Abstract:
One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is also a unique gas turbine engine. A further embodiment is a unique method for operating a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and bleed air systems therefor. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims benefit of U.S. Provisional Patent Application No. 61/428,714, filed Dec. 30, 2010, entitled GAS TURBINE ENGINE WITH BLEED AIR SYSTEM, which is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates to gas turbine engines, and more particularly, to gas turbine engines having bleed air systems. 
       BACKGROUND 
       [0003]    Gas turbine engine bleed air systems remain an area of interest. Some existing systems have various shortcomings, drawbacks, and disadvantages relative to certain applications. Accordingly, there remains a need for further contributions in this area of technology. 
       SUMMARY 
       [0004]    One embodiment of the present invention is a unique gas turbine engine. Another embodiment of the present invention is also a unique gas turbine engine. A further embodiment is a unique method for operating a gas turbine engine. Other embodiments include apparatuses, systems, devices, hardware, methods, and combinations for gas turbine engines and bleed air systems therefor. Further embodiments, forms, features, aspects, benefits, and advantages of the present application will become apparent from the description and figures provided herewith. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]    The description herein makes reference to the accompanying drawings wherein like reference numerals refer to like parts throughout the several views, and wherein: 
           [0006]      FIG. 1  schematically illustrates some aspects of a non-limiting example of a gas turbine engine in accordance with an embodiment of the present invention. 
           [0007]      FIG. 2  schematically illustrates some aspects of a non-limiting example of a three-spool gas turbine engine in accordance with an embodiment of the present invention. 
           [0008]      FIG. 3  schematically illustrates some aspects of a non-limiting example of a two-spool gas turbine engine in accordance with an embodiment of the present invention. 
           [0009]      FIG. 4  schematically illustrates some aspects of a non-limiting example of a two-spool gas turbine engine in accordance with an embodiment of the present invention. 
           [0010]      FIG. 5  schematically illustrates some aspects of a non-limiting example of a bleed system in accordance with an embodiment of the present invention. 
           [0011]      FIG. 6  illustrates some aspects of a non-limiting example of a vane/strut of a gas turbine engine in accordance with an embodiment of the present invention. 
       
    
    
     DETAILED DESCRIPTION 
       [0012]    For purposes of promoting an understanding of the principles of the invention, reference will now be made to the embodiments illustrated in the drawings, and specific language will be used to describe the same. It will nonetheless be understood that no limitation of the scope of the invention is intended by the illustration and description of certain embodiments of the invention. In addition, any alterations and/or modifications of the illustrated and/or described embodiment(s) are contemplated as being within the scope of the present invention. Further, any other applications of the principles of the invention, as illustrated and/or described herein, as would normally occur to one skilled in the art to which the invention pertains, are contemplated as being within the scope of the present invention. 
         [0013]    Referring to the drawings, and in particular  FIG. 1 , some aspects of a non-limiting example of a gas turbine engine  10  in accordance with an embodiment of the present invention are schematically depicted. In one form, engine  10  is a thrust-producing engine for aircraft propulsion. In other embodiments, engine  10  may be employed for other purposes, e.g., to power a generator, pump and/or other equipment. In one form, engine  10  is a turbofan engine. In other embodiments, engine  10  may be a turboprop engine, a turboshaft engine, a propfan engine, or any other type of thrust producing gas turbine engine. Gas turbine engine  10  includes or is coupled to a load absorber  12 . Gas turbine engine  10  includes a compressor system  14  including a compressor stage  16 ; a combustion system  18 ; a turbine system  20  including a turbine stage  22 ; and a bleed system  24 . In one form, load absorber  12  is a propulsor, e.g., a propulsor  12 . In one form, propulsor  12  is a fan, i.e., a turbofan, and in various embodiments may or may not include one or more low pressure compressor stages. In other embodiments, propulsor  12  may take other forms, and may be, for example, a propeller system or other thrust producing rotor system. In still other embodiments, load absorber  12  may be any type of load absorber, and may be, for example and without limitation, a generator, a pump or compressor, and/or may be configured to power a ground-based or water-borne vehicle. 
         [0014]    Compressor system  14  is a multistage gas turbine engine compressor system, of which compressor stage  16  is an intermediate compressor stage, i.e., fluidly disposed between the lowest pressure compressor stage (including any compressor stage that rotates with propulsor  12 , for such embodiments so equipped) and the highest pressure compressor stage of compressor system  14 . In various embodiments, compressor system  14  may include one or more rotors operating at the same or different speeds. 
         [0015]    Combustion system  18  is in fluid communication with compressor system  14 . Combustion system  18  may be any suitable gas turbine engine combustion system. Turbine system  20  is in fluid communication with combustion system  18 . Turbine system  20  is a multistage gas turbine engine turbine system, of which turbine stage  22  is an intermediate turbine stage, i.e., disposed downstream of the highest pressure turbine stage of turbine system  20 . In typical embodiments, turbine stage  22  is a low pressure turbine stage, although in other embodiments, turbine stage  22  may not be considered a low pressure turbine stage. Turbine stage  22  operates at a lower pressure than compressor stage  16 . Turbine stage  22  is drivingly coupled to load absorber  12 , e.g., a turbofan. Compressor system  14 , combustion system  18  and turbine system  20  operate as in conventional gas turbine engines, except for the use of bleed system  24 , some embodiments of which are described herein. 
         [0016]    Bleed system  24  is configured to bleed pressurized air from compressor stage  16  and to deliver the bleed air to the turbine stage  22  in response to a discharge temperature of compressor system  14  (e.g., the exit total temperature of the final compressor stage of compressor system  14 ) reaching a predetermined temperature limit. In various embodiments, bleed system  24  may deliver the bleed air to other turbine stages in addition to turbine stage  22 . By providing the bleed air to turbine stage  22 , turbine stage  22  and turbine stages downstream of turbine stage  22  receive an increased amount of engine  10 &#39;s working fluid under pressure, from which turbine stage  22  and turbine stages downstream of turbine stage  22  may extract additional power for driving load absorber  12  than had the bleed air not been provided. Thus, some aspects of the present invention include not only preventing an over-temperature condition in compressor system  14  by controlling compressor system  14  discharge temperature relative to the predetermined temperature limit, but also include simultaneously increasing the amount of power provided to load absorber  12 . In one form, the predetermined temperature limit represents a design upper temperature limit of the final stage of compressor system  14 , e.g., designated as such for life determination purposes for one or more components compressor system  14 . Bleed system  24  may be employed in many different gas turbine engine  10  configurations, only a few of which are illustrated in  FIGS. 2-4 . 
         [0017]    Referring to  FIG. 2 , some aspects of a non-limiting example of a three-spool gas turbine engine  100  in accordance with an embodiment of the present invention is schematically depicted. Engine  100  represents one of many possible configurations of engine  10  of  FIG. 1 . Engine  100  includes a turbofan  112 , a compressor system  114 , including an intermediate pressure (IP) compressor  116  and a high pressure (HP) compressor  117 ; a combustion system  118 , and a turbine system  120 , including an HP turbine  121 A, and IP turbine  121 B and a low pressure (LP) turbine  122 . It will be noted that some reference numerals in  FIG. 2  correspond to the reference numerals for like parts of  FIG. 1 , with the addition of a numerical value of 100 to the reference numerals of  FIG. 1  to achieve the reference numerals in  FIG. 2 . For example, load absorber  12  of  FIG. 1  is represented by fan  112  of  FIG. 2 ; compressor stage  16  of  FIG. 1  is represented in the embodiment of  FIG. 2  as IP compressor  116 ; and turbine stage  22  of  FIG. 1  is represented as LP turbine  122  in  FIG. 2 . In one form, IP compressor  116  includes a plurality of compressor stages. In other embodiments, IP compressor  116  may include only a single compressor stage. In one form, LP turbine  122  includes a plurality of turbine stages. In other embodiments, LP turbine  122  may include only a single turbine stage. 
         [0018]    IP compressor  116  is in fluid communication with fan  112 . HP compressor  117  is in fluid communication with IP compressor  116 . Combustor  118  is in fluid communication with HP compressor  117 . HP turbine  121 A is drivingly coupled to HP compressor  117  and in fluid communication with combustion system  118 . IP turbine  121 B is driving coupled to the IP compressor and is in fluid communication with HP turbine  121 A. LP turbine  122  is coupled to fan  112 , and is in fluid communication with IP turbine  121 B. Engine  100  also includes bleed system  24 , which in the embodiment of  FIG. 2  is configured to bleed pressurized air from the discharge of IP compressor  116  and/or other IP compressor  116  compressor stages, and to deliver the bleed air to one or more turbine stages of LP turbine  122  in response to the discharge temperature of HP compressor  117  reaching a predetermined temperature limit. By providing the bleed air to LP turbine  122 , LP turbine  122  receives an increased amount of engine  100 &#39;s working fluid under pressure, from which LP turbine  122  may extract additional power for driving fan  112  than had the bleed air not been provided. Thus, some aspects of the present invention include not only preventing an over-temperature condition in compressor system  114 , but also include simultaneously increasing the amount of power provided to fan  112 , thereby increasing the thrust output of engine  100 . 
         [0019]    Referring to  FIG. 3 , some aspects of a non-limiting example of a two-spool gas turbine engine  200  in accordance with an embodiment of the present invention is schematically depicted. Engine  200  represents one of many possible configurations of engine  10  of  FIG. 1 . Engine  200  includes a turbofan  212 , a compressor system  214 , including an LP compressor  216  and an HP compressor  217 ; a combustion system  218 , and a turbine system  220 , including an HP turbine  221  and an LP turbine  222 . It will be noted that some reference numerals in  FIG. 3  correspond to the reference numerals for like parts of  FIG. 1 , with the addition of a numerical value of 200 to the reference numerals of  FIG. 1  to achieve the reference numerals in  FIG. 3 . For example, load absorber  12  of  FIG. 1  is represented by fan  212  of  FIG. 3 ; compressor stage  16  of  FIG. 1  is represented in the embodiment of  FIG. 3  as LP compressor  216 ; and turbine stage  22  of  FIG. 1  is represented as LP turbine  222  in  FIG. 3 . In one form, LP compressor  216  includes a plurality of compressor stages. In other embodiments, LP compressor  216  may include only a single compressor stage. In one form, LP turbine  222  includes a plurality of turbine stages. In other embodiments, LP turbine  222  may include only a single turbine stage. 
         [0020]    LP compressor  216  is in fluid communication with fan  212 . HP compressor  217  is in fluid communication with LP compressor  216 . Combustor  218  is in fluid communication with HP compressor  217 . HP turbine  221  is drivingly coupled to HP compressor  217  and in fluid communication with combustion system  218 . LP turbine  222  is drivingly coupled to LP compressor  216  and fan  212 , and is in fluid communication with HP turbine  221 . Engine  200  also includes bleed system  24 , which in the embodiment of  FIG. 3  is configured to bleed pressurized air from the discharge of LP compressor  216  and/or other LP compressor  216  compressor stages, and to deliver the bleed air to one or more turbine stages of LP turbine  222  in response to the discharge temperature of HP compressor  217  reaching a predetermined temperature limit. By providing the bleed air to LP turbine  222 , LP turbine  222  receives an increased amount of engine  200 &#39;s working fluid under pressure, from which LP turbine  222  may extract additional power for driving fan  212  than had the bleed air not been provided. Thus, some aspects of the present invention include not only preventing an over-temperature condition in compressor system  214 , but also include simultaneously increasing the amount of power provided to fan  212 , thereby increasing the thrust output of engine  200 . 
         [0021]    Referring to  FIG. 4 , some aspects of a non-limiting example of a two-spool gas turbine engine  300  in accordance with an embodiment of the present invention is schematically depicted. Engine  300  represents one of many possible configurations of engine  10  of  FIG. 1 . Engine  300  includes a turbofan  312 , a compressor  314 , including an intermediate compressor stage  316 ; a combustion system  318 , and a turbine system  320 , including an HP turbine  321  and an LP turbine  322 . It will be noted that some reference numerals in  FIG. 4  correspond to the reference numerals for like parts of  FIG. 1 , with the addition of a numerical value of 300 to the reference numerals of  FIG. 1  to achieve the reference numerals in  FIG. 4 . For example, load absorber  12  of  FIG. 1  is represented by fan  312  of  FIG. 4 ; compressor stage  16  of  FIG. 1  is represented in the embodiment of  FIG. 4  as intermediate compressor stage  316 ; and turbine stage  22  of  FIG. 1  is represented as LP turbine  322  in  FIG. 3 . Compressor  314  includes a plurality of compressor stages, only one of which is intermediate compressor stage  316 , and another of which is the final compressor stage (not shown). In one form, LP turbine  322  includes a plurality of turbine stages. In other embodiments, LP turbine  322  may include only a single turbine stage. 
         [0022]    Compressor  314  is in fluid communication with fan  312 . Combustor  318  is in fluid communication with compressor  314 . HP turbine  321  is drivingly coupled to compressor  314  and in fluid communication with combustion system  318 . LP turbine  322  is drivingly coupled to compressor  314  and fan  312 , and is in fluid communication with HP turbine  321 . Engine  300  also includes bleed system  24 , which in the embodiment of  FIG. 4  is configured to bleed pressurized air from intermediate compressor stage  316  (alone or in combination with other intermediate compressor stages of compressor  314 ), and to deliver the bleed air to one or more turbine stages of LP turbine  322  in response to the discharge temperature of compressor  314  reaching a predetermined temperature limit. By providing the bleed air to LP turbine  322 , LP turbine  322  receives an increased amount of engine  300 &#39;s working fluid under pressure, from which LP turbine  322  may extract additional power for driving fan  312  than had the bleed air not been provided. Thus, some aspects of the present invention include not only preventing an over-temperature condition in compressor system  314 , but also include simultaneously increasing the amount of power provided to fan  312 , thereby increasing the thrust output of engine  300 . 
         [0023]    Referring to  FIG. 5 , some aspects of a non-limiting example of bleed system  24  in accordance with an embodiment of the present invention is described. Bleed system  24  includes passages  28 , a valve  30  and a controller  32 . Passages  28  may take one or more of many forms, including pipes, tubing, internal passages inside other engine structures and the like. One passage  28  is coupled to and in fluid communication with intermediate compressor stage  16  at one end, and with valve  30  at the other end. Compressor stage  16  may be, for example and without limitation, the final compressor stage and/or one or more other compressor stages of IP compressor  116  ( FIG. 2 ); the final compressor stage and/or one or more other compressor stages of LP compressor  216  ( FIG. 3 ); or intermediate compressor stage  316  and/or one or more other compressor stages of compressor  314  ( FIG. 4 ). A passage  28  is also coupled to and in fluid communication with turbine stage  22  at one end and with valve  30  at the other end. Turbine stage  22  may be, for example and without limitation, the initial and/or one or more subsequent turbine stages of LP turbine  122  ( FIG. 2 ); the initial and/or subsequent turbine stages of LP turbine  222  ( FIG. 3 ); or the initial and/or one or more subsequent turbine stages of LP turbine  322  ( FIG. 4 ). The pressure at the compressor stage from which the is extracted is higher than the pressure at the turbine stage(s) where the air is delivered. For example, in one form, the total pressure at intermediate compressor stage  16  where the bleed air is extracted is greater than the total and/or static pressure at turbine stage  22  at the location where the bleed air is delivered via passage  28  to turbine stage  22 . In other embodiments, the static pressure at intermediate compressor stage  16  where the bleed air is extracted is greater than the total and/or static pressure at the location where the bleed air is delivered via passage  28  to turbine stage  22 . 
         [0024]    Valve  30  is configured to regulate the flow rate of the bleed air that is extracted from compressor stage  16  and provided to turbine stage  22  via passages  28 . In one form, valve  30  is configured to selectively prevent or allow the flow of bleed air from compressor stage  16 . In one form, valve  30  is configured to regulate the bleed air flow rate to a desired level. Valve  30  is controlled by controller  32 . Valve  30  may take any suitable form, and may be, for example and without limitation, a butterfly valve, a gate valve, a poppet valve or any other suitable valve type. Valve  30  is actuated by an actuation mechanism (not shown) under the direction of controller  32 . 
         [0025]    Controller  32  is communicatively coupled to valve  30  via a communications link  34 . Communications link  34  may take any suitable form, and may be, for example, a wired and/or wireless and/or optical link capable of transmitting control signals (and feedback signals, depending upon the configuration of valve  30 ) to (and from) valve  30 . In some embodiments, link  34  may also provide electrical power for actuating valve  30 . Controller  32  is configured to execute program instructions to control valve  30  to turn the bleed air on and off, and to regulate the bleed air flow rate to a desired level, based on compressor  14  discharge temperature, which is the temperature at the final compressor stage of compressor  14  (e.g., the final compressor stage of HP compressor  117 , HP compressor  217  or compressor  314 ). In one form, controller  32  is microprocessor based and the program instructions are in the form of software stored in a memory (not shown). However, it is alternatively contemplated that controller  32  and the program instructions may be in the form of any combination of software, firmware and hardware, including state machines, and may reflect the output of discreet devices and/or integrated circuits, which may be co-located at a particular location or distributed across more than one location, including any digital and/or analog devices configured to achieve the same or similar results as a processor-based controller executing software or firmware based instructions. In one form, controller  32  is a gas turbine engine controller, such as a full authority digital electronic control (FADEC) unit. In other embodiments, controller  32  may take any suitable form, and in some embodiments may be a dedicated controller for operating valve  30 . 
         [0026]    Controller  32  is configured to control valve  30  based on comparing compressor  14  discharge temperature with a predetermined temperature limit for compressor  14  discharge temperature. In various embodiments, the compressor  14  discharge temperature may be determined based on a measured compressor  14  discharge temperature or a calculated compressor  14  discharge temperature. For example, in one form, controller  32  is communicatively coupled to a compressor discharge temperature sensor  36 , which provides data reflective of a measured compressor  14  discharge temperature to controller  32 . During operation, controller  32  compares the compressor  14  discharge temperature based on sensor  36  to the predetermined temperature limit for purposes of controlling valve  30 , including commanding valve  30  to open to allow the flow of bleed air to turbine stage  22  and in some embodiments, regulating the bleed air flow rate to a desired level. 
         [0027]    In another embodiment, controller  32  is communicatively coupled to a pressure sensor  38  and a temperature sensor  40 . Sensors  38  and  40  may be, for example, located at the engine inlet or otherwise provide data pertaining to engine inlet conditions. In such embodiments, controller  32  determines a calculated compressor  14  discharge temperature, e.g., based on the output of sensors  38  and  40 , and on known characteristics of compressor  14 . During operation, controller  32  compares the calculated compressor  14  discharge temperature with the predetermined temperature limit for purposes of controlling valve  30 , including commanding valve  30  to open to allow the flow of bleed air to turbine stage  22  and in some embodiments, regulating the bleed air flow rate to a desired level. In still other embodiments, controller  32  may obtain compressor  14  discharge temperature data via other means, e.g., directly and/or indirectly measured and/or calculated, and may control valve  30  based on comparing the obtained compressor  14  discharge temperature data with the predetermined temperature limit. In some embodiments, the determination of compressor  14  discharge temperature may be made prior to starting engine  10 , e.g., calculated based on sensed or anticipated engine inlet conditions, whereas in other embodiments, the determination of compressor  14  discharge temperature may be made after starting engine  10 . 
         [0028]    The bleed air extracted from intermediate compressor stage  16  may be delivered to turbine stage  22  in via any convenient manner, which may vary, e.g., depending upon the amount of bleed air that is delivered to turbine stage  22 . For example, in some embodiments, the bleed air may be dumped into the core flowpath upstream of turbine stage  22 , e.g., via openings in one or more turbine system components (not shown). As another example, in some embodiments, e.g., where the bleed air flow rate is in the range of 5%-15% of core engine flow or greater, the bleed air may be delivered to turbine stage  22  via a mixer  42  that is configured to mix core gas flow with the bleed air flow for provision of the bleed air flow to turbine stage  22 . 
         [0029]    Referring to  FIG. 6 , in another example, the bleed air may be provided to turbine stage  22  via openings in one or more turbine vanes and/or struts  44  upstream of or within turbine stage  22 . For example, where a relatively low amount of bleed air is provided to turbine stage  22 , e.g., a bleed flow rate in the range of 1%-5% of core flow, the bleed air may be provided via existing cooling air discharge openings  46  in turbine vanes and/or struts  44 . Where higher bleed air flow rates are desired, e.g., a bleed flow rate in the range of 2%-10% of core flow, additional openings, e.g., trailing edge openings  48  may be added to turbine vanes and/or struts  44  to accommodate the additional bleed air flow. It will be understood that the bleed flow rates mentioned herein are exemplary only, and that bleed rates in particular embodiments may or may not be within the ranges mentioned herein. 
         [0030]    Embodiments of the present invention include a gas turbine engine, comprising: a fan; an intermediate pressure (IP) compressor in fluid communication with the fan; a high pressure (HP) compressor in fluid communication with the IP compressor; a combustor in fluid communication with the HP compressor; an HP turbine coupled to the HP compressor and in fluid communication with the combustor; an IP turbine coupled to the IP compressor and in fluid communication with the HP turbine; an LP turbine coupled to the fan and in fluid communication with the IP turbine; and a bleed system configured to bleed pressurized air from the IP compressor and deliver the bleed air to the LP turbine in response to a discharge temperature of the HP compressor reaching a predetermined temperature limit. 
         [0031]    In a refinement, the bleed system includes a valve configured to regulate a flow rate of the bleed air from the IP compressor. 
         [0032]    In another refinement, the valve is configured to selectively prevent a flow of the bleed air from the IP compressor. 
         [0033]    In yet another refinement, the gas turbine engine further comprises a controller configured to execute program instructions to control the valve to regulate a bleed air flow rate. 
         [0034]    In still another refinement, the controller is configured to control the valve based on comparing the discharge temperature of the HP compressor with the predetermined temperature limit. 
         [0035]    In yet still another refinement, the controller is configured to control the valve based on determining a calculated HP compressor discharge temperature based on engine inlet conditions, and comparing the calculated discharge temperature with the predetermined temperature limit. 
         [0036]    In a further refinement, a total pressure at an IP compressor bleed location from which the bleed air is extracted is higher than the total pressure at an injection location in the LP turbine where the bleed air is delivered to the LP turbine. 
         [0037]    In a yet further refinement, the LP turbine includes turbine vanes having air discharge openings, and wherein the bleed air is discharged through the air discharge openings. 
         [0038]    In a still further refinement, the gas turbine engine further comprises a mixer positioned at the LP turbine, wherein the mixer is adapted to receive the bleed air and mix the bleed air with core gas flow passing through the LP turbine. 
         [0039]    Embodiments of the present invention include a gas turbine engine, comprising: a compressor system having a plurality of compressor stages including an intermediate compressor stage and culminating in a final compressor stage; a combustor in fluid communication with the final compressor stage; a turbine system having a plurality of turbine stages including a low pressure turbine stage, wherein the low pressure turbine stage operates at a lower pressure than the intermediate compressor stage, and including an initial turbine stage in fluid communication with the combustor; and a bleed system configured to bleed pressurized air from the intermediate compressor stage and deliver the bleed air to the low pressure turbine stage in response to a discharge temperature of the compressor system reaching a predetermined temperature limit. 
         [0040]    In a refinement, the gas turbine engine is configured as a three-spool engine, wherein the compressor system includes an intermediate pressure (IP) compressor; wherein the intermediate compressor stage is part of the IP compressor; wherein the turbine system includes a low pressure (LP) turbine; and wherein the low pressure turbine stage is part of the LP turbine. 
         [0041]    In another refinement, the gas turbine engine is configured as a two-spool engine, wherein the compressor system includes a high pressure (HP) compressor; wherein the intermediate compressor stage is part of the HP compressor; wherein the turbine system includes a low pressure (LP) turbine; and wherein the low pressure turbine stage is part of the LP turbine. 
         [0042]    In yet another refinement, the bleed system includes ducting configured to deliver the bleed air to the low pressure turbine stage. 
         [0043]    In still another refinement, the bleed system includes a valve configured to regulate a flow rate of the bleed air from the IP compressor. 
         [0044]    In yet still another refinement, the gas turbine engine further comprises a controller configured to execute program instructions to control the valve to regulate a bleed air flow rate. 
         [0045]    In a further refinement, the controller is configured to control the valve based on comparing the discharge temperature of the compressor system with the predetermined temperature limit. 
         [0046]    In a yet further refinement, the controller is configured to control the valve based on determining a calculated compressor system discharge temperature based on engine inlet conditions, and comparing the calculated discharge temperature with the predetermined temperature limit. 
         [0047]    Embodiments of the present invention include a method for operating a gas turbine engine, comprising: determining a compressor discharge temperature; comparing the compressor discharge temperature with a compressor discharge temperature limit; bleeding air from an intermediate compressor stage in response to the comparison; and delivering the bleed air to a turbine stage having a lower operating pressure than the intermediate compressor stage. 
         [0048]    In a refinement, the determination of the compressor discharge temperature includes measuring the compressor discharge temperature. 
         [0049]    In another refinement, the determination of the compressor discharge temperature includes calculating the compressor discharge temperature based on engine inlet conditions. 
         [0050]    In yet another refinement, the gas turbine engine is configured as a three-spool engine having a fan, and intermediate pressure (IP) compressor, a high pressure (HP) compressor, an HP turbine coupled to the HP compressor, an IP turbine coupled to the IP compressor, and an LP turbine coupled to the fan; wherein the intermediate compressor stage is part of the IP compressor; and wherein the turbine stage is part of the LP turbine. 
         [0051]    In still another refinement, the gas turbine engine is configured as a two-spool engine having a propulsor, a high pressure (HP) compressor, an HP turbine coupled to the HP compressor, and a low pressure (LP) turbine coupled to the propulsor; wherein the intermediate compressor stage is part of the HP compressor; and wherein the turbine stage is part of the LP turbine. 
         [0052]    In a further refinement, the gas turbine engine is configured as a two-spool engine having a low pressure (LP) compressor, a high pressure (HP) compressor, an HP turbine coupled to the HP compressor, and an LP turbine coupled to the LP compressor; wherein the intermediate compressor stage is part of the LP compressor; and wherein the turbine stage is part of the LP turbine. 
         [0053]    While the invention has been described in connection with what is presently considered to be the most practical and preferred embodiment, it is to be understood that the invention is not to be limited to the disclosed embodiment(s), but on the contrary, is intended to cover various modifications and equivalent arrangements included within the spirit and scope of the appended claims, which scope is to be accorded the broadest interpretation so as to encompass all such modifications and equivalent structures as permitted under the law. Furthermore it should be understood that while the use of the word preferable, preferably, or preferred in the description above indicates that feature so described may be more desirable, it nonetheless may not be necessary and any embodiment lacking the same may be contemplated as within the scope of the invention, that scope being defined by the claims that follow. In reading the claims it is intended that when words such as “a,” “an,” “at least one” and “at least a portion” are used, there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. Further, when the language “at least a portion” and/or “a portion” is used the item may include a portion and/or the entire item unless specifically stated to the contrary.