JP5039595B2 - Gas turbine and gas turbine operation stop method - Google Patents

Description

  The present invention relates to a gas turbine and a gas turbine operation stop method.

In a conventional gas turbine, during the turning operation that is performed in the middle of the shutdown operation, the temperature of the upper part of the casing and the temperature of the lower part of the casing are measured, and the temperature difference is obtained from these temperatures, and this temperature difference is set to the set value. When reaching, the turning operation is temporarily stopped, and instead, the spin operation is performed (see, for example, Patent Document 1).
That is, in the turning operation, there is almost no stirring and flow of the fluid inside the turbine, so the metal temperature at the upper part of the casing becomes higher than the metal temperature at the lower part of the casing, and the upper part of the casing extends in the rotor axial direction compared to the lower part of the casing (cat). Back) occurs. When such a phenomenon occurs, the clearance between the lower inner surface of the casing and the tip of the moving blade decreases, and the clearance between the upper inner surface of the casing and the tip of the moving blade tends to increase. In order to avoid the phenomenon that the metal temperature difference between the upper part and the lower part of the casing exceeds the limit value and the inner surface of the lower part of the casing comes into contact with the tip of the rotor blade, spin operation is sometimes performed to stir the turbine internal fluid, The metal temperature difference is made small.
JP-A-6-2570

  However, in this spin operation, the number of rotations of the rotor, which is rotated at a very low speed of about 2 to 3 rpm by the turning motor, is increased to about 600 rpm using the starter motor, and the starter motor is driven. Therefore, there is a problem that a large amount of electric power is required to increase the running cost.

  The present invention has been made in view of the above circumstances, and an object thereof is to provide a gas turbine and a gas turbine operation stop method that can eliminate the need for a spin operation during a turning operation performed after the operation.

The present invention employs the following means in order to solve the above problems.
A gas turbine according to the present invention includes a compressor that compresses combustion air, a combustor that generates high-temperature combustion gas by injecting and burning fuel into high-pressure air sent from the compressor, and the combustion A turbine located downstream of the combustor and driven by combustion gas exiting the combustor, a compressor bleed system for directing compressor bleed from the compressor into the nozzle guide vanes of the turbine, and the compressor bleed A gas turbine comprising: a turbine bypass system that is connected to a system and guides the compressor bleed air directly to an exhaust diffuser; and a controller that controls opening and closing of a turbine bypass valve connected to the turbine bypass system. When the operation stop signal is input, the generator outputs a command signal for fully opening the turbine bypass valve, and the turbine rotation speed is determined by the nozzle guide vane and the front Once reduced to a region on a turbine blade does not cause a stall, the turbine bypass valve, to the extent that ensures air flow to the turbine and to output a command signal to a closed state.

According to the gas turbine of the present invention, the combustion gas (hot gas) remaining inside the casing is discharged to the outside of the casing together with the air supplied to the combustion gas flow path via the compressor extraction system and the nozzle guide vane. At the same time, the inside of the casing is cooled substantially uniformly by the air passing through the combustion gas flow path.
Therefore, the temperature difference between the upper and lower metal parts of the casing on the compressor side can be kept lower than the limit value from the beginning to the end during the turning operation, and the temperature difference between the upper and lower metal parts of the casing on the turbine side. Can be maintained lower than the limit value from the beginning to the end of the turning operation, and the occurrence of a cat back phenomenon can be suppressed (reduced). Spin operation during turning operation can be eliminated.

  In the gas turbine, in the center and lower half of the inscribed circle of the split ring located in the upper half of the substantially ring-shaped split ring arranged so as to surround the tip side of the blade along the circumferential direction. When the split ring is arranged such that the center of the inscribed circle of the split ring located on the vertical axis passing through the center of the casing is decentered at the same distance across the horizontal joint surface of the casing Is preferred.

  According to such a gas turbine, the center of the inscribed circle of the split ring located in the upper half and the center of the inscribed circle of the split ring located in the lower half are on the vertical axis passing through the center of the casing. Since the split rings are arranged so that they are decentered at the same distance across the horizontal joint surface of the casing, they are likely to come into contact when the catback phenomenon occurs (clearance is reduced). The clearance between the blade tip (passing through the bottom) and the inner peripheral surface of the split ring can be maintained (secured), and the contact between the blade tip and the inner peripheral surface of the split ring can be reliably prevented. Can do.

  A gas turbine shutdown method according to the present invention includes a compressor that compresses combustion air, and a combustor that generates high-temperature combustion gas by injecting fuel into high-pressure air sent from the compressor and burning the fuel. A turbine located downstream of the combustor and driven by combustion gas exiting the combustor, and a compressor bleed system for directing compressor bleed from the compressor into the turbine nozzle guide vanes; A gas turbine operation stopping method comprising a turbine bypass system connected to the compressor bleed system and directly leading the compressor bleed gas to an exhaust diffuser, and a turbine bypass valve connected to the turbine bypass system, The fuel supplied to the combustor is cut, and at the same time, the turbine bypass valve is fully opened, and the rotational speed of the turbine is such that the nozzle guide vane and Once reduced to a region on the serial turbine blade does not cause a stall, the turbine bypass valve, to the extent that ensures air flow rate to the turbine is intended to be closed.

According to the gas turbine operation stop method according to the present invention, the combustion gas (hot gas) remaining in the casing, together with the air supplied to the combustion gas flow path via the compressor bleed system and the nozzle guide vane, is casing. The inside of the casing is cooled substantially uniformly by the air passing through the combustion gas flow path.
Therefore, the temperature difference between the upper and lower metal parts of the casing on the compressor side can be kept lower than the limit value from the beginning to the end during the turning operation, and the temperature difference between the upper and lower metal parts of the casing on the turbine side. Can be maintained lower than the limit value from the beginning to the end of the turning operation, and the occurrence of a cat back phenomenon can be suppressed (reduced). Spin operation during turning operation can be eliminated.

  According to the gas turbine and the gas turbine operation stop method according to the present invention, it is possible to eliminate the need for the spin operation during the turning operation performed after the operation.

Hereinafter, a first embodiment of a gas turbine according to the present invention will be described with reference to FIGS. 1 to 4.
FIG. 1 is a schematic system diagram of a gas turbine according to the present embodiment, and FIGS. 2 to 4 are graphs showing test results carried out using the gas turbine according to the present invention.

  As shown in FIG. 1, a gas turbine 1 according to this embodiment includes a compressor 2 that compresses combustion air, and fuel is injected into high-pressure air that is sent from the compressor 2 to burn it. A combustor (not shown) that generates combustion gas, a turbine 3 that is located downstream of the combustor and that is driven by the combustion gas exiting the combustor, a compressor discharge air system 4, and an intermediate compressor A stage bleed system (compressor bleed system) 5 is provided.

The compressed air compressed by the compressor 2 is supplied to a combustor and mixed with a separately supplied fuel (for example, liquefied natural gas (LNG)) and burned. The combustion gas generated by this combustion is supplied to the turbine 3 and causes the turbine 3 to generate a rotational driving force.
The turbine 3 includes a turbine stator 6 that is a stationary part and a turbine rotor 7 that is a rotating part.
The turbine stator 6 is generally called a turbine nozzle, and has nozzle guide vanes (static blades) 6a having a blade-shaped cross section arranged in an annular shape. The turbine stator 6 expands and depressurizes the gas, and also sets the flow direction so that the outflow gas from the nozzle guide vane 6a collides with the blade (moving blade) 7a of the turbine rotor 7 at an optimum angle. Has the function of giving.
The turbine rotor 7 is configured such that a plurality of blades 7a having an airfoil cross section are attached on the outer periphery of a disk-shaped turbine disk 7b.

  The nozzle guide vanes 6a and the blades 7a are alternately arranged along the axial direction (left and right direction in FIG. 1) of the turbine rotor 7, and are adjacent to each other between the nozzle guide vanes 6a adjacent in the circumferential direction and in the circumferential direction. Combustion gas passages 8 are formed between the blades 7a.

  The discharge air (compressed air) guided from the compressor 2 to the inside of the first stage nozzle guide vane 6a of the turbine 3 through the compressor discharge air system 4 is cooled after the first stage nozzle guide vane 6a is cooled. The combustion gas that passes through the gas flow path 8 merges.

Discharged air (compressed air) introduced from the compressor 2 into the nozzle guide vanes 6a in the second and subsequent stages of the turbine 3 through the compressor intermediate stage extraction system 5 passes through the nozzle guide vanes 6a in the second and subsequent stages. After cooling, it merges with the combustion gas passing through the combustion gas flow path 8.
A turbine bypass system 10 including a turbine bypass valve 9 is connected to the compressor intermediate stage bleed system 5. The turbine bypass valve 9 is an on-off valve that is opened and closed by a command signal from the controller 11. When the turbine bypass valve 9 is opened, the compressor intermediate stage bleed air bypasses the turbine 3 and is directly guided to the exhaust diffuser 12. It is like that.
Although FIG. 1 shows only the fourth stage nozzle guide vane 6a and the fourth stage blade 7a, this does not limit the present invention.

In the gas turbine 1 according to the present embodiment, when an operation stop signal is input to the controller 11, the fuel supplied to the combustor is cut, and at the same time, a command is sent from the controller 11 to the turbine bypass valve 9. While the signal is output, the turbine bypass valve 9 is fully opened, and the compressor intermediate stage bleed air is guided to the exhaust diffuser 12 through the compressor intermediate stage bleed system 5 and the turbine bypass system 10.
Next, when the rotational speed of the turbine rotor 7 decreases (decreases) to a region where the nozzle guide vane 6a and the blade 7a are not stalled or surging (for example, about 10% of the constant speed rotational speed). The signal from a detection means (not shown) (for example, a rotation sensor) that detects this is input to the controller 11. At the same time, a command signal is output from the controller 11 to the turbine bypass valve 9, and the turbine bypass valve 9 is closed to such an extent that an air flow rate to the turbine 3 can be secured. Thus, the compressor intermediate stage bleed air is guided into the nozzle guide vanes 6 a in the second and subsequent stages of the turbine 3.

According to the gas turbine 1 according to the present embodiment, in the turbine bypass valve 9 that is fully opened at the same time as the operation stop, the rotational speed of the turbine rotor 7 does not cause stall or surging in the nozzle guide vanes 6a and the blades 7a. When it is lowered (decreased) to the region, it is closed to such an extent that the air flow rate to the turbine 3 can be secured.
That is, the combustion gas (hot gas) remaining inside the casing is discharged to the outside of the casing together with the air supplied to the combustion gas flow path 8 through the compressor intermediate stage extraction system 5 and the nozzle guide vane 6a. The interior of the casing is cooled substantially uniformly by the air passing through the combustion gas flow path 8.
Therefore, as shown by a thick solid line in FIG. 2, the temperature difference between the upper and lower metal of the casing on the compressor 2 side can be kept lower than the limit value from the beginning to the end (over) during the turning operation. 3, the temperature difference between the upper and lower casing metals on the turbine 3 side can be maintained lower than the limit value from the beginning to the end during the turning operation, as shown by a thick solid line in FIG. Since the occurrence of the cat back phenomenon can be suppressed (reduced), the spin operation during the turning operation performed after the operation can be made unnecessary, as indicated by the one-dot chain line in FIG.
The thin solid line in FIG. 2, the thin solid line in FIG. 3, and the broken line in FIG. 4 indicate the turning operation without closing the turbine bypass valve 9 and without performing the spin operation. The test results when only the test is performed are shown.

A second embodiment of the gas turbine according to the present invention will be described with reference to FIGS. 5 and 6.
5 is an enlarged cross-sectional view of a main part of the gas turbine according to the present embodiment, and is a view seen in a cross section including the axis of the turbine rotor. FIG. 6 is a main part perspective view showing an assembled state of the nozzle guide vane shown in FIG. FIGS. 7A and 7B are views seen in a cross section orthogonal to the axis of the turbine rotor, where FIG. 7A is a view showing the arrangement of a heat shield ring of an annular heat shield ring type, and FIG. 7B is an oval heat shield. It is a figure which shows arrangement | positioning of the ring-type heat shield ring.

  The gas turbine 21 according to the present embodiment is different from that of the first embodiment described above in that it includes a heat shield ring 22. Since other components are the same as those of the first embodiment described above, description of these components is omitted here.

The heat shield ring 22 is a member formed to have an annular shape in a plane orthogonal to the axis of the turbine rotor 7, and surrounds the tip (tip) side of the blade 7 a around the circumferential direction. The ring-shaped divided ring 23 arranged on the support ring is supported and fixed to the blade ring 24. The heat shield ring 22 also supports and fixes the outer shroud 25 formed on the radially outer side of the nozzle guide vane 6 a with respect to the blade ring 26.
The heat shield ring 22 is divided into a plurality (20 in this embodiment) along the circumferential direction.
Moreover, the white arrow in FIG. 5 has shown the flow direction of combustion gas.

Now, the arrangement of the heat shield ring 22 and the split ring 23 in the gas turbine 21 according to the present embodiment will be described.
There are two methods for arranging the heat shield ring 22, an annular heat shield ring method and an oval heat shield ring method.
In the annular heat shield ring system, as shown in FIG. 7A, the heat shield ring 22 is arranged so as to draw a circle with a radius R inscribed in the split ring 23 with the center O of the passenger compartment as the center. Refers to cases. The center O of the casing refers to the center of the casing in a cross section perpendicular to the turbine rotor axis, and corresponds to the point where the horizontal joint surface A, which is a horizontal dividing surface of the casing, and the vertical axis B intersect.
On the other hand, in the oval type heat shield ring system, as shown in FIG. 7B, a heat shield ring (hereinafter referred to as “upper half heat shield ring”) 31 in which the heat shield ring 22 is located in the upper half portion. When divided into a heat shield ring (hereinafter referred to as “lower half heat shield ring”) 32 located in the lower half, the center positions of the upper half heat shield ring 31 and the lower half heat shield ring 32 are The vertical axis direction is shifted from each other. That is, the center O1 of the inscribed circle of the split ring 23 of the upper half heat shield ring 31 and the center O2 of the inscribed circle of the split ring 23 of the lower half heat shield ring 32 are both on the vertical axis B. In this case, the horizontal joint surfaces of the casing (which are the same as the center O of the casing) A are arranged so as to be decentered by the same distance T from the center O of the casing. Here, the upper half heat shield ring 31 and the lower half heat shield ring 32 are a plurality of heat shield rings 22 arranged above or below the horizontal joint surface A with the horizontal joint surface A of the casing as a boundary. Means a heat shield ring group (in the case of this embodiment, each of which is composed of 10 heat shield rings). Normally, the upper half heat shield ring 31 and the lower half heat shield ring 32 are attached and adjusted so that the predetermined length T is 1 mm or less. Further, in the oval type heat shield ring system, the center of the turbine rotor is located at the same position as the inscribed circle center O1 of the upper half heat shield ring 31 and is slightly decentered upward in the vertical axis B with respect to the center O of the casing. I am letting.

  With this arrangement of the heat shield ring 22, the clearance between the casing lower inner surface and the tip of the rotor blade can be set larger than in the conventional annular heat shield ring system.

  According to the gas turbine 21 according to the present embodiment, when the catback phenomenon occurs, the tip of the blade 7a and the inner peripheral surface of the split ring 23 that are likely to come into contact with each other (the clearance becomes smaller) located at the lower portion (bottom) of the casing. The clearance between the tip of the blade 7a and the inner peripheral surface of the split ring 23 can be reliably prevented. That is, as shown in FIGS. 2 and 3, substantially the same effect can be obtained as when the limit value of the temperature difference between the upper and lower metal parts of the casing on the compressor 2 side and the turbine 3 side is increased.

  Note that the present invention is not limited to the above-described embodiments, and modifications, changes, and combinations can be appropriately made as necessary without departing from the technical idea of the present invention.

1 is a schematic system diagram of a gas turbine according to a first embodiment of the present invention. It is a graph which shows the test result implemented using the gas turbine which concerns on this invention. It is a graph which shows the test result implemented using the gas turbine which concerns on this invention. It is a graph which shows the test result implemented using the gas turbine which concerns on this invention. It is the principal part expanded sectional view of the gas turbine which concerns on 2nd Embodiment of this invention, Comprising: It is the figure seen in the cross section containing the axis line of a turbine rotor. It is a principal part perspective view which shows the assembly state of the nozzle guide vane shown in FIG. It is the figure seen in the cross section orthogonal to the axis line of a turbine rotor, Comprising: (a) is a figure which shows arrangement | positioning of the heat shield ring of a ring type heat shield ring system, (b) is the heat shield of an oval type heat shield ring system It is a figure which shows arrangement | positioning of a ring.

Explanation of symbols

1 Gas Turbine 2 Compressor 3 Turbine 5 Compressor Intermediate Stage Extraction System (Compressor Extraction System)
6a Nozzle guide vane 7a Blade 9 Turbine bypass valve 10 Turbine bypass system 11 Controller 12 Exhaust diffuser 21 Gas turbine 22 Heat shield ring 23 Split ring 24 Blade ring

Claims (3)

A compressor that compresses combustion air, a combustor that injects and burns fuel into high-pressure air sent from the compressor, and generates high-temperature combustion gas; and a downstream side of the combustor, A turbine driven by the combustion gas exiting the combustor; a compressor bleed system for directing compressor bleed from the compressor into the nozzle guide vanes of the turbine; and the compressor bleed system connected to the compressor A gas turbine comprising a turbine bypass system that directly leads the machine bleed air to the exhaust diffuser, and a controller that controls opening and closing of a turbine bypass valve connected to the turbine bypass system,
When the operation stop signal is input, the controller outputs a command signal for fully opening the turbine bypass valve, and the rotational speed of the turbine does not cause the nozzle guide vane and the turbine blade to stall. A gas turbine that outputs a command signal that closes the turbine bypass valve to an extent that can secure an air flow rate to the turbine when the pressure falls to a region.   Of the substantially annular split ring arranged so as to surround the tip side of the blade in the circumferential direction, the center of the inscribed circle of the split ring located in the upper half and the split ring located in the lower half The split ring is arranged such that the center of the inscribed circle is eccentric to each other at the same distance across the horizontal joint surface of the casing on a vertical axis passing through the center of the casing. The gas turbine according to 1. A compressor that compresses combustion air, a combustor that injects and burns fuel into high-pressure air sent from the compressor, and generates high-temperature combustion gas; and a downstream side of the combustor, A turbine driven by the combustion gas exiting the combustor; a compressor bleed system for directing compressor bleed from the compressor into the nozzle guide vanes of the turbine; and the compressor bleed system connected to the compressor A gas turbine operation stop method comprising a turbine bypass system for directing machine bleed air directly to an exhaust diffuser, and a turbine bypass valve connected to the turbine bypass system,
When the fuel supplied to the combustor is cut, and the turbine bypass valve is fully opened at the same time, and the rotational speed of the turbine decreases to a region where the nozzle guide vanes and the blades of the turbine are not stalled, A gas turbine operation stopping method, wherein the turbine bypass valve is closed to such an extent that an air flow rate to the turbine can be secured.

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