Aero-engine turbine blade cooling channel rotational flow structure and method

Abstract

The invention relates to the technical field of aero-engine turbine blades, and discloses an aero-engine turbine blade cooling channel rotational flow structure and method, and the aero-engine turbine blade cooling channel rotational flow structure comprises an S-shaped flow channel located in a middle chord area of a turbine blade; the S-shaped runner comprises a plurality of circulating cavities which are communicated in sequence, and runner baffles are arranged among the circulating cavities; a spiral twisted piece is arranged in each circulation cavity, a guide vane is arranged above the flow channel baffle, the guide vane is of a C-shaped structure, an opening of the guide vane faces the flow channel baffle, and the two ends of the guide vane are connected with the ends of the spiral twisted pieces on the two adjacent sides respectively; the spiral twisted sheet comprises an inner wall surface in contact with fluid and an outer wall surface opposite to the inner wall surface; a plurality of arc-shaped protrusions are distributed on the inner wall face, and a plurality of arc-shaped pits are formed in the outer wall face. The contact distance between cooling gas and the wall face of the flow channel is prolonged, the flowing speed of the cooling gas in the channel is reduced, the cooling effect of the turbine blade is effectively improved, and then the overall performance of the turbine blade is improved.

Description

Technical Field

[0001] The present invention relates to the technical field of aircraft engine cooling, and in particular to a swirl structure and method for a cooling channel of an aircraft engine turbine blade. Background Art

[0002] In the design of aircraft engines and similar high-power density equipment, turbine blade cooling technology has always been a crucial research field. As a key component in the engine, turbine blades are subjected to extremely high temperatures and pressures, so their cooling effect directly affects the performance and life of the engine.

[0003] Traditional turbine blade cooling technology usually uses cold air impact to reduce the temperature of the blade by spraying cold air into the inside or surface of the blade. However, during the cold air impact process, due to the specific design of the flow channel structure, the flow of cold air in the flow channel often presents a faster speed. This high-speed flow state greatly reduces the contact time between the cold air and the flow channel surface, thereby reducing the heat exchange efficiency of the cold air, that is, the cooling effect.

[0004] Specifically, when the cold air passes through the mid-chord blade flow channel at a faster speed, it is difficult for the cold air to fully exchange heat with the flow channel wall due to the constraints of flow inertia and flow channel shape. This phenomenon not only weakens the cooling effect of the cold air on the blade, but may also cause flow instability in the flow channel, further affecting the working efficiency and safety of the engine.

[0005] In this regard, this program proposes a swirl structure and method for the cooling channel of an aircraft engine turbine blade, which prolongs the contact distance between the cooling gas and the channel wall, thereby making the contact time between the cooling gas and the wall longer, effectively improving the cooling effect of the turbine blade and improving the overall performance of the turbine blade. Summary of the invention

[0006] The present invention aims to provide a swirl structure and method for the cooling channel of an aircraft engine turbine blade. Without changing the flow channel structure of the turbine blade, the contact distance between the cooling gas and the flow channel wall is extended, thereby increasing the contact time between the cooling gas and the wall, reducing the flow velocity of the cooling gas, effectively improving the cooling effect of the turbine blade, and further improving the overall performance of the turbine blade.

[0007] To achieve the above-mentioned purpose, the present invention adopts the following technical scheme: a swirl structure of a cooling channel of an aircraft engine turbine blade, applied to the turbine blade or combustion chamber of an aircraft engine, including an S-shaped flow channel located in the chord area of ​​the turbine blade; the S-shaped flow channel includes a plurality of flow chambers connected in sequence, and a flow channel baffle is provided between each flow chamber; a spiral twisted piece is provided in each flow chamber along the flow direction of the fluid, and a guide blade is provided above the flow channel baffle, the guide blade is a C-shaped structure and the opening is facing the flow channel baffle, and the two ends of the guide blade are respectively connected to the ends of the spiral twisted piece on the adjacent two sides; the spiral twisted piece includes an inner wall surface in contact with the fluid, and an outer wall surface opposite to the inner wall surface; a plurality of arc-shaped protrusions are distributed on the inner wall surface, and a plurality of arc-shaped pits corresponding to the arc-shaped protrusions are provided on the outer wall surface.

[0008] At the same time, the present scheme also provides a swirl method for the cooling channel of an aircraft engine turbine blade, which is applied to the above-mentioned swirl structure of the cooling channel of an aircraft engine turbine blade, including arranging a spiral twisted blade in the flow cavity of the S-shaped flow channel of the vortex blade; arranging a guide vane at the end of the spiral twisted blade so that adjacent spiral twisted blades are connected through the guide vane; and evenly arranging a plurality of arc-shaped protrusions on the inner wall surface where the spiral twisted blade contacts the fluid, so that the fluid flows in a spiral in the spiral twisted blade and is disturbed by the arc-shaped protrusions.

[0009] The principles and advantages of this solution are:

[0010] In the cooling technology of turbine blades, the main methods used are plate cooling and convection-enhanced cooling of internal ribbed channels. Although plate cooling has high cooling efficiency and can reduce the consumption of cooling gas, it is mainly used in areas with large heat loads, and requires the addition of a large number of spoiler columns, which are separated into different plate units by partitions, making the structure complex. While the spoiler columns enhance heat exchange, they also increase the flow resistance in the channel, reducing the overall heat exchange efficiency of the turbine blades.

[0011] Arranging spoiler ribs on the suction side and pressure side of the channel can increase the heat exchange area in the channel. At the same time, the spoiler ribs have a disturbing effect on the incoming flow, destroying the development of the flow boundary layer and the thermal boundary layer, thereby enhancing the heat exchange efficiency. However, while the spoiler ribs enhance the heat exchange, they also increase the flow resistance in the channel. It is necessary to comprehensively analyze the cooling effect by studying the shape, angle, spacing and other factors of the spoiler ribs and combining them with the flow resistance characteristics. The design is difficult, there are many unstable factors, and it is very challenging and difficult to replicate.

[0012] This solution breaks through the existing cooling method of layer plates and spoiler ribs, and instead sets a spiral twisted blade in the circulation cavity to guide the cooling gas to flow in a spiral, which increases the contact distance between the cooling gas and the wall surface without increasing the flow resistance of the gas and reducing the pressure loss of the gas. At the same time, the spiral twisted blade is connected as a whole through the guide vane to reduce the complexity of the structure, and can also play a buffering role for the gas when turning, further improving the cooling effect. The structure is simpler, reducing the design difficulty, and effectively improving the cooling effect. BRIEF DESCRIPTION OF THE DRAWINGS

[0013] Figure 1 This is a schematic diagram of the structure of a swirl structure of a cooling channel of a turbine blade of an aero-engine according to the present invention;

[0014] Figure 2 A schematic diagram of the structure of an S-shaped flow channel of a swirl structure of a cooling channel of a turbine blade of an aero-engine according to the present invention;

[0015] Figure 3 The present invention is a schematic structural diagram of a flow cavity model of a swirl structure of a cooling channel of an aero-engine turbine blade. DETAILED DESCRIPTION

[0016] The following is further described in detail through specific implementation methods:

[0017] The reference numerals in the drawings of the specification include: S-shaped flow channel 1, flow cavity 2, spiral twisted sheet 3, guide vane 4, flow channel baffle 5, inner wall surface 6, outer wall surface 7, arc-shaped protrusion 8.

[0018] Embodiment 1

[0019] This embodiment is shown in the attached Figure 1 As shown, a swirl structure of the cooling channel of an aircraft engine turbine blade can be applied to the turbine blade or combustion chamber of an aircraft engine. Without changing the S-shaped flow channel structure of the turbine blade, a spiral twisting blade 3 and a guide vane 4 are added to enable the cooling gas to flow in a spiral shape, thereby extending the contact distance between the cooling gas and the flow channel wall surface, allowing the cooling gas to fully contact the flow channel wall surface, improving the heat exchange efficiency, improving the cooling effect of the turbine blade, and improving the engine performance and reliability. Specifically, it includes an S-shaped flow channel 1 located in the mid-chord area of ​​the turbine blade, wherein, as shown in the attached Figure 2 As shown, the S-shaped flow channel 1 includes a plurality of sequentially connected flow chambers 2. A flow channel baffle 5 is provided between the flow chambers 2, and the mid-chord area of ​​the turbine blade is divided into an S-shaped flow channel by the flow channel baffle 5. The cooling gas flows from the inlet to the outlet along the S-shaped flow channel to cool the turbine blade.

[0020] In this embodiment, for the convenience of description, a single flow chamber 2 is used as shown in the attached Figure 3 The model shown is represented.

[0021] Combined with Figure 3 As shown, a spiral twisted blade 3 is provided in the circulation cavity 2 along the fluid flow direction, and the width of both sides of the spiral twisted blade 3 is consistent with the width of the circulation cavity 2, so that the spiral twisted blade 3 can fill a single circulation cavity 2, thereby guiding the cooling gas to flow in a spiral along the spiral twisted blade 3 inside the channel as much as possible, increasing the contact distance between the cooling gas and the wall, and reducing the flow speed of the cooling gas.

[0022] In this embodiment, the number of periods of the spiral twisted blade 3 is set according to the length of the flow chamber 2, where the number of periods N = L / D, where L is the overall length of the flow chamber 2, and D is the distance between two adjacent spirals of the spiral twisted blade 3, and D is set according to the actual size of the turbine blade and the cooling flow velocity.

[0023] A guide vane 4 is provided above the flow channel baffle 5. In this embodiment, as shown in the attached Figure 3 As shown, the guide blade 4 is set to a C-shaped structure with its opening facing the flow channel baffle 5, so that the guide blade 4 is inverted on the flow channel baffle 5, and the front and rear ends of the guide blade 4 are respectively fixedly connected to the inner wall of the flow cavity 2, so that the guide blade 4 is stabilized above the flow channel baffle 5. At the same time, the C-shaped structure is conducive to the turning flow of the cooling gas, effectively reducing the impact pressure loss of the cooling gas, and further ensuring the cooling effect of the cooling gas.

[0024] In this embodiment, the two ends of the guide blade 4 are respectively connected to the ends of the spiral twisted blades 3 on the adjacent two sides, and the width of the guide blade 4 is consistent with the wall width of the spiral twisted blade 3, so that the spiral twisted blade 3 can be directly connected to the end of the guide blade 4, so that the spiral twisted blade 3 and the guide blade 4 are set as a whole, ensuring the symmetry and stability of the structure. The spiral twisted blades 3 on both sides are connected through the guide blade 4 to ensure the stability of the spiral twisted blade 3. At the same time, they are directly connected through the guide blade 4 to reduce additional connection structures and achieve the effect of guiding. In this embodiment, the thickness of the spiral twisted blade 3 is greater than 0mm and less than 1mm, and the thickness of the guide blade 4 is greater than 0mm and less than 1mm, so as to ensure that it can be installed in the circulation cavity without causing blockage to the fluid, while ensuring the guiding effect on the fluid to achieve a uniform guiding effect.

[0025] In this embodiment, as shown in the attached Figure 1 and attached Figure 3As shown, the spiral twisted blade 3 includes an inner wall surface 6 in contact with the fluid, and an outer wall surface 7 opposite to the inner wall surface 6. A plurality of arc-shaped protrusions 8 are distributed on the inner wall surface 6, and a plurality of arc-shaped pits corresponding to the arc-shaped protrusions 8 and concave inward are provided on the outer wall surface 7. In this embodiment, the arc-shaped protrusions 8 are arranged as a circular arc hemispherical structure. The arc-shaped pits and the arc-shaped protrusions 8 are arranged symmetrically on both sides. When the cooling gas flows, the gas is disturbed by the arc-shaped protrusions 8, thereby increasing the contact area between the cooling gas and the wall surface. At the same time, the flow velocity of the cooling gas can be reduced, the impact pressure loss can be reduced, and the cooling effect can be improved.

[0026] In this embodiment, as shown in the attached Figure 1 As shown, the rotation direction of the spiral lines of two adjacent spiral twisted sheets 3 is set in reverse. Specifically, the reverse setting is that according to the flow direction of the fluid in the circulation cavity 2, the rotation direction of the spiral twisted sheet 3 in the direction of the fluid inlet is clockwise, and the rotation direction of the spiral twisted sheet 3 in the direction of the outlet is counterclockwise, which helps to guide the heated air more effectively to the gap between the heat sinks. Driven by the reverse airflow, the hot air can be completely blown away from the heat sink, thereby achieving a better heat dissipation effect. In addition, the reverse flow can also make full use of the characteristics of air convection, allowing the hot air to flow along the curved surface of the heat sink, constantly mixing with the surrounding cold air, reducing the problems of insufficient mixing of hot air and cold air and slow flow speed, thereby achieving rapid cooling. The cooling gas can more fully cover all parts of the blades or heat sinks, thereby ensuring that all parts can be fully cooled. This uniformity helps to extend the service life of the blades or heat sinks and improve the reliability of the entire cooling system.

[0027] At the same time, the present scheme also provides a swirl method for the cooling channel of an aircraft engine turbine blade, which is applicable to the above-mentioned swirl structure of the cooling channel, specifically including setting a spiral twisted blade 3 in the flow cavity of the S-shaped flow channel of the vortex blade; setting a guide vane at the end of the spiral twisted blade 3, so that adjacent spiral twisted blades are connected through the guide vane; and evenly arranging a plurality of arc-shaped protrusions on the inner wall surface of the spiral twisted blade 3 in contact with the fluid, so that the fluid flows in a spiral in the spiral twisted blade and is disturbed by the arc-shaped protrusions. The rotation direction of the spiral twisted blade is set to the opposite direction or the same direction according to the flow direction of the fluid.

[0028] In this solution, a spiral twisted blade 3 is provided in the flow cavity 2 of the S-shaped flow channel 1, and is connected by a guide blade 4, so that the cooling gas can flow in a spiral along the spiral twisted blade 3 in the S-shaped flow channel 1, thereby increasing the contact distance between the cooling gas and the inner wall surface of the flow cavity 2. At the same time, the flowing gas is disturbed by the arc-shaped protrusion 8 on the spiral twisted blade 3, so as to further reduce the flow velocity of the flowing gas, increase the contact time of the gas, and improve the cooling effect. At the same time, a guide blade 4 is provided at the turning point of the flow cavity 2, and while connecting the spiral twisted blades 3 on both sides, the turning gas can be buffered to reduce the impact pressure loss of the cooling gas, reduce the complexity of the structure, and be more conducive to spatial arrangement, thereby ensuring the fluidity of the gas and improving the cooling effect of the cooling gas.

[0029] Embodiment 2

[0030] Different from the first embodiment, in this embodiment, the rotation directions of the spiral lines of two adjacent spiral twisted sheets 3 are set in the same direction. Specifically, the same direction is set according to the flow direction of the fluid in the circulation cavity 2, and the rotation directions of the spiral twisted sheets 3 in the fluid inlet direction and the spiral twisted sheets 3 in the outlet direction are both clockwise or counterclockwise, that is, counterclockwise. According to the flow direction of the cooling gas, the rotation reverse direction is reasonably set to ensure the fluidity of the cooling gas in the circulation cavity 2, reduce the flow difference between the channels, make the cooling gas more evenly distributed inside the blade, and help ensure that all parts of the blade can be fully cooled.

[0031] In this embodiment, the rotation direction of the spiral twisted blade 3 is set to ensure the fluidity of the cooling gas, and at the same time, the cooling gas is guided to flow in the same direction of the spiral twisted blade 3. The cooling gas rotating in the same direction can more effectively exchange heat with the blade wall. Due to the centrifugal force generated by the rotation, the cooling gas forms a thinner boundary layer near the blade wall, which is beneficial to the transfer of heat. At the same time, the rotating flow can also enhance the disturbance of the cooling gas and increase the heat transfer coefficient, thereby further improving the heat transfer efficiency. The cooling gas rotating in the same direction forms a relatively stable flow structure in the channel, reducing vortices and turbulence in the flow process, thereby reducing the flow resistance. This helps to reduce the energy consumption of the cooling system and improve the overall efficiency of the system.

[0032] The above are only embodiments of the present invention, and the common knowledge such as the known specific technical solutions and / or characteristics in the solution are not described in detail here. It should be pointed out that for those skilled in the art, several modifications and improvements can be made without departing from the technical solution of the present invention, which should also be regarded as the protection scope of the present invention, and these will not affect the effect of the implementation of the present invention and the practicality of the patent. The scope of protection required by this application shall be based on the content of its claims, and the specific implementation methods and other records in the specification can be used to interpret the content of the claims.

Claims

1. A swirl structure for cooling channel of turbine blades of an aircraft engine, characterized in that: The invention is applied to turbine blades or combustion chambers of aircraft engines, and includes an S-shaped flow channel located in the mid-chord area of ​​the turbine blade; the S-shaped flow channel includes a plurality of flow chambers connected in series, and a flow channel baffle is provided between each flow chamber; a spiral twisted sheet is provided in each flow chamber along the direction of fluid flow, and a guide blade is provided above the flow channel baffle, the guide blade is a C-shaped structure and its opening faces the flow channel baffle, and the two ends of the guide blade are respectively connected to the ends of the spiral twisted sheet on the adjacent two sides; the spiral twisted sheet includes an inner wall surface in contact with the fluid, and an outer wall surface opposite to the inner wall surface; a plurality of arc-shaped protrusions are distributed on the inner wall surface, and a plurality of arc-shaped pits corresponding to the arc-shaped protrusions are provided on the outer wall surface.

2. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 1, characterized in that: The rotation directions of the helical lines of two adjacent spiral twisted sheets are arranged in opposite directions.

3. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 2, characterized in that: The reverse setting is that, according to the flow direction of the fluid in the circulation cavity, the spiral twisted blades rotate in a clockwise direction in the fluid inlet direction, and the spiral twisted blades rotate in a counterclockwise direction in the outlet direction.

4. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 1, characterized in that: The rotation directions of the helical lines of two adjacent spiral twisted sheets are arranged in the same direction.

5. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 4, characterized in that: The same direction is set to that the rotation direction of the spiral twisted pieces is clockwise or counterclockwise.

6. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 1, characterized in that: The width of the guide vane is consistent with the wall width of the spiral twisted blade.

7. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 1, characterized in that: The thickness of the spiral twisted sheet is greater than 0 mm and less than 1 mm; the thickness of the guide blade is greater than 0 mm and less than 1 mm.

8. The swirl structure of the cooling channel of the turbine blade of an aircraft engine according to claim 1, characterized in that: The arc-shaped protrusion is an arc-shaped hemispherical structure.

9. A swirl method for cooling channel of a turbine blade of an aircraft engine, applicable to a swirl structure for cooling channel of a turbine blade of an aircraft engine as claimed in any one of claims 1 to 8, characterized in that: A spiral twisted blade is arranged in the flow cavity of the S-shaped flow channel of the vortex blade; a guide blade is arranged at the end of the spiral twisted blade so that adjacent spiral twisted blades are connected through the guide blade; a plurality of arc-shaped protrusions are evenly arranged on the inner wall surface of the spiral twisted blade in contact with the fluid so that the fluid flows in a spiral in the spiral twisted blade and is disturbed by the arc-shaped protrusions.

10. The swirl flow method for cooling passage of turbine blades of an aircraft engine according to claim 9, characterized in that: The rotation direction of the spiral twisted blades is set to opposite directions or the same direction according to the fluid flow direction.

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