US20090226327A1

US20090226327A1 - Gas Turbine Engine Including Temperature Control Device and Method Using Memory Metal

Info

Publication number: US20090226327A1
Application number: US12/044,017
Authority: US
Inventors: David A. Little; Hubertus E. Paprotna
Original assignee: Siemens Power Generations Inc
Current assignee: Siemens Energy Inc
Priority date: 2008-03-07
Filing date: 2008-03-07
Publication date: 2009-09-10

Abstract

The gas turbine engine includes a compressor section, a combustion section downstream from the compressor section and a turbine section downstream from the combustion section. The turbine section includes a rotor shaft, a plurality of turbine blades, and a plurality of discs coupling corresponding ones of the plurality of turbine blades to the rotor shaft. Each disc has a plurality of disc cooling fluid passages therein associated with cooling the turbine blades, and a respective thermal shape memory sleeve is in at least some of the disc cooling fluid passages. The thermal shape memory sleeve defines a sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough. Part load efficiency may be increased by supplying only the needed cooling fluid to areas of the blades.

Description

FIELD OF THE INVENTION

The present invention relates to the field of turbine engines, and, more particularly, to cooling air flow in turbines and related methods.

BACKGROUND OF THE INVENTION

A gas turbine typically includes a compressor section that produces compressed air. Fuel is then mixed with and burned in a portion of this compressed air in a combustion section having one or more combustors, thereby producing a hot compressed gas. The hot compressed gas is then expanded in a turbine section to produce rotating shaft power. The turbine section typically uses a plurality of alternating rows of stationary vanes and rotating blades. Each of the rotating blades has an airfoil portion and a root portion by which it is affixed to a rotor, e.g. via a disc. The root portion includes a platform from which the airfoil portion extends.
Blades and vanes in gas turbine engines may include cooling air passages leading to an outer surface of the airfoil that requires cooling. These cooling air passages are typically located in specific locations on the airfoil where extremely high temperatures exists during operation of the engine. Certain regions of the surface require larger amounts of cooling air than other areas that require less cooling air. When designing the size of the cooling air passages, the designer typically sizes the passages to be able to supply the amount of cooling air to cool the airfoil surface under the worst case situation of highest possible heat load. This design temperature is reached under normal operation of the engine. Also, the heat load varies on surfaces of the airfoil, so not every surface requires the same amount of cooling airflow. Thus, the amount of cooling air passing through the passage and onto the external surface of the airfoil is designed to adequately cool that area of the airfoil. Thus, no cooling air flow is wasted and overall engine performance and efficiency is optimized.
U.S. Pat. No. 6,408,610 issued to Caldwell et al. is directed to “A METHOD OF ADJUSTING GAS TURBINE COMPONENT COOLING AIR FLOW”, and shows an airfoil that includes a plurality of cooling holes having a thermal barrier coating applied at various thicknesses in the holes to provide a desired hole diameter. Under this method, the size of the cooling air passages can be designed to provide a desired amount of cooling air flow onto the surface of the airfoil, depending upon the air pressure within the blade and around the opening of the cooling air passage, such that a desired amount of cooling can occur. However, the sizes of the cooling holes do not vary based upon the operating conditions of the engine in the region of the specific cooling air passage. So, the size of the cooling air passage may be smaller than needed, resulting in less cooling air flow than required, or larger than needed, resulting in more cooling air flow than required. Either way, the engine performance or efficiency is reduced.
U.S. Pat. No. 6,416,279 issued to Weigand et al entitled “A COOLED GAS TURBINE COMPONENT WITH ADJUSTABLE COOLING” in which the cooling air passage includes different means to vary the amount of cooling air flow during engine operation. In one method, a restrictor having an opening of specific size is placed in the cooling air passage to regulate the cooling air flow during engine operation. In this method, the size of the restrictor cannot be changed during engine operation. In another method, a control system is used and includes a temperature sensor and a control valve, where the control valve regulates an amount of cooling air flow based upon a value from the temperature sensor.
U.S. Pat. No. 6,485,255 issued to Care et.al. entitled “A COOLING AIR FLOW CONTROL DEVICE FOR A GAS TURBINE ENGINE” is directed to a single shape memory metal valve disposed in a cooling passage upstream of the many cooling air passages that open out onto the outer surface of the airfoil. In the Care et al. device, the valve varies the air flow depending upon temperature, but all of the cooling air passages opening onto the airfoil surfaces are controlled by this single valve. The passages exposed to the hottest surface of the airfoil are regulated by the same valve and supply airflow as the openings exposed to the coolest airfoil surface so that cooling may be insufficient at some locations.
U.S. Pat. No. 7,241,107 issued to Spanks et al. entitled “GAS TURBINE AIRFOIL WITH ADJUSTABLE COOLING AIR FLOW PASSAGES” discloses an airfoil for a gas turbine engine, wherein the airfoil includes a plurality of cooling air passages to supply cooling air to an external surface of the airfoil. The cooled surface of the airfoil has a critical temperature in which any cooled surface of the airfoil should not exceed. The cooling air passages have a coating applied within the passages, and the coating is made of a material that has an oxidizing property such that the material oxidizes away and opens the passage to more flow when exposed to a temperature above the critical temperature. When the airfoil surface is not properly cooled by a flow passing through the passage, the material oxidizes away until the size of the passage increases to allow for the proper amount of cooling air to flow to cool the airfoil. Each passage is located in a different part of the airfoil that requires more or less cooling flow, and each passage will oxidize until the size of the passage is large enough to allow for the proper amount of cooling flow.

SUMMARY OF THE INVENTION

In view of the foregoing background, it is therefore an object of the present invention to provide a gas turbine engine with efficient controlled temperature regulation of the turbine section blades.
This and other objects, features, and advantages in accordance with the present invention are provided by a gas turbine engine including a compressor section, a combustion section downstream from the compressor section and a turbine section downstream from the combustion section and including a rotor shaft, a plurality of turbine blades, and a plurality of discs coupling corresponding ones of the plurality of turbine blades to the rotor shaft. Each disc has a plurality of disc cooling fluid passages therein associated with cooling the turbine blades, and a respective thermal shape memory sleeve is in at least some of the disc cooling fluid passages. The thermal shape memory sleeve defines a sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough.
Thus, the approach addresses blade tip clearance control that may be needed when the turbine inlet temperature is maintained at a high level during part load operation of a gas turbine engine, which may be done, for example, to reduce CO emissions. To that end, features may include ensuring a minimum blade tip clearance under such conditions.
Conventionally, during part load conditions when the rotor cooling air temperature is decreased to maintain the integrity of the blade tips, the blades are overcooled. The present approach may more efficiently cool the blades at the cooler rotor cooling air temperature (discs set tip clearances) by reducing the throat area of the memory metal sleeves to reduce the cooling flows to those required at the cooler rotor air temperature and turbine inlet temperature. Accordingly, part load efficiency may be improved by supplying only needed cooling air to all areas of the blades.
The thermal shape memory sleeve may be a thermal shape memory metal sleeve or a thermal shape memory polymer sleeve, for example. A cooling system may be included and have a cooler to cool at least a portion of compressed air from the compressor section and to provide cooling fluid to the plurality of disc cooling fluid passages. Each of the plurality of turbine blades may have a plurality of blade cooling fluid passages therein coupled in fluid communication with the disc cooling fluid passages.
The sleeve throat opening may have a cylindrical shape. The sleeve throat opening of the thermal shape memory sleeve may be larger when the temperature is relatively high, and the sleeve throat opening may be smaller when the temperature is relatively low. The sleeve throat opening of the thermal shape memory sleeve may have a shape transition temperature in a range of 175-400° C.
A method aspect is directed to controlling blade tip clearance in a gas turbine engine including a compressor section, a combustion section downstream from the compressor section, and a turbine section downstream from the combustion section and including a rotor shaft, a plurality of turbine blades, and a plurality of discs coupling corresponding ones of the plurality of turbine blades to the rotor shaft. The method includes forming a plurality of disc cooling fluid passages in each disc for cooling the turbine blades, and positioning a respective thermal shape memory sleeve in at least some of the cooling fluid passages to define a sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough.

BRIEF DESCRIPTION OF THE DRAWINGS

FIG. 1 is partly schematic cross-sectional view of an example of a portion of a turbine engine in accordance with the present invention.

FIG. 2 is a more detailed cross-sectional view of a rotor disc of the turbine engine of FIG. 1.

FIG. 3 is an enlarged detailed cross-sectional view of a portion of the rotor disc of FIG. 2 within the dashed circle B.

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS

The present invention will now be described more fully hereinafter with reference to the accompanying drawings, in which preferred embodiments of the invention are shown. This invention may, however, be embodied in many different forms and should not be construed as limited to the embodiments set forth herein. Rather, these embodiments are provided so that this disclosure will be thorough and complete, and will fully convey the scope of the invention to those skilled in the art. Like numbers refer to like elements throughout.
Referring initially to FIG. 1, a gas turbine engine 10 in accordance with features of the present invention will now be described. The turbine 10 includes a turbine section 11, a compressor section 12 and a combustion section 15. Turbine engines, such as single shaft industrial gas turbines, are designed to operate at a constant design turbine inlet temperature under any ambient air temperature (i.e., the compressor inlet temperature). This design turbine inlet temperature allows the engine to produce maximum possible power, known as base load. Any reduction from the maximum possible base load power is referred to as part load operation. In other words, part load entails all engine operation from 0% to 99.9% of base load power.
Part load operation may result in the production of high levels of carbon monoxide (CO) during combustion. One known method for reducing part load CO emissions is to bring the combustor exit temperature or the turbine inlet temperature near that of the base load design temperature. Also, some of the compressor exit air 14 from the combustion section 15 may be used to cool the stationary support structure 16 of the turbine near the first row of blades 20 a. The stationary support structure 16 can include the outer casing, blade rings, and ring segments. In addition, some compressed air may be piped directly out of the compressor section 12 through piping 19 a (additional pipes not shown). This compressor bleed air is used to cool the stationary support structure 16 near the second, third and fourth rows of blades 20 b, 20 c, 20 d and is supplied through piping 19 b, 19 c, 19 d.
The support structure 16 may contract or shrink in radius when exposed to the cooler compressor exit and bleed air. But, at the same time, the temperature of the hot gas leaving the combustor 18 and flowing over the turbine blades 20 (illustratively 20 a, 20 b, 20 c, 20 d) is kept at a high level, causing a constant radially outward thermal expansion of the blades 20.
The expansion of the blades 20 along with the shrinkage of the support structure 16 reduces the clearance C between the tips 21 of the blades 20 and the surrounding support structure 16, commonly referred to as the blade tip clearance C. While the clearance C is shown between the fourth row of blades 20 d and the adjacent support structure 16, similar clearances C exist between the first, second and third rows of blades 20 a, 20 b, 20 c and the stationary support structure 16. A minimal blade tip clearance C should be maintained so that the blades 20 do not rub against the support structure 16.
In this gas turbine engine 10, compressor exit air 14 from the combustion section 15 can be used to cool at least the turbine rotor 50, discs 52, and blades 20. The compressor exit air may routed out of the engine, passed through a cooling fluid loop 22, including a cooler 24, and is ultimately redelivered to the engine at a substantially constant design cooling air return temperature. The design temperature is held substantially constant so that the discs 52 and blades 20 temperatures are held substantially constant, e.g. to maintain the life of the discs 52 and blades 20.
The design cooling return temperature can be specific to a particular engine design. For example, in the Siemens W501G engine, the design cooling air return temperature is about 350° Fahrenheit for loads below 90%. When load reaches 90%, the temperature of the rotor cooling air can be increased to about 405° Fahrenheit to about 480° Fahrenheit. The rotor cooling temperature is increased at loads of 90% and above as an active way to reduce tip clearances C at full load to maximize engine power and efficiency. When rotor cooling air return temperature increases, the discs 52 and blades 20 expand radially outward, causing the clearance C between the tips 21 of the blades 20 and the nearby stationary support structure 16 to decrease. The smaller clearance C means less losses and, thus, more power extraction for the same fuel input, thereby increasing efficiency. Further, the rotor cooling air return temperature can be increased above 90% load because, by the time the engine reaches that level, most of the stationary components 16 of the engine have thermally grown to their final shapes. Thus, distortion and ovalization, which can cause blade tip rubbing, are minimized.
The Siemens Westinghouse W501G turbine engine is only one example of a design cooling air return temperature. In other engines, the design cooling air return temperature can range from about 350° Fahrenheit to about 750° Fahrenheit.
Whatever the range, the design cooling air return temperature can be supplied at a substantially constant temperature so that the rotor 50 and the discs 52 are always at substantially constant metal temperatures. However, the temperature of the compressor exit air 14 often exceeds the design cooling air return temperature. Therefore, to provide cooling air at the design temperature, a portion of compressor exit 14 air may be bled from the Combustion section 15 and cooled to the appropriate temperature in the external cooling loop 22, which is configured to reduce the cooling air to the design cooling air return temperature. The external cooling loop 22 can also include heat exchanger devices as well as valves for controlling the quantity of air passing through or bypassing the heat exchanger devices so as to achieve the design cooling air return temperature. Once treated, the cooled air can be returned to the engine to cool at least the rotor 50 and discs 52 at the substantially constant design cooling air return temperature.
As discussed above, features of the turbine engine 10 relate to ensuring adequate blade tip clearance C under part load conditions. Such features may include reducing the cooling air return temperature to below the design temperature level to maintain a minimum acceptable clearance C between a blade tip 21 and surrounding stationary support structure 16. For example, the minimum acceptable clearance C may be about 1 millimeter or about 0.040 inches. Because the temperature of the cooling return air may be lower than the design temperature, the discs 52 and blades 20 will tend to shrink when exposed to the cooling return air. In spite of the expansion of the turbine blades 20 due to the passing high temperature gases as well as the shrinkage of the stationary support structure 16 due to cooler compressor exit air 14 and compressor bleed air temperatures, as described earlier, an adequate blade tip clearance C may be maintained with the cooler temperature of the cooling air return causing the discs 52 and blades 20 to shrink, thus widening the gap between the tip 21 of the blade 20 and the nearby support structure 16 at part load.
The turbine section 11, downstream from the combustion section 15, includes the rotor shaft 50, the plurality of turbine blades 20, and a plurality of discs 52 coupling corresponding ones of the plurality of turbine blades to the rotor shaft. Referring additionally to FIGS. 2 and 3, each disc 52 has a plurality of disc cooling fluid passages 54 therein associated with cooling the turbine blades 20, and a respective thermal shape memory sleeve 56 is in at least some of the disc cooling fluid passages. The thermal shape memory sleeve 56 defines a sleeve throat opening A that changes based upon a temperature to adjust a flow of cooling fluid 60 therethrough.
Tip clearance control is obtained by operating with rotor cooling air return temperature low at lower loads (transient conditions, i.e. loading and unloading the turbine engine) to keep the blade tips from rubbing, then by increasing rotor cooling air temperature near base load. This increase tightens tip clearances (once the stationary components have stabilized) primarily because of the increase in radii of the discs. The increase in blade lengths is a secondary contributor to the reduction in tip clearance.
Blading is designed at base load conditions where both turbine inlet temperature and rotor cooling air are maximized. These worst case conditions will occur continuously across the ambient temperature range when the engine is on base load control (100% load). Cooling flow consumption is regulated, e.g. optimized or minimized, everywhere on each airfoil at base load conditions. However, in conventional systems, during part load conditions when the rotor cooling air temperature has been decreased to maintain the integrity of the blade tips, the blades may be overcooled.
The present approach, including the use of thermal shape memory sleeves 56 in disc cooling fluid passages 54, may more efficiently cool the blades at the cooler rotor cooling air temperature (discs set tip clearances) by reducing the throat area of the thermal shape memory sleeves 56 to reduce the cooling flows to those needed at the cooler rotor air temperature and turbine inlet temperature. Part load efficiency may be made more efficient by supplying only needed cooling air to all areas of the blades 20.
The thermal shape memory sleeve 56 may be a thermal shape memory metal sleeve or a thermal shape memory polymer sleeve, for example. A shape memory alloy (SMA, also known as a smart alloy or memory metal) is an alloy that “remembers” its geometry. After a sample of SMA has been deformed from its original crystallographic configuration, it regains its original geometry by itself during heating (one-way effect) or, at higher ambient temperatures, simply during unloading (pseudo-elasticity or superelasticity). These extraordinary properties are due to a temperature-dependent martensitic phase transformation from a low-symmetry to a highly symmetric crystallographic structure. Those crystal structures are known as martensite (at lower temperatures) and austenite (at higher temperatures).
Examples of types of SMAs include copper-zinc-aluminum-nickel, copper-aluminum-nickel, and nickel-titanium (NiTi) alloys. Other SMAs may include Ag—Cd, Au—Cd, Cu—Al—Ni, Cu—Sn, Cu—Zn, Cu—ZnS—i, Fe—Pt, Mn—Cu, Fe—Mn—Si, Pt alloys, Co—Ni—Al, Co—Ni—Ga, Ni—Fe—Ga and Ti—Pd, for example.
As discussed above, a cooling loop or system 22 may be included and have a cooler 24 to cool at least a portion of compressed air 14 from the compressor section 12 and to provide cooling fluid 60 to the plurality of disc cooling fluid passages 54. Each of the plurality of turbine blades 20 may have a plurality of blade cooling fluid passages 58 (FIG. 2) therein coupled in fluid communication with the disc cooling fluid passages 54, as would be appreciated by those skilled in the art.
The sleeve throat opening A may have a cylindrically shape. The sleeve throat opening A of the thermal shape memory sleeve may be larger when the temperature is relatively high, as indicated by the dashed line in FIG. 3, and the sleeve throat opening A may be smaller when the temperature is relatively low. The sleeve throat opening A of the thermal shape memory sleeve 56 may have a shape transition temperature in a range of 175-400° C., for example. (0036] A method aspect is directed to controlling blade tip clearance C in a gas turbine engine 10 including a compressor section 12, a combustion section 15 downstream from the compressor section, and a turbine section 11 downstream from the combustion section and including a rotor shaft 50, a plurality of turbine blades 20, and a plurality of discs 52 coupling corresponding ones of the plurality of turbine blades to the rotor shaft. The method includes forming a plurality of disc cooling fluid passages 54 in each disc 52 for cooling the turbine blades 20, and positioning a respective thermal shape memory sleeve 56 in at least some of the cooling fluid passages to define a sleeve throat opening A that changes based upon a temperature to adjust a flow of cooling fluid 60 therethrough.
Thus, features of the approach address blade tip clearance C control that may be needed when the turbine inlet temperature is maintained at a high level during part load operation of a gas turbine engine 10, which may be done, for example, to reduce CO emissions. To that end, features of the approach may include ensuring a minimum blade tip clearance C under such conditions. The present approach may more efficiently cool the blades at the cooler rotor cooling air temperature (discs set tip clearances) by reducing the throat area of the thermal shape memory sleeves 56 to reduce the cooling flows to those required at the cooler rotor air temperature and turbine inlet temperature. Accordingly, part load efficiency may be improved by supplying only needed cooling air to all areas of the blades.
Many modifications and other embodiments of the invention will come to the mind of one skilled in the art having the benefit of the teachings presented in the foregoing descriptions and the associated drawings. Therefore, it is understood that the invention is not to be limited to the specific embodiments disclosed, and that modifications and embodiments are intended to be included within the scope of the appended claims.

Claims

1. A gas turbine engine comprising:

a compressor section;

a combustion section downstream from said compressor section;

a turbine section downstream from said combustion section including

a rotor shaft,

a plurality of turbine blades, and

a plurality of discs coupling corresponding ones of the plurality of turbine blades to the rotor shaft, each disc having a plurality of disc cooling fluid passages therein associated with cooling the turbine blades, and

a respective thermal shape memory sleeve in at least some of the disc cooling fluid passages and defining a sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough.

2. The gas turbine engine according to claim 1, wherein the thermal shape memory sleeve comprises a thermal shape memory metal sleeve.

3. The gas turbine engine according to claim 1, wherein the thermal shape memory sleeve comprises a thermal shape memory polymer sleeve.

4. The gas turbine engine according to claim 1, further comprising a cooling system including a cooler to cool at least a portion of compressed air from the compressor section and to provide cooling fluid to the plurality of disc cooling fluid passages.

5. The gas turbine engine according to claim 1, wherein each of the plurality of turbine blades has a plurality of blade cooling fluid passages therein coupled in fluid communication with the disc cooling fluid passages.

6. The gas turbine engine according to claim 1, wherein the sleeve throat opening has a cylindrically shape.

7. The gas turbine engine according to claim 1, wherein the sleeve throat opening of the thermal shape memory sleeve is larger when the temperature is relatively high, and the sleeve throat opening is smaller when the temperature is relatively low.

8. The gas turbine engine according to claim 1, wherein the sleeve throat opening of the thermal shape memory sleeve has a shape transition temperature in a range of 175-400° C.

9. A gas turbine engine comprising:

a combustion section;

a turbine section downstream from said combustion section including

a rotor shaft,

a plurality of turbine blades, and

a respective thermal shape memory metal sleeve in each of the cooling fluid passages and defining a cylindrical sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough.

10. The gas turbine engine according to claim 9, further comprising a cooling system including a cooler to provide cooling fluid to the plurality of disc cooling fluid passages.

11. The gas turbine engine according to claim 9, wherein each of the plurality of turbine blades has a plurality of blade cooling fluid passages therein coupled in fluid communication with the disc cooling fluid passages.

12. The gas turbine engine according to claim 9, wherein the cylindrical sleeve throat opening of the shape memory sleeve is larger when the temperature is relatively high, and the cylindrical sleeve throat opening is smaller when the temperature is relatively low.

13. A method of controlling blade tip clearance in a gas turbine engine including a compressor section, a combustion section downstream from the compressor section, and a turbine section downstream from the combustion section and including a rotor shaft, a plurality of turbine blades, and a plurality of discs coupling corresponding ones of the plurality of turbine blades to the rotor shaft, the method comprising:

forming a plurality of disc cooling fluid passages in each disc for cooling the turbine blades; and

positioning a respective thermal shape memory sleeve in at least some of the cooling fluid passages to define a sleeve throat opening that changes based upon a temperature to adjust a flow of cooling fluid therethrough.

14. The method according to claim 13, wherein the thermal shape memory sleeve comprises a thermal shape memory metal sleeve.

15. The method according to claim 13, wherein the thermal shape memory sleeve comprises a thermal shape memory polymer sleeve.

16. The method according to claim 13, further comprising cooling at least a portion of compressed air from the compressor section to provide cooling fluid to the plurality of disc cooling fluid passages.

17. The method according to claim 13, further comprising providing each of the plurality of turbine blades with a plurality of blade cooling fluid passages therein coupled in fluid communication with the disc cooling fluid passages.

18. The method according to claim 13, wherein each the sleeve throat opening has a cylindrical shape.

19. The method according to claim 13, wherein the sleeve throat opening of the thermal shape memory sleeve is larger when the temperature is relatively high, and the sleeve throat opening is smaller when the temperature is relatively low.

20. The method according to claim 13, wherein the sleeve throat opening of the thermal shape memory sleeve has a shape transition temperature in a range of 175-400° C.