alternate embodiment where the engine  10  is a turboshaft, the propeller  16  is omitted and the rotational load (which may include, but is not limited to, helicopter main rotor(s) and/or tail rotor(s), propeller(s) for a tilt-rotor aircraft, pump(s), generator(s), gas compressor(s), marine propeller(s), etc.) is driven by the LP turbine section  21  via the RGB  31 , or the propeller  16  and RGB  31  are omitted such that the output of the engine  10  is provided by the output drive shaft  24 . 
     The drive shaft  24  extending forward of the LP turbine section  21  and the power shaft  23  extending aft of the LP turbine section  21  provide the engine  10  with bidirectional drive. Modularity criteria for gas turbine engines may require the use of distinct shafts  23 , 24  that are directly or indirectly connected together. Alternately, the power shaft  23  and the drive shaft  24  can be integral with one another, with a first segment of the integral output shaft extending between the LPC section  22  and the LP turbine section  21 , and a second segment extending between the rotatable load and the LP turbine section  21 . Whether the power shaft  23  is integral with the drive shaft  24  or distinct therefrom, the LP turbine section  21  provides rotational drive outputted at each end of the power shaft  23 . 
     In light of the preceding, it can be appreciated that the LP turbine section  21  drives both the rotatable load and the LPC section  22 . Furthermore, the rotatable load, when mounted to the engine  10 , and the LPC section  22  are disposed on opposite ends of the LP turbine section  21 . It can thus be appreciated that one or more low pressure turbines are used to drive elements in front of the low pressure turbines (e.g. propeller  16 , RGB  31 , etc.) as well as to drive elements to the rear of the low pressure turbines (e.g. LPC section  22 ). This configuration of the LP turbine section  21  allows it to simultaneously drive the rotatable load and the LPC section  22 , if desired. As will be discussed in greater detail below, this arrangement of the rotatable load, the LP turbine section  21 , and the LPC section  22  can contribute to improving the thermodynamic efficiency of the engine  10 . 
     Still referring to  FIG.  1   , the engine  10  includes a second spool  40  with at least one component to compress the air that is part of the compressor section  12 , and at least one component to extract energy from the combustion gases that is part of the turbine section  14 . The second spool  40  is also disposed along the center axis  17  and includes a high pressure turbine section  41  drivingly engaged (e.g. directly connected) to a high pressure compressor section  42  by a high pressure shaft  43  rotating independently of the power shaft  23 . Similarly to the LP turbine section  21  and the LPC section  22 , the high pressure turbine section  41  (sometimes referred to herein simply as “HP turbine section  41 ”) and the high pressure compressor section  42  (sometimes referred to herein simply as “HPC section  42 ”) can include axial rotary components. They can also each include one or more stages of rotors and stators, depending upon the desired engine thermodynamic cycle, for example. In the depicted embodiment, the HPC section  42  includes a centrifugal compressor  42 A or impeller and an axial compressor  42 B, both of which are driven by the HP turbine section  41 . During operation of the engine  10 , the HP turbine section  41  drives the HPC section  42 . 
     The HP turbine section  41  is aft of the LP turbine section  21 , and forward of the combustor  13 . The HPC section  42  is aft of the combustor  13 , and forward of the LPC section  22 . From this arrangement of the HP turbine section  41  and the HPC section  42 , it can be appreciated that during operation of the engine  10 , the LPC section  22  driven by the LP turbine section  21  feeds pressurized air to the HPC section  42 . Therefore, the pressurized air flow produced by the LPC section  22  is provided to the HPC section  42  and contributes to the work of both the LP turbine section  21  and the HP turbine section  41 . 
     It can thus be appreciated that the presence of the above-described first and second spools  20 , 40  provides the engine  10  with a “split compressor” arrangement. More particularly, some of the work required to compress the incoming air is transferred from the HPC section  42  to the LPC section  22 . In other words, some of the compression work is transferred from the HP turbine section  41  to the more efficient LP turbine section  21 . This transfer of work may contribute to higher pressure ratios while maintaining a relatively small number of rotors. In a particular embodiment, higher pressure ratios allow for higher power density, better engine specific fuel consumption (SFC), and a lower turbine inlet temperature (sometimes referred to as “T4”) for a given power. These factors can contribute to a lower overall weight for the engine  10 . The transfer of compression work from the HPC section  42  to the LPC section  22  contrasts with some conventional reverse-flow engines, in which the high pressure compressor (and thus the high pressure turbine) perform all of the compression work. 
     In light of the preceding, it can be appreciated that the LP turbine section  21  is the “low-speed” and “low pressure” turbine section when compared to the HP turbine section  41 , which is sometimes referred to as the “gas generator”. The LP turbine section  21  is sometimes referred to as a “power turbine” section. The turbine rotors of the HP turbine section  41  spin at a higher rotational speed than the turbine rotors of the LP turbine section  21  given the closer proximity of the HP turbine section  41  to the outlet of the combustor  13 . Consequently, the compressor rotors of the HPC section  42  may rotate at a higher rotational speed than the compressor rotors of the LPC section  22 . The engine  10  shown in  FIG.  1    is thus a “two-spool” engine  10 . 
     The HP turbine section  41  and the HPC section  42  can have any suitable mechanical arrangement to achieve the above-described split compressor functionality. For example, and as shown in  FIG.  1   , the second spool  40  includes a high pressure shaft  43  extending between the HPC section  42  and the HP turbine section  41 . The high pressure shaft  43  is coaxial with the power shaft  23  and rotatable relative thereto. The relative rotation between the high pressure shaft  43  and the power shaft  23  allow the shafts  23 , 43  to rotate at different rotational speeds, thereby allowing the HPC section  42  and the LPC section  22  to rotate at different rotational speeds. The high pressure shaft  43  can be mechanically supported by the power shaft  23  using bearings or the like. In the depicted embodiment, the power shaft  23  is at least partially disposed within the high pressure shaft  43 . 
     The split compressor arrangement also allows bleed air to be drawn from between the HPC section  42  and the LPC section  22 . More particularly, in the embodiment of  FIG.  1   , the engine  10  includes an inter-stage bleed  44  port or valve that is aft of the HPC section  42  and forward of the LPC section  22 , which may provide for increased flexibility in the available bleed pressures. In a particular embodiment, the bleed pressure design point of the inter-stage bleed  44  is selected based on the pressure ratio of the LPC section  22 , which runs independently from the HPC section  42 . For operability, variable inlet guide vanes (VIGV) and variable guide vanes (VGV) can be used on the LPC section  22  and at the entry of the HPC section  42 , together with the inter-stage bleed  44 . 
     Still referring to the embodiment shown in  FIG.  1   , the engine  10  also includes an accessory gearbox  50 . The accessory gearbox  50  (sometimes referred to herein simply as “AGB  50 ”) receives a rotational output and in turn drives accessories (e.g. fuel pump, starter-generator, oil pump, scavenge pump, etc.) that contribute to the functionality of the engine  10 . The AGB  50  can be designed with side-facing accessories, top-facing accessories, or rear-facing accessories depending on the installation needs. The AGB  50  is aft of the air inlet  11 . The engine  10  also has a tower shaft  51  that is mechanically coupled to a rear of the high pressure shaft  43  and driven thereby. An accessory gear box drive shaft  52  has a first geared end  52 A mechanically coupled to the tower shaft  51 , and a second geared end  52 B mechanically coupled to the AGB  50 . During operation of the engine  10 , the high pressure shaft  43  transmits a rotational drive to the tower shaft  51  which in turn drives the accessory gearbox drive shaft  52  to thereby drive the accessories of the AGB  50 . In the depicted embodiment, the accessory gearbox drive shaft  52  extends across the air inlet  11 . This configuration of the accessory gearbox drive shaft  52  can take different forms. For example, it can be located outside the air inlet  11 , or may be placed within the air inlet  11  along a strut of the air inlet  11 . It can thus be appreciated that the second end  52 B of the accessory gearbox drive shaft  52  meshes with an input gear of the AGB  50  to drive the AGB  50  across the air inlet  11 . 
     The AGB  50  can be arranged relative to the core  18  of the engine  10  differently than as described above. Referring to  FIG.  2   , the embodiment of the engine  110  has an architecture and arrangement of turbomachinery similar to the engine  10  of  FIG.  1   , where similar elements are identified by the same reference numerals and will not be described further herein.  FIG.  2    shows the AGB  150  mounted on the side of the engine  110 , and forward of the air inlet  11 . The circumferential angular position of the AGB  150  can be selected to suit specific installation needs. Other positions and arrangements for the AGB  150  are possible. 
     Referring to  FIG.  1   , there is also disclosed a method of operating the gas turbine engine  10 . Air is drawn into the core  18  of the engine  10  via the air inlet  11  with the LPC section  22 . The air from the LPC section  22  is further pressurized by the HPC section  42  disposed forward of the LPC section  22 . The air is mixed with fuel and ignited in the combustor  13  to generate combustion gases. The HPC section  42  is driven by the HP turbine section  41  disposed forward of the combustor  13 , and the HP turbine section  41  extracts energy from the combustion gases. The LPC section  22  is driven by the LP turbine section  21  disposed forward of the HP turbine section  41 . The output drive shaft  24  and rotatable load are driven by the LP turbine section  21 , and are located forward of the LP turbine section  21 . 
     It can thus be appreciated that at least some of the embodiments of the engine  10 , 110  disclosed herein provide a mechanical architecture of turbomachinery that allows for a split compressor system. Such a split compressor reverse-flow engine  10 , 110  may be used for aircraft nose installations, as well as for wing installations. 
     The above description is meant to be exemplary only, and one skilled in the art will recognize that changes may be made to the embodiments described without departing from the scope of the invention disclosed. For example, although the engine  10 , 110  is described above as being a turboprop or a turboshaft, it will be appreciated that the engine  10 , 110  can have suitable (through-flow from front to rear) by-pass ducting and be used as a turbofan as well. Still other modifications which fall within the scope of the present invention will be apparent to those skilled in the art, in light of a review of this disclosure, and such modifications are intended to fall within the appended claims.