JP2010203250A - Gas turbine blade - Google Patents

Abstract

<P>PROBLEM TO BE SOLVED: To reduce loss caused by a secondary flow even if sealing air is supplied, and to contribute to the improvement of the efficiency of a turbine. <P>SOLUTION: The gas turbine blade 20 includes a blade body 20 formed into a blade shape having a front end 20b and a rear end 20c and extended along a radial direction of a rotating shaft (L1 direction), and a retreat part 22 formed at the front end and making the front end gradually retreat to the rear end side toward a blade end 20a. The retreating of the retreat part is started from a position P separated from a root side for at least 75% of a blade height H. <P>COPYRIGHT: (C)2010,JPO&INPIT

Description

  The present invention relates to a gas turbine blade that functions as a stationary blade (stator vane) or a moving blade (rotor vane) of a gas turbine.

As is well known, a gas turbine engine is housed in an engine case, and is mainly composed of a compressor, a combustor, and a gas turbine. Among these, the gas turbine includes a rotating shaft, a plurality of moving blades, and stationary blades in a casing.
The moving blade is one of gas turbine blades, is fixed around the shaft, and plays a role of rotating the shaft under the pressure of combustion gas flowing in the axial direction of the shaft. On the other hand, the stationary blade is also one of the gas turbine blades, and is fixed to the split ring constituting the inner wall of the casing, and plays a role of changing the flow direction, pressure and speed of the combustion gas.

  These rotor blades and stationary blades are alternately arranged in the combustion gas flow path formed between the shaft and the inner wall surface of the casing, and rotate the combustion gas generated by the combustor to rotate the thermal energy. It is converted into energy to generate power.

  By the way, since it is necessary to rotate the moving blade while being accommodated in the casing, it is necessary to provide a tip clearance as a gap between the casing and the tip end surface of the moving blade. For this reason, during the operation of the gas turbine, a part of the combustion gas leaks to the trailing edge side of the moving blade through the tip clearance, resulting in a blade tip leakage vortex. This tip leakage vortex is one of the secondary flows that are different from the main flow that the rotor blades originally want to flow, disturbing the correct flow on the back side (negative pressure side) and lowering the turbine efficiency. It is considered as a factor.

  Thus, a moving blade is provided that can reduce the amount of leakage passing through the tip clearance and reduce the generation of blade tip leakage vortices (see Patent Document 1). According to this moving blade, it is possible to reduce the clearance loss and improve the turbine efficiency.

Japanese Patent Laid-Open No. 11-62506

However, although the amount of leakage can be reduced, it is impossible to completely close the tip clearance, and thus a leakage flow that passes through the tip clearance is inevitably generated. Thus, the problem of tip leakage vortices remains.
By the way, when configuring a gas turbine, normally, in order to prevent the combustion gas from leaking from the joint of the split ring to the outside, high-pressure seal air is supplied from the joint to the inside of the casing. .

  However, since the sealing air is supplied from the radially outer side (divided ring side) toward the radially inner side upstream of the moving blade, the combustion basically flows in the axial direction of the shaft. It flows in a direction different from the gas flow. In particular, the blade tip leakage vortex is easily affected by the flow of the seal air, and the direction of the vortex leaks is changed by joining the flow of the seal air. This point will be specifically described with reference to the drawings.

First, a general moving blade 100 is shown in FIGS. 10 is an external perspective view of the moving blade 100, FIG. 11 is a view of the moving blade 100 arranged in the circumferential direction viewed from the blade tip 101a side (tip side), and FIG. 12 is a moving blade. It is the figure which looked at 100 from the back side.
As shown in each figure, the moving blade 100 is integrally formed of a blade body 101 formed in an airfoil shape having a leading edge 101b and a trailing edge 101c, and a root portion 102 for fixing the moving blade 100. Has been.

  Next, when the gas turbine having the moving blade 100 configured as described above is operated, if no seal air is supplied, the flow flowing into the blade tip starts from the direction of the arrow R as shown in FIG. 100 hits the rear side front edge on the wing tip 101a side. Thereafter, the flow at the blade tip portion flows toward the trailing edge 101c having a low pressure along the back surface of the moving blade 100 as shown in FIGS. 13 is a view of the moving blade 100 as viewed from the blade tip 101a side, and FIG. 14 is a view of the moving blade 100 as viewed from the back side.

  On the other hand, when the sealing air is supplied, the direction of the flow flowing into the blade tip changes due to the influence of the sealing air as shown in FIG. 15 and hits the moving blade 100 from the arrow N direction instead of the arrow R direction. End up. That is, the direction is changed so as to hit the rear side, and it strikes the moving blade 100 in a back-flowing manner. Then, since the flow is concentrated on the rear side front edge on the blade tip 101a side of the moving blade 100, a stagnation pressure D in which the pressure is locally increased is generated in this portion.

  For this reason, the starting point of the secondary flow is shifted further toward the rear front edge side. Therefore, as shown in FIGS. 15 and 16, the range of the secondary flow is widened from the blade tip 101a side to the root side as compared with the case where the seal air is not supplied, and the flow of the secondary flow. It will become stronger. 15 is a view of the moving blade 100 as viewed from the blade tip 101a side, and FIG. 16 is a view of the moving blade 100 as viewed from the back side.

  Therefore, as shown in FIG. 17, when the total pressure loss coefficient of the moving blade 100 is compared between the case where the seal air is present (dotted line) and the case where the seal air is present (solid line), the loss is greater when the seal air is present. End up. Specifically, a loss occurs remarkably over a range of about 50% to 90% of the blade height from the root side of the moving blade 100.

  The present invention has been made in view of such circumstances, and the object thereof is to reduce the loss due to the secondary flow even if the seal air is supplied, and contribute to the improvement of the turbine efficiency. It is to provide a gas turbine blade that can.

The present invention provides the following means in order to solve the problems and achieve the object.
A gas turbine blade according to the present invention is a plurality of gas turbine blades arranged in a circumferential direction of a rotating shaft, and is formed into an airfoil having a leading edge and a trailing edge, and extends along a radial direction of the rotating shaft. A wing body formed on the leading edge, and a receding portion that gradually recedes the leading edge toward the trailing edge toward the wing tip, wherein the receding portion is at least 75% of the wing height from the root side. The backward movement is started from the separated position.

In the gas turbine blade according to the present invention, the leading edge of the blade body gradually retreats toward the trailing edge toward the blade tip (tip) by the retreating portion. For this reason, if the flow at the tip of the blade, which has changed its flow direction under the influence of sealing air, hits the vicinity of the front edge of the rear side of the blade tip side, the stagnation pressure that locally increased the pressure on the back side However, the point shifts to the downstream side (rear edge side) compared to the conventional case.
Therefore, since the starting point of the secondary flow is shifted to the downstream side, the range of the flow that spreads from the blade tip side toward the root side can be narrowed compared to the conventional case. In addition, since the leading edge gradually recedes toward the trailing edge toward the wing tip, the incidence angle (incident angle) and stagger angle at the wing tip (the line connecting the leading edge and the trailing edge and the axis of the rotation axis) Can be increased. Therefore, it is possible to facilitate the flow of the blade tip portion whose flow direction has changed to the back strike direction along the back surface, and to obtain an airfoil shape corresponding to the flow of the back strike appearance. Therefore, the stagnation pressure itself can be relaxed compared to the conventional case, and the secondary flow can be weakened.

As described above, the secondary flow itself can be weakened as compared to the conventional case, and the range of the secondary flow can be narrowed, so that the loss can be reduced and the turbine efficiency can be improved. .
Moreover, the retreating portion does not start retreating from the base side, but starts retreating from a position at least 75% of the blade height from the base side. Therefore, the above-described effects can be expected without disturbing the main flow of the gas flowing through the wing body.

  The gas turbine blade according to the present invention is characterized in that, in the gas turbine blade according to the present invention, the retracted amount of the retracted portion is 25% or less of the shaft cord of the blade body.

  In the gas turbine blade according to the present invention, the retracted amount of the retracted portion is 25% or less of the shaft code of the blade body (the length along the axial direction of the rotating shaft). Can be prevented from becoming shorter. If the shaft code of the blade tip is excessively short, the pressure of the combustion gas is received only by the remaining blade tip, which may increase the secondary flow such as blade tip leakage vortex. However, since the retraction amount is defined as described above, the secondary flow can be surely weakened.

  The gas turbine blade according to the present invention is characterized in that, in the gas turbine blade according to the present invention, the blade tip side of the blade body is curved so as to recede to the downstream side.

  In the gas turbine blade according to the present invention, since the blade tip side of the blade body including the retracted portion is curved so as to retract backward, the same effect as when the blade leading edge is retracted is the axial cord. Can be obtained without reducing the secondary flow. Therefore, loss can be further reduced.

  According to the gas turbine blade according to the present invention, loss due to the secondary flow can be reduced even if the seal air is supplied, and the efficiency of the turbine can be improved.

It is a simplified lineblock diagram of a gas turbine engine showing one embodiment concerning the present invention. It is an external appearance perspective view of the moving blade which comprises the gas turbine shown in FIG. It is the figure which looked at the state where the moving blade shown in FIG. 2 was located in the circumferential direction from the blade end side. It is the figure which looked at the moving blade shown in FIG. 3 from the back side. FIG. 3 is a diagram showing a secondary flow generated when a flow affected by seal air hits the blade tip of the blade shown in FIG. 2, and shows the blade from the blade tip side. is there. It is the figure which looked at the moving blade shown in FIG. 5 from the back side. FIG. 3 is a diagram comparing the degree of loss due to the secondary flow caused by the flow at the blade tip affected by the sealing air between the conventional blade and the blade shown in FIG. It is the figure which graphed how the total pressure loss coefficient changed over time. It is a figure which shows the modification which concerns on this invention, Comprising: It is an external appearance perspective view of the moving blade curvedly formed so that a blade end side might reverse | retreat to a downstream side. It is the figure which looked at the moving blade shown in FIG. 8 from the back side. It is an external appearance perspective view of the moving blade which comprises the conventional gas turbine. It is the figure which looked at the state where the moving blade shown in FIG. 10 was located in the circumferential direction from the blade end side. It is the figure which looked at the moving blade shown in FIG. 11 from the back side. FIG. 12 is a diagram showing a secondary flow generated when a flow hits the moving blade shown in FIG. 11 and is a view of the moving blade viewed from the blade tip side. It is the figure which looked at the moving blade shown in FIG. 13 from the back side. FIG. 11 is a diagram showing a secondary flow generated when the flow affected by the seal air hits the blade tip of the blade shown in FIG. 10, and shows the blade from the blade tip side. is there. It is the figure which looked at the moving blade shown in FIG. 15 from the back side. It is a diagram comparing how much the loss due to the secondary flow caused by the flow at the tip of the blade changes depending on the presence or absence of sealing air, and how the total pressure loss coefficient changed from the root to the tip of the blade FIG.

Hereinafter, an embodiment according to the present invention will be described with reference to FIGS.
As shown in FIG. 1, the gas turbine engine 1 is mainly composed of a compressor 2, a combustor 3, and a gas turbine 4, and the compressor 2 and the gas turbine 4 are connected by a shaft (rotary shaft) 5. The combustor 3 is disposed between the compressor 2 and the gas turbine 4.

  The compressor 2 plays a role of taking in and compressing a large amount of air. The combustor 3 mixes and burns fuel with the high-pressure air compressed by the compressor 2, and sends the combustion gas G to the flow path 6 connected to the gas turbine 4. The gas turbine 4 is mainly composed of a shaft 5 connected to the compressor 2, a plurality of moving blades (gas turbine blades) 10, a plurality of stationary blades (gas turbine blades) 11, and a casing 12 for housing them. Has been.

  The moving blades 10 are fixed to the shaft 5 in a state of being arranged around the shaft 5, and rotate the shaft 5 by receiving the pressure of the combustion gas G flowing in the axial direction (L1 direction) of the shaft 5. Is responsible. On the other hand, a plurality of stationary blades 11 are fixed to a not-shown split ring constituting the inner wall of the casing 12 in a state of being arranged around the shaft 5, and the role of changing the flow direction, pressure and speed of the combustion gas G is changed. Is responsible.

  The rotor blades 10 and the stationary blades 11 are alternately arranged in the flow path of the combustion gas G formed between the shaft 5 and the inner wall surface of the casing 12, while expanding the combustion gas G generated in the combustor 3. By flowing downstream, heat energy is converted into rotational energy of mechanical work to generate power. This power is used as power for the compressor 2 and is generally used as power for a generator of a power plant.

  By the way, the blade tip of the stationary blade 11 is provided with a seal mechanism 13 that closes the gap with the shaft 5. On the other hand, a slight gap called a tip clearance K is secured between the blade tip of the rotor blade 10 and the casing 12. Further, in order to prevent the combustion gas G flowing in the casing 12 from leaking to the outside from the joint of the split ring fixing the stationary blade 11, a high pressure is applied from the joint toward the inside in the radial direction of the casing 12. The seal air A can be supplied.

Next, the configuration of the stationary blade 11 and the moving blade 10 will be described in more detail. In addition, since the structure of the stationary blade 11 and the moving blade 10 is the same, it demonstrates taking the moving blade 10 as an example here.
As shown in FIGS. 2 to 4, the moving blade 10 is configured integrally with a blade body 20 and a root portion 21. 2 is an external perspective view of the moving blade 10, FIG. 3 is a view of the moving blade 10 arranged in the circumferential direction as viewed from the blade tip 20a side (tip side), and FIG. 4 is a moving blade. It is the figure which looked at 10 from the back side.

The root portion 21 is a portion for fixing the moving blade 10 and is formed in a Christmas tree-like key shape. The wing body 20 is formed in an airfoil shape having a leading edge 20b and a trailing edge 20c, and extends from the root side to the wing tip 20a side (tip side) along the radial direction (L2 direction) of the shaft 5. Has been.
The front edge 20b of the wing body 20 is formed with a retreating portion 22 that gradually retreats the front edge 20b toward the wing tip 20a toward the rear edge 20c. The retreating part 22 starts retreating from a position at least 75% of the blade height H from the base side of the blade body 20. In the present embodiment, the retreat is started from a position P that is 90% apart from the root side by 90% of the blade height H. Further, the retreating amount of the retreating portion 22 is retreated so that it is within 25% or less of the axial code C of the blade body 20, that is, the length along the axial direction (L1 direction) of the shaft 5. In this embodiment, it is formed so as to recede by a distance of 10% of the axis code C.

Next, the flow of the combustion gas G in the gas turbine 4 at the time of operating the gas turbine engine 1 configured as described above will be described.
During the operation of the gas turbine 4, the combustion gas G advances downstream while alternately flowing between the stationary blades 11 and the moving blades 10, and is expanded and depressurized as appropriate by the stationary blades 11 and the moving blades 10. Thereby, the thermal energy of the combustion gas G is converted into rotational energy.
By the way, apart from the original mainstream flow described above, a part of the combustion gas G passes through the tip clearance K opened between the blade tip 20a of the moving blade 10 and the casing 12, and the trailing edge 20c of the moving blade 10. Leaks to the side, and a wing tip leakage vortex is generated. At this time, under the influence of the seal air A supplied from the joint of the split ring, the flow direction changes to the backlash of the moving blade 10, and as shown in FIG. It hits the vicinity of the rear side front edge on the blade tip 20a side. Thereby, as shown in FIG. 6, the stagnation pressure D where the pressure increased locally is generated on the back side.

  However, in the moving blade 10 of the present embodiment, the leading edge 20b gradually retreats toward the trailing edge 20c toward the blade tip 20a by the retreating portion 22. Therefore, the position of the stagnation pressure D is shifted to the downstream side (rear edge 20c side) as compared with the conventional case. Therefore, since the starting point of the secondary flow is shifted to the downstream side, the range (region S) of the secondary flow that spreads from the blade tip 20a side to the root side can be narrowed compared to the conventional case.

  Further, since the leading edge 20b gradually retreats toward the trailing edge 20c toward the blade tip 20a, as shown in FIG. 3, the incidence angle (incident angle) and stagger angle θ at the blade tip 20a (the leading edge 20b) And the angle between the line M connecting the trailing edge 20c and the axis O of the shaft 5) can be increased. Therefore, it is possible to easily cause the blade tip leakage vortex, whose flow direction has changed to the backlash direction, to flow along the back surface, and to obtain an airfoil shape corresponding to the backflow-like flow. Therefore, the stagnation pressure D can be relaxed compared to the conventional case, and the secondary flow can be weakened.

  As described above, the secondary flow can be weakened compared to the conventional case, and the range of the secondary flow can be narrowed, so that the loss can be reduced. In this regard, as shown in FIG. 7, referring to a diagram comparing the total pressure loss coefficient between the conventional moving blade and the moving blade 10 of the present embodiment, the blade height H particularly from the root side under exactly the same conditions. It was confirmed that the loss was reduced in a range of 70% to 100% apart.

  As described above, according to the moving blade 10 of the present embodiment, the loss can be reduced as compared with the conventional one, which can contribute to the improvement of the turbine efficiency. Moreover, the retreating portion 22 does not start retreating from the root side, but starts retreating from a position P that is 90% away from the root side of the blade height H. Therefore, the above-described effects can be expected without disturbing the main flow of the combustion gas G flowing through the blade body 20.

  Further, since the retracting amount of the retracting portion 22 is 10% of the axis code C of the blade body 20, the axis code C of the blade tip 20a is not excessively shortened. If the shaft code C of the blade tip 20a is excessively short, the pressure of the combustion gas G is received only by the remaining blade tip 20a, and the secondary flow such as blade tip leakage vortex may become strong. is there. However, since the retraction amount is defined as described above, the secondary flow can be surely weakened.

  The technical scope of the present invention is not limited to the above embodiment, and various modifications can be made without departing from the spirit of the present invention.

  For example, in the embodiment described above, the start position P of the retreating portion 22 is set to a position that is 90% away from the root side of the blade height. . Even in this case, the loss can be reduced and the turbine efficiency can be improved. Further, although the retraction amount of the retreating portion 22 is 10% of the axis code C of the blade body 20, it is not limited to this case and may be 25% or less. Even in this case, the same effect can be obtained.

  Moreover, in the said embodiment, as shown in FIG.8 and FIG.9, you may make it curve so that the wing | blade tip 20a side of the wing | blade main body 20 containing the retreat part 22 may reverse | retreat to a downstream side. When configured in this manner, the same effect as that obtained when the leading edge of the blade tip is retracted can be obtained without shortening the shaft cord C, and the secondary flow can be further weakened. Therefore, the loss can be further reduced, which is more preferable.

  Further, the gas turbine blade according to the present invention is not limited to the moving blade 10 and may be the stationary blade 11. Even in the case of the stationary blade 11, the same effect can be obtained.

C ... Axis code 5 ... Shaft (Rotating shaft)
10 ... Rotor blade (gas turbine blade)
11 ... Static blade (gas turbine blade)
20 ... wing body 20a ... wing tip 20b ... front edge 20c ... rear edge 22 ... retracted part

Claims (3)

A plurality of gas turbine blades arranged in the circumferential direction of the rotation shaft,
A wing body formed in an airfoil having a leading edge and a trailing edge, and extending along a radial direction of the rotating shaft;
A receding portion that is formed on the leading edge and gradually recedes the leading edge toward the trailing edge toward the wing tip,
The gas turbine blade according to claim 1, wherein the retreating portion starts retreating from a position at least 75% of the blade height from the root side. The gas turbine blade according to claim 1,
The gas turbine blade according to claim 1, wherein the retracting portion has a retracting amount of 25% or less of the shaft code of the blade body. In the gas turbine blade according to claim 1 or 2,
The gas turbine blade according to claim 1, wherein a blade tip side of the blade body is curved so as to recede toward the downstream side.

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