CN112629840A - Aircraft engine double-rotor-support-casing tester and testing method thereof - Google Patents
Aircraft engine double-rotor-support-casing tester and testing method thereof Download PDFInfo
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- G—PHYSICS
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- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M13/00—Testing of machine parts
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01H—MEASUREMENT OF MECHANICAL VIBRATIONS OR ULTRASONIC, SONIC OR INFRASONIC WAVES
- G01H17/00—Measuring mechanical vibrations or ultrasonic, sonic or infrasonic waves, not provided for in the preceding groups
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01M—TESTING STATIC OR DYNAMIC BALANCE OF MACHINES OR STRUCTURES; TESTING OF STRUCTURES OR APPARATUS, NOT OTHERWISE PROVIDED FOR
- G01M1/00—Testing static or dynamic balance of machines or structures
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- Y02T—CLIMATE CHANGE MITIGATION TECHNOLOGIES RELATED TO TRANSPORTATION
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Abstract
The invention belongs to the technical field of aero-engine tests and tests, in particular to an aero-engine double-rotor-support-casing tester and a test method thereof, wherein the tester comprises a low-pressure rotor, a central bevel gear system, a high-pressure rotor, a low-pressure casing, an intermediate casing and a high-pressure casing; the low-pressure rotor comprises a low-pressure turbine shaft, and the low-pressure turbine shaft penetrates through the central bevel gear system and is in rotary connection with the high-pressure rotor; the low-voltage rotor is driven by a low-voltage driving motor, the central bevel gear system is connected with a high-voltage driving motor, and the high-voltage rotor is driven by the central bevel gear system. The invention has the advantages that the structure is simplified and reasonable, the design of a scale with fixed proportion (such as 2:1) is adopted according to the structure of a real engine, and the structures of the fan four-stage disc and the high-pressure nine-stage disc are consistent with those of the real engine. The structure of a squirrel cage, a conical shell, a bearing amplitude plate, a flexible casing and the like in a real engine is reserved, the structure of an engine rotor system and the vibration of the engine rotor system can be truly simulated, and the requirement of the whole-engine vibration research of an aeroengine is met.
Description
Technical Field
The invention belongs to the technical field of aero-engine tests and tests, and particularly relates to a dual-rotor-support-casing tester for an aero-engine and a test method thereof.
Background
Along with the increase of thrust-weight ratio of the aircraft engine, the structure of the rotor system is greatly changed, and the high rotating speed of the aircraft engine makes the vibration problem of the aircraft engine more prominent. The aerodynamic coupling characteristic of the rotor of the aircraft engine and the transmission of the unbalanced vibration of the rotor comprise the influence of vibration transmission rules of the inner transmission and the outer transmission of the aircraft engine and the influence of hanging points on vibration transmission, the coupling vibration characteristic of a disk drum, the influence of gear meshing excitation and the like on the unbalanced vibration of the rotor, the vibration rules of vibration participation of a casing and the like are more complex, and the mechanism and the rules have a plurality of uncertain places to be researched urgently. In particular, the study of the imbalance vibration between the supporting points of the rotor and the vibration transmission law of the rotor-support-casing is very important and urgent because the imbalance vector and the imbalance vibration law of the rotor are not clear after the rotor is assembled or operated for a period of time, and because the measurement cannot be directly performed on the internal rotor, but only on the aeroengine casing.
The internal rotor and the supporting structure of the aircraft engine are complex, and the existing testing means cannot directly obtain the vibration data of the rotor structure body on the real engine; the numerical simulation and emulation technology can not determine the internal vibration state data of the aircraft engine with high precision, and the rule that the vibration of the internal rotor of the engine is transmitted to the outside and the final vibration performance on an external measuring point are more difficult to accurately analyze.
The aviation real engine has a complex structure and more influence factors, and the experimental research of the aviation engine is limited; the development of the domestic rotor tester mostly adopts a structural form of rigid support of a disc shaft, the influence of a disc-drum coupling structure and an elastic support on the vibration characteristic of the rotor is not considered, a simplified model of the rotor tester is greatly different from a real machine structure, and the vibration characteristic of a real engine cannot be well simulated. Therefore, based on the theory of structural similarity and dynamics similarity, the double-rotor-support-casing tester and the testing method for the unbalanced vibration transmission rule research have innovation and important significance.
Disclosure of Invention
In order to solve the problems in the background art, the invention provides a double-rotor-bearing-casing tester of an aircraft engine and a testing method thereof, which can effectively realize vibration transmission and testing among a rotor, a supporting structure and a casing.
A first object of the present invention is to provide an aircraft engine double rotor-bearing-casing tester comprising a low pressure rotor, a central bevel gear system and a high pressure rotor arranged laterally in sequence, wherein: the low-pressure rotor, the central bevel gear system and the high-pressure rotor are fixed on the frame through a supporting structure, a low-pressure casing is arranged on the periphery of the low-pressure rotor, an intermediate casing is arranged on the periphery of the central bevel gear system, a high-pressure casing is arranged on the periphery of the high-pressure rotor, and the low-pressure casing, the intermediate casing, the high-pressure casing, the low-pressure rotor and the high-pressure rotor are coaxially arranged and supported by the supporting structure;
the low-pressure rotor comprises a low-pressure turbine shaft, and the low-pressure turbine shaft penetrates through the central bevel gear system and is in rotary connection with the high-pressure rotor and the central bevel gear system;
the low-voltage rotor is driven by a low-voltage driving motor, the central bevel gear system is connected with a high-voltage driving motor, and the high-voltage rotor is driven by the central bevel gear system.
Further, the low-pressure rotor comprises a low-pressure fan section rotor and a low-pressure turbine rotor, the low-pressure fan section rotor comprises a fan rotating shaft and a four-stage fan disc sleeved on the fan rotating shaft, the low-pressure turbine rotor comprises a low-pressure turbine shaft, a low-pressure turbine conical shell and a low-pressure turbine disc sleeved on the low-pressure turbine shaft, the low-pressure turbine conical shell is fixedly installed at the outer end of the low-pressure turbine disc, the fan rotating shaft is connected with the low-pressure turbine shaft through a coupler, and the low-pressure driving motor is connected.
Further, the high-pressure rotor comprises a high-pressure forearm shaft, a 9-stage wheel disc of the high-pressure compressor, a high-pressure turbine shaft, a high-pressure turbine disc, a high-pressure turbine conical shell and an intermediate bearing, wherein:
the high-pressure forearm shaft is a hollow shaft, the front end of the high-pressure forearm shaft is connected with the central bevel gear system, and the rear end of the high-pressure forearm shaft is connected with a 9-stage wheel disc of the high-pressure compressor; the high-pressure turbine disc and a part of 9-stage wheel disc of the high-pressure compressor are rigidly connected through a high-pressure turbine shaft, and a high-pressure turbine conical shell is fixedly connected to the outer end of the high-pressure turbine disc;
the high-pressure turbine cone shell is connected with the low-pressure turbine shaft through an intermediate bearing.
Further, the central bevel gear system comprises a speed increaser, a long coupler, a central bevel gear box, a driving bevel gear and a driven bevel gear, wherein: the speed increaser is connected with a high-voltage driving motor, the speed increaser drives a driving bevel gear through a long coupler, the driving bevel gear is meshed with a driven bevel gear, and the driven bevel gear is connected with a high-voltage front arm shaft in a matched manner so as to drive a high-voltage rotor to rotate.
Furthermore, the supporting structure comprises a first supporting structure, a second supporting structure, a third supporting structure and a fourth supporting structure, the first supporting structure and the fourth supporting structure are used for supporting the low-pressure rotor, the second supporting structure is used for supporting the low-pressure rotor and the central bevel gear box, the third supporting structure is used for supporting the high-pressure rotor, and the third supporting structure is an elastic supporting structure.
Furthermore, the first supporting structure comprises a No. 1 amplitude plate, a No. 1 bearing conical shell, a No. 1 bearing seat, a No. 1 squirrel cage and a No. 1 bearing, and one end of the rotating shaft of the fan is rotationally connected with the first supporting structure through the No. 1 bearing;
the second supporting structure comprises a No. 2 amplitude plate, a No. 2 bearing seat and a No. 2 bearing, and the other end of the fan rotating shaft is rotatably connected with the second supporting structure through the No. 2 bearing;
the third supporting structure comprises a No. 3 amplitude plate, a No. 3 bearing conical shell, a No. 3 squirrel cage, a No. 3 bearing seat and a No. 3 bearing, and the middle part of the high-pressure front shaft arm is rotationally connected with the third supporting structure through the No. 3 bearing;
the fourth supporting structure comprises a No. 4 amplitude plate, a No. 4 bearing conical shell, a No. 4 squirrel cage, a No. 4 bearing seat and a No. 4 bearing, and the low-pressure turbine shaft is rotatably connected with the fourth supporting structure through the No. 4 bearing.
Furthermore, the low-voltage casing is fixed on the No. 1 radial plate and the No. 2 radial plate, the middle casing is fixed on the No. 2 radial plate and the No. 3 radial plate, and the high-voltage casing is fixed on the No. 3 radial plate and the No. 4 radial plate.
Furthermore, the frame includes rotor system frame and high-pressure transmission system frame, and low pressure driving motor fixes in the rotor system frame, and high pressure driving motor and speed increaser fix in the high-pressure transmission system frame.
Furthermore, a plurality of bolt holes are formed in the periphery of the edge of the four-stage fan disc, the 9-stage wheel disc of the high-pressure compressor, the low-pressure turbine disc and the high-pressure turbine disc.
A second object of the present invention is to provide a testing method based on any one of the above-mentioned aircraft engine dual rotor-bearing-casing testers, comprising the following steps:
(1) adopting Latin hypercube group sampling to determine the amount and phase of unbalance to be added;
(2) setting the rotating speed of a low-voltage driving motor, enabling a low-voltage rotor to rotate and a high-voltage rotor not to rotate, and extracting vibration data of each test point;
(3) setting the rotating speed of a high-voltage driving motor to enable a high-voltage rotor to rotate and a low-voltage rotor not to rotate, and extracting vibration data of each test point;
(4) setting the rotating speed of a low-voltage driving motor and the rotating speed of a high-voltage driving motor at the same time, enabling the low-voltage rotor and the high-voltage rotor to rotate, and extracting vibration data of each test point;
(5) repeating the steps (1) to (4) until 100 groups of sampling unbalanced working conditions are completed;
(6) and (3) carrying out self-adaptive filtering on the measuring point signals, removing interference components, converting the acceleration signals into vibration speed through numerical integration, converting the displacement signals into vibration displacement through numerical differentiation, and unifying vibration data dimensions.
(7) And training the obtained test data on a neural network model by using a deep learning algorithm, and analyzing the vibration transfer coupling rule of the double rotors, the support and the casing.
Compared with the prior art, the invention has the following beneficial effects:
(1) the aircraft engine double-rotor-support-casing tester provided by the invention has a high-low pressure double-rotor coupling structure, and a low-pressure fan, a low-pressure turbine, a high-pressure compressor 9-stage wheel disc and a high-pressure turbine disc are mutually coupled through an intermediate bearing, so that the coupling characteristic of a double-rotor aircraft engine can be better reflected.
(2) The aircraft engine double-rotor-support-casing tester is simple and reasonable in structure and comprehensive in consideration. The design of a scale with a fixed ratio (such as 2:1) is adopted according to the structure of a real engine, and the structures of a fan four-stage disc and a high-pressure compressor 9-stage disc are consistent with those of the real engine. The structure of the squirrel cage, the conical shell, the bearing web plate, the flexible casing and the like in the real machine is reserved. Compared with the domestic rotor tester, most of the domestic rotor testers adopt a structural form of rigid support of a disc shaft, the influence of a disc-drum coupling structure, double-rotor high-low pressure coupling and an elastic support on the vibration characteristic of the rotor is not considered, the structure and the vibration of an engine rotor system can be truly simulated, and the requirement of the whole machine vibration research of an aeroengine is met.
(3) The double-rotor-support-casing tester for the aircraft engine adopts a central bevel gear system, a real aircraft engine high-pressure rotor is taken as a power output end, and a motor drives a central bevel gear so as to drive the high-pressure rotor to operate. The high-pressure rotor is driven to operate, and the operation of the central bevel gear system is simulated.
(4) The aircraft engine double-rotor-bearing-casing tester provided by the invention has a supporting structure of a squirrel cage-conical shell-amplitude plate, is similar to a real engine supporting structure, comprises an elastic squirrel cage with a vibration damping effect, simplifies the conical shell and amplitude plate structures in a real engine, and is consistent with a real engine vibration transmission path.
(5) The aircraft engine double-rotor-support-casing tester provided by the invention is convenient for applying unbalance. A plurality of bolt holes are designed in the edge of a four-stage fan disc, a 9-stage wheel disc of a high-pressure compressor, a low-pressure turbine disc and a wheel disc of a high-pressure turbine disc in the tester along the circumferential direction, so that unbalanced mass can be conveniently applied.
(6) The invention relates to an aircraft engine double-rotor-support-casing tester and a testing method thereof, which can measure vibration data of a plurality of positions relative to a real engine.
Drawings
In order to more clearly illustrate the embodiments of the present invention or the technical solutions in the prior art, the drawings used in the description of the embodiments or the prior art will be briefly described below, it is obvious that the drawings in the following description are only some embodiments of the present invention, and for those skilled in the art, other drawings can be obtained according to the drawings without creative efforts.
FIG. 1 is an isometric view of an aircraft engine dual rotor-bearing-case tester of the present invention (inorganic case);
FIG. 2 is a general block diagram (without a case) of the aircraft engine dual rotor-bearing-case tester of the present invention;
FIG. 3 is a top view of an aircraft engine dual rotor-bearing-case tester of the present invention (inorganic case);
FIG. 4 is a cross-sectional view of an aircraft engine dual rotor-bearing-case tester of the present invention;
FIG. 5 is a schematic view of a low-pressure rotor structure of a dual-rotor-bearing-casing tester for an aircraft engine according to the present invention;
FIG. 6 is a schematic view of a high-pressure rotor structure of a dual-rotor-bearing-casing tester of an aircraft engine according to the present invention;
FIG. 7 is a schematic view of a central bevel gear system of the dual rotor-bearing-casing tester of the aircraft engine of the present invention;
FIG. 8 is a schematic view of a first support structure of the dual rotor-bearing-casing tester of the aircraft engine of the present invention;
FIG. 9 is a schematic view of a second support structure of the dual rotor-bearing-casing tester for an aircraft engine according to the present invention;
FIG. 10 is a schematic view of a third support structure of the dual rotor-bearing-casing tester for an aircraft engine according to the present invention;
FIG. 11 is a schematic view of a fourth support structure of the dual rotor-bearing-casing tester for an aircraft engine according to the present invention;
FIG. 12 is a schematic view of a low-pressure case structure of a dual rotor-bearing-case tester for an aircraft engine according to the present invention;
FIG. 13 is a schematic view of an intermediate case structure of a dual rotor-bearing-case tester for an aircraft engine according to the present invention;
FIG. 14 is a schematic diagram of a high-pressure casing structure of a dual rotor-bearing-casing tester for an aircraft engine according to the present invention;
FIG. 15 is a schematic diagram of a frame structure of a dual rotor-bearing-casing tester for an aircraft engine according to the present invention;
wherein: 1-low pressure rotor, 11-fan rotating shaft, 12-fan four-stage disk, 13-low pressure turbine shaft, 14-low pressure turbine cone shell, 15-low pressure turbine disk, 16-low pressure driving motor, 17-set gear coupling, 2-high pressure rotor, 21-high pressure forearm shaft, 22-high pressure compressor 9-stage wheel disk, 221-1 stage sealing disk, 23-high pressure turbine shaft, 24-high pressure turbine disk, 25-high pressure turbine cone shell, 26-intermediate bearing, 3-central cone gear system, 31-speed increaser, 32-long coupling, 33-central cone gear box, 34-driving bevel gear, 35-driven bevel gear, 4-supporting mechanism, 41-first supporting structure, 411-1 amplitude plate, 412-1 number bearing cone shell, 413-1 bearing seat, 414-1 squirrel cage, 415-1 bearing, 416-1 hanging point, 42-second support structure, 421-2 amplitude plate, 422-2 bearing seat, 423-2 bearing, 424-2 hanging point, 43-third support structure, 431-3 amplitude plate, 432-3 bearing conical shell, 433-3 squirrel cage, 434-3 bearing seat, 435-3 bearing, 436-3 hanging point, 44-fourth support structure, 441-4 amplitude plate, 442-4 bearing conical shell, 443-4 squirrel cage, 444-4 bearing seat, 445-4 bearing, 446-4 hanging point, 5-frame, 51-rotor system frame, 511-foot, 512-low-voltage motor seat, 52-high-voltage transmission system frame, 521-a high-voltage motor base, 522-an accelerator base, 6-a low-voltage casing, 7-a middle casing, 8-a high-voltage casing and 9-a high-voltage driving motor.
Detailed Description
The technical solutions in the embodiments of the present invention will be described clearly and completely with reference to the accompanying drawings in the present invention, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all of the embodiments. All other embodiments, which can be derived by a person skilled in the art from the embodiments given herein without making any creative effort, shall fall within the protection scope of the present invention.
As shown in fig. 1 to 3, the present invention provides an aircraft engine double rotor-support-casing tester comprising a low pressure rotor 1, a central bevel gear system 3 and a high pressure rotor 2 arranged in transverse sequence, wherein: the low-voltage rotor 1, the central bevel gear system 3 and the high-voltage rotor 2 are fixed on a rack 5 through a supporting structure 4, a low-voltage casing 6 is arranged on the periphery of the low-voltage rotor 1, an intermediate casing 7 is arranged on the periphery of the central bevel gear system 3, a high-voltage casing 8 is arranged on the periphery of the high-voltage rotor 2, and the low-voltage casing 6, the intermediate casing 7, the high-voltage casing 8, the low-voltage rotor 1 and the high-voltage rotor 2 are coaxially arranged and supported by the supporting structure 4; the low-pressure rotor 1 comprises a low-pressure turbine shaft 13, the low-pressure turbine shaft 13 penetrates through the central bevel gear system 3 and is in rotary connection with the high-pressure rotor 2 and the central bevel gear system 3, and a 5# fulcrum and a 3# fulcrum are formed; the low voltage rotor 1 is driven by a low voltage drive motor 16, the central bevel gear system 3 is connected to a high voltage drive motor 9, and the high voltage rotor 2 is driven by the central bevel gear system 3. The low-pressure rotor 1 and the high-pressure rotor 2 are coupled with each other, so that the coupling characteristic of the double-rotor aircraft engine can be better reflected.
Specifically, as shown in fig. 5, the low-pressure rotor 1 of the invention includes a low-pressure fan section rotor and a low-pressure turbine rotor, the low-pressure fan section rotor includes a fan rotating shaft 11 and a four-stage fan disc 12 sleeved on the fan rotating shaft 11, and is similar to a prototype fan section, the tester low-pressure fan section rotor is also called a low-pressure compressor, and is mainly characterized in that a 1-stage wheel disc and a 1-stage drum are combined with each other to form a fan 1-stage drum, the 2-stage drum is similar to the 3-stage drum, the wheel disc is a hollow disc, and the drum is a thin-wall cylinder. The fourth-stage fan disc 12 comprises three sections, wherein the first section comprises a fan 1-stage disc drum and a fan 2-stage disc drum, the second section comprises a 3-stage disc drum, the third section comprises a 4-stage wheel disc, the first-stage disc drum is integrally welded to form a structural member, the second-stage disc drum is the same as the first-stage disc drum, and the first-stage disc drum, the second-stage disc drum and the third-stage disc drum are all connected through bolts; the low-pressure turbine rotor comprises a low-pressure turbine shaft 13, a low-pressure turbine conical shell 14 and a low-pressure turbine disc 15, wherein the low-pressure turbine conical shell 14 and the low-pressure turbine disc 15 are sleeved on the low-pressure turbine shaft 13, the low-pressure turbine conical shell 14 is fixedly installed at the outer end of the low-pressure turbine disc 15, the fan rotating shaft 11 is connected with the low-pressure turbine shaft 13 through a sleeve gear coupling 17, and the low-pressure driving motor 16 is connected with the fan rotating shaft 11 through a.
Specifically, as shown in fig. 6, the high-pressure rotor 2 of the present invention includes a high-pressure forearm shaft 21, a high-pressure compressor 9-stage wheel disc 22, a high-pressure turbine shaft 23, a high-pressure turbine disc 24, a high-pressure turbine cone shell 25 and a middle bearing 26, wherein: the high-pressure front arm shaft 21 is a hollow shaft, the front section of the shaft is provided with a spline and is connected with the central gear system 3 through the spline, and the rear section of the shaft is connected with the 9-stage wheel disc 22 of the high-pressure compressor through a bolt; the 9-stage wheel disc 22 of the high-pressure compressor is divided into three sections, wherein the first section comprises a front three-stage disc drum, the third-stage disc drum comprises a conical shell structure similar to the original machine type, the second section comprises 4, 5 and 6-stage disc drums, a drum barrel of the 6 th-stage disc drum comprises a conical shell structure similar to the original machine type, and the third section consists of 7, 8 and 9-stage disc drums and a 1-stage sealing disc 221; the first section of the disc drum is integrally welded to form a structural member, the second section is the same as the first section, the first section is connected with the second section in a bolt mode, and the second section is connected with the third section in a long bolt mode; the high-pressure turbine disc 24 is rigidly connected with a part of the high-pressure compressor 9-stage wheel disc 22 through a high-pressure turbine shaft 23, and the high-pressure turbine conical shell 25 is fixedly connected with the high-pressure turbine disc 24; the high-pressure turbine cone shell 25 and the low-pressure turbine shaft 13 are connected by an intermediate bearing 26, wherein a 5# fulcrum is formed, the 5# fulcrum is an intermediate fulcrum, and the intermediate bearing 26 is preferably a roller bearing. It will be appreciated that the low-pressure turbine shaft 13 is coaxial with the high-pressure rotor 2 and extends through the high-pressure rotor 2 and is coupled to the high-pressure rotor by means of the intermediate bearing 26, so as to better reflect the coupling characteristics of the twin-rotor aircraft engine.
Specifically, as shown in fig. 7, the central bevel gear system 3 of the present invention utilizes a central bevel gear structure to realize power driving of a high-pressure rotor, and mainly includes a speed increaser 31, a long coupler 32, a central bevel gear box 33, a driving bevel gear 34 and a driven bevel gear 35, wherein the driving direction of the central bevel gear system is opposite to that of the prototype, but the central bevel gear structure is similar to that of the prototype, and the transmission ratio of the central bevel gear is consistent with that of the prototype, so as to ensure that the stresses of the high-pressure front arm shaft 21 and the high-pressure turbine shaft 23 are consistent. Meanwhile, in consideration of the problem of limited structure of the tester, the midspan mode of the driving bevel gear 34 of the prototype is changed into a cantilever mode, a two-point supporting long-span shaft is utilized to ensure the stability of the driving bevel gear 34, and the central bevel gear box 33 is fixed on the supporting structure 4 through bolts and is also fixed on the frame 5. The speed increaser 31 is connected with a high-voltage driving motor 9 through a coupler, the speed increaser 31 drives a driving bevel gear 34 through a long coupler 32, the driving bevel gear 34 is meshed with a driven bevel gear 35, and the driven bevel gear 35 is matched and connected with a high-voltage forearm shaft 21 through a spline to drive a high-voltage rotor 2 to rotate so as to complete power transmission of a high-voltage part. The design of the central bevel gear system is consistent with the power transmission path of a real engine, and the direction is opposite.
Specifically, as shown in fig. 8 to 14, the support structure 4 mainly provides a fulcrum for the fan rotating shaft 11, the low-pressure turbine shaft 13, the high-pressure forearm shaft 21, and the like, and provides a hanging point for the low-pressure casing 6, the intermediate casing 7, the high-pressure casing 8, and the central bevel gear box 33 to fix the casings and the central bevel gear box. The support structure 4 comprises a first support structure 41, a second support structure 42, a third support structure 43 and a fourth support structure 44, the first support structure 41 and the fourth support structure 44 are used for supporting the low pressure rotor 1, the second support structure 42 is used for supporting the low pressure rotor 1 and the central bevel gear box 33, the third support structure 43 is used for supporting the high pressure rotor 2, and the third support structure 43 is an elastic support structure. The casing is a stator casing and is mainly used for simulating a thin-wall casing in a real engine.
Specifically, as shown in fig. 8 to 11, the first support structure 41 of the present invention includes a number 1 amplitude plate 411, a number 1 bearing cone shell 412, a number 1 bearing seat 413, a number 1 squirrel cage 414, a number 1 bearing 415, and a number 1 hanging point 416, wherein one end of the fan rotating shaft 11 is rotatably connected to the first support structure 41 through the number 1 bearing 415, which is a number 1 fulcrum; the second supporting structure 42 comprises a No. 2 web 421, a No. 2 bearing seat 422, a No. 2 bearing 423 and a No. 2 hanging point 424, and the other end of the fan rotating shaft 11 is rotatably connected with the second supporting structure 42 through the No. 2 bearing 422, which is a No. 2 fulcrum; the third supporting structure 43 comprises a 3# amplitude plate 431, a 3# bearing conical shell 432, a 3# squirrel cage 433, a 3# bearing seat 434, a 3# bearing 435 and a 3# hanging point 436, the high-pressure front shaft arm 21 is rotatably connected with the third supporting structure 43 through the 3# bearing 435, and a 4# fulcrum is arranged at the position; the preferable No. 3 bearing is a ball bearing; the fourth supporting structure 44 comprises a number 4 amplitude plate 441, a number 4 bearing conical shell 442, a number 4 squirrel cage 443, a number 4 bearing seat 444, a number 4 bearing 445 and a number 4 hanging point 446, and the low-pressure turbine shaft 13 is rotatably connected with the fourth supporting structure 44 through the number 4 bearing 445, which is a number 6 fulcrum. The first supporting structure 41 is fixed on the frame 5 through a number 1 hanging point 416, and the fixing mode can be bolt fixing; the second supporting structure 42 is fixed on the frame 5 through a number 2 hanging point 424, preferably by bolt; the third supporting structure 43 is fixed on the frame 5 through a number 3 hanging point 436, preferably in a bolt fixing manner; the fourth support structure 44 is fixed to the frame 5 by means of a number 4 hanger point 446, preferably by means of bolts. No. 1, No. 2, No. 3 and No. 1, No. 2 and No. 4 spokes 441 are large round discs with hollow middle parts, so that the rigidity is high and the disc is not easy to deform; the No. 1 bearing conical shell 412, the No. 3 bearing conical shell 432 and the No. 4 bearing conical shell 442 are all conical hollow structures, the rigidity is low, the larger end of the bearing conical shell is fixed on the amplitude plate through a bolt, the smaller end of the bearing conical shell is fixed with the squirrel cage through a bolt, the squirrel cage is of a cylindrical structure, the middle part of the squirrel cage is designed in a hollow mode and is changed into ribs, and therefore the squirrel cage is low in rigidity and good in elasticity; the other end of the squirrel cage is connected with the bearing seat through a bolt, and the elastic squirrel cage can play a role in damping vibration from the bearing seat. The supporting structure of the squirrel cage-conical shell-amplitude plate is similar to a supporting structure of a real engine, comprises the elastic squirrel cage with the vibration reduction function, simplifies the structures of the conical shell and the amplitude plate in the real engine, is consistent with the vibration transmission path of the real engine, and can better simulate the vibration condition of an aero-engine.
Specifically, as shown in fig. 4, the low-voltage casing 6 of the present invention is fixed on the number 1 web 411 and the number 2 web 421, the middle casing 7 is fixed on the number 2 web 421 and the number 3 web 431, and the high-voltage casing is fixed on the number 3 web 431 and the number 4 web 441.
Specifically, as shown in fig. 15, the rack 5 of the present invention includes a rotor system rack 51 and a high-pressure transmission system rack 52, and the rack 5 mainly provides an installation reference and a platform for the high-pressure rotor 2, the low-pressure rotor 1 and the transmission system thereof; the frame 5 is fixed on the ground through a machine leg 511 by using an expansion bolt, the low-voltage driving motor 16 is fixed on a low-voltage motor base 512 of the rotor system frame 51 through a bolt, the supporting structure 4 is fixed on the rotor system frame 51 through a bolt, the high-voltage driving motor 9 is fixed on a high-voltage motor base 521 of the high-voltage transmission system frame 52 through a bolt, and the speed increaser 31 is fixed on a speed increaser base 522 of the high-voltage transmission system frame 52 through a bolt. The frame 5 is of a frame structure, the frame is hollow, light in weight, and convenient to mount, and has certain rigidity.
Specifically, the four-stage fan disc 12, the high-pressure compressor 9 stage wheel disc 22, the low-pressure turbine disc 15 and the high-pressure turbine disc 24 are circumferentially provided with a plurality of bolt holes at the wheel disc edge, so that unbalanced mass can be conveniently applied, and the phase and the size of the unbalanced mass can be adjusted according to the positions of the added screws and the mass of the screws.
The aircraft engine double-rotor-support-casing tester is simple and reasonable in structure and comprehensive in consideration. The design of scaling (such as 2:1) is adopted according to the structure of a real engine, and the structures of a fan four-stage disc and a high-pressure nine-stage disc are consistent with those of the real engine. The structure of the squirrel cage, the conical shell, the bearing web plate, the flexible casing and the like in the real machine is reserved. Compared with the domestic rotor tester, most of the domestic rotor testers adopt a structural form of rigid support of a disc shaft, the influence of a disc-drum coupling structure, double-rotor high-low pressure coupling and an elastic support on the vibration characteristic of the rotor is not considered, the structure and the vibration of an engine rotor system can be truly simulated, and the requirement of the whole machine vibration research of an aeroengine is met.
The aircraft engine double-rotor-bearing-casing tester can effectively realize vibration transmission and testing among the rotor, the supporting structure and the casing, and has similarity with a real engine structure and similarity with dynamic characteristics compared with the existing tester.
The invention also provides a test method, which is mainly used for researching the birotor-support-casing vibration transfer coupling rule and comprises Latin hypercube group sampling, adaptive filtering, numerical calculus, deep learning algorithm and neural network model. The method mainly adopts Latin hypercube group sampling, randomly extracts 100 groups of unbalance and unbalance phase combinations, respectively arranges the unbalance and unbalance phase combinations at the positions of a four-stage fan disc 12, a high-pressure compressor 9-stage wheel disc 22, a high-pressure turbine disc 24 and a low-pressure turbine disc 15 for testing, extracts the vibration displacement from a 1# fulcrum to the position near a 6# fulcrum, extracts the vibration acceleration at the positions of a 1# amplitude plate 411, a 1# bearing conical shell 412, a 1# bearing seat 413, a 1# squirrel cage 414, a 2# amplitude plate 421, a 3# amplitude plate 431, a 3# bearing conical shell 432, a 3# squirrel cage 433, a 3# bearing seat 434, a 4# amplitude plate 441, a 4# bearing conical shell 442, a 4# squirrel cage 443 and a 4# bearing seat 444 and extracts the vibration acceleration at the positions of a low-pressure casing 6, a middle casing 7 and a high-pressure casing 8. The acceleration signal is measured and converted into a vibration speed signal, and the vibration displacement signal is converted into a vibration speed signal. And training the obtained 100 groups of test data on a neural network model by using a deep learning algorithm, and analyzing the vibration transfer coupling rule of the double rotors, the supports and the case.
The test method of the invention comprises the following steps:
(1) adopting Latin hypercube group sampling to determine the amount and phase of unbalance to be added;
(2) setting the rotating speed of a low-voltage driving motor, enabling a low-voltage rotor to rotate and a high-voltage rotor not to rotate, and extracting vibration data of each test point;
(3) setting the rotating speed of a high-voltage driving motor to enable a high-voltage rotor to rotate and a low-voltage rotor not to rotate, and extracting vibration data of each test point;
(4) setting the rotating speed of a low-voltage driving motor and the rotating speed of a high-voltage driving motor at the same time, enabling the low-voltage rotor and the high-voltage rotor to rotate, and extracting vibration data of each test point;
(5) repeating the steps (1) to (4) until 100 groups of sampling unbalanced working conditions are completed;
(6) and (3) carrying out self-adaptive filtering on the measuring point signals, removing interference components, converting the acceleration signals into vibration speed through numerical integration, converting the displacement signals into vibration displacement through numerical differentiation, and unifying vibration data dimensions.
(7) And training the obtained test data on a neural network model by using a deep learning algorithm, and analyzing the vibration transfer coupling rule of the double rotors, the support and the casing.
The test method provided by the invention can measure vibration data of a plurality of positions relative to a real machine.
The present invention has been further described with reference to specific embodiments, but it should be understood that the detailed description should not be construed as limiting the spirit and scope of the present invention, and various modifications made to the above-described embodiments by those of ordinary skill in the art after reading this specification are within the scope of the present invention.
Claims (10)
1. An aircraft engine birotor-bearing-casing tester is characterized by comprising a low-pressure rotor, a central bevel gear system and a high-pressure rotor which are transversely arranged in sequence, wherein:
the low-pressure rotor, the central bevel gear system and the high-pressure rotor are fixed on the frame through a supporting structure, a low-pressure casing is arranged on the periphery of the low-pressure rotor, an intermediate casing is arranged on the periphery of the central bevel gear system, a high-pressure casing is arranged on the periphery of the high-pressure rotor, and the low-pressure casing, the intermediate casing, the high-pressure casing, the low-pressure rotor and the high-pressure rotor are coaxially arranged and supported by the supporting structure;
the low pressure rotor includes a low pressure turbine shaft extending through the central bevel gear system and being in rotational communication with the high pressure rotor and the central bevel gear system;
the low pressure rotor is driven by a low pressure drive motor, the central bevel gear system is connected with a high pressure drive motor, and the high pressure rotor is driven by the central bevel gear system.
2. The aircraft engine dual-rotor-support-casing tester as claimed in claim 1, wherein the low-pressure rotor comprises a low-pressure fan section rotor and a low-pressure turbine rotor, the low-pressure fan section rotor comprises a fan rotating shaft and a four-stage fan disk sleeved on the fan rotating shaft, the low-pressure turbine rotor comprises the low-pressure turbine shaft, and a low-pressure turbine cone shell and a low-pressure turbine disk sleeved on the low-pressure turbine shaft, the low-pressure turbine cone shell is fixedly installed at the outer end of the low-pressure turbine disk, the fan rotating shaft is connected with the low-pressure turbine shaft through a coupling, and the low-pressure driving motor is connected with the fan rotating shaft.
3. The aircraft engine dual rotor-bearing-casing tester as claimed in claim 2, wherein the high pressure rotor comprises a high pressure forearm shaft, a high pressure compressor 9-stage wheel disc, a high pressure turbine shaft, a high pressure turbine disc, a high pressure turbine cone shell and an intermediate bearing, wherein:
the high-pressure forearm shaft is a hollow shaft, the front end of the high-pressure forearm shaft is connected with the central bevel gear system, and the rear end of the high-pressure forearm shaft is connected with the 9-stage wheel disc of the high-pressure compressor; the high-pressure turbine disc is rigidly connected with part of the 9-stage wheel disc of the high-pressure compressor through the high-pressure turbine shaft, and the high-pressure turbine conical shell is fixedly connected to the outer end of the high-pressure turbine disc;
the high-pressure turbine cone shell is connected with the low-pressure turbine shaft through the intermediate bearing.
4. The aircraft engine dual rotor-bearing-case tester as claimed in claim 3, wherein the central bevel gear system comprises a speed increaser, a long coupling, a central bevel gear box, a drive bevel gear and a driven bevel gear, wherein:
the speed increaser is connected with the high-pressure driving motor, the speed increaser drives the driving bevel gear through the long coupler, the driving bevel gear is meshed with the driven bevel gear, the driven bevel gear is connected with the high-pressure front arm shaft in a matched mode to drive the high-pressure rotor to rotate.
5. The aircraft engine dual rotor-bearing-casing tester as claimed in claim 4, wherein the support structure comprises a first support structure, a second support structure, a third support structure and a fourth support structure, the first and fourth support structures being configured to support the low pressure rotor, the second support structure being configured to support the low pressure rotor and the central bevel gear box, the third support structure being configured to support the high pressure rotor, the third support structure being a resilient support structure.
6. The aircraft engine dual-rotor-bearing-casing tester as claimed in claim 5, wherein the first supporting structure comprises a No. 1 amplitude plate, a No. 1 bearing conical shell, a No. 1 bearing seat, a No. 1 squirrel cage and a No. 1 bearing, and one end of the fan rotating shaft is rotatably connected with the first supporting structure through the No. 1 bearing;
the second supporting structure comprises a No. 2 amplitude plate, a No. 2 bearing seat and a No. 2 bearing, and the other end of the fan rotating shaft is rotatably connected with the second supporting structure through the No. 2 bearing;
the third support structure comprises a No. 3 amplitude plate, a No. 3 bearing conical shell, a No. 3 squirrel cage, a No. 3 bearing seat and a No. 3 bearing, and the middle part of the high-pressure front shaft arm is rotationally connected with the third support structure through the No. 3 bearing;
the fourth supporting structure comprises a No. 4 amplitude plate, a No. 4 bearing conical shell, a No. 4 squirrel cage, a No. 4 bearing seat and a No. 4 bearing, and the low-pressure turbine shaft is rotatably connected with the fourth supporting structure through the No. 4 bearing.
7. The aircraft engine dual rotor-bearing-case tester as claimed in claim 6, wherein said low pressure case is secured to said web No. 1 and said web No. 2, said intermediate case is secured to said web No. 2 and said web No. 3, and said high pressure case is secured to said web No. 3 and said web No. 4.
8. The aircraft engine dual rotor-bearing-casing tester as claimed in claim 4, wherein the frame includes a rotor system frame and a high-pressure drive train frame, the low-voltage drive motor being fixed to the rotor system frame, the high-pressure drive motor and the speed increaser being fixed to the high-pressure drive train frame.
9. The aircraft engine dual rotor-bearing-casing tester as claimed in claim 3, wherein the four-stage fan disk, the high-pressure compressor 9-stage wheel disk, the low-pressure turbine disk and the high-pressure turbine disk are provided with a plurality of bolt holes along the periphery of the wheel disk.
10. A test method based on the aircraft engine double rotor-support-casing tester of any one of claims 1 to 9, characterized by comprising the following steps:
(1) adopting Latin hypercube group sampling to determine the amount and phase of unbalance to be added;
(2) setting the rotating speed of a low-voltage driving motor, enabling a low-voltage rotor to rotate and a high-voltage rotor not to rotate, and extracting vibration data of each test point;
(3) setting the rotating speed of a high-voltage driving motor to enable a high-voltage rotor to rotate and a low-voltage rotor not to rotate, and extracting vibration data of each test point;
(4) setting the rotating speed of a low-voltage driving motor and the rotating speed of a high-voltage driving motor at the same time, enabling the low-voltage rotor and the high-voltage rotor to rotate, and extracting vibration data of each test point;
(5) repeating the steps (1) to (4) until 100 groups of sampling unbalanced working conditions are completed;
(6) and (3) carrying out self-adaptive filtering on the measuring point signals, removing interference components, converting the acceleration signals into vibration speed through numerical integration, converting the displacement signals into vibration displacement through numerical differentiation, and unifying vibration data dimensions.
(7) And training the obtained test data on a neural network model by using a deep learning algorithm, and analyzing the vibration transfer coupling rule of the double rotors, the support and the casing.
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