阅读说明：本技术 用于提高低压涡轮叶片气动效率的结构及其工作方法 (Structure for improving aerodynamic efficiency of low-pressure turbine blade and working method thereof ) 是由 饶宇 谢胤 程宇立 于 2021-10-15 设计创作，主要内容包括：本发明提供了一种用于提高低压涡轮叶片气动效率的结构,包括：吸力面、压力面、凹陷和叶片本体,吸力面为叶片本体的外凸面,压力面为叶片本体的内凹面；凹陷成对地设置在吸力面上,凹陷与气流之间具有倾角β；当气流流过叶片本体的表面时,凹陷的一端吸附低能流体,并使低能流体在凹陷内沿倾斜方向螺旋流动形成螺旋涡流,并从凹陷的另一端排出；其中,气流包括低能流体和高能流体。本发明通过涡轮叶片表面上的倾斜凹陷内部产生螺旋形移动的涡流,在凹陷的下游段上产生了高强度和大范围流动附着,延迟了叶片本体后部壁面上的流动分离,使其获得更好的涡轮叶片减阻效果。(The present invention provides a structure for improving aerodynamic efficiency of a low-pressure turbine blade, comprising: the blade comprises a suction surface, a pressure surface, a recess and a blade body, wherein the suction surface is an outer convex surface of the blade body, and the pressure surface is an inner concave surface of the blade body; the depressions are arranged in pairs on the suction surface, and an inclination angle beta is formed between the depressions and the airflow; when the airflow flows through the surface of the blade body, one end of the recess adsorbs low-energy fluid, and the low-energy fluid spirally flows in the recess along the oblique direction to form a spiral vortex and is discharged from the other end of the recess; wherein the gas stream comprises a low energy fluid and a high energy fluid. The invention generates spiral moving vortex in the inclined recess on the surface of the turbine blade, generates high-strength and large-range flow adhesion on the downstream section of the recess, delays flow separation on the wall surface of the rear part of the blade body, and obtains better turbine blade drag reduction effect.)

1. A structure for improving aerodynamic efficiency of low pressure turbine blades, comprising: the blade comprises a suction surface (10), a pressure surface (11), a recess (20) and a blade body, wherein the suction surface (10) is an outer convex surface of the blade body, and the pressure surface (11) is an inner concave surface of the blade body;

the recesses (20) being arranged in pairs on the suction surface (10), the recesses (20) having an angle of inclination β with respect to the air flow;

when the air flow flows through the surface of the blade body, one end of the recess (20) adsorbs low-energy fluid, and the low-energy fluid spirally flows in the recess (2) along the inclined direction to form a spiral vortex and is discharged from the other end of the recess (20); wherein the gas stream comprises a low energy fluid and a high energy fluid.

2. The structure for improving the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 1, wherein the depression (20) is provided at a set suction surface flow separation position near a position on the curved blade body wall surface with a chord length of 50% -90%.

3. The structure for improving the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 2, characterized in that the depression (20) is provided aft of 50% of the blade body chord length.

4. The structure for increasing the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 1, characterized in that said recess (20) comprises an upstream section (21) and a downstream section (23), said upstream section (21) being a half-sphere with a diameter D2; the downstream section (23) is a half spherical surface with a diameter D1; wherein the diameter D1 is greater than or equal to D2.

5. The structure for improving the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 4, characterized in that the recess (20) further comprises a middle section (25), the middle section (25) being a smoothly transitioning cylindrical or conical surface, the middle section (25) gradually increasing in diameter from the upstream section (21) to the downstream section (23).

6. The structure for improving the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 1, characterized in that the inclination angle β between the recess (20) and the air flow is 0-90 degrees.

7. The structure for improving the aerodynamic efficiency of low-pressure turbine blades as claimed in claim 4, characterized in that the narrowness L/D1 of the recess (20) is between 1 and 10, where L denotes the distance between the centers of the circles of the upstream (21) and downstream (23) sections and D1 denotes the diameter of the downstream section.

8. Structure for improving the aerodynamic efficiency of low-pressure turbine blades according to claim 4, characterized in that the depth ratio h1/D1, h2/D2 of the recess (20) is between 0 and 0.2, where h2 is the depth of the upstream section (21) and h1 is the depth of the downstream section (23).

9. The structure for improving the aerodynamic efficiency of low-pressure turbine blades according to claim 4, characterized in that the edges of the upstream section (21) are rounded and the edges of the downstream section (23) are rounded.

10. A method of operating a turbine blade, characterized in that a helical vortex is generated by means of the structure for improving the aerodynamic efficiency of a low pressure turbine blade according to any one of claims 1 to 9, the flow attachment of the air flow is generated in the downstream section (23) of the recess (20), the flow separation on the suction surface (10) is delayed and the drag is reduced.

Technical Field

The invention relates to the technical field of power blades of aero-engines, in particular to a structure for improving the aerodynamic efficiency of a low-pressure turbine blade and a working method thereof.

Background

Gas turbines are important components in aircraft engines. When the aircraft engine operates, the gas turbine takes high-temperature and high-pressure gas as a working medium, and the gas turbine generates mechanical work by utilizing the expansion of the working medium, and has the function of converting the heat energy of the gas into the mechanical work.

After the high-temperature and high-pressure gas flows through the channel between the turbine blades, the temperature and the pressure of the gas flow are reduced, and the conversion from the internal energy of the gas to the kinetic energy and then to the mechanical energy is realized; the gas flow process produces a force interaction with the turbine blades, which perform mechanical work externally. In turbofan aircraft engines, the low pressure turbine outputs work to drive the fan of the turbofan engine, which drives a large flow of air through the engine and generates the primary engine thrust. Therefore, the operating efficiency and aerodynamic performance of the low pressure turbine have a significant impact on engine performance.

The turbine blade cascade refers to a blade assembly formed by grouping static blades or moving blades in a turbine, and the inner arc and back arc molded surfaces and the upper and lower end walls of the adjacent static blades or moving blades form a through-flow path, namely a through-flow part, of gas. When the gas flows through the stationary blade grid channel, the heat energy is converted into kinetic energy, and when the gas flows through the movable blade grid channel, part of the heat energy is converted into the kinetic energy, and meanwhile, the kinetic energy of the gas is converted into mechanical work. When the gas turbine works, high-temperature and high-pressure gas expands and accelerates through the static blade channels of the through-flow part of the gas turbine and flows out in a certain direction, and then continues to expand in the moving blade channels and converts the kinetic energy of the gas into mechanical work.

Through search, patent document CN104314618A discloses a low-pressure turbine blade structure and a method for reducing blade loss, including a blade leading edge, a blade suction side, a blade pressure side and a blade trailing edge, wherein a rough belt is disposed on the surface of the blade suction side, and the starting and ending positions of the rough belt are determined according to the two-dimensional profile of the middle part of the blade height. The roughness of the surface of the blade is increased at the upstream of the separation point of the suction surface of the blade, and the roughness gradually changes along with the flow direction, so that the low-energy fluid on the surface of the blade is accelerated to transition, the working efficiency of the low-pressure turbine is improved, and the working margin of the low-pressure turbine is enlarged. However, the method brings extra flow loss on the suction surface of the blade under the condition of high Reynolds number, and cannot improve the aerodynamic efficiency of the low-pressure turbine.

Patent document CN112177680A discloses a high-pressure turbine blade structure with an array of drag-reducing dimples for reducing flow losses by arranging drag-reducing dimples at the mid-chord position of the suction surface of the high-pressure turbine blade, and a trailing edge surface for blade suction surface flow separation control. However, the prior art has the disadvantages that the turbine blade with the concave part can only be suitable for the working condition that the main flow separation occurs at the concave position; and for the obvious change (advance or delay) of the flow separation position on the surface of the blade under the variable working condition running of the actual engine, the flow control drag reduction effect of the expansion-shaped recess is limited, and even the opposite drag increase effect is realized under the main flow condition of high Reynolds number. The main reason is that the interior of the expanded recess on the wall surface of the turbine blade generates strong backflow vortex, the expanded recess increases the area of the backflow vortex, the backflow vortex scours the front edge of the expanded recess and rolls upwards to be mixed with the main flow to a large extent, the aerodynamic loss of the main flow is remarkably increased, and the backflow vortex in the recess consumes additional flow energy, so that the aerodynamic drag reduction effect of the blade is limited. Under the working condition of high Reynolds number, the position of the concave arranged on the surface of the blade does not generate large-area flow separation, but generates flow separation and backflow vortex inside the concave, the concave generates additional obvious flow loss, and the working applicable Reynolds number range is narrow.

One trend in the development of modern aircraft engines is that the development of turbine blades is moving towards high loads, and therefore blade bending is increasing. The flow separation is easy to occur on the back surface-suction surface of the turbine blade, particularly under the low Reynolds number flow condition, the Reynolds number is 5000-. The low pressure turbine low reynolds number operating condition may occur in a small turbofan aircraft engine, and the turbofan aircraft engine operates at high altitude.

Therefore, there is a need to develop a turbine blade structure that eliminates or reduces flow separation when the suction surface of the high-load low-pressure turbine blade operates at a low reynolds number, improves the aerodynamic efficiency of the high-load low-pressure turbine, and does not increase aerodynamic loss at a high reynolds number.

Disclosure of Invention

Aiming at the defects in the prior art, the invention aims to provide a structure for improving the aerodynamic efficiency of a low-pressure turbine blade and a working method thereof, which can overcome the adverse pressure gradient at the rear part of the suction surface of the turbine blade, inhibit flow separation or delay the flow separation of the suction surface of the turbine blade, improve the aerodynamic efficiency of the turbine blade under the condition of low Reynolds number and enlarge the stable working range of the turbine.

According to the present invention, there is provided a structure for improving aerodynamic efficiency of a low-pressure turbine blade, comprising: the blade comprises a suction surface, a pressure surface, a recess and a blade body, wherein the suction surface is an outer convex surface of the blade body, and the pressure surface is an inner concave surface of the blade body; the depressions are arranged in pairs on the suction surface, and an inclination angle beta is formed between the depressions and the airflow; when the airflow flows through the surface of the blade body, one end of the recess adsorbs low-energy fluid, and the low-energy fluid spirally flows in the recess along the oblique direction to form a spiral vortex and is discharged from the other end of the recess; wherein the gas stream comprises a low energy fluid and a high energy fluid.

Preferably, the recess is provided at a set suction surface flow separation location, which is near a location on the curved blade body wall surface where 50% -90% of the chord length is located.

Preferably, the recess is provided aft of 50% of the chord length of the blade body.

Preferably, the recess comprises an upstream section and a downstream section, the upstream section being one-half spherical and having a diameter D2; the downstream section is a spherical surface of one half, and the diameter is D1; wherein the diameter D1 is greater than or equal to D2.

Preferably, the recess further comprises a middle section which is a smooth transition cylindrical or conical surface, and the diameter of the middle section gradually increases from the upstream section to the downstream section.

Preferably, the inclination angle β between the recess and the air flow is 0-90 degrees.

Preferably, the narrowness L/D1 of the recess is between 1-10, where L denotes the distance between the centers of the upstream and downstream segments and D1 denotes the diameter of the downstream segment.

Preferably, the depth ratios h1/D1, h2/D2 of the depressions are each between 0 and 0.2, where h2 is the depth of the upstream segment and h1 is the depth of the downstream segment.

Preferably, the edges of the upstream section are rounded and the edges of the downstream section are rounded.

According to the working method of the turbine blade, the spiral vortex is generated by the turbine blade structure, the airflow generates flow adhesion at the downstream section of the recess, the flow separation on the suction surface is delayed, and the resistance is reduced.

Compared with the prior art, the invention has the following beneficial effects:

1. the invention generates spiral moving vortex in the inclined recess on the surface of the turbine blade, generates high-strength and large-range flow adhesion on the wall surface of the downstream blade of the recess, delays flow separation on the wall surface of the rear part of the blade body, and obtains better resistance reduction effect of the turbine blade.

2. According to the invention, through the design of the inclined recesses on the surface of the turbine blade, the chord length of the covered blade is longer, and under the condition of high Reynolds number, the flow separation position on the surface of the blade is delayed, so that the turbine blade has better flow control and drag reduction effects.

3. The invention solves the problem of low aerodynamic efficiency of the turbine blade under the condition of low Reynolds number, does not increase the flow resistance of the turbine blade under the condition of high Reynolds number, and enlarges the high-efficiency stable working range of the turbine.

Drawings

Other features, objects and advantages of the invention will become more apparent upon reading of the detailed description of non-limiting embodiments with reference to the following drawings:

FIG. 1 is a schematic view of the overall structure of the present invention;

fig. 2 is a cross-sectional view of the present invention.

Detailed Description

The present invention will be described in detail with reference to specific examples. The following examples will assist those skilled in the art in further understanding the invention, but are not intended to limit the invention in any way. It should be noted that it would be obvious to those skilled in the art that various changes and modifications can be made without departing from the spirit of the invention. All falling within the scope of the present invention.

As shown in fig. 1 and 2, the present invention provides a turbine blade structure including: the blade comprises a suction surface 10, a pressure surface 11, a recess 20 and a blade body, wherein the suction surface 10 is an outer convex surface of the blade body, and the pressure surface 11 is an inner concave surface of the blade body; the recesses 20 are arranged in pairs on the suction surface 10, the recesses 20 having an inclination β with respect to the air flow; when the air current flows through the surface of the blade body, one end of the recess 20 adsorbs low-energy fluid, and the low-energy fluid spirally flows in the recess 2 along the inclined direction to form a spiral vortex and is discharged from the other end of the recess 20; wherein the gas stream comprises a low energy fluid and a high energy fluid. Moreover, when the air flow passes through the recess 20 of the suction surface 10, because the shear stress on the suction surface 10 is reduced, the fluid above the suction surface 10 is accelerated and attached to the downstream suction surface of the recess 20, and the flow energy of the downstream boundary layer is increased; in addition, the spiral direction of the internal vortex of the recess 20 is consistent with the direction of the high-speed main flow above the suction surface, and the spiral vortex brings the upper main flow to the vicinity of the suction surface, so that the flow kinetic energy of the near suction surface in the region is improved, and the transition of the near suction surface is promoted.

Wherein the recess 20 is provided at a set suction surface flow separation position, which is a position near a curved blade body wall surface. The recess 20 comprises an upstream section 21, a downstream section 23 and a middle section 25, the upstream section 21 being one-half spherical with a diameter D2; the downstream section 23 is a one-half spherical surface with a diameter of D1; wherein the diameter D1 is greater than or equal to D2. The intermediate section 25 is a smooth transition cylindrical or conical surface, and the diameter of the intermediate section 25 gradually increases from the upstream section 21 to the downstream section 23. The angle of inclination beta between the recess 20 and the gas stream is 0-90 degrees.

The narrowness L/D1 of the recess 20 is between 1-10, where L denotes the distance between the centers of the upstream and downstream sections 21 and 23, and D1 denotes the diameter of the downstream section.

The depth ratios h1/D1, h2/D2 of the recess 20 are all between 0 and 0.2, where h2 is the depth of the upstream segment 21 and h1 is the depth of the downstream segment 23.

Preferred examples of the present invention will be further described.

Based on the above embodiments, the recess 20 in the present invention is provided behind 50% of the chord length of the blade body.

Based on the above embodiments, the inclination angle β between the recess 20 and the air flow in the present invention is 30 to 60 degrees.

Based on the above embodiments, the effect of the recess of the present invention is better when the narrowness L/D1 of the recess 20 is greater than 3.

Based on the above embodiment, the depth ratio of the recess 20 in the present invention is preferably 0.05 to 0.2, and the recess depth ratio varies such that the recess depth becomes shallower from downstream to upstream. The downstream concave portion is deeper with a concave depth h1 of 0-0.2 to diameter D1, while the upstream concave portion is shallower with a concave depth h2 of 0-0.2 to diameter D2.

The invention eliminates or reduces flow separation when the suction surface of the high-load low-pressure turbine blade works under the low Reynolds number, thereby improving the pneumatic efficiency of the high-load low-pressure turbine; and the low-pressure turbine blade does not increase aerodynamic loss under the condition of high Reynolds number, so that the working range of the turbine engine is expanded. The flow on the vane surface interacts with the inclined depression wall surface so that the low energy fluid at the near suction surface starts moving in a spiral inside the downstream section 23 of the depression 20 and is discharged from the other end 21 of the depression; the spiral vortex flow generated in the inclined recess of the present invention can be continuously discharged and draws the high energy fluid above the wall surface to adhere to the rear suction surface, which has a significant flow control advantage over other types of recesses in which the vortex flow resides.

Based on the above embodiment, the diameter D1 of the downstream section 23 is 2 times the diameter D2 of the upstream section 21.

Based on the above embodiment, the rounding of the edges of the upstream section 21 and the rounding of the edges of the downstream section 23 facilitates reducing flow losses of the upstream fluid adhering to the trailing edge of the recess, while facilitating helical vortex shedding inside the recess.

Based on the above embodiment, the dimples 20 are arranged in V-shaped pairs with the apex pointing upstream or downstream.

The invention also provides a working method of the turbine blade, which utilizes the turbine blade structure to generate spiral vortex, airflow generates flow adhesion on the downstream section 23 of the recess 20, flow separation on the suction surface 10 is delayed, and resistance is reduced.

The position of flow separation at the suction surface of the turbine blade is actually changed due to changes in the Reynolds number of the flow and the parameters of the incoming flow during actual turbine operation. Generally, when the Reynolds number of the incoming flow is lower, the flow separation position of the suction surface of the blade is closer to the upstream wall surface of the blade; when the incoming flow reynolds number is higher, the flow separation is closer to the downstream wall of the vane. The inclined recess provided by the invention can adapt to the change of the flow separation position on the surface of the blade in a wider range, and has a wider effective working range for inhibiting the flow separation.

The inclined depression on the suction surface of the turbine blade blocks the influence of the downstream flow separation or the adverse pressure gradient of the blade on the upstream flow, so that the surface flow of the blade on the upstream of the depression is less in flow separation, and the resistance reduction is facilitated.

The inclined recess 20 can generate a spiral vortex in the recess, the vortex reduces the shearing force of the external main flow and draws the external high-energy fluid to the wall surface of the blade, and the kinetic energy of the near-wall fluid is improved. The rotating direction of the vortex is consistent with the direction of the speed of the external main flow, and the shearing stress is reduced, so that the effect of accelerating the external flow close to the wall surface is achieved.

The downstream section 23 of the inclined recess 20 is larger and deeper, which is beneficial to introducing more low-energy fluid near the suction surface into the recess 20, and is beneficial to generating stronger interaction between the high-speed main flow above the suction surface and the wall surface of the recess, thereby generating stronger spiral vortex in the recess 20.

The narrower and shallower upstream section 21 of the inclined recess 20 is beneficial to reduce flow separation inside the recess 20 and to facilitate the vortex inside the recess 20 to flow out of the upstream and be carried away by the upstream energetic fluid. In addition, under the condition of high Reynolds number, when the upstream section 21 of the recess 20 is not subjected to flow separation, the upstream section 21 of the recess 20 does not bring extra flow loss, so that the adaptation range of the invention for realizing aerodynamic drag reduction of the turbine is enlarged.

In the description of the present application, it is to be understood that the terms "upper", "lower", "front", "rear", "left", "right", "vertical", "horizontal", "top", "bottom", "inner", "outer", and the like indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience in describing the present application and simplifying the description, but do not indicate or imply that the referred device or element must have a specific orientation, be constructed in a specific orientation, and be operated, and thus, should not be construed as limiting the present application.

The foregoing description of specific embodiments of the present invention has been presented. It is to be understood that the present invention is not limited to the specific embodiments described above, and that various changes or modifications may be made by one skilled in the art within the scope of the appended claims without departing from the spirit of the invention. The embodiments and features of the embodiments of the present application may be combined with each other arbitrarily without conflict.