JPH0814001A - Gas turbine blade - Google Patents

Abstract

PURPOSE:To enhance the efficiency of a gas turbine and release the thermal stress of a blade so as to improve its reliability, by uniformalizing the temperature distribution of the blade as much as possible through the processes of providing the optimum cooling flow distribution for the purpose of radial temperature distribution of providing the optimum passing order distribution or providing the optimum passing order of cooling air or providing the optimum thermal barrier coating in order to cool the blade by means of the minimum quantity of the cooling air. CONSTITUTION:A gas-turbine blade has cooling air paths 6a-6d for conventional cooling inside the blade, and cooling air exhaust nozzles 7, 8 connecting the outside of the blade to the cooling air paths. The quantity of cooling air ejected out of the cooling air exhaust nozzles 7, 8 is set in compliance with the temperature distribution 5 of the main flow gas in the gas passage, in such a way that the set value becomes relatively high at a high temperature zone of the main flow gas and relatively low at a low temperature zone thereof.

Description

Detailed Description of the Invention

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine blade,
In particular, it relates to a gas turbine blade having improved blade cooling and heat shield means.

[0002]

2. Description of the Related Art Generally, in a gas turbine, a portion where a combustion gas generated in a combustor expands to generate an output is called a gas passage. Gas turbine blades, that is, stationary blades and moving blades are arranged in this gas passage. The parts where these wings are arranged are collectively called paragraphs. The most basic technique for increasing the thermal efficiency of a gas turbine is to increase the combustion temperature in the combustor.

However, the inlet temperature of the gas turbine is already 13 times the inlet temperature of the first stage rotor blade in the turbine for power generation.
The temperature has reached nearly 00 ° C. At such temperatures, the heat resistance limit of existing metals is exceeded, so it is necessary to protect the blades with cooling technology and coating technology.

In general, for cooling the blades, a part of the air in the middle of the compressor or at the outlet is extracted and sent to the stationary blades or moving blades in a low temperature state by bypassing the combustor. The higher the temperature in the gas passage, the more cooling air is required. However, since the cooling air bypasses the combustor and the gas passage, the cooling air does not contribute to the generation of output until mixing in the gas passage after cooling the blades. Therefore, the thermal efficiency of the gas turbine decreases as the amount of the cooling air increases, and the high thermal efficiency obtained by the increase in the inlet temperature of the gas passage may be offset by the increase in the cooling air.

Therefore, in order to increase the thermal efficiency of the gas turbine, it is important to carry out efficient blade cooling with as little cooling air as possible.

Here, the tendency of the temperature distribution inside the gas turbine will be observed. FIG. 21 shows a general gas turbine, which is configured by connecting a compressor 24, a combustor 25, a turbine 26, etc. in order.

FIG. 22 is an enlarged view of the combustor 25 and the turbine 26. Fuel is injected into the air compressed inside the combustor 25, and a combustion reaction occurs to raise the temperature. This combustion generally occurs inside a component called flow sleeve 28, the cross section of which is shown in FIG. This time,
Since the combustion flame tends to have a high temperature near the center 28a of the flow sleeve 28, a temperature distribution 30 is formed inside.

Further, since the flow sleeve 28 itself is also protected from the hot combustion gas, the flow sleeve cooling air 29
Is required. Since this cooling air is supplied from the outer wall, this also causes the temperature outside to be lowered.

The temperature distribution 30 in the combustor thus formed is not attenuated immediately after the combustion gas exits from the combustor 25, and is generally the temperature distribution even after expansion through the turbine 26. Is known to exist. Therefore, in the gas passage of the gas turbine, a temperature distribution in the radial direction is formed in which the temperature is low at the root portion (root portion) and the tip portion (tip portion) of the blade and is high at the central portion.

The temperature distribution in the radial direction is expressed by the following equation called a pattern factor PF in a gas turbine for power generation.

[0011]

PF = (Tm-Tco) / (Ta-Tco) Formula (1) where Tm is the maximum temperature of the target gas passage section,
Tco is the compressor outlet temperature, and Ta is the average temperature. Modern combustors typically have a pattern factor PF of 1.1.
Take a value of about 5. Assuming that Ta = 1300 ° C. and Tco =
Assuming 430 ° C., the maximum temperature in the radial direction is Tm = 143
The temperature is 0 ° C., which is 130 ° C. higher than the average temperature in the paragraph.

By the way, the most important point in cooling the blade is not to create a local high temperature portion in the blade. In order to prevent the oxidation and deterioration of the blade surface, any local temperature needs to be lower than the allowable temperature of the blade material, and a large amount of cooling air needs to be supplied to a portion where the gas temperature is high.

In the opposite sense, in the region where the gas temperature is low, the surface temperature of the blade can be kept sufficiently low even if the cooling air is relatively small. Ideally, the cooling efficiency of the blade is optimized with respect to the temperature distribution in the radial direction, and it is desirable that the temperature distribution be as uniform as possible over the entire blade surface. If a uniform temperature distribution on the blade surface can be achieved, the blade can be cooled with the minimum cooling air, which not only improves the efficiency of the gas turbine but also reduces the thermal stress inside the blade.
It also leads to improved wing reliability.

Further, if there is a temperature distribution in the radial direction inside the mainstream gas, heat flows from the high temperature portion to the low temperature portion inside the mainstream so as to relax this temperature distribution, but this heat flow is inevitably entropy. It causes an increase and causes a loss. Also in this sense, it is desirable to eliminate the temperature distribution in the radial direction as early as possible.

[0015]

Problems with the prior art will be described with reference to FIG. FIG. 24A shows a conventional cooling configuration of a typical moving blade. The cooling air is divided and supplied to the front edge cooling air passage 6a, the axial center cooling air passage 6e, and the rear edge cooling air passage 6f. In the leading edge cooling air passage 6a, the cooling air flows from the blade body root portion 2c toward the blade body tip portion 2a, and is jetted out from the shower head hole 9 of the leading edge 10 to the outside. Inside the leading edge 10, impingement (impingement) cooling is performed by the impingement holes 23 that provide high thermal conductivity, and the blades are protected from the hot mainstream gas 4.

Next, in the axial center cooling air passage 6e, the cooling air flows from the blade passage root 2c to the blade passage tip 2a, and then again to the blade passage root 2c. After flowing along the path, it blows out from the film hole 7 and protects the blade from the hot mainstream gas 4. Trailing edge cooling air passage 6f
However, similarly, after flowing along the meandering flow path, it is ejected to the outside from the trailing edge blowing hole 8.

By the way, in the conventional gas turbine blade, it cannot be said that the cooling air is optimally circulated with respect to the radial temperature distribution 5 of the mainstream gas 4 generated in the combustor. That is, since the radial temperature distribution 5 is highest in the blade passage central portion 2b, supplying cooling air at a low temperature to this portion greatly contributes to equalizing the blade surface temperature. However, in the above-described conventional gas turbine blade, the temperature distribution 5 in the radial direction is ignored, and the cooling air is simply supplied from the blade passage root portion 2c side to the blade passage tip portion 2a side. Therefore, the temperature of the cooling air gradually rises due to the convection effect, and the cooling air having a low temperature is not preferentially supplied to the blade passage central portion 2b.

Further, the blade passage tip portion 2a and the blade passage central portion 2
Since b and the blade passage root portion 2c do not form independent flow passages, it is difficult to control the flow rate of the cooling air in each portion. Therefore, in the conventional example, the optimum way of passing the cooling air is not always carried out for the radial temperature distribution 5 of the mainstream gas 4.

Although not shown in FIG. 24, a thermal barrier coating called a thermal barrier coating is applied to the blade surface by a technique such as plasma spraying. In the conventional example, this coating was set to have a uniform thickness over the entire blade surface and was not optimized for the radial temperature distribution 5.

As described above, the temperature distribution in the radial direction generated in the combustor exists in the passage portion of the gas turbine through which the mainstream gas passes and expands. On the other hand, in order to improve the efficiency of the gas turbine, it is necessary to cool the blades with a minimum amount of cooling air.

An object of the present invention is to make the temperature distribution of the blade as uniform as possible by optimal distribution of the cooling flow rate with respect to the radial temperature distribution, optimal cooling air passage order, or optimal thermal barrier coating. By cooling the blades with a minimum amount of cooling air, the efficiency of the gas turbine is increased, and the thermal stress is relaxed to improve the reliability of the blades.

[0022]

According to the first aspect of the present invention,
In a gas turbine blade having a cooling air passage for convection cooling inside the blade and a hole for jetting cooling air that connects the outer surface of the blade and the cooling air passage, it corresponds to the temperature distribution of the mainstream gas in the gas passage. Then, the amount of cooling air ejected from the cooling air ejection hole is set, and the set value is relatively large in the high temperature region of the mainstream gas and small in the low temperature region.

According to a second aspect of the present invention, there is provided a gas turbine blade having a cooling air passage for convection cooling inside the blade, and a hole for jetting cooling air for communicating the outer surface of the blade with the cooling air passage. The temperature of the cooling air jetted from the cooling air jetting hole is set corresponding to the temperature distribution of the mainstream gas in the gas passage, and the set value is relatively low in the high temperature region of the mainstream gas and low. It is characterized by being heated to a high temperature in the area.

The invention according to claim 3 is the invention according to claim 1 or 2.
The described gas turbine blade is characterized in that the cooling air passage and the holes for jetting the cooling air are divided into a plurality of independent groups.

The invention according to claim 4 is the invention according to claim 1 or 2.
The gas turbine blade described above is characterized in that the upstream side of the cooling air passage is arranged so as to correspond to the high temperature region of the mainstream gas in the gas passage, and the downstream side thereof is arranged so as to correspond to the low temperature region of the mainstream gas.

According to a fifth aspect of the present invention, in the gas turbine blade according to the first to fourth aspects, the pitch of the holes for jetting the cooling air is dense in the high temperature region of the mainstream gas in the gas passage and is in the low temperature region. Or the diameter of the hole is set large in the high temperature region of the mainstream gas and small in the low temperature region.

According to a sixth aspect of the present invention, in a gas turbine blade having a thermal barrier coating for heat shielding on the outer surface of the blade, the thermal resistance of the thermal barrier coating is relatively large in the high temperature region of the mainstream gas in the gas passage. Also, it is characterized by being set small in the low temperature range.

[0028]

According to the first to fourth aspects of the present invention, the cooling air flow is classified according to the radial temperature distribution of the mainstream gas, so that the cooling efficiency of the blade passage central portion, the leading edge, etc. is increased. It is possible to preferentially or prematurely cool a necessary portion with a large amount of cooling air or with low temperature cooling air.

According to the fifth aspect of the invention, a film cooling hole through which cooling air flows out to the outer surface of the blade, a shower head hole,
It is effective to change the pitch, area, etc. of the trailing edge blowing holes in the central portion of the blade passage, the tip portion of the blade passage, and the root portion of the blade passage. This makes it possible to control and optimize the cooling air flow rate according to the temperature distribution in the radial direction.

According to the invention of claim 6, the thermal barrier coating applied to the outer surface of the blade is made thicker at the central portion of the blade passage, or the thermal conductivity is made smaller,
It is possible to increase the thermal resistance of the portion and reduce the temperature of the blade.

By implementing the above-mentioned means singly or in combination, the action of enhancing the cooling efficiency or the heat shielding effect of the central portion of the blade passage where the temperature distribution of the mainstream gas in the radial direction is high is This can be further improved depending on the type of turbine blade, which makes it possible to make the temperature distribution inside the blade uniform with a minimum of cooling air.

[0032]

Embodiments of the present invention will be described below with reference to the drawings. 1 to 16 show a first embodiment of a gas turbine blade according to the present invention.

In this embodiment, as shown in FIG.
The moving blade 1 is configured to have a blade main body portion 2 that comes into contact with hot mainstream gas to obtain a rotational force, and an implant portion 3 for fixing the blade main body portion 2 to a rotor disk. As shown in FIG. 1 (b), the mainstream gas 4 emitted from the vanes has a radial temperature distribution 5 created by the combustor. This temperature distribution 5 depends on the characteristics of the combustor as described above. Relatively, the blade body tip portion 2a and the blade body root portion 2c have a low temperature distribution, and the blade body central portion 2b has a highest temperature distribution.

The cooling air passages 6 are a plurality of independent passages, that is, the leading edge cooling air 6a and the blade body tip cooling air passages 6.
b, a blade main body cooling air passage 6c and a blade main body cooling air passage 6d. Cooling air is independently supplied to each of the passages 6a to 6d through a supply pipe individually connected to each of the passages 6a to 6d.

FIG. 2 shows a cross section taken along the line AA of FIG. Shower head holes 9 are formed in a plurality of rows in the lateral direction at the front edge 10 of the rotor blade 1 so as to face the mainstream gas 4. On the back side 11 and the ventral side 12, film holes 7 for ejecting cooling air along the flow direction of the mainstream gas 4 are formed.
Has been drilled.

In the leading edge cooling air passage 6a, cooling air is independently supplied toward the leading edge 10 where the blade surface pressure becomes a stagnation point, and faces the mainstream gas 4 through a large number of shower head holes 9. Drains in the direction.

The blade main body cooling air 6b, the blade main body central cooling air passage 6c, and the blade passage root cooling air passage 6d form a meandering passage in the blade main body 2 after passing through the implanting portion 3. Therefore, the inner surface of the blade is cooled over a wide range as much as possible. Then, a part of the cooling air is jetted from the many film holes 7 to the outer surface of the blade, and is mixed with the mainstream gas 4 to cover the outer surface of the blade on the film to protect the blade. In this way, the cooling air gradually decreases in flow rate, and finally, in either case, the cooling air finally flows out from the trailing edge blowing hole 8 and joins with the mainstream gas 4.

FIG. 3A shows the rotor blade 1 on the ventral side (lower side in FIG. 2).
The outer shape of the film hole 7 is shown in FIGS. As shown in these figures, the pitch of the film holes 7 is different between the blade central portion and the blade tip portion or blade root portion. That is, FIG.
Corresponding to the radial temperature distribution 5 of the mainstream gas 4 shown in (b), the cooling hole pitch 14a at the blade central portion is set smaller than the cooling hole pitch 14b at the blade tip portion and the blade root portion. It has a structure in which a large amount of cooling air flows in a portion where the temperature of the mainstream gas 4 is high. Not only the film hole 7, but also the shower head hole 9 and the trailing edge blowing hole 8 are similarly densely arranged in the blade central portion.

FIGS. 4 to 7 are views for explaining the thermal barrier coating 15. In these drawings, the thermal barrier coating 15 is shown in an enlarged manner, and is applied to the blade body 2 by means such as plasma spraying in order to prevent oxidation and deterioration of the blade surface at high temperature and to enhance the heat shield property. . The thickness of the thermal barrier coating 15 can be changed depending on the blade cross section or the temperature of the mainstream gas. 5, 6 and 7 are shown in FIG.
-E cross section and different radial positions are shown. That is, FIG. 5 is a sectional view taken along line BB in FIG.
That is, it is a diagram showing the thermal barrier coating thickness 15a at the tip portion. Similarly, FIG. 6 is a sectional view taken along the line AA in FIG. 1, that is, the central thermal barrier coating thickness 15b.
FIG. 7 is a diagram showing the CC cross-sectional position in FIG. 1, that is, the root thermal barrier coating thickness 15c.

In the present embodiment, as shown in FIGS. 5, 6 and 7, the thickness of the thermal barrier coating is changed according to the cross-sectional position of the blade, and the central thermal barrier coating thickness 15b is the largest, The partial thermal barrier coating thickness 15c and the tip thermal barrier coating thickness 15a are successively reduced.

Next, the operation of this embodiment having the above structure will be described with reference to FIGS. Leading edge 10 of rotor blade 1
As described above, since the leading edge cooling air passage 6a is formed independently of the cooling air passages 6b, 6c, 6d of the other parts, the low temperature cooling air is guided with a minimum pressure loss. Is possible.

FIGS. 8A and 8B show the blade cross section and the pressure distribution on the blade surface based on the flow of the mainstream gas 4. Since the leading edge 10 of the moving blade 1 becomes a stagnation point, the blade surface pressure becomes highest due to the dynamic pressure of the mainstream gas 4. Therefore, the leading edge 10
The part of is a part where the heat transfer coefficient of the outer surface of the blade is high and the temperature environment is severe. Therefore, spraying cooling air from the shower head hole 9 along the direction facing the mainstream gas 4 at this portion to protect the outer surface of the blade is a very effective cooling method. However, it should be noted here that, as shown in FIG. 8 (b), since the pressure of the outer surface of the blade is high at the leading edge 10, unless the pressure of the cooling air is set higher than that, the high temperature That is, the mainstream gas 4 may flow in and cause damage to the blade. In the present embodiment, as described above, the cooling air independent of the other parts is supplied to the leading edge cooling air passage 6a to guide the low temperature cooling air with a minimum pressure loss, so that the cooling is performed. It is very advantageous in ensuring a proper pressure difference between the air and the mainstream gas. Further, since the cooling air of this part is supplied independently, heat is not exchanged with the cooling air of the cooling air passages of other parts, and therefore the temperature of the cooling air rises due to the influence from other parts. It is possible to cool this site without it.

It is also a great advantage that the cooling air is independently supplied to each of the blade body tip cooling air passage 6b, the blade body central portion cooling air passage 6c and the blade body root cooling air passage 6d. For the most efficient cooling of the entire blade, ideally there should be no local hot spots.

In the present embodiment, a large amount of cooling air is supplied to a portion where the mainstream gas temperature 4 is high, or a low temperature cooling air is supplied to enhance cooling efficiency, and conversely, to a portion where the temperature of the mainstream gas 4 is low. Can be approached to the above ideal by supplying either a small amount of cooling air or a high temperature cooling air. That is, if each cooling air passage is independent, the flow rate of the cooling air can be adjusted according to the radial temperature distribution 5 of the temperature of the mainstream gas 4 shown in FIG.

In the case of this embodiment, specifically, FIG.
As shown in (c), the adjustment can be realized by changing the pitches of the film holes 7, the trailing edge blowing holes 8, the shower head holes 9 and the like in the radial direction. That is, in the blade body central portion 2b where the radial temperature distribution 5 is high, the cooling hole pitch 14a in the blade central portion is made dense as shown in FIG. 3C, and the amount of air used for the blade body central portion cooling air passage 6c is set. Do more.

On the other hand, at the blade body tip portion 2a and the blade body root portion 2c, as shown in the radial temperature distribution 5 in FIG. 1 (b),
The mainstream gas temperature 4 is not necessarily high. Therefore, in these portions, the cooling hole pitches 14b on the blade tip side and the blade root side are roughened as shown in FIG. 3 (b) to form the blade body tip cooling air passage 6b and the blade body root cooling air passage 6d. The amount of air supplied is relatively small.

Here, the operation of making the cooling air passage independent for each radial position of the blade will be further described. FIG. 9 is a diagram showing a concept of film cooling, and FIG. 10 is a diagram showing a general tendency of film efficiency. As shown in FIG. 9, the cooling air 17 having a low temperature flows out from the film hole 7 formed in the rotor blade 1, flows along the outer surface of the rotor blade 1, and protects the rotor blade 1 from the hot mainstream gas 4. . Regarding the heat transfer from the mainstream gas 4 to the moving blades 1, it is of course necessary to consider this film cooling. The effectiveness of this film cooling is represented by the film efficiency η and is generally defined by the following equation.

[0048]

[Equation 2] η = (Tg −Tf) / (Tg −Tc) Formula (2) where Tg: mainstream gas temperature, Tf: film temperature, T
c: Cooling air temperature.

The film efficiency is shown in FIG. 1 immediately after the film hole 7.
Although it takes a high value as shown in 0, it gradually becomes lower as the blade surface position becomes on the wake side. Further, the film efficiency is governed by various factors such as the shape of the film hole 7, the curvature of the blade surface, the surface roughness, etc. When the film efficiency and the mainstream gas temperature Tg are the same, the film temperature Tf
From equation (2).

[0050]

## EQU3 ## Tf = Tg-.eta. (Tg-Tc) Equation (3), and naturally, the lower the cooling air temperature, the lower the temperature and the greater cooling effect.

However, since the cooling air passes through the complicated cooling air passage in the moving blade 1, the cooling air is already cooled before reaching the film hole 7 in the central portion 2b of the blade body, which conventionally requires a large cooling effect. The temperature was rising due to convective heat transfer inside the body.

On the other hand, in the case of this embodiment, the cooling air passage is formed in accordance with the radial temperature distribution 5 of the mainstream gas 4 like the blade body tip portion 2a, the blade body central portion 2b, and the blade body root portion 2c. Since they are set independently, cooling air having a low temperature is supplied to only the blade passage central portion 2b having a particularly high temperature, and the cooling air having a low temperature is supplied to the blade passage central portion 2b, so that the film temperature can be kept low.

Next, the operation of optimizing the pitch of the film holes 7 in the radial direction will be described. 11 and 12 are enlarged views taken along the line D-D of FIG. 2, which compare the pitches of the film holes 7, and FIG. 11 shows a cross section taken along the line AA of FIG. 12 shows a portion having a coarse pitch in the BB cross section or the CC cross section in FIG.

After the cooling air 17 is blown from the film hole 7, it spreads to the outer surface of the blade while diffusing into the mainstream gas 4. At this time, the merging position 18 of the cooling air is closer to the film hole 7 in FIG. 11 where the pitch is denser than in FIG. 12 where the pitch is coarse. Therefore, the cooling air from the adjacent film holes 7 mixes quickly to cover the outer surface of the blade over a wide range, and the cooling efficiency is higher in FIG. 11.

Next, with reference to FIG. 13 and FIG. 14, description will be given of the effect that the cooling air flow rate increases in the radial direction toward the high temperature portion in the following paragraphs. 13 and 14 show a radial temperature distribution 5 of the mainstream gas 4 before entering the moving blade 1 (left side of the drawing) and a radial temperature distribution 5a after cooling air mixing (right side of the drawing). ing. FIG. 13 shows a conventional example in which the cooling air flow rate distribution 19 is constant in the radial direction as a comparative example, and FIG. 14 shows that the cooling air passages are independent in the radial direction and the film holes 7 and the shower head holes are provided. 9 shows the case of this embodiment in which the pitch of the trailing edge blowing holes 8 is changed in the radial direction.

FIG. 1 according to this embodiment, in which a large amount of cooling air is caused to flow through the central portion of the blade body where the temperature distribution of the mainstream gas in the radial direction is highest.
In the case of No. 4, the radial direction temperature distribution 5a after cooling air mixing after leaving the moving blade 1 has a more uniform temperature distribution as compared with the case of FIG. 13 showing the conventional example.

When the temperature distribution of the mainstream gas is non-uniform, thermal diffusion in the mainstream leads to a uniform temperature distribution, but at this time the entropy increases, and the paragraph efficiency decreases. Therefore, to make the temperature distribution after the next paragraph as uniform as possible,
It has the effect of improving the efficiency of the paragraph and facilitating the cooling design of the subsequent paragraphs.

The above-mentioned effect is obtained not only by the cooling air flow rate distribution 19 but also when the cooling air temperature 20 is low in the central portion of the blade body as shown in FIG.
When preferentially flowing low-temperature cooling air to the center of the blade body,
The temperature of the cooling air ejected from the cooling holes has a low central shape like the cooling air temperature distribution 20 in the figure, and the radial temperature distribution 5a after mixing the cooling air after exiting the moving blade 1 is shown in FIG.
A more uniform temperature distribution can be obtained as compared with the above.

In addition to the above cooling action, changing the thickness of the thermal barrier coating 15 in the radial direction also gives an optimum cooling effect to the radial temperature distribution 5 of the mainstream gas 4. This action will be described with reference to FIG.

FIG. 16 shows the thermal barrier coating part shown in FIGS. 5 to 7 with the orientation thereof changed to the vertical direction, and a temperature distribution 21 in the thickness direction is added. The flow is shown schematically. An infinite flat plate is assumed for simplification of description. In this case, the heat flux q per unit area is

## EQU4 ## q = K (Tf-Tc) Equation (4) is obtained. K is the heat transmission coefficient,

(Equation 5) Can be expressed as Here, Tf; film temperature 2
1a, T1; coating outer surface temperature 21b, T2; blade outer surface temperature 21c, T3; blade inner surface temperature 21d, Tc; cooling air temperature 21e. Further, δt: thermal barrier coating thickness 15d, δ: blade thickness 22, αo: heat transfer coefficient of outer surface of blade, αi: heat transfer coefficient of inner surface of blade, λt: heat conductivity of thermal barrier coating, λ; The thermal conductivity of the blade.

From the continuous condition, the heat flux q is

(Equation 6) At this time, T2; blade outer surface temperature 21c and T3; blade inner surface temperature 2
1d is

(Equation 7) It is expressed as From these equations, it is apparent that unless the thermal barrier coating thickness (δt) 15d is thick, the blade outer surface temperature (T2) 21c and the blade inner surface temperature (T3) 21d are kept low. Thermal barrier coating 15
Is thicker, the more time and difficulty of construction increases,
The radial temperature distribution 5 is high like the blade passage central portion 2b,
The thickening of only the part that requires a larger heat shield effect than other parts also contributes to uniform blade temperature distribution. It should be noted that a part of the thickness of the coating is basically to obtain an effect of increasing the thermal resistance, and a similar effect is obtained by controlling the porosity of the ceramic coating material or the additive component. You can also get

Based on the structure and operation of this embodiment described above, the following effects are obtained. First, with respect to the radial temperature distribution of the mainstream gas 4 having the highest temperature in the blade central portion, the air passages 6a to 6d for the cooling air supplied to the blade central portion, the blade tip portion, and the blade root portion are branched beforehand independently. The cooling air having a low temperature can be supplied to the blade central portion where the highest cooling effect is desired.

Further, by making the pitches of the shower head holes 9, the film holes 7, the trailing edge blowing holes 8 and the like in the central portion of the blade smaller in the central portion of the blade, a larger amount of cooling air can be made to flow than in other portions. It becomes possible and the cooling effect can be enhanced at this portion.

Further, since the pitch is small, the cooling air from the adjacent film holes 7 mixes more quickly,
By covering the outer surface of the blade over a wide range, a higher cooling efficiency can be obtained in this part compared to other parts.

In this way, the cooling air passages are independent and separated according to the mainstream temperature distribution in the radial direction, and the flow rate thereof is optimized by the pitch of the cooling holes. Since the temperature distribution on the blade surface is made uniform, it greatly contributes to improving the efficiency of the gas turbine.

Further, if the temperature distribution on the blade surface is made uniform,
The thermal stress can be kept small, which also contributes to the improvement of reliability. Furthermore, if a large amount of cooling air is made to flow in the central part of the blade where the temperature of the mainstream gas is high, the temperature distribution in the radial direction will be relaxed in the next paragraph, the efficiency in the next paragraph will increase, and the cooling design in the next paragraph will be easier. It also has the effect of

As described above, providing the optimum cooling method for the temperature distribution in the radial direction of the main flow can also be realized by the thermal barrier coating 15 applied to the outer surface of the blade. That is, if the thermal barrier coating 15 at the blade central portion is thickly applied or the porosity of the coating material or the additive component is controlled, the thermal resistance at this portion is increased and the temperature of the blade outer surface and the blade inner surface is lowered. When combined with the above-described optimization of the cooling air passage, the effect is further enhanced.

FIG. 17 shows a second embodiment of the present invention.
In this embodiment, the cooling air passage 6 extends from the implanting portion 3 to the front edge 10 so that a part of the cooling air flows out from the shower head hole 9 to the outer surface of the blade. The cooling air passage 6 meanders toward the trailing edge 13 side.

Although the cooling air passage 6 is not independent,
The order in which the cooling air flows is arranged in the order that requires a higher cooling effect. That is, first, it flows to the leading edge 10 having the highest priority, and then sequentially flows from the blade body central portion 2b to the blade body root portion 2c or the blade passage tip portion 2a. Therefore, the cooling air passage 6 passing through the leading edge 10 is supplied to the blade main body central portion 2b in which the temperature of the mainstream gas 4 is next higher, so that the film temperature of this portion is kept lower than that of other portions. The cooling efficiency can be improved.

Further, the pitch of the cooling holes such as the film holes 7, the trailing edge blowing holes 8 and the shower head holes 9 in the central portion 2b of the blade main body is made small so that a large amount of cooling air is supplied to this portion and the thermal barrier is provided. Similar to the previous embodiment, the thermal resistance of the coating is increased to enhance the heat shielding effect and the temperature of the entire blade can be made uniform.

FIG. 18 illustrates the third embodiment. In this embodiment, the leading edge cooling air passage 6a for cooling the leading edge 10 is made independent, and a collision flow (impingement flow) is applied to the inner surface of the blade of the leading edge 10 for cooling the portion. It has become. The cooling effect is enhanced by providing the impingement holes 23 that increase the heat transfer coefficient of the inner surface. The other cooling air passages 6 extend from the blade body central portion 2b to the blade body root portion 2c or the blade body tip portion 2a in descending order of priority of cooling. This point is almost the same as the embodiment of FIG.

Also in this embodiment, the cooling holes such as the film holes 7, the trailing edge blowing holes 8, the shower head holes 9 and the impingement holes 23 are made small in pitch, and a large amount of cooling air is supplied to these parts. In addition, the thermal resistance of the thermal barrier coating is increased to increase the heat shield effect,
The point that the temperature of the entire blade can be made uniform is the same as in each of the above embodiments.

FIG. 19 shows a fourth embodiment. In this embodiment, the leading edge cooling air passage 6a and the blade body central portion cooling air passage 6c are independent of the other cooling air 6 passages. The cooling structure of the front edge 10 is substantially the same as that of the embodiment shown in FIG.

In this embodiment, the cooling air passage 6 passing through the cooling air passage 6c in the central portion of the blade body flows out only from the film hole 7 in the central portion and is completely independent. Also, there is an advantage that the flow rate can be easily set.

Further, the entire flow of the cooling air is directed from the blade root direction to the blade tip direction over all the passages, and flows in the direction of the centrifugal force applied to the blade. Therefore, the flow field is simplified, the conduit loss is reduced, and the local nonuniform heat transfer coefficient such as flow reversal is reduced.

Also in this embodiment, the pitch of the cooling holes such as the film holes 7, the trailing edge blowing holes 8, the shower head holes 9 and the impingement holes 23 is made small, and a large amount of cooling air is supplied to this portion. Further, the thermal resistance of the thermal barrier coating is increased to enhance the heat shielding effect, and the temperature of the entire blade can be made uniform, which is the same as in each of the above embodiments.

FIG. 20 shows the fifth embodiment. In this embodiment, the amount of cooling air supplied to each part of the blade is optimized not by the pitch of the cooling holes but by the area of the cooling holes. That is, as shown in FIGS. 20B and 20C, in this embodiment, the cooling holes such as the film hole 7 and the shower head hole 9 in the blade body central portion 2b are the blade body tip portions 2a.
And has a larger area than the cooling holes of the blade body root portion 2c. With such a configuration, the blade temperature can be optimized according to the temperature distribution in the radial direction.

In addition to the moving blades described above, the present invention can be appropriately applied to stationary blades, nozzles and the like.

[0079]

As described in detail above, according to the present invention, the supply of the cooling air to the gas turbine can be made independent according to the temperature distribution of the mainstream gas, or can be supplied in order of priority. It has a very great effect on the uniformization of the blade surface temperature. In particular, the cooling air passages of the leading edge portion, the blade central portion, the blade tip portion and the blade root portion are independent of each other, or the cooling air is supplied in order from the leading edge having a high priority of cooling to the blade central portion, In addition, the pitch or area of the cooling holes is set so that the largest amount of cooling air flows in the central part of the blade where the temperature distribution is highest in each direction, and the thermal barrier coating is also thickest in the central part of the blade, or the porosity and By controlling the additive components to increase the heat shielding effect at the relevant portion, a cooling method optimized for the radial temperature distribution of the mainstream gas can be realized.

Therefore, there is an effect that an increase in cooling air, which causes a large decrease in the efficiency of the gas turbine, is suppressed to a minimum and a highly efficient gas turbine is provided. Further, by making the blade surface temperature distribution uniform, it is possible to contribute to the relaxation of thermal stress inside the blade, and the reliability of the blade is also improved.

Further, in the case of the configuration in which the cooling air field flow is caused to flow only in the direction of the centrifugal force acting during the rotation of the blade, there is no inversion part for the cooling air, so that the flow field in the cooling air passage part is simplified. In addition, the heat transfer coefficient inside the blade can be made uniform, and the temperature distribution can be made uniform.

[Brief description of drawings]

1A is a cross-sectional view of a blade showing a first embodiment of the present invention, and FIG. 1B is a view showing a temperature distribution.

FIG. 2 is a sectional view taken along line AA of FIG.

3 (a) is an outline view of the blade shown in FIG. 1 (a),
(B), (c) is a figure which shows the pitch of a cooling hole.

FIG. 4 is a view showing a thermal barrier coating in the above embodiment.

FIG. 5 is an enlarged view (BB in FIG. 1A) showing the thickness of the thermal barrier coating in the vicinity of the chip in the example.
Equivalent to the cross-sectional position).

FIG. 6 is an enlarged view (AA in FIG. 1A) showing the thickness of the thermal barrier coating in the vicinity of PCD in the example.
Equivalent to the cross-sectional position).

FIG. 7 is an enlarged view showing the thickness of the thermal barrier coating near the route in the above example (CC in FIG. 1A).
Equivalent to the cross-sectional position).

8A and 8B are views for explaining the blade surface pressure distribution in the above-described embodiment, where FIG. 8A is a blade sectional view and FIG. 8B is a distribution diagram.

FIG. 9 is a diagram showing a state of film cooling flow in the embodiment.

FIG. 10 is a graph showing film efficiency in the above example.

FIG. 11 is a view showing a film cooling state of cooling holes having a small pitch in the above-described embodiment (a view taken along the line D-D in FIG. 2).

FIG. 12 is a view showing a film cooling state of cooling holes having a large pitch in the above-described embodiment (a view taken along the line CC in FIG. 2).

FIG. 13 is a view showing a comparative example of the above-mentioned embodiment and showing a transition of the temperature distribution in the radial direction.

FIG. 14 is a diagram showing the transition of the temperature distribution in the radial direction in the above embodiment (a diagram showing the effect of unequal pitch).

FIG. 15 is a diagram showing the transition of the temperature distribution in the radial direction in the above embodiment (a diagram showing the effect of the supply air temperature).

FIG. 16 is a diagram showing the effect of the thickness of the thermal barrier coating in the above example.

FIG. 17 is a diagram showing a second embodiment of the present invention.

FIG. 18 is a diagram showing a third embodiment of the present invention.

FIG. 19 is a diagram showing a fourth embodiment of the present invention.

FIG. 20 is a diagram showing a fifth embodiment of the present invention.

FIG. 21 is a cross-sectional view showing the entire gas turbine.

FIG. 22 is an enlarged view showing a part of FIG. 21;

FIG. 23 shows a combustor liner.

FIG. 24 (a) is a diagram showing a conventional gas turbine blade;
(B) is a figure which shows temperature distribution.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 rotor blade 2 blade passage portion 22a blade body tip portion 2b blade body central portion 2c blade body root portion 3 implantation portion 4 mainstream gas 5 radial temperature distribution 5a radial temperature distribution after cooling air mixing 6 cooling air passage 6a leading edge Part cooling air passage 6b Blade body tip cooling air passage 6c Blade body central cooling air passage 6d Blade body root cooling air passage 6e Axial central cooling air passage 6f Trailing edge cooling air passage 7 Film hole 8 Trailing edge blowing Hole 9 Shower head hole 10 Leading edge 11 Dorsal side 12 Ventral side 13 Trailing edge 14a Cooling hole pitch at the blade central portion 14b Cooling hole pitch at the blade tip and root 15 Thermal barrier coating 15a Thermal barrier coating thickness 15b Thermal central portion Barrier coating thickness 15c Root thermal barrier coating thickness 15d Marubarrier coating thickness 16 Leading edge blade surface pressure 16a Dorsal blade surface pressure 16b Ventral blade surface pressure 17 Cooling air 18 Cooling air confluence position 19 Cooling air flow rate distribution 20 Cooling air temperature distribution 21 Thickness direction temperature distribution 21a Film temperature 21b Coating outer surface temperature 21c Blade outer surface temperature 21d Blade inner surface temperature 21e Cooling air temperature 22 Blade thickness 23 Impingement hole 24 Compressor 25 Combustor 26 Turbine 27 Stator blade 28 Flow sleeve 28a Flow sleeve center 29 Flow sleeve cooling air 30 Combustor inside Temperature distribution

Claims (6)

[Claims] 1. A gas turbine blade having a cooling air passage for convection cooling inside the blade, and a hole for jetting cooling air communicating between the outer surface of the blade and the cooling air passage, wherein a mainstream gas in the gas passage is provided. The amount of cooling air jetted from the cooling air jetting hole is set in accordance with the temperature distribution of, and the set value is relatively large in the high temperature region of the mainstream gas and small in the low temperature region. A gas turbine blade characterized by. 2. A gas turbine blade having a cooling air passage for convection cooling inside the blade, and a hole for jetting cooling air communicating the outer surface of the blade with the cooling air passage. The temperature of the cooling air ejected from the holes for ejecting the cooling air is set according to the temperature distribution of, and the set values are relatively low in the high temperature region of the mainstream gas and high in the low temperature region. A gas turbine blade characterized by. 3. The gas turbine blade according to claim 1, wherein the cooling air passage and the holes for jetting the cooling air are divided into a plurality of independent groups. 4. The gas turbine blade according to claim 1, wherein the cooling air passage is arranged such that an upstream side thereof corresponds to a high temperature region of the mainstream gas in the gas passage and a downstream side thereof corresponds to a low temperature region of the mainstream gas. A gas turbine blade characterized by the above. 5. The gas turbine blade according to any one of claims 1 to 4, wherein the pitch of the holes for jetting cooling air is set densely in the high temperature region of the mainstream gas in the gas passage and coarsely in the low temperature region, Alternatively, the diameter of the hole is set to be large in a high temperature region of the mainstream gas and small in a low temperature region thereof. 6. A gas turbine blade having a thermal barrier coating for heat shielding on the outer surface of the blade, wherein the thermal resistance of the thermal barrier coating is relatively large in the high temperature region of the mainstream gas in the gas passage and small in the low temperature region. A gas turbine blade characterized by being set.