CN113123833B - Turbine outer ring block air supply structure with separated air supply - Google Patents
Turbine outer ring block air supply structure with separated air supply Download PDFInfo
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- CN113123833B CN113123833B CN202110323611.XA CN202110323611A CN113123833B CN 113123833 B CN113123833 B CN 113123833B CN 202110323611 A CN202110323611 A CN 202110323611A CN 113123833 B CN113123833 B CN 113123833B
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- turbine
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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
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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
- 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
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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
- F01D5/186—Film cooling
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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
- F01D5/188—Convection cooling with an insert in the blade cavity to guide the cooling fluid, e.g. forming a separation wall
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention belongs to the field of thermal protection of aero-engines, and particularly relates to a turbine outer ring block air supply structure capable of accurately supplying air in a cavity. According to the invention, a layer of throttling cavity is additionally arranged between the air supply cavity and the impact cavity of the turbine outer ring block, the throttling cavity is axially divided into a plurality of independent composite cooling throttling sub-cavities along the turbine, and the throttling cavity corresponds to the throttling cavity, and the composite cooling impact cavity is also axially divided into a plurality of independent impact sub-cavities along the turbine, so that the whole outer ring block is axially divided into a plurality of cooling units which are not interfered with each other, the problem that the outer ring block is poor before cooling and rich after cooling due to a single impact cavity is avoided, a pointed cooling structure is formed, the utilization rate of cold air is improved, and the performance of an engine is enhanced.
Description
Technical Field
The invention belongs to the field of thermal protection of aero-engines, and particularly relates to a turbine outer ring block air supply structure capable of accurately supplying air in a cavity.
Background
At present, the turbine front temperature of the aero-engine reaches about 2000K, and according to thermodynamic cycle, the turbine front temperature of the aero-engine is continuously increased along with the improvement of performance, so that the cooling of the outer ring of the turbine is very important. The bleed air flow of the existing engine air system reaches 25%, the bleed air flow of the existing engine air system is possibly higher in the future, the improvement of the performance of the engine is greatly inhibited by the increase of the bleed air flow of the existing engine air system, and therefore how to reduce the bleed air flow of the existing engine air system becomes a key problem of the cooling design of the hot end component of the engine.
The engine outer ring block is arranged at the turbine movable blade outer casing, and during the operation of the engine, the pressure at the top of the turbine movable blade has a large pressure gradient along the axial direction of the engine, as shown in fig. 1, so that the outlet pressure of the outer ring block along the axial direction of the engine has a great pressure gradient. In the prior art, the same air inlet pressure is generally adopted at the inlet of the outer ring block, as shown in fig. 2, the outer ring block has an impact-turbulence-air film composite cooling three-layer structure, the pressure of the inlet of the impact hole is the same, the impact action on the impact target surface of the impact cavity is basically the same, the inlet of the air film hole in the impact cavity is under the condition of basically the same pressure boundary, the air blowing ratio at the outlet of the air film hole is lower than that at the front part and higher than that at the rear part, the air blowing ratio at the position of the outer ring block corresponding to the inlet of the turbine movable blade is too small, and the air blowing ratio at the outer ring block corresponding to the outlet of the movable blade is far greater than that at the inlet. The result of this is that the outer ring piece surface air film covering effect is good before and bad after, simultaneously because along the main stream temperature gradient of engine axial is great, the outer ring piece cooling is rich after poor before that. Therefore, if the cooling effect is to be improved, the cold air flow needs to be improved to compensate for the short plate for cooling the high-temperature area of the outer ring block. This measure, although ensuring adequate cooling of the outer ring block, increases the bleed air ratio of the air system, resulting in a reduction in engine performance.
Disclosure of Invention
In order to solve the problems that in the cooling process of an outer ring block, due to the fact that the axial pressure gradient of an outlet of the outer ring block is large, the temperature gradient of a main stream side is large, cold air distribution is unreasonable, cooling of the outer ring block is affected, and excessive loss of the cold air is avoided, the invention designs a turbine outer ring block air supply structure for accurately supplying air in a cavity division mode. According to the invention, the throttling function is added in front of the impact hole, so that the air inlet and exhaust pressure of the impact-turbulence-air film composite cooling along the axial direction of the turbine can have a tendency of reduction, and the problems of insufficient cooling caused by too large blowing ratio of the air film hole and performance reduction caused by too large flow due to too high air inlet pressure of the outer ring block corresponding to the outlet of the movable blade and the like are avoided.
In order to achieve the purpose, the invention provides a gas supply structure of a turbine outer ring block for supplying gas in a cavity-divided manner, which sequentially comprises a throttle plate, a throttle cavity, an impact plate, an impact cavity and a gas film plate along the flowing direction of a fluid; the throttling plate and the impact plate are arranged at intervals, the throttling cavity is formed between the throttling plate and the impact plate, and a plurality of throttling holes which are arranged in an array mode and used for allowing fluid to flow into the throttling cavity are formed in the throttling plate; a plurality of first column ribs are arranged in the throttling cavity at intervals in rows so as to divide the throttling cavity into a plurality of mutually independent throttling branch cavities; the impact plate and the air film plate are arranged at intervals, the impact cavity is formed between the impact plate and the air film plate, and a plurality of impact holes which are arranged in an array and used for allowing fluid to flow into the impact cavity from each throttling sub-cavity are formed in the impact plate; a plurality of second column ribs are arranged in the impact cavity at intervals in a row so as to divide the impact cavity into a plurality of independent impact sub-cavities; the air film plate is provided with a plurality of air film holes which are arranged in an array and used for fluid to flow out of each impact sub-cavity; each impact sub-cavity comprises a plurality of turbulence columns; the throttling subchambers and the impact subchambers are equal in number and correspond in a row, so that the gas supply structure is divided into a plurality of cooling units which are independent of each other; the length direction of each first column rib and each second column rib is perpendicular to the axial direction of the turbine.
In some embodiments, the total orifice area of each cooling unit is configured such that the flow rate of fluid entering each cooling unit varies in a gradient along the turbine axial direction.
In some embodiments, the total orifice area setting process for each cooling unit is as follows:
1) calculating and obtaining pressure and temperature distribution of an interface of a turbine main flow and an outer ring block along the axial direction of the turbine;
2) dividing the gas supply structure into n cooling units according to the width of each cooling unit; dividing the interface of the main flow of the turbine and the outer ring block into n units, wherein the n units correspond to the positions of film hole outlets of the n cooling units one by one, and determining the average temperature and the average pressure under each unit based on the pressure and the temperature distribution of the interface of the main flow of the turbine and the outer ring block of the turbine along the axial direction of the turbine, which are obtained in the step 1), namely determining the average temperature and the average pressure at the outlet of each cooling unit;
3) analyzing the flow characteristics of the single cooling unit, acquiring the relation between the inlet-outlet pressure ratio and the flow at different outlet temperatures of the single cooling unit, and determining the inlet-outlet pressure boundary and the temperature boundary of each cooling unit based on the average temperature and the average pressure at the outlet of each cooling unit determined in the step 2);
4) determining the flow rate of cold air flowing through each cooling unit according to the allowable temperature requirement of the material, and determining the total area of the throttling hole at the inlet of each cooling unit according to the relationship between the inlet-outlet pressure ratio and the flow rate of the cold air flowing through the single cooling unit at different outlet temperatures acquired in the step 3).
In some embodiments, the total orifice area of each cooling unit is configured to decrease gradually in the turbine axial direction.
In some embodiments, the orifice, impingement holes, and film holes are circular holes, the hole size being set according to the cooling air requirements of the cooling unit.
In some embodiments, the turbulence column is a cylindrical turbulence column or a chevron-shaped turbulence column.
The invention has the beneficial effects that:
1) the invention changes the defect of poor air distribution and rich air distribution of the outer ring block by accurately supplying air in the separated cavity, ensures the reasonable distribution of the cold air, and reduces the increase of air entraining quantity of the air system caused by the small quantity of the cold air in the high-temperature part;
2) the invention ensures the blowing ratio of the main flow merged into the outlet of the outer ring block in a more reasonable cooling mode, and avoids the influence on the main flow caused by overlarge blowing ratio;
3) the outer ring block is axially divided into a plurality of cooling structures which are not interfered with each other, namely bleed air of an air system is firstly distributed at an inlet through the throttling holes with different throttling capacities along the axial direction, and the flow resistance of the whole sub-cavity is adjusted, so that different cold air supplies are realized after the cold air flows through each cavity, a certain flow gradient is ensured along the axial direction of an engine, sufficient cooling of each position of the outer ring block is ensured, and the integral cooling effect of the outer ring block is enhanced.
Drawings
FIG. 1 is an outer ring block outlet turbine bucket tip pressure profile;
FIG. 2 is a schematic diagram of an outer ring block structure of a prior art impingement-turbulator-film hybrid cooling system;
FIG. 3 is a schematic illustration of a chambered gas feed turbine outer ring block gas feed configuration according to an embodiment of the present invention;
FIG. 4 is a schematic view of an orifice distribution according to an embodiment of the present invention;
FIG. 5 is a schematic view of a double-walled chevron turbulator structure in accordance with an embodiment of the present invention;
FIG. 6 is a cross-sectional view of a double-walled chevron turbulator structural element in accordance with an embodiment of the present invention.
Detailed Description
According to the invention, a layer of throttling cavity is additionally arranged between the air supply cavity and the impact cavity of the turbine outer ring block, the throttling cavity is axially divided into a plurality of independent composite cooling throttling sub-cavities along the turbine, the throttling cavity corresponds to the throttling cavity, and the composite cooling impact cavity is also axially divided into a plurality of independent impact sub-cavities along the turbine, so that the whole outer ring block is axially divided into a plurality of cooling units which are not interfered with each other along the turbine, the problem that the outer ring block is poor before cooling and rich after cooling due to a single impact cavity is avoided, a targeted cooling structure is formed, the utilization rate of cold air is improved, and the performance of the engine is improved on the premise of the same turbine front temperature.
The invention is further described below with reference to the accompanying drawings and examples, it being understood that the examples described below are intended to facilitate the understanding of the invention, and are not intended to limit it in any way.
As shown in fig. 3, the turbine outer ring block air supply structure for divided chamber air supply of the present embodiment includes a throttle plate 1, a throttle chamber 2, an impingement plate 3, an impingement chamber 4, and a gas film plate 5 in sequence along the flow direction of the cool air (shown from top to bottom in fig. 3).
The throttle plate 1 and the impact plate 3 are arranged at intervals, a throttle cavity 2 is formed between the throttle plate 1 and the impact plate 3, and a plurality of throttle holes 11 which are arranged in an array and used for allowing fluid to flow into the throttle cavity 2 are formed in the throttle plate 1. The throttle chamber 2 is provided with a plurality of first column ribs 6 at intervals in a row, and the throttle chamber 2 is divided into a plurality of mutually independent throttle subchambers 21 in a row along the turbine axial direction (the direction indicated by the arrow in fig. 4).
The impact plate 3 and the air film plate 5 are arranged at intervals, an impact cavity 4 is formed between the impact plate 3 and the air film plate 5, and a plurality of impact holes 31 which are arranged in an array and used for allowing fluid to flow into the impact cavity 4 from each throttling subchamber 21 are arranged on the impact plate 3. In particular, since the flow distribution in the axial direction of the turbine may be impaired in the impingement chamber, the present invention provides the same chamber division structure at the location of the impingement chamber corresponding to the throttle chamber to ensure the supply of flow and thus sufficient cooling. Specifically, a plurality of second column ribs 7 are arranged in the impact chamber 4 at intervals in a row, the impact chamber 4 is divided into a plurality of impact subchambers 41 which are independent of each other, a plurality of turbulence columns 42 are included in each impact subchamber 41, and the number of the throttling subchambers 21 and the number of the impact subchambers 41 are equal and correspond in a row. The air film plate 5 is provided with a plurality of air film holes 51 which are arranged in an array and used for fluid to flow out of each impact sub-cavity 41. The longitudinal direction of the first column rib 6 and the second column rib 7 is perpendicular to the turbine axial direction.
In particular, by making the number of the throttling subchambers 21 and the number of the impingement subchambers 41 equal and corresponding in rows, the whole air supply structure can be divided into a plurality of cooling units which are independent of each other, so that the problem of lean and rich before and after cooling of the outer ring block caused by a single impingement chamber is avoided, and a targeted cooling structure is formed.
In particular, in order to ensure sufficient cooling, the smaller the size of each cooling unit, the better the smaller the holes. Based on the existing conventional processing technology, all the hole patterns are set as circular holes, and the hole diameter is 0.4 mm.
In addition, in order to ensure effective cooling of the outer ring block, namely, reasonably distributing the limited cold air of the outer ring block, the invention designs the number of the throttling holes included in each cooling unit and the size of each throttling hole on the basis of dividing the outer ring block into a plurality of cooling units which are not interfered with each other along the axial direction of the turbine so as to ensure that the throttling holes have different throttling capacities along the axial direction of the turbine, thus realizing distribution of bleed air of an air system, and then impacting the outer ring block through a composite cooling structure of a turbulent air film. Advantageously, the number of the throttling holes of each cooling unit is set, so that the flow of the fluid entering each cooling unit is changed in a gradient manner along the axial direction of the turbine, the flow resistance of the whole sub-cavity is adjusted, the cold air flows through each sub-cavity and then has different cold air supplies, a certain flow gradient along the axial direction of the turbine is ensured, and sufficient cooling of each position of the outer ring block is ensured. Then according to the difference of the outlet pressure and the outlet temperature of the cooling units divided upwards along the turbine shaft, the total area of the air inlet throttle holes of the cooling units is designed to be adjusted to obtain different air inlet throttle areas, and the specific determination process is as follows:
1) calculating and obtaining pressure and temperature distribution of an interface of a turbine main flow and an outer ring block along the axial direction of the turbine;
2) dividing the gas supply structure into n cooling units according to the width of each cooling unit; dividing the interface of the main flow of the turbine and the outer ring block into n units, wherein the n units correspond to the positions of film hole outlets of the n cooling units one by one, and determining the average temperature and the average pressure under each unit based on the pressure and the temperature distribution of the interface of the main flow of the turbine and the outer ring block of the turbine along the axial direction of the turbine, which are obtained in the step 1), namely determining the average temperature and the average pressure at the outlet of each cooling unit;
3) analyzing the flow characteristics of the single cooling unit, acquiring the relation between the inlet-outlet pressure ratio and the flow at different outlet temperatures of the single cooling unit, and determining the inlet-outlet pressure boundary and the temperature boundary of each cooling unit based on the average temperature and the average pressure at the outlet of each cooling unit determined in the step 2);
4) determining the flow rate of cold air flowing through each cooling unit according to the allowable temperature requirement of the material, and determining the total area of the throttling hole at the inlet of each cooling unit according to the relationship between the inlet-outlet pressure ratio and the flow rate of the cold air flowing through the single cooling unit at different outlet temperatures acquired in the step 3).
The turbine main flow and the outer ring block interface are divided into n cells in the turbine axial direction, and the average pressure of the ith cell is p (i) (1 to n) and the average temperature is t (i) (1 to n). Through calculation, the flow characteristic of a single cooling unit under different temperatures and different inlet areas is m-f1(Tout，Pin/Pout，Ain)，ToutDenotes the outlet temperature, P, of the cooling unitinDenotes the inlet pressure, P, of the cooling unitoutDenotes the outlet pressure of the cooling unit, AinDenotes the inlet area of the cooling unit, f1Represents m and Tou、Pin/Pout、AinThe mapping relationship of (2); the maximum temperature of the outer ring block unit is Tmax＝f2(Tout，Pin/Pout，m)，f2Represents TmaxAnd Tout、Pin/PoutM, ToutIndicating coldThe outlet temperature of the cooling unit; the inlet area a of the cooling unitin＝g1(Tout，Pin/Pout，m)，Pin，TinKnown as g1Is represented by AinAnd Tout、Pin/PoutAnd m. When the required temperature of the material is known, the required flow rate is mneed＝g2(Tmax，Tout，Pin/Pout)，g2Represents mneedAnd Tmax、Tout、Pin/PoutThe mapping of (a) is that the inlet area of the ith cooling unit is Ain(i)＝g1(T(i)，Pin/P(i)，mneed)(i＝1～n)。
In the present embodiment, 14 first columnar ribs 6 are provided in the throttle chamber 2 at intervals in a row, and 14 second columnar ribs 7 are also provided in the surge chamber 4 at intervals in a row, and the throttle subchambers 21 and the surge subchambers 41 are formed at positions corresponding to each other and 15 in number.
As shown in fig. 4, the number of the throttle holes in the turbine axial direction in this embodiment is 7, 6, 5, 4, or 4, respectively. Under the design, the flow supplied by each cooling unit along the axial direction of the turbine is strictly divided according to the area ratio of the throttling hole, and the outer ring block is accurately cooled.
In particular, each impingement subchamber 41 comprises a plurality of cells arranged in series laterally. As shown in fig. 5 to 6, each cell includes a space formed by two second cylindrical ribs 7, and a turbulence column 42, 2 impingement holes 31, and 6 film holes 51, which are located in the space and have a herringbone shape in axial section. The turbulence column 42 includes a top 421 and two side wings 422 extending from the top 421 to the left and right sides and obliquely outward from top to bottom. The 6 air film holes 51 are uniformly distributed above the top 521 and close to one of the second ribs 7. Each flank 422 and two second ribs 7 form a reduction channel 423, and 2 impact holes 31 are arranged at the middle position under the two flanks 422, so that when cold air flowing into the element from the 2 impact holes 31 passes through the reduction channel 423, kinetic energy is continuously accelerated, a strong cross flow effect is ensured, the convective heat transfer coefficient is enhanced, and cooling is further enhanced.
It will be apparent to those skilled in the art that various modifications and improvements can be made to the embodiments of the present invention without departing from the inventive concept thereof, and these modifications and improvements are intended to be within the scope of the invention.
Claims (5)
1. The gas supply structure of the turbine outer ring block for supplying gas in a separated cavity is characterized by sequentially comprising a throttle plate, a throttle cavity, an impact plate, an impact cavity and a gas film plate in the flowing direction of a fluid; the throttling plate and the impact plate are arranged at intervals, the throttling cavity is formed between the throttling plate and the impact plate, and a plurality of throttling holes which are arranged in an array mode and used for allowing fluid to flow into the throttling cavity are formed in the throttling plate; a plurality of first column ribs are arranged in the throttling cavity at intervals in rows so as to divide the throttling cavity into a plurality of mutually independent throttling branch cavities; the impact plate and the air film plate are arranged at intervals, the impact cavity is formed between the impact plate and the air film plate, and a plurality of impact holes which are arranged in an array and are used for allowing fluid to flow into the impact cavity from each throttling sub-cavity are formed in the impact plate; a plurality of second column ribs are arranged in the impact cavity at intervals in a row so as to divide the impact cavity into a plurality of independent impact sub-cavities; the air film plate is provided with a plurality of air film holes which are arranged in an array and used for fluid to flow out of each impact sub-cavity; each impact sub-cavity comprises a plurality of turbulence columns; the throttling subchambers and the impact subchambers are equal in number and correspond in a row, so that the gas supply structure is divided into a plurality of cooling units which are independent of each other; the length direction of each first column rib and each second column rib is vertical to the axial direction of the turbine;
the total orifice area of each cooling unit is configured such that the flow rate of the fluid entering each cooling unit changes in a gradient manner in the turbine axial direction.
2. The air supply structure according to claim 1, wherein the total orifice area setting process of each cooling unit is as follows:
1) calculating and obtaining the pressure and temperature distribution of the interface of the turbine main flow and the outer ring block along the axial direction of the turbine;
2) dividing the gas supply structure into n cooling units according to the width of each cooling unit; dividing the interface of the main flow of the turbine and the outer ring block into n units, wherein the n units correspond to the positions of film hole outlets of the n cooling units one by one, and determining the average temperature and the average pressure under each unit based on the pressure and the temperature distribution of the interface of the main flow of the turbine and the outer ring block of the turbine along the axial direction of the turbine, which are obtained in the step 1), namely determining the average temperature and the average pressure at the outlet of each cooling unit;
3) analyzing the flow characteristics of the single cooling unit, acquiring the relation between the inlet-outlet pressure ratio and the flow at different outlet temperatures of the single cooling unit, and determining the inlet-outlet pressure boundary and the temperature boundary of each cooling unit based on the average temperature and the average pressure at the outlet of each cooling unit determined in the step 2);
4) determining the flow rate of cold air flowing through each cooling unit according to the allowable material temperature requirement, and determining the total area of the throttling holes at the inlets of the cooling units according to the relationship between the inlet-outlet pressure ratio and the flow rate of the cold air flowing through the single cooling unit at different outlet temperatures acquired in the step 3).
3. The air supply structure according to claim 2, wherein the total orifice area of each cooling unit is set to gradually decrease in the turbine axial direction.
4. The air supply structure according to any one of claims 1 to 3, wherein the orifice, the impingement holes and the film holes are circular holes having a diameter set according to a cold air demand of the cooling unit.
5. Air supply structure according to any one of claims 1 to 3, characterised in that the turbulence column is a cylindrical or chevron-shaped turbulence column.
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| CN202110323611.XA CN113123833B (en) | 2021-03-26 | 2021-03-26 | Turbine outer ring block air supply structure with separated air supply |
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| CN202110323611.XA CN113123833B (en) | 2021-03-26 | 2021-03-26 | Turbine outer ring block air supply structure with separated air supply |
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| CN114109514B (en) * | 2021-11-12 | 2023-11-28 | 中国航发沈阳发动机研究所 | Turbine blade pressure surface cooling structure |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107435563A (en) * | 2017-05-05 | 2017-12-05 | 西北工业大学 | A kind of case structure with tip clearance control and the flowing control of leaf top |
| CN107559090A (en) * | 2016-06-30 | 2018-01-09 | 中国航发商用航空发动机有限责任公司 | The cooling component of turbine outer ring |
| CN207813753U (en) * | 2017-12-26 | 2018-09-04 | 中国航发商用航空发动机有限责任公司 | Turbine outer ring cooling structure, turbine structure and aero-engine |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| WO2007099895A1 (en) * | 2006-03-02 | 2007-09-07 | Ihi Corporation | Impingement cooling structure |
| US8727704B2 (en) * | 2010-09-07 | 2014-05-20 | Siemens Energy, Inc. | Ring segment with serpentine cooling passages |
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| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN107559090A (en) * | 2016-06-30 | 2018-01-09 | 中国航发商用航空发动机有限责任公司 | The cooling component of turbine outer ring |
| CN107435563A (en) * | 2017-05-05 | 2017-12-05 | 西北工业大学 | A kind of case structure with tip clearance control and the flowing control of leaf top |
| CN207813753U (en) * | 2017-12-26 | 2018-09-04 | 中国航发商用航空发动机有限责任公司 | Turbine outer ring cooling structure, turbine structure and aero-engine |
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