CN111927560A - Low-position air inlet vane type pre-rotation nozzle structure - Google Patents
Low-position air inlet vane type pre-rotation nozzle structure Download PDFInfo
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- CN111927560A CN111927560A CN202010756608.2A CN202010756608A CN111927560A CN 111927560 A CN111927560 A CN 111927560A CN 202010756608 A CN202010756608 A CN 202010756608A CN 111927560 A CN111927560 A CN 111927560A
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Classifications
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D11/00—Preventing or minimising internal leakage of working-fluid, e.g. between stages
- F01D11/02—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type
- F01D11/04—Preventing or minimising internal leakage of working-fluid, e.g. between stages by non-contact sealings, e.g. of labyrinth type using sealing fluid, e.g. steam
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D5/00—Blades; Blade-carrying members; Heating, heat-insulating, cooling or antivibration means on the blades or the members
- F01D5/02—Blade-carrying members, e.g. rotors
- F01D5/08—Heating, heat-insulating or cooling means
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/041—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector using blades
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- F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
- F01—MACHINES OR ENGINES IN GENERAL; ENGINE PLANTS IN GENERAL; STEAM ENGINES
- F01D—NON-POSITIVE DISPLACEMENT MACHINES OR ENGINES, e.g. STEAM TURBINES
- F01D9/00—Stators
- F01D9/02—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles
- F01D9/04—Nozzles; Nozzle boxes; Stator blades; Guide conduits, e.g. individual nozzles forming ring or sector
- F01D9/047—Nozzle boxes
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- Engineering & Computer Science (AREA)
- Mechanical Engineering (AREA)
- General Engineering & Computer Science (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a low-level air inlet blade type pre-rotation nozzle structure, which comprises an outer ring, an inner ring, a flow deflector and a grate disc, wherein a ring cavity communicated with the outside is formed between the outer ring and the inner ring, the ring cavity comprises an air inlet, a ring cavity body and an air outlet, a plurality of mounting holes distributed in the radial direction are arranged on the inner ring, the mounting holes are arranged at the air outlet of the ring cavity, the number of the flow deflector is consistent with that of the mounting holes, the bottoms of the flow deflector are arranged in the mounting holes in a one-to-one correspondence manner, the top of the flow deflector is in interference fit with the inner wall surface of the outer ring, the grate disc comprises a disc body, an inner grate tooth and an outer grate tooth, the inner grate tooth and the outer grate tooth are both in annular structures, the inner grate tooth and the outer grate tooth are coaxially arranged on the disc body, a plurality of axial through holes are arranged in the circumferential direction between the inner grate tooth and the outer grate tooth, the air outlet of, the temperature of the cold air inlet is reduced by 30K, and the temperature requirement of the existing high-pressure turbine rotor blade in the full state is met.
Description
Technical Field
The invention belongs to the technical field of design of air systems of aircraft engines, and relates to a low-position air inlet blade type pre-rotation nozzle structure.
Background
The turbine rotor blade is a core part in the turbine, the working environment is very harsh, the long-term working temperature is more than 1000 ℃, the rotating speed is more than 10000r/min, and the turbine rotor blade bears huge centrifugal force, thermal stress and pneumatic bending moment load and is scoured and corroded by high-temperature gas. And with the increase of the thrust of the engine, the corresponding total temperature before the turbine is increased, which puts higher requirements on the strength and the service life of the turbine rotor blade. In order to solve the above problems, in addition to improving the thermal strength of the turbine blade material, the design of cooling the turbine blade is also a current development path. Because the cooling air flow of the turbine blades comes from a fan or an air compressor, the performance of an engine is reduced due to overlarge bleed air flow, and the cooling requirements of the turbine blades are not met due to the fact that the bleed air flow is too small. This places a certain flow and temperature demand on the cooling air flow entering the turbine blades.
The aircraft engine air system refers to: an airflow system outside the airflow in the main combustion gas passage of the engine. One of its primary functions is to direct the cool air from the compressor through the air system components to the high temperature components (turbines) that have certain flow and temperature requirements. Wherein an air system component refers to a part or device having certain flow resistance and heat exchange characteristics. The prewhirl nozzle is one of key elements of an air system, and the basic principle is that cooling airflow entering the rotor blade generates circumferential speed along the rotation direction of a turbine disc, so that the relative total temperature between the cold air and the rotor blade is reduced, and the functions of reducing the inlet temperature of the cold air of the blade and reducing the using amount of the cold air are further realized. So as to achieve the purposes of improving the cooling effect of the blades and reducing the air entraining amount of the air compressor.
Disclosure of Invention
The invention aims to solve the problem that the cooling design of the high-pressure turbine rotor blade requires the flow and temperature of inlet cold air by adopting a low-position arranged blade type pre-rotation nozzle structure aiming at the cooling requirement of the high-pressure turbine rotor blade of the existing aviation gas turbine fan engine.
The technical scheme of the invention is as follows:
a low-position air inlet blade type prewhirl nozzle structure, which comprises an outer ring, an inner ring, a flow deflector and a grate disc, an annular cavity communicated with the outside is formed between the outer ring and the inner ring, the annular cavity comprises an air inlet, an annular cavity body and an air outlet, the inner ring is provided with a plurality of mounting holes which are distributed in the radial direction and are positioned at the air outlet of the ring cavity, the number of the flow deflectors is consistent with that of the mounting holes, the bottoms of the flow deflectors are correspondingly arranged in the mounting holes one by one, the top of the flow deflector is in interference fit with the inner wall surface of the outer ring, the grate disc comprises a disc body, an inner grate and an outer grate, the inner and outer grates are both of annular structures, the inner and outer grates are coaxially arranged on the tray body, a plurality of axial through holes are circumferentially arranged between the inner and outer grates, and the air outlet of the annular cavity is opposite to the axial through holes.
Furthermore, the guide vane comprises a blade basin, a blade back, a front edge and a tail edge, the profile degrees of the blade basin and the blade back are both 0.15, the front edge is an arc with the radius of 0.45-0.5mm, and the tail edge is an arc with the radius of 0.17-0.19 mm.
Furthermore, the chord length of the flow deflector is 16-17 mm.
Furthermore, the number of the flow deflectors and the number of the mounting holes are 65.
Furthermore, the blade-shaped installation angle of the guide vane is 22 degrees 17-22 degrees 57'.
Further, the minimum distance between adjacent guide vanes is 2.3-2.34 mm.
Further, the inner ring is connected with the outer ring through a bolt.
Furthermore, the inner grid comprises 4 parallel grids, wherein the radius of 2 grids close to the tray body is larger than that of 2 grids far away from the tray body.
Furthermore, the outer grid comprises 4 parallel grids, wherein the radius of 2 grids close to the plate body is larger than that of 2 grids far away from the plate body.
Furthermore, the number of the axial through holes between the inner grate and the outer grate is 20, and the radius of the axial through holes is 13.3 mm.
The beneficial technical effects are as follows: the low-position air inlet blade type pre-rotation nozzle structure provided by the invention has the advantages that under the condition that the temperature field level of the original high-pressure turbine rotor blade is not changed, through calculation and experimental verification, the low-position air inlet blade type pre-rotation nozzle structure can reduce the flow of cold air by 30%, reduce the inlet temperature of the cold air by 30K, and meet the temperature requirement of the existing high-pressure turbine rotor blade in a full state; meanwhile, the amount of air entraining is reduced, and the performance of the engine is improved.
Drawings
FIG. 1 is a schematic view of the installation positions of the outer ring, the inner ring and the guide vanes of the present invention;
FIG. 2 is a schematic view of the mounting structure of the present invention;
FIG. 3 is a schematic view of a baffle of the present invention;
the device comprises a ring 1, an outer ring 2, an inner ring 3, a flow deflector 4, a labyrinth disc 5, a leaf basin 6, a leaf back 7, bolts 8, a ring cavity body 9, an air outlet 10, an air inlet 11, a disc body 12, an inner labyrinth 13, an outer labyrinth 14 and an axial through hole.
Detailed Description
The following description of the embodiments of the present invention, with reference to the accompanying drawings, will be made in further detail for the purpose of providing a more complete, accurate and thorough understanding of the concept and technical solutions of the present invention, by describing the embodiments, such as the shapes, structures, mutual positions and connection relationships of the components, the functions and operating principles of the components, the manufacturing processes and the operation and use methods of the components.
The invention belongs to an air system of an aircraft engine, is one of key elements of the system, and particularly relates to a low-position air inlet blade type prewhirl nozzle. To minimize the amount of cooling air required to meet the high pressure turbine rotor blade temperature levels.
Meanwhile, the requirements of the inlet temperature of the high-pressure turbine rotor blade cold air and the high-pressure turbine inter-stage sealing are considered, the low-position arrangement of the pre-rotation nozzle is favorable for controlling the inter-stage sealing leakage amount not to be too large, and the requirement of the inlet temperature of the high-pressure turbine rotor blade cold air can be met.
As shown in fig. 1 and 2, a low-position air inlet vane type pre-rotation nozzle structure comprises an outer ring 1, an inner ring 2, a flow deflector 3 and a grate disc 4, wherein an annular cavity communicated with the outside is formed between the outer ring 1 and the inner ring 2, the annular cavity comprises an air inlet 10, an annular cavity body 8 and an air outlet 9, a plurality of radially distributed mounting holes are arranged on the inner ring 2, the mounting holes are positioned at the air outlet 9 of the annular cavity, the number of the flow deflector 3 is the same as that of the mounting holes, the bottom of the flow deflector 3 is slightly larger than the size of the mounting holes, the bottoms of the flow deflector 3 are correspondingly arranged in the mounting holes one by one and are in interference fit with the mounting holes, so that the flow deflector 3 can be conveniently and firmly mounted and dismounted, the top of the flow deflector 3 supports against the inner wall surface of the outer ring 1 in an interference fit manner, and the top of the flow deflector 3 is assembled with the inner wall, the two ends of the flow deflector 3 are also perpendicular to the inner wall surfaces of the inner ring 1 and the outer ring 1 respectively, the grate disc 4 comprises a disc body 11, an inner grate 12 and an outer grate 13, the inner grate 12 and the outer grate 13 are both of an annular structure, the inner grate 12 and the outer grate 13 are coaxially arranged on the disc body 11, the inner grate 12 is positioned at a position close to the center of the disc body 11, the outer grate 13 is positioned at the outer side of the inner grate 12, 20 axial through holes 14 are circumferentially arranged between the inner grate 12 and the outer grate 13, the axial through holes 14 are uniformly distributed along the axial direction of the disc body 11, the diameter of the axial through holes 14 is 13.3mm, the mechanical strength of the disc body 11 is ensured to be sufficient, the air flow can be ensured to be normal, the air outlet 9 of the annular cavity is opposite to the axial through holes 14, and the majority of air flowing out from the air outlet 9 of the annular cavity is ensured to enter.
As shown in fig. 3, according to the theoretical calculation structure of the flow surface of S2S2, the guide vane 3 of the present invention includes a blade basin 4, a blade back 5, a leading edge and a trailing edge, wherein the profile degrees of the blade basin 4 and the blade back 5 are both 0.15, the leading edge is a circular arc chamfer with a radius of 0.45-0.5mm, preferably 0.45mm, and the trailing edge is a circular arc with a radius of 0.17-0.19mm, preferably 0.17 mm; the chord length of the flow deflector 3 is 16-17mm, preferably 16.5mm, and the number of the flow deflector 3 and the number of the mounting holes are 65; the vane profile installation angle of the guide vanes 3 is 22 degrees 17-22 degrees 57', and the minimum distance between adjacent guide vanes 3 is 2.3-2.34mm, preferably 2.32 mm. The design of the parameters of the blade type guide vane 3 has the characteristics of high flow coefficient and low pressure loss. To minimize the amount of cooling air required to meet the high pressure turbine rotor blade temperature levels.
The flow deflector 3 is made of titanium alloy or nickel alloy material, the titanium alloy has the advantages of high strength, good corrosion resistance, good fatigue resistance, small heat conductivity and expansion coefficient and the like, can be used for a long time below 350-450 ℃, and can be used at low temperature of-196 ℃.
The inner ring 2 and the outer ring 1 are connected or welded through the bolts 7, the bolts 7 can be conveniently installed and detached in a connected mode, the fixing is firm, the using amount of parts can be reduced through welding, the weight of the airplane is reduced, the weight of the airplane is further reduced, and the fixing effect is firmer.
As shown in fig. 2, the inner labyrinth 12 includes 4 parallel grates, wherein the radius of the 2 grates close to the tray body 11 is greater than the 2 grates far from the tray body 11, the outer labyrinth 13 includes 4 parallel grates, wherein the radius of the 2 grates close to the tray body 11 is greater than the 2 grates far from the tray body 11, and the grates with different radii are used for air sealing of a front cavity of the turbine rotor, and the sudden expansion and the sudden contraction of the channel are used for increasing the flow resistance and consuming the energy of the high-temperature gas flow.
Referring to fig. 2, when the low-position air inlet vane type pre-swirl nozzle structure of the present invention is installed, a nozzle assembly composed of an outer ring 1, an inner ring 2, and a guide vane 3 is installed on an inner support of a high-pressure turbine vane by a bolt 7. The grate plate 4 is installed on the high-pressure turbine disc through a long bolt and rotates at a high speed along with the high-pressure turbine disc, and the axial through hole 14 of the grate plate 4 is aligned with the outlet of the nozzle during assembly. In order to obtain a higher flow coefficient and a lower pressure loss coefficient, the vane profile parameters of the guide vane 3 are as follows: the radius of the front edge, the radius of the tail edge, the chord length, the number of blade grids, the installation angle of the blade profile and the like are designed into corresponding values. In order to ensure the flow of cooling air flow, the minimum distance between adjacent guide vanes 3 and the number and size of axial through holes of the labyrinth plate are controlled emphatically, four stepped upper and lower labyrinth teeth are adopted for sealing, and the air flow leakage is reduced.
According to the arrow direction in fig. 1, the cooling gas led from the two air flows of the combustion chamber enters an annular cavity formed by the outer ring 1 and the inner ring 2, the air flow velocity is increased after passing through the flow deflector 3, and the air flow direction is deflected. Most of the air flow enters the rotating disc cavity through the axial through hole 14 of the grate disc 4 to cool the wheel disc and the blades. Small parts respectively leak out from the inner grate 12 and the outer grate 13. The working principle is as follows: the airflow obtains high tangential speed and then enters the rotating disc cavity with high rotating speed through the prerotation action of the flow deflector 3 on the airflow. Because of the influence of pre-rotation, the turntable does not need to do work on the air flow, so that the relative total temperature of the air flow is reduced, and the effect of improving the cooling effect is achieved.
The invention adopts a flow-thermal coupling calculation method to design the nozzle, the 3-blade profile of the flow deflector is given by S2 flow surface pneumatic calculation, and the nozzle is arranged at a low position. Under the condition that the temperature field level of the original high-pressure turbine rotor blade is not changed, calculation and test verification prove that the low-position air inlet blade type pre-rotation nozzle structure can reduce the cold air flow by 30 percent, the flow coefficient of the nozzle reaches 0.91, the total temperature of a cold air inlet of the blade is reduced by about 30K, the requirements of the cold air flow and the temperature of a high-pressure turbine blade are met, the temperature of a high-vortex rotor is effectively reduced, and the strength reserve is improved. The invention passes a component test, and the flow, the flow coefficient, the pressure loss coefficient, the outlet airflow angle and the like meet the requirements of design indexes; the low-level air inlet blade type pre-rotation nozzle structure is assembled on an engine and is subjected to complete machine test run examination for many times, the performance of the engine in each examination reaches the standard, the reliability of a high-pressure turbine rotor meets the overall requirement, and the designed low-level air inlet blade type pre-rotation nozzle structure achieves the design target.
The invention has been described above with reference to the accompanying drawings, it is obvious that the invention is not limited to the specific implementation in the above-described manner, and it is within the scope of the invention to apply the inventive concept and solution to other applications without substantial modification.
Claims (10)
1. The utility model provides a low level is admitted air the blade profile type and is revolved nozzle structure in advance which characterized in that: the novel guide vane structure comprises an outer ring (1), an inner ring (2), guide vanes (3) and a grate disc (4), wherein an annular cavity communicated with the outside is formed between the outer ring (1) and the inner ring (2), the annular cavity comprises an air inlet (10), an annular cavity body (8) and an air outlet (9), a plurality of radially distributed mounting holes are formed in the inner ring (2), the mounting holes are located at the air outlet (9) of the annular cavity, the number of the guide vanes (3) is consistent with that of the mounting holes, the bottoms of the guide vanes (3) are arranged in the mounting holes in a one-to-one correspondence manner, the tops of the guide vanes (3) are in interference fit with the inner wall surface of the outer ring (1), the grate disc (4) comprises a disc body (11), inner grate teeth (12) and outer grate teeth (13), the inner grate teeth (12) and the outer grate teeth (13) are of an annular structure, and the inner grate teeth (12) and the outer grate teeth (13) are of an annular structure, The outer grate (13) is coaxially arranged on the plate body (11), a plurality of axial through holes (14) are circumferentially arranged between the inner grate (12) and the outer grate (13), and the air outlet (9) of the annular cavity is opposite to the axial through holes (14).
2. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the guide vane (3) comprises a vane basin (5), a vane back (6), a front edge and a tail edge, wherein the profile degrees of the vane basin (5) and the vane back (6) are both 0.15, the front edge is an arc with the radius of 0.45-0.5mm, and the tail edge is an arc with the radius of 0.17-0.19 mm.
3. The low inlet vane type pre-swirl nozzle structure of claim 2, wherein: the chord length of the flow deflector (3) is 16-17 mm.
4. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the number of the flow deflectors (3) and the number of the mounting holes are 65.
5. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the blade-shaped installation angle of the guide vane (3) is 22 degrees 17-22 degrees 57'.
6. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the minimum distance between the adjacent guide vanes (3) is 2.3-2.34 mm.
7. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the inner ring (2) is connected with the outer ring (1) through a bolt.
8. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the inner grid (12) comprises 4 parallel grids, wherein the radius of 2 grids close to the tray body (11) is larger than that of 2 grids far away from the tray body (11).
9. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the outer grid teeth (13) comprise 4 parallel grid teeth, wherein the radius of 2 grid teeth close to the tray body (11) is larger than that of 2 grid teeth far away from the tray body (11).
10. The low inlet vane type pre-swirl nozzle structure of claim 1, wherein: the number of the axial through holes (14) between the inner grate (12) and the outer grate (13) is 20, and the radius is 13.3 mm.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010756608.2A CN111927560A (en) | 2020-07-31 | 2020-07-31 | Low-position air inlet vane type pre-rotation nozzle structure |
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| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202010756608.2A CN111927560A (en) | 2020-07-31 | 2020-07-31 | Low-position air inlet vane type pre-rotation nozzle structure |
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| CN111927560A true CN111927560A (en) | 2020-11-13 |
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| CN202010756608.2A Pending CN111927560A (en) | 2020-07-31 | 2020-07-31 | Low-position air inlet vane type pre-rotation nozzle structure |
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Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115013069A (en) * | 2022-07-06 | 2022-09-06 | 中国航发湖南动力机械研究所 | Turbine rotor cooling system and aircraft engine |
| CN115853598A (en) * | 2022-11-29 | 2023-03-28 | 中国航空发动机研究院 | Turbine blade air conditioning supercharging impeller with axial air intake and pre-rotation supercharging air supply system |
| CN116537895A (en) * | 2023-07-04 | 2023-08-04 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
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Cited By (5)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN115013069A (en) * | 2022-07-06 | 2022-09-06 | 中国航发湖南动力机械研究所 | Turbine rotor cooling system and aircraft engine |
| CN115853598A (en) * | 2022-11-29 | 2023-03-28 | 中国航空发动机研究院 | Turbine blade air conditioning supercharging impeller with axial air intake and pre-rotation supercharging air supply system |
| CN115853598B (en) * | 2022-11-29 | 2023-09-22 | 中国航空发动机研究院 | Turbine blade cold air supercharging impeller for axial air intake and pre-rotation supercharging air supply system |
| CN116537895A (en) * | 2023-07-04 | 2023-08-04 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
| CN116537895B (en) * | 2023-07-04 | 2023-09-15 | 中国航发四川燃气涡轮研究院 | Pre-rotation air supply system with comb gap control |
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