Aircraft film cooling hole cooling structure for an aeroengine and turbine blade

Abstract

The application provides an aircraft film hole cooling structure for an aero-engine, which comprises an aircraft film hole core, the aircraft film hole core comprises a fixedly connected cylinder and a cone, the outlet shape of the cone is in the shape of an aircraft, compared with common special-shaped film holes, such as a dustpan-shaped hole, a conical hole and a water-drop-shaped hole, the outlet shape of the film hole is complicatedly designed, the cold air outflow flow can be controlled through the round hole, a better film hole type can be obtained, the cold air can be fully developed towards the blade span, and thus the span coverage area of the film is increased; the generation and development of the reverse kidney-shaped vortex at the outlet of the film hole can be effectively inhibited, and the cooling result downstream of the film hole is improved. In addition, the aircraft film hole structure provided by the application has higher film efficiency, compared with common round holes and dustpan-shaped special-shaped film holes, under the same cold air amount, the film efficiency is respectively increased by 100% and 50%.

Description

Technical Field

[0001] This invention belongs to the field of gas turbine blade technology, specifically relating to an aircraft-type film cooling structure and turbine blade for aero engines. Background Technology

[0002] Film cooling (FS) is one of the most important cooling methods for turbine blades in aero-engines. It involves drawing cool air from inside the blade through film orifices and attaching it to the blade wall, thus preventing direct contact between the high-temperature combustion gases and the blade. Advanced aero-engine turbine blades employ complex film cooling structures. Cylindrical film orifices are the most common and easiest to manufacture type of film cooling orifice in aero-engines. However, under high blow ratio conditions, the momentum of the cylindrical orifice film jet is too concentrated, making it prone to penetrating the main stream and detaching from the wall. Although this problem is somewhat improved at low blow ratios, the effective utilization rate of cool air remains low due to the limited spanwise coverage of the cylindrical orifice jet. Researchers have proposed various techniques to improve film cooling performance. These techniques can be broadly categorized into two types: one is to modify the flow structure of the film jet by adding auxiliary structures such as protrusions, ridges, and grooves at the cylindrical orifice outlet, allowing the film to adhere better to the wall; the other is to improve the film orifice structure to alter the flow structure of the film jet, thereby enhancing the film cooling effect.

[0003] Numerous research studies both domestically and internationally have shown that irregularly shaped film cooling orifices can effectively reduce the penetration of the film jet into the mainstream, making it easier for the film jet to adhere to the wall surface under the influence of the mainstream, thereby achieving better film cooling effects. Designing better orifice shapes to achieve better film cooling effects with the same or less cooling gas mass flow rate is currently a hot topic in film cooling research. Summary of the Invention

[0004] To address the aforementioned problems, the present invention aims to provide an aircraft-type film cooling structure suitable for air-cooled turbine blades of aero engines. Under a certain amount of cooling air, this structure can significantly improve the film cooling effect of the blades, thereby enhancing their heat resistance.

[0005] To achieve the above objectives, the present invention provides the following technical solution: an aircraft-type film cooling structure for aero engines, the cooling structure comprising an aircraft-type film cooling core, the aircraft-type film cooling core comprising a cylinder and a cone fixedly connected together, the outlet shape of the cone being aircraft-shaped, the aircraft shape comprising a first arc segment, a first straight segment, a second arc segment, a second straight segment, a third straight segment, a fourth straight segment, a fifth straight segment, a third arc segment, and a sixth straight segment connected in sequence.

[0006] The intersection of the third and fourth straight line segments lies on the same vertical line as the center of the first arc segment.

[0007] The aircraft shape is symmetrical about its axis.

[0008] The aircraft-type film cooling structure for aero engines provided by the present invention also has the following feature: the height of the cylinder is L1, the height of the cone is L2, and then L2 = (L1 + L2) * K, where K is a constant, K = 0.4 to 0.6.

[0009] The aircraft-type film cooling structure for aero engines provided by the present invention also has the following features: the included angle between the first straight segment and the sixth straight segment is α, and the included angle formed between the dot of the second arc segment, the dot of the third arc segment and the dot of the first arc segment is δ, wherein δ > α.

[0010] The aircraft-type film cooling structure for aero engines provided by the present invention also has the following feature: the included angle between the first straight segment and the second straight segment is β, where β = 25°-35°.

[0011] The angle between the third and fourth line segments is γ, where γ = 130° - 155°.

[0012] The aircraft-type film cooling structure for aero engines provided by the present invention also has the following feature: the diameter of the cylinder is D1, and the distance between the center of the first arc segment and the intersection of the third and fourth straight segments is L5, then L5 = D1.

[0013] The aircraft-type film cooling structure for aero engines provided by this invention also has the following feature: the intersection of the third and fourth straight segments is designated as point G, the cross section between the cylinder and the cone is designated as section 04, and the angle between the shortest straight line from point G to section 04 and the center line of the cylinder is the forward tilt angle θ of the cone, where θ = 9°-12°.

[0014] Another object of the present invention is to provide a turbine blade, the turbine blade comprising a blade body and a plurality of aircraft-type film cooling structures disposed on the blade body, wherein the aircraft-type film cooling structures are any of the aforementioned aircraft-type film cooling structures.

[0015] The angle between the central axis of the aircraft-type film cooling structure and the blade surface is ε, where ε = 35°-50°.

[0016] The air film vent cooling structure of the aircraft type has its vent openings oriented downstream of the airflow.

[0017] The turbine blade provided by the present invention also has the feature that the hole spacing of the aircraft-type film cooling structure is in the range of 5-10 times D1.

[0018] Beneficial effects

[0019] The aircraft-style film cooling structure for aero-engines provided by this invention features a complex design in the outlet shape compared to common irregular film cooling holes, such as scoop-shaped holes, conical holes, and teardrop-shaped holes. The structure formed by the arc segment 13, straight segments 15, and 16 allows the cool air to develop better along the spanwise direction, thereby increasing the spanwise coverage area of ​​the film cooling system. The structure formed by the straight segments 18 and 19 can effectively suppress the generation and development of the reverse kidney-shaped vortex at the outlet of the film cooling hole, improving the cooling effect downstream of the film cooling hole. The resulting film cooling efficiency is higher, increasing by 100% and 50% respectively compared to common circular holes and scoop-shaped irregular film cooling holes, under the same amount of cool air. The cool air utilization rate is high, meaning that less cool air is needed to achieve the same film cooling coverage effect.

[0020] In addition, the turbine blade provided by the present invention is based on aircraft-type film air holes arranged in an alternating manner on the blade, which have a good film air coverage effect within a range of 5 times the diameter D1 of the downstream circular body. Attached Figure Description

[0021] Figure 1 This is a plan view of an aircraft provided in an embodiment of the present invention;

[0022] Figure 2 A two-dimensional diagram of the air film pore structure provided in an embodiment of the present invention;

[0023] Figure 3 This is a schematic diagram showing the positional relationship between the aircraft-type film cooling structure and the blade surface provided in an embodiment of the present invention;

[0024] Figure 4 A plan view of a single aircraft-type air film vent provided in some embodiments of the present invention;

[0025] Figure 5 This is a diagram illustrating the arrangement of a multi-exhaust film perforation structure on a turbine blade.

[0026] Figure 6 This is a downstream streamline diagram of the aircraft-type film air hole provided in an embodiment of the present invention;

[0027] Figure 7 This is a comparison cloud diagram of the outlet streamlines of an aircraft-type air film orifice and a circular orifice provided in an embodiment of the present invention;

[0028] Figure 8 These are cloud maps comparing the air film coverage effects of different shaped air film pore structures. Detailed Implementation

[0029] The present invention will be further described in detail below with reference to the accompanying drawings and embodiments. However, it should be noted that these embodiments are not intended to limit the present invention. Equivalent changes or substitutions in function, method, or structure made by those skilled in the art based on these embodiments are all within the protection scope of the present invention.

[0030] In the description of the embodiments of the present invention, it should be understood that the terms "center", "longitudinal", "lateral", "up", "down", "front", "back", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", etc., indicate the orientation or positional relationship based on the orientation or positional relationship shown in the accompanying drawings. They are only for the convenience of describing the invention 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 limiting the invention.

[0031] Furthermore, the terms "first," "second," "third," etc., are used for descriptive purposes only and should not be construed as indicating or implying relative importance or implicitly specifying the number of technical features indicated. Thus, a feature defined with "first," "second," etc., may explicitly or implicitly include one or more of that feature. In the description of this invention, unless otherwise stated, "a plurality of" means two or more.

[0032] The terms "installation," "connection," and "linking" should be interpreted broadly. For example, they can refer to fixed connections, detachable connections, or integral connections; they can refer to mechanical connections or electrical connections; they can refer to direct connections or indirect connections through an intermediate medium; and they can refer to the internal connection between two components. Those skilled in the art will understand the specific meaning of these terms in this invention based on the specific circumstances.

[0033] like Figure 1-8 As shown, this embodiment provides an aircraft-type film cooling structure for an aero-engine. The cooling structure includes an aircraft-type film cooling core, which comprises a cylinder 2 and a cone 1 fixedly connected together. The outlet shape of the cone 1 is aircraft-shaped, and the aircraft shape includes a first arc segment 11, a first straight segment 14, a second arc segment 12, a second straight segment 17, a third straight segment 19, a fourth straight segment 18, a fifth straight segment 16, a third arc segment 13, and a sixth straight segment 15 connected in sequence.

[0034] The intersection point G of the third straight line segment 19 and the fourth straight line segment 18 is located on the same vertical line as the center O1 of the first arc segment 11, and the airplane shape is symmetrical about the axis.

[0035] In some embodiments, the height of the cylinder 2 is L1, and the height of the cone 1 is L2, then L2 = (L1 + L2) * K, where K is a constant, K = 0.4 to 0.6. Figure 2 The distance from section 05 to section 04 is the height of cylinder 2, and the distance from section 04 to section 01 is the height of cone 1.

[0036] In some embodiments, the included angle between the first straight segment 14 and the sixth straight segment 15 is α, and the included angle formed between the dot O3 of the second arc segment 12, the dot O2 of the third arc segment 13, and the dot O1 of the first arc segment 11 is δ, wherein δ > α. δ > α ensures the expansion effect of the air film outlet.

[0037] In some embodiments, the included angle between the first straight line segment 14 and the second straight line segment 17 is β, where β = 25°-35°; based on the fact that the airplane-shaped structure is a left-right symmetrical structure, the included angle between the fifth straight line segment 16 and the sixth straight line segment 15 is also β;

[0038] The angle between the third line segment 19 and the fourth line segment 18 is γ, where γ = 130° - 155°.

[0039] In some embodiments, the diameter of the cylinder 2 is D1, and the distance between the center O1 of the first arc segment 11 and the intersection of the third straight line segment 19 and the fourth straight line segment 18 is L5, then L5 = D1.

[0040] In some embodiments, the intersection of the third straight line segment 19 and the fourth straight line segment 18 is set as point G, and the cross section between the cylinder and the cone is set as cross section 04. Then, the angle between the shortest straight line from point G to cross section 04 and the center line of the cylinder is the forward tilt angle θ of the cone, where θ = 9°-12°.

[0041] In some embodiments, a turbine blade is provided, the turbine blade including a blade body and a plurality of aircraft-type film cooling structures disposed on the blade body, the aircraft-type film cooling structures being any of the aforementioned aircraft-type film cooling structures.

[0042] The angle between the central axis of the aircraft-type film cooling structure and the blade surface is ε, where ε = 35°-50°.

[0043] The opening direction of the air film vent in the conical structure of the aircraft-type air film vent cooling structure is oriented downstream of the airflow.

[0044] In some embodiments, the hole spacing of the aircraft-type film cooling structure ranges from 5 to 10 times D1.

[0045] In some embodiments, the fabrication steps of the turbine blade with an aircraft-style film cooling structure are as follows:

[0046] Step 1: Draw a circle (O1) with diameter D1 = 0.5 mm. Establish a circle O2 with center D4 = 0.75 mm away from center O1. Draw a circle (O2) with diameter D2 = 0.25 mm. Establish a circle O3 with center D5 = 0.75 mm away from center O1. Draw a circle (O3) with diameter D3 = 0.25 mm. Make the included angle δ = 110° between the three circles.

[0047] Step 2: Draw line segments 14 and 15 through the circular tangent points A, B, C, and D; then draw line segments 16 and 17 through the circular tangent points E and F at β = 25°.

[0048] Step 3: Establish point G at a distance L6 = 0.5 mm from the center O1. Draw straight line segments 18 and 19 with γ = 135°, intersecting with straight line segments 16 and 17 to establish points H and I. Finally, subtract the curved portions of circles (O1), (O2), and (O3) contained within the aperture area. The remaining curves 11-19 constitute the outlet shape of the aircraft-type air film aperture of this invention, as shown below. Figure 1 As shown.

[0049] Step four: Based on the completed planar shape 10 of the aircraft-type air film aperture, establish the center O4 of the cylindrical segment at a distance L2 = 0.6 mm from the center point O1 of circle (1), and draw a circle 21 with a diameter of 0.5 mm using this as the center point; connect the planar shape 10 of the aircraft-type air film aperture and circle 21 through a curve group to form a cone 1; then stretch circle (21) axially to a length L1 = 0.9 mm to section 22 to form a cylinder 2; cone 1 and cylinder 2 together constitute the unit body of the aircraft-type air film aperture of this invention, as shown in the figure. Figure 2 As shown.

[0050] Step 5: Place the aircraft-style film gas vent at an angle of ε = 35° to the blade surface, such as... Figure 3 As shown, the difference between the blade model and the aircraft-style film cooling hole core is used to obtain the aircraft-style film cooling hole pattern on the blade, as shown. Figure 4 As shown.

[0051] Step six: Arrange the aircraft-style film cooling pores obtained in step five at a longitudinal spacing of 3.0 mm and a transverse spacing of 3.0 mm to 4.5 mm, ultimately obtaining multiple rows of aircraft-style film cooling pores on the blade surface. For example... Figure 5 .

[0052] In another embodiment, the fabrication steps of the turbine blade with an aircraft-style film cooling structure are as follows:

[0053] Step 1: Draw a circle (O1) with diameter D1 = 0.73 mm. Establish a circle O2 with center D4 = 1.2 mm away from center O1. Draw a circle (O2) with diameter D2 = 0.3 mm. Establish a circle O3 with center D5 = 1.2 mm away from center O1. Draw a circle (O3) with diameter D3 = 0.3 mm. Make the included angle δ = 114° between the three circles.

[0054] Step 2: Draw line segments 14 and 15 through the circular tangent points A, B, C, and D; then draw line segments 16 and 17 through the circular tangent points E and F at β = 30°.

[0055] Step 3: Establish point G at a distance L6 = 0.73 mm from the center O1. Draw straight line segments 18 and 19 with γ = 140°, intersecting with straight line segments 16 and 17 to establish points H and I. Finally, subtract the curved portions of circles (O1), (O2), and (O3) contained within the aperture area. The remaining curves 11-19 constitute the outlet shape of the aircraft-type air film aperture of this invention, as shown below. Figure 1 As shown.

[0056] Step four: Based on the completed planar shape 10 of the aircraft-type air film aperture, establish the center O4 of the cylindrical segment at a distance L2 = 0.7 mm from the center point O1 of the circle (1), and draw a circle (21) with a diameter of 0.73 mm using this as the center point; connect the planar shape 10 of the aircraft-type air film aperture and the circle 21 through a curve group to form a cone (1); then stretch the circle 21 axially to a length L1 = 0.8 mm to the cross section 22 to form a cylinder 2; the cone 1 and the cylinder 2 together constitute the unit body of the aircraft-type air film aperture of the present invention, as shown in the figure. Figure 2 As shown.

[0057] Step 5: Place the aircraft-style film gas vent at an angle of ε = 35° to the blade surface, such as... Figure 3 As shown, the difference between the blade model and the aircraft-style film cooling hole core is used to obtain the aircraft-style film cooling hole pattern on the blade, as shown. Figure 4 As shown.

[0058] Step six: Arrange the aircraft-style film cooling pores obtained in step five at a longitudinal spacing of 4.38 mm and a transverse spacing of 4.38 mm to 5.84 mm, ultimately obtaining multiple rows of aircraft-style film cooling pores on the blade surface. For example... Figure 5 .

[0059] like Figure 6As shown, the aircraft-type film membrane structure provided in the aforementioned embodiment produces a slight combustion gas intrusion effect upstream of the film membrane outlet. The advantage of this is that the cold air ejected from the film membrane can be quickly compressed to the wall by the mainstream combustion gas, thereby forming a good wall coverage. At the same time, the structure formed by straight segments 18 and 19 can also effectively suppress the generation and development of the reverse kidney vortex at the outlet of the film membrane, further improving the cooling results downstream of the film membrane.

[0060] like Figure 7 As shown, the aircraft-type film membrane structure provided in the aforementioned embodiment is flatter than the reverse kidney-shaped vortex formed at the outlet of a conventional circular hole, confirming that the aircraft-type film membrane has better wall coverage. At the same time, due to the presence of the structure formed by the arc segment 13, the straight segments 15 and 16, the aircraft-type film membrane also forms two additional vortices on both sides of the reverse kidney-shaped vortex, allowing the cold air to develop better along the spanwise direction, thereby increasing the spanwise coverage area of ​​the film membrane.

[0061] like Figure 8 As shown in the figure, from top to bottom, the air film efficiency generated by a circular hole, a scoop-shaped irregular air film hole, and an aircraft-shaped air film hole under the same amount of cooling air is illustrated. As can be seen from the figure, the aircraft-shaped air film hole structure proposed in this embodiment of the invention produces a higher air film efficiency. Compared with the common circular hole and scoop-shaped irregular air film holes, the air film efficiency is increased by approximately 100% and 50%, respectively, under the same amount of cooling air. Furthermore, the aircraft-shaped air film hole structure proposed in this invention has a high utilization rate of cooling air, meaning that less cooling air is needed to achieve the same air film coverage effect.

[0062] The above description is merely a preferred embodiment of the present invention and is not intended to limit the present invention. Any modifications, equivalent substitutions, and improvements made within the spirit and principles of the present invention should be included within the protection scope of the present invention. The above description is only a preferred embodiment of the present invention. It should be noted that for those skilled in the art, several improvements and modifications can be made without departing from the technical principles of the present invention, and these improvements and modifications should also be considered within the protection scope of the present invention.

Claims

1. An aircraft-type film cooling structure for aero engines, characterized in that, The cooling structure includes an aircraft-type film cooling core, which comprises a cylinder and a cone fixedly connected together. The outlet shape of the cone is aircraft-shaped, and the aircraft shape includes a first arc segment, a first straight segment, a second arc segment, a second straight segment, a third straight segment, a fourth straight segment, a fifth straight segment, a third arc segment, and a sixth straight segment connected in sequence. The intersection of the third and fourth straight line segments lies on the same vertical line as the center of the first arc segment. The aircraft shape is symmetrical about its axis. If the height of the cylinder is L1 and the height of the cone is L2, then L2 = (L1 + L2). K is a constant, K = 0.4 to 0.

6. The angle between the first line segment and the second line segment is β, where β = 25° - 35°; The angle between the third and fourth line segments is γ, where γ = 130° - 155°. The diameter of the cylinder is D1, and the distance between the center of the first arc segment and the intersection of the third and fourth straight line segments is L5. Therefore, L5 = D1. The intersection of the third and fourth straight line segments is designated as point G, and the cross section between the cylinder and the cone is designated as section 04. The angle between the shortest straight line from point G to section 04 and the center line of the cylinder is the forward tilt angle θ of the cone, where θ = 9°-12°.

2. The aircraft-type film cooling structure for aero engines according to claim 1, characterized in that, The angle between the first and sixth straight line segments is α, and the angle between the dot of the second arc segment, the dot of the third arc segment, and the dot of the first arc segment is δ, wherein δ > α.

3. A turbine blade, characterized in that, The turbine blade includes a blade body and a plurality of aircraft-type film cooling structures disposed on the blade body, wherein the aircraft-type film cooling structure is an aircraft-type film cooling structure for an aero-engine as described in any one of claims 1-2. The angle between the central axis of the aircraft-type film cooling structure and the surface of the blade is ε, where ε = 35°-50°. The air film vent cooling structure of the aircraft type has its vent openings oriented downstream of the airflow.

4. The turbine blade according to claim 3, characterized in that, The hole spacing of the aircraft-type film cooling structure is in the range of 5-10 times D1.

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