Turbine guide vane for improving blade air film cooling effect under rotational flow influence and cooling method

Abstract

The turbine guide vane capable of improving the blade air film cooling effect under the rotational flow influence comprises a turbine guide vane body and air film holes formed in the guide vane front edge of the turbine guide vane body in a matrix mode, and the air film holes are used for restraining air film deviation of the turbine guide vane body under the rotational flow condition and increasing the air film coverage area. The air film hole is a discontinuous annular hole composed of at least two fan ring holes and fan ring partitions used for connecting every two fan ring holes. The invention also provides a cooling method of the turbine guide vane, which comprises the following steps: the gas main flow containing rotational flow impacts on the wall surface of the front edge of the guide vane, the cold air flows out of the gas film holes to form gas film outflow, and on one hand, the gas film outflow is mixed with the gas main flow to reduce the temperature of the gas on the wall surface; and on the other hand, the cold air forms a gas film covering and blocking the scouring of the fuel gas to the wall surface on the front edge wall surface of the guide vane. The device is suitable for gas film outflow under the rotational flow condition, gas film deviation is restrained, the gas film coverage area is increased, and effective gas film coverage is formed.

Description

Technical Field

[0001] The present invention relates to the field of gas turbine engines and relates to a turbine guide vane and a cooling method for the guide vane. Background Art

[0002] Increasing the inlet temperature of the turbine can improve the thermal efficiency and performance of the gas turbine engine. The cooling technology of the turbine blades can ensure the safe and stable operation of the blades when the gas temperature far exceeds the temperature resistance limit of the material. It is a key technology that determines the performance level of the engine. Film cooling is an efficient cooling technology and is widely used.

[0003] In recent years, the influence of the combustion chamber outlet flow field on the aerodynamic and heat transfer performance of turbine blades has been widely studied. Previous research results show that swirl causes the turbine interstage efficiency to decrease by up to 3.1%, and the heat load of the guide vanes and rotor blades caused by swirl increases by 200% and 40%, respectively. The cooling design of the turbine needs to consider the influence of the uneven flow field at the outlet of the upstream combustion chamber. Residual swirl is prone to occur at the outlet of the low-pollution combustion chamber. The swirl vortex has a strong interactive effect with the mainstream of the turbine guide vane cascade channel, making it difficult for the cooling structure of the turbine guide vane to meet the cooling requirements under swirl inlet conditions under uniform inflow.

[0004] When the mainstream velocity at the leading edge of the blade is the lowest, the proportion of the swirl velocity in the combined velocity is the largest, and the swirl has the most significant effect on the cooling of the leading edge film. Strong swirl can cause the film to shift significantly. The mechanism of this phenomenon is: the swirl inlet changes the inlet attack angle of the blade channel, distorts the pressure stagnation line at the leading edge of the blade, and brings additional radial pressure gradients, which causes the film to shift seriously; at the same time, the cold air converges in the low-pressure area, and the high-pressure area cannot be covered by the film, making the cooling efficiency of the local area close to zero, and the average value of the cooling efficiency surface is reduced by as much as 23%.

[0005] For air film cooling under the influence of swirl flow, the outflow direction of all air film holes at the leading edge can be made opposite to the radial pressure gradient, that is, it can resist the deflection effect of the swirl flow on the air film and expand the coverage area of ​​the air film; or reducing the inclination angle of the leading edge air film holes and changing the cylindrical holes into traditional expanded special-shaped holes can also improve the cooling efficiency of the leading edge air film, but it will not have any inhibitory effect on the deflection of the air film. Summary of the invention

[0006] The purpose of the present invention is to avoid the shortcomings of the prior art and provide a turbine guide vane and cooling method that can overcome the influence of swirl, increase the effective air film coverage and coverage area, reduce the wall temperature of the turbine guide vane, and improve the blade air film cooling effect under the influence of swirl.

[0007] To achieve the above object, the technical solution adopted by the present invention is as follows: A turbine guide vane for improving the film cooling effect of blades under the influence of swirl flow, including a turbine guide vane and film holes arranged in a matrix on the leading edge of the guide vane. The film holes are used to suppress film offset under swirl flow conditions of the turbine guide vane and increase the film coverage area;

[0008] The film hole is a discontinuous annular hole composed of at least two sector ring holes and sector ring partitions for connecting two adjacent sector ring holes;

[0009] The inner diameter of the annular inner ring formed by the film holes is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the sector ring hole. Then:

[0010] On the leading edge of the guide vane, the width B of the film hole < 0.5D is provided to increase the outlet momentum of the film hole, overcome air flow stagnation and large pressure loads, and improve the film cooling effect of the leading edge of the guide vane;

[0011] On the blade body of the turbine guide vane, the width B of the film hole is taken as: 0.1D < B < 0.2D, for obtaining the cold air coverage effect;

[0012] The matrix arrangement of the film holes on the leading edge of the guide vane is specifically as follows:

[0013] Map the leading edge of the guide vane to a plane and divide it into multiple identical area units. The film holes are arranged in the area units. Then the longitudinal unit length Sy of the area unit in the blade height direction of the leading edge of the guide vane, and the transverse unit length Sx in the leading edge arc length direction in the spanwise direction. The opening ratio of the film holes in the area unit is:

[0014] Film hole area / Sx × Sy = 5 - 15%,

[0015] The film holes are continuously or staggeredly extended row by row on the leading edge of the guide vane from the blade root to the blade tip.

[0016] Further, the film holes are continuously or staggeredly extended row by row, and at least four rows are provided.

[0017] Further, the film holes include 2 - 6 sector ring holes.

[0018] Further, the areas of the multiple sector ring holes of the film holes are equal or different, but the total area of the multiple sector ring holes is the same as the area of a circular hole with a diameter d ’ of 0.5 - 2 mm.

[0019] Further, the center point of the annular shape formed by the film holes coincides with the midpoint of the area unit, that is, the film holes are arranged at the center position of the area unit.

[0020] Furthermore, the distance between the air film holes arranged in the blade height direction is P1, the distance between the air film holes arranged in the leading edge arc length direction is P2, and the diameter of the circular hole with the same area as the fan ring hole is d ’ , then:

[0021] P1=2.5~15d'; P2=1.5~10d'.

[0022] Furthermore, the air film holes are normal through holes or backward expansion holes of the cooling wall surface;

[0023] When it is a normal through hole, a pair of inclination angles α1=α2=90° are set correspondingly in the cross section of the air film hole;

[0024] When the hole is expanded backward, the sum of the pair of inclination angles of the cross section of the air film hole corresponding to each other is 90 degrees, and:

[0025] 30°≤α1≤90°, 30°≤α2≤90°.

[0026] Further, the fan ring partition has a pair of arc-shaped connecting edges and a pair of parallel or inclined edges;

[0027] When the parallel sides are described, the distance L between the two sides is 0.05-0.25 mm;

[0028] When it is the inclined edge, the central angle corresponding to the inclined edge is θ, and the range of θ is 15 to 45 degrees.

[0029] Furthermore, the air film hole is a discontinuous elliptical annular hole;

[0030] The elliptical inner ring surrounded by the air film holes has a longer chord length a and a shorter chord length b, and then: b<2a.

[0031] The present invention also provides a cooling method for the turbine guide vane as described above, comprising: a swirling gas mainstream C2 impacts the guide vane leading edge wall surface, and cold air C1 flows out from the air film hole to form an air film outflow C3, and the air film outflow C3 is mixed with the gas mainstream C2 to reduce the temperature of the wall gas; on the other hand, the cold air C1 forms an air film on the guide vane leading edge wall surface to cover and block the gas from scouring the wall surface;

[0032] In this process, since the gas mainstream C2 contains swirl, the air film is distorted, but since the air film hole is composed of multiple fan ring holes, the cold air C1 enters the air film hole from a wider range and a narrower hole channel, which improves the multi-directional expansion of the air film outflow C3, enhances the normal momentum of the air film outflow C3, and weakens the effect of the mainstream C2 with swirl on the air film cold air, so that the air film hole has the ability to suppress the distortion of the air film caused by the swirl, and realizes uniform outflow under swirl conditions, thereby forming a stable air film and curbing the deviation of the air film;

[0033] At the same time, multiple fan ring holes drive the mainstream gas C2 to reduce the outlet momentum of the air film holes, effectively improving the downstream air film cooling efficiency and cold air utilization rate.

[0034] The beneficial effects of the present invention are as follows: the present invention proposes a cooling structure for high-temperature components of aircraft engines under swirl conditions, using a special structure air film hole set on the leading edge of the guide vane, the air film hole is composed of multiple short rings, on the one hand, it is suitable for air film outflow under swirl conditions, curbs the deviation of the air film, increases the coverage area of ​​the air film, forms effective air film coverage, and avoids the consequences of reduced cooling efficiency caused by the change of air film direction under swirl in traditional circular air film holes, slot-shaped holes and other single special-shaped holes; on the other hand, compared with a single hole, the air film coverage area composed of multiple short ring air film holes is increased, the air film cooling efficiency is improved, the wall temperature is effectively reduced, and the cold air utilization rate is improved, that is, the new air film hole can effectively improve the air film cooling effect under swirl conditions. BRIEF DESCRIPTION OF THE DRAWINGS

[0035] Figure 1 A schematic diagram of the structure of the turbine guide vane of the present invention;

[0036] Figure 2 A schematic diagram of the arrangement of the air film holes with the guide vane leading edge mapped as a plane according to the present invention;

[0037] Figure 3 Schematic diagram of the top view of the air film hole of the present invention;

[0038] Figure 4 Schematic diagram of the cross-sectional structure of the air film hole of the present invention;

[0039] Figure 5 Another schematic diagram of the top view of the air film plate of the present invention;

[0040] Figure 6 It is the streamline diagram of a single annular hole under the swirl condition of the present invention;

[0041] Figure 7 Schematic diagram of the cooling efficiency distribution of cylindrical film holes under swirl conditions;

[0042] Figure 8 The present invention provides a schematic diagram of the distribution of film hole cooling efficiency under swirl conditions;

[0043] Fig. 9 Schematic diagram of the radial average cooling efficiency distribution under swirling flow conditions;

[0044] Fig.10 Schematic diagram of the spanwise average cooling efficiency distribution under swirling flow conditions. Specific implementation manners

[0045] The principles and features of the present invention will be described below with reference to the accompanying drawings. The examples given are only for explaining the present invention and are not intended to limit the scope of the present invention.

[0046] To achieve the above object, the present invention provides the following specific implementation manners:

[0047] Example 1: As shown in Figure 1-10 A turbine guide vane for improving the film cooling effect of a blade under the influence of swirling flow includes a turbine guide vane 1 and film holes 2 arranged in a matrix on the leading edge 11 of the guide vane of the turbine guide vane 1. The film holes 2 are used to curb the film offset of the turbine guide vane 1 under swirling flow conditions and increase the film coverage area;

[0048] As shown in Figure 1 The film hole 2 is a discontinuous annular hole composed of at least two sector ring holes 21 and sector ring partitions 22 for connecting two adjacent sector ring holes 21;

[0049] The inner diameter of the annular inner ring surrounded by the film hole 2 is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the sector ring hole 21. Then:

[0050] On the leading edge 11 of the guide vane, the width B of the film hole 2 < 0.5D, which is used to increase the outlet momentum of the film hole, overcome the airflow stagnation and large pressure load, and improve the film cooling effect of the leading edge 11 of the guide vane;

[0051] On the blade body of the turbine guide vane 1, the value of the width B of the film hole 2 is: 0.1D < B < 0.2D, which is used to obtain the cold air coverage effect;

[0052] As shown in Figure 2 The matrix arrangement of the film holes 2 on the leading edge 11 of the guide vane is specifically:

[0053] The leading edge 11 of the guide vane is mapped to a plane and divided into multiple identical area units. The film holes 2 are arranged in the area units. The longitudinal unit length Sy of the leading edge 11 of the guide vane in the height direction and the transverse unit length Sx of the leading edge arc length in the spanwise direction are defined. The opening ratio of the film holes 2 in the area unit is:

[0054] Area of film hole 2 / Sx × Sy = 5 - 15%,

[0055] The film holes 2 are arranged in rows continuously or staggered on the leading edge 11 of the guide vane from the blade root to the blade tip, and at least four rows are provided. The center point of the ring surrounded by the film holes 2 coincides with the midpoint of the area unit, that is, the film holes 2 are arranged at the center position of the area unit.

[0056] like Figure 3-5 As shown, the air film hole 2 includes 2 to 6 fan ring holes 21. The areas of the multiple fan ring holes 21 of the air film hole 2 are equal or different, but the total area of ​​the multiple fan ring holes 21 is proportional to the diameter d ’ The area of ​​the circular holes of 0.5 to 2 mm is the same.

[0057] The hole spacing of the air film holes 2 arranged in the blade height span direction is P1, the hole spacing of the air film holes 2 arranged in the leading edge arc length span direction is P2, and the diameter of the circular hole with the same area as the fan ring hole 21 is d ’ , then:

[0058] P1=2.5~15d'; P2=1.5~10d'.

[0059] like Figure 4 As shown, the air film hole 2 is a normal through hole or a backward expansion hole of the cooling wall surface 1;

[0060] When it is a normal through hole, a pair of inclination angles α1=α2=90° are set correspondingly in the cross section of the air film hole 2;

[0061] When the hole is expanded backward, the sum of the pair of inclination angles of the cross section of the air film hole 2 corresponding to each other is 90 degrees, and:

[0062] 30°≤α1≤90°, 30°≤α2≤90°.

[0063] like Figure 5 As shown, the fan ring partition 22 has a pair of arc-shaped connecting edges and a pair of parallel or inclined edges;

[0064] When the parallel sides are described, the distance L between the two sides is 0.05-0.25 mm;

[0065] When it is the inclined edge, the central angle corresponding to the inclined edge is θ, and the range of θ is 15 to 45 degrees.

[0066] Embodiment 2: Same as embodiment 1, except that Figure 3 As shown in (d), the air film hole 2 is a discontinuous elliptical ring hole; the air film hole 2 forms an elliptical inner ring, the longer chord length of the elliptical inner ring is a and the shorter chord length is b, then: b<2a.

[0067] Example 3: Figure 6As shown, the present invention also provides a cooling method for the turbine guide vane as described above, comprising: the swirling gas mainstream C2 impacts the guide vane leading edge 11 wall surface, and the cold air C1 flows out from the air film hole 2 to form an air film outflow C3, and the air film outflow C3 is mixed with the gas mainstream C2 to reduce the temperature of the wall gas; on the other hand, the cold air C1 forms an air film on the guide vane leading edge 11 wall surface to cover and block the gas from scouring the wall surface;

[0068] In this process, since the gas mainstream C2 contains swirl, the air film is distorted, but since the air film hole 2 is composed of multiple fan ring holes 21, the cold air C1 enters the air film hole 2 from a wider range and a narrower hole channel, which improves the multi-directional expansion of the air film outflow C3, enhances the normal momentum of the air film outflow C3, and weakens the effect of the mainstream C2 with swirl on the air film cold air, so that the air film hole 2 has the ability to suppress the distortion of the air film caused by the swirl, and realizes uniform outflow under swirl conditions, thereby forming a stable air film and curbing the deviation of the air film;

[0069] At the same time, the plurality of fan ring holes 21 drive the main flow of gas C2 to reduce the outlet momentum of the air film hole 2, thereby effectively improving the downstream air film cooling efficiency and cold air utilization rate.

[0070] The simulation results of the turbine guide vanes provided by the present invention are as follows: Figure 7-10 As shown:

[0071] The present invention uses the air film holes 2 composed of a plurality of semicircular ring structures in the turbine guide vane, so as to overcome the problem that the air film holes of the turbine guide vane have air film outflows C3 with multiple directions of expansion under swirl conditions, and can form effective air film coverage, thereby avoiding the consequences of reduced cooling efficiency caused by the change of air film direction under swirl conditions in traditional circular air film holes, slot-shaped holes and other single special-shaped holes;

[0072] On the other hand, compared with a single hole, the air film coverage area composed of multiple short annular air film holes used in the present invention is increased, the air film cooling effect is improved, the wall temperature is effectively reduced, and the cold air utilization rate is improved. That is, the new air film hole can effectively improve the air film cooling effect under swirl conditions.

[0073] Figure 7 The distribution of film cooling efficiency of the cylindrical hole is shown. It can be found from the figure that due to the effect of swirl, the area covered by the cold air flowing out of the film hole is distorted. The film cooling efficiency near the hub (lower end wall) on the pressure side is significantly enhanced, and the film cooling efficiency near the casing (upper end wall) is slightly improved, while the opposite is true for the suction side. A low film cooling efficiency area appears in the middle blade height area of ​​the leading edge, and a large area of ​​low cooling efficiency also appears in the middle blade height area between the suction surface and the pressure surface. This not only makes some areas almost completely lose the protection of the air film, but also increases the unevenness of the air film distribution, resulting in a significant increase in thermal stress.

[0074] Figure 8 The film cooling efficiency distribution of the present invention is demonstrated. Compared with the cylindrical film hole, the film cooling effect of the annular film hole under swirl conditions is significantly enhanced. In the leading edge area, the area of ​​the high film cooling efficiency area of ​​the blade using the annular film hole is larger and the value is higher. This is because the annular hole has high adaptability to the incoming flow direction and can achieve relatively uniform film outflow under various swirl conditions. On the other hand, compared with a single hole, the combination of multiple short-sector film holes can significantly reduce the outlet momentum of the film hole, effectively improving the downstream film cooling efficiency and cold air utilization rate.

[0075] Under the condition of swirl number S=0.3, the radial pressure gradient caused by the swirl makes the fluid on the suction surface tend to deflect upward as a whole, while the pressure surface is offset to the middle and lower part. The distribution of film cooling efficiency with leading edge cylindrical film holes is significantly affected by the swirl, and the high film cooling efficiency area is concentrated in the leading edge near the upper end wall area and the front of the suction surface near the upper end wall area, while the film cooling efficiency distribution with leading edge annular film holes is less affected by the swirl.

[0076] Fig. 9 The schematic diagram of radial average cooling efficiency distribution under swirl conditions is shown in Figure 2. Fig. 9 It can be found that the radial average cooling efficiency of the leading edge cylindrical film holes under swirl conditions is the highest at the upper end wall and decreases radially, while the radial average cooling efficiency distribution uniformity of the leading edge annular film holes is significantly improved.

[0077] Fig.10 is a schematic diagram of the spanwise average cooling efficiency distribution under swirl conditions. Fig.10 It can be found that compared with traditional cylindrical film holes, the film cooling effect of the suction surface of the blade with annular film holes is significantly improved, the leading edge area is slightly improved, the pressure surface area is slightly decreased, and the overall film cooling efficiency of the entire surface is improved by 5.41%.

[0078] The above description is only a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, improvements, etc. made within the spirit and principle of the present invention should be included in the protection scope of the present invention.

Claims

1. A turbine guide vane for improving the film cooling effect of blades under the influence of swirl, characterized in that, it includes a turbine guide vane (1) and film holes (2) arranged in a matrix on the leading edge (11) of the guide vane of the turbine guide vane (1). The film holes (2) are used to contain the film offset under the swirl condition of the turbine guide vane (1) and increase the film coverage area; the film hole (2) is a discontinuous annular hole composed of at least two sector ring holes (21) and a sector ring partition (22) for connecting two adjacent sector ring holes (21); the inner diameter of the annular inner ring surrounded by the film hole (2) is d, the outer diameter is D, and the distance between the outer ring and the inner ring is the width B of the sector ring hole (21), then: on the leading edge (11) of the guide vane, the width B of the film hole (2) < 0.5D, which is used to increase the outlet momentum of the film hole, overcome the air flow stagnation and large pressure load, and improve the film cooling effect of the leading edge (11) of the guide vane; on the blade body of the turbine guide vane (1), the width B of the film hole (2) is taken as: 0.1D < B < 0.2D, which is used to obtain the cold air coverage effect; the matrix arrangement of the film holes (2) on the leading edge (11) of the guide vane is specifically: mapping the leading edge (11) of the guide vane to a plane and dividing it into multiple identical area units. The film holes (2) are arranged in the area units. Then, in the leading edge (11) of the guide vane, the longitudinal unit length in the height direction is Sy, and the transverse unit length in the arc length direction of the leading edge is Sx. The opening ratio of the film holes (2) in the area unit is: area of the film hole (2) / Sx × Sy = 5 - 15%, the film holes (2) are continuously or staggeredly extended and arranged row by row on the leading edge (11) of the guide vane from the blade root to the blade tip.

2. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, the film holes (2) are continuously or staggeredly extended and arranged row by row, and at least four rows are provided.

3. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, the film hole (2) includes 2 - 6 sector ring holes (21).

4. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, The areas of the plurality of fan ring holes (21) of the air film hole (2) are equal or different, but the total area of ​​the plurality of fan ring holes (21) is proportional to the diameter d ’ The area of ​​the circular holes of 0.5 to 2 mm is the same.

5. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, the center point of the ring surrounded by the film hole (2) coincides with the midpoint of the area unit, that is, the film hole (2) is arranged at the center position of the area unit.

6. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, The hole spacing of the air film holes (2) arranged in the blade height direction is P1, the hole spacing of the air film holes (2) arranged in the leading edge arc length direction is P2, and the diameter of the circular hole with the same area as the fan ring hole (21) is d ’ , then: P1 = 2.5 - 15d'; P2 = 1.5 - 10d'.

7. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, characterized in that, the film hole (2) is a normal through hole or a backward expanding hole of the cooling wall surface (1); When the air film hole (2) is a normal through hole, a pair of inclination angles α are set correspondingly in the cross section of the air film hole (2) 1 =α 2 =90°; when it is a backward expanding hole, the sum of the angles of a pair of inner inclination angles corresponding to the cross section of the film hole (2) is 90 degrees, and: 30°≤α 1 ≤90°，30°≤α 2 ≤90°。 8. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl according to claim 1, It is characterized in that The fan ring partition (22) has a pair of arc-shaped connecting edges and a pair of parallel or inclined edges; When the parallel sides are described, the distance L between the two sides is 0.05-0.25 mm; When it is the inclined edge, the central angle corresponding to the inclined edge is θ, and the range of θ is 15 to 45 degrees.

9. The turbine guide vane for improving the film cooling effect of blades under the influence of swirl as claimed in any one of claims 1 to 8, It is characterized in that The air film hole (2) is a discontinuous elliptical annular hole; The elliptical inner ring formed by the air film hole (2) has a longer chord length a and a shorter chord length b, and then: b<2a.

10. A method for cooling a turbine guide vane according to claims 1 to 9, It is characterized in that The main stream of gas C2 containing the swirl flow impacts the wall surface of the guide vane leading edge (11), and the cold gas C1 flows out from the gas film hole (2) to form a gas film outflow C3. On the one hand, the gas film outflow C3 is mixed with the main stream of gas C2 to reduce the temperature of the gas on the wall surface; on the other hand, the cold gas C1 forms a gas film on the wall surface of the guide vane leading edge (11) to cover and block the gas from scouring the wall surface; In this process, since the main gas flow C2 contains a swirl flow, the air film is distorted. However, since the air film hole (2) is composed of a plurality of fan ring holes (21), the cold air C1 enters the air film hole (2) from a wider range and a narrower hole channel, thereby increasing the multi-directional expansion of the air film outflow C3, enhancing the normal momentum of the air film outflow C3, and weakening the effect of the main gas flow C2 with a swirl flow on the air film cold air, so that the air film hole (2) has the ability to suppress the distortion of the air film caused by the swirl flow, and realizes uniform outflow under swirl flow conditions, thereby forming a stable air film and curbing the deviation of the air film. At the same time, the plurality of fan ring holes (21) drive the main flow of the gas C2 to reduce the outlet momentum of the air film hole (2), thereby effectively improving the downstream air film cooling efficiency and cold air utilization rate.

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