KR101317443B1 - A cooled blade of gas turbine - Google Patents

A cooled blade of gas turbine Download PDF

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Publication number
KR101317443B1
KR101317443B1 KR1020120112364A KR20120112364A KR101317443B1 KR 101317443 B1 KR101317443 B1 KR 101317443B1 KR 1020120112364 A KR1020120112364 A KR 1020120112364A KR 20120112364 A KR20120112364 A KR 20120112364A KR 101317443 B1 KR101317443 B1 KR 101317443B1
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South Korea
Prior art keywords
inlet
flow path
cooling
gas turbine
inner flow
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KR1020120112364A
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Korean (ko)
Inventor
곽재수
최은영
박정신
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한국항공대학교산학협력단
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Priority to KR1020120112364A priority Critical patent/KR101317443B1/en
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    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/18—Hollow blades, i.e. blades with cooling or heating channels or cavities; Heating, heat-insulating or cooling means on blades
    • F01D5/187—Convection cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D25/00—Component parts, details, or accessories, not provided for in, or of interest apart from, other groups
    • F01D25/08—Cooling; Heating; Heat-insulation
    • F01D25/12—Cooling
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
    • F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
    • F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
    • F01D5/12—Blades
    • F01D5/14—Form or construction
    • F01D5/147—Construction, i.e. structural features, e.g. of weight-saving hollow blades
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F02—COMBUSTION ENGINES; HOT-GAS OR COMBUSTION-PRODUCT ENGINE PLANTS
    • F02C—GAS-TURBINE PLANTS; AIR INTAKES FOR JET-PROPULSION PLANTS; CONTROLLING FUEL SUPPLY IN AIR-BREATHING JET-PROPULSION PLANTS
    • F02C7/00—Features, components parts, details or accessories, not provided for in, or of interest apart form groups F02C1/00 - F02C6/00; Air intakes for jet-propulsion plants
    • F02C7/12—Cooling of plants
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2220/00—Application
    • F05D2220/30—Application in turbines
    • F05D2220/32—Application in turbines in gas turbines
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F05—INDEXING SCHEMES RELATING TO ENGINES OR PUMPS IN VARIOUS SUBCLASSES OF CLASSES F01-F04
    • F05D—INDEXING SCHEME FOR ASPECTS RELATING TO NON-POSITIVE-DISPLACEMENT MACHINES OR ENGINES, GAS-TURBINES OR JET-PROPULSION PLANTS
    • F05D2260/00—Function
    • F05D2260/20—Heat transfer, e.g. cooling

Abstract

PURPOSE: A cooling blade of a gas turbine is provided to increase a heat transfer efficient by strengthening secondary flow by a lib because cold air flows into the inside by leaning to one side, by slantly installing multiple libs in an inner flow path and installing a structure on an inlet of the inner flow path. CONSTITUTION: A cooling blade of a gas turbine includes an airfoil (20) and an inner flow path (30). In the inner flow path of the airfoil, multiple libs (31) are slantly installed. On an inlet of the inner flow path, a structure (32) is installed to disturb the flow of cold air which flows into through an inlet port (11a). The structure is formed as multiple pins which are installed to make one side of the inlet of the inner flow path have higher density than the other side, or as a rectangular block shape which is installed to lean to one side of the inlet of the inner flow path, in order to make the inlet flow lean to one side. The inlet flow which leans to one side strengthens secondary flow by the libs in order to increase a heat transfer coefficient.

Description

가스터빈의 냉각블레이드{A Cooled Blade of Gas Turbine}Cooling Blade of Gas Turbine {A Cooled Blade of Gas Turbine}

본 발명은 가스터빈의 냉각블레이드에 관한 것으로서, 특히 내부유로의 열전달 계수를 증가시켜 열전달을 촉진함으로써 냉각성능을 향상시킬 수 있는 가스터빈의 냉각블레이드에 관한 것이다.
The present invention relates to a cooling blade of a gas turbine, and more particularly, to a cooling blade of a gas turbine which can improve cooling performance by increasing a heat transfer coefficient of an internal flow path to promote heat transfer.

최근의 가스터빈은 출력과 효율을 향상시키기 위해 터빈으로 유입되는 연소가스의 온도를 높이고 있고, 따라서 터빈 내 각 부품들의 열부하 또한 증가하고 있다. 이렇게 열부하가 증가됨으로써 터빈 블레이드 재질의 한계로 인해 터빈 블레이드의 파손이 야기된다.Recent gas turbines are increasing the temperature of the combustion gases entering the turbine to improve power and efficiency, and thus the heat load of each component in the turbine is also increasing. This increased heat load causes damage to the turbine blades due to the limitations of the turbine blade material.

이러한 문제를 개선하기 위해서 최근에 표면 처리 등 가공적인 측면에서 요소의 내열성을 향상시키기 위한 연구들이 진행되고 있다.In order to improve such a problem, researches to improve the heat resistance of the element in terms of processing such as surface treatment have recently been conducted.

그러나, 이러한 방법에만 의존하여 블레이드를 보호하고자 하는 것 역시 근본적인 문제해결에 한계를 보이기 있기 때문에 다양한 냉각 구조를 적용하는 시도가 병행되고 있다.However, attempts to apply various cooling schemes have been made in parallel because attempting to protect the blades only by this method also presents a limitation in fundamental problem solving.

이때의 냉각방식으로는 막냉각, 충돌제트 냉각, 강제대류 냉각 등이 고려될 수 있다. 이 중에서 특히 강제 대류 냉각은 다른 냉각 방법에 비해서 블레이드 표면에 추가적인 가공을 하지 않고 블레이드 내부에 있는 유로를 통해서 냉각시키는 방법이지만, 아직까지는 그 효과가 충분치 않은 문제점이 지적되고 있다.
At this time, as the cooling method, film cooling, impingement jet cooling, forced convection cooling, etc. may be considered. Among these, forced convection cooling is a method of cooling through the flow path inside the blade without further processing on the blade surface as compared to other cooling methods, but the problem is not pointed out enough so far.

본 발명은 상기한 종래기술의 문제점을 해결하기 위하여 안출된 것으로서, 내부유로에 복수개의 리브를 경사지게 설치함과 아울러 내부유로의 입구부분에 구조물을 설치하여 냉각공기가 한쪽으로 치우쳐 유입되도록 함으로써 리브에 의한 이차 유동을 강화시켜 열전달 계수를 증가시키고, 이런 결과로 열전달을 촉진하여 냉각 성능을 형상시켜 수명을 연장할 수 있는 가스터빈의 냉각블레이드를 제공하는데 그 목적이 있다.
The present invention has been made in order to solve the above problems of the prior art, by installing a plurality of ribs inclined in the inner flow passage and by installing a structure at the inlet portion of the inner flow passage so that the cooling air flows to one side inflow into the rib The purpose of the present invention is to provide a cooling blade of a gas turbine that can increase the heat transfer coefficient by enhancing the secondary flow, thereby promoting heat transfer to shape the cooling performance and extending the life.

상기한 과제를 해결하기 위한 본 발명에 의한 가스터빈의 냉각블레이드는 터빈에 결합되는 루트부의 상측에 고온 고압의 연소가스와 충돌하는 에어포일이 일체로 형성되고, 상기 루트부에 형성된 유입구를 통해 유입된 냉각공기가 상기 에어포일의 내부에 형성된 내부유로를 통과한 후 상기 에어포일에 형성된 분사홀을 통하여 냉각공기가 배출되도록 하는 가스터빈의 냉각블레이드에 관한 것으로서, 상기 에어포일의 내부유로에는 복수개의 리브가 일정간격으로 경사지게 설치되고, 상기 내부유로의 입구에는 상기 유입구를 통하여 유입된 냉각공기의 유동을 교란시키는 구조물이 설치된다.Cooling blades of the gas turbine according to the present invention for solving the above problems is formed integrally with the airfoil colliding with the combustion gas of the high temperature and high pressure on the upper side of the root portion coupled to the turbine, inflow through the inlet formed in the root portion The cooling blade of the gas turbine for allowing the cooling air is discharged through the injection hole formed in the air foil after the cooled air passes through the internal flow path formed in the air foil, the plurality of internal flow paths of the air foil Ribs are installed to be inclined at a predetermined interval, and a structure for disturbing the flow of the cooling air introduced through the inlet is installed at the inlet of the internal flow path.

여기서, 상기 구조물은 핀(Pin) 형태로 형성될 수도 있는데, 이때 상기 핀 형태의 구조물은 1열 또는 복수의 열로 설치되고, 상기 내부유로의 입구부분 일측이 타측보다 높은 밀도를 갖도록 설치된다.Here, the structure may be formed in the form of a pin (pin), wherein the structure of the pin shape is installed in one row or a plurality of rows, one side of the inlet portion of the inner flow passage is installed to have a higher density than the other side.

그리고, 상기 구조물은 사각 형태의 블록 형상으로 형성될 수도 있는데, 이때 상기 블록 형상의 구조물은 상기 내부유로의 입구부분 일측에 치우쳐 설치된다.
In addition, the structure may be formed in a block shape of a rectangular shape, wherein the block-shaped structure is installed on one side of the inlet portion of the internal flow path.

상기와 같이 구성되는 본 발명의 가스터빈의 냉각블레이드는 내부유로의 입구에 설치된 구조물에 의하여 냉각공기가 한쪽으로 치우쳐 내부유로에 유입되면 내부유로에 설치된 리브에 의한 냉각공기의 2차 유동교란이 강화되어 열전단 계수가 증가되고, 이러한 열전달 계수의 증가에 의하여 열전달이 촉진되어 냉각성능이 향상되고 블레이드의 수명이 연장되는 이점이 있다.
Cooling blades of the gas turbine of the present invention constituted as described above is reinforced by the secondary flow disturbance of the cooling air by the rib installed in the inner flow path when the cooling air flows to one side by the structure installed at the inlet of the inner flow path. The heat transfer coefficient is increased, and heat transfer is promoted by the increase of the heat transfer coefficient, thereby improving cooling performance and extending the life of the blade.

도 1은 본 발명에 의한 가스터빈의 냉각블레이드의 일 실시예를 보인 단면도.
도 2는 본 발명에 의한 가스터빈의 냉각블레이드의 다른 실시예를 보인 단면도.
도 3a 내지 도 3b는 도 1에 도시된 가스터빈의 냉각블레이드의 요부와 냉각공기 유동을 보인 단면도.
도 4a 내지 도 4b는 도 2에 도시된 가스터빈의 냉각블레이드의 요부와 냉각공기 유동을 보인 단면도.1 is a cross-sectional view showing an embodiment of a cooling blade of a gas turbine according to the present invention.
Figure 2 is a cross-sectional view showing another embodiment of a cooling blade of the gas turbine according to the present invention.
3a to 3b are cross-sectional views showing the main portion and the cooling air flow of the cooling blade of the gas turbine shown in FIG.
4a to 4b are cross-sectional views showing the main portion and the cooling air flow of the cooling blade of the gas turbine shown in FIG.

이하, 본 발명에 의한 가스터빈의 냉각블레이드의 실시 예를 첨부된 도면을 참조하여 상세히 설명한다.Hereinafter, an embodiment of a cooling blade of a gas turbine according to the present invention will be described in detail with reference to the accompanying drawings.

도 1은 본 발명에 의한 가스터빈의 냉각블레이드의 일 실시예를 보인 단면도이고, 도 2는 본 발명에 의한 가스터빈의 냉각블레이드의 다른 실시예를 보인 단면도이다.1 is a cross-sectional view showing an embodiment of a cooling blade of a gas turbine according to the present invention, Figure 2 is a cross-sectional view showing another embodiment of a cooling blade of a gas turbine according to the present invention.

그리고, 도 3a 내지 도 3b는 도 1에 도시된 가스터빈의 냉각블레이드의 요부와 냉각공기 유동을 보인 단면도이며, 도 4a 내지 도 4b는 도 2에 도시된 가스터빈의 냉각블레이드의 요부와 냉각공기 유동을 보인 단면도이다.
3A to 3B are cross-sectional views illustrating main parts and cooling air flow of the cooling blades of the gas turbine shown in FIG. 1, and FIGS. 4A to 4B are main parts and cooling air of the cooling blades of the gas turbine shown in FIG. 2. It is sectional drawing which shows flow.

본 발명에 의한 가스터빈의 냉각블레이드에 관하여 설명하기 전에, 가스터빈의 냉각블레이드는 고정된 것과 회전하는 것이 있는데, 본 발명은 이 고정형과 회전형 모두에 적용가능한 것임을 미리 밝혀두고 본 발명에 의한 가스터빈의 냉각블레이드에 관한 실시예에 대하여 설명하도록 한다.Before describing the cooling blades of the gas turbine according to the present invention, the cooling blades of the gas turbine are fixed and rotating, but the gas according to the present invention is known in advance that the present invention is applicable to both the fixed and the rotating type. An embodiment of a cooling blade of a turbine will be described.

본 발명에 의한 가스터빈의 냉각블레이드는 터빈의 외측면에 결합되어 고온 고압의 연소가스에 의해 터빈이 회전 작동되도록 하는 것으로서, 생크(11)와 플랫폼(12)으로 구성되어 터빈에 견고하게 결합되는 루트부(10)와, 상기 루트부(10)의 상측에 일체로 형성되어 전후면의 압력차에 의해 터빈이 회전되도록 하는 에어포일(20)과, 상기 에어포일(20)의 내부에 형성되는 내부유로(30)를 포함하여 구성된다.The cooling blade of the gas turbine according to the present invention is to be coupled to the outer surface of the turbine so that the turbine is rotated by the combustion gas of high temperature and high pressure, it is composed of the shank 11 and the platform 12 is firmly coupled to the turbine Root portion 10, an airfoil 20 formed integrally with the upper portion of the root portion 10 so that the turbine is rotated by the pressure difference between the front and rear surfaces, and is formed inside the airfoil 20 It is configured to include an internal flow path (30).

상기 루트부(10)의 생크(11)에는 하나 이상의 유입구(11a)가 형성된다. 상기 생크(11)에 형성된 유입구(11a)를 통하여 연소가스보다 낮은 온도의 냉각가스가 상기 에어포일(20) 내부로 유입된다.One or more inlets 11a are formed in the shank 11 of the root portion 10. Cooling gas having a lower temperature than the combustion gas is introduced into the airfoil 20 through the inlet port 11a formed in the shank 11.

상기 에어포일(20)은 상기 루트부(10)의 상측에 일체로 형성되는 것으로서, 고온 고압의 연소가스와 충돌한다. 이러한 에어포일(20)은 연소가스가 유입되는 전면으로는 외측방으로 볼록한 곡면을 이루며 돌출된 흡입면(21)이 형성되고, 후면으로는 상기 흡입면(21) 측으로 오목하게 함몰된 곡면을 이루는 압력면(22)이 형성된다. 따라서, 상기 에어포일(20)의 전후의 압력차, 즉 흡입면(21)과 압력면(22)의 압력차가 극대화되면서 원활한 공기의 흐름이 이루어진다.The air foil 20 is formed integrally with the root portion 10, and collides with the combustion gas of high temperature and high pressure. The airfoil 20 has a curved surface convex toward the outside toward the front surface where the combustion gas is introduced, and a protruding suction surface 21 is formed, and a rear surface of the air foil 20 concave concave recessed toward the suction surface 21 side. The pressure surface 22 is formed. Therefore, while the pressure difference between the front and rear of the air foil 20, that is, the pressure difference between the suction surface 21 and the pressure surface 22 is maximized, smooth air flow is achieved.

그리고, 상기 에어포일(20)에는 분사홀(23)이 형성되어, 상기 루트부(10)의 생크(11)에 형성된 유입구(11a)를 통하여 유입된 냉각공기가 에어포일(20)의 내부에 형성된 내부유로(30)를 통과한 후 분사홀(23)을 통하여 외부로 배출된다.
In addition, an injection hole 23 is formed in the air foil 20, and cooling air introduced through the inlet 11 a formed in the shank 11 of the root part 10 is formed in the air foil 20. After passing through the formed internal flow path 30 is discharged to the outside through the injection hole (23).

상기 내부유로(30)는 연소가스보다 낮은 온도의 냉각공기가 유입되어 이 냉각공기가 블레이드의 내측면과 열교환하면서 블레이드의 전체적인 온도가 하강되도록 하는 것으로서, 상기 에어포일(20)의 내부에 형성된다. 이러한 내부유로(30)에는 복수개의 리브(31)가 일정간격으로 설치된다. 그리고, 이 리브(31)는 내부유로(30)에 일정각도 경사지게 설치된다.The internal flow path 30 is a cooling air at a lower temperature than the combustion gas flows in such a way that the overall temperature of the blade is lowered while the cooling air exchanges with the inner surface of the blade, and is formed inside the airfoil 20. . A plurality of ribs 31 are provided in the internal flow path 30 at regular intervals. The rib 31 is inclined at an angle to the internal flow path 30.

이렇게 리브(31)가 내부유로(30)에 경사지게 설치됨으로써 상기 루트부(10)의 생크(11)에 형성된 유입구(11a)를 통하여 내부유로(30)에 유입된 냉각공기는 리브(31)에 충돌되면서 그 유동에 교란이 발생되어 블레이드의 내측면과 열교환이 더욱 활발하게 일어나게 된다. 내부유로(30)에 리브(31)를 설치하는 것만으로도 냉각공기의 유동에 교란을 발생시켜 내부 열전달을 촉진할 수 있는데, 리브(31)를 일정각도 경사지게 설치하면 유동 교란이 더 확대되어 열전달이 더욱 촉진되는 효과가 있다.Thus, the ribs 31 are inclined in the inner flow path 30 so that the cooling air introduced into the inner flow path 30 through the inlet 11a formed in the shank 11 of the root portion 10 is transferred to the ribs 31. As the collision occurs, disturbances occur in the flow, and heat exchange with the inner surface of the blade occurs more actively. The installation of the ribs 31 in the internal flow path 30 may only cause disturbances in the flow of the cooling air, thereby facilitating internal heat transfer. When the ribs 31 are inclined at an angle, the flow disturbances are further enlarged. This has a further promoting effect.

그리고, 상기 내부유로(30)의 입구에는 상기 유입구(11a)를 통하여 유입된 냉각공기가 상기 내부유로(30) 상으로 진입하기 전에 미리 1차적으로 유동을 교란시키기 위하여 구조물(32)이 설치된다. In addition, a structure 32 is installed at the inlet of the internal flow path 30 to firstly disturb the flow before the cooling air introduced through the inlet 11a enters the internal flow path 30. .

따라서, 상기 루트부(10)의 생크(11)에 형성된 유입구(11a)를 통하여 유입된 냉각공기는 상기 내부유로(30)의 입구에 설치된 구조물(32)에 충돌하면서 1차 유동교란이 발생되고, 연이어 유동이 교란된 상태로 내부유로(30) 상으로 진입한 냉각공기는 다시 경사진 리브(31)에 충돌하여 2차 유동교란이 발생된다. 이러한 냉각공기의 유동이 교란됨으로써 내부유로의 열전달 계수는 더욱 증가된다.Therefore, the cooling air introduced through the inlet (11a) formed in the shank 11 of the root portion 10 impinges on the structure 32 installed at the inlet of the internal flow path 30, the primary flow disturbance is generated , The cooling air entering the internal flow path 30 in a state in which the flow is subsequently disturbed collides with the inclined rib 31 again to generate secondary flow disturbance. As the flow of cooling air is disturbed, the heat transfer coefficient of the internal flow path is further increased.

여기서, 상기 구조물(32)은 핀(Pin) 형태로 형성될 수 있다. 이렇게 구조물(32)을 작은 지름을 갖는 원기둥 형태의 핀 형상으로 형성시킬 경우에는 1열 또는 복수의 열로 내부유로(30)의 입구에 설치한다. 그리고, 이렇게 구조물(32)을 복수개의 핀 형태로 구성할 경우에는 상기 내부유로(30)의 입구부분 일측이 입구부분의 타측보다 높은 밀도로 설치한다. 이렇게 핀 형태의 구조물(32)을 설치할 때 일측의 밀도가 타측의 밀도보다 높게 설치하는 이유는 냉각공기의 입구 유동이 한쪽으로 치우치게 유도함으로써 내부유로(30)에 유입된 후 리브(31)에 의한 2차 유동교란을 더욱 강화시켜 내부 열전달 계수를 더욱 증가시키기 위함이다.Here, the structure 32 may be formed in the shape of a pin. When the structure 32 is formed in a cylindrical pin shape having a small diameter, the structure 32 is installed at the inlet of the internal passage 30 in one row or a plurality of rows. When the structure 32 is configured in the form of a plurality of fins, one side of the inlet portion of the inner passage 30 is installed at a higher density than the other side of the inlet portion. The reason why the density of one side is installed higher than the density of the other side when installing the fin-shaped structure 32 is induced by the inlet flow of the cooling air to one side and introduced into the internal flow path 30 by the ribs 31. This is to further increase the internal heat transfer coefficient by further strengthening the secondary flow disturbance.

또한, 상기 구조물(32)은 사각 형태의 블록 형상으로 형성시킬 수도 있다. 이렇게 구조물(32)을 블록 형상으로 형성시킬 때는 이 구조물(32)을 내부유로(30)의 입구에 설치할 때 입구부분의 중앙에 설치하지 않고 입구부분의 일측에 치우치게 설치한다. 이렇게 블록 형상의 구조물(32)을 내부유로(30)의 입구 일측에 치우치게 설치하는 이유는 냉각공기의 입구 유동이 한쪽으로 치우치게 유입되도록 유도함으로써 내부유로(30)에 유입된 후 리브(31)에 의한 2차 유동을 더욱 강화시켜 내부 열전달 계수를 더욱 증가시키기 위함이다.
In addition, the structure 32 may be formed in a rectangular block shape. When the structure 32 is formed in a block shape like this, when the structure 32 is installed at the inlet of the internal flow path 30, it is installed at one side of the inlet part without being installed at the center of the inlet part. The reason why the block-shaped structure 32 is installed on one side of the inlet side of the inner passage 30 is to induce the inlet flow of the cooling air to one side so that the inlet flows to the inner passage 30 and then to the rib 31. This is to further increase the internal heat transfer coefficient by further strengthening the secondary flow.

10: 루트부 11: 생크
11a: 유입구 12: 플랫폼
20: 에어포일 21: 흡입면
22: 압력면 23: 분사홀
30: 내부유로 31: 리브
32: 구조물10: root part 11: shank
11a: inlet 12: platform
20: airfoil 21: suction surface
22: pressure surface 23: injection hole
30: Internal Euro 31: Rib
32: Structure

Claims (7)

터빈에 결합되는 루트부(10)의 상측에 고온 고압의 연소가스와 충돌하는 에어포일(20)이 일체로 형성되고, 상기 루트부(10)에 형성된 유입구(11a)를 통해 유입된 냉각공기가 상기 에어포일(20)의 내부에 형성된 내부유로(30)를 통과한 후 상기 에어포일(20)에 형성된 분사홀(23)을 통하여 냉각공기가 배출되도록 하는 가스터빈의 냉각블레이드에 있어서,
상기 에어포일(20)의 내부유로(30)에는 복수개의 리브(31)가 일정간격으로 경사지게 설치되고, 상기 내부유로(30)의 입구에는 상기 유입구(11a)를 통하여 유입된 냉각공기의 유동을 교란시키는 구조물(32)이 설치되며,
상기 구조물(32)이 상기 내부유로(30)의 입구부분 일측이 타측보다 높은 밀도를 갖도록 설치되는 다수의 핀(Pin), 또는 상기 내부유로(30)의 입구부분 일측에 치우쳐 설치되는 사각 형태의 블록 형상으로 형성되어 입구 유동을 한쪽으로 치우치도록 함으로써 한쪽으로 치우친 입구 유동이 상기 리브(31)에 의한 이차유동을 강화시켜 내부 열전달 계수를 증가시킬 수 있도록 한 것을 특징으로 하는 가스터빈의 냉각블레이드.The airfoil 20 colliding with the combustion gas of high temperature and high pressure is integrally formed on the root portion 10 coupled to the turbine, and cooling air introduced through the inlet port 11a formed in the root portion 10 is integrated. In the cooling blade of the gas turbine for passing the cooling air through the injection hole 23 formed in the air foil 20 after passing through the internal flow path 30 formed in the air foil 20,
A plurality of ribs 31 are installed to be inclined at a predetermined interval in the internal flow path 30 of the airfoil 20, and the inlet of the internal flow path 30 allows the flow of cooling air introduced through the inlet port 11a. Disturbing structure 32 is installed,
The structure 32 has a plurality of pins (Pin) is installed so that one side of the inlet portion of the inner passage 30 has a higher density than the other side, or of a square shape installed to be biased to one side of the inlet portion of the inner passage (30). Cooling blades for gas turbines, which are formed in a block shape and have an inlet flow biased to one side so that the inlet flow biased to one side can enhance the secondary flow by the ribs 31 to increase the internal heat transfer coefficient. . 삭제delete 삭제delete 삭제delete 삭제delete 삭제delete 삭제delete
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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
KR101670230B1 (en) 2016-06-03 2016-11-09 (유)한성산기 Pump capable of operating without water
KR101715604B1 (en) 2016-06-03 2017-03-13 (유)한성산기 Vibration reduced pump
WO2019245237A1 (en) * 2018-06-21 2019-12-26 한국기계연구원 Gas turbine
US10801332B2 (en) 2016-05-20 2020-10-13 Hanwha Aerospace Co., Ltd. Core for casting turbine blade, method of manufacturing the core, and turbine blade manufactured using the core
KR102180396B1 (en) * 2019-06-10 2020-11-18 두산중공업 주식회사 Airfoil and gas turbine comprising it

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JPH10212903A (en) * 1997-01-28 1998-08-11 Mitsubishi Heavy Ind Ltd Gas turbine blade
JPH1172003A (en) * 1997-06-26 1999-03-16 Soc Natl Etud Constr Mot Aviat <Snecma> Turbine blade cooled by spiral gradient, cascade shock and fastener mechanism in double surface
KR20020069462A (en) * 2001-02-27 2002-09-04 조형희 Discrete rib arrangements in turbine blade cooling passage
JP2010190198A (en) 2009-02-20 2010-09-02 Mitsubishi Heavy Ind Ltd Turbine blade

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* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
JPH10212903A (en) * 1997-01-28 1998-08-11 Mitsubishi Heavy Ind Ltd Gas turbine blade
JPH1172003A (en) * 1997-06-26 1999-03-16 Soc Natl Etud Constr Mot Aviat <Snecma> Turbine blade cooled by spiral gradient, cascade shock and fastener mechanism in double surface
KR20020069462A (en) * 2001-02-27 2002-09-04 조형희 Discrete rib arrangements in turbine blade cooling passage
JP2010190198A (en) 2009-02-20 2010-09-02 Mitsubishi Heavy Ind Ltd Turbine blade

Cited By (5)

* Cited by examiner, † Cited by third party
Publication number Priority date Publication date Assignee Title
US10801332B2 (en) 2016-05-20 2020-10-13 Hanwha Aerospace Co., Ltd. Core for casting turbine blade, method of manufacturing the core, and turbine blade manufactured using the core
KR101670230B1 (en) 2016-06-03 2016-11-09 (유)한성산기 Pump capable of operating without water
KR101715604B1 (en) 2016-06-03 2017-03-13 (유)한성산기 Vibration reduced pump
WO2019245237A1 (en) * 2018-06-21 2019-12-26 한국기계연구원 Gas turbine
KR102180396B1 (en) * 2019-06-10 2020-11-18 두산중공업 주식회사 Airfoil and gas turbine comprising it

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