Patent Application
20150198094
The present disclosure generally pertains to gas turbine engines, and is more particularly directed toward a ceramic pedestal and shield for gas path temperature measurement in the gas turbine engine.
Gas turbine engines include compressor, combustor, and turbine sections. During operation, the turbine section is subjected to high temperatures. Temperature sensors are often used to measure the gas path temperature in the turbine, and in particular the first stage of the turbine.
A shield is often located around the temperature sensor used to measure the gas path temperature. U.S. Pat. No. 4,187,434 to F. Pater, Jr. discloses a suction pyrometer radiation shield comprising an elongated first alumina refractory tube, a series of smaller alumina refractory tubes arranged around and bonded to the inside surface of said first tube forming central passageway, an outer fracture resistant alumina refractory tube surrounding said first tube and an alumina refractory washer closely surrounding said first tube in abutting contact with said outer alumina refractory tube.
The present disclosure is directed toward overcoming one or more of the problems discovered by the inventors or that is known in the art.
A mounting stand for a temperature sensor is disclosed. The mounting stand includes a base, a pedestal, and a shield. The pedestal extends from the base to a pedestal end distal to the base at a pedestal length. The pedestal includes a first hollow cylinder shape forming a bore there through. The bore is sized to hold the temperature sensor. The shield includes a second hollow cylinder shape extending from the base about the pedestal and extending beyond the pedestal to a shield end distal to the base forming a flow region with an annular shape between the pedestal and the shield. The shield also includes a first aspiration hole extending through the second hollow cylinder shape. The base, the pedestal, and the shield are formed of a ceramic material.
The systems and methods disclosed herein include a mounting stand for attaching a temperature sensor to a nozzle segment of a gas turbine engine. In embodiments, the mounting stand includes a base, a shield and a pedestal within the shield, each made of a ceramic material. The pedestal locates a temperature sensor within the shield. The overall design of the mounting stand including the length of the pedestal, the length of the shield beyond the pedestal, the size of the annular space between the pedestal and the shield, and the use of a ceramic material may reduce conduction between the nozzle segment and the mounting stand, may reduce radiation errors, may reduce convection errors, may promote a time accurate reading, and may reduce the manufacturing cost of the mounting stand.
In addition, the disclosure may generally reference a center axis 95 of rotation of the gas turbine engine, which may be generally defined by the longitudinal axis of its shaft 120 (supported by a plurality of bearing assemblies 150). The center axis 95 may be common to or shared with various other engine concentric components. All references to radial, axial, and circumferential directions and measures refer to center axis 95, unless specified otherwise, and terms such as “inner” and “outer” generally indicate a lesser or greater radial distance from, wherein a radial 96 may be in any direction perpendicular and radiating outward from center axis 95.
A gas turbine engine 100 includes an inlet 110, a shaft 120, a compressor 200, a combustor 300, a turbine 400, an exhaust 500, and a power output coupling 600. The gas turbine engine 100 may have a single shaft or a dual shaft configuration.
The compressor 200 includes a compressor rotor assembly 210, compressor stationary vanes (stators) 250, and inlet guide vanes 255. The compressor rotor assembly 210 mechanically couples to shaft 120. As illustrated, the compressor rotor assembly 210 is an axial flow rotor assembly. The compressor rotor assembly 210 includes one or more compressor disk assemblies 220. Each compressor disk assembly 220 includes a compressor rotor disk that is circumferentially populated with compressor rotor blades. Stators 250 axially follow each of the compressor disk assemblies 220. Each compressor disk assembly 220 paired with the adjacent stators 250 that follow the compressor disk assembly 220 is considered a compressor stage. Compressor 200 includes multiple compressor stages. Inlet guide vanes 255 axially precede the compressor stages.
The combustor 300 includes one or more fuel injectors 310 and includes one or more combustion chambers 390.
The turbine 400 includes a turbine rotor assembly 410 and turbine nozzles 450. The turbine rotor assembly 410 mechanically couples to the shaft 120. As illustrated, the turbine rotor assembly 410 is an axial flow rotor assembly. The turbine rotor assembly 410 includes one or more turbine disk assemblies 420. Each turbine disk assembly 420 includes a turbine disk that is circumferentially populated with turbine blades. A turbine nozzle 450, such as a nozzle ring, axially precedes each of the turbine disk assemblies 420. Each turbine nozzle 450 includes multiple nozzle segments 451 grouped together to form a ring. Each turbine disk assembly 420 paired with the adjacent turbine nozzle 450 that precede the turbine disk assembly 420 is considered a turbine stage. Turbine 400 includes multiple turbine stages.
The turbine 400 may also include a turbine housing 430 and turbine diaphragms 440. Turbine housing 430 may be located radially outward from turbine rotor assembly 410 and turbine nozzles 450. Turbine housing 430 may include one or more cylindrical shapes. Each nozzle segment 451 may be configured to attach, couple to, or hang from turbine housing 430. Each turbine diaphragm 440 may axially precede each turbine disk assembly 420 and may be adjacent a turbine disk. Each turbine diaphragm 440 may also be located radially inward from a turbine nozzle 450. Each nozzle segment 451 may also be configured to attach or couple to a turbine diaphragm 440.
The exhaust 500 includes an exhaust diffuser 510 and an exhaust collector 520. The power output coupling 600 may be located at an end of shaft 120.
Upper shroud 452 may also include upper forward rail 454 and upper aft rail 455. Upper forward rail 454 extends radially outward from upper endwall 453. In the embodiment illustrated in
Upper aft rail 455 may also extend radially outward from upper endwall 453. In the embodiment illustrated in
Lower shroud 456 is located radially inward from upper shroud 452. Lower shroud 456 may also be located adjacent and radially outward from turbine diaphragm 440 when nozzle segment 451 is installed in gas turbine engine 100. Lower shroud 456 includes lower endwall 457. Lower endwall 457 is located radially inward from upper endwall 453. Lower endwall 457 may be a portion of an annular shape, such as a sector. For example, the sector may be a portion of a nozzle ring. Multiple lower endwalls 457 are arranged to form the annular shape, such as a toroid, and to define the radially inner surface of the flow path through a turbine nozzle 450. Lower endwall 457 may be coaxial to upper endwall 453 and center axis 95 when installed in the gas turbine engine 100.
Lower shroud 456 may also include lower forward rail 458 and lower aft rail 459. Lower forward rail 458 extends radially inward from lower endwall 457. In the embodiment illustrated in
Lower aft rail 459 may also extend radially inward from lower endwall 457. In the embodiment illustrated in
Airfoil 460 extends between upper endwall 453 and lower endwall 457. Airfoil 460 includes leading edge 461, trailing edge 462, pressure side wall 463, and suction side wall 464. Leading edge 461 extends from upper endwall 453 to lower endwall 457 at the most upstream axial location where highest curvature is present. Leading edge 461 may be located near upper forward rail 454 and lower forward rail 458. Trailing edge 462 may extend from upper endwall 453 axially offset from and distal to leading edge 461, adjacent the axial end of upper endwall 453 opposite the location of leading edge 461 and from lower endwall 457 adjacent the axial end of upper endwall 453 opposite and axially distal to the location of leading edge 461. When nozzle segment 451 is installed in gas turbine engine 100, leading edge 461, upper forward rail 454, and lower forward rail 458 may be located axially forward and upstream of trailing edge 462, upper aft rail 455, and lower aft rail 459. Leading edge 461 may be the point at the upstream end of airfoil 460 with the maximum curvature and trailing edge 462 may be the point at the downstream end of airfoil 460 with maximum curvature. In the embodiment illustrated in
Pressure side wall 463 may span or extend from leading edge 461 to trailing edge 462 and from upper endwall 453 to lower endwall 457. Pressure side wall 463 may include a concave shape. Suction side wall 464 may also span or extend from leading edge 461 to trailing edge 462 and from upper endwall 453 to lower endwall 457. Suction side wall 464 may include a convex shape. Leading edge 461, trailing edge 462, pressure side wall 463 and suction side wall 464 may contain a cooling cavity 469 (partially shown in
Airfoil 460 includes multiple cooling holes or apertures, such as pressure side cooling apertures 466, suction side cooling apertures 467, and showerhead cooling apertures 465. Each cooling hole or aperture may be a channel extending through a wall of the airfoil 460. Each set of cooling apertures may be grouped together in a pattern, such as in a row or in a column.
Airfoil 460 may further include slots 468. Slots 468 may be located on pressure side wall 463 and may be adjacent trailing edge 462. Slots 468 may be rectangular and may be aligned in the radial direction between upper endwall 453 and lower endwall 457. Slots 468 may extend from cooling cavity 469 to trailing edge 462.
In the embodiment illustrated in
The various components of nozzle segment 451 including upper shroud 452, lower shroud 456, airfoil 460, and second airfoil 470 may be integrally cast or metalurgically bonded to form a unitary, one piece assembly thereof.
One or more of the above components (or their subcomponents) may be made from stainless steel and/or durable, high temperature materials known as “superalloys”. A superalloy, or high-performance alloy, is an alloy that exhibits excellent mechanical strength and creep resistance at high temperatures, good surface stability, and corrosion and oxidation resistance. Superalloys may include materials such as HASTELLOY, alloy x, INCONEL, WASPALOY, RENE alloys, HAYNES alloys, alloy 188, alloy 230, INCOLOY, MP98T, TMS alloys, Mar, M247, and CMSX single crystal alloys.
As illustrated in
Base 710, pedestal 730, and shield 720 may be coaxial and may revolve about mounting stand axis 701. In the embodiment illustrated in
In the embodiment illustrated in
In the embodiment illustrated, the bore 715 extends through pedestal 730 and through base 710. In other embodiments, bore 715 may not extend through base 710. Bore 715 may be sized to a bore diameter 708 to receive a temperature sensor 705 at the end distal to base 710. Bore diameter 708 may be sized and configured so that temperature sensor 705 is flush with the bore 715. Temperature sensor 705 may be within 1.27 millimeters (0.050 inches) from the end of pedestal 730 distal to base 710. Temperature sensor 705 may be secured to pedestal 730 with a bonding agent, such as a high temperature ceramic adhesive. The bonding agent may include an adhesive suitable for high temperature and high pressure applications and may have thermal expansion coefficients similar to the material of the nozzle segment 451.
Pedestal 730 includes a pedestal end 737 distal to base 710. In the embodiment illustrated in
Pedestal 730 may include a pedestal thickness 738, the radial thickness of pedestal 730 between pedestal inner surface 733 and pedestal outer surface 734. In one embodiment, pedestal thickness 738 is at least 0.381 millimeters (0.015 inches). In another embodiment, pedestal thickness 738 is from 0.381 millimeters (0.015 inches) to 1.02 millimeters (0.040 inches). In other embodiments, pedestal thickness 738 is up to 1.02 millimeters (0.040 inches).
In the embodiment illustrated in
Shield 720 includes a shield end 727 distal to base 710. In one embodiment, shield 720 extends axially beyond pedestal 730 at an entrance length 729, the axial distance between shield end 727 and pedestal end 737, at least 1.9 millimeters (0.075 inches). In another embodiment, shield 720 extends axially beyond pedestal 730 at an entrance length 729 from 1.9 millimeters (0.075 inches) to 5.08 millimeters (0.2 inches). In some embodiments, shield 720 extends axially beyond pedestal 730 at an entrance length 729 up to 2.80 millimeters (0.11 inches). In other embodiments, shield 720 extends axially beyond pedestal 730 at an entrance length 729 up to 5.08 millimeters (0.2 inches). In yet other embodiments, shield 720 extends axially beyond pedestal 730 at an entrance length 729 of 2.54 millimeters (0.10 inches).
Shield 720 may include a shield thickness 728, the radial thickness of shield 720 between shield inner surface 723 and shield outer surface 724. In one embodiment, shield thickness 728 is at least 0.635 millimeters (0.025 inches). In another embodiment, shield thickness 728 is from 0.635 millimeters (0.025 inches) to 1.905 millimeters (0.075 inches). In other embodiments, shield thickness 728 is up to 1.905 millimeters (0.075 inches).
Shield 720 may include multiple aspiration holes, such as a first aspiration hole 721 and a second aspiration hole 722. Each aspiration hole may extend through the hollow cylinder shape of shield 720 from shield inner surface 723 to shield outer surface 724 and may be proximal base 710. In some embodiments, such as the embodiment in
Shield 720 is located radially outward from pedestal 730. Shield 720 and pedestal 730 may be spaced apart at an offset distance 718, the distance between shield inner surface 723 and pedestal outer surface 724, forming a flow region 735 there between. Flow region 735 is the annular space between shield 720 and pedestal 730 formed by shield inner surface 723 and pedestal outer surface 724. In one embodiment, the offset distance 718 is up to 1.27 millimeters (0.050 inches). In another embodiment, the offset distance 718 is from 0.635 millimeters (0.025 inches) to 1.27 millimeters (0.050 inches). In other embodiments, the offset distance 718 is at least 0.635 millimeters (0.025 inches).
Referring again to
Referring to
Temperature sensor 705 may be made into various shapes, such as a cylinder, cube, sphere, etc., and have dimensions at the micrometer scale. Temperature sensor 705 may be suitable for measuring temperatures in a wide temperature range and in high temperature and high pressure environments, such as the conditions experienced within a gas turbine engine during operation. For example, temperature sensor 705 may be configured to measure temperatures between 150 degrees Celsius and 1450 degrees Celsius.
Temperature sensor 705 may be made of an irradiated crystal, such as silicon carbide or Izmeritel Maximalnoi Temperaturi Kristalincheskii (IMTK) crystal. Temperature sensor 705 may record and provide temperature information through microstructural changes without the need for wires. The microstructural changes may be deformations to crystal lattice structures. Temperature sensor 705 may microstructurally change as a function of the temperature surrounding the temperature sensor 705. The temperature sensed by temperature sensor 705 may be read wirelessly by an external receiver/detector, such as an X-ray defractometry, and may be converted to temperature data.
In some embodiments, temperature sensor 705 is configured to retain the microstructural changes and hence the information of the temperature. In these embodiments, temperature sensor 705 is removed from pedestal 730 to collect the temperature information. In other embodiments, temperature sensor 705 may partially transform from solid to liquid at a particular temperature. The temperature may be determined by checking/detecting the phase change of temperature sensor 705.
In the embodiments illustrated in
In the embodiment illustrated in
In some two piece embodiments, such as the embodiment shown in
The various components, shapes, and sizes of mounting stand 800, such as shield 820, first bleed hole 821, second bleed hole 822, shield entrance length 829, flow region 835, pedestal thickness 838, outer edge chamfer 812, bore 815, bore diameter 808, offset distance 818, flow region 835, first base surface 813, second base surface 814, pedestal inner surface 833, pedestal outer surface 834, shield inner surface 823, shield outer surface 824, pedestal end 837, shield end 827, and mounting stand axis 801 may be the same or similar as those described in conjunction with the mounting stand 700 of
Each mounting stand 700 and 800 and its various components may be made from a ceramic material resistant to high temperatures, such as alumina, zirconium oxide, or silicon carbide.
Gas turbine engines may be suited for any number of industrial applications such as various aspects of the oil and gas industry (including transmission, gathering, storage, withdrawal, and extraction of oil and natural gas), the power generation industry, cogeneration, aerospace, and other transportation industries.
Referring to
Once compressed air 10 leaves the compressor 200, it enters the combustor 300, where it is diffused and fuel is added. Air 10 and fuel are injected into the combustion chamber 390 via fuel injector 310 and combusted. Energy is extracted from the combustion reaction via the turbine 400 by each stage of the series of turbine disk assemblies 420. Exhaust gas 90 may then be diffused in exhaust diffuser 510, collected and redirected. Exhaust gas 90 exits the system via an exhaust collector 520 and may be further processed (e.g., to reduce harmful emissions, and/or to recover heat from the exhaust gas 90).
Operating efficiency and output power of a gas turbine engine generally increases with a higher combustion temperature. Thus, there is a trend in gas turbine engines to increase the combustion temperatures. Gas reaching forward stages of a turbine from a combustion chamber 390 may be 1000 degrees Fahrenheit or more. It may be desirable to measure the temperature of the combustion gases within the gas turbine engine including upstream of the first stage turbine nozzle 450.
While measuring the temperature of the combustion gases along the leading edge 461 of a nozzle segment 451, errors in the measurements may be introduced from conduction, radiation, and convection. Conduction from the nozzle segment 451 through the pedestal 730 may affect the temperature measurement. Providing a pedestal 730 formed of a ceramic material, such as alumina, zirconium oxide, or silicon carbide may reduce the conduction. Extending the pedestal 730 beyond the leading edge 461 to the extruding length 704 may further reduce or prevent the conduction error as the conduction error may correlate to the pedestal length 739. The improvements in conduction error over the length of the pedestal 730 may get incrementally smaller as the length increases. The pedestal length 739 may be increased up to the point where the increase in material cost and potential for durability failure outweighs the incrementally smaller improvement in the conduction error.
The radiation error may be reduced surrounding the pedestal 730 and in particular the temperature sensor 705 with shield 720. A reduction in the radiation error may correlate to the entrance length 729, the distance that shield 720 extends beyond pedestal 730. The entrance length 729 may be increased by increasing the length of shield 720 up to the point where the increase in material cost and potential for durability failure outweighs the improvement in the radiation error.
The convection error may be reduced by increasing the flow area 736 defined by the offset distance 718 between shield 720 and pedestal 730 and/or by decreasing the size of aspiration holes 721 and 722. The flow area 736 can be varied until the flow through the flow region 735 begins to stagnate. Increasing the flow area 736 and decreasing the size of aspiration holes 721 and 722 may reduce the temperature error and may promote a time accurate reading.
Use of a ceramic material may reduce the cost of the mounting stand 700 over the use of a high temperature metal alloy. Use of the two piece mounting stand 800 may further reduce the cost.
Use of the single integral piece mounting stand 700 may reduce or prevent any tolerance issues between pieces and may improve the strength of the part as it may not require any bonds.
The preceding detailed description is merely exemplary in nature and is not intended to limit the invention or the application and uses of the invention. The described embodiments are not limited to use in conjunction with a particular type of gas turbine engine. Hence, although the present disclosure, for convenience of explanation, depicts and describes a particular mounting stand for a temperature sensor, it will be appreciated that the mounting stand in accordance with this disclosure can be implemented in various other configurations, can be used with various other types of gas turbine engines, and can be used in other types of machines. Furthermore, there is no intention to be bound by any theory presented in the preceding background or detailed description. It is also understood that the illustrations may include exaggerated dimensions to better illustrate the referenced items shown, and are not consider limiting unless expressly stated as such.
