Patent Application
20140060001
The invention relates to industrial gas turbine engines with can annular combustion assemblies configured in a manner that enables a shorter rotor shaft, and a radial diffuser that improves aerodynamics in the shortened gas turbine engines.
Industrial gas turbines are used primarily to generate electrical power, where other gas turbine engines may have other primary purposes. For example, aero gas turbine engines are designed to be lightweight, and as small as possible in order to provide propulsion for aircraft. Aeroderivative gas turbine engines are aero gas turbine engines that have been modified to generate electrical power. Due to their original aero purpose, the aeroderivative engines are lighter than the industrial gas turbine engines, and therefore portable, but they are less robust and generate less electrical power. With little need to be light weight, portable, or aerodynamic, industrial gas turbines are generally made of heavy duty components with a primary concern being long engine life and electrical power output. This often results in an industrial gas turbine engine that is heavier and bulkier than aero or aeroderivative counterparts. This bulk yields longer engine life and greater output capacity, but also creates design and maintenance complexity and cost.
The invention is explained in the following description in view of the drawings that show:
The present inventor has identified a way to shorten a rotor length in an industrial gas turbine engine using a can annular combustor system. The can combustors of a can annular combustion system can be reconfigured to a position more radially outward, axially closer to the turbine, and having a combustor longitudinal axis at a lesser angle with respect to a plane defined by a turbine inlet annulus. By reorienting the combustor cans in this manner, the diameter of the combustion assembly (which includes all of the combustors and the structures between the combustors and the turbine) increases. The present inventor has recognized that a length along an axis of the engine taken up after reorienting the combustion assembly may be decreased in comparison to a length taken up using traditionally oriented can annular combustors. This decrease in combustor section length, and associated reduction in engine length, can be significant. For example, in smaller industrial gas turbine engines having engine lengths of 5-6 meters from a leading edge of a first row of compressor airfoils (either vanes or blades, whichever is first) to a trailing edge of a last row of turbine airfoils (either vanes or blades, whichever is last), the reduction in axial length of the combustion assembly, and therefore the engine length, may be about ½ meter in certain emerging technology designs. In larger industrial gas turbine engines having engine lengths of 10-12 meters, the reduction in axial length may be about 1 meter. Other engine sizes, including industrial gas turbine engines having an engine length on the order of less than two meters will experience a comparable length reduction as well. As used herein, a combustor section length is a length between a trailing edge of a last row of compressor airfoils and a leading edge of first row of turbine blades. In an industrial gas turbine engine with more than one turbine, it is the first row of turbine blades of the first turbine that is being referred to herein. The first row of stationary vanes upstream of and adjacent to the first row of rotating turbine blades is considered herein to be part of the combustion section.
The focus on heavy duty components in industrial gas turbine engines has resulted in long, heavy rotor shafts and associated bearings. As rotor shaft length increases, so do rotor shaft dynamics. Increased rotor dynamics require increasingly complex rotor shaft design as well as bearings sized to handle the rotor shafts. Consequently, any reduction in rotor shaft size will result in smaller rotor shafts, less complex rotor shaft designs, smaller bearings, and smaller clearances at blade tips, each of which reduce cost and complexity. The result of applying the teachings herein is an industrial gas turbine engine that is shorter but which retains longevity and power output yet will cost less to build and maintain.
It is further recognized that compressed air aerodynamics, and thus engine performance, can be improved by utilizing a radial diffuser to direct compressed air exiting the compressor from an axial travel direction to a more radial travel direction. The radial diffuser may be particularly useful in one emerging combustor technology design for can annular gas turbine engines that includes structures that direct combustion gas from combustion to a first row of turbine blades without a need for a row of vanes at the end of the structure (i.e. upstream of the first row of turbine blades), where the structures properly orient and accelerate the combustion gas. Each flow directing structure includes a combustor can and a respective flow duct that directs combustion gas from the combustor to the first row of turbine blades along a straight flow path and at a proper speed and orientation without a first row of vanes. The combustion assembly includes all of the flow directing structures, one for each point of combustion. One such combustion assembly is disclosed in U.S. Pat. No. 7,721,547 to Bancalari et al. issued May 25, 2010, incorporated in its entirety herein by reference. Another such combustion assembly that also includes an annular chamber immediately upstream of the first row of turbine blades is disclosed in U.S. Pre-Grant Publication Number 2010/0077719 to Wilson et al., filed Apr. 8, 2009, which is incorporated in its entirety herein by reference. In particular, because the combustor inlet is disposed so much farther radially outward in the reoriented configuration, and because the compressor outlet is disposed closer to an aft wall of the compressor section in this emerging combustor technology, a radial diffuser may provide a dramatic improvement in aerodynamic performance. Any industrial gas turbine engine utilizing such an emerging technology combustor can benefit from the radial diffuser. This includes smaller industrial gas turbine engines having an engine length of less than two meters and a rated electrical power output of less than 1 MW to those having an engine length of over twelve meters and a rated electrical power output of over 100 MW.
The conventional combustion section 14 includes a combustion assembly 40 including individual combustor cans 42 and respective conventional transition ducts 44 configured to receive combustion gas from the respective combustor cans 42 and deliver it to the turbine section 16. The conventional combustion section 14 also includes a plenum 46 defined by a conventional rotor combustion section casing 48, where the plenum 46 receives diffused compressed air from the diffuser 32, and acts as a pressure vessel of sorts to contain the diffused compressed air as it works its way toward a combustor inlet 50 of each combustor can 42. As used herein, a conventional combustion section length 52 spans from the trailing edge 28 of (a base of) the last row of compressor airfoils, (be they compressor vanes 20 or compressor blades 22) to a leading edge 54 of (a base of) a first row 56 of turbine blades. The conventional combustion section length 52 includes a row of vanes 58 at the end of the conventional transition duct 44, upstream of and adjacent to the first row 56 of turbine blades.
The turbine section 16 includes turbine vanes 60 and turbine blades 62. As used herein, a turbine section length 64 spans from the leading edge 54 of (a base of) the first row 56 of turbine blades to a trailing edge 66 of (a base of) the last row of turbine airfoils (be they turbine vanes 60 or turbine blades 62).
In the prior art industrial gas turbine engine shown it can be seen that the conventional combustion section length 52 is approximately 23% of a conventional engine length 68 which, as used herein, spans from the leading edge 26 of the first row of compressor airfoils to the trailing edge 66 of the last row of turbine airfoils. For example, a gas turbine engine with a 350 MW output may have a length of 10 meters between bearings, where a first bearing is disposed proximate the first row of compressor airfoils and the second bearing is disposed proximate the last row of turbine airfoils. The inventor is unaware of any prior art industrial gas turbine engine using a can annular combustion arrangement about a single rotor shaft and which produces at least 75 MW where the conventional combustion section length 52 is less than approximately 23% of the conventional engine length 68. (As used herein, a can annular combustion arrangement about a single rotor shaft refers to industrial gas turbine engines not having concentric rotor shafts between the first row of compressor airfoils and the last row of turbine airfoils.) The inventor proposes to reduce that percentage to not more than 20%. For a given industrial gas turbine engine 10, the inventor proposes to reduce the engine length by up to 8% to 10% by moving the conventional combustor cans to a more radial position. For example, in the 350 MW gas turbine engine originally having a 10 meter length between the bearings, the inventor proposed to reduce the length from 10 meters to 9 meters. In one exemplary embodiment described in further detail herein the inventors replace the conventional combustion assembly 40 with the emerging technology (a.k.a. reconfigured) combustion assembly but leaving the compressor section 12 and the turbine section 16 the same in relevant part. When using the emerging technology combustion assembly the 350 MW gas turbine engine originally having a 10 meter length between the bearings may have its length be further reduced from 10 meters to 8.8 meters.
The conventional combustion assembly 40 takes up a conventional combustion system axial length 74 (from a fore end of the combustor inlet 50 to the leading edge 54 of the first row 56 of turbine blades), along a gas turbine engine longitudinal axis 76. It can be seen that that the conventional combustion system axial length 74 is almost as long as the conventional combustion section length 52 and has a great degree of influence on the magnitude of the conventional combustion section length 52. It can also be seen that the row of vanes 58 at the end of the conventional transition duct 44 occupy a vane length 78 which the conventional rotor shaft 18 must accommodate.
Also visible in
The inventor has recognized that the orientation of the combustor cans 52 in the reconfigured combustor assembly 80 is more radially outward, closer to the turbine, and with a smaller angle β between a combustor longitudinal axis 87 and the plane defined by the turbine inlet annulus 72. This smaller angle β, from 35 degrees and below, results in the reconfigured combustion assembly 80 with a reconfigured combustion system axial length 96. It can be seen that the reconfigured combustion system axial length 96 takes up a significantly smaller portion of the conventional combustion section length 52 (shown as a dotted line) than did the conventional combustion system axial length 74 of the prior art. This leaves a remnant length 98 of the conventional combustion section length 52. The reconfigured combustion assembly is not limited to that which is shown in
It is expected that this configuration would operate with the conventional rotor shaft 18 and combustor section 48 casing without any problems brought on by the reconfigured combustion assembly 80. However, the remnant length 98 and the vane length 78 are lengths that contribute to a conventional engine length 68 that the conventional rotor shaft 18 must accommodate. (Although it is known that in certain prior art gas turbine engines the conventional rotor shaft 18 may extend beyond the compressor section 12 and the turbine section 16, for purposes simplicity in explanation, as used herein the conventional rotor engine length equals the conventional engine length 68.) The present inventor has recognized that if the remnant length 98 and/or the vane length 78 can be removed from the design, the conventional rotor shaft 18, the conventional combustion section 14, and the conventional combustion section casing 48 can be shortened, and this will reduce the rotor shaft dynamics, and associated design, manufacturing, and maintenance costs.
The reconfigured combustion section 110 necessarily moves the diffuser 32 axially closer to an aft side 116 of the reconfigured combustion section 110. Consequently, compressed air exiting a diffuser outlet 118 is likely to continue traveling somewhat axially until encountering an obstruction such as the IEP 86 or the reconfigured combustion section aft side 116 itself. Combustion efficiency is very dependent on a smooth, predictable, and efficient flow of compressed air through the plenum 46. Consequently, any obstructions introduce turbulence, local variations in pressure, and pressure loss, each of which reduce combustion efficiency and may increase harmful emissions.
While it is believed that the gas turbine engine would operate with the diffuser 32 so located, in one exemplary embodiment the inventor has applied a radial diffuser wall 130 configured to receive axially flowing compressed air exiting the diffuser outlet 118 and redirect it in a more radial direction. The radial diffuser wall 130 may direct the compressed air anywhere from slightly radially, to a destination radially outside of the annular chamber 85, to perpendicular to the gas turbine engine longitudinal axis 76. The radial diffuser wall 130 may even direct the flow of compressed air more than 90 degrees such that it flows radially outward and backwards with respect to the axial flow exiting the diffuser outlet 118, such that direction of travel is toward the compressor section 12 as well as radially outward. In this manner the radial diffuser wall 130 can direct circumferential portions of the flow of compressed air directly into the top hat openings 92. In an exemplary embodiment shown by a dashed line, the radial diffuser wall 130 directs the flow of compressed air parallel to the combustor longitudinal axis 87. In an exemplary embodiment the radial diffuser wall 130 may be a single sheet and may undulate axially along its circumference from between adjacent IEP's 86 to upstream of the IEP's 86. In another exemplary embodiment the radial diffuser wall may be disposed primarily between the adjacent IEP's 86 and have openings or simply not extend proximate the IEP's 86.
Further, a combination of the curved diffuser and the conical diffuser may be used. For example, one wall may be curved and another straight, or one or both walls may include a curved and/or a straight section. Various exemplary embodiments using straight and/or curved walls may be used without limitation so long as the compressed air is redirected from essentially axial to more radially outward.
The shortened industrial gas turbines disclosed herein may be used in a wide variety of applications. One application includes being part of an industrial gas turbine engine assembly where the shortened gas turbine engine is used in conjunction with a free power turbine. In such cases exhaust from the turbine of the shortened industrial gas turbine engine disclosed herein may be directed into the free power turbine. Remaining energy present in the exhaust from the turbine of the shortened industrial gas turbine engine is used to turn the free power turbine. As a result the free power turbine, which is not mechanically connected to the rotor shaft of the shortened industrial gas turbine engine, is able to convert the remaining energy into electrical power, and this, in turn, increases the amount of electrical power that can be extracted from the combustion gas which might otherwise go unutilized.
While various embodiments of the present invention have been shown and described herein, it will be obvious that such embodiments are provided by way of example only. Numerous variations, changes and substitutions may be made without departing from the invention herein. Accordingly, it is intended that the invention be limited only by the spirit and scope of the appended claims.
