JP7249096B2 - turbo cooling vane of gas turbine engine - Google Patents

Description

[0001]
CROSS-REFERENCE TO RELATED APPLICATIONS This application claims priority from U.S. Provisional Application No. 62/112,263, filed February 5, 2015. , 403. This application also claims priority to US Provisional Patent Application No. 62/201,031, filed Aug. 4, 2015. Each of these applications is incorporated herein by reference in its entirety.

[0002]
This invention relates generally to power systems including the power generation capabilities of gas turbine engines, and more particularly to providing an alternative source of cooling air for components of gas turbine engines.

[0003]
Gas turbine engines are widely understood to be utilized in conjunction with electrical generators that convert mechanical shaft power into electrical power. Referring initially to FIG. 1A, a schematic diagram of a simple cycle gas turbine engine 100 commonly utilized in power generation power plants is shown. Gas turbine engine 100 includes a compressor 102 connected to turbine 104 by shaft 106 . Air from compressor 102 is directed to one or more combustors 108 where fuel 110 is added to the air. The fuel and air mixture is ignited to form hot combustion gases that drive the turbine 104 which in turn drives the compressor 102 . Shaft 106 is also connected to a generator 112 that produces electrical power 114 . FIG. 1B shows gas turbine performance as a function of specific power output, corresponding to thermal efficiency for both simple cycle efficiency and power output for various gas turbine pressure ratios and firing temperatures. As will be appreciated by those skilled in the art, the combustion temperature of a gas turbine engine regulates and limits the operation of the overall engine, and the pressure ratio is directly proportional to the efficiency of the gas turbine. For a combined cycle gas turbine, plant efficiency is directly proportional to combustion temperature, as shown in FIG. 2B. That is, as the combustion temperature increases, the power output of the simple cycle gas turbine increases, assuming that the mass flow rate remains constant while operating in the combined cycle, increasing the efficiency of the same gas turbine.

[0004]
Gas turbine counterparty brand manufacturers generally improve combustion by improving materials and coating techniques in the turbine section to allow higher temperature gases to pass through the turbine while maintaining the capabilities of the turbine components. raising the temperature.

[0005]
Referring now to FIG. 2A, a schematic diagram of a combined cycle power plant 200 is shown comprising a compressor 202 connected by a shaft 206 to a turbine 204 . Air from compressor 202 is directed to one or more combustors 208 where fuel 210 is added to the air from compressor 202 . The fuel and air mixture is ignited to form hot combustion gases that power turbine 204 and drive compressor 202 . Shaft 206 is also connected to a generator 212 that produces electrical power 214 . Combined cycle power plant 200 also includes a heat recovery steam generator or HRSG, 216 , which receives hot exhaust from turbine 204 and heats a water source to generate steam 218 . Steam turbine 220 is powered by steam from HRSG 216 and steam turbine 220 drives a second generator 222 to generate additional electrical power 224 . FIG. 2B shows gas turbine performance as a function of combustion temperature, corresponding to efficiency in terms of both combined cycle efficiency and power output. Figures 1B and 2B are similar to those disclosed in GE Gas Turbine Performance Characteristics (GER3567) and are included herein for reference purposes.

[0006]
As will be appreciated by those skilled in the art, combustion temperature is defined as the temperature of the combustion gases immediately downstream of the first stage turbine nozzle. Due to the different terminology used in the field of gas turbine engines, the first stage turbine nozzle may also be referred to as the first stage turbine vane. Referring to FIG. 3, a cross-section of a portion of a gas turbine engine is shown showing standard temperature parameters utilized in the gas turbine industry. FIG. 3 is also similar to the diagram disclosed in the GE Gas Turbine Performance Characteristics (GER3567) referenced above. As shown in FIG. 3, turbine inlet temperature (T _A ) is measured upstream of first stage turbine nozzle 300, indicated by plane AA. The combustion temperature of the engine (T _B ) is measured just after the first stage turbine nozzle, indicated by plane BB.

[0007]
As noted above, turbine inlet temperature and turbine firing temperature are important considerations due to the operation of gas turbine engines. These temperature readings are obtained upstream and downstream of the first stage turbine nozzle, respectively. As such, because gas turbine engine control is based on these temperatures, it is important to maintain turbine nozzle metal temperatures within acceptable material operating limits.

[0008]
The high operating temperatures of turbine nozzles require active cooling of the turbine nozzles to maintain metal temperatures at acceptable levels. A cooling fluid, such as compressed air, bypasses the turbine cooling and leakage air (TCLA), or combustion process, and is provided to the turbine nozzle as part of the overall compressed air utilized for cooling. TCLA is typically obtained from multiple locations within the compressor, including the discharge plenum of the gas turbine engine, in amounts required to cool the turbine components which vary by component and engine type. However, for the General Electric Frame 7FA engine, approximately 20% of the compressed air produced by the engine compressor is utilized as TCLA. That is, utilizing 20% of the compressed air for cooling means that this air cannot pass through the combustion system or be fired through the turbine, thereby being converted into engine lost energy and means that it contributes to the poor thermal efficiency of For example, the gas turbine engine mentioned above has a thermal efficiency of approximately 37 percent.

[0009]
FIG. 4, which is similar to that disclosed in GE Gas Turbine Performance Characteristics (GER3567), shows a typical cooling scheme for a first stage turbine nozzle 400. As shown in FIG. In such cooling schemes, compressed air is supplied to the internal passages of the turbine vanes and is often directed to multiple passages within the nozzle, some of which may be serpentine. Air for cooling the first stage turbine nozzle is typically generated by the compressor and is obtained from the compressor discharge plenum and is therefore at the outlet pressure and temperature of the engine compressor. This first stage nozzle, where the hottest gases from the combustor are seen, is also fed by a source of maximum pressure cooling air from the compressor discharge plenum (CDP). That is, the gas path pressure is only a few pounds per square inch (psi) less than that of the combustor. Therefore, as will be appreciated by those skilled in the art, the pressure of the cooling air supplied to the leading edge 402 of the first stage nozzle 400 is just high enough to force the air out of the series of holes in the airfoil. be. Although the spacing and orientation of the cooling holes can vary, one such common type has holes in the leading edge 402 of the nozzle 400, also referred to as a showerhead pattern. Additionally, obtaining air from the engine compressor to cool the turbine components reduces power output from the engine, thereby reducing the amount of machine work that the turbine can produce.

[0010]
Referring now to FIG. 5, a cross-sectional view of a portion of a gas turbine engine with a prior art cooling scheme is shown. Gas turbine engine 500 includes a compressor 502 that channels compressed air to a discharge plenum 504 . Most of the air from compressor 502 passes through combustor case 508, end cap 510, combustion liner 512, swirler assembly 514, transition piece 516, and transition piece 516 to a portion of the turbine frame, here first stage vane outer ring 520. It passes through one or more combustors 506 that have brackets 518 that hold them in place. Air is received within combustor 506 and mixed with fuel from one or more fuel nozzles 522 to generate hot combustion gases that pass through transition piece 516 to the turbine. In this embodiment, the first stage vane outer ring 520 is secured to a compressor discharge plenum (CDP) case 524 .

[0011]
Air is maintained in the compressor discharge plenum by a seal 526 between the rotor 528 and the inner casing 530 so that most of the air is directed to the combustor 506 or TCLA. The inner casing 530 has a mechanical interface 532 with a first stage turbine nozzle 531 that provides the necessary structural axial and torsional support. Inner casing 530 is generally supported within compressor discharge plenum case 524 by ID struts 534 located between adjacent combustors 506 . Rotor 528 has bearings 536 that connect rotor 528 to the casing through struts 534 .

[0012]
Cooling air 541 is supplied to the outer diameter of the first turbine nozzle 531 as the first vane outer ring feeds the vanes 531 with compressed air from the compressor discharge plenum 504 , the first outer vane ring 520 and the compressor discharge. It passes between the plenum case 524 and flows into the holes on the first vane outer ring 543 . In this embodiment of the invention, the compressed air from compressor discharge plenum 504 is approximately 750°F (approximately 399°C) at ISO conditions and base load. Similarly, the inner diameter of the first stage nozzle 542 is fed with turbine cooling and leakage air (TCLA) 552 from the compressor discharge plenum 504 . Both first stage nozzle cooling air 541 and 552 flow through the internal passages 531 of the vanes to provide the necessary cooling to the first stage nozzle 542 as disclosed in FIG. This TCLA eventually joins the hot combustion gases passing between the first stage nozzles 542 and acts as a coolant to reduce the temperature of the hot gases until the first stage blades 511 are exposed. In subsequent nozzle and rotor stages, the second stage nozzles are sealed to the rotor by a second stage inner support ring 554 and similarly in the third stage by a third stage inner support ring 553 .

[0013]
The following description relates to a General Electric Frame 7FA gas turbine engine at ISO conditions and base load and is presented for illustrative purposes only as an acceptable engine in which the present invention may be utilized and the scope of the invention as set forth below. is not intended to limit Most of the compressed air (approximately 80%) from the compressor passes through the combustion system where fuel is added and the mixture is ignited, raising the temperature of the hot combustion gases to approximately 2700°F (approximately 1482°C). . A pressure drop of typically 2 to 3 pounds per square inch (907 to 1361 g) (psi) occurs as the compressed air passes through the combustor. Therefore, due to this mechanism, there is very little pressure limit for cooling the nozzle, especially its leading edge. For Class F gas turbine engines, typically about 10% of the cooling air is diverted from the combustion process and utilized to cool the vanes. For example, for a 7FA engine, compressor discharge air at about 750°F (about 399°C) and 220 psi is utilized to cool the first stage nozzle. During the cooling process, this air increases in temperature to about 250°F (about 121°C) and is then discharged into the gas flow path thereby increasing the temperature coming from the combustion process (~2700°F (1482°C)). Dilute the gas in C)) to generate combustion temperatures. A typical combustion temperature for a 7FA engine is approximately 2450°F (approximately 1343°C) (as obtained at plane BB in FIG. 3) and approximately 2700°F (approximately 1482°C) from the combustion process. C) 900 lb/sec of hot combustion gases at temperature and 100 lb/sec of air at about 1000°F (about 538°C) from the cooling air for the nozzle. This therefore produces a combustion temperature of 2540°F (1393°C) at plane BB [(2700*900+100*1000)/1000 = 2540°F (1393°C)]. The basis for the higher temperature in the calculation (2540°F (1393°C) > 2450°F (1343°C)) is also to mix and reduce the combustion dilution and the actual temperature entering from the combustor. , the presence of cooling air in the plane BB thereby reducing the temperature. To estimate the effective combustion diluent and leakage air at compressor exit temperature (approximately 750°F (approximately 399°C)), (2700*900+100*1000+Flow*750)/(1000+Flow) = 2450, which gives Solving gives Flow=5. Thus, if the compressor inlet flow rate is about 1005 lb/sec, 900 lb/sec goes to the combustion process, about 5 lb/sec leaks to dilute the combustion process, and 100 lb/sec goes to the first stage nozzle to cool. These figures show that in a gas turbine compressor about 10% of the 1005 lb/sec going to the turbine inlet is removed before it exits the combustor to cool the rotating and then stationary parts of the turbine. does not reflect the fact that Therefore, for the example above, all of the flow figures are reduced by 10% or the combustor flow is about 810 lb/sec, the first stage nozzle flow is about 90 lb/sec, and the combustor flow is about 90 lb/sec. The dilution and leak rate is 4.3 lb/sec. As those skilled in the art will appreciate, these numbers are approximations, but if leak and cooling air are mixed at plane BB, it will result in a mixing temperature (firing temperature) of 2450°F (1343°C). .

[0014]
The industry standard for judging the cooling benefits achieved through cooling air is its cooling effectiveness. Cooling effectiveness is understood to be the ratio of the difference between the hot combustion gas temperature of the turbine nozzle and the mean metal temperature divided by the difference between the hot combustion gas and cooling air temperature. As an example, the cooling effect of the first stage turbine vanes of the 7FA engine described above is about 0.59 (hot combustion gases (~2700 (1482°C)) and average metal temperature (~1550 (843°C)) is the ratio of the temperature difference between the hot combustion gases divided by the difference between the hot combustion gas and cooling air temperatures (~750°F (399°C)).

[0015]
Cooling the hottest components, typically the first stage nozzle and first stage blades, is a costly technology for all original equipment manufacturers (OEMs) of gas turbine engines. For example, over the past two years, large frame gas turbine engines have been improved, but thermal efficiency improvements have only increased from about 33% to 37%.

[0016]
SUMMARY OF THE INVENTION The present invention presents several embodiments for improving the cooling efficiency of gas turbine components, including first stage turbine nozzles.

[0017]
In an embodiment of the invention, a system for directing cooling air to turbine vanes comprising an auxiliary source of compressed air having a heated engine, an auxiliary compressor, and a recuperator for supplying the heated auxiliary compressed air. and methods are presented. Heated auxiliary compressed air is supplied to the plurality of turbine vanes through conduits such that the auxiliary supply of compressed air provides a dedicated supply of cooling air for cooling the turbine vanes.

[0018]
In an alternative embodiment of the invention, a system and method are presented for selectively supplying cooling air to turbine vanes. A plurality of air-cooled turbine vanes, an auxiliary source of compressed air with a heated engine, an auxiliary compressor, and a recuperator are provided. Auxiliary compressed air is supplied to the plurality of turbine vanes through conduits through which the air is selectively directed to cool the turbine vanes. If a supplemental source of compressed air is not utilized, cooling air for the turbine vanes is supplied from the gas turbine engine compressor.

[0019]
In one embodiment of the present invention, at least a portion of the required turbine cooling and leakage air (TCLA) is supplied by an auxiliary source of compressed air having a lower temperature than prior art cooling designs, thereby reducing the required TCLA. Reduce volume and improve overall efficiency.

[0020]
In yet another embodiment of the present invention, a system and method for supplying cooling air and selecting paths for turbine vanes is disclosed. Cooling air is generated by an auxiliary compressor and passed through the leading edge region of the turbine vanes, with a portion of the air being supplied to the leading edge and then directed to cool another portion of the turbine nozzle.

[0021]
In another embodiment of the present invention, a system and method for supplying cooling air and selecting paths for turbine vanes is disclosed. Cooling air is generated by an auxiliary compressor and the distribution of the cooling air is varied to the turbine nozzles according to predetermined control parameters.

[0022]
Additional advantages and features of the invention will be set forth in part in the description that follows, and in part will be apparent to those skilled in the art from consideration of the following, or may be learned from practice of the invention. The invention will now be described with particular reference to the accompanying drawings. Although a first stage nozzle is utilized as an example of this embodiment, it is contemplated that the techniques outlined in this invention can be applied to other components within the turbine section.

[0023]
The present invention will be described in detail below with reference to the accompanying drawings.

[0024]
1 shows a schematic diagram of a simple cycle gas turbine engine; FIG. 1B shows the relationship between combustion temperature and thermal efficiency and power output for the engine of FIG. 1A; 1 shows a schematic diagram of a combined cycle gas turbine engine; FIG. 2B shows the relationship between combustion temperature and thermal efficiency and power output for the engine of FIG. 2A; 1 illustrates a partial cross-sectional view of a gas turbine engine showing axial locations at which standard temperatures are measured; FIG. 1 is a perspective view of a typical gas turbine nozzle showing its cooling pattern; FIG. 1 is a partial cross-sectional view of a gas turbine engine providing a method of directing cooling air to first stage turbine vanes according to the prior art; FIG. 1 is a partial cross-sectional view of a gas turbine engine providing a method of directing cooling air to first stage turbine vanes in accordance with one embodiment of the present invention; FIG. 1 is a schematic diagram of an auxiliary compressed air source, according to one embodiment of the present invention; FIG. 4 is a partial cross-sectional view of a gas turbine engine providing a method of selectively directing cooling air to first stage turbine vanes in accordance with an alternative embodiment of the present invention; FIG. 4 is a partial cross-sectional view of a gas turbine engine providing an alternative method of directing dedicated cooling air for cooling first stage turbine vanes, in accordance with an alternative embodiment of the present invention; FIG.

[0025]
The present invention relates to methods and systems for supplying cooling air to multiple gas turbine engine components such as turbine vanes and, more particularly, to first stage turbine vanes for improving overall gas turbine engine efficiency. PowerPHASE, LLC, the assignee of the present invention, has developed a separate technology, known as Turbophase®, in which waste heat from the engine is utilized to heat the compressed air prior to injection into the gas turbine engine. It has a patent pending supplemental compression system that delivers air to the compressor discharge area through a compression and heating process driven by the fueled engine. Prior art air compression and delivery systems are unable to provide compressed air at the temperatures and pressures required to provide adequate cooling and improve the thermal efficiency of gas turbine engines.

[0026]
Referring now to FIG. 6, a system 600 for providing an alternative cooling source to first stage turbine vanes 631 is shown. System 600 includes a compressor 602 that directs compressed air into a discharge plenum 604 . Most of the air from compressor 602 passes through combustor case 608, end cap 610, combustion liner 612, swirler assembly 614, transition piece 616, and transition piece 616 to a portion of the turbine frame, here first stage vane outer ring 620. It passes through one or more combustors 606 that have brackets 618 that hold them in place. Air is received within combustor 606 and mixed with fuel from one or more fuel nozzles 622 . In this embodiment, the first stage vane outer ring 620 is secured to a compressor discharge plenum (CDP) case 624 .

[0027]
Air in the compressor discharge plenum is sealed between rotor 628 and inner casing 630 by seal 626 so that most of the air either goes to combustor 606 or TCLA (turbine cooling and leakage air). . The inner casing 630 has a mechanical interface 632 with a first stage nozzle 631 that provides the necessary structural axial and torsional support. Inner casing 630 is generally supported within compressor discharge plenum case 624 by ID struts 634 located between adjacent combustors 606 . Rotor 628 has bearings 636 that connect rotor 628 to the casing by struts 634 .

[0028]
Still referring to FIG. 6, the system 600 also provides an alternative source of TCLA to the first stage nozzle 631 of the gas turbine engine. An air supply is provided at A to flange 650 for case 624 . This air source A originates from an auxiliary source, as shown in FIG. More particularly, referring to FIG. 7, a supplemental source of compressed air 700 comprises a fuel engine 702 that receives air 704 and engine fuel 706 and produces mechanical shaft power 708 and hot exhaust 710 . Engine fuel 706 may be natural gas or liquid fuel. Machine shaft power 708 is utilized to drive a multi-stage intercooling compressor 712 in which ambient air 714 is taken in, compressed and cooled at each stage of the compressor 712 . Compressor 712 provides warm compressed air 716 that is directed to recuperator 718 and heats compressed air 716 with hot exhaust 710 from fuel engine 702 , thereby generating heated compressed air 720 and warm exhaust 722 . . This heated compressed air has a temperature of approximately 400 degrees Fahrenheit and warm exhaust 722 . The auxiliary source of compressed air 700 may also include a valve 724 for regulating the flow of heated compressed air 720 .

[0029]
One such auxiliary source of compressed air, depicted in FIG. 7, that can be utilized with the present invention is the patent-pending Turbophase® system manufactured by PowerPHASE LLC of Jupiter, Florida. In this system, air is compressed, heated to an intermediate temperature of approximately 400 F, and supplied at a pressure slightly above the compressor discharge pressure of compressor 602 . Heated compressed air 720 is generated approximately 25% more efficiently than compressed air from compressor 602 due to the system's patent-pending generation process.

[0030]
Referring again to FIG. 6, a supplemental source of compressed air 700, designated A in FIG. injected into Seal 654 further includes an air supply hole 656 for supplying TCLA air. This plenum 652 also includes a swirler 658 designed to serve multiple functions. That is, when the heated compressed air is being delivered at A, the tangential vortices of the air reduce the actual flow rate of air that can enter the first stage nozzle 631, causing a fraction of the air from the compressor 602 to flow. aerodynamically prevents fluid from exiting through feed hole 656 . The feed holes 656 are large enough to supply the required level turbine nozzles 631 of cooling air when no heated compressed air is being supplied at A. Air is then supplied to vanes 631 through inlets 643 . If the supply of compressed air at A is most reliable, the supply hole 656 can be eliminated.

[0031]
Compressed air for cooling may also be supplied to the inner diameter region of the first stage nozzle 631 . More specifically, referring to FIG. 6, compressed air is taken from plenum 652 and directed through a plurality of pipes 660 to inner diameter plenum 662 to the inner diameter region of first stage nozzle 631 . Also located in the inner diameter plenum 662 is a seal 664 located between the first stage nozzle inner diameter platform and the inner case 641 . A TCLA feed hole 666 is positioned within this seal 664 . This plenum 662 also includes a swirler 668 designed to serve two functions. First, when the compressed air from the auxiliary source of compressed air 700 is being delivered at A, a tangential swirl is imparted, reducing the actual flow rate of air that can be provided to the first stage nozzles 631 . , aerodynamically prevents a portion of the compressor discharge air from exiting through the TCLA feed hole 666 . The TCLA supply holes 666 are large enough to supply the current level of TCLA to the first stage nozzles 631 when the auxiliary source of compressed air 700 is not delivering air. If Turbophase® TCLA is most likely, the TCLA feed hole 666 can be eliminated.

[0032]
Referring now to Figure 8, an alternative embodiment of the present invention is shown. In this embodiment of the invention, compressed air from an auxiliary source of compressed air, indicated as A, is fed into inlet flange 802 . Control valve 804 is located adjacent inlet flange 802 . When control valve 804 is closed, pipe 660 feeding air to stage one nozzle 631 forces all of the air into stage one nozzle outer diameter region 652 and stage one nozzle inner diameter region 662 . do.

[0033]
As will be appreciated by those skilled in the art, valve 804 may be a control valve or a check valve. If the auxiliary source of compressed air is not operating and is supplying air, then control valve 804 opens and air flows from gas turbine compressor discharge plenum 604 through compressor discharge flange 806 to outer diameter plenum 652 and inner diameter plenum 652 . Entering plenum 662 via pipe 660 , air may be supplied to first stage nozzle 631 . When valve 804 is open and there is air being supplied at A, depending on the pressure and flow rate of the added air, air from the compressor discharge case of the gas turbine will flow into or out of flange 806. may flow out from If the flow exits flange 806, then the resulting temperature of the mixed air flow, which is a mixture of air and air from the gas turbine compressor discharge case from auxiliary compressor source A, is the mixing temperature. Since the gas turbine compressor outlet temperature is typically 750°F (399°C), the air being supplied from the auxiliary compressor is below 750°F (399°C) and the mixture temperature is below the compressor discharge temperature. If no air is supplied from auxiliary compressor source A, then compressor discharge air exits flange 806 to supply cooling air to the nozzles.

[0034]
Having high pressure air available at A from an external compressor can accomplish other functions. In gas turbines, the clearance between the rotating blade inner diameter platform, also known as the rim cavity, and the adjacent upstream and downstream nozzles is typically a very delicate and difficult area to keep cool. There is By supplying the TCLA to the rim cavity where the TCLA has a higher pressure than the pressurized gas in the channel, the pressurized gas in the channel is prevented from flowing into the rim cavity. Some current gas turbines have very low pressure limits in the rim cavity that either limit their operation or force the TCLA to increase significantly to prevent adequate rim cavity temperatures. to maintain An auxiliary source of compressed air can provide air at a higher pressure than the engine compressor 602, or at a TCLA pressure, thereby reducing current TCLA utilization, resulting in increased engine efficiency.

[0035]
A characteristic of a typical gas turbine engine is that as the coolant temperature decreases, less air is required to provide the same level of cooling to maintain the lowest metal temperatures of cooled components within the turbine. . This may improve efficiency. For example, alternative counterparty brand manufacturers, including Siemens Westinghouse and Mitsubishi Heavy Industries, utilize TCLA's cooling systems, which are also utilized in some of their turbines. This system is referred to as a Rotor Air Cooler (RAC) system and cools a portion of the TCLA outside the gas turbine engine at air temperatures between approximately 750°F (approximately 399°C) and approximately 450°F (approximately 232°C). C) leads to a cooler that reduces to C). This temperature reduction is sufficient to reduce the amount of cooling air required, but still hot enough to eliminate the risk of thermal shock to the components receiving the cooled air. RAC air is returned to the rotating parts of the gas turbine engine through a pipe after the cooler due to the pressure sensitivity described above.

[0036]
These performance improvements are based on the fact that air from an auxiliary source of compressed air is used in the cooling system of the first stage turbine nozzle so that the control system of the gas turbine can be adjusted to properly maintain the same first stage nozzle temperature. can be achieved using a passive cooling system directed to the inlet of the Utilizing this passive system, when the auxiliary supply of compressed air is not operating, the combustion temperature is unaffected, but as the cooler cooling airflow is directed to the first stage turbine nozzle, gas Fuel flow to the combustors can be proportionally increased to improve the power and efficiency of the turbine system.

[0037]
All of the cooling air supplied to the first stage nozzles comes from a supplemental source of compressed air, and as a result also utilizes non-passive or dedicated systems which are the most reliable systems to implement. can do. With this configuration, a higher pressure, different cooling scheme can be used to improve the cooling effect of the first stage nozzle. For example, if it were possible to improve the cooling effectiveness of 0.59 to 0.65 by about 10%, then the amount of cooling air could be reduced by about 10 lb/sec, thereby giving a 170 MW gas turbine about An additional 4 MW of power is added, or an increase in power and efficiency of approximately 2.4%. This incremental power and efficiency is additive to the cooler cooling air and the constant cooling effect described above.

[0038]
Referring now to FIG. 9, an alternative embodiment of dedicated cooling system 900 includes a closed loop system in which air is extracted from compressor discharge plenum 902, cooled by cooler 904, and then increased in pressure by compressor 906. . Pressurized air 908 is then piped through inlet 910 to a dedicated cooling system for cooling first stage nozzle 931 . Instead of discharging the cooling air into the hot gas path, as is conventionally done in air-cooled nozzles, some or all of the cooling air is returned to the compressor discharge plenum 902 where it passes through the combustion process and releases the cooling air. Circulate efficiently. One major advantage of this process is that no new air is added to the gas turbine cycle and thus the mass flow rate of the gas turbine exhaust remains relatively unchanged, making it much easier to pass through the gas turbine exhaust mass flow rate. can be kept relatively constant. For example, current combined cycle power plants may utilize duct burners that produce much higher emissions than the gas turbine itself for incremental power. Since the auxiliary source of compressed air operates on top of the gas turbine and has the emissions characteristics of a gas turbine, incremental emissions are much lower for incremental megawatts of power generated.

[0039]
Another advantage realized by the closed-loop cooled first vane of FIG. Being able to use the system. Currently, auxiliary sources of compressed air are primarily power augmentation systems and can provide some load benefits, but are somewhat limited to very low loads due to gas turbine compressor surge limitations. The closed-loop cooling system shown in FIG. 9 increases or decreases the cooling applied to the air as it leaves the compressor discharge plenum, thereby increasing the temperature of the air being returned from the cooled first vane 931. can be effectively controlled, thereby allowing the gas turbine to operate with lower operating limits.

[0040]
However, by utilizing cooler air (approximately 400°F (approximately 204°C)) to cool the nozzle, the air exiting the nozzle is much cooler (1000°F (538°C) (approximately 700° F. (approximately 371° C.)), and therefore the cooler nozzle cooling air is mixed with the gases in the hot gas path to effectively reduce the combustion temperature. By maintaining the same cooling effect and reducing the coolant temperature, the combustion temperature can be effectively increased. For example, for one embodiment of the present invention, the cooling effect is approximately 0.59 [(2700-1550)/(2700-750)=0.59]. Keeping this at a higher combustor temperature and lower coolant temperature increases the combustion temperature as follows. 0.59=(2700+x-1550)/(2700+x-400), x=504F. Therefore, if the air for cooling the first stage nozzle is cooler, the effective combustion temperature will be reduced to about 500° C. while maintaining nozzle metal temperature and life and increasing power and efficiency of the gas turbine system. F (approximately 260°C) can be raised.

[0041]
In prior art gas turbines, static components such as first stage nozzles (referred to as turbine vanes) are air cooled through differential air pressure in the nozzles. The nozzle is cooled by compressor discharge air, and with similar pressure outside the nozzle, there is almost no pressure limit at the leading edge of the nozzle. For example, if the combustor pressure drop is 2.5%, the compressor discharge pressure is 220 psig and then the pressure seen at the leading edge of the nozzle is about 214.5 psi, with only about 5.5 psi of pressure for air flow. forced through the vane's cooling system and out through its leading edge. For this reason, the air supply to the leading edge of the vane is normally done with as little pressure drop as possible. For example, air can be obtained from the inner diameter region of the transition piece to attempt to obtain a portion of the total pressure associated with the flow rate coming from the compressor diffuser. Similarly, within the nozzle, the leading edge, which typically consumes a large amount of cooling air, is responsible for most of the heat transfer utilized to keep the nozzle cool as the air passes through a series of leading edge showerhead holes. Transpiration and film cooling, which is a combination of heat conduction to cool the Advanced gas turbines typically have hundreds of cooling holes densely packed in the leading edge of the nozzle to provide this function. After passing through these holes in the nozzle leading edge, the air is directed to spread over the surface of the nozzle airfoil as a film cooling layer to dilute the hot gases that strike the nozzle directly.

[0042]
The present invention provides cooling air at a pressure that can be adjusted above the compressor discharge pressure to provide a different and more efficient cooling scheme to the leading edge of the nozzle. The use of a separately driven compressor, either electrically powered or powered by an auxiliary engine, directs this air through a piping and manifold network to the nozzles to provide a source of compressed air with means dedicated to supplying air to the nozzles. be done. Reduces the amount of conduction and film required to cool the nozzle leading edge by first inward impinging directly on the leading edge to improve heat transfer instead of utilizing prior art conduction and film cooling schemes. A large pressure drop can be used for this purpose. Sufficient to prevent nozzle catastrophic events such as complete airfoil burn-through in the event that an unforeseen event such as an inferior fuel induced foreign object damage (FOD) leaves a burnt hole in the leading edge of the nozzle. Several other unique features that allow the air supply to the impingement holes to be engineered so that the pressure that impacts the leading edge can be engineered or adjusted in real time to provide cooling. can also be added.

[0043]
Additionally, as will be appreciated by those skilled in the art, turbine nozzles typically include multiple cooling circuits. One such circuit discharges the cooling air at the exit plane of the nozzle after the pressure drop associated with the nozzle has occurred, so that much less pressure is required to drive the cooling flow. edge circuit. As a result, a portion of the air utilized for impingement cooling of the nozzle leading edge can be channeled from the interior into the nozzle along its path to effect cooling of the nozzle trailing edge region, where it is located behind the nozzle. The air can be used to cool the edges. This differs from prior art nozzles where the air utilized to cool the leading edge of the nozzle is dedicated to the leading edge region. Here, air can be utilized to cool the leading and midsection of the nozzle and/or the trailing edge of the nozzle as the pressure builds above the compressor discharge pressure. This versatile use of cooling air results in significantly less cooling air required to cool the nozzles, thereby increasing the efficiency of the gas turbine system.

[0044]
Additionally, as will be appreciated by those skilled in the art, nozzle cooling systems are designed for inspection intervals that are typically 24,000 hours between inspections. The design point is the hottest condition, usually base load operation, and at part load where the combustion temperature drops, the nozzle metal temperature also drops below the design condition. For separately cooled nozzle systems, pressure, temperature and/or flow are varied to increase metal temperature at part load conditions, thereby further reducing cooling air to the nozzle and improving part load efficiency. be able to.

[0045]
Similarly, there are typically hot spots within the nozzle, ie, areas of the nozzle operating at higher metal temperatures. These regions sometimes relate to the nozzle tangential position relative to the transition piece. For example, in one engine, such as the Siemens Westinghouse 501F gas turbine, there are 16 transition pieces and 32 first stage nozzles. Sixteen of the nozzles are located on the sidewalls of the transition piece and the remaining sixteen nozzles are located in the center of the discharge frame of the transition piece. As a result, nozzles located on the sidewalls of the transition piece experience lower hot gas path temperatures due to the cooling and leakage flow of the transition piece sidewalls. As such, these nozzles typically run at a much lower temperature than the nozzles and are directly exposed to the hot combustion gases exiting the transition piece. As disclosed herein, with a dedicated nozzle cooling system, the cooling air supply is bifurcated into two regions and the metal temperature of the nozzle, and thus life, is controlled by the nozzle and transition located near the sidewalls of the transition piece. It can be separately controlled to be the same for the nozzles in the piece discharge passage.

[0046]
As will be appreciated by those skilled in the art, adjustment of the cooling air flow rate can be accomplished by a variety of means. For example, exemplary means for regulating the flow of cooling air to the nozzles may include various engine control algorithms and mechanical means including, but not limited to, flow control valves and metering plates.

[0047]
This unique cooling configuration and process can also be applied to the turbine nozzle sector. In many cases, the hot gas temperature from the combustor varies around the gas turbine nozzle inlet region. With a dedicated nozzle cooling system that is divided into sectors, each sector can be adjusted to provide constant cooling temperature and life as the gas temperature changes. With this mechanism, if an unforeseen event such as FOD causes premature failure of a component, the cooling air temperature, flow rate, and/or Or you can adjust the pressure. As will be appreciated by those skilled in the art, multiple combinations of cooling air pressure, temperature and flow rate can be individually adjusted to achieve similar results in order to achieve the desired cooling effect. As a result, in particular, it is also envisaged that in some cases the pressure may not rise to cool the nozzle components. Although first stage turbine nozzles have been utilized herein, application of the present invention to first stage nozzles is only one expression of the potential uses of the present invention. The invention is also applicable to other static components including other turbine nozzles and shroud blocks.

[0048]
As will be appreciated by those skilled in the art, the principles described for reducing cooling air to the first turbine vane lead directly to increased efficiency and can be applied to other turbine components. For example, a first stage blade outer air seal is a seal located radially outward of the first stage turbine blades. It is also a difficult part to cool due to the operating pressures and temperatures. Therefore, when using a separate source of cooling air whose air pressure can be controlled above the pressure available in the gas turbine, the cooling air first cools the inside using some impingement mechanism and then the film It becomes possible to take advantage of alternative cooling technologies spread as

[0049]
As described above, the present invention provides a method for cooling turbine nozzles in which cooling air is supplied through a separate process external to the gas turbine engine, such as through a supplemental source of compressed air 700 as shown in FIG. . The cooling air so compressed has a pressure in excess of the air in the compressor discharge plenum and is directed to the leading edge of the turbine nozzle. In one embodiment of the invention, a portion of the air from the leading edge is then directed to cool a portion of the turbine nozzle behind the leading edge, such as the trailing edge or midsection of the turbine nozzle. This circulation or reuse of cooling air is made possible by the lower temperature and higher pressure of the air generated by the auxiliary source of compressed air.

[0050]
In one embodiment of the invention, the distribution of compressed air from the auxiliary source of compressed air is controlled to vary the flow rate to the turbine nozzles according to predetermined control parameters. Various control parameters may be utilized including air pressure, temperature, air flow rate, or a combination of these control parameters. That is, the amount of cooling flow supplied to the turbine nozzles generated by the individual external processes is regulated by the respective air pressure, temperature, or cooling air flow rate. This process is regulated by a system that measures the control parameters of the air generated by the auxiliary source of compressed air, as well as the temperature and pressure at the turbine nozzle and consequently adjusts the flow of cooling air to the turbine nozzle.

[0051]
While the invention has been described in what are presently known as the preferred embodiments, the invention is not intended to be limited to the disclosed embodiments, but rather various modifications and equivalents within the scope of the following claims. It should be understood that it is intended to cover mechanisms. The present invention has been described in relation to particular embodiments which are intended in all respects to be illustrative rather than restrictive. Specifically, a first stage nozzle is utilized as an example in this application. However, the principles apply to other rotating and stationary turbine components commonly referred to as hot gas path components.

[0052]
From the foregoing description, it will be seen that the present invention is well suited to attain all of the objectives set forth above, along with other advantages which are obvious and inherent to the present system and method. It is understood that certain features and subcombinations are useful and can be employed without reference to other features and subcombinations. This is contemplated by and is within the scope of the claims.

Claims (10)

A method of cooling a turbine nozzle in a gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section fluidly coupled together, comprising:
cooling air for the turbine nozzle through a separate process external to the compression process of the compressor section of the gas turbine engine to form a supply of compressed air having a pressure exceeding the air in the compressor discharge plenum; compressing a portion; and
directing the supply of compressed air to a leading edge of the turbine nozzle;
directing a portion of the compressed air supply from the leading edge to a portion of the turbine nozzle aft of the leading edge;
including
The method, wherein the supply of compressed air is based on a predetermined control parameter, the predetermined control parameter being any one of air pressure, flow rate, and any combination of air pressure, temperature and flow rate. 2. The method of claim 1, wherein the supply of compressed air is utilized to cool the trailing edge of the turbine nozzle after internally cooling the leading edge of the turbine nozzle. 2. The method of claim 1, wherein the compressed air supply is utilized to cool the turbine nozzle center after cooling the turbine nozzle leading edge. A method of cooling a plurality of turbine nozzles in a gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section all fluidly connected to each other, comprising:
compressing a portion of the cooling air to form compressed air for the turbine nozzle, wherein the cooling air is compressed in a separate process external to the gas turbine engine process;
controlling the distribution of the compressed air around a stage of turbine nozzles to vary the compressed air supplied to each turbine nozzle in the stage;
a method comprising
The method, wherein the distribution of the compressed air is based on a predetermined control parameter, the predetermined control parameter being any one of air pressure, flow rate, and any combination of air pressure, temperature and flow rate. A method for cooling static components in a gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section fluidly coupled together, comprising:
compressing a portion of cooling air for the static components of the turbine section in a separate process external to the gas turbine engine process;
controlling distribution of the compressed cooling air around the static component based on control parameters;
including
The method, wherein the control parameter is any one of air pressure, flow rate, or any combination of air pressure, temperature, and flow rate. A method of cooling components of a gas turbine engine having a compressor section, a compressor discharge plenum, a combustor section, and a turbine section fluidly connected together, the method comprising: compressing air having a higher pressure than air in the compressor discharge plenum. generating at least a portion of cooling air for the components in a process external to the gas turbine engine process to form a supply;
The method, wherein the supply of compressed air is based on a predetermined control parameter, the predetermined control parameter being any one of air pressure, flow rate, and any combination of air pressure, temperature and flow rate. 7. The method of claim 6, wherein said component is a turbine nozzle. 8. The method of claim 7, wherein at least a portion of said air supplied from said process external to said gas turbine engine process is directed to a leading edge of said turbine nozzle. at least a portion of the air supplied from the process external to the gas turbine engine process is first directed to the leading edge of the turbine nozzle; 9. The method of claim 8, wherein at least a portion is then directed to a central or trailing edge region of the turbine nozzle after cooling of the leading edge of the turbine nozzle. The component is a blade outer air seal, wherein at least a portion of the air supplied from the process external to the process of the gas turbine engine first enters the blade outer air seal. 7. The method of claim 6, having a configuration to direct it to impinge inwardly and then to cooling holes for film cooling of said blade outer air seal.

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