An aircraft engine turbine rear support plate blade and turbine rear casing

Abstract

An aircraft engine turbine rear support plate blade and a turbine rear casing, wherein the rear support plate blade includes a blade body, wherein each cross section on the blade body has a zero angle of attack; radially, the chord length and the installation angle of each cross section on the blade body are first increased and then decreased; each cross section is linearly stacked along the line connecting the centers of the leading edge arcs. The turbine rear casing includes a casing inner ring and a casing outer ring, and the aircraft engine turbine rear support plate blade is arranged between the casing inner ring and the casing outer ring. In the present invention, the middle portion of the trailing edge of the blade body forms a local reverse bend shape that bends toward the back side of the blade, i.e., a "quasi-reverse bend blade" is formed. The local reverse bend can effectively control the secondary flow near the back side of the support plate, overcome the large adverse pressure gradient caused by the large expansion of the channel, suppress or eliminate the large separation flow, and improve performance.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft engine blades, and in particular to an aircraft engine turbine rear support plate blade and a turbine rear casing provided with the aircraft engine turbine rear support plate blade. Background Art

[0002] The turbine assembly is one of the key components of afterburner aircraft engines and has a significant impact on the performance of the entire engine. The turbine rear support plate and rear casing are important components of the turbine assembly. Their main aerodynamic function is to deflect the deflected airflow to the axial direction, creating reasonable inlet airflow conditions for the afterburner combustion chamber and achieving efficient operation of the engine. In a mixed exhaust afterburner turbofan engine, the turbine rear support plate is located between the low-pressure turbine rotor and the afterburner combustion chamber. During the aerodynamic design, while considering the angle of the airflow at the outlet of the low-pressure turbine rotor, it is also necessary to consider the low-speed and axial requirements of the afterburner combustion chamber for its inlet airflow. Therefore, the aerodynamic design of the turbine rear support plate is an important part of the turbine assembly design. Especially as modern aircraft engines develop towards a high thrust-to-weight ratio, an important technical route is to increase the stage load of the turbine, thereby reducing the number of turbine stages or blades and thus improving the thrust-to-weight ratio. However, this also creates a new problem, namely, the angle at which the low-pressure turbine rotor outlet airflow deviates from the axial direction will continue to increase. In the second and third generation aircraft, the low-pressure turbine outlet airflow deviates from the axial direction by about 10° to 20°, and after the load increases, it will reach 30° to 50°, presenting a high residual swirl flow state (airflow angle above 30°), which brings greater difficulty to the low-loss aerodynamic design of the turbine rear support plate.

[0003] In the prior art, a patent application with publication number CN112031879A discloses a turbine rear support plate blade and an aircraft engine thereof, wherein a large support plate is used to ensure the assembly of lubricating oil pipelines and load-bearing structures, and a number of small support plates are installed behind the large support plate to achieve a rectifying effect, thereby solving the technical problem that the engine turbine rear support plate is heavy and easily causes airflow blockage.

[0004] For example, the publication number CN114542207A discloses a method for designing the outer profile of a turbine rear casing support plate. By setting the leading and trailing edge small circle diameters of the support plate, the leading edge adopts a larger small circle diameter to accommodate a wider incoming flow direction, and the trailing edge adopts a moderate small circle diameter to ensure the thickness of the trailing edge, which can both meet casting requirements and minimize trailing edge flow losses. By setting the angle between the straight line connecting the center of the leading edge small circle and the center of the trailing edge small circle and the horizontal line of the profile, and setting the support plate inlet structural angle and outlet structural angle to form the center arc line of the support plate, the main airflow before and after the rear casing support plate can meet the airflow turning requirements of a certain angle, the guiding effect is enhanced, which is beneficial to flow loss control, and is beneficial to the axial exhaust design behind the support plate when the incoming flow deviates greatly from the axial direction, and can also reduce the risk of flow separation.

[0005] However, when the turbine rear support plate operates in a high swirl inflow environment, the engine requires a small number of support plates, light weight, and low overall ring consistency. The lubricating oil pipeline, support structure, and test structure pass through the interior, and a certain blade thickness must be guaranteed. In addition, the afterburner combustion chamber of the afterburner turbofan engine requires a lower inflow Mach number, which forces the turbine rear support plate meridian channel to be designed in an expansion form. Under the conditions of high swirl inflow and large channel expansion, large-scale flow separation is very likely to occur on the back side of the turbine rear support plate, resulting in serious losses, reducing engine thrust, and affecting the performance of the entire machine. Under the conditions of high swirl inflow and large channel expansion, the support plate structure and blade design methods in the above-mentioned existing patent documents can no longer achieve the purpose of realizing low-loss rectification design with a limited number of support plates, and there is no disclosure on how to solve the problem of support plate separation flow in high swirl inflow and large expansion channels. Summary of the Invention

[0006] The main purpose of the present invention is to propose an aircraft engine turbine rear support plate blade and a turbine rear casing to solve the problems of high residual swirl incoming flow and support plate separation flow in a large expansion channel, reduce flow losses, provide a reasonable airflow state for the afterburner, and meet the requirements of high thrust and low fuel consumption of the entire engine.

[0007] To achieve the above-mentioned objectives, on the one hand, the present invention proposes an aircraft engine turbine rear support plate blade, comprising a blade body, wherein each cross section on the blade body has a zero angle of attack; along the radial direction of the blade body, the blade chord length and the installation angle of each cross section on the blade body are first increased and then decreased; and each cross section is linearly stacked according to the line connecting the centers of the leading edge arcs.

[0008] Preferably, the radii of the leading edge arcs of each cross section of the blade body are equal, and the line connecting the centers of the leading edge arcs of each cross section is a straight line.

[0009] Preferably, the radii of the trailing edge arcs of each cross section of the blade body are equal.

[0010] Preferably, the blade body has a tip section, a mid-section and a root section in sequence in the radial direction; the leading edge structural angle of each cross section on the blade body is the same as the incoming airflow angle at the corresponding upstream position.

[0011] Furthermore, the leading edge structural angle of the blade tip section is 66°, the leading edge structural angle of the blade mid-section is 58°, and the leading edge structural angle of the blade root section is 57°.

[0012] Preferably, the blade chord length of the blade root section is the same as the blade chord length of the blade tip section, and the blade chord length of the blade tip section is preferably 1.11 to 1.15 times the blade chord length of the blade root section or the blade chord length of the blade tip section.

[0013] Furthermore, the blade chord lengths of the blade tip section and the blade root section are 130 mm, and the blade chord length of the blade mid-section is 147 mm.

[0014] Furthermore, the installation angle of the blade tip section is 99°, the installation angle of the blade mid-section is 100°, and the installation angle of the blade root section is 97°.

[0015] The aircraft engine turbine rear support blade according to claim 1, wherein the blade profile angle of each cross section of the blade body is not less than 20 degrees.

[0016] The present invention also provides a turbine rear casing, comprising an inner casing ring and an outer casing ring, characterized in that the aforementioned aircraft engine turbine rear support blades are disposed between the inner and outer casing rings, and that 12 blade bodies are evenly distributed throughout the turbine rear casing. One end of each blade body is plugged into the outer casing ring, and the other end is plugged into the inner casing ring. An outer sleeve is welded between the outer casing ring and the blade body, and an inner sleeve is welded between the inner casing ring and the blade body.

[0017] Due to the adoption of the above technical solution, the beneficial effects of the present invention are as follows:

[0018] (1) In the present invention, each cross section of the blade body adopts a zero angle of attack design, and along the radial direction, the blade chord length and installation angle of each cross section are in a shape that first increases and then decreases, and each cross section is linearly stacked according to the center line of the leading edge arc, and the center line of the leading edge arc of the blade is a straight line; the angle between the suction surface and the end wall at the trailing edge of the blade body is an acute angle, and the angle between the pressure surface and the end wall is an obtuse angle, which together form a local reverse bend shape in the middle of the trailing edge of the blade body that bends toward the back side of the blade, that is, a "quasi-reverse bend blade" is formed. The local reverse bend can effectively control the secondary flow near the back side of the support plate, overcome the large adverse pressure gradient caused by the large expansion of the channel, inhibit or eliminate the large separation flow, and improve performance.

[0019] (2) In the present invention, the blade body is defined by three cross sections: root, middle and tip. The leading edge points are stacked in a straight line. The leading edge of each cross section adopts a zero angle of attack design according to the direction of the incoming airflow. By adjusting the relative sizes of the chord lengths and installation angles of the three cross sections of the blade, the chord length and installation angle of the middle part are made larger. After stacking, the blade presents a three-dimensional shape in which the leading edge is a straight blade, the pressure surface of the trailing edge is concave toward the suction surface, and the trailing edge in the middle of the blade is convex backward. This shape generates radial secondary flow on the back side of the trailing edge of the support plate, sucking the fluid near the root and tip to the middle part and rectifying it, which can effectively control flow separation. BRIEF DESCRIPTION OF THE DRAWINGS

[0020] In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the following briefly introduces the drawings required for use in the embodiments or the description of the prior art. Obviously, the drawings described below are only some embodiments of the present invention. For ordinary technicians in this field, other drawings can be obtained based on the structures shown in these drawings without paying any creative work.

[0021] Figure 1 is a schematic cross-sectional view of the blade body;

[0022] Figure 2 It is the three-dimensional view of the blade body;

[0023] Figure 3 Schematic diagram of the geometric parameters of the blade section on the blade body;

[0024] Figure 4 This is a schematic diagram of the turbine rear casing assembly;

[0025] Figure 5 This is the front view of the turbine rear casing (from the leading edge to the trailing edge);

[0026] Figure 6 This is the rear view of the turbine rear casing (from trailing edge to leading edge);

[0027] Figure 7 This is the streamline diagram of the blade root section of the conventionally designed blade support plate;

[0028] Figure 8 This is the streamline diagram of the blade mid-section of the conventionally designed blade support plate;

[0029] Figure 9 This is the streamline diagram of the blade tip section of the conventionally designed blade support plate;

[0030] Figure 10 Streamline diagram of the blade root section of the blade body of the present invention;

[0031] Figure 11 Streamline diagram of the leaf midsection;

[0032] Figure 12 Tip cross-sectional streamline diagram.

[0033] Description of Figure Numbers:

[0034] 100, blade body;

[0035] 1. Blade tip cross section; 2. Blade mid cross section; 3. Blade root cross section;

[0036] 4. Support plate suction surface; 5. Support plate pressure surface; 6. Blade leading edge; 7. Blade trailing edge

[0037] 8. Forehead line; 9. Leading edge construction angle; 10. Mounting angle; 11. Blade chord length; 12. Blade bend angle;

[0038] 14. Receiver inner ring; 15. Receiver outer ring; 16. Inner bushing; 17. Outer bushing; 18. Gasket. DETAILED DESCRIPTION

[0039] The following will clearly and completely describe the technical solutions in the embodiments of the present invention in conjunction with the accompanying drawings. Obviously, the described embodiments are only part of the embodiments of the present invention, not all of the embodiments. All other embodiments obtained by ordinary technicians in this field based on the embodiments of the present invention without making any creative efforts shall fall within the scope of protection of the present invention.

[0040] It should be noted that all directional indications in the embodiments of the present invention (such as up, down, left, right, front, back, etc.) are only used to explain the relative position relationship, movement status, etc. between the various components under a certain specific posture (as shown in the accompanying drawings). If the specific posture changes, the directional indication will also change accordingly.

[0041] For the design of aero-engine turbine rear support plate blades, the afterburner combustion chamber of the afterburner turbofan engine requires a lower incoming flow Mach number, so the turbine rear support plate meridian channel has to be designed in an expansion form. The degree of expansion is usually defined by the equivalent expansion angle. The equivalent expansion angle is defined as: the inlet and outlet areas of the flow channel annular channel are equivalent to the two circular areas of a cone, and the axial length of the cone is consistent with the axial length of the transition flow channel. At this time, the angle formed by the waistline of the cone and the axis is the equivalent expansion angle of the flow channel. Generally, the equivalent expansion angle is designed to be around 9°, and a value higher than this value is defined as a large expansion angle.

[0042] Combine Figure 3 The figure shows the geometric parameters of the blade section on the blade body. In the blade design, the main geometric parameters include: forehead line 8, leading edge structural angle 9, installation angle 10, blade chord length 11, and blade bend angle 12.

[0043] Combine Figure 1 、 Figure 2The present invention shows an aircraft engine turbine rear strut blade, comprising a blade body 100. Each cross section of the blade body 100 has a zero angle of attack. Along the radial direction of the blade body 100, the chord length and installation angle of each cross section of the blade body 100 first increase and then decrease. Each cross section is linearly stacked along a line connecting the centers of the leading edge arcs. This structure creates a localized recurved shape in the middle of the trailing edge of the blade body 100, curving toward the back of the blade, forming a "quasi-recurved blade." This localized recurvature effectively controls secondary flow near the back of the strut, overcomes the large adverse pressure gradient generated by the large channel expansion, suppresses or eliminates large separation flows, and improves performance.

[0044] Preferably, the radii of the leading edge arcs of each cross section of the blade body 100 are equal, and the line connecting the centers of the leading edge arcs of each cross section is a straight line. The radii of the trailing edge arcs of each cross section of the blade body 100 are equal. When designing the blade body 100, the midline thickness method is used to perform parametric modeling design on the blade body 100. Based on the blade primitive level design, the blade profiles of each section from the root to the tip are modeled separately, and two-dimensional calculations are performed to verify whether the load distribution on the blade profile surface meets the expected requirements. Subsequently, a three-dimensional blade is stacked in a leading edge center stacking manner. Then, with the installation angle and chord length of each section blade profile as optimization variables, and the total pressure recovery coefficient and outlet airflow angle as optimization targets, a three-dimensional cycle iteration from modeling design to three-dimensional flow field calculation is performed on the three-dimensional blade until satisfactory aerodynamic performance is obtained. The final design is determined by combining geometric constraint adjustment, typical operating condition performance verification, and strength design verification.

[0045] Combine Figure 1 As shown, in the radial direction of the blade body 100, there are a blade tip section 1, a blade mid-section 2 and a blade root section 3 in sequence; the leading edge structural angle of each cross section on the blade body 100 is the same as the incoming airflow angle at the corresponding upstream position; the leading edge structural angle of the blade tip section 1 is 66°, the leading edge structural angle of the blade mid-section 2 is 58°, and the leading edge structural angle of the blade root section 3 is 57°.

[0046] Furthermore, the blade profile chord length of the blade root section 3 is the same as the blade profile chord length of the blade tip section 1, and the blade profile chord length of the blade tip section 1 is preferably 1.11 to 1.15 times the blade profile chord length of the blade root section 3 or the blade profile chord length of the blade tip section 1. Specifically, the blade profile chord lengths of the blade tip section 1 and the blade root section 3 are 130 mm, and the blade profile chord length of the blade mid-section 2 is 147 mm. The installation angle of the blade tip section 1 is 99°, the installation angle of the blade mid-section 2 is 100°, and the installation angle of the blade root section 3 is 97°. The blade profile bending angle of each cross section of the blade body 100 is not less than 20 degrees.

[0047] Through the above parameter design, it is verified by numerical simulation that under the conditions of high residual swirl flow (30° away from the axial direction) and large expansion channel (equivalent expansion angle of 16°, the equivalent expansion angle of conventional design is about 9°), the total pressure recovery coefficient of the conventional design turbine rear equal cross-section straight large and small support plate blades is 0.95, the average outlet airflow angle is 16° (with the axial direction), and the maximum outlet airflow angle is 25° (with the axial direction); while the turbine rear support plate blades proposed by the present invention achieve a total pressure recovery coefficient of 0.98, an average outlet airflow angle of 3° (with the axial direction), and a maximum outlet airflow angle of 9° (with the axial direction), and the total pressure recovery coefficient is improved by 3% relative to the conventional design. Figures 7 to 12 shown. Figures 7 to 9 The velocity streamline diagram of the existing conventional design shows that a large separation flow occurs under the conditions of high residual swirl flow and large channel expansion. Figures 10 and 11 This is a velocity streamline diagram of the blade support plate of the present invention. It can be seen from the figure that no large separation flow occurs under the conditions of high residual swirl incoming flow and large channel expansion.

[0048] On the other hand, this embodiment further provides a turbine rear casing, comprising a casing inner ring 14 and a casing outer ring 15, characterized in that the above-mentioned aircraft engine turbine rear strut blades are disposed between the casing inner ring 14 and the casing outer ring 15. A plurality of gaskets 18 are disposed on the casing outer ring 15.

[0049] Specifically, 12 blade bodies 100 are evenly distributed on the turbine rear casing.

[0050] The blade basin side of the blade body 100 is the support plate pressure surface 5, and the blade back side is the support plate suction surface 4. Since the blade body 100 has a local curved shape at the trailing edge, the angle between the support plate suction surface 4 and the wall surfaces of the inner ring 14 and the outer ring 15 of the casing is an acute angle, and the angle between the support plate pressure surface 5 and the wall surfaces of the inner ring 14 and the outer ring 15 of the casing is an obtuse angle.

[0051] Furthermore, one end of the blade body 100 is plugged into the outer ring 15 of the casing, and the other end is plugged into the inner ring 14 of the casing. An outer sleeve 17 is provided between the outer ring 15 of the casing and the blade body 100 and welded; an inner sleeve 16 is provided between the inner ring 14 of the casing and the blade body 100 and welded.

[0052] The above description is only a preferred embodiment of the present invention and does not limit the patent scope of the present invention. Any equivalent structural transformation made by using the contents of the present invention specification and drawings under the inventive concept of the present invention, or directly / indirectly applied in other related technical fields, is included in the patent protection scope of the present invention.

Claims

1. An aircraft engine turbine rear support blade, comprising a blade body (100), characterized in that: Each cross section on the blade body (100) has a zero attack angle; Along the radial direction of the blade body (100), the blade chord length and the installation angle of each cross section on the blade body (100) are in a shape that first increases and then decreases; each cross section is linearly stacked along a line connecting the centers of the leading edge arcs; The radii of the leading edge arcs of each cross section of the blade body (100) are equal, and the line connecting the centers of the leading edge arcs of each cross section is a straight line; The radii of the trailing edge arcs of each cross section of the blade body (100) are equal; The blade body (100) has a blade tip section (1), a blade mid-section (2), and a blade root section (3) in sequence in the radial direction; the leading edge structural angle of each cross section on the blade body (100) is the same as the incoming airflow angle at the corresponding upstream position; The leading edge structural angle of the blade tip section (1) is 66°, the leading edge structural angle of the blade mid-section (2) is 58°, and the leading edge structural angle of the blade root section (3) is 57°; The blade chord length of the blade root section (3) is the same as the blade chord length of the blade tip section (1); The blade tip section (1) and the blade root section (3) have a blade profile chord length of 130 mm, and the blade mid-section (2) has a blade profile chord length of 147 mm; The installation angle of the blade tip section (1) is 99°, the installation angle of the blade mid-section (2) is 100°, and the installation angle of the blade root section (3) is 97°; The blade profile angle of each cross section of the blade body (100) is not less than 20 degrees.

2. A turbine rear casing, comprising a casing inner ring (14) and a casing outer ring (15), characterized in that: The aircraft engine turbine rear support blade according to claim 1 is arranged between the casing inner ring (14) and the casing outer ring (15), and 12 blade bodies (100) are evenly distributed on the entire ring of the turbine rear casing; one end of the blade body (100) is plugged into the casing outer ring (15), and the other end is plugged into the casing inner ring (14); an outer sleeve (17) is arranged between the casing outer ring (15) and the blade body (100) and is welded; an inner sleeve (16) is arranged between the casing inner ring (14) and the blade body (100) and is welded.

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