Cooling structure and pre-cooling method for turbine blade of aero-engine

Abstract

The invention relates to the technical field of aero-engine turbine blade cooling, and discloses an aero-engine turbine blade cooling structure and a pre-cooling method.The aero-engine turbine blade cooling structure comprises an inner wall and an outer wall which are of a double-layer structure, the inner wall is arranged in the outer wall, the inner wall comprises a wall face, and a plurality of impact holes are formed in the wall face in parallel in a penetrating mode; a plurality of cooling channels are arranged in the wall surface in parallel; the cooling channels are arranged among the impact holes; each cooling channel comprises an inlet end and an outlet end; the inlet end is connected with a fuel oil system and is used for feeding fuel oil into the cooling channel; the outlet end is connected with a fuel oil system and used for discharging fuel oil in the channel to the fuel oil system. The diameter of the cooling channel is within 1mm; and a plurality of air film holes are formed in the outer wall. Cooling gas in the turbine is pre-cooled through aviation fuel, so that the cooling effect of the turbine is further improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft engine cooling, and in particular to an aircraft engine turbine blade cooling structure and a pre-cooling method. Background Art

[0002] As we all know, the aircraft engine is the core of the entire aircraft, and its performance directly determines the overall performance of the aircraft. In the aircraft gas turbine engine, the turbine blades have the worst working environment and the most complex stress. At the same time, the turbine blades are also the key to the aircraft engine's high performance under the condition of small size and light weight.

[0003] At the same time, with the increasingly tight global energy supply and the urgent need for environmental protection, improving the performance of aircraft engines to achieve energy conservation and emission reduction has become a top priority. As the core of aircraft engines, the performance improvement of turbine blades has become the key. Due to the high temperature, high pressure, high speed and high load working environment, it is well known that the turbine inlet gas temperature has a direct impact on the overall performance of the turbine. At present, the turbine inlet temperature of advanced aircraft engines has reached about 1850-1950K, while the melting point of the metal material of high-pressure turbine blades can only reach about 1423K. Excessive temperature will cause thermal deformation, thermal fatigue, thermal corrosion and other problems of turbine blades, which will seriously affect their safety and stability.

[0004] Therefore, the improvement of cooling technology for turbine blades becomes the key. Existing cooling technologies are roughly divided into internal cooling technology and external cooling technology. The typical turbine blade cooling method currently used is a composite cooling structure, that is, external impact cooling and film cooling technologies are used in combination with internal cooling units for simultaneous cooling. However, because the non-uniformly distributed wall temperature will aggravate the thermal stress concentration of the blades, the cooling effect is low, and the complex cooling structure will increase the flow loss of the internal cooling gas. At the same time, due to the small space of the turbine blades, high-performance internal cooling units are too complex and difficult to implement, and the advanced cooling structure has a long design cycle, high cost, and great processing difficulty, which makes it difficult to further improve the cooling technology of turbine blades.

[0005] In this regard, this program proposes an aircraft engine turbine blade cooling structure and pre-cooling method, which effectively improves the cooling effect of the turbine blade, improves the overall performance of the turbine blade, and reduces the processing difficulty. Summary of the invention

[0006] The present invention aims to provide an aircraft engine turbine blade cooling structure and a pre-cooling method to improve the cooling effect of the turbine blade, improve the overall performance of the turbine blade, and reduce the difficulty of processing.

[0007] To achieve the above-mentioned purpose, the present invention adopts the following technical scheme: an aircraft engine turbine blade cooling structure, applied to the aircraft engine turbine blade, including an inner wall and an outer wall in a double-layer structure, wherein the inner wall is arranged inside the outer wall, and the inner wall includes a wall surface, and a plurality of impact holes are arranged in parallel through the wall surface, and a plurality of cooling channels are arranged in parallel inside the wall surface; the cooling channels are arranged between the impact holes; each of the cooling channels includes an inlet end and an outlet end; the inlet end is connected to the fuel system for delivering the fuel into the cooling channel; the outlet end is connected to the fuel system for discharging the fuel in the channel to the fuel system; the diameter of the cooling channel is within 1 mm; and a plurality of air film holes are arranged on the outer wall.

[0008] The principles and advantages of this solution are:

[0009] In the traditional composite cooling technology of turbine blades, the research is mainly divided into two directions: one is the mechanism research under the influence of gas-heat parameters, and the other is the innovative structural design to improve the cooling performance. The size and design parameters of the composite cooling structure have an important influence on the performance and weight of the blade. The size of the high-pressure turbine blade of an aircraft engine is usually in the millimeter level, with a general blade height of about 85mm and a chord length of about 55mm. The interior adopts a double-wall composite cooling structure design, in which the inner wall thickness is about 0.3-1.0mm, the outer wall thickness is about 0.7-1.2mm, and the aperture of the internal cooling impact hole is about 0.5-2.0mm. The innovative design of the composite cooling structure is limited by the narrow space, and it is very difficult to further improve the performance of the composite cooling structure. Therefore, this scheme proposes a pre-cooling cooling structure, which improves the cooling effect by pre-cooling without upgrading the composite cooling structure, thereby getting rid of the problem of limited internal space.

[0010] However, the thickness of the inner wall is relatively small. In order to design the cooling channel within such a size range and ensure the cooling effect of the cooling channel, it is necessary to ensure the circulation of the cooling channel and the coverage of the cooling area. The diameter that the cooling channel can reach is limited by the thickness of the inner wall. Therefore, while ensuring the diameter of the cooling channel, it is necessary to ensure that the coolant flows smoothly in the cooling channel without causing blockage and affecting the cooling effect. Therefore, this solution uses fuel as the coolant and controls the temperature of the incoming fuel at 300-400K to avoid the coking of aviation fuel and blockage of the channel, ensuring the normal flow of the coolant. BRIEF DESCRIPTION OF THE DRAWINGS

[0011] Figure 1 This is a schematic structural diagram of an aircraft engine turbine blade cooling structure in Embodiment 1 of the present invention;

[0012] Figure 2 This is a schematic diagram of the model structure of the cooling structure in the first embodiment of the present invention;

[0013] Figure 3 for Figure 2 Schematic diagram of the transverse cross-sectional structure;

[0014] Figure 4 This is a schematic flow chart of a method for precooling a turbine blade of an aircraft engine in Embodiment 1 of the present invention;

[0015] Figure 5 This is a schematic diagram of the model structure of the cooling structure in the second embodiment of the present invention;

[0016] Figure 6 This is a schematic diagram of the model structure of the cooling structure in Example 3 of the present invention. DETAILED DESCRIPTION

[0017] The following is further described in detail through specific implementation methods:

[0018] The reference numerals in the drawings of the specification include: turbine blade 1 , inner wall 2 , outer wall 3 , impact hole 4 , cooling channel 5 , inlet end 51 , outlet end 52 , U-shaped end 53 , air film hole 6 , and spoiler column 7 .

[0019] Embodiment 1

[0020] like Figure 1 As shown, an aircraft engine turbine blade cooling structure can be applied to the turbine blades or combustion chambers of aircraft engines. When cooling the turbine blades, the impact cooling gas is pre-cooled first, thereby reducing the impact cooling gas temperature and improving the cooling effect, thereby overcoming the limitation of only being able to innovate and improve the composite cooling structure, and effectively improving the overall performance of the aircraft engine. Specifically, it includes an inner wall 2 and an outer wall 3 in a double-layer structure. The inner wall 2 is installed on the inner side of the outer wall 3, and the whole structure is a teardrop shape. For the convenience of description, in this embodiment, the inner wall 2 and the outer wall 3 are used as shown in the attached Figure 2 The square model shown is represented.

[0021] Specifically, as attached Figure 2 As shown, a plurality of impact holes 4 are arranged side by side on the inner wall 2, the spacing between each row of impact holes 4 is 3-5mm, and the diameter of each impact hole 4 is 2mm. A plurality of air film holes 6 are arranged side by side on the outer wall 3, the spacing between each row of air film holes 6 is 3mm, and the diameter of the air film holes 6 is 3mm. In this embodiment, the air film holes 6 are through holes with a certain inclination angle to the outer wall surface. Figure 3As shown, the cooling gas rushing into the inner wall 2 from below the turbine blades is impacted toward the outer wall 3 through the impact hole 4, and the impact cooling gas forms a cooling film on the outer wall 3 through the air film hole 6, thereby isolating the high-temperature gas to achieve a cooling effect. At the same time, a spoiler column 7 is also provided between the outer wall 3 and the inner wall 2 to expand the flow range and flexibility of the cooling gas, reduce the blockage and accumulation of the cooling gas, and ensure the smooth flow of the cooling gas.

[0022] Among them, as attached Figure 2 and attached Figure 3 As shown, a plurality of cooling channels 5 are vertically integrated in the wall surface of the inner wall 2, and each cooling channel 5 is arranged between each row of impact holes 4 to ensure that the coolant in the cooling channel 5 can fully cover the wall surface of the inner wall 2 and pre-cool the entire inner wall 2 area. In this embodiment, the distance between each cooling channel 5 is 5mm to achieve a cooling effect. In this embodiment, each cooling channel 5 is set as a U-shaped channel to expand the coverage area of ​​the coolant and improve the pre-cooling effect. As shown in the attached Figure 2 As shown, the ends of the U-shaped channel are an inlet end 51 and an outlet end 52, wherein the inlet end 51 is connected to the fuel system, so that the aviation kerosene (fuel) enters the cooling channel 5 from the inlet end 51, and flows out from the outlet end 52 through the U-shaped channel and enters the fuel system. In this embodiment, the inlet end 51 and the outlet end 52 are arranged at the bottom of the inner wall 2, and the U-shaped end 53 is arranged at the top of the inner wall 2, and is 1-3mm higher than the impact hole at the highest position, so as to ensure that the coolant in the U-shaped channel can pass through the entire inner wall 2, ensure the coverage and effect of pre-cooling, and prevent uneven temperature differences from affecting the cooling effect.

[0023] In this embodiment, according to the specific structure and thickness of the inner wall 2, the diameter of the cooling channel 5 is set to 1 mm, the distance between the inlet end 51 and the outlet end 52 of the U-shaped channel is 3 mm, and the curvature of the elbow is 180° to prevent the coolant from accumulating at the U-shaped end 53, causing poor circulation and affecting the pre-cooling efficiency of the coolant.

[0024] In this solution, the composite cooling structure is no longer improved and upgraded, and the cooling effect of the composite cooling structure is not limited by the narrow space of the composite cooling structure. Instead, the cooling structure is designed in a pre-cooling manner. By setting a cooling channel 5 in the inner wall 2, the fuel is sent into the wall surface of the inner wall 2 through the cooling channel 5, and the impact cooling gas is pre-cooled to improve the cooling effect. At the same time, according to the specific structure of the inner wall 2 of the composite cooling structure, the number and shape of the cooling channels can be flexibly set to improve the efficiency of pre-cooling.

[0025] In order to ensure the normal flow of the coolant and achieve the pre-cooling effect in the micro-scale cooling pipe 5, the structural design and size design of the cooling pipe 5 are also very important, so as to ensure comprehensive pre-cooling coverage and uniform temperature, and effectively improve the cooling effect.

[0026] Therefore, the present invention also provides a method for precooling turbine blades of an aircraft engine, which is applicable to the above-mentioned cooling structure, such as the attached Figure 4 As shown, when the engine is working, aviation fuel is used to continuously cool the cooling gas from the high pressure compressor. The temperature of the aviation fuel entering the cooling channel is maintained at 300-400K, and then the aviation kerosene is uniformly fed from the inlet end 51 into the cooling channel 5 at a certain flow rate to continuously pre-cool the impact cooling gas.

[0027] In this embodiment, the fuel needs to be pretreated so that the temperature of the fuel in the cooling pipe 5 is continuously maintained at 300-400K, thereby effectively avoiding the formation of coking in the micro-scale cooling pipe 5, avoiding the blockage of the fuel in the cooling pipe 5, ensuring that the aviation fuel can flow smoothly, and achieving the effect of pre-cooling. At the same time, using fuel as a coolant will not increase the burden of excessive coolant space occupation, reduce aviation load, and can make more efficient use of fuel, further improving the cooling effect of turbine blades. At the same time, using it as a coolant can also ensure the smooth flow of coolant and reduce blockage and other problems, which are difficult to meet the application and implementation in micro-scale structures. During pre-cooling, the fuel is input at a uniform speed to ensure that the temperature of each cooling channel 5 is consistent, effectively avoiding uneven temperature from aggravating the thermal stress concentration of the blades and improving the cooling effect. By pre-cooling, the temperature of the impact cooling gas can be effectively reduced, and the cooling efficiency can be improved.

[0028] Embodiment 2

[0029] Different from the first embodiment, in this embodiment, the cooling channel 5 is different from the structure of the inner wall 2. In this embodiment, the size of the inner wall is 1 mm, and the diameter of the impact hole is less than 1 mm. Therefore, the cooling channel 5 is designed to be linear. Figure 5 As shown, the linear channel is vertically arranged in the wall surface of the inner wall 2, and its upper end is the outlet end 52, which is arranged at the top of the wall surface, connected to the outside world and connected to the fuel system, and the lower end of the channel is the inlet end 51, which is arranged at the bottom of the wall surface, connected to the outside world at the lower end, and connected to the fuel system. In this embodiment, the distance between each cooling channel 5 is 3-5mm. According to the structure of the inner wall 2, the diameter of the linear channel can be set to 1mm to ensure the maximum flow rate of the coolant and improve the pre-cooling effect.

[0030] In this embodiment, the linear channel can be applied to a typical double-wall turbine blade cooling structure, which has a simple structure and low processing difficulty, and can effectively reduce the production cycle and reduce production costs.

[0031] Embodiment 3

[0032] Different from the first embodiment, in this embodiment, the cooling channel 5 is different from the structure of the inner wall 2. In this embodiment, the size of the inner wall is 1 mm, and the diameter of the impact hole is 1 mm. Therefore, the cooling channel 5 is designed to be a serpentine channel. Figure 6 As shown, the serpentine channels are vertically spaced within the wall surface of the inner wall 2 and are located between the rows of impact holes 4, with the upper end being the outlet end 52, which is disposed at the top of the wall surface, connected to the outside world and connected to the fuel system; the lower end of the channel is the inlet end 51, which is disposed at the bottom of the wall surface, connected to the outside world at the lower end, and connected to the fuel system, so that the fuel is circulated and reused without waste.

[0033] In this embodiment, the distance between each cooling channel 5 is 3-5mm. According to the structure of the inner wall 2, the diameter of the serpentine channel can be set to 1mm to ensure the maximum flow rate of the coolant and improve the pre-cooling effect. At the same time, in order to avoid blockage at the bends of the serpentine channel, each bend of the serpentine channel is a right-angle bend, and the width of the bend is 2mm and the horizontal length is 2mm, so as to avoid the bends affecting the flow rate of the coolant and causing uneven distribution of the coolant, so as to avoid uneven temperature and affect the pre-cooling effect.

[0034] In this embodiment, the serpentine channel can be applied to various cooling structures such as typical double-wall turbine blades, effectively increasing the effective area of ​​pre-cooling, improving the pre-cooling coverage, further reducing the temperature of the impact cooling gas, and improving the cooling performance.

[0035] The above are only embodiments of the present invention, and the common knowledge such as the known specific technical solutions and / or characteristics in the solution are not described in detail here. It should be pointed out that for those skilled in the art, several modifications and improvements can be made without departing from the technical solution of the present invention, which should also be regarded as the protection scope of the present invention, and these will not affect the effect of the implementation of the present invention and the practicality of the patent. The scope of protection required by this application shall be based on the content of its claims, and the specific implementation methods and other records in the specification can be used to interpret the content of the claims.

Claims

1. An aircraft engine turbine blade cooling structure, characterized in that: It comprises an inner wall and an outer wall in a double-layer structure, wherein the inner wall is arranged inside the outer wall, the inner wall comprises a wall surface, a plurality of impact holes are arranged in parallel through the wall surface, and a plurality of cooling channels are arranged in parallel inside the wall surface; the cooling channels are arranged between the impact holes; each of the cooling channels comprises an inlet end and an outlet end; the inlet end is connected to the fuel system for delivering the fuel into the cooling channel; the outlet end is connected to the fuel system for discharging the fuel in the channel to the fuel system; the diameter of the cooling channel is within 1 mm; a plurality of air film holes are arranged on the outer wall.

2. The cooling structure of an aircraft engine turbine blade according to claim 1, characterized in that: The cooling channels are arranged at intervals between the rows of impact holes, and the distance between two adjacent cooling channels is 3-5 mm.

3. The cooling structure of an aircraft engine turbine blade according to claim 2, characterized in that: Each cooling channel is a U-shaped channel, which is vertically arranged in the wall surface of the inner wall; the inlet end and the outlet end are both arranged at the bottom of the wall surface.

4. The cooling structure of an aircraft engine turbine blade according to claim 3, characterized in that: The distance between the two channels of the U-shaped channel is 3-5 mm; the diameter of the U-shaped channel is less than 1 mm.

5. The cooling structure of an aircraft engine turbine blade according to claim 2, characterized in that: Each cooling channel is a linear channel, which is vertically arranged in the wall surface of the inner wall, the inlet end is arranged at the bottom end of the wall surface, and the outlet end is arranged at the top end of the wall surface.

6. The cooling structure of an aircraft engine turbine blade according to claim 2, characterized in that: Each cooling channel is a serpentine channel, which is vertically arranged in the wall surface of the inner wall, the inlet end is arranged at the bottom end of the wall surface, and the outlet end is arranged at the top end of the wall surface; the diameter of the serpentine channel is less than 1mm.

7. A method for precooling an aircraft engine turbine blade, applicable to an aircraft engine turbine blade cooling structure as claimed in any one of claims 1 to 6, characterized in that: When the engine is running, aviation fuel is used to continuously cool the cooling gas from the high-pressure compressor.

8. The method for precooling a turbine blade of an aero-engine according to claim 7, characterized in that: The temperature of aviation fuel entering the cooling channel is 300-400K.

9. The method for precooling a turbine blade of an aero-engine according to claim 8, characterized in that: It also involves pre-treatment of aviation fuel to a temperature of 300-400K.

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