JP7558128B2 - Two-shaft gas turbine - Google Patents

Description

This disclosure relates to a two-shaft gas turbine.

Two-shaft gas turbines have been known in the past, in which the high-pressure turbine and the low-pressure turbine are arranged on separate shafts, and the combustion gas that has passed 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 from this type of two-shaft gas turbine. Here, when using a two-shaft gas turbine, especially for power generation, the rotation speed of the low-pressure turbine side needs to be kept constant (3000 rpm as an example). In order to improve the output without changing the rotation speed, it is possible to increase the average diameter of the moving blades (first stage moving blades) of the low-pressure turbine. If the average diameter of the low-pressure turbine is increased, the flow path cross-sectional area of the low-pressure turbine also increases, so the flow velocity of the combustion gas decreases. On the other hand, if the high-pressure turbine, which is difficult to change in design due to its high temperature, is not improved, the diameter difference between the high-pressure turbine and the low-pressure turbine increases, and the intermediate flow path connecting the two becomes curved. As a result, the flow velocity distribution in the radial direction of the moving blade becomes non-uniform, and the flow loss also increases at the same time. In addition, the centrifugal stress on the moving blade increases as the average diameter increases. As a configuration for reducing such centrifugal stress, for example, the one described in the following Patent Document 1 is known.

The rotor blade described in the following Patent Document 1 is configured so that the direction in which the stacking lines of the rotor blade extend is different between the radially outer portion and the radially inner portion.

International Publication No. 2020/095470

However, even if the shape of the rotor blades is changed as described above, it is difficult to reduce the centrifugal stress on the hub side. As a result, there is a problem in that it is not possible to achieve a higher peripheral speed of the low-pressure turbine.

This disclosure has been made to solve the above problems, and aims to provide a two-shaft gas turbine that can achieve even higher peripheral speeds.

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 connected to the rotating shaft, a low-pressure turbine having a second shaft different from the first shaft and disposed at a distance from the high-pressure turbine in the axial direction, and an intermediate flow passage provided between a final stage rotor blade of the high-pressure turbine and a first stage rotor blade of the low-pressure turbine, for guiding combustion gas from the high-pressure turbine to the low-pressure turbine, wherein an inner diameter of the first stage rotor blade in a radial direction relative to the axis is 1.2 to 1.4 times the inner diameter of the final stage rotor blade in the radial direction, a length of the first stage rotor blade in the radial direction is 1.0 to 1.6 times the length of the final stage rotor blade in the radial direction, and the radial length of the first stage rotor blade is 1.2 to 1.6 times the length of the final stage rotor blade in the radial direction . a maximum thickness at a hub side end which is an end on the radially inner side is 4 to 8 times the maximum thickness at a tip side end which is an end on the radially outer side, a chord length at the hub side end of the first stage blade is 1.0 to 1.3 times the chord length at the tip side end, a length in the axial direction of a blade root of the first stage blade is 1.2 to 1.5 times the chord length at the hub side end, a maximum thickness at a 75% position in the radial direction of the first stage blade is 2 to 4 times the maximum thickness at a 25% position in the radial direction of the first stage blade, when the radial position of the tip side end of the first stage blade is taken as 100% with respect to the hub side end of the first stage blade , and the intermediate flow passage is curved due to a diameter difference between the low-pressure turbine and the high-pressure turbine.

This disclosure makes it possible to provide a two-shaft gas turbine that can achieve even higher peripheral speeds.

FIG. 1 is a schematic diagram illustrating a 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 portion of a two-shaft gas turbine according to an embodiment of the present disclosure. FIG. 2 is a radial view of a first stage rotor blade of a low-pressure turbine according to an embodiment of the present disclosure. FIG. 2 is a circumferential view of a first stage rotor blade of a low-pressure turbine according to an embodiment of the present disclosure.

(Configuration of two-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 Fig. 1 to Fig. 4. The two-shaft gas turbine according to this embodiment is suitable for use as a heavy-structure gas turbine for various industrial machines, power generation devices, etc. The two-shaft gas turbine according to this embodiment can also be used as a gas turbine for other applications, such as an aircraft (aircraft engine).

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

The compressor drive turbine section 2 includes a compressor 4 that compresses air R1 taken in from the atmosphere to generate compressed air R2, a combustor 5 that mixes fuel with the compressed air R2 sent from the compressor 4 and burns it to generate 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 the rotor of the high-pressure turbine 6. In the compressor drive 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 via the first shaft 7 to drive the compressor 4. The first shaft 7 also serves as the rotor of the high-pressure turbine 6.

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

The output turbine section 3 has a low-pressure turbine 8. The low-pressure turbine 8 and the load equipment 10 are connected via a second shaft (power turbine shaft) 9, which 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 sent from the high-pressure turbine 6, whose pressure has been reduced by driving the high-pressure turbine 6. The rotational power obtained by the low-pressure turbine 8 is transmitted to the load equipment 10, which drives the load equipment 10. The combustion gas R5 that drives the low-pressure turbine 8 is discharged as exhaust gas.

As shown in Figures 1 and 2, between the compressor drive turbine section 2 and the output turbine section 3, i.e., between the high-pressure turbine 6 and the low-pressure turbine 8 in the direction of the axis O1, there is provided an intermediate flow passage section (intermediate flow passage section) 12 equipped with an intermediate duct 11 (intermediate flow passage 13) for feeding combustion gas R4 from the high-pressure turbine 6 to the low-pressure turbine 8. The intermediate duct 11 has a circular double-pipe structure equipped with an inner pipe 11a and an outer pipe 11b arranged coaxially along the axis O1. The space between the inner pipe 11a and the outer pipe 11b of the intermediate duct 11 is the intermediate flow passage 13 for circulating the combustion gas R4.

The intermediate flow passage 13 (intermediate duct 11) is provided between the final stage moving blades 14 of the high-pressure turbine 6 and the first stage moving blades 15 of the low-pressure turbine 8 in the direction of the axis O1, and supplies combustion gas R4 from the high-pressure turbine 6 to the low-pressure turbine 8. In addition, in the two-shaft gas turbine 1 of this embodiment, the struts 16 arranged in the intermediate flow passage 13 are configured to also serve as the first stage stator blades of the low-pressure turbine 8. Multiple struts 16 are arranged radially at intervals in the circumferential direction of the axis O1 within the intermediate flow passage 13.

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

The blade body 15h extends radially relative to the axis O1 and has an airfoil-shaped cross-sectional shape when viewed from the radial direction. As shown in FIG. 3, the radially outer end of the blade body 15h is the tip end 15b, and the radially inner end is the hub end 15a. Of the two circumferential surfaces of the blade body 15h with an airfoil-shaped cross section, the surface that is concave toward one circumferential side is the pressure surface 15p, and the surface facing the opposite side is curved convexly toward one circumferential side to form the negative pressure surface 15n.

The platform 15c is provided at the radially inner end of the blade body 15h. The dimension of the platform 15c in the axis O1 direction is larger than the dimension of the blade body 15h in the axis O1 direction. 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 multiple serration-like projections and recesses. Here, as shown in FIG. 4, when the chord length at the hub side end 15a is C2 and the dimension of the blade root 15d in the axis O1 direction is L1, the value of L1 is 1.2 times or more and 1.5 times or less of C2.

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 set to be 1.0 to 1.6 times w1. Furthermore, when the inner diameter (radial dimension based on the axis O1) of the final stage rotor blade 14 is d1 and the inner diameter of the first stage rotor blade 15 is d2, the value of d2 is set to be 1.2 to 1.4 times d1.

In addition, as shown in FIG. 3, the chord length C2 of the blade body 15h at the hub end 15a is 1.0 to 1.3 times the chord length at the tip end 15b. Furthermore, when the maximum thickness of the blade body 15h at the hub end 15a is D2 and the maximum thickness at the tip end 15b is D1, the value of D2 is 4 to 8 times D1. Furthermore, at the 75% position in the radial direction based on the hub end 15a, the maximum thickness of the blade body 15h is 2 to 4 times the maximum thickness at the 25% position.

(Action and Effect)
In recent years, there has been an increasing demand for a two-shaft gas turbine 1 with higher output. Here, when the two-shaft gas turbine 1 is used for power generation, the rotation speed of the low-pressure turbine 8 must be kept constant (for example, 3000 rpm). In order to improve the output without changing the rotation speed, it is possible to increase the average diameter of the moving blades (first-stage moving blades 15) of the low-pressure turbine 8. When the average diameter is increased, the flow path cross-sectional area of the low-pressure turbine 8 also increases, so the flow velocity of the combustion gas decreases. On the other hand, the diameter difference with the high-pressure turbine 6 increases, and the flow path curves as shown in FIG. 3. As a result, the flow velocity distribution in the radial direction of the first-stage moving blades 15 becomes non-uniform, and the flow loss increases. In addition, the centrifugal stress acting on the moving blades increases as the average diameter increases.

The two-shaft gas turbine 1 according to this embodiment has the above-mentioned dimensional conditions for the first stage rotor blade 15. According to the above-mentioned configuration, the inner diameter d2 of the first stage rotor blade 15 is set to be 1.2 to 1.4 times the inner diameter d1 of the final stage rotor blade 14 of the high-pressure turbine 6. Furthermore, the length w2 of the first stage rotor blade 15 is set to be 1.0 to 1.6 times the length w1 of the final stage rotor blade 14. If the length of the first stage rotor blade 15 is kept at 1.6 times or less, there is little bias in the distribution of the combustion gas flow in the intermediate flow passage 13, and therefore the increase in pressure loss can be mitigated compared to when the blade length is increased further.

Furthermore, by adopting this configuration, the length of the first stage moving blade 15 does not become excessively large, and the constraints on the natural frequency are relaxed. In general, in order to deal with an increase in centrifugal stress, measures are often taken to increase the cross-sectional area by expanding the chord length on the hub side, thereby reducing stress. This is because in many cases, the natural first mode becomes a vibration mode in the chord direction, and resonance with a low-order excitation source becomes a problem, so it is necessary to ensure the natural frequency as much as possible. On the other hand, in this embodiment, the upper limit of the length of the first stage moving blade 15 is restricted, so that the natural frequency is already sufficiently high, and it is not necessary to expand the chord length. Instead of expanding the chord length, a measure is taken to ensure the cross-sectional area on the hub side by making the maximum thickness D2 of the hub side end 15a of the first stage moving blade 15 4 to 8 times the maximum thickness D1 of the tip side end 15b. This makes it possible to expand the cross-sectional area at the hub side end 15a while suppressing the blade weight, and to reduce the centrifugal stress acting on the hub cross section of the first stage moving blade 15.

In addition, by increasing the maximum thickness of the hub-side end 15a, the curvature of the pressure surface 15p is reduced. This reduces the possibility of the combustion gas separating from the pressure surface 15p, further improving the efficiency of the low-pressure turbine 8.

Furthermore, the chord length C2 at the hub end 15a of the first stage rotor blade 15 is 1.0 to 1.3 times the chord length C1 at the tip end 15b. With this configuration, there is no significant difference in chord length between the hub end 15a and the tip end 15b, and since the blade length is limited as described above, it is possible to control the natural frequency of the first stage rotor blade 15 to an appropriate value.

The length L1 of the blade root 15d of the first stage rotor blade 15 in the direction of the axis O1 is 1.2 to 1.5 times the chord length C2 at the hub-side end 15a. With this configuration, the blade root 15d can be formed relatively large, which reduces the centrifugal stress in the blade root 15d. As a result, it is possible to extend the life of the blade root 15d and the turbine disk.

Other Embodiments
Although the embodiments of the present disclosure have been described in detail above with reference to the drawings, the specific configuration is not limited to this embodiment, and design changes and the like that do not deviate from the gist of the present disclosure are also included.

<Additional Notes>
The two-shaft gas turbine described in each embodiment can be understood, for example, as follows.

(1) The two-shaft gas turbine 1 according to the first aspect includes a compressor 4 having a rotating shaft rotatable around an axis O1, 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 arranged at a distance from the high-pressure turbine 6 in the direction of the axis O1, and a rotor blade 14 of the high-pressure turbine 6 and a rotor blade 15 of the low-pressure turbine 8 arranged between the final stage rotor blade 14 of the high-pressure turbine 6 and the first stage rotor blade 15 of the low-pressure turbine 8, and a rotor blade 16 of the high-pressure turbine 6 and a rotor blade 17 of the low-pressure turbine 8. and an intermediate flow passage 13 that guides the combustion gas to the turbine 8. The inner diameter of the first stage rotor blade 15 in the radial direction relative to the axis O1 is 1.2 to 1.4 times that of the final stage rotor blade 14, the length of the first stage rotor blade 15 in the radial direction is 1.0 to 1.6 times, and the maximum thickness of the hub side end 15a, which is the radially inner end of the first stage rotor blade 15, is 4 to 8 times the maximum thickness of the tip side end 15b, which is the radially outer end.

According to the above configuration, the inner diameter of the first stage moving blade 15 of the low pressure turbine 8 is 1.2 to 1.4 times that of the final stage moving blade 14 of the high pressure turbine 6, and the length of the first stage moving blade 15 is 1.0 to 1.6 times. If the blade height is 1.6 times or less, the distribution of the combustion gas flow in the intermediate flow passage 13 does not become biased, and an increase in pressure loss can be avoided. Furthermore, by adopting this configuration, it is not necessary to increase the chord length on the hub side to avoid resonance of the vibration mode in the chord direction. Furthermore, the maximum thickness of the hub side end 15a of the first stage moving blade 15 is set to 4 to 8 times the maximum thickness of the tip side end 15b, and the cross-sectional area at the hub side end 15a is increased, thereby reducing the centrifugal stress acting on the first stage moving blade 15. In addition, by increasing the maximum thickness of the hub side end 15a, the curvature of the pressure surface 15p is reduced. This reduces the possibility of the combustion gas separating from the pressure surface 15p, further improving the efficiency of the low-pressure turbine 8.

(2) In the two-shaft gas turbine 1 according to the second aspect, the chord length at the hub end 15a of the first stage rotor blade 15 is 1.0 to 1.3 times the chord length at the tip end 15b.

With the above configuration, there is no significant difference in chord length between the hub end 15a and the tip end 15b, and the blade length is limited as described above, making it possible to control the natural frequency of the first stage rotor blade 15 to an appropriate value.

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

The above configuration allows the blade root 15d to be formed relatively large, which reduces the centrifugal stress in the blade root 15d. As a result, it is 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 rotor blade is between two and four times as large at a 75% position in the radial direction based on the hub end compared to a 25% position.

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

1 Two-shaft gas turbine 2 Compressor drive turbine section (gas generator section)
3. Output turbine section (power turbine section)
Reference Signs List 4 Compressor 5 Combustor 6 High-pressure turbine 7 First shaft 7a One end 8 Low-pressure turbine 9 Second shaft 10 Load device 11 Intermediate duct 11a Inner pipe 11b Outer pipe 12 Intermediate flow passage section (Intermediate flow passage section)
13 Intermediate flow passage 14 Last stage rotor blade 15 First stage rotor 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 (1)

A compressor having a rotating shaft that is rotatable around an axis line;
a high-pressure turbine having a first shaft connected to the rotating shaft;
a low pressure turbine having a second shaft different from the first shaft and spaced apart from the high pressure turbine in the axial direction;
an intermediate flow passage provided between a final stage rotor blade of the high-pressure turbine and a first stage rotor blade of the low-pressure turbine, for guiding combustion gas from the high-pressure turbine to the low-pressure turbine;
Equipped with
an inner diameter of the first stage rotor blade in a radial direction relative to the axis is 1.2 times or more and 1.4 times or less an inner diameter of the final stage rotor blade in the radial direction ;
a length of the first stage rotor blade in the radial direction is 1.0 times or more and 1.6 times or less of a length of the final stage rotor blade in the radial direction ,
a maximum thickness of a hub side end portion, which is an end portion on the radial inner side of the first stage blade, is 4 to 8 times a maximum thickness of a tip side end portion, which is an end portion on the radial outer side,
a chord length at the hub side end of the first stage blade is 1.0 times or more and 1.3 times or less than a chord length at the tip side end,
a length of the blade root of the first stage blade in the axial direction is 1.2 times or more and 1.5 times or less of a chord length at the hub side end,
when the radial position of the tip side end of the first stage rotor blade is taken as 100% with respect to the hub side end of the first stage rotor blade, a maximum thickness at a 75% position in the radial direction of the first stage rotor blade is between two and four times a maximum thickness at a 25% position in the radial direction of the first stage rotor blade ,
A two-shaft gas turbine , wherein the intermediate flow passage is curved due to a difference in diameter between the low-pressure turbine and the high-pressure turbine .

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