Abstract:
A gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications.

Description:
BACKGROUND 
       [0001]    This application relates to accessory air flow for use on an aircraft. 
         [0002]    Gas turbine engines typically need a good deal of accessory air. Air is utilized for various purposes such as cooling components on the engine. Also, gas turbine engines utilized on the aircraft also supply air for use in the cabin of the aircraft. All of these applications require relatively high volumes of air. 
         [0003]    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 in such engines, the fan rotates at a slower speed compared to the low pressure compressor. With such engines, the air available for accessory use is moving at a slower speed than in the past, and there may not be sufficient volume as would be desirable. 
       SUMMARY 
       [0004]    In a featured embodiment, a gas turbine engine has a first source of air to be delivered into a core of the engine, and a second source of air, distinct from the first source of air and including separately controlled first and second fans, each delivering air into respective first and second conduits connected to distinct auxiliary applications. 
         [0005]    In another embodiment according to the previous embodiment, the first and second fans are positioned to be downstream of a heat exchanger. 
         [0006]    In another embodiment according to any of the previous embodiments, the heat exchanger is an air to oil cooler. 
         [0007]    In another embodiment according to any of the previous embodiments, one of the applications is for cooling a pitch control mechanism for a propeller included in the gas turbine engine. 
         [0008]    In another embodiment according to any of the previous embodiments, at least one of the applications is for cooling a gear reduction incorporated into the gas turbine engine to drive a propulsor. 
         [0009]    In another embodiment according to any of the previous embodiments, the air to oil cooler receives oil which is utilized to cool the gear reduction for driving the propulsor. 
         [0010]    In another embodiment according to any of the previous embodiments, the first and second fans may be caused to deliver distinct amounts of air to first and second conduits each leading to one of the distinct auxiliary locations. 
         [0011]    In another embodiment according to any of the previous embodiments, at least one of the applications is for an environmental control system. 
         [0012]    In another embodiment according to any of the previous embodiments, at least one of the applications is for an environmental control system. 
         [0013]    In another embodiment according to any of the previous embodiments, the first and second fans are separately controlled such that they may be caused to deliver distinct amounts of air into the first and second conduits. 
         [0014]    In another embodiment according to any of the previous embodiments, at least one of the applications is for cooling a gear reduction incorporated into the gas turbine engine to drive a propulsor. 
         [0015]    In another embodiment according to any of the previous embodiments, one of the applications is for cooling a pitch control mechanism for a propeller included in the gas turbine engine. 
         [0016]    In another embodiment according to any of the previous embodiments, a propulsor is provided in the gas turbine engine. 
         [0017]    In another embodiment according to any of the previous embodiments, the propulsor is driven by a propulsor turbine through a propulsor drive shaft that is downstream of a turbine section driving a compressor section. 
         [0018]    In another embodiment according to any of the previous embodiments, the propulsor turbine drives a fan at an upstream end of the engine. 
         [0019]    In another embodiment according to any of the previous embodiments, the turbine section includes a first and second turbine rotor. The compressor section includes a first and second compressor rotor. The first turbine rotor drives the first compressor rotor, and the second turbine rotor drives the second compressor rotor. 
         [0020]    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. 
         [0021]    In another embodiment according to any of the previous embodiments, the propulsor is at least one propeller. 
         [0022]    In another embodiment according to any of the previous embodiments, the first turbine rotor drives the first compressor rotor through a first shaft and the second turbine rotor drives the second compressor rotor through a second shaft. The first shaft surrounds the second shaft. The propulsor drive shaft is spaced axially further into the engine relative to the first and second shafts. 
         [0023]    In another embodiment according to any of the previous embodiments, the propulsor is a propeller. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0024]      FIG. 1  schematically shows a three spool gas turbine engine. 
           [0025]      FIG. 2A  shows a second embodiment. 
           [0026]      FIG. 2B  shows another embodiment. 
           [0027]      FIG. 3  shows a first embodiment air supply system. 
           [0028]      FIG. 4  shows a second embodiment. 
           [0029]      FIG. 5  shows a schematic system. 
       
    
    
     DETAILED DESCRIPTION 
       [0030]    A gas turbine engine  19  is schematically illustrated in  FIG. 1 . A core engine, or gas generator  20 , includes high speed shaft  21  is 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 . 
         [0031]    Engine  19  also includes a free turbine  34  is shown positioned downstream of the low pressure turbine rotor  32  and serves to drive a propeller  36 . 
         [0032]    Various embodiments are within the scope of the disclosed engine. These include embodiments in which: 
         [0033]    a good deal more work is down by the low pressure compressor rotor  30  than is done by the high pressure compressor rotor  26 ; 
         [0034]    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; 
         [0035]    the low pressure compressor rotor  30  includes eight stages and has a pressure ratio at cruise conditions of 14.5; 
         [0036]    the high pressure compressor rotor  26  had six stages and an overall pressure ratio of 3.6 at cruise; 
         [0037]    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; 
         [0038]    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; 
         [0039]    even more narrowly, the ratio of the two pressure ratios is greater than about 3.5. 
         [0040]    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. 
         [0041]    With the lower compressor, the high pressure turbine rotor  28  may include a single stage. In addition, the low pressure turbine rotor  32  may include two stages. 
         [0042]    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. 
         [0043]    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. 
         [0044]    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. 
         [0045]    The engine  60  as shown in  FIG. 2A  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  65 ), and toward the exit  65 . A gear reduction  63  may be placed between the fan drive turbine  74  and fan  62 . 
         [0046]    The core engine  120 , as utilized in the engine  60 , may have characteristics similar to those described above with regard to the core engine  20 . 
         [0047]    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. 
         [0048]    Another engine embodiment  400  is illustrated in  FIG. 2B . 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. 
         [0049]    All of the engines illustrated in  FIGS. 1 ,  2 A, and  2 B lack a high speed fan delivering air into the inlet of the engine. As such, they all face the challenges with regard to receiving sufficient air volume. 
         [0050]    Further details of the bleed line or port  40  and an associated air supply system are shown in  FIGS. 3 and 4 . 
         [0051]    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 . 
         [0052]    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. 
         [0053]    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   a  and  208   b  to an air supply line  211 . Air supply line  211  may communicate through a valve  212  to a mixing box  198 . 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. 
         [0054]    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. 
         [0055]    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. 
         [0056]    When the valve  212  is open, air flows from the source  211  through the mixing box  198 . 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 . 
         [0057]      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   a  and  308   b.  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. 
         [0058]    As mentioned above, with an embodiment such as shown in  FIG. 2B , there may not be sufficient air delivered for all of the uses anticipated by  FIGS. 3 and 4 . The same is true with the engines shown in  FIGS. 1 and 2A . 
         [0059]    Thus, the present invention utilizes two fans  208 A and  208 B to assist in driving the air flow. The two fans  208 A and  208 B are shown in  FIG. 5  downstream of the heat exchanger  306 . They will serve to induce air flow into two conduits  219 A and  219 B, which will go to distinct applications, such as are shown, for example, in  FIGS. 3 and 4 . Impellers  209 A and  209 B are shown associated with each fan. A control  400  is shown schematically for controlling the speed of the impellers  209 A and  209 B. Now, by controlling the relative speeds of the two fans  208 A and  208 B, the amount of air delivered into the two conduits  219 A and  219 B can be controlled. 
         [0060]    As can be appreciated, the control  400  can control the fan impellers  209 A and  209 B to rotate at distinct speeds. Alternatively, the fans  208 A and  208 B may also be provided with distinct sizes such that they deliver distinct volumes of airflow into conduits  219 A and  219 B. Should the location receiving air from the conduit  219 A require more air than the location receiving air from the conduit  219 B, than the impeller  209 A may be driven at a higher speed than the impeller  209 B to deliver increased airflow to the conduit  219 A. 
         [0061]    In addition, the required volume by the various locations and systems receiving air will vary during flight operation. Thus, the control  400  will be programmed to anticipate the change in airflow volume needs of the system, and to modify the speed and hence the volume of airflow provided by the impellers  209 A and  209 B, as appropriate. Thus, a sufficient quantity of air will be provided for the various applications that may be required on an aircraft application. 
         [0062]    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.