JP5705608B2 - Rotating machine blade design method - Google Patents

Description

The present invention relates to a method of designing a blade body of the rotating machine.

  A gas turbine compresses air in a compressor to generate compressed air, and a combustor generates the combustion gas by mixing the compressed air and fuel for combustion. Then, the combustion gas is injected into the turbine flow path in which the stationary blades and rotor blades, which are blade bodies, are alternately arranged, and the rotor blades and the rotor are rotated to output the combustion gas energy as rotational energy. And powering the compressor.

  Here, the blade body such as the stationary blade and the moving blade includes a blade body having a convex back surface and a concave belly surface and having a sectional blade shape, and a blade end wall provided at an end of the blade body. And. And the combustion gas which distribute | circulates the inside of a turbine flow path is divided | segmented in the blade body front edge of these blade bodies, and flows into the back surface side and the abdominal surface side. At this time, a horseshoe-shaped vortex may be formed along the back surface and the abdominal surface from the leading edge of the wing body. Such a horseshoe-shaped vortex rolls up along with the boundary layer flow from the abdominal surface to the back surface, causing a loss of total pressure called a secondary flow loss and reducing turbine efficiency.

  In order to prevent a decrease in turbine efficiency due to such secondary flow loss, a blade body having the following inclined surface formed on the abdominal surface side of the blade body in the blade tip wall has been proposed. Specifically, the slope of the blade end wall terminates in the axial direction near the leading and trailing edges on the ventral side of the blade body and smoothly continues to the reference outer surface of the blade end wall. The flat surface has a relatively large width in the circumferential direction and is connected to the blade belly surface near the center in the axial direction of the blade belly surface (see, for example, Patent Document 1).

JP 2006-291949 A

    However, the blade body of Patent Document 1 cannot effectively reduce the pressure loss, and a technique for improving the turbine efficiency by reducing the pressure loss more effectively is desired. Further, in the wing body of Patent Document 1, the height of the wing body portion that exhibits the performance as the wing is reduced by forming the wing tip wall to have a slope and rise, and as a result, There was a problem that the turbine efficiency might decrease.

The present invention was made to solve the above problems, there is provided a method of designing the wing body capable rotating machine can be effectively suppressed pressure loss.

In order to solve the above problems, the present invention proposes the following means.
The present invention relates to a wing body of a rotary machine comprising a wing body having a convex back surface and a concave belly surface between a leading edge and a trailing edge, and a wing end wall to which an end of the wing body is connected. In the design method, a reference wing body having a convex reference back surface and a concave reference abdominal surface between a front edge and a rear edge based on preset conditions, and an end portion of the reference wing body A reference wing body determining step for determining a reference wing body having a connected reference wing end wall, and a surface of the wing end wall on the reference abdominal surface side of the reference wing body determined in the reference wing body determining step, A base end line determining step for determining a base end line that forms a smooth concave curve that is continuous to the leading edge and the trailing edge within a range inside the chord and that is concave on the chord side. And the base line determined in the base line determination step is continuous with the surface of the reference blade end wall and A blade thickness portion abdominal surface setting step for setting a blade thickness portion abdominal surface that is a curved surface continuous with the reference abdominal surface of the reference wing body, so as to form a smooth concave curved surface in a cross section orthogonal to the vertical direction; The portion where the blade thickness portion abdominal surface is set in the blade thickness portion abdominal surface setting step is the blade thickness portion abdominal surface, and the portion where the blade thickness portion abdominal surface is not set in the blade thickness portion abdominal surface setting step is the reference abdominal surface And a wing body determining step for determining the wing tip wall by determining the wing body and having a surface continuously connected to the reference abdominal surface.

  According to this method, by executing the reference wing body determination step, the base line determination step, the blade thickness portion abdominal surface setting step, and the wing body determination step, the cross section perpendicular to the direction along the chord, Are formed in a smooth concave curved surface that is continuous with the abdomen surface portion and the surface of the blade tip wall, and is continuous with the leading edge and the trailing edge in a cross section perpendicular to the blade height direction. Thus, it is possible to design a wing body having a portion where the thickness of the wing having the abdominal surface of the wing thick portion formed in a smooth concave curved surface is increased. In such a wing body, as described above, it is possible to suppress an increase in pressure loss by suppressing cross flow while effectively exhibiting wing performance. Further, when another wing body is arranged on the downstream side, generation of a horseshoe vortex at the leading edge of the other wing body can be suppressed, and pressure loss in the other wing body can also be suppressed. .

According to the rotating machine of the present invention, pressure loss can be effectively suppressed.
Further, according to the gas turbine of the present invention, it is possible to improve the turbine efficiency.
Further, according to the method for designing a wing body of a rotary machine of the present invention, a wing body capable of effectively suppressing pressure loss can be designed.

1 is a schematic diagram of a gas turbine according to an embodiment of the present invention. In the gas turbine concerning the embodiment of the present invention, it is a fragmentary perspective view showing details of a turbine blade and a turbine stationary blade. It is sectional drawing of the turbine rotor blade which concerns on embodiment of this invention, Comprising: (a) Sectional drawing in sectional line AA in FIG. 2, (b) Sectional drawing in sectional line BB in FIG. c) It is sectional drawing in the cutting | disconnection line CC in FIG. It is sectional drawing of the turbine rotor blade which concerns on embodiment of this invention, Comprising: (a) Sectional drawing in the cutting line DD in FIG. 2, (b) Sectioning drawing in the cutting line EE in FIG. c) It is sectional drawing in the cutting line FF in FIG. In the gas turbine which concerns on embodiment of this invention, it is sectional drawing which looked at the turbine rotor blade and turbine stationary blade explaining the state of the flow of combustion gas to radial direction. FIG. 6 is a cross-sectional view taken along a cutting line GG in FIG. 5 in the gas turbine according to the embodiment of the present invention, in which a turbine rotor blade and a turbine stationary blade for explaining a flow state of combustion gas are viewed in the circumferential direction. In the turbine rotor blade design method according to the embodiment of the present invention, it is an explanatory diagram for explaining a reference blade body design process, (a) a cross-sectional view viewed in the radial direction, (b) a cutting line H- in (a). It is sectional drawing in H. In the turbine blade design method according to the embodiment of the present invention, it is an explanatory view for explaining a base line determination step, (a) a sectional view viewed in the radial direction, (b) a cutting line H- in (a). It is sectional drawing in H. In the turbine rotor blade design method according to the embodiment of the present invention, it is an explanatory diagram for explaining a blade thickness portion abdominal surface setting step, (a) a sectional view viewed in the radial direction, (b) a cutting line H in (a) It is sectional drawing in -H. In the turbine rotor blade design method according to the embodiment of the present invention, it is an explanatory diagram for explaining a blade body determination step, (a) a cross-sectional view viewed in the radial direction, (b) a cutting line HH in (a). FIG. It is a streamline figure in the 2nd stage turbine blade vicinity which shows the CFD analysis result of an Example and a comparative example, (a) A comparative example, (b) It is an Example. It is a graph showing the relationship between the blade height direction position which shows the CFD analysis result of an Example and a comparative example, and the magnitude | size of a pressure (total pressure) loss, (a) Two-stage turbine blade, (b) Three-stage turbine It is a graph about a stationary blade. It is a graph showing the relationship between the blade height direction position which shows the CFD analysis result of an Example and a comparative example, a blade height direction position, and an axial P flow velocity, (a) Two-stage turbine blade, (b ) It is a graph about a three-stage turbine stationary blade.

Hereinafter, embodiments according to the present invention will be described with reference to the drawings.
An embodiment according to the present invention will be described with reference to FIGS. The gas turbine 1 includes a compressor 2 that generates compressed air, a combustor 3 that generates a combustion gas M by mixing fuel with the compressed air generated by the compressor 2, and a combustor 3. And a turbine 4 that rotationally drives the combustion gas M as a working fluid. A rotary shaft 5 is inserted through the compressor 2 and the turbine 4. The compressor 2 includes a compressor casing 2a through which the rotating shaft 5 is inserted, a compressor moving blade 2b that can rotate together with the rotating shaft 5, and a compressor stationary blade 2c fixed to the compressor casing 2a. A plurality of compressor blades 2b and compressor stationary blades 2c are provided radially in the circumferential direction R, respectively. The compressor rotor blades 2b and the compressor stationary blades 2c are alternately provided in the axial direction P, and each of the compressor rotor blades 2b and the compressor stationary blades 2c is provided in a plurality of stages. The sucked air flows between the compressor stationary blades 2c and is compressed by the rotation of the compressor blades 2b on the downstream side, whereby the compressed air is generated.

  The turbine 4 includes a turbine casing 10 through which the rotating shaft 5 is inserted, a turbine blade 20 that can rotate together with the rotating shaft 5, and a turbine stationary blade 30 that is fixed to the turbine casing 10. A plurality of turbine blades 20 and turbine stationary blades 30 are provided radially in the circumferential direction R. Further, the turbine rotor blades 20 and the turbine stationary blades 30 are alternately provided in the axial direction P, and each of the plurality of blades provided in the circumferential direction R constitutes one stage, and a plurality of stages are provided. And the combustion gas M which is the working fluid which flowed in from the combustor 3 circulates between the turbine stationary blades 30, and the turbine blades 20 are fixed by repeatedly rotating the turbine blades 20 on the downstream side. The rotating shaft 5 is rotated by applying torque.

  FIG. 2 is a perspective view showing details of the turbine rotor blade 20 and the turbine stationary blade 30 located on the downstream side in the axial direction P. FIG. As shown in FIG. 2, the turbine rotor blade 20 includes a blade body 25 having a convex back surface 23 and a concave belly surface 24 between a leading edge 21 and a trailing edge 22, and an inner periphery of the blade body 25. And a platform 26 which is a wing tip wall to which a base end portion on the side is connected. A split ring 10 a that forms a part of the turbine casing 10 with a gap is disposed on the outer peripheral side of the tip of the blade body 25 of the turbine rotor blade 20.

  Further, the turbine stationary blade 30 has a blade body 35 having a convex back surface 33 and a concave belly surface 34 between a leading edge 31 and a trailing edge 32, and a base end portion serving as an outer peripheral side of the blade body 35. An outer shroud 36 that is a connected wing end wall and an inner shroud 37 that is a wing end wall to which a tip portion on the inner peripheral side of the wing body 25 is connected are provided. The platform 26 of the turbine rotor blade 20 and the inner shroud 37 of the turbine stationary blade 30 are alternately arranged in the axial direction P with a gap therebetween, and the combustion gas M 4 circulates in the combustion gas flow path 4b. A peripheral partition is formed. The split ring 10a and the outer shroud 36 of the turbine stationary blade 30 are alternately arranged in the axial direction P with a gap therebetween, and constitute a partition wall on the outer peripheral side of the combustion gas passage 4b. .

  As shown in FIGS. 2 to 4, the blade main body 25 of the turbine rotor blade 20 is provided between the main body portion 251 and the main body portion 251 and the platform 26, and compared with the main body portion 251, the leading edge 21. To the rear edge 22, the blade thickness portion 252 is formed on the side of the abdominal surface 24 so as to increase the thickness of the blade. Here, in the cross section orthogonal to the blade height direction shown in FIG. 3, a line segment indicated by an alternate long and short dash line is in contact with each of the leading edge 21 and the trailing edge 22 at the ends and connects the leading edge 21 and the trailing edge 22. Is a chord 28. Further, in the abdominal surface 24 and the back surface 23 of the wing body 25, those in the main body portion 251 are referred to as a main body portion abdominal surface 241 and main body portion back surface 231, and those in the wing thick portion 252 are referred to as a wing thick portion abdominal surface 242 and wing thick portion back surface 232. Called. The curved surface shape along the blade chord 28 in the main body part abdominal surface 241 and the main body part back surface 231 and the blade thick part back surface 232 and the curved surface shape along the blade height direction which is the radial direction R in the rotating shaft 5 are required. It is designed by a known method based on the turbine performance, the characteristics of the working fluid, and the like. In addition, the blade thickness portion ventral surface 242 has a blade thickness reduced from a curved surface shape designed by a known method based on the required turbine performance, working fluid characteristics, etc., through a process described later. It is designed into a shape that has been processed to increase its thickness. As a result, as shown in FIGS. 3 and 4, the blade thick portion abdominal surface 242 is a cross-section (FIG. 4) orthogonal to the direction along the chord 28, and the body portion abdominal surface 241 and the platform 26 which is the blade end wall. It is formed in a smooth concave curved surface shape that is continuous with the surface 26a, and is smooth so as to be continuous with the leading edge 21 and the trailing edge 22 in a cross section perpendicular to the blade height direction (FIG. 3). It is formed in a concave curved surface. Here, the height of the blade thick portion abdominal surface 242 gradually increases from the front edge 21 and the rear edge 22 at the end toward the center of the abdominal surface 24 to form a vertex T24. The position of the apex T24 in the blade height direction is set from 5% to 50% of the chord length L28a at the surface 26a position (0% position in the blade height direction) from the surface 26a of the platform 26. Is preferred.

Next, the operation of this embodiment will be described.
As shown in FIGS. 5 and 6, the combustion gas M flowing from the upstream toward the downstream in the combustion gas flow path 4 b is divided at the front edge 21 of the turbine rotor blade 20, and the abdominal surface 24 side and the back surface 23. And flows between the adjacent turbine rotor blades 20. The blade thickness portion 252 is formed so that the thickness of the blade is increased on the side of the abdominal surface 24 from the front edge 21 to the rear edge 22 as compared with the main body portion 251, so that the distribution of the combustion gas M The area decreases, the speed increases, and the static pressure decreases. For this reason, the static pressure difference between the mutually opposite abdominal surface 24 and the back surface 23 of the turbine blades 20 adjacent to each other in the circumferential direction R is reduced. The cross flow M1 flowing toward the rear surface 23 side of the turbine rotor blade 20 is relieved. Further, by relaxing the cross flow M1, the secondary flow M2 generated when the cross flow M1 collides with the back surface 23 of the turbine rotor blade 20 and winds up can be suppressed.

  Here, as described above, the blade thickness portion abdominal surface 242 of the blade thickness portion 252 has a cross section perpendicular to the direction along the chord 28 and a main body portion abdominal surface 241 which is the abdominal surface 24 of the main body portion 251, and the surface 26 a of the platform 26. Are formed in a smooth concave curved surface shape that is continuous with each other, and a smooth concave curved surface that is continuous with the leading edge 21 and the trailing edge 22 in a cross section orthogonal to the blade height direction (radial direction Q). It is formed in a shape. For this reason, the function as the abdominal surface 24 on which positive pressure acts can be maintained, and the performance as a wing can be effectively exhibited over the entire blade height direction. Further, as described above, the static pressure of the combustion gas M decreases in the blade thickness portion 252, so that a partial flow M3 of the combustion gas M flowing in the vicinity of the blade thickness portion 252 in the main body portion 251 causes the trailing edge 22 side. , The flow rate of the combustion gas M on the blade thick portion 252 trailing edge 22 side can be increased. Therefore, the turbine stationary blade 30 on the downstream side of the turbine rotor blade 20 includes an inner shroud 37 serving as a blade end wall of the flow rate of the combustion gas M flowing from the trailing edge 22 of the turbine rotor blade 20 to the turbine stationary blade 30. The flow rate in the vicinity can be increased, whereby the formation of the horseshoe-shaped vortex M4 at the leading edge 31 of the downstream turbine vane 30 can be suppressed. As a result, the pressure loss caused by the formation of the horseshoe-shaped vortex M4 in the turbine stationary blade 30 on the downstream side can also be suppressed. For this reason, in the gas turbine 1, the efficiency of the turbine 4 can be improved.

  Here, the blade thickness portion 252 is formed in a range within 50% of the chord length L28a from the surface 26a of the platform 26 in the blade height direction, and more effectively, over the entire blade height direction. The performance as a wing can be demonstrated. Further, the blade thickness portion 252 is formed such that the position where the blade tip portion T24 is located at the apex T24 in the blade height direction from the surface 26a of the platform 26 is within a range of 5% or more of the chord length L28a from the surface 26a of the platform 26. The cross flow M1 can be suppressed more effectively. Further, the formation of the horseshoe-shaped vortex M4 at the front edge 31 of the turbine stationary blade 30 can be more effectively suppressed in the downstream turbine stationary blade 30.

Next, a method for designing such a turbine stationary blade 30 will be described.
The design method of this embodiment includes a reference blade body determination step, a base line determination step, a blade thickness portion abdominal surface setting step, and a blade body determination step, and the turbine stationary blade 30 is designed by executing these steps. To do. Details are shown below.

  As shown in FIG. 7, first, as a reference wing body determination step, a convex reference back surface 123 and a concave reference abdominal surface 124 are provided between a front edge 121 and a rear edge 122 based on preset conditions. A reference wing body 120 having a reference wing body 125 having a reference wing end wall 126 to which an end portion of the reference wing body 125 is connected is determined. Here, the preset conditions are the required turbine output, the temperature of the combustion gas M, the flow rate of the combustion gas M, and the like. The reference wing body 120 is determined by performing, for example, CFD analysis.

  Next, as shown in FIG. 8, the base chord 128 is applied to the surface 126a of the reference blade end wall 126 on the reference abdominal surface 124 side of the reference wing body 120 determined in the reference wing body determination step as the base line determination step. A base line 129 that is a smooth concave curve that is concave on the side of the chord 128 is determined so as to be continuous with the leading edge 121 and the trailing edge 122 in the inner range. Here, the specifications such as the position and curvature of the base line 129 are determined by performing CFD analysis on the flow in the vicinity of the surface 126a of the reference blade end wall 126, for example.

  Next, as shown in FIG. 9, as the blade thick portion abdominal surface setting step, the base end line 129 determined in the base end line determination step is continuous with the surface 126a of the reference blade end wall 126 and the chord 128 The blade thickness portion abdominal surface 242 which is a curved surface continuous with the reference abdominal surface 124 of the reference wing body 125 is set so as to have a smooth concave curved surface shape in a cross section orthogonal to the direction along.

Then, as shown in FIG. 10, the final turbine rotor blade 20 is determined as the blade body determination step. That is, the abdominal surface 24 is defined as the blade thickness portion abdominal surface 242 for the portion where the blade thickness portion abdominal surface 242 is set in the blade thickness portion abdominal surface setting step, and the portion where the wing thickness portion abdominal surface 242 is not set in the blade thickness portion abdominal surface setting step. Determines the turbine blade 20 as a reference ventral surface 124 and the platform 26 as having a surface that is continuously connected to the reference ventral surface 124.
The turbine rotor blade 20 is designed as described above.

  Next, CFD analysis is performed on an example of the turbine 4 including the two-stage turbine rotor blade 20 having the blade thickness part 252 and a comparative example of the conventional two-stage turbine 20 ′ not having the blade thickness part 252. did. Here, in both the example and the comparative example, the turbine stationary blade 30 (30 ′) and the turbine rotor blade 20 (20 ′) are configured in four stages. In the two-stage turbine rotor blade 20 of the embodiment, the two-stage turbine rotor blade 20 ′ in the turbine of the comparative example has a blade thickness portion 252. The apex T24 of the blade thickness portion 252 was set at a position of 50% of the chord length L28a at the position of 0% from the surface 26a of the platform 26 in the blade height direction.

  FIG. 11 is a streamline diagram obtained by CFD analysis, showing a two-stage turbine blade, and (a) is a comparative example, that is, a configuration having no blade thickness portion 252 ( b) is an example, that is, a configuration having a blade thickness portion 252. As shown in FIG. 11, compared with the turbine blade 20 ′ of the comparative example, the turbine blade 20 of the embodiment is improved so that the direction of the cross flow M1 is along the axial direction P. This is because, as described above, in the turbine rotor blade 20 of the embodiment, the difference in static pressure between the abdominal surface 24 side and the back surface 23 side is reduced by having the blade thickness portion 252.

  FIG. 12 is a graph showing the relationship between the blade height direction position and the pressure (total pressure) loss obtained by CFD analysis. (A) Two-stage turbine blade, (b) 3 The relationship in a stage turbine stationary blade is shown, respectively, and a solid line shows an example and a dotted line shows a comparative example. FIG. 13 is a graph showing the relationship between the blade height direction position and the axial flow velocity obtained by CFD analysis. (A) Two-stage turbine blade, (b) Three-stage turbine stationary blade The solid line indicates an example, and the dotted line indicates a comparative example. 12 and 13, the blade height direction position means that the surface 26a of the platform 26 which is the blade end wall serving as the base end in the turbine rotor blade 20 is 0 (%) and the tip is 100 ( %), From the proximal end to the distal end, expressed as a percentage of the total height.

  As shown in FIG. 12, in the two-stage turbine rotor blade 20 having the blade thickness portion 252, it is apparent that the pressure loss is suppressed in the range of the blade height position of 30% or less. This is because, as described above, the cross flow M1 is suppressed, and further, the secondary flow generated by the winding due to the cross flow M1 is suppressed. Furthermore, as shown in FIG. 12, even in a three-stage turbine stationary blade that is downstream of a two-stage turbine rotor blade having a blade thickness portion 252, pressure loss is suppressed within a blade height position of 40% or less. It is clear that As described above, the flow rate of the combustion gas M from the trailing edge of the two-stage turbine rotor blade having the upstream blade thickness portion 252 increases near the surface of the platform, thereby generating a horseshoe-shaped vortex near the leading edge. This is due to the suppression. This is apparent from the fact that, in the three-stage turbine stationary blade, as shown in FIG. 13, the axial flow velocity is increased in the blade height direction of 10% or less in the vicinity of the surface of the platform.

  As described above, the gas turbine 1 according to the present embodiment includes the turbine rotor blade 20 having the blade thickness portion 252, thereby suppressing the pressure loss in the turbine rotor blade 20 and reducing the turbine static on the downstream side. Pressure loss at the blades 30 can also be suppressed, thereby improving turbine efficiency.

  As mentioned above, although embodiment of this invention was explained in full detail with reference to drawings, the concrete structure is not restricted to this embodiment, The design change etc. of the range which does not deviate from the summary of this invention are included.

  In the above embodiment, the turbine blade 20 has the blade thickness portion 252 in two stages, but may have the blade thickness portion 252 in all stages. Further, the turbine stator blade 30 has the same effect even if it has a blade thickness portion. In this case, the turbine stationary blade 30 is provided with an inner shroud 37 as a blade end wall on the inner peripheral side and an outer shroud 36 as a blade end wall on the outer peripheral side. It is good also as what provides a part. That is, in a rotary machine equipped with a wing body, the wing body has a blade thickness portion, thereby suppressing pressure loss in the wing body and improving efficiency.

1 Gas turbine (rotary machine)
20 Turbine blade (wing body)
21 Front edge 22 Rear edge 23 Back surface 24 Abdominal surface 25 Wing body 26 Platform (wing end wall)
28 Blade chord 30 Turbine stationary blade 120 Reference blade body 123 Reference back surface 124 Reference abdominal surface 125 Reference blade body 126 Reference blade end wall 128 Base line 241 Body portion abdominal surface 242 Blade thick portion abdominal surface 251 Body portion 252 Blade thickness portion L28 Blade chord length P axis direction Q radial direction R circumferential direction T24 apex

Claims (1)

In a wing body design method for a rotary machine comprising: a wing body having a convex back surface and a concave abdominal surface between a leading edge and a trailing edge; and a wing end wall to which an end of the wing body is connected. ,
Based on preset conditions, a reference wing body having a convex reference back surface and a concave reference abdominal surface between the leading edge and the trailing edge, and a reference wing to which an end of the reference wing body is connected A reference wing body determining step for determining a reference wing body having an end wall;
On the surface of the wing tip wall on the reference abdominal surface side of the reference wing body determined in the reference wing body determination step so as to be continuous with the leading edge and the trailing edge within the range inside the chord, And, a base line determination step for determining a base line that becomes a smooth concave curve that is concave on the chord side,
The base line determined in the base line determination step is continuous with the surface of the reference blade end wall, and has a smooth concave curved surface in a cross section perpendicular to the direction along the chord, A blade thickness portion abdominal surface setting step for setting a blade thickness portion abdominal surface that is a curved surface continuous with the reference abdominal surface of the reference wing body;
For the portion where the blade thickness portion stomach surface is set in the blade thickness portion stomach surface setting step, the blade thickness portion stomach surface is the blade thickness portion stomach surface, and for the portion where the blade thickness portion stomach surface is not set in the blade thickness portion stomach surface setting step A blade for a rotary machine, comprising: a blade body determining step for determining the blade body as the reference abdominal surface and determining the blade end wall as having a surface continuously connected to the reference abdominal surface. Body design method.

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