JP2023025389A - Two-shaft gas turbine - Google Patents

Abstract

To provide a two-shaft gas turbine capable of realizing further increase in peripheral speed.SOLUTION: A two-shaft gas turbine comprises: a compressor which has a rotary shaft rotatable around a shaft line; a high-pressure turbine connected to the rotary shaft; a low-pressure turbine arranged with a distance from the high-pressure turbine in a shaft line direction; and an intermediary flow passage which is arranged between a last stage blade of the high-pressure turbine and a first stage blade of the low-pressure turbine and guides fuel gas from the high-pressure turbine to the low-pressure turbine. An inner diameter of the first stage blade in a radial direction is equal to or larger than 1.2 times and equal to or less than 1.4 times the inner diameter of the last stage blade. A length of the first stage blade is equal to or larger than 1.0 time and equal to or less than 1.6 times the length of the last stage blade. A maximum thickness at a hub side edge section is equal to or larger than 4 times and equal to or less than 8 times the maximum thickness at a chip side edge section.SELECTED DRAWING: Figure 4

Description

The present disclosure relates to twin shaft gas turbines.

2. Description of the Related Art Conventionally, a twin-shaft gas turbine is known in which a high-pressure turbine and a low-pressure turbine are provided on separate shafts, and combustion gas after passing through the high-pressure turbine is supplied to the low-pressure turbine through an intermediate duct.

In recent years, there has been an increasing demand for higher output of this type of two-shaft gas turbine. Here, especially when using a two-shaft gas turbine for power generation, the rotation speed on the low-pressure turbine side must be kept constant (eg, 3000 rpm). In order to improve the output without changing the rotation speed, it is conceivable to increase the average diameter of the rotor blades (first stage rotor blades) of the low-pressure turbine. When the average diameter of the low-pressure turbine is increased, the cross-sectional area of the flow path of the low-pressure turbine is also increased, which reduces the flow velocity of the combustion gas. On the other hand, if the high-pressure turbine, whose design is difficult to change due to its high temperature, is not improved, the difference in diameter between the high-pressure turbine and the low-pressure turbine will increase, and the intermediate flow path that connects the two will bend. For this reason, the flow velocity distribution in the radial direction of the moving blade becomes non-uniform, resulting in an increase in flow loss at the same time. In addition, as the average diameter increases, the centrifugal stress applied to the moving blade also increases. As a configuration for reducing such centrifugal stress, for example, one described in Patent Document 1 below is known.

The rotor blade according to Patent Document 1 below adopts a configuration in which the direction in which the stacking line of the rotor blade extends is different between the radially outer portion and the radially inner portion.

WO2020/095470

However, even if the rotor blade shape is changed as described above, it is difficult to reduce the centrifugal stress on the hub side. As a result, there has been a problem that a further increase in peripheral speed of the low-pressure turbine cannot be achieved.

The present disclosure has been made to solve the above problems, and an object of the present disclosure is to provide a two-shaft gas turbine capable of realizing a further increase in peripheral speed.

In order to solve the above problems, a two-shaft gas turbine according to the present disclosure includes: a compressor having a rotating shaft rotatable about an axis; a high-pressure turbine having a first shaft coupled to the rotating shaft; a low pressure turbine having a second shaft different from the first shaft and spaced axially relative to the high pressure turbine; a last stage rotor blade of the high pressure turbine and a first stage rotor blade of the low pressure turbine; and an intermediate passage for guiding combustion gas from the high-pressure turbine to the low-pressure turbine, wherein the inner diameter of the first stage rotor blade in a radial direction with respect to the axis is 1.2 times that of the last stage rotor blade 1.0 times or more and 1.4 times or less, and the length of the first stage moving blade in the radial direction is 1.0 time or more and 1.6 times or less, and is the radially inner end portion of the first stage moving blade. The maximum thickness of the hub-side end is 4 to 8 times the maximum thickness of the tip-side end, which is the radially outer end.

Advantageous Effects of Invention According to the present disclosure, it is possible to provide a two-shaft gas turbine capable of realizing a further increase in peripheral speed.

1 is a schematic diagram showing the configuration of a two-shaft gas turbine according to an embodiment of the present disclosure; FIG. 1 is an enlarged cross-sectional view of a main part of a two-shaft gas turbine according to an embodiment of the present disclosure; FIG. 1 is a radial view of a first stage rotor blade of a low pressure turbine according to an embodiment of the present disclosure; FIG. 1 is a circumferential view of a first stage rotor blade of a low pressure turbine according to an embodiment of the present disclosure; FIG.

(Configuration of 2-shaft gas turbine)
A two-shaft gas turbine 1 according to a first embodiment of the present disclosure will be described below with reference to FIGS. 1 to 4. FIG. The two-shaft gas turbine of the present embodiment is suitably used as a heavy-structure gas turbine for various industrial machines, power generators, and the like. Note that the two-shaft gas turbine according to the present embodiment can also be used as a gas turbine for other uses such as for aircraft (for aircraft engines).

The two-shaft gas turbine 1 includes a compressor driving side turbine section (gas generator section) 2 and an output side turbine section (power turbine section) 3, as shown in FIG. The two-shaft gas turbine 1 is configured such that the output-side turbine section 3 drives a load device 10 such as an industrial machine or a generator-motor. The two-shaft gas turbine 1 also includes a control device (not shown), a turbine casing enclosing a compressor drive side turbine section 2 and an output side turbine section 3, and the like.

The compressor driving side turbine section 2 includes a compressor 4 that compresses air R1 taken from the atmosphere to generate compressed air R2, and fuel mixed with the compressed air R2 sent from the compressor 4 and combusted, It includes a combustor 5 that generates combustion gas R3, and a high pressure turbine 6 that is coaxially connected to the compressor 4 via a first shaft (gas generator shaft) 7 that also serves as a rotor of the high pressure turbine 6 . In the compressor drive side turbine section 2, the high-pressure turbine 6 is rotated by the high-temperature, high-pressure combustion gas R3 sent from the combustor 5, and the rotational power of the high-pressure turbine 6 is transmitted to the compressor 4 through the first shaft 7. Compressor 4 is driven. The first shaft 7 also serves as the rotor of the high pressure turbine 6 .

An air intake port of the compressor 4 is provided with an IGV (inlet guide vane) (not shown). The IGV is driven by an IGV driving device, and the air intake amount of the compressor can be adjusted by adjusting the opening of the IGV.

The output side turbine section 3 has a low pressure turbine 8 . The low-pressure turbine 8 and load equipment 10 are connected via a second shaft (power turbine shaft) 9 that also serves as the rotor of the low-pressure turbine 8 . The low-pressure turbine 8 is driven to rotate by the combustion gas R4, which is sent from the high-pressure turbine 6 with combustion gas R4 whose pressure has been lowered by driving the high-pressure turbine 6. Rotational power obtained by the low-pressure turbine 8 is transmitted to the load equipment 10 to drive the load equipment 10 . The combustion gas R5 that drives the low-pressure turbine 8 is discharged as exhaust gas.

As shown in FIGS. 1 and 2, between the compressor drive side turbine section 2 and the output side turbine section 3, that is, between the high pressure turbine 6 and the low pressure turbine 8 in the direction of the axis O1, there is a 8 is provided with an intermediate channel section (intermediate channel section) 12 having an intermediate duct 11 (intermediate channel 13) for feeding the combustion gas R4. The intermediate duct 11 has an annular double-tube structure including an inner tube 11a and an outer tube 11b coaxially arranged along the axis O1. In the intermediate duct 11, a space between the inner tube 11a and the outer tube 11b serves as an intermediate flow path 13 for circulating the combustion gas R4.

The intermediate flow path 13 (intermediate duct 11) is provided between the final stage moving blade 14 of the high pressure turbine 6 and the first stage moving blade 15 of the low pressure turbine 8 in the direction of the axis O1, and flows the combustion gas R4 from the high pressure turbine 6 to the low pressure turbine 8. supply. Further, in the two-shaft gas turbine 1 of the present embodiment, the strut 16 arranged in the intermediate flow path 13 is configured to also serve as the first-stage stator vane of the low-pressure turbine 8 . A plurality of struts 16 are radially arranged within the intermediate flow path 13 at intervals in the circumferential direction of the axis O1.

(Structure of first-stage moving blade of low-pressure turbine)
Next, the configuration of the first stage moving blade 15 of the low pressure turbine 8 will be described in detail with reference to FIGS. 2 to 4. FIG. As shown in FIG. 4, the first stage rotor blade 15 has a blade body 15h, a platform 15c, and a blade root 15d.

The blade main body 15h extends in a radial direction with respect to the axis O1 and has an airfoil cross-sectional shape when viewed from the radial direction. As shown in FIG. 3, the radially outer end of the blade main body 15h is a tip-side end 15b, and the radially inner end is a hub-side end 15a. Further, of both surfaces facing the circumferential direction of the blade main body 15h having an airfoil cross section, the surface that is recessed on one side in the circumferential direction serves as a pressure surface 15p, and the surface that faces on the opposite side is convexly curved on the one side in the circumferential direction. Thus, a negative pressure surface 15n is formed.

The platform 15c is provided at the radially inner end of the blade body 15h. The dimension of the platform 15c in the direction of the axis O1 is larger than the dimension of the blade main body 15h in the direction of the axis O1. The blade root 15d is a portion for fixing the first-stage rotor blade 15 to a blade groove formed in a disk (not shown), and has a plurality of serration-like irregularities. Here, as shown in FIG. 4, when the cord length at the hub-side end portion 15a is C2 and the dimension of the blade root 15d in the direction of the axis O1 is L1, the value of L1 is 1.2 times or more of C2. .5 times or less.

Further, as shown in FIG. 2, when the blade length (length) of the final stage rotor blade 14 of the high pressure turbine 6 is w1 and the blade length (length) of the first stage rotor blade 15 of the low pressure turbine 8 is w2, The value of w2 is 1.0 times or more and 1.6 times or less of w1. Furthermore, when the inner diameter of the final stage rotor blade 14 (the radial dimension based on the axis O1) is d1 and the inner diameter of the first stage rotor blade 15 is d2, the value of d2 is 1.2 times or more of d1. .4 times or less.

In addition, as shown in FIG. 3, the cord length C2 of the blade main body 15h at the hub-side end portion 15a is 1.0 to 1.3 times the cord length at the tip-side end portion 15b. Furthermore, when the maximum thickness of the blade main body 15h at the hub-side end portion 15a is D2, and the maximum thickness at the tip-side end portion 15b is D1, the value of D2 is set to 4 times or more and 8 times or less of D1. . In addition, at the 75% position in the radial direction with respect to the hub-side end portion 15a, the thickness is set to be two to four times the maximum thickness of the blade main body 15h as compared to the 25% position.

(Effect)
In recent years, there has been an increasing demand for higher output of the two-shaft gas turbine 1 as described above. Here, especially when the two-shaft gas turbine 1 is used for power generation, the rotation speed on the low-pressure turbine 8 side must be kept constant (eg, 3000 rpm). In order to improve the output without changing the rotational speed, it is conceivable to increase the average diameter of the rotor blades (first stage rotor blades 15) of the low-pressure turbine 8. FIG. When the average diameter is increased, the cross-sectional area of the flow path of the low-pressure turbine 8 is also increased, so the flow velocity of the combustion gas is decreased. On the other hand, the diameter difference with the high-pressure turbine 6 increases, and the flow path curves as shown in FIG. As a result, the flow velocity distribution in the radial direction of the first-stage rotor blade 15 becomes non-uniform, resulting in an increase in flow loss. In addition, as the average diameter increases, the centrifugal stress applied to the moving blade also increases.

Therefore, the two-shaft gas turbine 1 according to the present embodiment has the dimensional conditions for the first stage moving blade 15 as described above. According to the above configuration, the inner diameter d2 of the first stage rotor blade 15 is 1.2 times or more and 1.4 times or less as large as the inner diameter d1 of the last stage rotor blade 14 of the high pressure turbine 6 . Furthermore, the length w2 of the first-stage rotor blade 15 is set to 1.0 times or more and 1.6 times or less the length w1 of the last-stage rotor blade 14 . If the length of the first-stage rotor blade 15 is kept to 1.6 times or less, the distribution of the combustion gas flow in the intermediate passage 13 is not so biased. can mitigate the increase in

Furthermore, by adopting this configuration, the length of the first-stage moving blade 15 is not excessively increased, and restrictions on the natural frequency are relaxed. In general, in response to an increase in centrifugal stress, it is often the case that the stress is reduced by increasing the length of the cord on the hub side to increase the cross-sectional area. This is because, in many cases, the natural first-order mode is the vibration mode in the chord direction, and resonance with the low-order excitation source becomes a problem, so it is necessary to secure the natural frequency as much as possible. . On the other hand, in the present embodiment, since the upper limit of the length of the first-stage moving blade 15 is restricted, the natural frequency is already sufficiently high, and extension of the cord length is unnecessary. Instead of enlarging the cord length, the maximum thickness D2 of the hub-side end portion 15a of the first-stage rotor blade 15 is set to 4 times or more and 8 times or less of the maximum thickness D1 of the tip-side end portion 15b. Take measures to secure the area. As a result, the cross-sectional area of the hub-side end portion 15a can be increased while the blade weight is suppressed, and the centrifugal stress acting on the hub cross-section of the first-stage rotor blade 15 can be reduced.

Further, by increasing the maximum thickness of the hub-side end portion 15a, the curvature of the pressure surface 15p is reduced. This reduces the possibility that the combustion gas will separate from the pressure surface 15p, and the efficiency of the low-pressure turbine 8 can be further improved.

Furthermore, the cord length C2 at the hub-side end portion 15a of the first-stage rotor blade 15 is 1.0 to 1.3 times the cord length C1 at the tip-side end portion 15b. According to this configuration, there is no large difference in cord length between the hub-side end portion 15a and the tip-side end portion 15b, and the blade length is limited as described above. It becomes possible to control the 15 natural frequencies to appropriate values.

Further, the length L1 of the blade root 15d of the first stage moving blade 15 in the direction of the axis O1 is 1.2 times or more and 1.5 times or less of the cord length C2 at the hub-side end portion 15a. With this configuration, the blade root 15d can be formed relatively large, so that the centrifugal stress in the blade root 15d can be reduced. As a result, it becomes possible to extend the life of the blade root 15d and the turbine disk.

(Other embodiments)
As described above, the embodiments of the present disclosure have been described in detail with reference to the drawings, but the specific configuration is not limited to these embodiments, and design changes etc. within the scope of the present disclosure are also included. .

<Appendix>
The two-shaft gas turbine described in each embodiment is understood, for example, as follows.

(1) A two-shaft gas turbine 1 according to a first aspect includes a compressor 4 having a rotating shaft rotatable around an axis O1, and a high-pressure turbine 6 having a first shaft 7 connected to the rotating shaft. , a low-pressure turbine 8 having a second shaft 9 different from the first shaft 7 and spaced from the high-pressure turbine 6 in the direction of the axis O1; and a final stage moving blade of the high-pressure turbine 6. 14 and a first stage rotor blade 15 of the low pressure turbine 8, and an intermediate passage 13 for guiding combustion gas from the high pressure turbine 6 to the low pressure turbine 8, the first stage in a radial direction with respect to the axis O1. The inner diameter of the rotor blades 15 is 1.2 times or more and 1.4 times or less that of the final stage rotor blades 14, and the length of the first stage rotor blades 15 in the radial direction is 1.0 times or more and 1.6 times. The maximum thickness of the hub-side end portion 15a, which is the radially inner end portion of the first-stage rotor blade 15, is 4 times the maximum thickness of the tip-side end portion 15b, which is the radially outer end portion. It is more than twice and less than 8 times.

According to the above configuration, the inner diameter of the first stage moving blade 15 of the low-pressure turbine 8 is 1.2 times or more and 1.4 times or less that of the last stage moving blade 14 of the high pressure turbine 6, and the length of the first stage moving blade 15 is 1.0 times or more and 1.6 times or less. If the blade height is 1.6 times or less, the uneven distribution of the combustion gas flow in the intermediate passage 13 is not increased, and an increase in pressure loss can be avoided. Furthermore, by adopting this configuration, it becomes unnecessary to increase the length of the cord on the hub side in order to avoid resonance in the vibration mode in the cord direction. Further, the maximum thickness of the hub-side end portion 15a of the first-stage rotor blade 15 is set to be four to eight times the maximum thickness of the tip-side end portion 15b, thereby increasing the cross-sectional area of the hub-side end portion 15a, It becomes possible to reduce the centrifugal stress acting on the blade 15 . Further, by increasing the maximum thickness of the hub-side end portion 15a, the curvature of the pressure surface 15p is reduced. This reduces the possibility that the combustion gas will separate from the pressure surface 15p, and the efficiency of the low-pressure turbine 8 can be further improved.

(2) In the two-shaft gas turbine 1 according to the second aspect, the cord length at the hub-side end portion 15a of the first stage rotor blade 15 is 1.0 times or more the cord length at the tip-side end portion 15b. .3 times or less.

According to the above configuration, there is no large difference in cord length between the hub-side end portion 15a and the tip-side end portion 15b, and the blade length is limited as described above. It becomes possible to control the 15 natural frequencies to appropriate values.

(3) In the two-shaft gas turbine 1 according to the third aspect, the length of the blade root 15d of the first-stage moving blade 15 in the direction of the axis O1 is 1.2 times the cord length of the hub-side end portion 15a. 1.5 times or less.

According to the above configuration, the blade root 15d can be formed relatively large, so the centrifugal stress in the blade root 15d can be reduced. As a result, it becomes possible to extend the life of the blade root 15d and the turbine disk.

(4) In the two-shaft gas turbine 1 according to the fourth aspect, the maximum thickness of the first-stage moving blade is greater at a position 75% in the radial direction than at a position 25% with respect to the hub-side end. is 2 times or more and 4 times or less.

According to the above configuration, by increasing the maximum thickness of the hub-side end portion 15a, a large cross-sectional area of the first stage moving blade 15 can be ensured. This makes it possible to further reduce the centrifugal stress.

1 2-shaft gas turbine 2 Compressor driving side turbine section (gas generator section)
3 Output side turbine section (power turbine section)
4 Compressor 5 Combustor 6 High Pressure Turbine 7 First Shaft 7a One End 8 Low Pressure Turbine 9 Second Shaft 10 Load Equipment 11 Intermediate Duct 11a Inner Pipe 11b Outer Pipe 12 Intermediate Channel Section (Intermediate Channel Section)
13 intermediate flow path 14 final stage moving blade 15 first stage moving blade 15a hub side end 15b tip side end 15c platform 15d blade root 15h blade body 16 strut 17, 18 bearing O1 axis

Claims (4)

a compressor having a rotating shaft rotatable around its axis;
a high pressure turbine having a first shaft coupled with the rotating shaft;
a low pressure turbine having a second axis different from the first axis and spaced axially relative to the high pressure turbine;
an intermediate passage provided between the last stage rotor blade of the high pressure turbine and the first stage rotor blade of the low pressure turbine for guiding combustion gas from the high pressure turbine to the low pressure turbine;
with
The inner diameter of the first stage moving blade in the radial direction with respect to the axis is 1.2 times or more and 1.4 times or less that of the last stage moving blade,
The length of the first stage rotor blade in the radial direction is 1.0 times or more and 1.6 times or less,
The maximum thickness of the hub-side end, which is the radially inner end of the first-stage rotor blade, is 4 times or more and 8 times or less than the maximum thickness of the tip-side end, which is the radially outer end. 2-shaft gas turbine. 2. The twin-shaft gas turbine according to claim 1, wherein the cord length at the hub-side end of the first-stage rotor blade is 1.0 to 1.3 times the cord length at the tip-side end. 3. The twin-shaft gas turbine according to claim 1, wherein the length in the axial direction of the blade root of the first-stage moving blade is 1.2 times or more and 1.5 times or less than the cord length at the hub-side end portion. . 4. The maximum thickness of the first-stage moving blade at a 75% position in the radial direction with respect to the hub-side end as a reference is two or more times or more and four times or less than at a 25% position. A two-shaft gas turbine according to any one of the preceding paragraphs.

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