Sealing and cooling structure and method for adjusting axial force of aero-engine rotor

Abstract

The invention relates to the technical field of aero-engines, in particular to a sealing and cooling structure and method for adjusting the axial force of an aero-engine rotor, which comprises a turbine disc provided with a rear baffle, and further comprises a low-pressure turbine guide vane arranged opposite to the turbine disc, a low guide front inner supporting ring and a low guide rear inner supporting ring are arranged in the middle of the low-pressure turbine guide vane; labyrinth teeth correspondingly matched with the low-guide front inner supporting plate are formed on the rear baffle, radial labyrinth tooth gaps are formed, and when airflow in an air collection cavity enters a rotor cavity from a low-pressure pre-rotation nozzle, the airflow is embedded into an air supply channel from a front pre-rotation hole at the same time; one part of airflow entering the air supply channel enters the disc rear cavity, and the other part of airflow enters the rotor cavity. By adjusting the sealing labyrinth structure, it is ensured that the axial force of the rotor does not exceed the allowable maximum value in the large-state working process and is not subjected to light load in the small-state working process, meanwhile, sealing gas used behind a turbine disc is ensured, turbine hot gas is prevented from flowing into an inner disc cavity from a main channel, and it is ensured that an engine works reliably.

Description

TECHNICAL FIELD

[0001] The present application relates to the technical field of aero-engines, in particular to an aero-engine air system and axial force design technology, and more particularly to a sealing cooling structure and method for adjusting the axial force of an aero-engine rotor. BACKGROUND

[0002] As the heart of an airplane, the working reliability of an aero-engine is of great importance, and the axial force has a direct impact on the working reliability of the aero-engine. In the design process of the aero-engine, the determination of the axial force is in the last link, and this stage faces many challenges.

[0003] On the one hand, there are many and extremely complex factors affecting the axial force. From the parts of the engine, it is difficult to achieve absolute precision in the structure of each part, and deviations will inevitably occur in the manufacturing process. These deviations accumulate and have a chain reaction on multiple key performance indicators of the engine. For example, it may cause a gap between the working efficiency of each part and the design expectation, the pressure ratio of the compressor cannot reach the theoretical calculation value, the expansion ratio of the turbine also has a deviation, and the actual value of the cavity pressure of the air system is inconsistent with the design calculation result.

[0004] On the other hand, due to the existence of the above-mentioned multiple deviations, there is a large deviation between the calculation result of the axial force and the design result. If this deviation cannot be effectively solved, it will directly affect the reliability of the engine and pose a hidden danger to the safe operation of the engine.

[0005] Therefore, under the background of the prior art, the adjustment of the axial force of the aero-engine lacks a systematic method. At present, only partial adjustment can be made after the completion of the whole machine test, which is not only low in efficiency, but also difficult to fundamentally solve the problem of axial force deviation.

[0006] It can be seen that the current axial force design process of the aero-engine rotor still has room for improvement, and should be optimized to improve the efficiency of the axial force adjustment design, improve the accuracy of the design scheme, reduce repeated modification and adjustment, and thus improve the reliability of the design. Therefore, a more reasonable technical scheme needs to be proposed to solve the technical problems in the prior art. SUMMARY

[0007] At least to overcome one of the above-mentioned defects, the present application proposes an aero-engine air system and axial force design technology, which adjusts the high-pressure turbine disc rear pressure by increasing the sealing grate and optimizing the grate gap, ensures that the rotor axial force does not exceed the maximum allowable value in large state work, does not overload in small state work, and at the same time ensures the sealing air after the turbine disc, prevents the turbine hot gas from flowing into the internal disc cavity from the main channel, and ensures the working reliability of the engine.

[0008] To achieve the above object, the sealing cooling structure disclosed by the present application can adopt the following technical solutions. The sealing cooling structure for adjusting the axial force of an aero-engine rotor comprises a turbine disc provided with a back baffle, and a low-pressure turbine guide vane arranged opposite to the turbine disc, wherein a middle part of the low-pressure turbine guide vane is provided with a low-guide front inner support ring and a low-guide rear inner support ring, a gas collection cavity is formed between the low-guide front inner support ring and the low-guide rear inner support ring, and a gas supply channel is formed between the low-guide front inner support ring, the back baffle, the turbine disc and the low-guide rear inner support ring; the back baffle is formed with a labyrinth corresponding to the low-guide front inner support ring and a radial labyrinth gap, when the gas flow collected in the gas collection cavity reaches a set amount, the gas flow enters a rotor cavity from a low-pressure pre-whirl nozzle of the gas collection cavity, and simultaneously enters the gas supply channel from a front pre-whirl hole; the gas flow entering the gas supply channel is partially output outward into a disc rear cavity through the labyrinth gap, and the other part of the gas flow enters the rotor cavity through a brush seal assembly at the low-guide rear inner support ring.

[0009] In the sealing cooling structure disclosed above, the gas flow is guided to enter the low-pressure turbine guide vane and then is respectively delivered to the rotor cavity and the gas supply channel from the gas collection cavity, the gas flow entering the rotor cavity is used to help the rotor dissipate heat, and the gas flow entering the gas supply channel is then divided into two routes, one of which enters the disc rear cavity to form a positive pressure seal at the labyrinth gap, and the other of which enters the rotor cavity from the brush seal assembly to help the rotor dissipate heat.

[0010] In the sealing cooling structure disclosed above, the low-pressure turbine guide vane is a hollow structure, and the gas flow passes through the hollow structure.

[0011] In some solutions, the front pre-whirl hole is arranged on the low-guide front inner support ring, and the function of the front pre-whirl hole is to introduce high-pressure cooling air, expand and accelerate the gas flow, and increase the tangential velocity of the cooling gas flow, so as to significantly reduce the total temperature of the gas flow felt by the turbine rotor, and control the flow of the cooling gas entering the disc rear cavity of the high-pressure turbine disc; preferably, the front pre-whirl hole is a circular hole, the diameter is 5 mm, there are 42 holes in the circumferential direction, the hole axis is at an angle of 70° with the engine axial direction, and the inlet is rounded or chamfered ≮0.1 mm.

[0012] In some solutions, the low-pressure pre-whirl nozzle is arranged on the low-guide rear inner support ring, and the function of the low-pressure pre-whirl nozzle is to introduce high-pressure cooling air, expand and accelerate the gas flow, and increase the tangential velocity of the cooling gas flow, so as to significantly reduce the total temperature of the gas flow felt by the turbine rotor, and control the flow of the cooling gas entering the disc front cavity of the low-pressure turbine disc.

[0013] Further, the rear baffle is used to help form a positive pressure seal to the rear cavity of the disc, the teeth on the rear baffle are used to cooperate with the tooth gap of the low guide front inner support ring, thereby allowing the airflow from the gas supply channel to pass through and form a positive pressure seal, the rear baffle can be constructed in various forms, and its structure is not uniquely limited, and one of the feasible options is optimized and proposed herein: the rear baffle is formed with a tooth plate extending towards the low guide front inner support ring, a plurality of teeth are formed on the tooth plate, and the radial tooth gap between the tooth plate and the low guide front inner support ring is 0.6mm-0.7mm. When the above scheme is adopted, the tooth gap can be set according to the actual air flow and sealing cooling requirements.

[0014] In some schemes, the tooth sealing structure is formed by the double teeth on the rear baffle and the honeycomb welded on the low guide front inner support ring, which is mainly used to adjust the turbine disc rear pressure and ensure that the rotor axial force is within a suitable range, and the preferred tooth shape of the tooth is inclined tooth, and the number of teeth is 2.

[0015] Further, in some schemes, the radial tooth gap is 0.65mm.

[0016] Further, on the passage of the airflow, the airflow is guided by setting corresponding guide structures, so that the airflow can maintain good sealing when flowing in different cavities, the guide structure can adopt various schemes, and its structure is not uniquely limited, and one of the feasible options is optimized and proposed herein: a guide pipe communicating with the low-pressure turbine guide vane is arranged between the low guide front inner support ring and the low guide rear inner support ring, and the airflow in the low-pressure turbine guide vane enters the plenum chamber through the guide pipe. When the above scheme is adopted, the guide pipes are distributed at intervals on the circumference of the low guide front inner support ring and the low guide rear inner support ring. The guide pipes are welded on the low guide front inner support ring and the low guide rear inner support ring, and simultaneously form the plenum chamber.

[0017] Further, when part of the airflow in the gas supply channel flows to the rotor cavity, the structure at the brush seal assembly can be constructed in various forms, which is not uniquely limited, and one of the feasible options is optimized and proposed herein: a rear shaft is arranged on the turbine disc, and a sealing surface cooperating with the brush seal assembly is formed on the rear shaft. When the above scheme is adopted, the rear shaft and the turbine disc are connected and fixed by fasteners.

[0018] In some schemes, the brush seal assembly is fixed to the low-pressure pre-rotation nozzle assembly by bolts, brush wires are arranged on the mounting seat of the brush seal assembly, and a brush seal device is formed with the rear shaft for controlling the downward cold gas leakage flow in the turbine disc rear passage C and ensuring that most of the cold gas after the outlet of the front pre-rotation hole is used for disc rear sealing.

[0019] Further, the air flow delivered to the inside of the low-pressure turbine guide vane is supplied by the external compressor, and the specific air supply structure can adopt various forms, which are not uniquely limited. Here, one feasible option is optimized and proposed: the outer edge of the low-pressure turbine guide vane is connected and matched with the bleed air pipe, and an annular cavity is formed between the low-pressure turbine guide vane and the bleed air pipe. The bleed air pipe is used to communicate with the compressor and deliver air flow into the annular cavity. When the above scheme is adopted, the bleed air pipe is circumferentially uniformly welded on the low-pressure turbine casing, and the low-pressure turbine casing is connected with the low-pressure turbine guide vane through bolts to form the annular cavity, thereby ensuring that the air flow at the outlet of the bleed air pipe uniformly enters the inner cavity of the low-pressure turbine guide vane.

[0020] According to the above disclosed sealing cooling system, the air flow with appropriate pressure and temperature is led out from a certain position of the compressor, enters the inner cavity of the low guide through several bleed air pipes, then enters the gas collecting cavity through the flow guide pipe, and is divided into two air flows in the inner cavity. One of the two air flows enters the rotor cavity through the low-pressure pre-rotation nozzle to cool the low-pressure turbine rotor, and the other air flow enters the air supply channel through the front pre-rotation hole and is then divided into two parts. Most of the air flow enters the disc rear cavity through the grid to prevent the turbine hot combustion gas from flowing into the internal disc cavity through the main channel. A small part of the air flow enters the rotor cavity through the brush seal and cools the low-pressure turbine rotor after being mixed with the air flow at the outlet of the low-pressure pre-rotation nozzle.

[0021] The above disclosure discloses a sealing cooling structure, and the present application also discloses a sealing cooling method.

[0022] A sealing cooling method for adjusting the axial force of an aero-engine rotor adopts the sealing cooling structure described above, and comprises the following steps. Determine the analysis working conditions, and take the minimum state and the maximum state in the engine performance recording and air system cavity temperature and cavity pressure measurement test as the key analysis working conditions of the axial force; Obtain the design requirements of the engine rotor axial force, including the maximum axial force threshold Fmax and the meshing force F1 in the intermediate state, and the minimum axial force threshold Fmin and the meshing force F2 in the slow vehicle state; Calibrate the air system calculation model before improvement to determine the flow area S of the front pre-rotation hole A , and calculate the disc rear cavity pressure PC; the calculation is specifically performed according to the following method:

[0023] In the above formula, π is the circular constant, R2 is the tooth tip radius of the grid, and R1 is the rear shaft radius corresponding to the brush seal assembly; Increase the grid sealing structure, preliminarily determine the tooth tip radius R2 of the grid, and determine the maximum threshold PC1 and the minimum threshold PC2 of the turbine disc rear pressure; The calculation is specifically performed according to the following method: PC1=F15 / S A PC2=F16 / SA F15 = F max - F1 - F3 - F5 + F7 + F9 - F11 - F13 F16 = F min - F2 - F4 - F6 + F8 + F10 - F12 - F14 In the above formula, F13 is the intermediate state turbine disk rear axial force, F14 is the slow state turbine disk rear axial force, F3 is the intermediate state compressor flow passage axial force, F4 is the slow state compressor flow passage axial force, F5 is the intermediate state compressor disk cavity axial force, F6 is the slow state compressor disk cavity axial force, F7 is the intermediate state turbine flow passage axial force, F8 is the slow state turbine flow passage axial force, F9 is the intermediate state turbine disk front axial force, F10 is the slow state turbine disk front axial force, F11 is the intermediate state turbine rear shaft axial force, F12 is the slow state turbine rear shaft axial force, F13 is the intermediate state turbine disk rear cavity axial force, F14 is the slow state turbine disk rear cavity axial force, PC1 is the maximum valve value of the disk rear pressure, and PC2 is the minimum valve value of the disk rear pressure. Obtain the minimum working clearance δ1 of the intermediate state labyrinth min Determine that the turbine disk rear PC meets the maximum valve value design requirement; Obtain the maximum working clearance δ2 of the slow state labyrinth max Determine that the slow state disk rear pressure PC meets the minimum valve value requirement; Calculate the theoretical working clearance difference △δ of the labyrinth in the intermediate state and the slow state, so as to meet: δ2 max -△δ<δ1<δ1 min Determine the optimal design clearance δ1 of the intermediate state labyrinth and the labyrinth tooth tip radius R2; Calculate the cold state clearance δ of the labyrinth, and determine the cold state size S3 of the labyrinth corresponding honeycomb.

[0024] Further, the calibration improves the air system calculation model, determines the flow area S of the front pre-swirl hole A Calculate the disk rear cavity pressure PC, wherein the turbine disk rear sealing flow G1 and the turbine disk front minimum critical sealing flow G1 min If G1>G1 min Indicates that the turbine disk rear sealing flow meets the design requirement, if G1 min Indicates that the disk rear sealing flow does not meet the design requirement, gradually expand the flow area S of the front pre-swirl hole A , until G1>G1 min Obtain new air system calculation results, and form new air system calculation network model and disk rear cavity pressure PC.

[0025]

[0026]

[0027] where N is the rotating speed, is the gas flow density, is the turbine disk rear flange sealing radius, is the dynamic viscosity, is the swirl coefficient of the gas flow, and K is an empirical coefficient.

[0028] Further, the minimum working gap δ1 of the intermediate state gullet is obtained min , and the turbine disk rear PC satisfies the maximum threshold design requirement; the method is as follows: Directly to the intermediate state gullet, if PC < PC1, it indicates that the intermediate state axial force satisfies the design requirement, if PC > PC1, it indicates that the intermediate state axial force exceeds the maximum axial force requirement of the bearing, and the gap δ1 of the gullet needs to be gradually increased to reduce the PC pressure until PC < PC1; when the adjustment is made to make PC < PC1, the new calculation model of the intermediate state air system and the minimum working gap δ1 of the gullet in the intermediate state are obtained min , and the rotor deformation corresponding to the intermediate state gullet is obtained by simultaneously carrying out the part temperature and deformation calculation and ; The maximum working gap δ2 of the slow speed state gullet is obtained max , and the turbine disk rear pressure PC in the slow speed state satisfies the minimum threshold requirement; the method is as follows: for the slow speed state gullet, if PC > PC2, it indicates that the slow speed state axial force satisfies the design requirement, if PC < PC2, it indicates that the disk rear pressure PC does not satisfy the slow speed state pressure requirement, the slow speed state axial force is light load, and the gap δ2 of the gullet needs to be gradually reduced to increase the PC pressure until PC > PC2, and the new calculation model of the slow speed state air system and the maximum working gap δ2 of the gullet in the slow speed state are obtained max , and the rotor deformation corresponding to the slow speed state gullet is obtained by simultaneously carrying out the part temperature and deformation calculation and ; The theoretical working gap difference △δ of the gullet in the intermediate state and the slow speed state is calculated as follows:

[0029] Wherein: is the rotor deformation amount of the intermediate state sealing gullet, is the rotor deformation amount of the slow speed state sealing gullet, is the stator deformation amount corresponding to the intermediate state sealing gullet, The stator deformation corresponding to the seal damper in the slow speed state.

[0030] Further, the calculation of the cold state gap δ of the seal damper and the determination of the cold state size S3 of the seal damper corresponding to the honeycomb are performed as follows:

[0031]

[0032] In the above formula, δ1 is the optimal design gap of the seal damper in the intermediate state, is the stator deformation corresponding to the seal damper in the intermediate state, is the rotor deformation of the seal damper in the intermediate state, and R2 is the tooth tip radius of the seal damper.

[0033] Compared with the prior art, some beneficial effects of the technical scheme of the present application include: The seal cooling structure and method provided by the present application adjust the seal damper structure between the turbine disc and the low guide front inner support ring, optimize the seal damper gap, adjust the high pressure turbine disc rear pressure, ensure that the rotor axial force does not exceed the maximum allowable value in large state operation and does not cause light load in small state operation, and at the same time, ensure the seal gas after the turbine disc, prevent the turbine hot gas from flowing into the internal disc cavity from the main passage, and ensure the reliable operation of the engine. BRIEF DESCRIPTION OF DRAWINGS

[0034] In order to more clearly illustrate the technical scheme of the embodiments of the present application, the following will briefly introduce the drawings needed to be used in the embodiments. It should be understood that the following drawings only represent some embodiments of the present application, and therefore should not be regarded as a limitation on the scope. For those skilled in the art, other related drawings can also be obtained without creative labor on the basis of these drawings.

[0035] Figure 1 is a partial cooling structure diagram for adjusting the rotor axial force of the present application.

[0036] Figure 2 is a seal cooling flow path diagram for adjusting the rotor axial force of the present application.

[0037] Figure 3 is a seal cooling analysis method flow for adjusting the rotor axial force of an aero-engine.

[0038] In the above drawings, the meanings of the respective marks are as follows: 1, bleed air pipe; 2, low pressure turbine casing; 3, low pressure turbine guide vane; 4, flow guide pipe; 5, low guide front inner support ring; 6, low guide rear inner support ring; 7, brush seal assembly; 8, turbine disc; 9, rear baffle; 10, moving blade; 11, low pressure pre-rotation nozzle; 12, front pre-rotation hole; 13, seal damper; 14, rear shaft; 15, seal surface.

[0039] A, annular cavity; B, gas collection cavity; C, gas supply channel; D, rotor cavity; E, disc rear cavity. DETAILED DESCRIPTION

[0040] The present embodiment will be further explained in combination with the accompanying drawings and specific embodiments.

[0041] In view of the deficiencies in the design control of the axial force of the rotor structure of the existing aviation engine, the following embodiments optimize and overcome the defects in the existing technology.

[0042] Embodiment 1 As shown in Figure 1 , Figure 2 , the present embodiment discloses a sealing cooling structure for adjusting the axial force of an aviation engine rotor, comprising a turbine disc 8, a back baffle 9 arranged on the turbine disc 8, and a low-pressure turbine guide vane 3 arranged opposite to the turbine disc 8, wherein the middle part of the low-pressure turbine guide vane 3 is provided with a low-guide front inner support ring 5 and a low-guide rear inner support ring 6, a gas collection cavity B is formed between the low-guide front inner support ring 5 and the low-guide rear inner support ring 6, and a gas supply channel C is formed between the low-guide front inner support ring 5, the back baffle 9, the turbine disc 8, and the low-guide rear inner support ring 6; the back baffle 9 is formed with a labyrinth 13 corresponding to the low-guide front inner support ring 5 and a radial labyrinth 13 gap, when the airflow gathered in the gas collection cavity B reaches a set amount, the airflow enters a rotor cavity D from a low-pressure pre-rotation nozzle 11 of the gas collection cavity B, and simultaneously enters the gas supply channel C from a front pre-rotation hole 12; part of the airflow entering the gas supply channel C is output outward into a disc rear cavity E through the labyrinth 13 gap, and the other part enters the rotor cavity D through a brush seal assembly 7 at the low-guide rear inner support ring 6.

[0043] The sealing cooling structure disclosed in the present embodiment guides the airflow to enter the low-pressure turbine guide vane 3 into the gas collection cavity B, and respectively transports it to the rotor cavity D and the gas supply channel C from the gas collection cavity B, the airflow entering the rotor cavity D is used to help the rotor dissipate heat, and the airflow entering the gas supply channel C is then divided into two routes, one route enters the disc rear cavity E to form a positive pressure seal at the labyrinth 13 gap, and the other route enters the rotor cavity D from the brush seal assembly 7 to help the rotor dissipate heat.

[0044] In the sealing cooling structure disclosed in the present embodiment, the low-pressure turbine guide vane 3 is a hollow structure, and the airflow passes through the hollow structure.

[0045] In some schemes, the front pre-swirl hole 12 is arranged on the low-guide front inner support ring 5. The function of the front pre-swirl hole 12 is to introduce high-pressure cooling air, accelerate the expansion of the airflow and increase the tangential velocity of the cooling airflow, thereby significantly reducing the total temperature of the airflow felt by the turbine rotor, and at the same time controlling the flow of cold air entering the rear cavity of the high-pressure turbine disk 8. The preferred hole type of the front pre-swirl hole 12 is a circular hole with a diameter of 5 mm, a total of 42 circumferential holes, an angle of 70° between the hole axis and the engine axis, and an inlet fillet or chamfer ≮0.1 mm.

[0046] In some embodiments, the low-pressure pre-swirl nozzle 11 is arranged on the low-guide rear inner support ring 6. The function of the low-pressure pre-swirl nozzle 11 is to introduce high-pressure cooling air, accelerate the expansion of the airflow and increase the tangential speed of the cooling airflow, thereby significantly reducing the total temperature of the airflow felt by the turbine rotor, and at the same time controlling the flow of cold air entering the front cavity of the low-pressure turbine disc 8.

[0047] The rear baffle 9 is used to help form a positive pressure seal against the rear disk cavity E. The grate teeth 13 on the rear baffle 9 are used to match the gap between the grate teeth 13 of the low-guide front inner support ring 5, thereby allowing airflow from the air supply channel C to pass through and form a positive pressure seal. The rear baffle 9 can be constructed in various forms, and its structure is not limited to a single form. This embodiment optimizes and adopts one feasible option: the rear baffle 9 forms a grate tooth 13 plate extending toward the low-guide front inner support ring 5, and the grate tooth 13 plate is formed with a plurality of grate teeth 13. The radial grate tooth 13 gap between the grate tooth 13 plate and the low-guide front inner support ring 5 is 0.6mm to 0.7mm. When adopting this solution, the grate tooth 13 gap can be set according to the actual airflow and sealing cooling requirements.

[0048] In some schemes, a grate tooth 13 sealing structure is formed by welding the double teeth on the rear baffle 9 and the honeycomb on the low-guide front inner support ring 5, which is mainly used to adjust the pressure behind the turbine disc 8 to ensure that the axial force of the rotor is within an appropriate range. The preferred tooth shape of the grate tooth 13 is a bevel grate tooth 13 with 2 teeth.

[0049] Preferably, in this embodiment, the gap between the radial comb teeth 13 is 0.65 mm.

[0050] On the passage of the airflow, the airflow is guided by setting corresponding guide structure, so that the airflow can maintain good sealing when transferring in different cavities. The guide structure can adopt various schemes, and its structure is not uniquely limited. In the embodiment, one of the feasible options is optimized and adopted: the guide pipe 4 connected with the low-pressure turbine guide vane 3 is arranged between the low-guide front inner support ring 5 and the low-guide rear inner support ring 6. The airflow in the low-pressure turbine guide vane 3 enters the gas collection cavity B through the guide pipe 4. When the above scheme is adopted, the guide pipe 4 is distributed on the circumference of the low-guide front inner support ring 5 and the low-guide rear inner support ring 6. The guide pipe 4 is welded on the low-guide front inner support ring 5 and the low-guide rear inner support ring 6, and simultaneously constitutes the gas collection cavity B.

[0051] When part of the airflow in the gas supply passage C flows to the rotor cavity D, the structure at the brush seal assembly 7 can be configured in various forms, which is not uniquely limited. In the embodiment, one of the feasible options is optimized and adopted: the rear shaft 14 is arranged on the turbine disc 8, and the sealing surface 15 cooperating with the brush seal assembly 7 is formed on the rear shaft 14. When the above scheme is adopted, the rear shaft 14 is connected and fixed with the turbine disc 8 by the fastener.

[0052] In some schemes, the brush seal assembly 7 is fixed on the low-pressure pre-rotation nozzle 11 assembly by bolts. The brush wire is arranged on the mounting seat of the brush seal assembly 7, and forms a brush seal device with the rear shaft 14, which is used to control the leakage flow of the downward cold gas in the rear passage of the turbine disc 8, and ensures that most of the cold gas after the outlet of the front pre-rotation hole 12 is used for sealing behind the disc.

[0053] The airflow delivered into the low-pressure turbine guide vane 3 is supplied by an external compressor. The specific air supply structure can adopt various forms, which is not uniquely limited. In the embodiment, one of the feasible options is optimized and adopted: the outer edge of the low-pressure turbine guide vane 3 is connected and matched with the bleed air pipe 1. The annular cavity A is formed between the low-pressure turbine guide vane 3 and the bleed air pipe 1. The bleed air pipe 1 is used to communicate with the compressor and deliver the airflow into the annular cavity A. When the above scheme is adopted, the bleed air pipe 1 is circumferentially and uniformly welded on the low-pressure turbine casing 2. The low-pressure turbine casing 2 is connected with the low-pressure turbine guide vane 3 by bolts, forming the annular cavity A, and ensuring that the airflow at the outlet of the bleed air pipe 1 uniformly enters the inner cavity of the low-pressure turbine guide vane 3.

[0054] According to the sealed cooling system disclosed above, an air flow with suitable pressure and temperature is drawn out from a certain position in the compressor, passes through several air bleed pipes 1, enters the low-pressure guide cavity, and then passes through the guide pipe 4 to enter the air collecting cavity B, and is divided into two air flows in the cavity. One of the air flows enters the rotor cavity D through the low-pressure pre-swirl nozzle 11 to cool the low-pressure turbine rotor, and the other air flows through the front pre-swirl hole 12, enters the air supply channel C, and is further divided into two parts. Most of the air flows through the grate teeth 13 to enter the rear cavity E of the disk to prevent the hot combustion gas of the turbine from flowing into the internal disk cavity from the main channel; a small part of the air flows into the rotor cavity D through the brush seal, mixes with the air flow at the outlet of the low-pressure pre-swirl nozzle 11, and then cools the low-pressure turbine rotor.

[0055] Example 2 The content of the above embodiment 1 discloses a sealing cooling structure, and this embodiment also discloses a sealing cooling method.

[0056] like Figure 3 As shown, a seal cooling method for regulating the axial force of an aircraft engine rotor adopts the seal cooling structure described above, comprising: Determine the analysis working conditions, and use the minimum and maximum states in the engine performance recording and air system cavity temperature and pressure measurement tests as the key analysis conditions for axial force; Obtain the design requirements for the engine rotor axial force, including the maximum axial force threshold Fmax and meshing force F1 in the intermediate state and the minimum axial force threshold Fmin and meshing force F2 in the idling state; Calibrate the calculation model of the air system before improvement and determine the flow area S of the front pre-swirl hole A , calculate the retrodiscal pressure PC; the specific calculation method is as follows:

[0057] In the above formula, π is the ratio of pi, R2 is the radius of the grate tooth tip, and R1 is the rear axle radius corresponding to the brush seal assembly; Add a grate sealing structure, preliminarily determine the grate tip radius R2, and determine the maximum and minimum pressure thresholds PC1 and PC2 behind the turbine disc; The calculation is as follows: PC1=F15 / S A PC2=F16 / S A F15=F max -F1-F3-F5+F7+F9-F11-F13 F16=F min -F2-F4-F6+F8+F10-F12-F14 In the above formula, F13 is the intermediate state turbine disk rear axial force, F14 is the slow state turbine disk rear axial force, F3 is the intermediate state compressor flow passage axial force, F4 is the slow state compressor flow passage axial force, F5 is the intermediate state compressor disk cavity axial force, F6 is the slow state compressor disk cavity axial force, F7 is the intermediate state turbine flow passage axial force, F8 is the slow state turbine flow passage axial force, F9 is the intermediate state turbine disk front axial force, F10 is the slow state turbine disk front axial force, F11 is the intermediate state turbine rear shaft axial force, F12 is the slow state turbine rear shaft axial force, F13 is the intermediate state turbine disk rear cavity axial force, F14 is the slow state turbine disk rear cavity axial force, PC1 is the maximum valve value of the disk rear pressure, and PC2 is the minimum valve value of the disk rear pressure; obtaining the minimum working gap δ1 of the intermediate state labyrinth min determining that the turbine disk rear PC meets the maximum valve value design requirement; obtaining the maximum working gap δ2 of the slow state labyrinth max determining that the slow state disk rear pressure PC meets the minimum valve value requirement; calculating the theoretical working gap difference △δ of the labyrinth in the intermediate state and the slow state, so as to meet: δ2 max -△δ<δ1<δ1 min determining the optimal design gap δ1 of the intermediate state labyrinth and the labyrinth tooth tip radius R2; calculating the cold state gap δ of the labyrinth, and determining the cold state size S3 of the labyrinth corresponding honeycomb.

[0058] The calibration improved air system calculation model is used to determine the flow area S of the front pre-swirl hole A calculating the disk rear cavity pressure PC, wherein the turbine disk rear sealing flow G1 and the turbine disk front minimum critical sealing flow G1 min need to be determined, if G1>G1 min , it indicates that the turbine disk rear sealing flow meets the design requirement, if G1 min , it indicates that the disk rear sealing flow does not meet the design requirement, the flow area S of the front pre-swirl hole is gradually expanded A , until G1>G1 min , a new air system calculation result is obtained, and a new air system calculation network model and the disk rear cavity pressure PC are formed.

[0059]

[0060]

[0061] wherein N is the rotating speed, is the air flow density, is the turbine disk rear rim sealing radius, viscosity of the fluid, K is the empirical coefficient.

[0062] the minimum working gap δ1 of the intermediate state labyrinth min , determine the turbine disc rear PC meets the maximum threshold design requirements; Specifically according to the following method is realized: directly to the intermediate state labyrinth, if PC < PC1, indicating that the intermediate state axial force meets the design requirements, if PC > PC1, indicating that the intermediate state axial force exceeds the maximum axial force bearing requirements, need to gradually increase the labyrinth gap δ1 to reduce PC pressure, until PC < PC1; When adjusting so that PC < PC1, obtain the intermediate state air system of new calculation model and the minimum working gap δ1 of the labyrinth in the intermediate state min , while carrying out parts temperature and deformation calculation, obtain the intermediate state labyrinth corresponding rotor and stator deformation and ; the maximum working gap δ2 of the slow car state labyrinth max , determine the slow car state disc rear pressure PC meets the minimum threshold requirements; Specifically according to the following method is realized: for the slow car state labyrinth, if PC > PC2, indicating that the slow car state axial force meets the design requirements, if PC < PC2, indicating that the disc rear pressure PC does not meet the slow car state pressure requirements, slow car state axial force light load, need to gradually reduce the labyrinth gap δ2 to increase PC pressure, until PC > PC2, obtain the slow car state air system of new calculation model and the maximum working gap δ2 of the labyrinth in the slow car state max , while carrying out parts temperature and deformation calculation, obtain the slow car state labyrinth corresponding rotor and stator deformation and ; the difference Δδ of the theoretical working gap of the labyrinth in the intermediate state and the slow car state, according to the following method is calculated:

[0063] , wherein: the rotor deformation of the intermediate state sealing labyrinth, the rotor deformation of the slow car state sealing labyrinth, the stator deformation corresponding to the intermediate state sealing labyrinth, the stator deformation corresponding to the slow car state sealing labyrinth.

[0064] the cold state gap δ of the labyrinth, determine the cold state size S3 of the labyrinth corresponding honeycomb, according to the following:

[0065]

[0066] In the above formula, δ1 is the optimum design gap of the intermediate state labyrinth, is the rotor deformation of the intermediate state labyrinth, and R2 is the labyrinth tip radius. In the above formula, δ1 is the optimum design gap of the intermediate state labyrinth,

[0067] The above is the embodiment of the enumerated embodiments, but the embodiment is not limited to the above optional embodiments, and those skilled in the art can obtain other various embodiments by arbitrarily combining the above described modes with each other. Any person can obtain other various forms of embodiments under the inspiration of the embodiment. The above specific embodiments should not be understood as a limitation on the protection scope of the embodiment, and the protection scope of the embodiment should be defined by the claims.

Claims

1. A sealing cooling structure for regulating the axial force of an aircraft engine rotor, characterized by: The invention comprises a turbine disc (8), a rear baffle (9) being provided on the turbine disc (8), and a low-pressure turbine guide vane (3) being provided opposite to the turbine disc (8), a low-pressure front inner support ring (5) and a low-pressure rear inner support ring (6) being provided in the middle of the low-pressure turbine guide vane (3), an air collecting cavity being formed between the low-pressure front inner support ring (5) and the low-pressure rear inner support ring (6), and an air supply channel being formed between the low-pressure front inner support ring (5), the rear baffle (9), the turbine disc (8) and the low-pressure rear inner support ring (6); The rear baffle (9) is formed with grate teeth (13) that correspond to the low-conductance front inner support plate and form a radial grate tooth (13) gap. When the airflow gathered in the air collecting chamber reaches a set amount, the airflow enters the rotor chamber from the low-pressure pre-swirl nozzle (11) of the air collecting chamber and is embedded in the air supply channel from the front pre-swirl hole (12). A part of the airflow entering the air supply channel is output outward through the grate tooth (13) gap and enters the disc rear chamber, and the other part enters the rotor chamber through the brush seal assembly (7) at the low-conductance rear inner support ring (6).

2. The sealing cooling structure for regulating the axial force of an aircraft engine rotor according to claim 1, characterized in that: A grate (13) plate extending toward the low-guide front inner support ring (5) is formed on the rear baffle (9), a plurality of grate teeth (13) are formed on the grate (13) plate, and a radial grate (13) gap between the grate (13) plate and the low-guide front inner support ring (5) is 0.6 mm to 0.7 mm.

3. The sealing cooling structure for regulating the axial force of an aircraft engine rotor according to claim 2, characterized in that: The radial grate teeth (13) have a gap of 0.65 mm.

4. The sealing cooling structure for regulating the axial force of an aircraft engine rotor according to claim 1, characterized in that: A guide pipe (4) communicating with the low-pressure turbine guide vane (3) is provided between the low-pressure front inner support ring (5) and the low-pressure rear inner support ring (6), and the airflow in the low-pressure turbine guide vane (3) enters the air collecting cavity through the guide pipe (4).

5. The sealing cooling structure for regulating the axial force of an aircraft engine rotor according to claim 1, characterized in that: A rear shaft (14) is provided on the turbine disc (8), and a sealing surface (15) is formed on the rear shaft (14) to cooperate with the brush seal assembly (7).

6. The sealing cooling structure for regulating the axial force of an aircraft engine rotor according to claim 1, characterized in that: The outer edge of the low-pressure turbine guide vane (3) is connected to the air bleed pipe (1), and an annular cavity is formed between the low-pressure turbine guide vane (3) and the air bleed pipe (1). The air bleed pipe (1) is used to connect to the compressor and transport air flow into the annular cavity.

7. A seal cooling method for regulating the axial force of an aircraft engine rotor, using the seal cooling structure according to any one of claims 1 to 6, characterized in that: include: Determine the analysis working conditions, and use the minimum and maximum states in the engine performance recording and air system cavity temperature and pressure measurement tests as the key analysis working conditions for axial force; Obtain the design requirements for the engine rotor axial force, including the maximum axial force threshold Fmax and meshing force F1 in the intermediate state and the minimum axial force threshold Fmin and meshing force F2 in the idle state; Calibrate the calculation model of the air system before improvement and determine the flow area S of the front pre-swirl hole A , calculate the retrodiscal pressure PC; the specific calculation method is as follows: In the above formula, π is the ratio of pi, R2 is the radius of the grate tooth tip, and R1 is the rear axle radius corresponding to the brush seal assembly; Add a grate sealing structure, preliminarily determine the grate tip radius R2, and determine the maximum and minimum pressure thresholds PC1 and PC2 behind the turbine disc; The calculation is as follows: PC1=F15 / S A PC2=F16 / S A F15=F max -F1-F3-F5+F7+F9-F11-F13 F16=F min -F2-F4-F6+F8+F10-F12-F14 In the above formula, F13 is the axial force behind the turbine disk in the intermediate state, F14 is the axial force behind the turbine disk in the idle state, F3 is the axial force of the compressor flow path in the intermediate state, F4 is the axial force of the compressor flow path in the idle state, F5 is the axial force of the compressor disk cavity in the intermediate state, F6 is the axial force of the compressor disk cavity in the idle state, F7 is the axial force of the turbine flow path in the intermediate state, F8 is the axial force of the turbine flow path in the idle state, F9 is the axial force in front of the turbine disk in the intermediate state, F10 is the axial force in front of the turbine disk in the idle state, F11 is the axial force of the turbine rear shaft in the intermediate state, F12 is the axial force of the turbine rear shaft in the idle state, F13 is the axial force of the turbine disk rear cavity in the intermediate state, F14 is the axial force of the turbine rear cavity in the idle state, PC1 is the maximum pressure threshold behind the disk, and PC2 is the minimum pressure threshold behind the disk; Get the minimum working gap δ1 of the intermediate state grate teeth min , confirm that the PC behind the turbine disc meets the maximum threshold design requirements; Get the maximum working clearance δ2 of the grate teeth at slow speed max , confirm that the idle state after-disc pressure PC meets the minimum threshold requirement; Calculate the theoretical working clearance difference △δ between the grate teeth in the intermediate state and the slow running state to satisfy: d2 max -△δ<δ1<δ1 min Determine the optimal design gap δ1 and grate tooth tip radius R2 of the intermediate state grate teeth; Calculate the cold gap δ of the grate teeth and determine the cold size S3 of the honeycomb corresponding to the grate teeth.

8. The seal cooling method for regulating the axial force of an aircraft engine rotor according to claim 7, characterized in that: The calibration improves the air system calculation model before determining the flow area S of the front pre-swirl hole A , calculate the rear cavity pressure PC, which requires determining the turbine disc rear seal flow G1 and the turbine disc front minimum critical seal flow G1 min , if G1>G1 min Indicates that the seal flow rate after the turbine disc meets the design requirements. If G1<G1 min Indicates that the flow rate after the disc seal does not meet the design requirements, and gradually expand the flow area S of the front pre-swirl hole A , until G1>G1 min , obtain new air system calculation results, and at the same time form a new air system calculation network model and rear disc cavity pressure PC; Where N is the rotation speed, is the air flow density, is the sealing radius of the rear rim of the turbine disc, is the dynamic viscosity, is the swirl coefficient of the airflow, and K is the empirical coefficient.

9. The seal cooling method for regulating the axial force of an aircraft engine rotor according to claim 7, characterized in that: The minimum working gap δ1 of the intermediate state grate teeth is obtained min , confirm that the PC behind the turbine disc meets the maximum threshold design requirements; this is achieved specifically by the following method: Facing the intermediate state grate teeth, if PC < PC1, it means that the intermediate state axial force meets the design requirements. If PC > PC1, it means that the intermediate state axial force exceeds the maximum axial force requirement of the bearing. It is necessary to gradually increase the gap δ1 of the grate teeth to reduce the PC pressure until PC < PC1. When the adjustment makes PC < PC1, the new calculation model of the intermediate state air system and the minimum working gap δ1 of the grate teeth in the intermediate state are obtained. min , and simultaneously carry out the calculation of the temperature and deformation of the parts to obtain the rotor and stator deformation corresponding to the intermediate state of the grate teeth and ; The maximum working clearance δ2 of the grate teeth in the slow state is obtained max , determine whether the pressure PC behind the disc in the slow-running state meets the minimum threshold requirement; specifically, this is achieved as follows: for the slow-running state grate, if PC>PC2, it means that the axial force in the slow-running state meets the design requirements; if PC<PC2, it means that the pressure PC behind the disc does not meet the slow-running state pressure requirements, and the axial force in the slow-running state is lightly loaded. It is necessary to gradually reduce the gap δ2 of the grate to increase the PC pressure until PC>PC2, and obtain a new calculation model of the slow-running state air system and the maximum working gap δ2 of the grate in the slow-running state. max , and simultaneously carry out part temperature and deformation calculations to obtain the rotor and stator deformation corresponding to the slow workshop state grate teeth and ; The theoretical working clearance difference Δδ between the grate teeth in the intermediate state and the slow running state is calculated as follows: in: The rotor deformation of the sealing grate teeth in the intermediate state, The rotor deformation of the sealing grate teeth in the slow running state, is the stator deformation corresponding to the sealing grate teeth in the intermediate state, It is the stator deformation corresponding to the sealing grate teeth in the slow running state.

10. The seal cooling method for regulating the axial force of an aircraft engine rotor according to claim 7, characterized in that: The calculation of the cold clearance δ of the grate teeth and the determination of the cold size S3 of the honeycomb corresponding to the grate teeth are carried out as follows: In the above formula, δ1 is the optimal design gap of the grate teeth in the intermediate state, is the stator deformation corresponding to the sealing grate teeth in the intermediate state, is the rotor deformation of the sealing grate teeth in the intermediate state, and R2 is the grate tooth tip radius.

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