Aero-engine turbine blade pulse jet flow structure and method

Abstract

The invention relates to the technical field of aero-engine turbine blade cooling, and discloses an aero-engine turbine blade pulse jet flow structure and method.The aero-engine turbine blade pulse jet flow structure comprises an inner layer wall and an outer layer wall, and a plurality of impact holes are formed in the inner layer wall; the impact hole comprises a nozzle structure and a through hole structure, and the nozzle structure is arranged on the front portion of the through hole structure and faces the outer layer wall. The through hole structure comprises a pulse cavity and an air inlet channel; the air inlet channel is of a conical structure and comprises an inlet end and an outlet end, the diameter of the inlet end is larger than that of the outlet end, the outlet end is arranged at one end of the pulse cavity, and the diameter of the inlet end is set according to pulse frequency. By controlling the size of the air inlet channel in the impact hole and the size of the pulse cavity, the air injection frequency is controlled, the self-oscillation pulse effect is achieved, heat can be rapidly taken away by injected air in time, the problem that the air is accumulated in the interlayer can be effectively solved, the cooling efficiency is improved, and the performance of the turbine blade is integrally improved.

Description

Technical Field

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

[0002] As the heart of an aircraft, aircraft engines are directly related to the safety, reliability and economy of an aircraft, and are an important indicator of the advancement of an aircraft. In the current context of increasingly tight global energy supply and increasingly stringent environmental protection, it is imperative to improve the performance of aircraft engines to achieve energy conservation and emission reduction. As the core of aircraft engines, turbine blades face increasingly urgent demands for technological upgrades and breakthroughs due to high temperature, high pressure, high speed and high load. The turbine inlet gas temperature has a direct impact on the overall performance of the turbine. Excessive temperature can cause thermal deformation, thermal fatigue, thermal corrosion and other problems in the turbine blades, which in turn seriously affects their safety and stability.

[0003] Therefore, improving the cooling effect of turbine blades has become the key. The typical turbine blade cooling method currently used is a composite cooling structure, that is, external impingement cooling and air film cooling technologies are used in combination with internal cooling units for simultaneous cooling. However, since the existing cooling structure uses a double-wall structure, by introducing impingement cooling gas into the cooling structure, the gas is forced from the impingement holes of the inner wall to the outer wall, and the heat is taken away by the gas flow to achieve the cooling effect.

[0004] However, due to the narrow space of the cooling structure, the gap between the inner wall and the outer wall is very small, and the gas rushing out from the impact hole will diffuse around the hole due to the airflow. Under the action of multiple impact holes, when the cooling gas continues to rush out, it is easy to cause gas accumulation between the two wall surfaces, disrupting the gas flow and forming a blockage, making it difficult for the cooling gas to be quickly discharged to take away the heat, thereby affecting the cooling effect.

[0005] In this regard, this program proposes a pulse jet structure and method for aircraft engine turbine blades to avoid the accumulation of cooling gas in the cooling structure, effectively improve the cooling effect of the turbine blades, and improve the overall performance of the turbine blades. Summary of the invention

[0006] The present invention aims to provide a pulse jet structure and method for aircraft engine turbine blades to avoid accumulation of impact cooling gas and blockage of airflow, thereby improving the fluidity of the impact cooling gas, thereby enhancing the cooling effect of the turbine blades and improving the overall performance of the turbine blades.

[0007] To achieve the above-mentioned purpose, the present invention adopts the following technical scheme: a pulse jet structure of an aircraft engine turbine blade, applied to the turbine blade of an aircraft engine, comprising an inner wall and an outer wall, wherein a plurality of impact holes are arranged on the inner wall; the impact hole comprises a nozzle structure and a through-hole structure which are interconnected, wherein the nozzle structure is arranged at the front of the through-hole structure and faces the direction of the outer wall; the through-hole structure comprises a pulse chamber and an air inlet channel which are interconnected; the pulse chamber is connected to the nozzle structure and is arranged at the front end of the air inlet channel; the air inlet channel is a conical structure, comprising an inlet end and an outlet end, wherein the diameter of the inlet end is larger than the diameter of the outlet end, wherein the outlet end is arranged at one end of the pulse chamber, the diameter of the inlet end is set according to the pulse frequency, and the length of the pulse chamber is less than 0.5 mm.

[0008] At the same time, the present invention also provides an aero-engine turbine blade pulse jet method, which is applied to the above-mentioned aero-engine turbine blade pulse jet structure, and specifically includes the following steps:

[0009] Step 1: setting the diameter of the inlet end of the air inlet channel and the size of the pulse cavity according to the required pulse frequency;

[0010] Step 2, the impact cooling gas is introduced into the pulse chamber from the inlet end at a set flow rate, so that the impact cooling gas is continuously gathered in the pulse chamber;

[0011] Step 3, when the pulse cavity is filled with the impact cooling gas, the gas in the pulse cavity is ejected from the nozzle structure by the impact cooling gas that subsequently enters;

[0012] Step 4: Repeat this process over and over to form a pulse cooling gas with a certain frequency.

[0013] The principles and advantages of this solution are:

[0014] The size of the high-pressure turbine blades of aircraft engines is usually in the millimeter range, with a blade height of about 85mm and a chord length of about 55mm. The internal structure 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 size and design parameters of the composite cooling structure have an important impact on the performance and weight of the blade. Due to the limited design space of the cooling structure, the interlayer gap between the inner wall and the outer wall is very small. The gas rushing out of the inner wall impact hole can easily spread around in the interlayer after hitting the outer wall, thereby blocking the surroundings of other impact holes, causing blockage of the gas in the interlayer and affecting the flow of gas.

[0015] To solve this problem, most of them will adopt the design of gas diversion and other structures in the relatively large interlayer, so that the gas can flow away quickly without affecting the speed of the impact cooling gas, ensuring that the fast-flowing gas can take away the heat in time to achieve the cooling effect. However, due to the limitation of the interlayer space, its structure will always be affected. At the same time, because the size of the impact hole is smaller, in order to ensure the rapid impact of the gas, the structure of the impact hole is more difficult to design, and the complex structure is more likely to cause the impact hole to be blocked.

[0016] However, the present solution does not adopt the method of rapid gas impact for cooling, but reduces the impact frequency of the gas so that it can achieve the effect of impact cooling, thereby slowing down the phenomenon of gas blockage caused by excessive speed. Secondly, the present application proposes a method of changing the internal structure of the impact hole in the impact hole so that the impact hole can achieve self-excited oscillation, thereby achieving the effect of gas pulsing at a certain frequency without increasing the internal structure of the impact hole, thereby avoiding the blockage of the impact hole. By controlling the size of the air inlet channel and the size of the pulse cavity in the impact hole, the time of gas collection is controlled, and then the frequency of the jet is controlled to achieve the effect of self-excited oscillation pulse, that is, the ejected gas can take away the heat quickly and promptly, and continuously spray to the outer wall, and can effectively solve the problem of gas accumulation in the interlayer, improve the cooling efficiency, and improve the performance of the turbine blades as a whole. BRIEF DESCRIPTION OF THE DRAWINGS

[0017] Figure 1 This is a schematic structural diagram of an aircraft engine turbine blade in Embodiment 1 of the present invention;

[0018] Figure 2 This is a schematic structural diagram of a pulse jet structure of an aircraft engine turbine blade in Embodiment 1 of the present invention;

[0019] Figure 3 Schematic diagram of the cross-sectional structure of the impact hole in the first embodiment of the present invention;

[0020] Figure 4 Schematic diagram of the process of the pulse jet method in Example 1 of the present invention. DETAILED DESCRIPTION

[0021] The following is further described in detail through specific implementation methods:

[0022] The figure marks in the drawings of the specification include: turbine blade 1, inner wall 2, outer wall 3, impact hole 4, nozzle structure 41, through hole structure 42, pulse chamber 401, air inlet channel 402, inlet end 501, outlet end 502, air film hole 6.

[0023] Embodiment 1

[0024] like Figure 1As shown, a pulse jet structure of an aircraft engine turbine blade can be applied to the turbine blade 1 or combustion chamber of an aircraft engine. The impact cooling gas entering the cooling structure is rushed from the impact hole to the outer wall in the form of pulses at a set frequency, thereby avoiding gas accumulation between the interlayers and causing gas blockage, thereby improving the cooling effect.

[0025] Specifically, the pulse jet structure includes an inner wall 2 and an outer wall 3. The inner wall 2 is sleeved in the cavity formed by the outer wall 3, and a sandwich with a gap of 3-7 mm is formed between the inner wall 2 and the outer wall 3. A plurality of impact holes 4 are arranged side by side through the inner wall 2, and the spacing between each row of impact holes 4 is 3-5 mm, and the diameter of each impact hole 4 is within 2 mm. A plurality of air film holes 6 are arranged at intervals through the outer wall 3, and the diameter of each air film hole 6 is 3-5 mm. The impact cooling gas reaching the outer wall 3 forms a layer of cooling air film through the air film holes 6 to isolate the temperature and achieve a cooling effect.

[0026] Specifically, as attached Figure 2 As shown, the impact hole 4 is a pulse jet structure. The cooling gas entering from the bottom of the turbine blade enters from the impact hole 4 and reaches the outer wall 3 after passing through the interlayer to achieve a cooling effect on the turbine blade. Figure 3 As shown, the impact hole 4 includes a nozzle structure 41 and a through hole structure 42, the nozzle structure 41 is connected to the through hole structure 42, and the nozzle structure 41 is arranged at the front of the impact hole 4, so that the nozzle of the nozzle structure 41 faces the outer wall 3. In this embodiment, the nozzle structure 41 is a conical cylinder structure channel with a length of 0.2-0.4 mm and a diameter within 1 mm, and the nozzle end is connected to the formed interlayer.

[0027] The rear end of the nozzle structure 41 is connected to the through-hole structure 42 . In this embodiment, the through-hole structure 42 includes a pulse chamber 401 and an air inlet channel 402 that are integrally formed and connected to each other. The pulse chamber 401 is disposed between the nozzle structure 41 and the air inlet channel 402 .

[0028] In this embodiment, the pulse chamber 401 includes a nozzle portion at the front end, an airflow chamber at the middle section, and an air inlet end at the rear end. The nozzle structure 41 is disposed in the nozzle portion to communicate with the airflow chamber. The through hole structure 42 is disposed in the air inlet end to communicate with the airflow chamber.

[0029] Specifically, the pulse chamber 401 is a cylindrical cavity, and its length is less than 0.5 mm. The length and size of the pulse chamber can be designed according to the amount of gas to be accommodated, so as to control the time of filling the gas in the pulse chamber 401, and then control the frequency of gas ejection, so as to form a pulse jet gas. In the prior art, in order to improve the cooling effect, it is considered that only when the flow of cooling gas is fast enough can the heat be taken away more quickly to achieve the cooling effect, but it is limited by the interlayer space between the inner wall 2 and the outer wall 3. In order to deal with the problem of easy accumulation of gas, most of them adopt diversion, design of air flow grooves, etc. to make the gas rushing out flow quickly along the air flow groove, and reduce the accumulation phenomenon, but the design of the diversion structure is always limited, and the more the structure is added, the more it will affect the flow of gas. As a result, the technology is limited. However, this scheme does not think from the perspective of diversion, but reversely controls the impact speed of the impact cooling gas, gathers the gas through the structure of the pulse chamber 401, so that the frequency of the gas rushing out just reaches the required cooling effect, and reduces the amount of gas rushing out to avoid the phenomenon of gas accumulation, effectively solving the blockage problem.

[0030] An air inlet channel 402 is provided at the air inlet end of the pulse chamber 401, and the impact cooling gas is sent into the air flow chamber at a certain flow rate through the air inlet channel 402. In this embodiment, the air inlet channel 402 is a conical structure and is located on the same center line as the nozzle structure 41. So that when the gas in the pulse chamber 401 is filled, the gas entering through the air inlet channel 402 can directly eject the gas in the pulse chamber 401 from the nozzle structure 41 to form a jet gas. The impact cooling gas can be gathered and compressed in the pulse chamber 401, effectively controlling the gas to be ejected outward all the time, while preventing the gas from continuously rushing out from the impact hole 4. The air inlet channel 402 can control the speed of the gas entering the air flow chamber, thereby controlling the time for the air flow chamber to be filled, so as to achieve the effect of controlling the pulse frequency. In this embodiment, the air inlet channel 402 includes an inlet end 501 and an outlet end 502, the inlet end 501 is located on the inner side of the inner wall 2, and the outlet end 502 is located in the air inlet end of the pulse chamber 401. The diameter of the inlet end 501 is larger than the diameter of the outlet end 502. In this embodiment, the diameter of the inlet end 501 is selected and set according to the flow rate of the cooling gas, and can be specifically set to 3-5 mm. The diameter of the outlet end 502 is smaller than the diameter of the inlet end 501, and the outlet end 502 can be set to 2-4 mm according to the diameter of the inlet end 502, so that the impact cooling gas can enter the pulse cavity 401 at a certain rate to form a stable pulse frequency.

[0031] At the same time, the present invention also provides a pulse jet method for a turbine blade of an aero-engine, which is applicable to the above-mentioned pulse jet structure, such as the attached Figure 4 As shown, the specific steps include:

[0032] S1, according to the required pulse frequency, select the designed diameter of the air inlet channel 402 and the size of the pulse chamber 401.

[0033] In this embodiment, according to the size of the cooling structure, the required cooling efficiency, and the flow speed of the cooling gas, the diameter of the inlet end 501 of the air inlet channel 402 can be set to 3 mm, the diameter of the outlet end 502 can be set to 2 mm, and the length can be set to 0.6 mm. The size of the cooling structure includes the size of the turbine blade, the thickness of the inner wall, the thickness of the outer wall, and the thickness of the middle layer, so as to evaluate the flow of the cooling gas.

[0034] S2, sending the impact cooling gas into the pulse chamber 401 from the inlet end 501 of the impact hole 4 at a uniform speed, so that the cooling gas is continuously gathered in the pulse chamber 401.

[0035] S3, when the cooling gas in the pulse chamber 401 fills the entire chamber, the gas rushing in again from the inlet end 501 will push the gas in the pulse chamber 401 to be ejected from the nozzle structure 41, thereby forming a pulse gas sprayed onto the outer wall 3 for cooling.

[0036] S4, repeats over and over again, forming pulse cooling gas with a certain interval frequency. Reduce the accumulation of gas in the interlayer, and avoid the accumulation of gas ejected from the impact hole 4. According to the cooling effect, the pulse gas is set to be ejected at a certain frequency through the buffer of the pulse cavity 401, which can ensure that the gas will not be interrupted in the interlayer, thereby achieving the effect of continuously taking away heat, and prevent the gas from squeezing each other and affecting the flow of the cooling gas behind.

[0037] At the same time, in this embodiment, the diameter of the nozzle structure can be set to control the gas volume and speed of the ejected gas to meet different cooling requirements, improve the cooling efficiency, reduce the risk of gas blockage, and ensure the normal flow of cooling gas.

[0038] In this embodiment, we will not consider how to improve the fluidity of the gas that is ejected, but will control the gas ejection time to slow down the frequency of gas ejection so that the gas is ejected in the form of interval pulses, solve the problem of gas accumulation, improve the fluidity of the gas, effectively take away the heat, and improve the cooling effect. However, since the cooling structure itself is relatively small, the impact hole 4 on the inner wall 2 has an even smaller structural space. Due to the limited operating space, we will not consider making structural improvements inside the impact hole 4, which is more likely to cause blockage in the hole. Therefore, structural improvements are generally made outside the impact hole 4, but the structural space of the interlayer is not large, which greatly limits the overall improvement.

[0039] This solution does not increase the internal structure of the impact hole 4, thereby avoiding the problem of blockage in the hole due to small space and complex structure. Instead, it changes the structure of the impact hole 4 itself to form a structure similar to self-excited oscillation, so that the gas can gather in the pulse cavity 401 of the impact hole 4 to form a pulse gas with a certain frequency, thereby ensuring the smoothness of the gas and solving the problem of easy accumulation and blockage of the gas, greatly improving the cooling effect of the impact cooling gas, and effectively improving the overall performance of the turbine blades.

[0040] 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 pulse jet structure, characterized in that: The invention comprises an inner wall and an outer wall, wherein a plurality of impact holes are arranged on the inner wall; the impact hole comprises a nozzle structure and a through hole structure which are interconnected, wherein the nozzle structure is arranged at the front of the through hole structure and faces the outer wall; the through hole structure comprises a pulse chamber and an air inlet passage which are interconnected; the pulse chamber is connected with the nozzle structure and is arranged at the front end of the air inlet passage; The air inlet passage is a conical structure, including an inlet end and an outlet end. The diameter of the inlet end is larger than the diameter of the outlet end. The outlet end is arranged at one end of the pulse cavity. The diameter of the inlet end is set according to the pulse frequency. The length of the pulse cavity is less than 0.5 mm.

2. The pulse jet structure of an aircraft engine turbine blade according to claim 1, characterized in that: The pulse cavity is a cylindrical cavity, with a nozzle portion at the front end of the cavity, an airflow cavity in the middle of the cavity, and an air inlet end at the rear end of the cavity; the nozzle structure is arranged in the nozzle portion and connected to the outside; the outlet end of the air inlet channel is arranged in the air inlet end.

3. The pulse jet structure of an aircraft engine turbine blade according to claim 1, characterized in that: The nozzle structure is a conical tube structure with a length of 0.2-0.4 mm and a diameter within 1 mm.

4. The aircraft engine turbine blade pulse jet structure according to claim 1, characterized in that: The diameter of the inlet end is selected and set according to the flow velocity of the cooling gas.

5. The pulse jet structure of an aircraft engine turbine blade according to claim 1, characterized in that: The center line of the air inlet passage and the center line of the nozzle structure are on the same axis.

6. A pulse jet method for an aircraft engine turbine blade, applicable to an aircraft engine turbine blade pulse jet structure as claimed in any one of claims 1 to 5, characterized in that: The following steps are involved: Step 1: setting the diameter of the inlet end of the air inlet channel and the size of the pulse cavity according to the required pulse frequency; Step 2, the impact cooling gas is introduced into the pulse chamber from the inlet end at a set flow rate, so that the impact cooling gas is continuously gathered in the pulse chamber; Step 3, when the pulse cavity is filled with the impact cooling gas, the gas in the pulse cavity is ejected from the nozzle structure by the impact cooling gas that subsequently enters; Step 4: Repeat this process over and over to form a pulse cooling gas with a certain frequency.

7. The method for pulse jetting of a turbine blade of an aero-engine according to claim 6, characterized in that: In step one, the sizes of the inlet and outlet are set according to the size of the cooling structure, the required cooling efficiency, and the flow rate of the cooling gas.

8. The method for pulse jetting of a turbine blade of an aero-engine according to claim 6, characterized in that: In step 2, the gas flows into the pulse chamber at a uniform flow rate.

9. The method for pulse jetting of a turbine blade of an aero-engine according to claim 6, characterized in that: It also includes setting the diameter of the nozzle structure to control the amount and speed of the gas ejected.

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