CN213478401U - Turbine blade cooling system and aircraft engine - Google Patents
Turbine blade cooling system and aircraft engine Download PDFInfo
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- CN213478401U CN213478401U CN202022044908.6U CN202022044908U CN213478401U CN 213478401 U CN213478401 U CN 213478401U CN 202022044908 U CN202022044908 U CN 202022044908U CN 213478401 U CN213478401 U CN 213478401U
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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; 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, configured in an annular shape, 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 air guide pipe is communicated with the air collection cavity and is configured to introduce cooling gas into the air collection cavity; and the flow guide pipe is arranged in the gas collection cavity, the air inlet of the flow guide pipe is connected with the outlet end of the gas guide pipe, the air outlet of the flow guide pipe deviates from the outlet end of the gas guide pipe, and the flow guide pipe is configured to guide the flow of the cooling gas so that the cooling gas has a component which flows along the gas collection cavity and rotates around the central axis of the gas collection cavity. The utility model provides a honeycomb duct is circumferential motion with the radial flow direction of air current, 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;
a plurality of turbine blades disposed within the turbine casing and distributed along a circumferential direction of the turbine casing;
a gas collection chamber disposed within the turbine casing, the gas collection chamber configured in an annular shape disposed along an outer circumference of the plurality of turbine blades, 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 honeycomb duct is arranged in the gas collection cavity, the air inlet of the honeycomb duct is connected with the outlet end of the gas introduction pipe, the air outlet of the honeycomb duct is deviated from the outlet end of the gas introduction pipe, and the honeycomb duct is configured to guide cooling gas to flow so that the cooling gas has components which flow along the gas collection cavity and rotate around the central axis of the gas collection cavity.
In some embodiments, the draft tube is configured to extend in a direction offset from a central axis of the outlet end of the bleed air tube.
In some embodiments, the flow conduit comprises a straight tube or an arcuate tube.
In some embodiments, the angle a between the straight tube and the central axis of the outlet end of the bleed air tube is in the range 0 ° < a <90 °.
In some embodiments, the arced tube projects in a direction proximate to a central axis of the turbine case.
In some embodiments, the arced tube projects away from a central axis of the turbine case.
In some embodiments, the area of the draft tube corresponds to at least one turbine blade, and the position of the draft tube corresponding to the turbine blade is provided with a vent hole, and the diameter of the vent hole is smaller than that of the outlet end of the bleed air pipe.
In some embodiments, the position of the flow guide pipe on the central axis of the outlet end of the bleed air pipe is provided with a vent hole, and the diameter of the vent hole is smaller than that of the outlet end of the bleed air pipe.
In some embodiments, at least two bleed air ducts are provided around the turbine casing, each bleed air duct being connected to a flow guide duct.
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 flow guide pipe is arranged in the gas collecting cavity, the flow guide pipe is communicated with the outlet end of the gas guide pipe, and the flow guide pipe is configured to guide the flow of the cooling gas, so that the cooling gas has a component which makes a rotating flow along the gas collecting cavity and around the central axis of the gas collecting cavity, that is, the flow guide pipe guides the radial flow of the airflow jetted by the outlet end of the gas guide pipe to a circumferential motion and dissipates the speed, the impact of the radial flow of the airflow at the jetting position of the gas guide pipe on the opposite turbine blades is eliminated, and the uniformity of the cooling of the turbine.
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 flow guide pipe; 51-air vent.
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 the airflow of the air guide pipe 4, and the airflow enters the downstream after being subjected to certain pressure stabilization 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, the turbine blade cooling system includes a turbine case 1, a plurality of turbine blades 2, a plenum 3, a bleed air duct 4, and a draft tube 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 draft tube 5 is arranged in the gas collecting cavity 3, the air inlet of the draft tube 5 is connected with the outlet end of the gas guiding tube 4, the air outlet of the draft tube 5 deviates from the outlet end of the gas guiding tube 4, and the draft tube 5 is configured to guide the flow of cooling gas, so that the cooling gas has a component which flows along the gas collecting cavity 3 and rotates 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 honeycomb duct 5 in gas collecting cavity 3, honeycomb duct 5 intercommunication bleed pipe 4's exit end, honeycomb duct 5 is configured to guide the cooling gas flow, make the cooling gas have along gas collecting cavity 3 and around the axis of gas collecting cavity 3 do the component of rotatory flow, that is to say, honeycomb duct 5 leads the radial flow direction of the air current of bleed pipe 4's exit end spun for circumferential motion, and dissipate this speed, eliminated the impact that the turbine blade 2 that just faces is caused to the air current radial flow that bleed pipe 4 jetted department, alleviated turbine blade 2 and cooled inhomogeneous problem.
In some embodiments, the draft tube 5 is configured to extend in a direction offset from the central axis of the outlet end of the bleed air tube 4.
Guide the exhaust cooling gas of the exit end of bleed air pipe 4 through setting up honeycomb duct 5, make the air current flow direction deviate from the axis direction of the exit end of bleed air pipe 4, and along the circumferential direction rotary motion of gas collecting chamber 3, therefore, can avoid the efflux impact that just receives the turbine blade 2 of the exit end of bleed air pipe 4, improve the admission pressure of the turbine blade 2 of the exit end of keep away from bleed air pipe 4 simultaneously, alleviate the condition that the cold air flow that leads to just being higher than other blades to the turbine blade 2 of the exit end of bleed air pipe 4 because of the radial fluidic shock effect of the exit end of bleed air pipe 4 is showing, the refrigerated inhomogeneity of circumferential turbine blade 2 has been reduced, it is corresponding, the inhomogeneity of turbine casket 1 circumferential deformation also can be alleviated.
Optionally, the flow guiding tube 5 is a tube with uniform diameter along the direction from the air inlet to the air outlet.
In some embodiments, as shown in fig. 2, the draft tube 5 comprises a straight tube.
In some embodiments the angle a between the straight tube and the centre axis of the outlet end of the bleed air tube 4 is in the range 0 < a <90 °.
In some embodiments, the diameter of the draft tube 5 is equal to or greater than the diameter of the outlet end of the bleed air tube 4.
On the inside of the turbine casing 1 a draft tube 5 is mounted, which draft tube 5 has an angle a with the centre axis of the outlet end of the bleed air duct 4, which angle a is 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 connected with a flow guide pipe 5, the flow guide pipes 5 are arranged around the central axis of the turbine casing 1, the inclination angles of the flow guide pipes 5 in the circumferential direction of the turbine casing 1 are the same, and the flow guide pipes are distributed along the turbine casing 1 in a central symmetry manner.
After entering the draft tube 5 through the outlet end of the bleed air tube 4, the airflow enters the air collecting cavity 3 in the direction forming an included angle a with the radial direction, the airflow does not directly impact the turbine blade 2 which is opposite to the outlet end of the bleed air tube 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 tube 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 draft tube 5 comprises an arcuate tube.
In some embodiments, the direction of the outlet of the draft tube 5 coincides with the tangential direction of the turbine casing 1 at the corresponding position.
The turbine casing 1 is internally provided with a flow guide pipe 5, the flow guide pipe 5 is an arc-shaped pipe, the periphery of the turbine casing 1 is provided with a plurality of air guide pipes 4, each air guide pipe 4 is correspondingly connected with one flow guide pipe 5, the flow guide pipe 5 is arranged around the central axis of the turbine casing 1, airflow enters the flow guide pipe 5 through the outlet ends of the air guide pipes 4 and then enters the air collection cavity 3 according to the radian of the flow guide pipe 5, 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 large component in the circumferential direction, so that the airflow in the air collection cavity 3 obtains the circumferential velocity, and the airflow can be better conveyed to the turbine blades 2 away from the outlet ends of the air guide pipes; 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 arced tube projects in a direction proximate to the central axis of the turbine case 1.
In some embodiments, the arced tube projects away from the central axis of the turbine case 1.
In some embodiments, the area of the draft tube 5 corresponds to at least one turbine blade 2, the draft tube 5 is provided with a vent hole 51 at a position corresponding to the turbine blade 2, the diameter of the vent hole 51 is smaller than the diameter of the exhaust port of the draft tube 5, and the diameter of the vent hole 51 is smaller than the diameter of the outlet end of the bleed air tube 4. The diameter of the draft tube 5 is larger than the diameter of the outlet end of the bleed air tube 4.
In some embodiments, the draft tube 5 is provided with a vent hole 51 at a position on the central axis of the outlet end of the bleed air tube 4, the diameter of the vent hole 51 is smaller than the diameter of the outlet end of the draft tube 5, and the diameter of the vent hole 51 is smaller than the diameter of the outlet end of the bleed air tube 4. The diameter of the draft tube 5 is larger than the diameter of the outlet end of the bleed air tube 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 blades 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 blades 2, the position facing the turbine blades 2 on the draft tube 5 is provided with a vent hole 51, and the vent hole 51 has the function of compensating the air flow. The cooling gas in the draft tube 5 is partially discharged from the vent holes 51, and most of the remaining cooling gas is discharged from the exhaust ports of the draft tube 5, so as to avoid the uneven phenomenon caused by the diversion of the air flow of the turbine blade 2 facing the draft tube 5.
In some embodiments, the flow area of the vent hole 51 is smaller than the flow area of the exhaust of the draft tube 5. The flow area of the air outlet of the draft tube 5 is more than or equal to the flow area of the outlet end of the air guide tube 4. The cooling gas in the draft tube 5 is partially discharged from the vent holes 51, and most of the remaining cooling gas is discharged from the exhaust port of the draft tube 5, so as to avoid the non-uniformity caused by the diversion of the air flow of the turbine blade 2 facing the draft tube 5.
Airflow enters the draft tube 5 at the outlet end of the bleed air tube 4 radially inwards, the airflow speed is directly guided into the circumferential direction by the draft tube 5, the radially inwards impact speed is completely eliminated, the airflow at the outlet end of the bleed air tube 4 does not directly face the turbine blades 2, and the airflow can be better conveyed to the turbine blades 2 far away from the outlet end of the bleed air tube 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 pipe is more obvious, and the gas turning loss is less.
In some embodiments, at least two bleed air ducts 4 are provided around the turbine casing 1, and a flow duct 5 is provided for each bleed air duct 4.
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 flow guiding pipe 5. A turbine blade 2 is arranged on the central axis of the outlet end of each bleed air duct 4. The position of the honeycomb duct 5, which is located on the central axis of the outlet end of the bleed air duct 4, is provided with a vent hole 51, and the flow area of the vent hole 51 is smaller than that of the outlet end of the bleed air duct 4.
As shown in fig. 2, the draft tube 5 comprises a straight tube. The angle range of the included angle a between the straight pipe and the central axis of the outlet end of the air guide pipe 4 is 0 degrees < a <90 degrees.
As shown in fig. 3, the draft tube 5 comprises an arc-shaped tube. The arced tube projects in a direction closer to the central axis of the turbine case 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 pipe or the arc pipe, and the turbine blade 2 directly opposite to the outlet end of the air guide pipe 4 is not directly impacted; 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);
a plurality of turbine blades (2) 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
honeycomb duct (5), locate in gas collecting chamber (3), the air inlet of honeycomb duct (5) is connected the exit end of bleed pipe (4), the gas vent of honeycomb duct (5) is deviated the exit end of bleed pipe (4), honeycomb duct (5) are configured to guide cooling gas to flow, make cooling gas have along gas collecting chamber (3) and around the axis of gas collecting chamber (3) is the component of rotatory flow.
2. The turbine blade cooling system according to claim 1, characterized in that the flow guide tube (5) is configured to extend in a direction deviating from the central axis of the outlet end of the bleed air tube (4).
3. The turbine blade cooling system of claim 1, characterized in that the flow guide tube (5) comprises a straight tube or an arc tube.
4. The turbine blade cooling system according to claim 3, characterized in that the angle a between the straight tube and the centre axis of the outlet end of the bleed air tube (4) is in the range 0 ° < a <90 °.
5. The turbine blade cooling system of claim 3, characterized in that the arced tube projects 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 arced tube projects away from a central axis of the turbine casing (1).
7. The turbine blade cooling system according to claim 1, characterized in that the area of the flow guide tube (5) corresponds to at least one turbine blade (2), and that the flow guide tube (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 tube (4).
8. The turbine blade cooling system according to claim 1, characterized in that the flow guide tube (5) is provided with a vent hole (51) at a position on the central axis of the outlet end of the bleed air tube (4), the diameter of the vent hole (51) being smaller than the diameter of the outlet end of the bleed air tube (4).
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), each bleed air duct (4) being connected to a flow duct (5) in correspondence.
10. An aircraft engine, characterized in that it comprises a turbine blade cooling system according to any one of claims 1 to 9.
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| CN202022044908.6U CN213478401U (en) | 2020-09-17 | 2020-09-17 | Turbine blade cooling system and aircraft engine |
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| CN202022044908.6U CN213478401U (en) | 2020-09-17 | 2020-09-17 | Turbine blade cooling system and aircraft engine |
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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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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Cited By (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| 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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