Abstract:
One embodiment according to the present invention is a unique system for gas turbine engine control. Other embodiments include unique apparatuses, systems, devices, and methods relating to gas turbine engines. Further embodiments, forms, objects, features, advantages, aspects, and benefits of the present invention shall become apparent from the following description and drawings.

Description:
CROSS REFERENCE TO RELATED APPLICATIONS 
       [0001]    The present application claims the benefit of U.S. Provisional Patent Application No. 61/204,039, filed Dec. 31, 2008, and is incorporated herein by reference. 
     
    
     FIELD OF THE INVENTION 
       [0002]    The present invention relates generally to gas turbine engines and more particularly to systems, apparatuses, and methods of gas turbine engine control. 
       BACKGROUND 
       [0003]    Gas turbine engines are an efficient source of energy and have proven useful to propel and power aircraft, for electricity generation, as well as for other uses. One aspect of gas turbine engines is that they include systems, subsystems, and elements, such as, mechanical, electrical, and electro-mechanical systems, subsystems, and elements that must be controlled during operation. Numerous gas turbine control issues exist. During some gas turbine engine operating conditions there is a delay or lag in mechanical response to commanded engine operation. Such delays or lags can result in transient operating conditions which briefly exceed prescribed engine operation limits or desired operation ranges, but which do not require operator intervention because corrective engine operation has already been commanded. There is a concern that operators may take inappropriate or unsafe action not understanding that corrective action is underway. Proposed approaches involve limiting or clipping a signal indicating that a prescribed engine operation limits or desired operation ranges has been exceeded. Such approaches are unsatisfactory because they mask true operation issues requiring operator intervention. Thus, there is a need for systems, apparatuses, and methods of gas turbine engine control disclosed herein. 
       SUMMARY 
       [0004]    One embodiment according to the present invention is a unique system for gas turbine engine control. Other embodiments include unique apparatuses, systems, devices, software, hardware, methods, and combinations of these and/or other aspects relating to gas turbine engines. Further embodiments, forms, objects, features, advantages, aspects, and benefits of the present invention shall become apparent from the following description and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]      FIG. 1  is a perspective view of an aircraft propelled by two gas turbine engines. 
           [0006]      FIG. 2  is an illustrative representation of a gas turbine engine. 
           [0007]      FIG. 3  is a schematic representation of a gas turbine engine subsystem and a control subsystem. 
           [0008]      FIG. 4  is a flow diagram representation of steps in a control process. 
           [0009]      FIG. 5  is a flow diagram representation of steps in a control process. 
           [0010]      FIG. 6  is a graph of signal information as a function of time. 
       
    
    
     DETAILED DESCRIPTION 
       [0011]    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 nevertheless be understood that no limitation of the scope of the invention is thereby intended, such alterations and further modifications in the illustrated device, and such further applications of the principles of the invention as illustrated therein being contemplated as would normally occur to one skilled in the art to which the invention relates. 
         [0012]    With reference to  FIG. 1  there is shown airplane  100  including gas turbine engine engines  110  and  120  which operate to propel airplane  100 . Airplane  100  is one example of a use to which gas turbine engines can be put. There are a variety of additional applications for gas turbine engines, including, for example, electricity generation, pumping sets for gas and oil transmission lines, land and naval propulsion, and still other applications. It should be appreciated that systems, apparatuses, and methods and other embodiments according to the present invention can be used in connection with the gamut of gas turbine engine applications including aircraft applications which include helicopters, airplanes, missiles, unmanned space devices and any other substantially similar devices. Thus, while the following description is in the context of a gas turbine engine suitable for aircraft propulsion, the invention broadly applies to the aforementioned applications and others. 
         [0013]    With reference to  FIG. 2  there is illustrated a schematic view of a gas turbine engine  200  which includes a compression section  215 , a combustor section  223 , and a turbine section  224  which are integrated together to produce an aircraft flight propulsion engine. In one form, the compression system  215  includes a fan section  221  and a compressor section  222 . This type of gas turbine engine is generally referred to as a turbo-fan. One alternate form of a gas turbine engine includes a compressor, a combustor, and a turbine that have been integrated together to produce an aircraft flight propulsion engine without-the fan section. 
         [0014]    The compressor section  222  includes a rotor  219  having a plurality of compressor blades  228  coupled thereto. The rotor  219  is affixed to a shaft  225  that is rotatable within the gas turbine engine  220 . A plurality of compressor vanes  229  are positioned within the compressor section  222  to direct the fluid flow relative to blades  228 . Turbine section  224  includes a plurality of turbine blades  230  that are coupled to a rotor disk  231 . The rotor disk  231  is affixed to the shaft  225 , which is rotatable within the gas turbine engine  220 . Energy extracted in the turbine section  224  from the hot gas exiting the combustor section  223  is transmitted through shaft  225  to drive the compressor section  222 . Further, a plurality of turbine vanes  232  are positioned within the turbine section  224  to direct the hot gaseous flow stream exiting the combustor section  223 . 
         [0015]    The turbine section  224  provides power to a fan shaft  226 , which drives the fan section  221 . The fan section  221  includes a fan  218  having a plurality of fan blades  233 . Air enters the gas turbine engine  220  in the direction of arrows A and passes through the fan section  221  into the compressor section  222  and a bypass duct  227 . The term airfoil refers to fan blades, fan vanes, compressor blades, turbine blades, compressor vanes, and turbine vanes unless specifically stated otherwise. Further details related to the principles and components of a conventional gas turbine engine will not be described herein as they are known to one of ordinary skill in the art. 
         [0016]    It is important to appreciate that there are a multitude of ways in which the gas turbine engine components can be linked together. For example, additional compressors and turbines could be added with intercoolers connecting between the compressors and reheat combustion chambers could be added between the turbines. A wide variety of additional configurations and variations are also possible as would occur to skilled artisans. 
         [0017]    With reference to  FIG. 3  there is illustrated system  300  which includes control subsystem  310  and engine subsystem  320  and could also include a variety of additional subsystems as would occur to skilled artisans. Control subsystem  310  includes a control unit  310  which can include a full authority digital electronic control (“FADEC”)  311  or any other logic, program, software, hardware, or combination of these and/or other elements operable to receive information and to output control signals. Subsystem  310  is interconnected to engine subsystem  320  with information interconnection  330  and control interconnection  340 . 
         [0018]    As indicated by ellipsis N there could also be a greater number of interconnections (or fewer in the case of a shared single interconnection) between subsystems  310  and  320 . Furthermore, a wide variety of interconnections are contemplated, including, wire, wireless, mechanical, electro-mechanical, electro-magnetic, optical, and combinations of these and other types of interconnections. As illustrated, any or all of the elements of subsystems  310  and/or  320  could be interconnected via one or more interconnections. Additional sub-systems and elements not illustrated could also be similarly interconnected with the illustrated elements and with one another. 
         [0019]    Engine subsystem  320  receives information or input from one or more interconnections, such as those just described or others. There can be interconnection between some or all of the various elements and/or subsystems of engine subsystem  320 , for example, compressor(s)  321 , combustor(s)  322 , turbine(s)  323 , and/or a load, such as, turbofan  324 . 
         [0020]    Subsystem  320  also includes shafts  350  and  360  which are driven by turbines and whose output connects to other elements of engine subsystem  320 . For example, in the case of a turbo fan, outer shaft  350  outputs to turbofan  324  and inner shaft  360  outputs to compressor  321 , a variety of other interconnections, such as those of the alternatives mentioned elsewhere herein and others, are within the scope of the present invention. The foregoing and other elements, and others, can be housed in a housing  325 . 
         [0021]    With reference to  FIG. 4  there is shown flow diagram  400  which begins at start operation  410  and proceeds to operation  420  where a signal is received from a sensor, for example, a signal of information relating to an engine temperature, such as inter turbine temperature (“ITT”) or another turbine temperature, could be received during an engine operational state where a high pressure output increase is occurring or is commanded. From operation  420  diagram  400  proceeds to operation  430  where a controller, such as a FADEC, calculates a reduction coefficient intended to reduce engine speed at N 1  (or simply calculates a reduced engine speed at N 1 ). From operation  430  diagram  400  can proceed to operation  440  where the FADEC sends a signal to a fuel pump metering unit or FPMU to reduce fuel flow. From operation  430 , diagram  400  can also proceed to operation  450  where a FADEC or other control means calculates a predictive reduction amount, X, such as an amount to reduce ITT or another temperature, as an objective for the reduction of engine speed at N 1 . From operation  450 , diagram  400  proceeds to operation  460  where the calculated predictive reduction amount, X, is applied to reduce ITT, or another operational aspect, and is phased out over a time period, such as, 1.5 seconds. From either operations  440  and  460 , diagram  400  proceeds to state  490  where it either ends or is reset. 
         [0022]    With reference to  FIG. 5  there is shown flow diagram  500  which begins at the START operation and proceeds to operation  510  where information from a sensor (such as a physical sensor, a virtual sensor, or other type of sensor) is received. From operation  510  diagram  500  proceeds to conditional  520  where the identity of a state of a gas turbine engine is monitored. From operation  520  diagram  500  proceeds to conditional  530  where the monitored state is analyzed to determine whether it is predictive or precedes a state indicating a need for engine control. From operation  530  diagram  500  proceeds to operation  540  where control is initiated, for example, by generating and/or outputting a control signal. Also at operation  540 , an adjustment to information received from a sensor is made (such as a decrease in received temperature information). From operation  540  diagram  500  proceeds to a state where it is reset or ends. Let it be understood that a variety of alternatives to the foregoing methods are contemplated within the scope of the invention, including methods where the order of events, operations, and/or conditionals is different, such as, by reordering or substantial overlap or concurrent operation. 
         [0023]    With reference to  FIG. 6  there is shown a graph  600  with time on its horizontal axis and inter-turbine temperature (in Celcius) on its vertical axis. Graph  600  shows the relationship of inter-turbine temperature (“ITT”) as it varies with time for a number of different signals  610 ,  620 ,  630 ,  640 ,  650 ,  660 ,  670  and  680 . Each of these signals exhibits a spike in ITT before a period of relatively lower and relatively stable ITT. In some cases, spikes such as those of graph  600  are attributable, at lease in part, to mechanical delay in response to engine control signals. The foregoing apparatuses, systems, methods, and other embodiments according to the present invention can be operable to predict such spikes and predict for, correct for, and/or smooth them, among other capabilities. 
         [0024]    Exemplary embodiments include control systems, methods, and controllers for gas turbine engines which anticipate changes in turbine temperature based on commanded engine operation and apply predictive adjustments to sensed turbine temperature information. An exemplary adjustment to sensed turbine temperature information includes adjusting for a transient turbine temperature increase, for example, a turbine temperature spike. During gas turbine engine operation an increased load may be applied to the engine. In some embodiments an increased load is applied to an engine spool, for example, the high pressure spool, to drive aircraft systems or other auxiliary systems. In some embodiments an increased load is applied by extracting bleed air for anti-ice operation or for other operations. Some embodiments include multiple increased loads applied to a gas turbine engine. If commanded engine speed remains unchanged the overall load on the engine increases and turbine temperature can increase above a defined level such as an engine yellow band or a red line level. To avoid operating the engine above the defined turbine temperature level, an engine controller calculates a reduction in engine output to offset the increased load applied to the engine and commands a reduced level of engine output. In some embodiments the reduction in engine output is a reduction in engine thrust. In some embodiments the reduction in engine output is a reduction in rotational engine speed. In some embodiments the reduction in engine output is a reduction in rotational speed of the high pressure spool. In some embodiments the reduction in engine output is a reduction in fuel provided to the engine. 
         [0025]    In exemplary embodiments there is a lag in the mechanical response of the engine to the commanded output reduction which results in a transient turbine temperature increase, for example, a turbine temperature spike. A sensor outputs turbine temperature information which reflects the transient increase in turbine temperature. The output turbine temperature information is provided to the engine controller. The engine controller adjusts the turbine temperature information provided by the sensor to offset, counteract or reduce the transient temperature increase. Adjustment of the turbine temperature information provided by the sensor may be accomplished in a number of manners. 
         [0026]    In some embodiments the engine controller calculates an adjustment to the sensed turbine temperature. In some embodiments the engine controller obtains an adjustment to the sensed turbine temperature from a look up table based upon a number of variables, for example, throttle position, altitude, and/or throttle change. In some embodiments the engine controller predicts a post-transient decrease in turbine temperature based upon a commanded decrease in engine output and reduces the magnitude of the turbine temperature information provided by the sensor based upon the magnitude of the predicted decrease. In some embodiments the reduction is applied for a specified duration. In some embodiments the reduction is applied for a period based upon a lag in mechanical response of the engine to commanded reduction in engine speed. In some embodiments the magnitude of the reduction is limited. Some embodiments limit the magnitude adjustment based upon values in a look up table and may account for multiple variables including altitude and engine throttle selected by a pilot. In some embodiments the magnitude of the reduction is scaled over a time period to approximate an inverse of an expected transient temperature increase. In some embodiments the reduction is offset in time to account for a lag response time. 
         [0027]    Some exemplary embodiments compensate for transient turbine temperature increases attributable to multiple loads imposed on the engine, for example, multiple bleeds may contribute to an overall observed temperature spike. One such embodiment calculates an adjusted turbine temperature by measuring a turbine temperature and subtracting a first bleed reduction value and subtracting a second bleed reduction value. The first bleed reduction value is offset by a first time constant which accounts for a delay in the effect created by the first bleed, and is also scaled as a function of time to produce an offset which is or approximates the inverse of the portion of the transient temperature spike attributable to the first bleed. The second bleed reduction value is offset by a second time constant which accounts for a delay in the effect created by the second bleed, and is also scaled as a function of time to produce an offset which is or approximates the inverse of the portion of the transient temperature spike attributable to the second bleed. By subtracting both the first bleed reduction value and the second bleed reduction value from a measured turbine temperature, an adjusted turbine temperature is provided which reduces the transient temperature spikes associated with two bleeds. Some embodiments may include additional bleed reduction values for additional bleeds or other loads. 
         [0028]    Some exemplary embodiments compensate for transient temperature spikes while simultaneously accounting for increase throttle commands from a pilot. One such embodiment calculates a transient temperature increase from a measured increase in turbine temperature and subtracts a temperature increase expected from a commanded throttle increase. The temperature increase expected from a commanded throttle increase is calculated by scaling the change in commanded throttle by a term that accounts for the starting throttle position, the aircraft altitude, and/or the magnitude of the throttle change. The scalar may be provided from a look up table is provided which specifies the scalar in one or more dimensions, for example, throttle position, change in throttle position, change throttle position normalized by the change in turbine temperature, and/or altitude. Another such embodiment determines a transient temperature increase from a look up table and subtracts a temperature increase expected from a commanded throttle increase. The temperature increase expected from a commanded throttle increase is calculated by scaling the change in commanded throttle by a term that accounts for the starting throttle position, the aircraft altitude, and/or the magnitude of the throttle change. The scalar may be provided from a look up table which specifies the scalar in one or more dimensions, for example, throttle position, change in throttle position, change throttle position normalized by the change in turbine temperature, and/or altitude. 
         [0029]    Some exemplary embodiments impose a limit on the magnitude of the adjustment or reduction of measured turbine temperature. In one such embodiment, a first adjustment value is calculated by subtracting a current turbine temperature value from an initial turbine temperature value. The first adjustment value may be scaled to account for a partial contribution to an overall transient temperature increase in the case of multiple loads. A term to account for a change in pilot commanded throttle change may also be subtracted. A second adjustment value is determined from a look up table which specifies subtracting a current turbine temperature value from an initial turbine temperature value. The second adjustment value may be scaled to account for a partial contribution to an overall transient temperature spike in the case of multiple loads. A term to account for a change in pilot commanded throttle may also be subtracted. A limit on the magnitude of the adjustment of measured turbine temperature is calculated by taking the minimum of the first adjustment value and the second adjustment value. A further limit may be imposed to set the adjustment to zero if the result of this calculation is a negative value. 
         [0030]    While the invention has been illustrated and described in detail in the drawings and foregoing description, the same is to be considered as illustrative and not restrictive in character, it being understood that only the preferred embodiments have been shown and described and that all changes and modifications that come within the spirit of the inventions are desired to be protected. It should be understood that while the use of words such as preferable, preferably, preferred or more preferred utilized in the description above indicate that the feature so described may be more desirable, it nonetheless may not be necessary and embodiments lacking the same may be contemplated as within the scope of the invention, the 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,” or “at least one portion” are used there is no intention to limit the claim to only one item unless specifically stated to the contrary in the claim. When the language “at least a portion” and/or “a portion” is used the item can include a portion and/or the entire item unless specifically stated to the contrary.