JPS5925090B2 - Shroud for gas turbine - Google Patents

Description

DETAILED DESCRIPTION OF THE INVENTION The present invention relates to a shroud for a gas turbine, and is applied to a cast turbine having cooling air holes provided in the rotor blades.

FIG. 1 shows the configuration around the combustor of a conventionally commonly used gas turbine.

The compressor section 1 pressurizes atmospheric air and feeds it under pressure to the turbine casing 3 as supply air 2 .

A combustor outer cylinder 4 is attached to the turbine casing, and a combustor inner cylinder 5 is provided therein.

The supply air 2 flows in the direction of the arrows A and B and flows into the combustor inner cylinder 5, and together with this, fuel is injected from the fuel nozzle 6 in the convex direction of the arrow, and is ignited by the spark plug 7 to be combusted.

The combustion gas flows within the transition case 8 in the two directions of the arrows, and flows into the turbine 9 to drive it.

Figure 2 is a diagram showing details of the turbine section, showing the combustion gas flow 10
is rectified by the nozzle 11 and speeded up to the front rotor blade 12.
It collides with the rotor blade 14 to impart rotational force thereto, and is further rectified by the nozzle 13 and collides with the rear rotor blade 14 to impart rotational force thereto.

FIG. 3 is a detailed view of the vicinity of the front first-stage rotor blade 12. In order to reduce the leakage of combustion gas from the gap between the rotor blade 12, which is a rotating body, and the casing 3, which is a stationary body, the casing 3 is
A shroud 15 is provided on the inner peripheral surface of the shroud 15.
The inner circumferential surface 15a of the rotor blade 12 and the tip surface 16 of the rotor blade 12 are opposed to each other with an appropriate gap interposed therebetween, but setting this gap poses a difficult technical problem as described below.

In other words, if this gap is made larger, the leakage of combustion gas will increase and the performance of the gas turbine will be degraded.On the other hand, if this gap is made small, the rotor blades 12 will be affected by thermal expansion and other factors.
There is a risk that the tip end surface 16 of the shroud 15 may come into contact with the inner circumferential surface 15a of the shroud 15 and be damaged.

FIG. 4 is a graph showing the relationship between the gap size δ and the efficiency reduction rate of the gas turbine, and it is clearly shown that the efficiency decrease increases in proportion to the increase in the gap size δ.

FIG. 5A is an explanatory diagram regarding the thermal expansion of the rotor blade 12 and the shroud 15. The opening in the figure is a main view with a cross section X-, and FIG.

Since the rotor blades 12 must rotate at high speed in a high-temperature gas flow of nearly 1000° C. and withstand large centrifugal forces, they are made of high-grade heat-resistant metal, and because of their relatively large coefficient of thermal expansion, the temperature rise during operation is also large.

As shown in cross section X- in the opening of FIG. 5, it is relatively thin, and in particular, the gas outlet side 18 is thinner than the gas inlet side 17, so that the temperature tends to be close to that of the combustion gas.

In this way, a slight degree of collar gradient also occurs inside the rotor blade.

On the other hand, since the shroud 15 is a stationary body and is not subject to separate centrifugal force, it is also desirable that it also serve as a heat insulating material for the casing 3, so a material with relatively poor thermal conductivity is used.

As a result, the temperature rise in the casing 3 is suppressed to approximately 300℃,
The casing 3 has less thermal expansion than the rotor blades 12.

Therefore, the amount of increase in the diameter of the outer circumferential surface of the shroud 15 connected to the inner circumferential surface of the casing 3 due to thermal expansion is greater than the amount of increase due to thermal expansion in the diameter of the rotation trajectory drawn by the rotor blade tip surface 16. Noticeably small.

FIG. 5C shows a cross section of a rotor blade in which cooling air holes are bored, and even in a rotor blade having such air cooling holes, its thermal expansion is significantly larger than that of the turbine casing 3.

In actual design and manufacturing, the diameter of the tip surface 16 of the rotor blade 12 has an elongation δ due to thermal expansion compared to when it is stationary at room temperature.
1 and elongation δ2 due to centrifugal force, and further taking into account elongation δ3 due to thermal expansion of the turbine casing 3 and elongation δ4 due to thermal expansion of the shroud, the following δ1 and δ are calculated.

Calculate by taking into account.

In other words, as mentioned above, the rotor blades 12 (inlet side 11 and outlet side 1
8, the elongation δ1 due to thermal expansion of the rotor blade is divided into the elongation δ1 due to thermal expansion on the inlet side and the elongation δ1 due to thermal expansion on the exit side.

Thermal expansion calculations are performed separately.

However, as a practical matter, it is extremely difficult to accurately calculate the amount of elongation due to thermal expansion of the casing 3 and the amount of elongation due to thermal expansion of the rotor blade 12, and furthermore, when operating for a long period of time, deformation occurs in the casing 3 etc. over time. Therefore, the above design calculation becomes even more difficult.

Due to these circumstances, the gap between the tip 16 of the rotor blade 12 and the inner circumferential surface 16 of the shroud 15 has to be made slightly larger in consideration of safety. (See FIG. 4) is unavoidable, so there has been a strong desire for an invention that can suppress the aforementioned thermal expansion, as well as an invention that can prevent leakage of combustion gas from the gap to prevent a decrease in efficiency.

An object of the present invention is to provide a gas turbine having rotor blades provided with cooling air holes as shown in FIG. An object of the present invention is to provide a shroud for a gas turbine, which improves the accuracy of design calculations by suppressing the elongation δ4 due to thermal expansion of the shroud, and which can prevent combustion gas from leaking from the gap by forming an air wall with a cooling air flow. It is something.

Next, an embodiment of the present invention will be described with reference to FIG. 6, FIG.

As shown in FIG. 6, the turbine casing 2 is a stationary body.
A front-stage nozzle 21 and a rear-stage nozzle 22, which are stationary bodies, are incorporated in the rotor 0, and a rotor blade 23, which is a rotary body, is located between the front-stage nozzle and the rear-stage nozzle.

A shroud 24 is fixed to the inner peripheral surface of the turbine casing 20 by a key-shaped groove 25.

This shroud 24 has a plurality of cylindrical shroud segments 24a, 524b, 24c, etc. as shown in FIG.
It is divided into ... etc. and installed consecutively.

The continuously installed shroud segments form a ring around the inner surface of the turbine casing 20, and FIG. 7A is a front view showing three of them, 24at24btt24e.

As shown in FIG. 7A, the inner circumferential surface 25 of the shroud segment has grooves 27 parallel to the turbine axis direction.
27.

27 is formed.

The length Ll of this groove 27 is the length Ll of the shroud segment 24a.
, 24b... width L2 in the turbine axial direction
It is suitably shorter than the shroud segment, and does not reach either the inlet end 28a or the outlet end 28b of the shroud segment.

In other words, a grooveless portion having a length L3 is left at both ends.

In this way, the grooves 27 are arranged as parallel grooves at equal intervals all over the inner circumference of the shroud 24.

FIG. 7 is a side view of the portion shown in FIG. 7A.

FIG. 8A is a sectional view showing the relationship between the shroud according to the present invention and the rotor blades in a gas turbine to which the shroud according to the present invention is applied, and the rotor blade 33 used in this embodiment has air holes for cooling. It is of the type with 34 holes.

Cooling air 35 is sent from the root of the rotor blade 33 and flows out from the tip.

The opening in Figure 8 is a cross-sectional z-z' perspective view.

The combustion gas 36 is blown against the rotor blades 33 to impart rotational force, and then flows away as exhaust gas 37.

A part of this gas flow attempts to leak in the same direction as arrow E through the gap between the rotor blade 33 and the shroud 24, but this leakage flow is caused by the groove 2 formed in the shroud 24 as described later.
It is suppressed by the action of 7.

Note that the gap size t between the rotor blade 33 and the shroud 24 is determined by deformation of the casing due to aging, thermal expansion of each member during operation, elongation deformation of the rotor blade during operation due to centrifugal force, and the turbine shaft (Fig. Minimum ∆ considering the deviation etc.
It is designed and manufactured to ensure t.

From the standpoint of the efficiency of the cast turbine, the value of Δt is preferably as close to zero as possible, but it is determined by design to ensure a safety margin.

FIG. 9 shows a shroud 24 according to the present invention, which includes a plurality of cooling air holes 34, 34, 34 bored in a rotor blade 33.
The air stream 35 ejected from the shroud 24 is blown onto the aforementioned groove 2γ provided in the shroud 24.

Since the rotor blades 33 are rotating, the continuous air flow 35 is blown onto the multiple grooves 27, 27, 27 one after another, and this process is repeated.

Since the shroud 24 is provided with grooves 27, the surface area that comes into contact with the cooling air flow is increased and the cooling effect is improved.
In addition, there are the following effects.

That is, regarding the other groove 27, since the rotor blade 33 facing this groove rotates, each of the plurality of rotor blades intermittently receives the cooling air flow blown one after another.

However, after one rotor blade passes by while blowing cooling air, the cooling air flows into the groove 27 for a minute time (about tens of microseconds) before the next rotor blade approaches and blows cooling air. The inner surface of the groove is prevented from coming into direct contact with the combustion gas flow.

Applying the present invention to provide a groove on the inner peripheral surface 25 of the shroud and causing the cooling air flow from the rotor blades to collide with this groove increases the effect of increasing the cooling area of the shroud and directing the shroud to the combustion gas. This creates an insulating effect that prevents contact, and suppresses the temperature rise of the shroud.

When the temperature rise of the shroud is suppressed, external demand for turbine casings is also suppressed, and the amount of thermal expansion and contraction of these members is reduced, so the amount of dimensional change in these members due to changes in the load state of the gas turbine is small. Become.

Therefore, design calculations related to setting the spacing between the rotor blades and the shroud become easier, and as a result, it becomes possible to accurately estimate changes in the spacing that may occur during operation of the gas turbine. Accuracy and efficiency can be improved, and even if the gas turbine overheats for some reason, the amount of deformation due to thermal expansion of the shroud or turbine casing is not large, so there is less risk of contact damage between the rotor blades and the shroud. produce a variety of direct effects.

Furthermore, as described above, by suppressing the temperature rise of the turbine casing, there is also an indirect effect of preventing the turbine casing from changing its dimensions or losing its shape over time.

Next, the effects of various embodiments devised in the above embodiments will be explained.

FIG. 10 is a schematic cross-sectional view for explaining the operation, and the imaginary arrow mark indicates the gas flow path that is about to leak through the gap between the rotor blade 33 and the shroud 24.

The airflows 35, 35, 35 blown out from the cooling air holes 34, 34, 34 bored in the rotor blades 33 are blown into the inside of the groove 27 formed in the shroud 24, so that they form an airflow wall. Creates an air curtain effect that blocks leakage gas flow.

More specifically with reference to FIG. 11, the cooling air flow 35, 35° 35 exiting from the tip of the rotor blade 33 forms an air flow wall.

As a result, the part 36a of the gas flow that was about to leak is blocked by the air wall and changes its direction, becoming a flow in the direction of the arrow H, and merges with the mainstream of the combustion gas flow, preventing leakage. Become.

Although the air streams 35, 35, 35 described above form air stream walls, they are not completely continuous, resulting in some leakage flow 42 from the gaps between each air stream, but as they flow through narrow gaps, there is no leakage. The state is such that it becomes a vortex 43.43 and drifts.

Furthermore, since the rotor blades 33 and the air flow 35 blown out from the rotor blades are rotating at high speed in the direction of the arrow N shown in FIG. This is a path that passes through the groove 27 several times like the leakage flow path 41 shown in .

For this reason, the leakage gas flow is attenuated every time it passes through the groove 27,
The amount of combustion gas leakage that has finally passed through becomes aesthetically negligible.

FIG. 12 shows a cross section of the shroud 24 and the tip (imaginary line) of the rotor blade 33 opposing the shroud 24.

The length Ll of the groove 27 is smaller than the width L2 of the shroud 24, as previously explained with reference to FIG. 7A, leaving a non-grooved portion of dimension L3 at both ends.

Therefore, the combustion gas cannot leak directly through the groove 27 as shown by the imaginary arrow.

As shown in FIG. 7A, grooves 27, 27, etc. of a plurality of tubes are formed in advance in each of the shroud segments 24a, 24b, which are divided into a plurality of tubes, and these grooves are The configuration in which the segments are attached to the turbine casing not only facilitates the manufacture and assembly of each component, but also eliminates dimensional deviations due to differences in thermal expansion coefficients and uneven temperature distribution between the turbine casing and shroud, thereby preventing the generation of thermal stress. There is a potential effect.

As a result of the effects described above, when the present invention is applied, the temperature rise of the shroud and turbine casing of a gas turbine equipped with rotor blades having cooling air holes is suppressed, thereby improving the performance, durability, and reliability of the gas turbine. The thermal efficiency of the gas turbine is improved by reducing the leakage gas flow rate.

[Brief explanation of the drawing]

Figure 1 is a cross-sectional view showing the structure of a combustor in a gas turbine, Figure 2 is a cross-sectional view of the vicinity of the tip of the rotor blade of the gas turbine,
Fig. 3 is a cross-sectional view to explain the relative position between the moving blades and shroud of a gas turbine, and Fig. 4 is a diagram showing the influence of the gap size between the moving blades and shroud of a gas turbine on the efficiency of the gas turbine. Figure 5 A is a cross-sectional view for explaining dimensional changes of the gas turbine rotor blades and shroud, Figure 5 opening is a cross-sectional view of a turbine rotor blade without cooling air holes, and Figure 5 C is a turbine with cooling air holes. 6 is a cross-sectional view of the rotor blade of a gas turbine according to an embodiment of the present invention, FIG. Fig. 7A is a side view of the members shown, Fig. 8A is a dimensional explanatory diagram of the relationship between the rotor blades and the shroud of a gas turbine in one embodiment of the present invention, and the opening of Fig. 10 is a cross-sectional view taken along line z-z', the parent view is a perspective view showing a part of the tip of a turbine rotor blade and a shroud in an embodiment of the present invention, and FIG. 10 is a perspective view of a turbine rotor blade in an embodiment of the present invention. FIG. 11 is a perspective view of the tip of a turbine rotor blade in an embodiment of the present invention, and FIG. 12 is a cross-sectional view of the tip of the turbine blade in an embodiment of the present invention. FIG. 3 is a cross-sectional view of the shroud. 3.20...Turbine casing, 10,26°36
... Combustion gas, 12, 14, 23, 33... Moving blade,
15.24--Shroud, 24a, 24bt24c
... Shroud segment, 25 ... Shroud segment inner peripheral surface, 28a, 28b ... Shroud segment end, 34 ... Cooling air hole, 35 ... Air flow, 37 ... Exhaust gas, 41 ...Leak gas flow path, 4
2... Leakage gas, 43... Eddy flow of leakage gas. ...Vortex flow of leaking gas.

Claims (1)

[Scope of Claims] 1. In a gas turbine having rotor blades equipped with cooling air holes opened on the outer peripheral tip surface of the rotor blades, the turbine blades are provided on the inner peripheral surface of a shroud that is provided facing the outer periphery of the rotor blades. By arranging a plurality of grooves extending in a direction substantially parallel to the axis and not reaching both ends of the shroud over almost the entire circumference of the shroud, the air flow blown out from the cooling air holes of the rotor blades is directed toward the shroud. A gas turbine shroud characterized by a shroud that collides with a groove on an inner circumferential surface. 2. The shroud for a gas turbine according to claim 1, wherein the shroud has a shape in which a plurality of shroud segments are arranged in series.