Turbomachine and turbine blade for this

Abstract

Rotor blade (36) comprising a blade (37), wherein the rotor blade (36) is configured for use with a turbomachine (10), the blade (37) comprising: a constriction distribution (D0), measured in a narrowest region in a passage (38) between adjacent rotor blades (36), in which adjacent rotor blades (36) extend transversely across the passage (38) between opposing walls (40, 42) to interact aerodynamically with a fluid flow; and wherein the constriction distribution (D0) is defined by a trailing edge (46) of the blade (37), wherein the constriction distribution (D0) is configured to reduce aerodynamic losses and improve aerodynamic loads on the blade (37);characterized in that the constriction distribution (D0) extends essentially linearly from a constriction / constriction mid-span value of approximately 82% at approximately 5% span to a constriction / constriction mid-span value of approximately 115% at approximately 90% span, a constriction / constriction mid-span value of approximately 110% at approximately 95% span, and a constriction / constriction mid-span value of approximately 82.5% at approximately 100% span; and the span at 0% is located on a radially inner section of the airfoil (37), while a span at 100% is located on a radially outer section of the airfoil (37), and the constriction / constriction mid-span value is 100% at approximately 50% to 55% span.

Description

BACKGROUND TO THE INVENTION The subject matter disclosed herein relates to turbomachinery and in particular to a rotor blade in a turbomachine. A turbomachine, such as a gas turbine, can contain a compressor, a combustion chamber, and a turbine. In the compressor, air is compressed. The compressed air is fed into the combustion chamber. The combustion chamber combines a fuel with the compressed air and then ignites the gas / fuel mixture. The exhaust gases, which have a high temperature and high energy, are then fed to the turbine, where the energy of the fluids is converted into mechanical energy. The turbine contains several stages of guide vanes and stages of rotor blades. The guide vanes are stationary components, and the rotor blades rotate on a rotor. US 2014 / 0041395A1 discloses a running blade with the features of the preamble of claim 1. Based on this, it is an object of the invention to further develop a rotor blade for use in a turbomachine and a turbomachine with such a rotor blade in such a way that aerodynamic losses can be reduced and aerodynamic loads on the blade of the rotor blade can be improved. This problem is solved by a rotor blade having the features of independent claim 1 and a turbomachine according to claim 9. Particularly preferred embodiments of the invention are the subject of the dependent claims. BRIEF DESCRIPTION OF THE INVENTION Certain embodiments that correspond to the scope of the originally claimed subject matter are briefly summarized below. These embodiments are not intended to limit the scope of protection of the claimed subject matter, but rather to provide a brief description of possible forms of the claimed subject matter. In fact, the claimed subject matter can take many forms that may be similar to or differ from the aspects / embodiments described below. In a first aspect of the invention, a rotor blade comprises a fan blade, and the rotor blade is configured for use with a turbomachine. The fan blade incorporates a constriction distribution, measured in a narrowest region in a passage between adjacent rotor blades, where adjacent rotor blades extend transversely across the passage between opposing walls to interact aerodynamically with a fluid flow. The constriction distribution is defined by a trailing edge of the fan blade, and the constriction distribution is configured to reduce aerodynamic losses and improve aerodynamic loads on the fan blade.The constriction distribution extends essentially linearly from a constriction / constriction mid-span value of approximately 82% at approximately 5% span to a constriction / constriction mid-span value of approximately 115% at approximately 90% span, a constriction / constriction mid-span value of approximately 110% at approximately 95% span, and a constriction / constriction mid-span value of approximately 82.5% at approximately 100% span. The span at 0% is located on a radially inner section of the airfoil, while a span at 100% is located on a radially outer section of the airfoil, and the constriction / constriction mid-span value is 100% at approximately 50% to 55% span. In particular, the bottleneck / bottleneck mid-span value can be 100% at approximately 54% of the span. In preferred embodiments of any of the aforementioned rotor blades, the constriction distribution can be defined by values ​​specified in Table 1, wherein the constriction distribution values ​​can be within a tolerance of + / -10% of the values ​​specified in Table 1. Additionally or as an alternative, the trailing edge of the blade can have a projection at approximately 50% of its span. Additionally, or as a further alternative, the trailing edge of the blade can have an offset of approximately 0 at approximately 0% span, approximately 100% at approximately 50% span, and approximately 0 at approximately 100% span. In some preferred embodiments of any of the aforementioned rotor blades, a trailing edge of the blade may have an offset as defined by the values ​​given in Table 2. Additionally or as an alternative, the blade can have a thickness distribution (Tmax / Tmax_midspan) as defined by the values ​​given in Table 3. Additionally, or as a further alternative, the blade can have a dimensionless thickness distribution according to the values ​​given in Table 4. In addition, or as a further alternative, the blade can have a dimensionless axial chord distribution according to the values ​​given in Table 5. In another aspect of the invention, a turbomachine is created which includes opposing walls defining a passage in which a fluid flow can be received in order to flow through the passage, and several impeller blades according to the first aspect as described herein. BRIEF DESCRIPTION OF THE DRAWINGS These and other features, aspects, and advantages of the present disclosure will be better understood if the following detailed description is read with reference to the accompanying drawings, in which the same reference numerals denote the same parts throughout the drawings, wherein: Fig. 1 shows a schematic representation of a turbomachine according to aspects of the present disclosure; Fig. 2 shows a perspective view of a rotor blade according to aspects of the present disclosure; Fig. 3 shows a top view of two adjacent rotor blades according to aspects of the present disclosure; Fig. 4 shows a graphical representation of the constriction distribution according to aspects of the present disclosure; Fig. 5 shows a graphical representation of a trailing edge offset according to aspects of the present disclosure; Fig. 6 shows a graphical representation of a maximum thickness distribution according to aspects of the present disclosure; Fig.7 a graphical representation of the maximum thickness, divided by the axial chord distribution, according to aspects of the present disclosure; and Fig. 8 a graphical representation of the axial chord, divided by the axial chord at the midspan, according to aspects of the present disclosure. DETAILED DESCRIPTION OF THE INVENTION One or more specific embodiments of the present disclosure are described below. In an effort to provide a concise description of these embodiments, the description may not include all features of an actual implementation. It should be recognized that, as in any engineering or design project, the development of any such actual implementation requires numerous implementation-specific decisions to be made in order to achieve the specific goals of the developers, such as compliance with system-related and company-related constraints, which may vary from one implementation to another.Furthermore, it should be recognized that while such a development effort may be complex and time-consuming, for experts in the field who benefit from this disclosure it would nevertheless represent a routine undertaking for design, manufacture and production. When elements of different embodiments of the present object are introduced, the articles "a", "an", and "the" or "a" shall mean that one or more of the elements may be present. The expressions "have", "contain", and "have" shall be meant inclusively and mean that, in addition to the elements listed, further elements may be present. Fig. 1 shows a diagram of an embodiment of a turbomachine 10 (e.g., a gas turbine and / or a compressor). The turbomachine 10, as illustrated in Fig. 1, comprises a compressor 12, a combustion chamber 14, a turbine 16, and a diffuser 17. Air, or any other gas, is compressed in the compressor 12, fed into the combustion chamber 14, mixed with a fuel, and subsequently combusted. The exhaust fluids are fed to the turbine 16, where energy from the exhaust fluids is converted into mechanical energy. The turbine 16 comprises several stages 18, including a single stage 20. Each stage 18 comprises a rotor (i.e., a rotating shaft) with an annular arrangement of axially aligned rotor blades rotating about an axis of rotation 26, and a stator with an annular arrangement of guide vanes. Accordingly, the stage 20 may comprise a guide vane stage 22 and a rotor blade stage 24.For clarity, Fig. 1 includes a coordinate system comprising an axial direction 28, a radial direction 32, and a circumferential direction 34. A radial plane 30 is also illustrated. The radial plane 30 extends in the axial direction 28 (along the axis of rotation 26) in one direction and then extends outwards in the radial direction 32. Fig. 2 shows a perspective view of a rotor blade 36. The rotor blades 36 in stage 20 extend in a radial direction 32 between a first wall (or platform) 40 and a second wall 42. The first wall 40 faces the second wall 42, and both walls define a path or passage through which a fluid flow can be received. The rotor blades 36 are arranged circumferentially 34 around a hub. Each rotor blade 36 has a blade 37, and the blade 37 is configured to interact aerodynamically with the exhaust fluids from the combustion chamber 14, while the exhaust fluids flow substantially downstream through the turbine 16 in the axial direction 28. Each impeller blade 36 has a leading edge 44, a trailing edge 46 which is arranged downstream of the leading edge 44 in the axial direction 28, a pressure side 48 and a suction side 50.The pressure side 48 extends in the axial direction 28 between the leading edge 44 and the trailing edge 46, and in the radial direction 42 between the first wall 40 and the second wall 42. The suction side 50 extends in the axial direction 28 between the leading edge 44 and the trailing edge 46, and in the radial direction 32 between the first wall 40 and the second wall 42, opposite the pressure side 48. The rotor blades 36 in stage 20 are arranged such that the pressure side 48 of one rotor blade 36 faces the suction side 50 of an adjacent rotor blade 36. As the exhaust fluids flow to and through the passage between the rotor blades 36, they interact aerodynamically with the rotor blades 36 such that they flow with an angular momentum or velocity relative to the axial direction 28.A rotor blade stage 24, equipped with rotor blades 36 featuring a special constriction distribution designed to exhibit reduced aerodynamic losses and improved aerodynamic loads, can result in improved machine efficiency and improved part durability. The mounting section 39 of the rotor blade 36 is shown with dashed lines and may include a dovetail section, angel wing seals, or other features as may be desired in the specific embodiment or application. Fig. 3 shows a top view of two adjacent rotor blades 36. Note that the suction side 50 of the lower rotor blade 36 faces the pressure side 48 of the upper rotor blade 36. The axial chord 56 is the dimension of the rotor blade 36 in the axial direction 28. The chord 57 is the distance between the leading and trailing edges of the blade. The passage 38 between two adjacent rotor blades 36 of a stage 18 defines a constriction distribution Do, which is measured in the narrowest part of the passage 38 between adjacent rotor blades 36. A fluid flows through the passage 38 in the axial direction 28. This constriction distribution Do over the span from the first wall 40 to the second wall 42 is explained in greater detail below with reference to Fig. 4. The maximum thickness of each rotor blade 36 at a given percentage span is illustrated as Tmax.The Tmax distribution over the height of the rotor blade 36 is explained in greater detail below with reference to Fig. 4. Fig. 4 shows a graphical representation of the constriction distribution Do, defined by adjacent rotor blades 36 and depicted as a curve 60. The vertical axis 62 represents the percentage span between the first annular wall 40 and the second annular wall 42, or the opposite end of the blade 37, in the radial direction 32. That is, 0% span essentially denotes the first annular wall 40, and 100% span denotes the opposite end of the blade 37, and any point between 0% and 100% corresponds to a percentage of the distance between the radially inner and radially outer parts of the blade 37 in the radial direction 32 along the height of the blade.The horizontal axis 64 represents Do(narrowness), the shortest distance between two adjacent rotor blades 36 at a given percentage span, divided by Do_midspan(narrowness_midspan), which is Do at approximately 50% to approximately 55% of the span. Dividing Do by Do_midspan makes the graphical representation 58 dimensionless, so that the curve 60 remains the same when the rotor blade stage 24 is scaled up or down for different applications. A similar graphical representation could be created for a single turbine size, in which the horizontal axis is Do. As can be seen in Fig. 4, the constriction distribution, as defined by the trailing edge of the rotor blade, is essentially linear, progressing from a constriction / constriction mid-span value of approximately 82% at about 5% span (point 66) to a constriction / constriction mid-span value of approximately 115% at about 90% span (point 70) and a constriction / constriction mid-span value of approximately 110% at about 95% span. The span at 0% is located on a radially inner part of the blade, and the span at 100% is located on a radially outer part of the blade. The constriction / constriction mid-span value is 115% at about 50% to 55% span (point 68). The constriction distribution illustrated in Fig. 4 can help improve performance in two ways. First, the distribution of constrictions helps to generate desired outlet flow profiles. Second, the one shown in Fig.Figure 4 illustrates the constriction distribution to help manipulate secondary flows (e.g., flows perpendicular to the main flow direction) and / or flushing flows near the first annular wall 40 (e.g., the hub). Table 1 lists the constriction distribution and various values ​​for the trailing edge shape of the blade 37 at several points along the span. Figure 4 shows a graphical representation of the constriction distribution. It should be understood that the constriction distribution values ​​can vary by approximately + / -10%. Table 1 1000,825 951,116 911,155 821,119 731,077 641,039 541,000 440,963 340,928 230,888 120,848 60,827 00.808 Fig. 5 shows a graphical representation of the trailing edge offset of the blade 37 of the rotor blade 36. The trailing edge 46 has a projection 500 at a span of approximately 50%. The vertical axis represents the percentage span between the first annular wall 40 and the opposite end of the blade 37 in the radial direction 32. The horizontal axis represents the trailing edge offset to a straight line extending from a line 510 (see Fig. 2) running from a radially inner section of the trailing edge to a radially outer section of the trailing edge. The projection 500 is greatest at approximately 50% of the span (i.e., 1 or 100%) and then gradually decreases to an offset value of 0 at approximately 0% of the span and approximately 100% of the span.Furthermore, a rotor blade 36 with a trailing edge offset increased by approximately 50% of its span can help to tune the rotor blade's resonant frequency to avoid interference with drive components. If the rotor blade's resonant frequency is not carefully adjusted to avoid interference with drive components, operation may result in impermissible stress on the rotor blade 36 and potential structural failure. Accordingly, a rotor blade 36 design with the projection 500 or the increased trailing edge offset, as illustrated in Fig. 5, can extend the rotor blade 36's service life. Table 2 lists the trailing edge offset and projection shape for various values ​​of the blade's trailing edge 37 at several points along its span. 1000 94,60,116 83,60,332 72,60,567 61,60,821 50,51,000 39,40,918 28,30,660 17,20,284 6,10,030 00 Fig. 6 shows a graphical representation of the thickness distribution Tmax / Tmax_midspan as defined by the thickness of the rotor blade 37. The vertical axis represents the percentage span between the first annular wall 40 and the opposite end of the rotor blade 37 in the radial direction 32. The horizontal axis represents Tmax divided by the Tmax_midspan value. Tmax is the maximum blade thickness at a given span, and Tmax_midspan is the maximum blade thickness at the midspan (e.g., at approximately 50% to 55% of the span). Dividing Tmax by Tmax_midspan makes the graphical representation dimensionless, so the curve remains the same when the rotor blade stage 24 is scaled up or down for different applications.Referring to Table 3, a mid-span value of approximately 53% has a Tmax / Tmax_mid-span value of 1, because at this span Tmax equals Tmax_mid-span. Table 3 1000.91 950.79 910.80 820.83 720.89 630.95 531.00 431.04 321.08 221.11 111.16 61.18 01.22 Fig. 7 shows a graphical representation of the blade thickness (Tmax) divided by the axial chord of the blade at various span values. The vertical axis represents the percentage span between the first annular wall 40 and the opposite end of the blade 37 in the radial direction 32. The horizontal axis represents Tmax divided by the axial chord value. Dividing the blade thickness by the axial chord makes the graphical representation dimensionless, so the curve remains the same when the rotor blade stage 24 is scaled up or down for different applications. A rotor blade design with the Tmax distribution illustrated in Figs. 6 and 7 can help to tune the rotor blade's resonant frequency to avoid interference with drive components. Accordingly, a rotor blade design 36 with the distribution shown in Figs. 6 and 7 can be used to achieve a resonant frequency of the rotor blade 36 to avoid interference with drive components.Figure 7 illustrates how the Tmax distribution increases the service life of the rotor blade 36. Table 4 lists the Tmax / axial chord value for various span values, where the dimensionless thickness is defined as the ratio of Tmax to the axial chord at a given span. 1000,375 950,323 910,326 820,333 720,348 630,361 530,374 430,382 320,390 220,397 110,408 60,415 00.427 Fig. 8 shows a graphical representation of the axial chord of the airfoil, divided by the axial chord value at the mid-span at various span values. The vertical axis represents the percentage span between the first annular wall 40 and the opposite end of the airfoil 37 in the radial direction 32. The horizontal axis represents the axial chord divided by the axial chord at the mid-span value. Referring to Table 5, a mid-span value of approximately 53% has an axial-chord / axial-chord_mid-span value of 1, because at this span the axial chord is equal to the axial chord at the mid-span. Dividing the axial chord by the axial chord at the mid-span makes the graphical representation dimensionless, so the curve remains the same when the rotor blade stage 24 is scaled up or down for different applications.Table 5 lists the values ​​for the axial chord of the blade, divided by the axial chord value at the mid-span, along various span values, where the dimensionless axial chord is defined as a ratio of the axial chord at a given span to the axial chord at the mid-span. 1000,905 950,910 910,918 820,938 720,959 630,980 531,000 431,018 321,034 221,048 111,060 61,066 01,072 A blade design with the axial chord distribution illustrated in Fig. 8 can help to tune the blade's resonant frequency to avoid interference with drive components. For example, a blade with a linear shape may have a resonant frequency of 400 Hz, while a blade 36 with increased thickness at some spans may have a resonant frequency of 450 Hz. If the blade's resonant frequency is not carefully adjusted to avoid interference with the drive components, operation may result in impermissible stress on the blade 36 and potential structural failure. Accordingly, a blade design 36 with the axial chord distribution illustrated in Fig. 8 can extend the blade's service life. The technical effects of the disclosed embodiments include an improvement in the turbine's performance in several different ways. First, the design of the rotor blades 36 and the constriction distribution, as illustrated in Fig. 4, can help to manipulate secondary flows (i.e., flows perpendicular to the main flow direction) and / or scavenging flows near the hub (e.g., the first annular wall 40). Second, a rotor blade 36 with a projection 500 around the 50% span can help to tune the rotor blade's resonant frequency to avoid interference with drive components. If the rotor blade's resonant frequency is not carefully tuned to avoid interference with drive components, operation can result in overloading of the rotor blade 36 and potential structural failure.Accordingly, a design of the guide vane 36 with the increased thickness at special clamping points can extend the service life of the guide vane 36. This written description uses examples to disclose the subject matter of the invention, including the best embodiment, and also to enable any person skilled in the art to carry out the subject matter of the invention, including the creation and use of any devices or systems and the performance of any methods contained therein. The patentable scope of the invention is defined by the claims. A turbomachine contains multiple rotor blades, and each rotor blade has a foil. The turbomachine has opposing walls that define a passage through which a fluid flow can be accommodated. A constriction distribution is measured at the narrowest point in the passage between adjacent rotor blades, where adjacent blades extend transversely across the passage between the opposing walls to interact aerodynamically with the fluid flow. The foil defines the constriction distribution, and this distribution reduces aerodynamic losses and improves aerodynamic loads at each foil. Parts list: 10 Turbomachine 12 Compressor 14 Combustion chamber 16 Turbine 17 Diffuser 18 Stages 20 Stage 22 Guide vane stage 24 Rotor blade stage 26 Axis of rotation 28 Axial direction 30 Radial plane 32 Radial direction 34 Circumferential direction 36 Rotor blade 37 Blade 38 Passage 39 Mounting section 40 First wall or platform 42 Second wall 44 Leading edge 46 Trailing edge 48 Pressure side 50 Suction side 56 Axial chord 57 Chord 58 Graphical representation 60 Curve 62 Vertical axis 64 Horizontal axis 66 Point 68 Point 70 Point 500 Projection 510 Line

Claims

Rotor blade (36) comprising a blade (37), wherein the rotor blade (36) is configured for use with a turbomachine (10), the blade (37) comprising: a constriction distribution (D0), measured in a narrowest region in a passage (38) between adjacent rotor blades (36), in which adjacent rotor blades (36) extend transversely across the passage (38) between opposing walls (40, 42) to interact aerodynamically with a fluid flow; and wherein the constriction distribution (D0) is defined by a trailing edge (46) of the blade (37), wherein the constriction distribution (D0) is configured to reduce aerodynamic losses and improve aerodynamic loads on the blade (37);characterized in that the constriction distribution (D0) extends essentially linearly from a constriction / constriction mid-span value of approximately 82% at approximately 5% span to a constriction / constriction mid-span value of approximately 115% at approximately 90% span, a constriction / constriction mid-span value of approximately 110% at approximately 95% span, and a constriction / constriction mid-span value of approximately 82.5% at approximately 100% span; and the span at 0% is located on a radially inner section of the airfoil (37), while a span at 100% is located on a radially outer section of the airfoil (37), and the constriction / constriction mid-span value is 100% at approximately 50% to 55% span. Guide vane (36) according to claim 1, wherein the constriction distribution (D0) is defined by values ​​specified in Table 1 and wherein the values ​​of the constriction distribution (Do) are within a tolerance of + / -10% of the values ​​specified in Table 1. Guide vane (36) according to claim 1 or 2, wherein the trailing edge (46) of the blade (37) has a projection (500) at approximately 50% of its span. Guide vane (36) according to claim 3, wherein the trailing edge (46) of the blade (37) has an offset of about 0 at about 0% span, about 100% at about 50% span and about 0 at 100% span. Guide vane (36) according to any one of claims 2-4, wherein the trailing edge (46) of the blade (37) has an offset as defined by values ​​specified in Table 2. Guide vane (36) according to any one of claims 2-5, wherein the blade (37) has a thickness distribution (Tmax / Tmax_midspan) as defined by values ​​given in Table 3. Guide vane (36) according to any one of claims 2-6, wherein the blade (37) has a dimensionless thickness distribution according to values ​​specified in Table 4. Guide vane (36) according to any one of claims 2-7, wherein the blade (37) has a dimensionless axial chord distribution according to values ​​specified in Table 5. Turbomachine (10) comprising opposing walls (40, 42) defining a passage (38) in which a fluid flow can be received to flow through the passage (38) and multiple impeller blades (36) according to any one of the preceding claims.

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