Abstract:
A gas turbine engine has a first particle separator stage including a surface for impacting air at outer periphery of an air flow passage, capturing impacted particles at the outer periphery, and routing captured particles towards a second particle separator stage. Air inward of the first particle separator stage passes towards a core of the engine. Cleaner air upstream of the second particle separator stage is utilized for an air function at a location other than the core engine. A particle discharge is disposed downstream of said second particle separator stage.

Description:
BACKGROUND 
       [0001]    This application relates to a filter particle separator for use in a gas turbine engine. 
         [0002]    Gas turbine engines typically need a good deal of air, such as for core air flow to support the combustion of fuel. In addition, air is utilized for various purposes such as cooling components on the engine. Finally, gas turbine engines utilized on the aircraft also supply air for use in the cabin of the aircraft. All of these applications require relatively clean air. 
         [0003]    It is known to provide a particle separator that will separate impurities from ambient air. Generally, as the air is driven, impurities will tend to be thrown outwardly, and a particle separator is then positioned to remove those particles. 
         [0004]    Historically, a fan drove air into the gas turbine engine. This fan has typically been driven at the same speed as a lower pressure compressor which is downstream of the fan. More recently, a gear reduction has been incorporated between the fan and the low pressure compressor, and the fan rotates at a slower speed compared to the low pressure compressor. With such engines, the air approaching the particle separator is moving at a slower speed than in the past, and there may not be particle separation as efficient as would be desirable. 
       SUMMARY 
       [0005]    In a featured embodiment, gas turbine engine has a first particle separator stage including a surface for impacting air at outer periphery of an air flow passage, capturing impacted particles at the outer periphery, and routing captured particles towards a second particle separator stage. Air inward of the first particle separator stage passes towards a core of the engine. Cleaner air upstream of the second particle separator stage is utilized for an air function at a location other than the core engine. A particle discharge is disposed downstream of the second particle separator stage. 
         [0006]    In another embodiment according to the previous embodiment, the particle separator includes a ring which provides the first particle separator stage, and which extends generally circumferentially to direct separated particles towards the second particle separator stage. 
         [0007]    In another embodiment according to any of the previous embodiments, the second particle separator stage is generally located at a location spaced by 180° from a mount surface for an associated gas turbine engine. 
         [0008]    In another embodiment according to any of the previous embodiments, the cleaner air for the air function passes over an air oil cooler heat exchanger. 
         [0009]    In another embodiment according to any of the previous embodiments, the particle separator surrounds a nose cone. A path into the core engine is downstream and radially inwardly of the surface, and outwardly of the surface of the nose cone. 
         [0010]    In another embodiment according to any of the previous embodiments, the cleaner air for the air function passes over an air oil cooler heat exchanger. 
         [0011]    In another embodiment according to any of the previous embodiments, the particle separator surrounds a nose cone. A path into the core engine is downstream and radially inwardly of the surface, and outwardly of the surface of the nose cone. 
         [0012]    In another embodiment according to any of the previous embodiments, the particle separator surrounds a nose cone. A path into the core engine is downstream and radially inwardly of the surface, and outwardly of the surface of the nose cone. 
         [0013]    In another embodiment according to any of the previous embodiments, the core engine includes a turbine section, a compressor section and a propulsor section. There is at least two turbine rotors in the turbine section, a first and second compressor rotor, and a propulsor. The at least two turbine rotors each drive one of the first and second compressor rotors. The propulsor is not driven at the same speed as either of the first and second turbine rotors. 
         [0014]    In another embodiment according to any of the previous embodiments, the propulsor is driven by one of the first and second turbine rotors which drives a lower stage one of the first and second compressor rotors. There is a gear reduction between one of the first and second turbine rotors and the propulsor. 
         [0015]    In another embodiment according to any of the previous embodiments, the propulsor is driven by a third propulsor turbine rotor, which is positioned to be downstream of the at least two turbine rotors. 
         [0016]    In another embodiment according to any of the previous embodiments, the second compressor rotor is downstream of the first compressor rotor. The second compressor rotor has a first overall pressure ratio. The first compressor rotor has a second overall pressure ratio. A ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 2.0. 
         [0017]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 3.0. 
         [0018]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio is greater than or equal to about 3.5. 
         [0019]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio is less than or equal to about 8.0. 
         [0020]    In another embodiment according to any of the previous embodiments, the propulsor turbine drives a fan at an upstream end of the engine. 
         [0021]    In another embodiment according to any of the previous embodiments, the ratio of the first overall pressure ratio to the second overall pressure ratio being less than or equal to about 8.0. 
         [0022]    In another embodiment according to any of the previous embodiments, the propulsor turbine drives a fan at an upstream end of the engine. 
         [0023]    In another embodiment according to any of the previous embodiments, an axially outer position is defined by the fan. The propulsor turbine is positioned between the fan and the first and second turbine rotors. The first and second compressor rotors are positioned further into the engine relative to the first and second turbine rotors. 
         [0024]    In another embodiment according to any of the previous embodiments, the propulsor is a plurality of propellers. 
         [0025]    These and other features can be best understood from the following specification and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0026]      FIG. 1A  shows a detail of a particle separator. 
           [0027]      FIG. 1B  shows another detail. 
           [0028]      FIG. 1C  shows features of the particle separator. 
           [0029]      FIG. 1D  shows a further detail. 
           [0030]      FIG. 2A  schematically shows a three spool gas turbine engine. 
           [0031]      FIG. 2B  shows a second engine type. 
           [0032]      FIG. 2C  schematically shows another engine type. 
           [0033]      FIG. 3  shows a first embodiment air supply system. 
           [0034]      FIG. 4  shows a second embodiment. 
       
    
    
     DETAILED DESCRIPTION 
       [0035]    The air to be delivered to the compressor of a gas turbine engine and to various uses, examples of which follow, must be relatively clean. As mentioned, the style of engines such as disclosed below raise challenges to providing clean air. 
         [0036]    Thus, as shown in  FIGS. 1A and 1B , a particle separator  198  has an internal cavity that receives core air flow C. The air flows around the outer periphery of the nose cone  321 . 
         [0037]    As can be appreciated from  FIG. 1B , an internal passage  302  receives the core air flow downstream of the particle separator  198 . As can also be appreciated from  FIG. 1B , the separator  198  includes a radially outer surface  197  that includes a radially inner wall  367  to define a capture chamber  369  that catches heavier particles around the periphery of the particle separator  198 . Air at the inlet location  202  would be cleaner towards radially inner locations, and include more dirt particles at radially outer locations, and those particles would tend to catch in the chamber  369 . 
         [0038]    As shown in  FIG. 1C , the chamber  369  within the surface  197  would tend to gather and force particles around the periphery of the particle separator  198  to gather in a lower location particle separator  204 . Another flow path  314  will pass the heat exchanger  306 , shown schematically, and such as disclosed below in the  FIGS. 3 and 4  embodiments. 
         [0039]    The lower location particle separator  204  provides a particle separator at a vertically lower location such that gravity will assist in removing the particles. 
         [0040]    As shown in  FIG. 1D , the heat exchanger  306  will receive air that is somewhat dirtier than the core air delivered at  302 , however, the dirtiest air would be delivered outwardly at particle separator portion  204 . Downstream of particle separator  204 , the air may be delivered for nacelle ventilation. Thus, there are two stages of particle separation, with the first stage directing cleaner air into the core than that delivered outwardly, and a second stage delivering cleaner air to the heat exchanger  306  than is directed into the particle separator  204 . 
         [0041]    As shown in  FIG. 1A , a mount surface  800  may be associated with the engine to mount the engine to an aircraft. The particle separator  204  may be spaced by 180° relative to mount surface  800 . 
         [0042]    A gas turbine engine  19  is schematically illustrated in  FIG. 2A . A core engine, or gas generator  20 , includes a high speed shaft  21  as part of a high speed spool along with a high pressure turbine rotor  28  and a high pressure compressor rotor  26 . A combustion section  24  is positioned intermediate the high pressure compressor rotor  26  and the high pressure turbine rotor  28 . A shaft  22  of a low pressure spool connects a low pressure compressor rotor  30  to a low pressure turbine rotor  32 . 
         [0043]    Engine  19  also includes a free turbine  34  shown positioned downstream of the low pressure turbine rotor  32  and serving to drive a propeller  36 . 
         [0044]    Various embodiments are within the scope of the disclosed engine. These include embodiments in which: 
         [0045]    a good deal more work is down by the low pressure compressor rotor  30  than is done by the high pressure compressor rotor  26 ; 
         [0046]    the combination of the low pressure compressor rotor  30  and high pressure compressor rotor  26  provides an overall pressure ratio equal to or above about 30; 
         [0047]    the low pressure compressor rotor  30  includes eight stages and has a pressure ratio at cruise conditions of 14.5; 
         [0048]    the high pressure compressor rotor  26  had six stages and an overall pressure ratio of 3.6 at cruise; 
         [0049]    a ratio of the low pressure compressor pressure ratio to the high pressure compressor ratio is greater than or equal to about 2.0, and less than or equal to about 8.0; 
         [0050]    more narrowly, the ratio of the two pressure ratios is between or equal to about 3.0 and less than or equal to about 8; 
         [0051]    even more narrowly, the ratio of the two pressure ratios is greater than about 3.5. 
         [0052]    In the above embodiments, the high pressure compressor rotor  26  will rotate at slower speeds than in the prior art. If the pressure ratio through the fan and low pressure compressor are not modified, this could result in a somewhat reduced overall pressure ratio. The mechanical requirements for the high pressure spool, in any event, are relaxed. 
         [0053]    With the lower pressure compressor  30  performing more of the compression, the high pressure turbine rotor  28  may include a single stage. In addition, the low pressure turbine rotor  32  may include two stages. 
         [0054]    By moving more of the work to the low pressure compressor rotor  30 , there is less work being done at the high pressure compressor rotor  26 . In addition, the temperature at the exit of the high pressure compressor rotor  26  may be higher than is the case in the prior art, without undue challenges in maintaining the operation. 
         [0055]    A bleed line or port  40  is positioned between the low pressure compressor rotor  30  and the high pressure compressor rotor  26 . Details of this porting are disclosed below. 
         [0056]    Variable vanes are less necessary for the high pressure compressor rotor  26  since it is doing less work. Moreover, the overall core size of the combined compressor rotors  30  and  26  is reduced compared to the prior art. 
         [0057]    An alternative engine  60  as shown in  FIG. 2B  includes a two spool core engine  120  including a low pressure compressor rotor  30 , a low pressure turbine rotor  32 , a high pressure compressor rotor  26 , and a high pressure turbine rotor  28 , and a combustor  24  as in the prior embodiments. This core engine  120  is a so called “reverse flow” engine meaning that the compressor  30 / 26  is spaced further into the engine than is the turbine  28 / 32 . Air downstream of the fan rotor  62  passes into a bypass duct  64 , and toward an exit  65 . However, a core inlet duct  66  catches a portion of this air and turns it to the low pressure compressor  30 . The air is compressed in the compressor rotors  30  and  26 , combusted in a combustor  24 , and products of this combustion pass downstream over the turbine rotors  28  and  32 . The products of combustion downstream of the turbine rotor  32  pass over a fan drive turbine  74 . Then, the products of combustion exit through an exit duct  76  back into the bypass duct  64  (downstream of inlet  66  such that hot gas is not re-ingested into the core inlet  66 ), and toward the exit  65 . A gear reduction  63  may be placed between the fan drive turbine  74  and fan  62 . 
         [0058]    The engines  19  and  60  are similar in that they have what may be called a propulsor turbine ( 34  or  74 ) which is spaced to be axially downstream of the low pressure turbine rotor  32 . Further, the high pressure spool radially surrounds the low pressure spool, but neither of the spools surround the propulsor turbine, nor the shaft  100  connecting the propulsor turbine to the propellers  36  or fan  62 . In this sense, the propulsor rotor is separate from the gas generator portion of the engine. 
         [0059]    Another engine embodiment  400  is illustrated in  FIG. 2C . In embodiment  400 , a fan rotor  402  is driven by a fan drive turbine  408  through a gear reduction  404 . A lower pressure compressor  406  is also driven by the fan drive turbine  408 . A high pressure turbine  412  drives a high pressure compressor  410 . A combustor section  414  is located between the compressor sections  406 / 410  and turbine sections  412 / 408 . In such engines, the fan  402  now rotates at a slower speed than it would have in a direct drive engine. 
         [0060]    All of the engines illustrated in  FIGS. 2A ,  2 B, and  2 C lack a high speed fan delivering air into the inlet of the engine. As such, they all face the challenges with regard to particle separation mentioned above. 
         [0061]    Further details of the bleed line or port  40  and an associated air supply system are shown in  FIGS. 3 and 4 . 
         [0062]    Particularly with an engine as disclosed above, the low pressure compressor  30  is supplying a higher pressure than is typically been the case in the past. As such, this compressor can be utilized as a source of air for environmental control systems on an associated aircraft. In the past, a higher pressure source has typically been required resulting in taps from the high pressure compressor. 
         [0063]    As shown in  FIG. 3 , an air supply system  190  incorporates a manifold  192  provided with a plurality of bleed lines or ports  194  and which communicate with an intermediate compressor case  200 . The intermediate compressor case  200  is positioned between the low pressure compressor  30  and the high pressure compressor  26 . 
         [0064]    The pressure of the air supplied by the low pressure compressor  30  will vary dramatically during operation of an associated engine. Thus, at some point, the air pressure delivered from the ports  194  may be undesirably high. 
         [0065]    A supply of lower pressure air is used to address this concern. An inlet  202  to a low pressure manifold  199  communicates through a heat exchanger  206 . The heat exchanger  206  may be utilized to cool oil at other locations. A particle separator  204  is positioned to filter dirt particles out of an air supply stream being delivered downstream through fans  208  to an air supply line  211 . Air supply line  211  may communicate through a valve  212  to a mixing box  398 . The valve  212  is controlled in combination with a valve  196  associated with the manifold  192 , such that the flow of air from the higher pressure manifold  192  and the lower pressure source  211 , are properly mixed to achieve a desired pressure at an outlet  310 . The outlet  310  leads to an environmental control system  400  for supplying air for use on an associated aircraft. 
         [0066]    A control, such as a full authority digital engine control, may control the valves  196  and  212 , based upon the pressure, temperature and any other variables within the operation of the associated engine. 
         [0067]    A worker of ordinary skill in the art would recognize how to achieve a desired pressure at the outlet  310 . The desired pressure at the outlet  310  may be dictated by the aircraft manufacturer. 
         [0068]    When the valve  212  is open, air flows from the source  211  through the mixing box  398 . However, as the valve  212  is moved toward a more closed position, that air is delivered through an outlet  214  downstream of the high pressure compressor  26 . 
         [0069]      FIG. 4  shows an alternative embodiment  250 . Alternative embodiment  250  is generally the same as the embodiment  190 . An inlet  302  leads into a low pressure supply manifold  300 . There is a dirt separator  304 , a heat exchanger  306  and fans  308 . Valves  312  and  296  are controlled to control the pressure of the air reaching a mixing box  298  which communicates with an outlet  311 , and eventually the environmental control system  400 . A pipe  510  communicating a lower pressure air supply into the mixing box  298  passes air through a one-way valve  420  and to the outlet  512 , similar to the first embodiment. 
         [0070]    The particle separator disclosed in  FIGS. 1A-1D  is particularly beneficial when used in an engine such as disclosed in  FIGS. 2A-2B , and providing the additional functions as shown, for example, in  FIGS. 3 and 4 . 
         [0071]    Although an embodiment of this invention 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 true scope and content of this disclosure.