Gas turbine stationary blade cooling

BACKGROUND OF THE INVENTION:

Field of the Invention:

The present invention relates to a gas turbine stationary blade, and more specifically to a gas turbine stationary blade having a cooling structure for applying air cooling to a second stage stationary blade with a high cooling efficiency.

Description of the Prior Art:

In Fig. 1, a cross sectional view of a typical structure of gas turbine is shown and an outline thereof will be described first. In Fig. 1, numeral 1 designates a compressor portion, numeral 2 designates a combustor portion and numeral 3 designates a turbine portion. Numeral 4 designates a rotor, which extends in a turbine axial direction from the compressor portion 1 to the turbine portion 3. Numeral 6 designates an inner housing and numerals 7, 8 designate cylinders of compressor portion 1, which surround an outer circumference of a compressor. Numeral 9 designates a cylindrical shell forming a chamber, numeral 10 designates an outer shell of the turbine portion 3, numeral 11 designates an inner shell of same, numeral 12 designates a stationary blade of compressor and a plurality of the stationary blades are disposed along a compressor circumferential direction with equal spacing between each of the blades and in multi-stage along a compressor axial direction and numeral 13 designates a moving blade of compressor and a plurality of the moving blades are fixed around the rotor 4 and disposed alternately with the stationary blades 12 along the compressor axial direction.

Numeral 14 designates a chamber surrounded by the cylindrical shell 9 and numeral 15 designates a combustor, disposed in the chamber 14, into which fuel 35 is injected from a fuel nozzle 34 for combustion. Numeral 16 designates a duct for leading a high temperature combustion gas 30 generated in the combustor 15 into the turbine portion 3. Numeral 17 designates a second stage stationary blade of gas turbine, which is the object of the present invention. In the case shown in Fig. 1, the gas turbine is constructed of four stage stationary blades and four stage moving blades disposed alternately therewith, and the high temperature combustion gas 30 passes through the blades and is discharged as an expanded gas 31. Numeral 21 designates a manifold of the compressor portion 1 and numeral 22 designates a manifold of the turbine portion 3 and cooling air is supplied from the manifold 21 of the compressor portion 1 to the manifold 22 of the turbine portion via a pipe 32 and an air piping 19.

In the gas turbine constructed as mentioned above, the fuel 35 is injected into the combustor 15 from the fuel nozzle 34 to be burnt to generate the high temperature combustion gas 30 and then flows into the turbine portion 3 to pass through a passage where the stationary blades and the moving blades are disposed alternately and to expand to rotate the moving blades and the rotor 4 and is discharged as the expanded gas 31.

On the other hand, while a portion of the cooling air is supplied from the compressor portion into the moving blades of the gas turbine for cooling thereof via rotor discs, a portion of the cooling air is also supplied from the manifold 21 of the compressor portion 1 into the manifold 22 of the turbine portion 3 for cooling of the second stage stationary blade 17 as well as to be used as a seal air via the pipe 32 and the air piping 19.

Next, the second stage stationary blade 17 will be described in detail. Fig. 6 is a cross sectional view of the second stage stationary blade 17 of the prior art gas turbine, said stationary blade being cut along a turbine axial direction at approximately a central portion of its inner shroud and seen from an inner side thereof, that is, on a rotor 4 side, Fig. 7 is a cross sectional view taken on line D-D of Fig. 6, Fig. 8 is a cross sectional view taken on line E-E of Fig. 6, Fig. 9 is a cross sectional view taken on line F-F of Fig. 6, Fig. 10 is a cross sectional view taken on line G-G of Fig. 6, Fig. 11 is a cross sectional view taken on line H-H of Fig. 6 and Fig. 12 is a cross sectional view taken on line J-J of Fig. 6.

In Fig. 6, numeral 26 designates an inner shroud and therein provided are a rib 40, and a leading edge passage 42 and a trailing edge passage 44 mutually separated by the rib 40, and a projection portion 95 provided therearound. Numerals 96, 97 designate rails of both side edge portions of the inner shroud 26 and numerals 93, 94 designate passages of cooling air provided in the rails 96, 97, respectively. A passage 88 is provided in an leading edge portion 41 of the inner shroud 26 and a multiplicity of passages 92 are provided in a trailing edge portion 43 of the inner shroud 26. There are provided a multiplicity of needle-like fins in the passage 88, so that a convection is accelerated and a heat transfer efficiency is enhanced. Numeral 100 designates a recess portion formed by the projection portion 95 and numerals 83, 84 designate impingement plates, each having a multiplicity of small holes 101 provided therein as passages of air.

Numerals 81, 82 designate a front flange and a rear flange, respectively, and there are provided passages 90, 91 in the front flange 81. Cooling air 57 which has entered the recess portion 100 passes through the passage 90 in the front flange 81 and the passage 88 in the leading edge portion 41 and then through the passage 91 in the front flange 81 and enters a chamber formed by the impingement plate 83. Also, a portion 58 of the cooling air which has entered the passage 88 passes through the passages 93, 94 in the rails 96, 97 of the side edge portions for cooling therearound and is discharged outside as a cooling air 61. The cooling air which has flown through the small holes 101 of the impinge plates 83, 84 and the cooling air which has flown through the passage 91 gather together in the chamber to further flow through the multiplicity of passages 92 of the trailing edge portion 43 and to be discharged outside as a cooling air 60.

In Fig. 7, being a cross sectional view taken on line D-D of Fig. 6, the passage 88 is formed in the leading edge portion 41 of the inner shroud 26 and the multiplicity of needle-like fins 89 are provided therein. In a space between the front flange 81 and the rear flange 82 provided are the recess portion 100 in the front of the projection portion 95 and a recess portion 99 in the rear of same. On the inner side thereof, the impingement plate 84 is provided so as to form the chamber 78 on an outer side of the impingement plate 84. In the front flange 81, there is provided a passage 90 which connects to the passage 88 and a portion 57 of the cooling air, flowing through the passages 90 and 88, and another portion 59 of the cooling air, passing through the small holes 101 of the impingement plate 84, gather together in the chamber 78 to further flow through the multiplicity of passages 92 of the trailing edge portion 43 and is then discharged as the cooling air 60.

In Fig. 8, being a cross sectional view taken on line E-E of Fig. 6, the second stage stationary blade 17 has the inner shroud 26 and the outer shroud 27 and a blade portion 25 is formed therebetween. The leading edge passage 42 in front of the rib 40 and the trailing edge passage 44 in the rear of same are formed between a leading edge portion 28 and a trailing edge portion 29 of the blade portion 25 and cylindrical members 46, 47 are inserted into these passages 42, 44, respectively. There are provided a multiplicity of cooling air holes 70, 71 in side walls of the cylindrical members 46, 47, respectively, and also cooling air holes 72, 73 in bottom walls of the cylindrical members 46, 47, respectively. Further, there are provided a multiplicity of pins 62 in the trailing edge portion 29.

In the leading edge portion 41 of the inner shroud 26, the passage 88 and the needle-like fins 89 in the passage 88 are provided, and in the trailing edge portion 43 of the inner shroud 26, the passages 92 are provided so as to connect to a cavity 45 which is formed by the front and rear flanges 81, 82 and a seal support portion 66. A chamber 77 is formed by the impingement plate 84 in the cavity 45. On the inner side of the cavity 45, the seal support portion 66 supports a seal 33, by which a seal mechanism between the inner shroud 26 and rotor side arm portions 48 is constructed.

Cooling air 19' from the air piping 19 flows into the cylindrical members 46, 47 to be injected through the cooling air holes 70, 71 to impinge on walls of the leading edge passage 42 and the trailing edge passage 44 and to flow toward the inner side thereof as well as to be injected through the cooling air holes 72, 73 of the bottom walls of the cylindrical members 46, 47 to flow into opening portions 68, 69 and then the cooling air as shown by numerals 75, 76 flows into the cavity 45. The cooling air then flows into a space between the inner shroud 26 and a front stage moving blade thereof and a space between the inner shroud 26 and a rear stage moving blade thereof via the seal 33 to thereby maintain said spaces in a higher pressure than in a passage of the high temperature combustion gas 30 and to prevent the high temperature combustion gas 30 from coming into said spaces.

In Fig. 9 being a cross sectional view taken on line F-F of Fig. 6, a recess portion 98 and the chamber 77 are formed by the impingement plate 83 between the front flange 81 and the rear flange 82, and the passage 91 provided in the front flange 81 connects to the passage 88 and the passages 92 provided in the trailing edge portion 43 connect to the chamber 77. Cooling air 59 in the cavity 45 is injected into the chamber 77 through the small holes 101 of the impingement plate 83 for cooling therearound, as shown by arrows of the air 59. On the other hand, cooling air which has flown through the passage 88 enters the passage 91 of the front flange 81 to join with the cooling air 59 in the chamber 77 both to be then discharged as the cooling air 60 through the passages 92 of the trailing edge portion 43.

In Fig. 10, being a cross sectional view taken on line G-G of Fig. 6, the recess portions 98, 99 are provided around the blade portion 25 and the passages 93, 94 are provided in the rails 96, 97, respectively. Also, the chambers 77, 78 are formed by the impingement plates 83, 84, respectively. Cooling air 75 flows into the cavity 45 from the leading edge passage 42 and flows therefrom into the chambers 77, 78 through the small holes 101 of the impingement plates 83, 84.

In Fig. 11, being a cross sectional view taken on line H-H of Fig. 6, the passages 90, 91 of the front flange 81 and the passages 93, 94 of the side edge portions are provided in both of the side edge portions of the inner shroud 26 and the passages 90, 91 connect respectively to the passage 88 of the leading edge portion 41.

In Fig. 12, being a cross sectional view taken on line J-J of Fig. 6, the passage 94 of the rail 97 is provided extending through the trailing edge portion 43 so that the cooling air 61 is discharged therefrom and the impingement plate 83 is provided between the front flange 81 and the rear flange 82.

In the second stage stationary blade of gas turbine described as above, the cooling air 57 from the recess portion 100 flows into the passage 88 of the leading edge portion 41 through the passage 90 of the front flange 81. There are provided the multiplicity of needle-like fins 99 in the passage 88, thereby cooling effect of the cooling air 57 is enhanced so that portions therearound are cooled efficiently. Then, the cooling air 57 bends approximately orthogonally at the passage 91 and flows into the chamber 77 formed by the impingement plate 83 to join with the cooling air flowing thereinto through the small holes 101 of the impingement plate 83 and flows together through the trailing edge portion 43 for cooling thereof and is discharged through the passages 92. Also, the cooling air which has been injected through the small holes 101 of the impingement plate 84 to enter the chamber 78 is likewise discharged through the passages 92.

Further, the portion 58 of the air which has entered the passage 88 passes through the passages 93, 94 in the rails 96, 97, respectively, of the side edge portions for cooling therearound and are discharged as the cooling air 61 from the trailing edge portion 43. Thus, the cooling air 75, 76 in the cavity 45 is portioned to be made effective use thereof respectively to flow through the passage 88, of which heat transfer is enhanced by the needle-like fins 89, the passages 93, 94 in the rails 96, 97 and the multiplicity of passages 92 in the trailing edge portion 43, thereby entire cooling of the inner shroud 26 is aimed to be performed efficiently.

That is, according to the air cooled system of the second stage stationary blade of gas turbine in the prior art as described above, in order to ensure the entire cooling effect of the inner shroud 26 thereof, the cooling air passes through the passage 88 and the needle-like fins 89 provided therein for enhancement of the cooling effect to further flow portionally into the chamber 77 formed by the impingement plate 83 through the passage 91 of the front flange 81 and also the cooling air is injected into the chambers 77, 78 through the small holes 101 of the impingement plates 83, 84 for cooling of the central portion and then both the cooling air joins together to flow through the multiplicity of passages 92 of the trailing edge portion 43 for cooling therearound and further the cooling air from the passage 88 of the leading edge portion 41 portionally flows through the passages 93, 94 of the rails 96, 97 of the side edge portions for cooling therearound.

According to the cooling structure mentioned above, however, while the entire inner shroud is cooled efficiently, the cooling air which has entered the passage 88 portionally flows out of the passage 91 for cooling of the central portion, hence in the leading edge portion 41 and the side edge portions which are exposed especially to the high temperature combustion gas, there occurs a shortage of the cooling air flowing in the passages 93, 94 of the side edge portions, resulting in insufficiency of cooling in the side edge portions.

Also, the cooling air entering the passage 88 of the leading edge portion 41 is a part of the cooling air entering the cavity 45 and comes from the recess portion 100 through the passage 90 and in order to further enhance the cooling effect of the leading edge portion 41, it is expected that amount of the cooling air flowing therein and flow velocity thereof are increased so as to enhance the cooling effect further.

SUMMARY OF THE INVENTION:

In view of the foregoing problem in the prior art, it is an object of the present invention to provide a gas turbine stationary blade in which an entire cooling effect of inner shroud is further enhanced by a construction made such that amount of cooling air entering a leading edge portion and flow velocity thereof are increased with cooling effect thereof being further enhanced by agitation of the cooling air and also cooling air flowing in both side edge portions is increased.

In order to attain said object, the present invention provides following means mentioned in (1) to (3): <ul id="ul0001" list-style="none" compact="compact"><li>(1) A gas turbine stationary blade constructed such that air from compressor is led into an outer shroud to be further led into a leading edge side passage and a trailing edge side passage, both provided in said stationary blade, as a cooling air of said stationary blade and then to be led into a cavity formed in an inner shroud to be portionally led from said cavity into spaces formed between said stationary blade and front and rear moving blades adjacent thereto as a seal air as well as to be portionally led from said cavity into said inner shroud to flow through a leading edge portion, both side edge portions, a central portion and a trailing edge portion, all of said inner shroud, as a cooling air of said inner shroud to be then discharged outside, characterized in that said gas turbine stationary blade comprises; a bottom plate for closing a passage connecting from said leading edge side passage to said cavity; and a passage for causing entire amount of the cooling air from said leading edge side passage to flow into a passage of said leading edge portion along said bottom plate, and the cooling air flown into the passage of said leading edge portion is caused to flow through said both side edge portions and said trailing edge portion to be then discharged outside.</li><li>(2) A gas turbine stationary blade as mentioned in (1) above, characterized in that there is provided an adjusting plate for adjusting a flow passage cross sectional area in the passage of said leading edge portion.</li><li>(3) A gas turbine stationary blade as mentioned in (1) above, characterized in that there are provided a plurality of turbulators in the passage of said leading edge portion.</li></ul>

In the gas turbine stationary blade mentioned in (1) above, the cooling air which has flown through the leading edge side passage for cooling the blade interior enters in its entire amount into the passage of the leading edge portion of the inner shroud for cooling of the leading edge portion and then is separated to flow into the passages of the side edge portions. The cooling air which has flown through the passages of the side edge portions for cooling thereof enters the trailing edge portion for cooling thereof and is then discharged outside.

Thus, the entire amount of the cooling air which has flown through the leading edge side passage for cooling of the blade interior enters the passage of the leading edge portion so that the leading edge portion which is exposed to a high temperature combustion gas and is in a severe temperature condition is cooled efficiently. The cooling air in the passage of the leading edge portion is then separated to flow through the respective side edge portions, thereby the side edge portions which are also exposed to the high temperature combustion gas are cooled efficiently. Then, the cooling air is discharged out of the trailing edge portion.

Also, the cooling air from the trailing edge side passage enters the central portion of the inner shroud to spread therearound for cooling the central portion and then is discharged outside through the trailing edge portion. In the prior art case, the construction is such that the cooling air to enter the leading edge portion once flows into a cavity to be then portionally flown to be used as a seal air and portionally flown into the leading edge portion to be used as a cooling air thereof, but in the present invention, the entire amount of the cooling air from the leading edge side passage flows directly into the leading edge portion, hence the air of high pressure can be supplied as it is with an increased amount of the air as compared with the prior art case.

The cooling air which has flown into the leading edge portion in the prior art case further flows portionally into the central portion of the inner shroud, but in the invention mentioned in (1) above, there is eliminated the passage connecting from the leading edge portion to the central portion and the entire amount of the air in the passage of the leading edge portion flows separatedly into the side edge portions, hence the leading edge portion and the side edge portions, both being in the severe temperature condition, are cooled efficiently as compared with the prior art case.

In the invention of (2) above, the flow passage cross sectional area of the leading edge portion is made narrower appropriately by the adjusting plate, hence the flow velocity of the cooling air therein is increased. Also, in the invention of (3) above, the turbulators are provided, hence the cooling effect of the leading edge portion is increased greatly by the agitating action of the turbulators as compared with the prior art case.

BRIEF DESCRIPTION OF THE DRAWINGS:

<ul id="ul0002" list-style="none" compact="compact"><li>Fig. 1 is a cross sectional view of a gas turbine which includes a stationary blade of an object of the present invention.</li><li>Fig. 2 is a cross sectional view of a gas turbine stationary blade of an embodiment according to the present invention, said gas turbine stationary blade being cut at its inner shroud portion along a turbine axial direction and seen from inner side thereof.</li><li>Fig. 3 is a cross sectional view taken on line A-A of Fig. 2.</li><li>Fig. 4 is a cross sectional view taken on line B-B of Fig. 2.</li><li>Fig. 5 is a cross sectional view taken on line C-C of Fig. 2.</li><li>Fig. 6 is a cross sectional view of a prior art gas turbine stationary blade, said gas turbine stationary blade being cut at its inner shroud portion along a turbine axial direction and seen from inner side thereof.</li><li>Fig. 7 is a cross sectional view taken on line D-D of Fig. 6.</li><li>Fig. 8 is a cross sectional view taken on line E-E of Fig. 6.</li><li>Fig. 9 is a cross sectional view taken on line F-F of Fig. 6.</li><li>Fig. 10 is a cross sectional view taken on line G-G of Fig. 6.</li><li>Fig. 11 is a cross sectional view taken on line H-H of Fig. 6.</li><li>Fig. 12 is a cross sectional view taken on line J-J of Fig. 6.</li></ul>

DESCRIPTION OF THE PREFERRED EMBODIMENTS:

Herebelow, description will be made on embodiments according to the present invention concretely with reference to figures. The present invention relates to a gas turbine stationary blade, and more specifically to a cooling structure of inner shroud of a second stage stationary blade of gas turbine. Fig. 1 is an entire cross sectional view of a gas turbine and a second stage stationary blade 17 shown there is the object of the present invention. Structure of other portions than the second stage stationary blade 17 of the present invention is same as that described in the column of the prior art with repeated description thereof being omitted here and featured portion of the present invention will be described with reference to Figs. 2 to 5.

Fig. 2 is a cross sectional view of the second stage stationary blade 17 which is cut along a turbine axial direction at approximately a central portion of its inner shroud 126 and seen from an inner side thereof, that is, on a rotor side. In Fig. 6, provided between a front flange 81 and a rear flange 82 of the inner shroud 126 are a rib 40 at a central portion, and a leading edge passage 42 and a trailing edge passage 44 mutually separated by the rib 40, and impingement plates 83, 84 therearound having a multiplicity of small holes 101. There are provided also rails 96, 97 on both side edge portions of the inner shroud 126 and passages 93, 94 in the rails 96, 97, respectively, and a multiplicity of passages 92 in a trailing edge portion 43. Structures of these component parts are same as those of the prior art shown in Fig. 6.

In a leading edge portion 41, there is provided a passage 188, which connects to a passage 90' provided in the front flange 81 so as to lead a cooling air thereinto. The passage 188 has a width of flow passage which is narrower than the prior art one, as described later, and has a plurality of turbulators 200 provided therein for enhancing further an agitating effect of internal air flow than the prior art needle-like fins.

Further, the prior art passage 91 (see Fig. 6) for outflow of the air is eliminated so that entire amount of the cooling air which has entered the passage 188 flows out into the passages 93, 94 provided in the rails 96, 97 of the side end portions with result that a cooling effect of the side edge portions is enhanced further.

Furthermore, a bottom plate 150, as described later, is provided at a bottom portion of the leading edge passage 42, so that entire amount of the cooling air flowing from the leading edge passage 42 flows into the passage 188 through the passage 90'. Thus, as compared with the prior art structure in which the air comes from the cavity, the cooling air of higher pressure can be supplied directly from the leading edge passage 42 and both the flow amount and the flow velocity of the air can be increased.

In the inner shroud 126 mentioned above, the entire amount of the cooling air supplied from the leading edge passage 42 enters the passage 188 through the passage 90' to be agitated by the turbulators 200 for cooling of the leading edge portion 41 with an enhanced heat transfer and is separated to flow into the respective passages 93, 94 in the rails 96, 97 of the side edge portions for cooling the side edge portions to be then discharged out of the passages 92 of the trailing edge portion 43 as air 61 after used for cooling.

On the other hand, as described later, cooling air supplied from the trailing edge passage 44 flows into a cavity 45 to be then injected through the small holes 101 of the impingement plates 83, 84 for cooling of a central portion of the inner shroud 126 with an impingement effect and is discharged out of the multiplicity of passages 92 of the trailing edge portion 43 as air 60 after used for cooling.

Fig. 3, being a cross sectional view taken on line A-A of Fig. 2, shows interiors of the stationary blade and the inner shroud. In Fig. 3, the second stage stationary blade 17 consists of a blade portion 25, an outer shroud 27 and the inner shroud 126. In the blade portion 25, there are provided the rib 40, and the leading edge passage 42 and the trailing edge passage 44 mutually separated by the rib 40. A cylindrical member 46 is provided in the leading edge passage 42 and a cylindrical member 47 is provided in the trailing edge passage 44, and a multiplicity of cooling holes 70, 71 are provided in side walls of the cylindrical members 46, 47, respectively. Also, cooling air holes 72, 73 are provided in bottom walls of the cylindrical members 46, 47, respectively.

In the inner shroud 126, the front flange 81 and the rear flange 82 are provided so as to form therebetween the cavity 45. In the cavity 45, the impingement plate 83 is provided so as to form a chamber 78 and also a bottom plate 150 is provided so as to close a bottom portion of the leading edge passage 42 to thereby form an opening portion 68. In the trailing edge portion 43, there are provided the multiplicity of passages 92 connecting to the cavity 45.

The opening portion 68 connects to the passage 90' of the front flange 81 so that entire amount of the cooling air from the leading edge passage 42 may flow into the passage 188. In the passage 188, an adjusting plate 151 is provided so as to make narrower a cross sectional area of flow passage of the passage 188 and to increase flow velocity of the cooling air. Also, in the passage 188, there are provided the turbulators 200 as mentioned above.

Cooling air 19' enters the cylindrical members 46, 47 and flows through the cooling air holes 70, 71 to impinge on wall surfaces of the leading edge and trailing edge passages 42, 44 for cooling of the wall surfaces with an increased heat transfer effect. The cooling air which has cooled the wall surface of the leading edge passage 42 flows to the opening portion 68 to join with the cooling air which has flown through the cooling air hole 72 of the bottom portion of the cylindrical member 46. The cooling air which has entered the cylindrical member 47 flows portionally into the cavity 45 through the cooling air hole 73 and portionally flows through the cooling air holes 71 for cooling of the wall surface of the trailing edge passage 44. The cooling air which has cooled the wall surface of the trailing edge passage 44 flows portionally through a trailing edge portion 29 of the blade portion 25 to be discharged outside therefrom and portionally flows into the cavity 45 to join with the cooling air which has entered there through the cooling air hole 73 and then enters the chamber 78 or chambers (not shown) through the impingement plates 83, 84 for cooling of a central portion of the inner shroud 126 and is discharged outside through the multiplicity of passages 92 of the trailing edge portion 43.

Also, as described in the prior art example, the cooling air in the cavity 45 flows out portionally through a hole 67 of a seal supporting portion 66 as shown by air 85 and 86. The air 85 flows into a space between the inner shroud 126 and a front stage moving blade thereof, thereby said space is maintained at a higher pressure tan in a passage through which an outside high temperature combustion gas 30 passes so that the high temperature gas is prevented from coming thereinto. Also, the air 86 flows through a seal 33 to enter a space between the inner shroud 126 and a rear stage moving blade thereof, thereby this space is maintained likewise at a higher pressure and the high temperature gas is prevented from coming thereinto.

As described above, the cooling air which has been supplied through the leading edge passage 42 for cooling of the blade portion 25 enters the opening portion 68 and entire amount of this air flows into the passage 188 through the passage 90' because of the bottom plate 150. In the passage 188, the cross sectional area thereof is adjusted by the adjusting plate 151 so as to become narrower and to increase the flow velocity of the air therein. Further, the air flow is agitated by the turbulators 200 and thereby the cooling effect is increased. Hence, as mentioned with respect to Fig. 2, both the leading edge portion 41 and the trailing edge portion 43 are cooled efficiently.

In Fig. 4, being a cross sectional view taken on line B-B of Fig. 2, between the front flange 81 and the rear flange 82 of the inner shroud 126, there are provided the impingement plate 84 having the multiplicity of small holes 101 and the bottom plate 150 for closing the bottom portion of the leading edge passage 42. In the leading edge portion 41, the passage 90' provided in the front flange 81 and a recess portion 100 connect to each other and the entire amount of the cooling air flown from the leading edge passage 42 flows into the passage 188 of the leading edge portion 41 through the passage 90'. In the passage 188, the adjusting plate 151 and the turbulators 200 are provided, as mentioned before. Also, the cooling air flown from the trailing edge passage 44 is injected through the small holes 101 of the impingement plate 84 into the chamber 78 formed by the impingement plate 84 and a recess portion 99, thus all these portions are cooled with enhanced cooling effect.

In Fig. 5, being an enlarged cross sectional view taken on line C-C of Fig. 2, the adjusting plate 151 is provided in the passage 188, thereby the flow passage cross sectional area is made narrower than the prior art case and the flow velocity of the air there is increased. Also, the turbulators 200 are provided to upper and lower wall surfaces of the passage 188, thereby the heat transfer effect by convection is increased.

In the gas turbine stationary blade having the air cooled structure of the second stage stationary blade 17 as described above, the passage 91 of the cooling air which has been provided in the front flange 81 of the leading edge portion 41 in the prior art case is eliminated and the entire amount of the cooling air in the passage 188 of the leading edge portion 41 is caused to flow through the passages 93, 94 provided in the rails 96, 97 of the side edge portions. In order to lead the entire amount of the cooling air which has flown through the leading edge passage 42 for cooling of the blade portion 25 into the passage 188, the bottom plate 150 is provided so as to close the bottom portion of the leading edge passage 42. Further, in order to increase the flow velocity of the air in the passage 188 of the leading edge portion 41, the adjusting plate 151 is provided. Also, the turbulators 200 are provided for increasing the cooling effect. Thus, following effects of the invention can be obtained.

The entire amount of the cooling air from the leading edge passage 42 flows into the passage 188 of the leading edge portion 41 and this air is used in its entire amount for cooling of the leading edge portion 41 without a portion thereof being taken for cooling of the central portion as has been done in the prior art case, hence the cooling effect of the leading edge portion 41, which is exposed to a high temperature gas and is in a severe temperature condition, is enhanced greatly as compared with the prior art one.

The adjusting plate 151 is provided in the passage 188 of the leading edge portion 41 so that the flow passage cross sectional area is made narrower and the flow velocity is increased, as compared with the prior art case, and further the turbulators 200 are provided in the passage 188, hence the cooling effect of the passage 188 is enhanced greatly as compared with the prior art case where only the needle-like fins are provided in the passage 88.

The entire amount of the cooling air which has entered the passage 188 of the leading edge portion 41 further flows separatedly into the passages 93, 94, respectively, of the rails 96, 97 of the side edge portions so that the air amount flowing in the passages 93, 94 increases as compared with the prior art case, hence the cooling effect of the side edge portions which are exposed to the high temperature gas increases. In the prior art case, the air which has entered the passage 88 portionally flows into the passage 91 of the front flange 81 for cooling of the central portion and portionally flows into the passages 93, 94. In the present invention, however, the passage 91 is eliminated, hence the amount of the cooling air flowing in the passages 93, 94 increases by that degree.

It is understood that the invention is not limited to the particular construction and arrangement herein illustrated and described but embraces such modified forms thereof as come within the scope of the following claims.