Film and impingement combined cooling structure suitable for turbine blades with low exhaust margin

Abstract

This application discloses a combined film cooling and impact cooling structure suitable for turbine blades with low exhaust margins, comprising: an impact plate with multiple impact holes; a hollow turbine blade with an inner cavity, the impact plate being placed within the inner cavity, forming an impact chamber between the impact plate and the turbine blade; multiple film cooling holes on the surface of the turbine blade to be cooled, with the airflow outlet of each impact hole facing the airflow inlet of a corresponding film cooling hole, suitable for allowing the core of the impact jet to enter the corresponding film cooling hole. The combined film cooling and impact cooling structure for turbine blades with low exhaust margins in this application has low pressure loss. Under the same cooling gas supply conditions, when placed at a position with low exhaust margins on the turbine blade surface, the risk of no cooling gas flowing out of the film cooling holes or even backflow of the main high-temperature combustion gas into the turbine blade is less likely to occur at this position.

Description

Technical Field

[0001] This application relates to the field of turbine blade cooling technology, specifically to a composite cooling structure of film cooling and impact cooling suitable for turbine blades with low exhaust margin. Background Technology

[0002] With the continuous improvement of gas turbine power and efficiency, the gas temperature at the turbine blade inlet has long exceeded the melting point of the turbine blade material. To prevent the turbine blade from melting during use, complex internal and external cooling structures are often arranged on the turbine blade. Among them, the internal cooling efficiency is the highest, which is impingement cooling, while external cooling is mainly film cooling. The composite cooling structure combining the two is widely used in the design of turbine blades. The existing composite cooling structure arranges the film cooling hole inlet and the impingement hole outlet alternately, so that the cold gas is internally cooled in the impingement chamber and then flows out through the nearby film cooling hole to achieve external cooling. However, although the staggered arrangement of the film cooling hole and the impingement hole has a high cold gas utilization rate and a good cooling effect, the impingement jet will first form a velocity stagnation area on the target surface after flowing out of the impingement hole. The impingement jet velocity in this area is extremely low, and the impact point will have a high heat transfer coefficient due to the sharp drop in velocity. This process will also result in a high pressure loss. The higher the pressure loss, the higher the inlet pressure P required to discharge the same amount of cold gas. s，c The higher the value, the greater the exhaust margin.

[0003] Exhaust margin refers to the pressure margin on the cold air side reserved to ensure that all outflow orifices of the turbine blades have cold air flowing out under any operating conditions. Its definition is shown in the following calculation formula:

[0004]

[0005] In the formula: P s，c The static pressure of the cold air inside the wall at the outlet orifice, MPa; P t，g The exhaust margin is the total pressure of the main flow at the outlet orifice, in MPa. Insufficient exhaust margin will lead to poor cooling effect and blade ablation, while too high exhaust margin will waste cool air and reduce turbine efficiency. The empirical value for exhaust margin of the cooling structure of the turbine blade body is about 1%.

[0006] Limited by casting costs and technology, existing composite cooling structures typically involve drilling holes in the surface of turbine blades and inserting impact plates inside. Due to the small size and thin blade body of turbine blades, the impact plates are cylindrical inserts to simultaneously meet the cooling requirements of both the suction and pressure sides of the blades. Since the cooling structures on both sides share the same cold air inlet, the static pressure P of the cold air within the walls at the outlet holes on both sides is the same. s，c The same applies, but the main flow pressure P at the outlet orifice on the pressure side... t，g Significantly higher than the mainstream total pressure P on the suction side t，gThis results in a lower exhaust margin on the pressure side compared to the exhaust margin on the suction side. Consequently, the cold air outflow on the pressure side is more affected by changes in turbine blade operating conditions, and it is also more likely to cause the danger of no cold air outflow or even backflow of mainstream high-temperature combustion gas into the turbine blades.

[0007] Therefore, it is necessary to design a new composite cooling structure for turbine blades that combines film cooling and impact cooling to meet the cold air outflow requirements at locations with low exhaust margins on the turbine blade surface.

[0008] Application content

[0009] Therefore, the technical problem to be solved by this application is to overcome the defect of the existing turbine blade film-impact composite cooling structure, which is prone to no cold air outflow or even backflow of mainstream high-temperature combustion gas at the low exhaust margin position, so as to provide a film-impact composite cooling structure suitable for turbine blades at the low exhaust margin position.

[0010] To solve the above-mentioned technical problems, the technical solution of this application is as follows:

[0011] A combined film and impingement cooling structure suitable for turbine blades with low exhaust margins includes:

[0012] An impact plate, wherein the impact plate is provided with multiple impact holes;

[0013] The turbine blade is hollow and has an inner cavity. The impact plate is placed in the inner cavity, and an impact cavity is formed between the impact plate and the turbine blade. Multiple air film holes are opened on the cooling wall surface of the turbine blade. The airflow outlet of the impact hole is directly opposite to the airflow inlet of one of the air film holes, which is suitable for allowing the core of the impact airflow to enter the directly opposite air film hole.

[0014] Furthermore, the angle between the axis of the air film hole and the wall surface to be cooled is α, where 0° < α ≤ 40°.

[0015] Furthermore, the diameter of the impact hole is larger than the diameter of the air film hole.

[0016] Furthermore, the ratio of the diameter of the impact hole to the diameter of the air film hole ranges from 2:1 to 3:1.

[0017] Furthermore, the air film pore is a cylindrical pore or a fan-shaped pore, and the diameter D of the air film pore is... f The thickness is 0.2mm to 1.5mm.

[0018] Furthermore, the impact hole is a cylindrical hole, and the diameter D of the impact hole is... i The thickness ranges from 0.5mm to 3mm.

[0019] Furthermore, the height H of the impact cavity is 0.5mm-5mm.

[0020] Furthermore, the thickness T of the impact plate i The thickness is 0.5mm to 5mm.

[0021] Furthermore, the thickness T of the cooling wall surface f The thickness ranges from 0.2mm to 5mm.

[0022] Furthermore, the flow pitch and spanwise pitch of both the impact hole and the film gas hole are 2D. i -5D i .

[0023] The technical solution of this application has the following advantages:

[0024] 1. The film cooling and impingement combined cooling structure for turbine blades with low exhaust margin provided in this application has the following advantages: the air outlet of the impingement hole is directly opposite the air inlet of a corresponding film cooling hole. The core of the impingement jet will be quickly discharged through the film cooling hole to cool the outer wall of the turbine blade. In this way, the impingement jet achieves a good heat transfer effect with a small pressure loss at the impact position. The high-pressure cooling gas can achieve a good film cooling and impingement combined cooling effect with a low initial pressure. This makes the exhaust margin required for the film cooling and impingement combined cooling structure for turbine blades provided in this application low. This means that when it is arranged at a low exhaust margin position on the blade surface, it is not easy to cause the danger of no cold air flowing out from the film cooling hole or even the backflow of the mainstream high-temperature combustion gas into the turbine blade.

[0025] 2. The film cooling and impingement combined cooling structure for turbine blades with low exhaust margin provided in this application has a smaller diameter for the film cooling orifice than for the impingement orifice. This increases the velocity of the impingement jet within the film cooling orifice, resulting in better heat transfer at the impingement location with less pressure loss. Consequently, the high-pressure cooling gas can achieve a better film cooling and impingement combined cooling effect with a lower initial pressure. In other words, the film cooling and impingement combined cooling structure for turbine blades provided in this application has a lower exhaust margin requirement, reducing the risk of insufficient cooling gas flow from the film cooling orifice or even backflow of high-temperature combustion gas into the turbine blade. Simulation results show that, under the same cooling gas outflow rate, the exhaust margin required by this application is only 81% of that of the traditional film-impingement combined cooling structure, while the overall cooling efficiency is only reduced by 3%. This demonstrates that this application can effectively reduce the exhaust margin requirement while maintaining the combined cooling effect, making it more suitable for turbine blade surfaces with low exhaust margins. Attached Figure Description

[0026] To more clearly illustrate the technical solutions in the specific embodiments of this application or the prior art, the drawings used in the description of the specific embodiments or the prior art will be briefly introduced below. Obviously, the drawings described below are some embodiments of this application. For those skilled in the art, other drawings can be obtained from these drawings without creative effort.

[0027] Figure 1 This is a three-dimensional schematic diagram of the turbine blade and impact plate in this application;

[0028] Figure 2 This is a top view of the turbine blade and impact plate in this application;

[0029] Figure 3 This is a cross-sectional schematic diagram of the turbine blade and impact plate in this application;

[0030] Figure 4 This is a schematic diagram of the simulation calculation domain for the combined film and impact cooling structure applicable to turbine blades with low exhaust margins, as described in this application.

[0031] Figure 5 The velocity distribution cloud map (partial) of a traditional air film and impact composite cooling structure;

[0032] Figure 6 This is a partial velocity distribution cloud map of the film and impact combined cooling structure applicable to turbine blades with low exhaust margins, as described in this application.

[0033] Explanation of reference numerals in the attached figures:

[0034] 1. Impact plate; 11. Impact hole; 2. Turbine blade; 21. Film cooling hole; 3. Impact chamber; 4. Main channel; 41. Main channel inlet; 42. Main channel outlet; 5. Cold air inlet; 51. Air supply chamber. Detailed Implementation

[0035] The technical solutions of this application will now be clearly and completely described with reference to the accompanying drawings. Obviously, the described embodiments are only some, not all, of the embodiments of this application. Based on the embodiments of this application, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of this application.

[0036] In the description of this application, it should be noted that the terms "center," "upper," "lower," "left," "right," "vertical," "horizontal," "inner," and "outer," etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are used only for the convenience of describing this application and simplifying the description, and do not indicate or imply that the device or element referred to must have a specific orientation, or be constructed and operated in a specific orientation. Therefore, they should not be construed as limitations on this application. Furthermore, the terms "first," "second," and "third" are used for descriptive purposes only and should not be construed as indicating or implying relative importance.

[0037] In the description of this application, it should be noted that, unless otherwise expressly specified and limited, the terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to a fixed connection, a detachable connection, or an integral connection; they can refer to a mechanical connection or an electrical connection; they can refer to a direct connection or an indirect connection through an intermediate medium; and they can refer to the internal connection between two components. Those skilled in the art can understand the specific meaning of the above terms in this application based on the specific circumstances.

[0038] Furthermore, the technical features involved in the different embodiments of this application described below can be combined with each other as long as they do not conflict with each other.

[0039] Example

[0040] like Figures 1 to 6 As shown, this embodiment provides a combined film and impact cooling structure suitable for turbine blades with low exhaust margins, including an impact plate 1 and a turbine blade 2.

[0041] like Figure 3 As shown, the thickness T of the impact plate 1 i The thickness is 0.5mm-5mm, preferably T. i The diameter is 1mm. Multiple impact holes 11 are formed on the impact plate 1. Each impact hole 11 is cylindrical and has a diameter D. i The thickness is 0.5mm-3mm, preferably D. i The diameter is 1mm. Of course, the impact hole 11 can also be elliptical or other shapes.

[0042] like Figure 3 The diagram shows a partial schematic of the turbine blade 2, which is partially considered to be flat. The actual turbine blade 2 is hollow, with an inner cavity. The impact plate 1 is placed within the inner cavity, and an impact cavity 3 is formed between the outer wall of the impact plate 1 and the inner wall of the turbine blade 2. The height H of the impact cavity 3 is 0.5mm–5mm, preferably 1mm. The extension direction of the impact hole 11 is perpendicular to the impact plate 1 and faces the outer wall of the turbine blade 2.

[0043] The surface on the turbine blade 2 that needs to be cooled is called the cooling surface, and the thickness T of the cooling surface is... f The thickness is 0.2mm-5mm, preferably T f The diameter is 1 mm, and multiple air film holes 21 are formed on the cooling wall surface. The angle between the axis of the air film hole 21 and the cooling wall surface is α, where 0° < α ≤ 40°, preferably α is 20°. The air film holes 21 are cylindrical or fan-shaped holes, and the diameter D of the air film holes 21 is... f The thickness is 0.2mm-1.5mm, preferably D. f It is 0.5mm.

[0044] like Figure 3 As shown, the airflow outlet of the impact hole 11 is positioned opposite the airflow inlet of a corresponding film gas hole 21, which is suitable for allowing the core of the impact jet (the core of the impact jet refers to the portion of the impact jet that maintains a constant flow velocity after flowing out of the impact hole) to enter the oppositely positioned film gas hole 21. In addition, in order to increase the velocity of the impact jet flowing out of the film gas hole 21, the diameter D of the film gas hole 21 is... f Smaller than the diameter D of the impact hole 11 i And the diameter D of the impact hole 11 i The diameter D of the air film pore 21 f The ratio range is 2:1 to 3:1, preferably, the diameter ratio is 2:1, and the flow pitch of the impact hole 11 and the air film hole 21 is ( Figure 4 (direction of A-A) and spanwise pitch ( Figure 4 The direction of B-B is 2D. i -5D i Preferably, the pitch in both directions is 4D. i .

[0045] See Figure 4 The following describes the cooling process of turbine blade 2 using a combined film and impact cooling structure suitable for turbine blades with low exhaust margins:

[0046] High-pressure cooling gas is ejected from the cooling gas inlet 5 into the air supply chamber 51, and then enters the impact hole 11 of the impact plate 1 to form an impact jet. After exiting the impact hole 11, the impact jet enters the impact chamber 3 to perform impact cooling on the inner wall of the turbine blade 2 near the impact chamber 3. Since the air outlet of the impact hole 11 is directly opposite the air inlet of a corresponding film cooling hole 21, the core of the impact jet in the impact chamber 3 is rapidly discharged from the film cooling hole 21 (the reason why the impact jet accelerates is because the diameter of the film cooling hole 21 is smaller than the diameter of the impact hole 11; when the gas flow channel narrows, the gas velocity increases). Therefore, the impact jet achieves a good heat exchange effect at the impact position with a small pressure loss. In this way, high-pressure cooling... The gas can achieve a good combined cooling effect of film cooling and impact cooling with a relatively low initial pressure. That is, the combined cooling structure of film cooling and impact cooling of turbine blades in this embodiment is suitable for positions with low exhaust margin. At the position of high mainstream pressure, it is not easy to cause the danger of no cold gas flowing out from the film cooling hole or even the backflow of high-temperature combustion gas into the turbine blade. The cooling gas discharged from the film cooling hole 21 forms a film cooling and flows along the outer wall of the turbine blade 2 under the action of the mainstream pressure gradient, cooling the outer wall of the turbine blade 2. In addition, since the extension direction of the film cooling hole 21 is inclined relative to the cooling wall surface, and 0°<α≤40°, it further promotes the flow of the cooling film toward the outer wall of the turbine blade 2, achieving a good cooling effect on the outer wall of the turbine blade 2.

[0047] To better illustrate the advantages of having the airflow outlet of the impact hole 11 aligned with the airflow inlet of the corresponding film cooling hole 21, velocity distribution cloud maps of a traditional film cooling and impact cooling composite structure and a film cooling and impact cooling composite structure suitable for turbine blades with low exhaust margins were specially created.

[0048] like Figure 5 The figure shows the velocity distribution cloud diagram of a traditional film cooling and impact cooling composite structure. As can be seen from the figure, due to the misaligned arrangement of the impact hole 11 and the film cooling hole 21, after the impact jet exits the impact hole 11 and enters the impact chamber 3, a velocity stagnation region (e.g., ...) is first formed at the impacted location on the inner wall of the turbine blade 2. Figure 5 (The area circled in dashed lines) In this region, the velocity of the impact jet is extremely low. As a result, the impact jet will have a high heat transfer coefficient due to the sharp drop in velocity. Furthermore, the impact jet will generate a high pressure loss in this region, which will result in insufficient pressure of the cooling gas discharged from the impact chamber 3 through the film gas hole 21. If normal cold gas outflow is to be ensured and the danger of no cold gas flowing out from the film gas hole or even the backflow of the mainstream high-temperature combustion gas into the turbine blade 2 is to be avoided, the initial pressure of the high-pressure cooling gas entering the impact plate 1 needs to be increased, making it unsuitable for positions with low exhaust margin.

[0049] like Figure 6The figure shows the velocity distribution cloud map of the film cooling and impact cooling composite structure applicable to the turbine blade with low exhaust margin in this embodiment. As can be seen from the figure, since the airflow outlet of the impact hole 11 is directly opposite to the airflow inlet of a corresponding film cooling hole 21, most of the impact jet is directly drawn out by the film cooling hole 21. A small portion of the impact jet flows at a relatively high speed on the turbine blade 2 near the inner wall of the impact chamber 3 and then flows out from the upstream film cooling hole 21. Furthermore, in the film and impact combined cooling structure applicable to turbine blades with low exhaust margin in this embodiment, the impact jet will accelerate rather than stagnate near the position on the turbine blade 2 where it is impacted (the reason why the impact jet accelerates is because the diameter of the film hole 21 is smaller than the diameter of the impact hole 11, and the gas velocity increases when the gas flow channel narrows). Therefore, the impact jet achieves a better heat exchange effect with a smaller pressure loss at the impact position. In this way, the high-pressure cooling gas can achieve a better film and impact combined cooling effect with a lower initial pressure, and the exhaust margin required to meet the same cold gas output is lower.

[0050] Additionally, refer to Figure 4 The diagram shows the interaction model between the film cooling and impingement cooling structure and the mainstream combustion gas and cooling gas in a turbine blade with low exhaust margin. Simulation calculations show that, in this embodiment, the film cooling and impingement cooling structure for turbine blades with low exhaust margin reduces the required exhaust margin by 19% while maintaining the same cold gas outflow rate, while the overall cooling efficiency only decreases by 3%. This indicates that the film cooling and impingement cooling structure for turbine blades with low exhaust margin can effectively reduce the exhaust margin requirement while maintaining the combined cooling effect. The formula for calculating the overall cooling efficiency is as follows:

[0051]

[0052] In the formula: T t，g The main inlet temperature is K; T is the wall temperature, K; T t，c The total temperature at the air inlet is K.

[0053] Obviously, the above embodiments are merely illustrative examples for clear explanation and are not intended to limit the implementation. Those skilled in the art will recognize that other variations or modifications can be made based on the above description. It is neither necessary nor possible to exhaustively list all possible implementations here. However, obvious variations or modifications derived therefrom are still within the scope of protection of this application.

Claims

1. A combined film and impact cooling structure suitable for turbine blades with low exhaust margin, characterized in that, include: Impact plate (1), wherein the impact plate (1) is provided with a plurality of impact holes (11); Turbine blade (2) is hollow and has an inner cavity. The impact plate (1) is placed in the inner cavity. An impact cavity (3) is formed between the impact plate (1) and the turbine blade (2). Multiple air film holes (21) are opened on the cooling wall surface of the turbine blade (2). The airflow outlet of the impact hole (11) is directly opposite to the airflow inlet of one of the air film holes (21) at the corresponding position, which is suitable for allowing the core of the impact jet to enter the air film hole (21) directly opposite to it. The angle between the axis of the air film hole (21) and the cooling wall surface is α, where 0° < α ≤ 40°; The diameter D of the impact hole (11) i Larger than the diameter D of the air film pore (21) f ; The diameter D of the impact hole (11) i The diameter D of the air film pore (21) f The ratio ranges from 2:1 to 3:

1.

2. The film cooling and impingement combined cooling structure for turbine blades with low exhaust margin as described in claim 1, characterized in that, The air film pore (21) is a cylindrical pore or a fan-shaped pore, and the diameter D of the air film pore (21) is... f The thickness ranges from 0.2mm to 1.5mm.

3. The film cooling and impact cooling composite structure for turbine blades with low exhaust margin as described in claim 1, characterized in that, The impact hole (11) is a cylindrical hole, and the diameter D of the impact hole (11) is... i The thickness ranges from 0.5mm to 3mm.

4. The film cooling and impact cooling composite structure for turbine blades with low exhaust margin as described in claim 1, characterized in that, The height H of the impact chamber (3) is 0.5mm-5mm.

5. The film cooling and impact cooling composite structure for turbine blades with low exhaust margin as described in claim 4, characterized in that, The thickness T of the impact plate (1) i The thickness ranges from 0.5mm to 5mm.

6. The film cooling and impact cooling composite structure for turbine blades with low exhaust margin as described in claim 5, characterized in that, The thickness T of the cooling wall surface f The thickness ranges from 0.2mm to 5mm.

7. The film cooling and impingement combined cooling structure for turbine blades with low exhaust margins as described in any one of claims 1-6, characterized in that, The flow pitch and spanwise pitch of both the impact hole (11) and the film gas hole (21) are 2D. i -5D i .

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