Patent Publication Number: US-8979470-B2

Title: Gas turbine engine and method for cooling the compressor of a gas turbine engine

Description:
RELATED APPLICATION 
     This application claims priority under 35 U.S.C. §119 to European Patent Application No. 10172376.5 filed in Europe on Aug. 10, 2010, the entire content of which is hereby incorporated by reference in its entirety. 
     FIELD 
     The present disclosure relates to a gas turbine engine and a method for cooling the compressor of a gas turbine engine. 
     BACKGROUND INFORMATION 
     Gas turbine engines are known to include a compressor wherein air is compressed to be then fed into a combustion chamber. Within the combustion chamber a fuel is injected into the compressed air and is combusted, generating high temperature and pressure flue gases that are expanded in a turbine. 
     A known gas turbine engine has a rotor shaft that carries at one end a compressor drum (carrying compressor rotor blades), and at the opposite end, turbine disks (carrying turbine rotor blades). The combustion chamber is provided between the compressor drum and the turbine disks. 
     The compressor drum has circumferential seats (shaped like circumferential dove tale slots) into which the compressor rotor blades are housed. 
     A casing is provided, which carries guide vanes for the compressor (compressor guide vanes) and for the turbine (turbine guide vanes). 
     The last stages of the compressor (where the air pressure is higher) can be thermally highly stressed. 
     The temperature of the compressed air at the outlet of the compressor can be high and the components at the last stages of the compressor can be cooled via cooling air injected into a gap between the compressor drum and the combustion chamber. The cooling air can be compressed air extracted downstream of the compressor before it enters the combustion chamber. 
     Therefore an equilibrium exists, which can allow a high lifetime for the parts concerned for the expected operating temperatures and stress, in particular, the compressor rotor, disk and blades that are the most stressed components of the compressor. 
     In order to increase power output and efficiency, it is desirable to increase the air mass flow through the compressor in order to increase the fuel mass flow that can be injected into the combustion chamber. This can increase the mass flow and temperature of the flue gases through the turbine. 
     Increasing the mass flow through the compressor can cause the temperature of the compressed air, for example, at the outlet of the compressor, to increase. 
     Such a temperature increase (tests showed that it could be as large as 20-30° C.) can influence the lifetime of the components affected. 
     With reference to  FIG. 10  (curve A), the dependence of the lifetime of the parts, for example, the compressor, rotor, disk and blades, from the temperature of the compressed air at the compressor outlet is shown. From this diagram it is clear that also a small temperature increase (e.g., an increase of about 20-30° C.) can cause a large lifetime decrease. Such a lifetime decrease may not be acceptable, because it can cause the expected lifetime of the affected components to fall below the minimum admissible lifetime. 
     SUMMARY 
     A gas turbine engine is disclosed, comprising a compressor including a compressor drum and rotor blades having roots connected into seats of a compressor drum, wherein at least one of the rotor blade roots and the compressor drum include longitudinal passages for a cooling fluid, the longitudinal passages connecting higher pressure areas to lower pressure areas of the gas turbine engine. 
     A method is disclosed for cooling a compressor of a gas turbine engine, the compressor including a compressor drum and rotor blades having roots connected into seats of the compressor drum, the method comprising: forming at least one of the blade roots and the compressor drum with longitudinal passages for a cooling fluid, the longitudinal passages connecting higher pressure areas to lower pressure areas of the gas turbine engine; and passing a cooling fluid through the longitudinal passages. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Further, characteristics and advantages of the disclosure will be more apparent from the description of exemplary embodiments of the gas turbine engine and method illustrated by way of non-limiting example in the accompanying drawings, in which: 
         FIG. 1  is a schematic view of an exemplary embodiment of compressor rotor blades connected to a rotor drum; 
         FIG. 2  is a schematic cross section through line II-II of  FIG. 1 ; 
         FIGS. 3 and 4  are cross sections respectively through lines III-Ill and IV-IV of  FIG. 2 ; 
         FIGS. 5 and 6  show different exemplary embodiments of root blade passages; 
         FIGS. 7 through 9  show respectively an exemplary embodiment of a compressor rotor blade, an exemplary embodiment of a compressor rotor spacer and an exemplary embodiment of compressor rotor blade; and 
         FIG. 10  shows the relationship between lifetime and temperature at the compressor outlet for a known gas turbine engine (curve A) and a gas turbine engine in an exemplary embodiment of the disclosure (curve B). 
     
    
    
     DETAILED DESCRIPTION 
     The disclosure provides an engine and a method for allowing a gas turbine compressor to compress air until it reaches a temperature higher than in known gas turbines, without unacceptably reducing the lifetime of the components affected, for example, without unacceptably reducing the compressor rotor, disk and blade lifetime. 
     With reference to the figures, an exemplary gas turbine engine includes a compressor, one or more combustion chambers (according to the configuration), and a turbine. In different exemplary embodiments, the engine may also be a sequential combustion gas turbine engine and include a compressor, one or more combustion chambers (according to the configuration), a high pressure turbine, one or more further combustion chambers (according to the configuration), and a low pressure turbine. 
     The compressor  1  can be an axial compressor having a compressor drum  2  with compressor rotor blades  3  and compressor guide vanes  5 . 
     The rotor blades  3  have roots  7  connected into seats  8  of the compressor drum  2 . 
     As shown in  FIG. 1 , the blade roots  7  define longitudinal passages  9  and/or the compressor drum  2  defines longitudinal passages  10  for a cooling fluid. The longitudinal passages  9 ,  10  connect higher pressure areas  13  to lower pressure areas  14  of the gas turbine engine. 
     The differential pressure between the higher and lower pressure areas  13 ,  14  can allow cooling air circulation. 
     The seats  8  can be defined by longitudinal slots into which the blade roots  7  are inserted. 
     The passages  9  of the blade roots  7  can be defined by longitudinal channels  11  provided in the blade roots  7 . All the blade roots  7  inserted into the same seat  8  have their channels connected together to define the passage  9  running over at least a portion of the compressor drum  2 . 
     In a first exemplary embodiment ( FIG. 9 ), the blades  3  have a structure with a platform  15  larger in the longitudinal direction (e.g., the direction of the passages  9 ) than the longitudinal size of the airfoil  16  carried by it. This can allow the rotor blades  3  to be directly connected one next to the other and, at the same time, can leave a gap between two next airfoils  16 , for a guide vane  5 . 
     In an exemplary embodiment, the rotor blades  3  have a structure with a platform  15  substantially as large in the longitudinal direction (e.g., in the direction of the passages  9 ) as the longitudinal size of the airfoils  16 . 
     In this case spacers  18  between two adjacent blade roots  7  housed into the same seat  8  can be provided. The spacers  18  have a spacer root  19  and a platform  20  defining, with the platforms  15  of the blades  3 , a compressed air path  22 . 
     Also the spacer&#39;s roots  19  have longitudinal channels  23  that can be connected to the channels  11  of the blade roots  7  to define the longitudinal passages  9 . 
     The higher and lower pressure areas can be defined in different positions of the engine. 
     For example, downstream of the compressor drum  2 , a gap  25  separating it from a combustion chamber  26  can be provided. 
     Within this gap  25  a protrusion  27  can be provided, to close the compressed air path  22 . 
     The higher pressure areas  13  can be defined between the protrusion  27  and the compressed air path  22  and the lower pressure areas  14  can be defined by areas of the gap  25  below the protrusion  27 . 
     In an exemplary embodiment, the higher pressure areas  13  can be defined between the protrusion  27  and the compressed air path  22  (as in the embodiment above described), and the lower pressure areas  14  can be defined in the inside of a holed compressor drum  2 . 
     The longitudinal passages  9 ,  10  can be provided over the whole compressor drum longitudinal length or only over a portion thereof. For example, the latter is desirable, because at the first stages of the compressor a large cooling may not be needed. 
     In order to connect the passages  9 ,  10  between the higher and lower pressure areas  13 ,  14 , a circumferential chamber  28  extending at an intermediate position of the compressor drum  2  can be provided. 
     The circumferential chamber  28  can be connected to the longitudinal passages  9  of the blade roots  7  and/or to the longitudinal passages  10  of the compressor drum  2  (e.g., according to the particular cooing scheme). 
     In a exemplary embodiment, both longitudinal passages  9 ,  10  of the blade roots  7  and rotor drum  2  can be provided. These longitudinal passages  9 ,  10  have axes parallel to an engine longitudinal axis  30  and have the same radial distance from it. 
     The longitudinal passages  9  of the blade roots  7  can be connected to the lower pressure areas  14  and the longitudinal passages  10  of the compressor drum  2  can be connected to the higher pressure areas  13 . 
     In the following, exemplary embodiments of the disclosure are described in detail with reference to the figures. 
     In a first exemplary embodiment ( FIGS. 1 through 4 ), both the longitudinal passages  9 ,  10  of the blade roots  7  and compressor drum  2  are provided. 
     In this case, the passages  10  can be straight passages over their whole length (i.e., they are parallel to the engine longitudinal axis  30 ) and have one end opening in the high pressure areas  13  of the gap  25  and the opposite end opening in the circumferential chamber  28 . 
     The longitudinal passages  9  have one end opening in the circumferential chamber  28  and extend straight (i.e., parallel to the axis  30 ) within the blade roots  7 . Then, a terminal portion  32  provided within the compressor drum  2  is bent to the straight part and opens in the lower pressure areas  14  of the gap  25 . In a exemplary embodiment, the bent portion  32  can be connected to a radial or bent portion  32   a  realised within the root  7  of the last blade  3  (i.e., the blade  3  that is closest to the combustion chamber  26 ). 
     In this embodiment, the seats  8  extend up to the border of the drum  2  facing the combustion chamber  26  and a locking element  34  is provided, to lock the blades  3  therein. 
     The operation of the compressor in this embodiment is the following. 
     Air passes through the compressed air path  22  and is compressed. Downstream of the compressor, a part of the compressed air is extracted and is cooled (in a cooler, not shown) to be then fed into the gap  25  as cooling air. 
     From the gap  25  (for example, its higher pressure areas  13 ) the cooling air enters the longitudinal passages  10  and passes through them reaching the circumferential chamber  28 . This lets the compressor drum  2  be cooled. 
     Then from the circumferential chamber  28 , the cooling air enters the longitudinal passages  9  of the blade roots  7  and passes through them, cooling them down. 
     From the longitudinal passage  9  of the last blade  3 , the cooling air enters the portion  32   a  and then the bent terminal portion  32 , to be discharged into the lower pressure areas  14  of the gap  25 . 
     This embodiment allows cooling of the compressor drum  2  and rotor roots  7 . 
     This embodiment may be implemented either with the rotor blades and spacers shown in  FIGS. 7 and 8 , or with the rotor blades shown in  FIG. 9  or combination thereof. 
     Different embodiments in which the passages  9  are connected to the higher pressure areas  13  and the passages  10  are connected to the lower pressure areas  14  or embodiments implementing even further cooling schemes are possible. 
     In a second exemplary embodiment, only the longitudinal passages  9  of the rotor blades  7  are provided. 
     For example, in this case, some of the longitudinal passages  9  may have a bent terminal portion (as shown in  FIG. 3 ) opening into the lower pressure areas  14  of the gap  25  and an opposite end opening in the circumferential chamber  28 , and other passages  9  (see  FIG. 5 ) may have an end opening in the circumferential chamber  28  and an opposite straight terminal portion  33  that may be realised within the locking element  34  (e.g., the terminal portion is not bent to the channels  11 , but it is coaxial with them and parallel to the axis  30 ) opening in the higher pressure areas  13  of the gap  25 . 
     The passages with bent terminal portions  32  can be alternated to passages with straight terminal portions  33 . 
     This embodiment can be implemented either with the rotor blades and spacers shown in  FIGS. 7 and 8 , with the rotor blades shown in  FIG. 9  or combination thereof. 
     This embodiment can be useful in case a limited cooling is desired. Additionally it can allow an easy machining. 
     In a third exemplary embodiment, only the passages  10  of the compressor drum  2  are provided. 
     Also in this case, some of the longitudinal passages  10  can have a bent terminal portion opening into the lower pressure areas  14  of the gap  25  and an opposite end opening in the circumferential chamber  28 , and other longitudinal passages  10  can have an end opening in the circumferential chamber  28  and an opposite straight terminal portion opening in the higher pressure areas  13  of the gap  25 . Passages with bent terminal portions can be alternated to passages with straight terminal portions. 
     This embodiment may be useful in case a limited cooling, for example, for the rotor drum  2 , is desired. 
     The operation of the compressor in the second and third embodiments can be substantially the same as the first embodiment described and, with particular reference to the second embodiment, it is the following. 
     The cooling air enters into the passages  9  with straight terminal portion  33  and passes through them, cooling the roots  7  and the rotor drum  2 , to then enter the circumferential chamber  28 . 
     From the circumferential chamber  28  it enters the passages  9  having the bent terminal portion  32 , to further cool the roots  7  and rotor drum  2 . 
     Then the cooling air is discharged into the lower pressure areas  14  of the gap  25 . 
     In exemplary embodiments (see  FIG. 6 ), the compressor can have the passages  9  of the blades root, or the passages  10  of the compressor drum  2  or both the passages  9  and  10  that have a straight terminal portion opening in the higher pressure areas  13  of the gap  25  and an opposite end opening into the circumferential chamber  28 . 
     The circumferential chamber  28  has a hole or duct  35  connecting it to the inside  36  of the rotor drum  2 . Further holes or duct  37  can then be provided, connecting the inside  36  of the rotor drum  2  (or inside of a hollow rotor shaft that is connected to the hollow rotor drum) to lower pressure areas  13  of the engine. 
     For example, a hole or duct  37  can be provided connecting the inside  36  of the compressor drum  2  to the gap  25 . In exemplary embodiments such holes or ducts can be provided in positions of the rotor shaft further downstream, to use the cooling air from the compressor  1  as cooling air for the turbine. 
     The operation of the compressor in this embodiment is as follows. 
     The cooling air enters the passages  9  and/or  10  and passes through them cooling the compressor drum  2  and blade roots  7  down. The cooling air enters the circumferential chamber  28 , to then enter (via the hole or duct  35 ) the inside  36  of the compressor drum  2 . 
     From the inside  36  of the compressor, drum  2  the cooling air enters the gap  25  via the hole or duct  37  or other position according to the cooling scheme. 
     The present disclosure also relates to a method for cooling the compressor of a gas turbine engine. 
     The method includes making a cooling fluid pass through the longitudinal passages  9 ,  10  of the blade roots  7  and/or compressor drum  2 , to cool them down. 
       FIG. 10  shows the dependence of the lifetime of the parts on the temperature at the compressor outlet. Respectively curve A refers to a known gas turbine engine and curve B refers to a gas turbine engine of an exemplary embodiment of the disclosure. 
       FIG. 10  shows that curve B is shifted towards the high temperatures and, thus, for the same compressor outlet temperature, the engine in the embodiments of the disclosure have a much longer lifetime or, for the same lifetime, the engine in embodiments of the disclosure can operate with a higher temperature, allowing a higher compression degree at the compressor and, thus, larger power generation and higher efficiency than in known gas turbine engines. 
     The features described may be independently provided from one another. 
     In practice, the materials used and the dimensions can be chosen at will according to specification, and to the state of the art. 
     Thus, it will be appreciated by those skilled in the art that the present invention can be embodied in other specific forms without departing from the spirit or essential characteristics thereof. The presently disclosed embodiments are therefore considered in all respects to be illustrative and not restricted. The scope of the invention is indicated by the appended claims rather than the foregoing description and all changes that come within the meaning and range and equivalence thereof are intended to be embraced therein. 
     Reference Numbers 
     
         
           1  compressor 
           2  compressor drum 
           3  compressor rotor blades 
           5  compressor guide vanes 
           7  roots of  3   
           8  seats 
           9  longitudinal passages of  7   
           10  longitudinal passages of  2   
           11  channels of  7   
           13  higher pressure areas 
           14  lower pressure areas 
           15  platform of  3   
           16  airfoil of  3   
           18  spacers 
           19  roots of  18   
           20  platforms of  18   
           22  compressed air path 
           23  channel of  18   
           25  gap 
           26  combustion chamber 
           27  protrusion 
           28  circumferential chamber 
           30  engine longitudinal axis 
           32  bent terminal portion of  9   
           32   a  portion of  9   
           33  straight terminal portion of  9   
           34  locking element 
           35  hole of  2   
           36  inside of  2   
           37  hole of  2   
         A dependence of the lifetime on the temperature at the compressor outlet for a known gas turbine engine 
         B dependence of the lifetime on the temperature at the compressor outlet for a gas turbine engine in an exemplary embodiment.