CN213298058U - Turbine blade cooling system and aircraft engine - Google Patents
Turbine blade cooling system and aircraft engine Download PDFInfo
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- CN213298058U CN213298058U CN202022044906.7U CN202022044906U CN213298058U CN 213298058 U CN213298058 U CN 213298058U CN 202022044906 U CN202022044906 U CN 202022044906U CN 213298058 U CN213298058 U CN 213298058U
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Abstract
The utility model relates to a turbine blade cooling system and aeroengine. Wherein the turbine blade cooling system comprises: a turbine case; a plurality of turbine blades disposed within the turbine case; a gas collecting chamber configured in a ring shape and disposed along an outer circumferential direction of the plurality of turbine blades; the air guide pipe is communicated with the air collection cavity and is configured to introduce cooling gas into the air collection cavity; and the first guide plate is arranged in the gas collection cavity, the first end of the first guide plate is connected with the edge of the outlet end of the gas guide pipe, the second end of the first guide plate extends to the direction of the central axis deviating from the outlet end of the gas guide pipe through the outlet end of the gas guide pipe, and the first guide plate is configured to guide the cooling gas discharged from the outlet end of the gas guide pipe to flow so that the cooling gas has a component which rotates and flows along the gas collection cavity and around the central axis of the gas collection cavity. The utility model provides a first guide plate leads the radial flow of air current for the circumferential motion, has eliminated the impact that causes the just turbine blade, has improved the cooling homogeneity.
Description
Technical Field
The utility model relates to an aerospace equipment field especially relates to a turbine blade cooling system and aeroengine.
Background
Turbines are important parts of aircraft engines and the like, and since the operating environment in which each part is located is at a high temperature during operation, cooling of the relevant part is required. For example, the blades are cooled to extend the life of the turbine and improve the operational reliability of the turbine.
In order to cool the turbine blades, a plurality of bleed ducts are generally provided in the turbine casing to guide cooling gas from a compressor or the like to the turbine blades distributed along the circumferential direction of the turbine casing, thereby cooling the turbine blades.
The inventors have found that the above cooling method has a problem of uneven cooling of the turbine blade.
Disclosure of Invention
Some embodiments of the utility model provide a turbine blade cooling system and aeroengine for alleviate the inhomogeneous problem of turbine blade cooling.
Some embodiments of the present invention provide a turbine blade cooling system, comprising:
a turbine case;
the turbine blades are arranged in the turbine casing and distributed along the circumferential direction of the turbine casing;
a gas collection chamber disposed within the turbine casing, the gas collection chamber configured in an annular shape and disposed along an outer circumference of the plurality of turbine blades, the cooling gas within the gas collection chamber configured to flow toward the turbine blades;
the bleed pipe is arranged outside the turbine casing and communicated with the gas collecting cavity, and the bleed pipe is configured to introduce cooling gas into the gas collecting cavity; and
the first guide plate is arranged in the gas collecting cavity, the first end of the first guide plate is connected with the edge of the outlet end of the gas introducing pipe, the second end of the first guide plate is extended towards the direction of the central axis deviated from the outlet end of the gas introducing pipe, and the first guide plate is configured to guide the flow of cooling gas discharged from the outlet end of the gas introducing pipe, so that the cooling gas has components which are in rotary flow along the gas collecting cavity and around the central axis of the gas collecting cavity.
In some embodiments, the turbine blade cooling system includes a second baffle disposed opposite the first baffle, the second baffle cooperating with the first baffle to form a flow passage for directing a flow of cooling gas.
In some embodiments, the first baffle comprises a straight plate or an arcuate plate.
In some embodiments, the angle a between the straight plate and the central axis of the outlet end of the bleed air duct is in the range 0 ° < a <90 °.
In some embodiments, the arcuate plate projects in a direction proximate a central axis of the turbine case.
In some embodiments, the arcuate plate projects away from a central axis of the turbine case.
In some embodiments, the area where the first flow guide plate is located corresponds to at least one turbine blade, and a vent hole is formed in the first flow guide plate at a position corresponding to the turbine blade, and the diameter of the vent hole is smaller than that of the outlet end of the bleed air pipe.
In some embodiments, the first baffle is provided with a vent hole at a position on the central axis of the outlet end of the bleed air pipe, and the diameter of the vent hole is smaller than that of the second end of the first baffle.
In some embodiments, at least two bleed ducts are provided around the turbine casing, each bleed duct being associated with a first baffle.
Some embodiments of the present invention provide an aircraft engine comprising a turbine blade cooling system as described above.
Based on the technical scheme, the utility model discloses following beneficial effect has at least:
in some embodiments, a first guide plate is disposed in the gas collecting cavity, a first end of the first guide plate is connected to an edge of the outlet end of the bleed air pipe, a second end of the first guide plate extends to a direction deviating from a central axis of the outlet end of the bleed air pipe through the outlet end of the bleed air pipe, and the first guide plate is configured to guide the flow of the cooling gas so that the cooling gas has a component that rotates along the gas collecting cavity and around the central axis of the gas collecting cavity, that is, the first guide plate guides the radial flow of the airflow jetted from the outlet end of the bleed air pipe to a circumferential motion, thereby eliminating impact of the radial flow of the airflow jetted from the bleed air pipe on the turbine blades facing each other and improving the uniformity of the.
Drawings
The accompanying drawings, which are included to provide a further understanding of the invention and are incorporated in and constitute a part of this application, illustrate embodiment(s) of the invention and together with the description serve to explain the invention without undue limitation to the invention. In the drawings:
FIG. 1 is a turbine blade cooling system and airflow schematic provided in accordance with some embodiments of the present invention;
FIG. 2 is a schematic cross-sectional view of a turbine air supply system provided in accordance with some embodiments of the present invention;
fig. 3 is a schematic cross-sectional view of a turbine air supply system provided in accordance with further embodiments of the present invention.
The reference numbers in the drawings illustrate the following:
1-a turbine case; 2-turbine blades; 3-a gas collection cavity; 4-a gas-guiding pipe; 41-a mounting seat; 5-a first baffle; 51-a vent hole; 6-second baffle.
Detailed Description
The technical solutions in the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings in the embodiments of the present invention. It is obvious that the described embodiments are only some of the embodiments of the present invention, and not all of them. Based on the embodiments of the present invention, all other embodiments obtained by a person of ordinary skill in the art without creative efforts belong to the protection scope of the present invention.
In the description of the present invention, it is to be understood that the terms "center", "longitudinal", "lateral", "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 of description and simplicity of description, and do not indicate or imply that the device or element so referred to must have a particular orientation, be constructed and operated in a particular orientation, and therefore, should not be taken as limiting the scope of the invention.
Referring to fig. 1-3, some embodiments provide a turbine blade cooling system that includes a turbine casing 1, a plurality of turbine blades 2, a plenum 3, and a bleed duct 4.
A plurality of turbine blades 2 are disposed within the turbine casing 1 and distributed along the circumferential direction of the turbine casing 1. Optionally, the turbine blades 2 comprise turbine vanes or fairing blades of a turbine interstage casing.
A gas collecting chamber 3 is provided in the turbine casing 1, the gas collecting chamber 3 is configured in an annular shape and is provided along the outer circumferential direction of the plurality of turbine blades 2, and the cooling gas in the gas collecting chamber 3 is configured to flow toward the turbine blades 2.
The bleed air pipe 4 is arranged outside the turbine casing 1 and is communicated with the gas collecting cavity 3. The bleed air duct 4 is configured to introduce cooling gas into the gas collecting chamber 3. Optionally, the outlet end of the bleed air duct 4 is provided to the turbine casing 1 by means of a mounting seat 41. Optionally, a bleed air duct 4 leads cooling gas from a compressor or the like into the gas collecting chamber 3.
Referring to fig. 2 and 3, four bleed air pipes 4 are arranged in the circumferential direction of the turbine casing 1, and cooling gas is introduced into the gas collection chamber 3 through the four bleed air pipes 4. The four bleed air ducts 4 can be arranged uniformly in the circumferential direction of the turbine casing 1. The circumferential direction of the turbine casing 1 is not limited to the provision of four bleed air ducts 4.
Twelve turbine blades 2 are circumferentially arranged inside the turbine casing 1, and the twelve turbine blades 2 can be uniformly arranged along the inner circumference of the turbine casing 1. However, the interior of the turbine casing 1 is not limited to being provided with twelve turbine blades 2 in the circumferential direction.
The aero-engine conducts air bleed from the blade tip of the air compressor, the led cooling gas enters the air bleed pipe 4 and is led into the gas collecting cavity 3 through the air bleed pipe 4, and the cooling gas in the gas collecting cavity 3 enters the turbine blades 2 to cool the turbine blades 2. The flow direction of the cooling gas is shown in fig. 1 to 3.
Referring to fig. 2 and 3, the central axis direction of the outlet end of the bleed air pipe 4 is consistent with the radial direction of the turbine casing 1, the cooling gas is introduced into the gas collecting cavity 2 from the outlet end of the bleed air pipe 4 along the radial direction of the turbine casing 1, and the turbine blades 2 are arranged at the downstream of the gas collecting cavity 2 along the flow direction of the cooling gas. The air collecting cavity 2 is used for collecting air flow discharged from the outlet end of the air guide pipe 4, and the air flow enters the downstream after being stabilized in the air collecting cavity 2 so as to cool the turbine blades 2.
Ideally, the dynamic pressure of the air flow in the air collecting cavity 2 is completely dissipated, and the air flow can uniformly enter the downstream. However, because the size of the gas collecting cavity 2 is limited, the circumferential number of the air guide pipes 4 is limited, and the airflow at the outlet ends of the air guide pipes 4 is in a jet state and cannot be completely dissipated; and the structures of the turbine blades 2 and the like at the downstream of the air collecting cavity 2 are also discrete in the circumferential direction, and the number of the turbine blades 2 is generally larger than that of the air guide pipes 4, so that the turbine blades 2 facing the outlet ends of the air guide pipes 4 have high airflow impact on the turbine blades 2 facing the outlet ends of the air guide pipes 4 with high probability, and therefore, the turbine blades 2 have the condition of uneven cooling flow distribution in the circumferential direction due to high pressure caused by the outlet jet flow of the air guide pipes 4.
The cooling air flow at the outlet end of the bleed air pipe 4 flows along the radial direction of the turbine casing 1, and impacts on the turbine blade 2 which is opposite to the turbine blade 2, and the flow rate of the turbine blade 2 which is opposite to the outlet end of the bleed air pipe 4 is far larger than that of the rest turbine blades 2, so that the cooling is uneven, and the turbine blade 2 deforms unevenly.
Based on this, the present disclosure provides a turbine blade cooling system for mitigating the problem of turbine blade cooling non-uniformity.
In some embodiments, a turbine blade cooling system includes a turbine casing 1, a plurality of turbine blades 2, a plenum 3, a bleed air duct 4, and a first baffle 5.
A plurality of turbine blades 2 are disposed within the turbine casing 1 and distributed along the circumferential direction of the turbine casing 1. A gas collecting chamber 3 is provided in the turbine casing 1, the gas collecting chamber 3 is configured in an annular shape and is provided along the outer circumferential direction of the plurality of turbine blades 2, and the cooling gas in the gas collecting chamber 3 is configured to flow toward the turbine blades 2. The outlet end of the bleed air duct 4 communicates with the gas collecting chamber 3 and is configured to introduce cooling gas into the gas collecting chamber 3.
The first guide plate 5 is arranged in the gas collecting cavity 3, the first end of the first guide plate 5 is connected with the edge of the outlet end of the gas guiding pipe 4, the second end of the first guide plate 5 extends towards the direction of the central axis deviating from the outlet end of the gas guiding pipe 4 through the outlet end of the gas guiding pipe 4, and the first guide plate 5 is configured to guide the cooling gas discharged from the outlet end of the gas guiding pipe 4 to flow, so that the cooling gas has a component which rotates and flows along the gas collecting cavity 3 and around the central axis of the gas collecting cavity 3.
Because the number of the bleed air pipes 4 of the engine is limited, the air flow enters the air collecting cavity 3 through the openings of the mounting seats 41 after passing through the bleed air pipes 4, and the outlet ends of the bleed air pipes 4 have a large radial jet velocity, so that the turbine blades 2 facing the air flow are impacted, the flow of the entering cold air is large, the impact pressure on the turbine blades 2 not facing the air flow is small, the flow of the cold air is small, the temperature field of each turbine blade 2 in the circumferential direction is not uniform due to the difference of the flow of the cold air, and further the turbine blades 2 and even the turbine casing 1 are not uniformly deformed. The related unevenness may cause a rub-and-rub condition between the rotors and the stators of the turbine blade 2, affecting the engine safety.
Based on this, set up first guide plate 5 in gas collecting cavity 3, first guide plate 5 begins the exit end through bleed pipe 4 from the exit end edge of bleed pipe 4, and extend to the direction of the axis of the exit end of skew bleed pipe 4, the cooling gas impact of bleed pipe 4's exit end is on first guide plate 5, flow under the guide effect of first guide plate 5, first guide plate 5 makes cooling gas have along gas collecting cavity 3 and do the component of rotatory flow around the axis of gas collecting cavity 3, that is to say, first guide plate 5 leads the radial flow direction of bleed pipe 4's exit end spun air current to circumferential motion, and dissipate this speed, the impact that the air current radial flow that has eliminated bleed pipe 4 and jetted the department caused to just right turbine blade 2, the inhomogeneous problem of turbine blade 2 cooling has been alleviated.
In some embodiments, the turbine blade cooling system includes a second baffle 6, the second baffle 6 is disposed opposite the first baffle 5, and the second baffle 6 cooperates with the first baffle 5 to form a flow channel for guiding the flow of the cooling gas.
Optionally, the first end of the first baffle 5 is connected to the radial edge of the outlet end of the bleed air pipe 4, the first end of the second baffle 6 is connected to the radial edge of the outlet end of the bleed air pipe 4, and the first end of the first baffle 5 and the first end of the second baffle 6 are located in the same radial direction of the outlet end of the bleed air pipe 4.
Optionally, the second baffle 6 is identical in structure and shape to the first baffle 5.
The cooling gas discharged from the outlet end of the air guide pipe 4 is guided by the cooperation of the second guide plate 6 and the first guide plate 5, the air flow direction deviates from the axis direction of the outlet end of the air guide pipe 4, and the circumferential rotating motion of the air collection cavity 3 is carried out, therefore, the jet impact on the turbine blades 2 at the outlet end of the air guide pipe 4 can be avoided, the air inlet pressure of the turbine blades 2 at the outlet end of the air guide pipe 4 is improved, the condition that the cold air flow of the turbine blades 2 at the outlet end of the air guide pipe 4 is obviously higher than that of other blades due to the impact effect of the radial jet flow at the outlet end of the air guide pipe 4 is relieved, the cooling nonuniformity of the circumferential turbine blades 2 is reduced, correspondingly, the nonuniformity of the circumferential deformation of the turbine box 1 can also.
In some embodiments, as shown in fig. 2, the first baffle 5 comprises a straight plate.
In some embodiments the angle a between the straight plate and the centre axis of the outlet end of the bleed air duct 4 is in the range 0 < a < 90.
In some embodiments, the width of the first flow guide plate 5 is equal to or greater than the diameter of the outlet end of the bleed air duct 4, the first flow guide plate 5 can completely cover the outlet end of the bleed air duct 4, and the air flow exiting the outlet end of the bleed air duct 4 impinges on the first flow guide plate 5 and then flows along the guide of the first flow guide plate 5.
A first flow deflector 5 is mounted on the inside of the turbine casing 1, the first flow deflector 5 having an angle a with the centre axis of the outlet end of the bleed air duct 4, the angle a being in the range 0 ° < a <90 °, preferably 45 ° < a <90 °.
The periphery of the turbine casing 1 is provided with a plurality of air guide pipes 4, each air guide pipe 4 is correspondingly provided with a first guide plate 5, the first guide plates 5 are arranged around the central axis of the turbine casing 1, the inclination angles of the first guide plates 5 in the circumferential direction of the turbine casing 1 are the same, and the first guide plates are distributed along the turbine casing 1 in a central symmetry manner.
After the airflow collides with the first guide plate 5 through the outlet end of the bleed air pipe 4, the airflow enters the air collecting cavity 3 along the direction forming an included angle a with the radial direction, the airflow does not directly impact the turbine blade 2 which is directly opposite to the outlet end of the bleed air pipe 4, the jet flow velocity has a larger component in the circumferential direction, so that the airflow in the air collecting cavity 3 obtains the circumferential velocity, and the airflow can be better conveyed to the turbine blade 2 away from the outlet end of the bleed air pipe 4 in the circumferential rotation process of the airflow; and because the main speed of the air flow is the circumferential speed, the inlet pressure obtained by each turbine blade 2 in the circumferential direction is basically equivalent, and the uniformity of the cooling flow of the turbine blades 2 is improved, so that the uniformity of a temperature field and deformation is improved, and the consequences of rotor and stator collision and abrasion and the like caused by the nonuniform deformation due to nonuniform cooling are prevented.
In some embodiments, as shown in fig. 3, the first baffle 5 comprises an arcuate plate.
In some embodiments, the second end of the first baffle 5 extends in a direction that coincides with a tangential direction of the turbine casing 1 at the corresponding location.
The inner side of the turbine casing 1 is provided with a first guide plate 5, the first guide plate 5 is an arc-shaped plate, the periphery of the turbine casing 1 is provided with a plurality of air guide pipes 4, each air guide pipe 4 is correspondingly provided with a first guide plate 5, the first guide plate 5 is arranged around the central axis of the turbine casing 1, airflow enters the air collection cavity 3 along the radian of the first guide plate 5 after impacting the first guide plate 5 through the outlet ends of the air guide pipes 4, the airflow does not directly impact the turbine blades 2 facing the outlet ends of the air guide pipes 4, the jet flow velocity has a larger component in the circumferential direction, so that the airflow in the air collection cavity 3 obtains a circumferential velocity, and the airflow can be better conveyed to the turbine blades 2 away from the outlet ends of the air guide pipes 4 in the circumferential rotation process of the airflow; and because the main speed of the air flow is the circumferential speed, the inlet pressure obtained by each turbine blade 2 in the circumferential direction is basically equivalent, and the uniformity of the cooling flow of the turbine blades 2 is improved, so that the uniformity of a temperature field and deformation is improved, and the consequences of rotor and stator collision and abrasion and the like caused by the nonuniform deformation due to nonuniform cooling are prevented.
In some embodiments, the arcuate plate projects in a direction proximate to the central axis of the turbine case 1.
In some embodiments, the arcuate plate projects away from the central axis of the turbine case 1.
In some embodiments, the area of the first flow guiding plate 5 corresponds to at least one turbine blade 2, a vent hole 51 is formed in the first flow guiding plate 5 at a position corresponding to the turbine blade 2, the width of the first flow guiding plate 5 is larger than the diameter of the outlet end of the bleed air pipe 4, and the diameter of the vent hole 51 is smaller than the diameter of the outlet end of the bleed air pipe 4.
In some embodiments, the first flow guiding plate 5 is provided with a vent hole 51 at a position on the central axis of the outlet end of the bleed air pipe 4, the width of the first flow guiding plate 5 is larger than the diameter of the outlet end of the bleed air pipe 4, and the diameter of the vent hole 51 is smaller than the diameter of the outlet end of the bleed air pipe 4.
Because the radial air flow at the outlet end of the bleed air pipe 4 is guided to the circumferential direction, the amount of cooling air obtained by the turbine blade 2 facing the outlet end of the bleed air pipe 4 in the radial direction is reduced, and in order to prevent uneven cooling of the turbine blade 2, a vent hole 51 is arranged on the first guide plate 5 at a position facing the turbine blade 2, and the vent hole 51 has the function of compensating the air flow. The cooling air in the first flow guiding plate 5 is partially discharged from the vent holes 51, and most of the remaining cooling air is discharged from the second end of the first flow guiding plate 5, so as to avoid the uneven phenomenon caused by the diversion of the air flow of the turbine blade 2 facing the first flow guiding plate 5.
In some embodiments, the width of the first baffle 5 is greater than the diameter of the outlet end of the bleed air duct 4 and the diameter of the vent hole 51 is less than the diameter of the outlet end of the bleed air duct 4. The part of the cooling air guided by the first guide plate 5 is discharged from the vent hole 51, and most of the rest of the cooling air flows around the central axis of the turbine casing 1 along the first guide plate 5, so as to avoid the uneven phenomenon caused by the diversion of the air flow of the turbine blade 2 facing the first guide plate 5.
The airflow radially and inwards impacts the first guide plate 5 at the outlet end of the air guide pipe 4, the airflow is guided into the circumferential direction by the first guide plate 5, the radially inward impact speed is completely eliminated, the airflow at the outlet end of the air guide pipe 4 does not directly face the turbine blade 2, and the airflow can be better conveyed to the turbine blade 2 far away from the outlet end of the air guide pipe 4; and because the main speed of the air flow is the circumferential speed, the inlet pressure obtained by each turbine blade 2 in the circumferential direction is basically equivalent, and the uniformity of the cooling flow of the turbine blades 2 is improved, so that the uniformity of a temperature field and deformation is improved, and the consequences of rotor and stator collision and abrasion and the like caused by the nonuniform deformation due to nonuniform cooling are prevented. Meanwhile, the flow guide effect of the arc-shaped plate is more obvious, and the gas turning loss is less.
In some embodiments, at least two bleed ducts 4 are provided around the turbine casing 1, each bleed duct 4 being provided with a first deflector plate 5. Each first guide plate 5 is correspondingly matched with one second guide plate 6 to form a flow channel for guiding the cooling gas.
As shown in fig. 2 and 3, in some embodiments, four bleed air pipes 4 are circumferentially arranged on the outer periphery of the turbine casing 2, outlet ends of the four bleed air pipes 4 are mounted on the turbine casing 2 through a mounting seat 41, twelve turbine blades 2 are circumferentially arranged inside the turbine casing 2, a gas collecting chamber 3 is circumferentially arranged on the outer periphery of the twelve turbine blades 2, and outlet ends of the four bleed air pipes 4 are communicated with the gas collecting chamber 3. In the gas collecting cavity 3, the outlet ends of the four gas guiding pipes 4 are respectively provided with a first guide plate 5. A turbine blade 2 is arranged on the central axis of the outlet end of each bleed air duct 4. The first guide plate 5 is provided with a vent hole 51 at the central axis of the outlet end of the bleed air pipe 4, and the flow area of the vent hole 51 is smaller than that of the outlet end of the bleed air pipe 4.
As shown in fig. 2, the first baffle 5 comprises a straight plate. The angle range of the included angle a between the straight plate and the central axis of the outlet end of the bleed air pipe 4 is 0 degrees < a <90 degrees.
As shown in fig. 3, the first baffle 5 comprises an arc-shaped plate. The arc-shaped plate projects in a direction close to the central axis of the turbine casing 1.
After the air flows through the outlet end of the air guide pipe 4, most of the flow is changed into circumferential flow through the flow guiding effect of the straight plate or the arc-shaped plate, and the turbine blade 2 directly opposite to the outlet end of the air guide pipe 4 is not directly impacted any more; a small portion of the airflow exits through the bleed holes 51 for cooling the turbine blades 2 opposite the outlet end of the bleed air duct 4.
Some embodiments provide an aircraft engine comprising the turbine blade cooling system described above.
In the description of the present invention, it should be understood that the terms "first", "second", "third", etc. are used to define the components, and are only used for the convenience of distinguishing the components, and if not stated otherwise, the terms have no special meaning, and thus, should not be construed as limiting the scope of the present invention.
Furthermore, the technical features of one embodiment may be combined with one or more other embodiments advantageously without explicit negatives.
Finally, it should be noted that the above embodiments are only used for illustrating the technical solutions of the present invention and not for limiting the same; although the present invention has been described in detail with reference to preferred embodiments, it should be understood by those skilled in the art that: the invention can be modified or equivalent substituted for some technical features; without departing from the spirit of the present invention, it should be understood that the scope of the claims is intended to cover all such modifications and variations.
Claims (10)
1. A turbine bucket cooling system, comprising:
a turbine casing (1);
the turbine blades (2) are arranged in the turbine casing (1) and distributed along the circumferential direction of the turbine casing (1);
a gas collection chamber (3) provided in the turbine casing (1), the gas collection chamber (3) being configured in an annular shape and arranged along an outer circumferential direction of the plurality of turbine blades (2), the cooling gas in the gas collection chamber (3) being configured to flow toward the turbine blades (2);
a bleed air pipe (4) arranged outside the turbine casing (1) and communicated with the gas collecting cavity (3), wherein the bleed air pipe (4) is configured to introduce cooling gas into the gas collecting cavity (3); and
first guide plate (5), locate in gas collection chamber (3), the first end of first guide plate (5) is connected the edge of the exit end of bleed pipe (4), the second end warp of first guide plate (5) the exit end of bleed pipe (4) is to deviating the axis direction of the exit end of bleed pipe (4) extends, first guide plate (5) are configured to the guide the exhaust cooling gas of exit end of bleed pipe (4) flows, makes cooling gas have the edge gas collection chamber (3) and centers on the axis of gas collection chamber (3) is the component of rotatory flow.
2. The turbine blade cooling system of claim 1 comprising a second baffle (6), the second baffle (6) being disposed opposite the first baffle (5), the second baffle (6) cooperating with the first baffle (5) to form a flow path for directing the flow of cooling gas.
3. The turbine blade cooling system of claim 1, characterized in that the first flow guide plate (5) comprises a straight plate or an arc-shaped plate.
4. The turbine blade cooling system according to claim 3, characterised in that the angle a between the straight plate and the centre axis of the outlet end of the bleed air duct (4) is in the range 0 ° < a <90 °.
5. The turbine blade cooling system of claim 3, characterized in that the arc-shaped plate is convex in a direction close to the central axis of the turbine casing (1).
6. The turbine blade cooling system of claim 3, characterized in that the arc-shaped plate projects in a direction away from the central axis of the turbine casing (1).
7. The turbine blade cooling system according to claim 1, characterized in that the area of the first flow deflector (5) corresponds to at least one turbine blade (2), and that the first flow deflector (5) is provided with a vent hole (51) at a position corresponding to the turbine blade (2), and the diameter of the vent hole (51) is smaller than the diameter of the outlet end of the bleed air duct (4).
8. The turbine blade cooling system according to claim 1, characterized in that the first flow deflector (5) is provided with a vent hole (51) at a position on the central axis of the outlet end of the bleed air duct (4), the diameter of the vent hole (51) being smaller than the diameter of the second end of the first flow deflector (5).
9. The turbine blade cooling system according to claim 1, characterized in that at least two bleed air ducts (4) are provided around the turbine casing (1), a first deflector plate (5) being associated with each bleed air duct (4).
10. An aircraft engine, characterized in that it comprises a turbine blade cooling system according to any one of claims 1 to 9.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022044906.7U CN213298058U (en) | 2020-09-17 | 2020-09-17 | Turbine blade cooling system and aircraft engine |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202022044906.7U CN213298058U (en) | 2020-09-17 | 2020-09-17 | Turbine blade cooling system and aircraft engine |
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| CN213298058U true CN213298058U (en) | 2021-05-28 |
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Cited By (2)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114508394A (en) * | 2021-12-29 | 2022-05-17 | 东方电气集团东方汽轮机有限公司 | Turbine steam extraction cavity structure |
| CN114542287A (en) * | 2022-02-17 | 2022-05-27 | 中国航发沈阳发动机研究所 | Air entraining structure for reducing circumferential temperature nonuniformity of casing wall surface |
-
2020
- 2020-09-17 CN CN202022044906.7U patent/CN213298058U/en active Active
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114508394A (en) * | 2021-12-29 | 2022-05-17 | 东方电气集团东方汽轮机有限公司 | Turbine steam extraction cavity structure |
| CN114508394B (en) * | 2021-12-29 | 2023-11-10 | 东方电气集团东方汽轮机有限公司 | Turbine steam extraction cavity structure |
| CN114542287A (en) * | 2022-02-17 | 2022-05-27 | 中国航发沈阳发动机研究所 | Air entraining structure for reducing circumferential temperature nonuniformity of casing wall surface |
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