CN112254966B - Assembly quantity measuring method for aircraft engine rotor simulation assembly process - Google Patents
Assembly quantity measuring method for aircraft engine rotor simulation assembly process Download PDFInfo
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- CN112254966B CN112254966B CN202010766675.2A CN202010766675A CN112254966B CN 112254966 B CN112254966 B CN 112254966B CN 202010766675 A CN202010766675 A CN 202010766675A CN 112254966 B CN112254966 B CN 112254966B
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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
- G01M15/00—Testing of engines
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01B—MEASURING LENGTH, THICKNESS OR SIMILAR LINEAR DIMENSIONS; MEASURING ANGLES; MEASURING AREAS; MEASURING IRREGULARITIES OF SURFACES OR CONTOURS
- G01B11/00—Measuring arrangements characterised by the use of optical techniques
- G01B11/02—Measuring arrangements characterised by the use of optical techniques for measuring length, width or thickness
- G01B11/06—Measuring arrangements characterised by the use of optical techniques for measuring length, width or thickness for measuring thickness ; e.g. of sheet material
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01B—MEASURING LENGTH, THICKNESS OR SIMILAR LINEAR DIMENSIONS; MEASURING ANGLES; MEASURING AREAS; MEASURING IRREGULARITIES OF SURFACES OR CONTOURS
- G01B17/00—Measuring arrangements characterised by the use of infrasonic, sonic or ultrasonic vibrations
- G01B17/02—Measuring arrangements characterised by the use of infrasonic, sonic or ultrasonic vibrations for measuring thickness
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- G—PHYSICS
- G01—MEASURING; TESTING
- G01L—MEASURING FORCE, STRESS, TORQUE, WORK, MECHANICAL POWER, MECHANICAL EFFICIENCY, OR FLUID PRESSURE
- G01L5/00—Apparatus for, or methods of, measuring force, work, mechanical power, or torque, specially adapted for specific purposes
- G01L5/24—Apparatus for, or methods of, measuring force, work, mechanical power, or torque, specially adapted for specific purposes for determining value of torque or twisting moment for tightening a nut or other member which is similarly stressed
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Abstract
An assembly quantity measuring method for simulating an assembly process of an aircraft engine rotor adopts a temperature difference method to realize that a rear shaft of a high-pressure compressor, a rear sealing disc of the compressor and a front shaft of a high-pressure turbine form a test assembly, a combination part of the test assembly is used as a measuring object in the measuring method, and a test point is positioned beside a bolt hole of the combination part; and representing the assembling quantity change of the assembling process by the thickness change of the joint part. The invention has the advantages that: the assembly process digitization can be realized, the empirical data is made into a database, and an assembly reference basis is provided for actual assembly.
Description
Technical Field
The invention relates to the field of aircraft engine assembly, in particular to an aircraft engine rotor assembly simulation test system.
Background
An aero-engine (aero-engine) is a highly complex and precise thermal machine, is used as the heart of an airplane, is not only the power for flying the airplane, but also an important driving force for promoting the development of aviation industry, and each important change in human aviation history is inseparable from the technical progress of the aero-engine.
Often, engine failures are most directly manifested by inadequate material properties, such as, for example, blade rupture events within the engine, and more severe turbine rupture, containment failure events within the engine. Why does a part crack or break entirely? The most direct shallowest surface is that the material of the part is not able to withstand the externally applied loads to the part, resulting in cracking or breaking.
However, aircraft engine operating conditions are very complex, and the engine is momentarily in a vibrating environment. The forces to which the parts are subjected are dynamic, and can be neglected, and the material is much easier to break under such vibrational stresses. Therefore, if high vibration stress occurs in the engine, the material is easily cracked and broken due to the failure to receive large vibration stress. The problem of vibration is not solved, and only the development of materials is not useful. And vibration of the engine rotor is caused in part by the assembly process.
The rotor of the aircraft engine is formed by assembling multiple stages of units along the axial direction, for example, a front shaft part of a high-pressure turbine, a rear sealing disc of a gas compressor and a rear shaft of the high-pressure gas compressor are fastened and connected by bolts. The torque wrench needs to penetrate through the space between the disks of the high-pressure turbine and extend into the joint of the front shaft part of the high-pressure turbine, the rear sealing disk of the gas compressor and the rear shaft of the high-pressure gas compressor, and the bolt is screwed by the long-transmission torque wrench under the condition of a deep-hole blind cavity. That is to say, when the aeroengine rotor is assembled through the bolt, the assembly process is invisible, the assembly result is not measurable, and after the assembly is accomplished, the bolt pretightning force and the degree of connection of being connected portion are unknown, and the great problem of connection divergence is comparatively outstanding.
Moreover, because the assembly process is invisible and the assembly result is not measurable, the assembly process cannot be quantitatively evaluated, the installation experience cannot be accumulated, and guidance cannot be provided for subsequent installation.
Disclosure of Invention
The invention aims to provide an aircraft engine rotor assembly simulation test system which enables the assembly process of an aircraft engine rotor to be visible and the assembly result to be measurable and enables process data of an assembly process to be quantized so as to provide experience guidance for subsequent assembly.
A simulation test system for the assembly of an aircraft engine rotor comprises a simulation part mechanism for simulating the assembly working condition of the rotor, a test assembly and a measuring mechanism for measuring the assembly process;
the simulation part mechanism comprises a high-pressure turbine simulation part and a tool, and a test component accommodating area is arranged between the high-pressure turbine simulation part and the tool; the size of an inlet of the high-pressure turbine simulation piece is equal to the rear shaft diameter of the high-pressure turbine wheel disc, and the depth from the inlet of the high-pressure turbine simulation piece to an aviation bolt assembling part is equal to the depth of the high-pressure turbine;
the test assembly comprises a high-pressure turbine front shaft part connected with the high-pressure turbine simulation part, a gas compressor rear sealing disc and a high-pressure gas compressor rear shaft part, and the aviation bolt is used for connecting the high-pressure turbine front shaft part, the gas compressor rear sealing disc and the high-pressure gas compressor rear shaft part; the test assembly is detachably arranged in the test assembly accommodating area; the measuring mechanism measures each measuring point of the test assembly.
According to the invention, the inlet aperture of a real high-pressure turbine and the depth from the inlet of the high-pressure turbine to an aviation bolt assembly part are simulated through the high-pressure turbine simulation part, so that the assembly simulation test system has the working condition of a real deep-hole blind cavity of an aircraft engine rotor during assembly. The high-pressure turbine simulation part only needs to be matched with a real assembly environment in terms of the aperture and the depth of an inlet, and the appearance of the high-pressure turbine simulation part is not limited. The high-pressure turbine simulator can be made of transparent materials, and the internal operation conditions can be observed by adopting the transparent materials.
Furthermore, a window is formed in the side of the high-pressure turbine simulation part, and operation is convenient.
The test component can be a 1:1 composite original of a connecting part of an aircraft engine rotor, or a connecting part cut out of a front shaft of a high-pressure turbine and a connecting part cut out of a rear shaft of a high-pressure compressor. In addition, in the assembly simulation test system, the test assembly is detachably matched with the high-pressure turbine simulation part and the tool, so that the system can be repeatedly tested for many times. The high-pressure turbine simulation part and the tool with different sizes are manufactured according to different sizes of the aeroengine rotor, and the assembly simulation test device can be suitable for assembly simulation tests of different models of aeroengine rotors.
In the rotor of the aero-engine, the assembly of the front shaft part of the high-pressure turbine, the rear sealing disc of the air compressor and the rear shaft part of the high-pressure air compressor is realized through a plurality of aero bolts, the aero bolts are distributed on a graduated circle at equal intervals, and the fastening process of each bolt is measured in the simulated assembly process. The parts of the front shaft part of the high-pressure turbine, the rear sealing disc of the compressor and the rear shaft part of the high-pressure compressor which are connected by the bolts are called as a test assembly joint part. And measuring the variation of the joint part when the fit degree of the joint part is changed every time one bolt is screwed.
During simulated assembly, the measuring mechanism is inserted into the test assembly, and the assembly of the aircraft engine rotor is changed from invisible and undetectable to visible in the assembly process.
In a first aspect, a specific measurement mechanism configuration is provided. Preferably, the measuring mechanism comprises a measuring bracket and a measuring assembly, wherein the measuring bracket comprises a base, an upright post and a sleeve; the sleeve is rotatably connected with the upright post; the measurement assembly includes a pair of distance probes.
The specific structure of the measuring support and the indexing process steps is as follows: preferably, there is a step-by-step indexing-rotation assembly between the upright and the sleeve, the step of the indexing-rotation assembly being equal to 360 °/number of aviation bolts. The step type is that when no external force is applied, the indexing rotary component can be locked at the current position, and can rotate to the next position only when the external force is input, and the included angle between the previous position and the next position is one step. For example, a stepper motor may be used as an indexing element.
Preferably, the sleeve is provided with a handle, the handle is perpendicular to the sleeve, and the base, the upright post and the sleeve are coaxial.
The specific structural scheme of the measuring assembly is as follows: preferably, the measuring assembly comprises a camera and/or a light source, and the camera is aligned with the test assembly combining part; or the measuring assembly comprises two cameras, one camera is aligned with the joint surface of the front shaft part of the high-pressure turbine and the rear sealing disc of the compressor, and the other camera is aligned with the joint surface of the rear sealing disc of the compressor and the rear shaft part of the high-pressure compressor.
When the camera and the light source are used as measuring components, the camera is aligned to the joint of the test component to take a picture, then the area of the joint of the test component is extracted from the image, or the areas of two joint surfaces are extracted, and then the thickness of the joint of the test component and the thickness of the two joint surfaces are identified.
When the distance measuring head is used as the measuring assembly, the measuring assembly comprises the distance measuring head and the lower distance measuring head, the distance between the two distance measuring heads is known, during measurement, the distance between the upper distance measuring head and the upper surface of the combining part can be measured and obtained, and the distance between the lower distance measuring head and the lower surface of the combining part can be measured and obtained, so that the distance between the two distance measuring heads, namely the distance between the upper distance measuring head and the upper surface of the combining part, and the distance between the lower distance measuring head and the lower surface of the combining part, are equal to the thickness of the region to be measured. The distance measuring head adopts laser ranging or ultrasonic ranging.
Preferably, when the measuring assembly selects a pair of distance measuring heads, an upper cross beam and a lower cross beam are arranged on a sleeve of the measuring support, one end of the upright post is fixed with the base, the other end of the upright post is rotatably connected with the sleeve, one end of each of the upper cross beam and the lower cross beam is connected with the sleeve, and the other end of each of the upper cross beam and the lower cross beam is a free end; the free end of the upper cross beam is provided with an upper distance measuring head, and the free end of the lower cross beam is provided with a lower distance measuring head. A joint between the upper distance measuring head and the lower distance measuring head can accommodate the test assembly.
Preferably, the upper cross beam is rotatably connected with the sleeve, and the centers of the front shaft part of the high-pressure turbine, the rear sealing disc of the compressor and the rear shaft part of the high-pressure compressor are provided with communicated shaft holes. And the upper beam is turned right angle upwards and penetrates through the upright column from top to bottom to be sequentially placed into the rear shaft part of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft part of the high-pressure turbine.
Preferably, the tool comprises a hollow cylinder, the stand column is arranged in the cylinder, the upper cross beam penetrates through the shaft hole and is located above the aviation bolt, and the lower cross beam is located below the aviation bolt.
Preferably, the cylindrical surface of the cylinder is provided with an opening. The measuring mechanism is arranged in the cylinder body, and the indexing rotary component of the measuring mechanism is convenient to operate through the opening. Preferably, the openings are symmetrically arranged two by two along the axis of the cylinder. The number of openings is not limited to two, and may be asymmetrically provided.
Preferably, the cylinder body is fixed with the rear shaft part of the high-pressure compressor through a flange, and a reinforcing block is arranged between the flange and the side wall of the cylinder body; the cylinder is concentric with the rear shaft part of the high-pressure compressor. And a precision bolt is adopted to realize the accurate installation of the test assembly and the fixture which are coaxial in centering. The top flange is used for supporting the high-pressure turbine simulation part and the test assembly, and the supporting strength of the flange is improved by arranging the reinforcing blocks.
Preferably, the tool comprises a working platform, the base comprises an upper base and a lower base, the upper base is fixed with the stand column, the lower base is fixed with the working platform, waist circular holes are formed in the circumferential directions of the upper base and the lower base, the upper waist circular hole and the lower waist circular hole are arranged in a transverse-vertical cross mode, and the barrel is fixed with the working platform. The upper base and the lower base of the measuring mechanism are positioned and finely adjusted through the oval hole, so that the measuring mechanism and the tool (cylinder) are coaxially positioned.
A front high-pressure turbine shaft of the aircraft engine is matched with a rear sealing disc of the gas compressor as a hole shaft, the rear sealing disc of the gas compressor is used as a hole, and the front high-pressure turbine shaft is used as a shaft; the rear shaft of the high-pressure compressor is matched with the rear sealing disc of the compressor as a hole shaft, the rear sealing disc of the compressor is used as a hole, and the rear shaft of the high-pressure compressor is used as a shaft. However, the two holes are matched with each other, and the hole diameter is smaller than the shaft diameter at normal temperature.
The elongation of the bolt can be measured, and the joint degree of the front shaft part of the high-pressure turbine and the joint surface of the rear sealing disc of the compressor can be seen and measured; the joint degree of the rear sealing disc of the gas compressor and the joint surface of the rear shaft part of the high-pressure gas compressor is visible and measurable.
In a second aspect of the invention, a method for measuring the thickness variation of the test assembly joint as the assembling process of the aircraft engine rotor is provided, wherein the test assembly joint is used as a measuring object.
An assembly quantity measuring method for an aircraft engine rotor simulation assembly process comprises the following steps:
1. the shaft is installed into the hole by a temperature difference method, a rear shaft of the high-pressure compressor, a rear sealing disc of the compressor and a front shaft of the high-pressure turbine form a test assembly, and then the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine are tightly pressed by process bolts, wherein the number of the process bolts is multiple, and the multiple process bolts are uniformly distributed along the hole; after the test assembly recovers the room temperature, measuring the thickness of the joint part of the test assembly, wherein each bolt hole corresponds to one measuring point which is positioned beside the bolt hole; measuring the thickness of each bonding part at each measuring point, wherein the thickness of each bonding part obtained at each measuring point is measured once, and the thickness of the bonding part obtained by the measurement is taken as the thickness of the first group of bonding parts;
2. disassembling the process bolts, sequentially measuring the thickness of the joint of each measuring point after all the process bolts are disassembled, and taking the thickness of the joint obtained by the measurement as the thickness of the second group of joints;
3. sequentially applying bolt pre-tightening to the test assembly according to a tightening process scheme, and measuring the thickness of the sequential joint part at a measuring point corresponding to each pre-tightening bolt when each pre-tightening bolt is tightened; sequentially assembling all the pre-tightening bolts, and obtaining the thickness of a third group of combined parts; the tightening process was pre-designed before the start of the test.
4. After all the pre-tightening bolts are tightened, the thicknesses of the joint parts of all the measuring points are sequentially obtained, and the thickness of the fourth group of joint parts is obtained;
and (4) each joint thickness obtained in the step 1-4 comprises a respective measuring point identifier and a thickness value of the measuring point. For example, the first hole location corresponds to a first measurement point C1 having a joint thickness of Xmm.
Preferably, in the step 1, the assembly by the temperature difference method is to heat the rear sealing disc of the compressor, and after the hole of the rear sealing disc of the compressor expands, the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are/is installed in the rear sealing disc of the compressor.
Or in the step 1, the assembly by the temperature difference method is to cool the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine, and then the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are installed in the rear sealing disc of the compressor.
The working process of the invention is as follows: firstly, the cylinder and the base are respectively fixed on a working platform, and the upper base and the lower base of the measuring mechanism are positioned and finely adjusted through the oval hole, so that the measuring mechanism and the tool (cylinder) are coaxial. Then, the upper cross beam is turned upwards, the rear shaft part of the high-pressure compressor penetrates through the stand column and is positioned in a matched mode through a precise bolt, the rear shaft part of the high-pressure compressor is coaxial with the cylinder, then a rear sealing disc of the compressor and the front shaft part of the high-pressure turbine are stacked on the rear shaft part of the high-pressure compressor respectively, bolt holes are aligned, an aviation bolt is placed in the bolt holes correctly, it is guaranteed that an elastic sheet of the bolt head is clamped in place, the elastic sheet is welded or bonded to the end of the bolt head in a clamped mode, and the bolt is prevented from falling off under the condition that a nut is not screwed through clamping with the rear shaft part of the high-pressure compressor. And after the installation is finished, measuring the joint surface of each measuring point in sequence.
The first group of measurement reflects the deformation condition of the joint part under the constraint conditions of seam allowance interference and process bolt preloading after the completion of the assembling procedure of the seam allowance cylindrical surface temperature difference method; the second group of measurement reflects the deformation condition of the joint part under the action of residual stress of the assembly process by the temperature difference method after the preloading constraint of the process bolt is released; the third group of measurement is that the deformation condition of a joint part near each loading point position in the screwing process of the flange bolt group is tracked by taking the assembled spigot of the temperature difference method as an initial boundary condition; the fourth group of measurement is to finally detect the deformation condition of the joint part of each monitoring point after all the assembly processes are finished; through four sets of measurements, the process monitoring of the whole process flow of rotor assembly can be comprehensively realized, and therefore basic test data support is provided for the optimization design of the multi-process assembly process of the rotor. Such as: temperature control and optimization of mechanical preload may be provided for a temperature-differential assembly based on the evaluation of the first set of measurements and the second set of measurements; and the evaluation of the third set of measurement and the fourth set of measurement can provide a basis for the optimization (elastic interaction under the constrained condition) of the bolt set screwing process. In a word, through measurement of a whole process of a plurality of processes of rotor assembly, the purpose is to disclose an internal action relation among the plurality of processes, and to develop related process optimization design attempts for achieving a control target of circumferential rigidity uniformity of rotor connection.
In a third aspect of the invention, a bolt assembled at the joint of the test assembly is taken as a measuring object, and the elongation of the bolt is taken as the assembling process of the aeroengine rotor.
An assembly quantity measuring method for an aircraft engine rotor simulation assembly process comprises the following steps:
1. the shaft is installed into the hole by a temperature difference method, a rear shaft of the high-pressure compressor, a rear sealing disc of the compressor and a front shaft of the high-pressure turbine form a test assembly, then the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine are tightly pressed by process bolts, the process bolts are multiple and are uniformly distributed along the hole, and after the test assembly is recovered to the room temperature; disassembling the process bolt;
2. sequentially applying bolt pre-tightening to the test assembly according to a tightening process scheme, and measuring the original length of each bolt before pre-tightening; after pre-tightening, measuring the pre-tightened length of the bolt once; the elongation of each bolt was calculated. In this scheme, the extension of using the bolt represents the bolt pretightning force. After the elongation of all bolts is measured, the distribution condition of the bolt pretightening force of the joint of the test assembly can be reflected.
Preferably, in the scheme of measuring the elongation of the bolt, the joint degree of the front shaft part of the high-pressure turbine and the joint surface of the rear sealing disc of the compressor and the joint surface of the rear shaft part of the high-pressure compressor are also measured.
The elongation of the bolt represents the actual pretightening force applied to the bolt, the bolt is taken as a measuring object,
accumulating the empirical data of the elongation of the bolt and the joint level of the rear sealing disc of the compressor and the rear shaft part of the high-pressure compressor as well as the joint level of the front shaft part of the high-pressure turbine and the rear sealing disc of the compressor; converting the elongation of the bolt into bolt pretightening force, and establishing corresponding relation empirical data of torque wrench output and actual bolt pretightening force; and the empirical data is made into a database to provide an assembly reference basis for actual assembly, and the aircraft engine rotor assembly process is expected to be evaluated and screened according to the empirical database.
The invention has the beneficial effects that:
1. the method uses the high-pressure turbine simulation part to simulate the aperture of the inlet of the high-pressure turbine and the depth from the inlet of the high-pressure turbine to the aviation bolt assembly part, so that the assembly simulation system has the working condition of a deep hole blind cavity of an actual aircraft engine rotor during assembly, a test component is a local part of the aircraft engine rotor, and the actual assembly of a front shaft of the high-pressure turbine, a rear sealing disc of a gas compressor and a rear shaft of the high-pressure gas compressor can be truly reproduced by using the simulation test system; the thickness of the joint part is given by the measuring mechanism, the deformation of the circumferential point position of the structural part can be measured, and further the circumferential rigidity distribution or the tight-fit distribution of the rotor connection is reflected, so that the whole assembly process is visible, and the parameters of the assembly process can be measured.
2. In the assembly simulation test system, the tool plays a role in supporting and fixing the test assembly, and the test assembly is detachably matched with the high-pressure turbine simulation part and the tool, so that the system can be repeatedly tested for multiple times; the high-pressure turbine simulation part and the tool with different sizes are manufactured according to different sizes of the aeroengine rotor, and the assembly simulation test device can be suitable for assembly simulation tests of different models of aeroengine rotors.
3. A method for representing the assembling process of a rotor by using the thickness change of a combined part with the combined part as a measuring object is provided. The measurement method is combined with a temperature difference assembly method and an actual assembly process of the rotor, assembly process digitization is achieved, empirical data are made into a database, assembly reference bases are provided for actual assembly, and assessment and screening of an aircraft engine rotor assembly process are hopefully performed according to the empirical database.
4. A method for representing a rotor assembling process by using bolt elongation by taking the bolt as a measuring object is provided. And the stress change of the joint part is reflected by the elongation of the bolt in the rotor assembling process. In the assembly process, the stress change process of the joint part is digitalized, the connection between the screwing process and the rotor assembly process is established, and a basis is provided for the evaluation and screening of the aircraft engine rotor assembly process.
Drawings
FIG. 1 is a schematic view of the structure of the present invention.
Fig. 2 is a schematic structural view of the measuring mechanism.
Fig. 3 is a schematic structural diagram of an upper beam and a lower beam of the measuring mechanism.
FIG. 4 is a structural diagram of a measuring mechanism for measuring an aircraft bolt.
FIG. 5 is a top view of a simulation testing system.
Fig. 6 is a structural view in the direction of a-a in fig. 5.
Fig. 7 is a structural view of the measuring mechanism fixed to the work platform.
The labels in the figure are: a high-pressure turbine simulation member 1; a test assembly 2; a high-pressure turbine front shaft part 21, a compressor rear sealing disc 22, a high-pressure compressor rear shaft part 23 and a shaft hole 24; the measuring mechanism 3, the distance measuring head 31, the first distance measuring head 311, the second distance measuring head 312, the first camera 32, the second camera 33, the base 34, the upper base 341, the lower base 342, the oval hole 343, the upright column 35, the sleeve 36, the upper cross beam 37, the lower cross beam 38, the indexing rotary component 39 and the handle 30; an aircraft bolt 4; the tool 5, the opening 51, the flange 52, the reinforcing block 54 and the working platform 55.
Detailed Description
The present invention will be further described with reference to the structures or terms used herein. The description is given for the sake of example only, to illustrate how the invention may be implemented, and does not constitute any limitation on the invention.
The invention is further described with reference to the following figures and detailed description. In the description of the present invention, it is to be understood that the terms "upper", "lower", "front", "rear", "left" and "right", etc., indicate orientations or positional relationships based on those shown in the drawings, and are only for convenience of description and simplification of description, but do not indicate or imply that the positions or elements referred to must have specific orientations, be constructed in specific orientations, and be operated, and thus are not to be construed as limitations of the present invention. Furthermore, the terms "first," "second," and the like are used for descriptive purposes only and are not to be construed as indicating or implying relative importance.
Example 1
As shown in FIG. 1, the aircraft engine rotor assembly simulation test system comprises a simulation part mechanism for simulating the assembly working condition of a rotor, a test component 2 mechanism and a measuring mechanism 3 for measuring the assembly process;
the simulation mechanism comprises a simulation mechanism body and a simulation device, wherein the simulation mechanism body comprises a high-pressure turbine simulation part 1 and a tool, and a test assembly 2 accommodating area is arranged between the high-pressure turbine simulation part 1 and the tool; the size of the inlet of the high-pressure turbine simulation part 1 is equal to the rear shaft diameter of the high-pressure turbine wheel disc, and the depth from the inlet of the high-pressure turbine simulation part 1 to the aviation bolt assembling part is equal to the depth of the high-pressure turbine.
According to the invention, the inlet aperture of the high-pressure turbine and the depth from the inlet of the high-pressure turbine to the aviation bolt assembly part are simulated through the high-pressure turbine simulation part, so that the assembly simulation test system has a real deep-hole blind cavity working condition when the aircraft engine rotor is assembled. The high-pressure turbine simulation part 1 only needs to be matched with a real assembly environment in terms of the aperture and the depth of an inlet, and the appearance of the high-pressure turbine simulation part 1 is not limited. The high-pressure turbine simulation piece 1 can be formed by 3D printing of a transparent resin material, and the internal operation condition can be observed by the transparent material. The side part of the high-pressure turbine simulation part 1 is provided with a window, so that the operation is convenient.
The test component 2 is a 1:1 composite original of a connecting part of an aircraft engine rotor, or a connecting part cut out of a front shaft of a high-pressure turbine and a connecting part cut out of a rear shaft of a high-pressure compressor. In addition, in the assembly simulation test system, the test component 2 is detachably matched with the high-pressure turbine simulation part 1 and the tool, so that the system can be repeatedly tested for many times. The high-pressure turbine simulation part 1 and the tool are manufactured according to different sizes of the aeroengine rotors, and the assembly simulation test method can be suitable for assembly simulation tests of different models of aeroengine rotors.
Example 2
With reference to FIGS. 2 to 6
This embodiment gives a detailed structure of the measuring mechanism 3 on the basis of embodiment 1. The simulation test system includes a measuring mechanism 3, and as shown in fig. 3, the measuring mechanism 3 includes a pair of distance measuring heads 31.
In the rotor of the aircraft engine, the assembly of a front shaft part 21 of a high-pressure turbine, a rear sealing disc 22 of a compressor and a rear shaft part 23 of the high-pressure compressor is realized through a plurality of aviation bolts 4, the aviation bolts 4 are distributed on a graduated circle at equal intervals, in the simulation assembly process, the fastening process of each bolt needs to be measured, and the elongation of each bolt is measured.
During simulated assembly, the measuring mechanism 3 can be inserted into the test component 2, referring to fig. 4, the assembly of the aircraft engine rotor is changed from invisible and undetectable to visible in the assembly process, the bolt tightening process is visible, and the bolt elongation is measurable. The bolt tightening process is visual, and can be realized by adding a camera and a light source on the measuring assembly and collecting images and videos in the assembling process. Or the high-pressure turbine simulation part can be made of transparent materials, so that the bolt tightening process can be visible.
Converting the elongation of the bolt into bolt pretightening force, and establishing corresponding relation empirical data of torque wrench output and actual bolt pretightening force; and the empirical data is made into a database to provide an assembly reference basis for actual assembly, and the aircraft engine rotor assembly process is expected to be evaluated and screened according to the empirical database.
As shown in fig. 2-3, the measuring mechanism 3 includes a base 34, a column 35, a sleeve 36, an upper beam 37 and a lower beam 38, wherein one end of the column 35 is fixed to the base 34, the other end is rotatably connected to the sleeve 36, one end of each of the upper beam 37 and the lower beam 38 is connected to the sleeve 36, and the other end is a free end; the free end of the upper beam 37 is fitted with a first distance measuring head 311 and the free end of the lower beam 38 is fitted with a second distance measuring head 312. When the aviation bolt 4 is not assembled with the nut, measuring the initial length of the aviation bolt 4; and after the aviation bolt 4 is screwed down, measuring the pre-tightened length of the aviation bolt 4, and subtracting the initial length from the pre-tightened length to obtain the elongation of the aviation bolt 4. It should be noted that the connection manner between the upper cross beam 37 and the sleeve 36 and the connection manner between the lower cross beam 38 and the sleeve 36 are not limited, and the upper cross beam 37 and the sleeve 36 are of an integral structure, and are welded or screwed.
The column 35 and the sleeve 36 have a step-wise indexing-rotation assembly 39, the step of the indexing-rotation assembly 39 being equal to 360 °/number of aviation bolts. By step-wise, it is meant that the indexing rotary member 39 can be locked in the current position in the absence of an external force and can be rotated to the next position only when an external force is applied, the angle between the previous position and the next position being one step.
The indexing rotation member 39 is an indexing pin hole fit, or a stepper motor. Therefore, the simulation test system can conveniently and accurately carry out video acquisition and bolt length measurement on the simulation assembly condition, and only needs to adjust and center for the first time and then only needs to rotate in a stepping type indexing manner. The cooperation of index pin hole specifically means that the hole the same with aviation bolt quantity is set up to circumference equidistant on stand 53, and the sleeve sets up the bolt hole, and the bolt pulls out the round pin when aiming at an aviation bolt, needs to rotate to next aviation bolt.
As shown in FIG. 2, the sleeve 36 has a handle 30 thereon, the handle 30 is perpendicular to the sleeve 36, and the base 34, the post 35 and the sleeve 36 are coaxial.
Specifically, the first distance measuring head 311 and the second distance measuring head 312 adopt laser distance measuring sensors, and the first distance measuring head 311 and the second distance measuring head 312 are arranged in a centering manner, as shown in fig. 6, are respectively located at two ends of the aviation bolt 4, and are perpendicular to the end face of the aviation bolt. The length measured by the screw is equal to the vertical distance between the upper beam 37 and the lower beam 38 for two distance measuring heads, distance measuring head 1L 1 and distance measuring head 2L 2.
The measurement of the elongation of the bolt is not limited to a laser distance measuring sensor, for example, ultrasonic measurement, and the principle and method of ultrasonic measurement also belong to the prior art, that is, a piezoelectric ceramic patch is attached to the head of the bolt, and an ultrasonic probe is used to realize the interface action with the bolt under the action of ultrasonic excitation. During measurement, timing is started when the head of the bolt sends ultrasonic waves to the rod part, and the time difference is calculated when the ultrasonic waves reflected back from the rod end are received. The time difference of the ultrasonic wave represents the length of the bolt, an original time difference Ts is recorded in a free state of the bolt, the ultrasonic equipment continuously sends an ultrasonic wave signal and continuously records the time difference Tt during the bolt tightening process, and the bolt elongation is (Tt-Ts)/Ts, so that the bolt elongation is obtained.
As shown in fig. 2, the upper cross member 37 is rotatably connected to the sleeve 36, and the high-pressure turbine front shaft portion 21, the compressor rear seal disc 22 and the high-pressure compressor rear shaft portion 23 have a shaft hole 24 communicating with each other at the center thereof. The upper cross beam 37 is turned upwards by ninety degrees and penetrates through the upright post 35 from top to bottom to be sequentially placed into the high-pressure compressor rear shaft part 23, the compressor rear sealing disc 22 and the high-pressure turbine front shaft part 23.
Preferably, the first camera head 32 is rotatably connected to the upper beam 37, and the second camera head 33 is rotatably connected to the lower beam 38. In this way, not only can the monitored field of view be adjusted, but the monitored field of view is also larger.
As shown in fig. 1, the tooling 5 comprises a hollow cylinder, the upright post 35 is arranged in the cylinder 5, the upper cross beam 37 passes through the shaft hole 24 and is positioned above the aviation bolt 4, and the lower cross beam 38 is positioned below the aviation bolt 4.
As shown in fig. 1, the cylindrical surface of the cylinder 5 is provided with an opening 51. The indexing and rotating assembly 39 of the measuring mechanism 3 is conveniently operated by placing the measuring mechanism 3 in the barrel 5 through the opening 51. Preferably, the openings 51 are symmetrically arranged in two along the axis of the cylinder 5. The number of the openings 51 is not limited to two, and may be asymmetrically provided.
As shown in fig. 1, the cylinder 5 is fixed with the rear shaft part 23 of the high-pressure compressor through a flange 52, and a reinforcing block 54 is arranged between the flange 52 and the side wall of the cylinder; the barrel 5 is concentric with the rear shaft portion 23 of the high pressure compressor. And a precision bolt is adopted to realize the accurate installation of the test assembly 2 and the tool 5 which are coaxial in centering. The top flange 23 is used for supporting the high-pressure turbine simulator 1 and the test assembly 2, and the supporting strength of the flange 52 is increased by arranging the reinforcing block 54.
As shown in fig. 1 and 7, the tooling 5 includes a working platform 55, the base 34 includes an upper base 341 and a lower base 342, the upper base 341 is fixed to the column 35, the lower base 342 is fixed to the working platform 55, waist-shaped holes 343 are circumferentially opened on the upper base 341 and the lower base 342, the upper waist-shaped hole 343 and the lower waist-shaped hole 343 are arranged in a transverse and vertical crossing manner, and the barrel 5 is fixed to the working platform 55. The upper base 341 and the lower base 342 of the measuring mechanism 34 are positioned and finely adjusted through a waist circular hole 343, so as to coaxially position the measuring mechanism 3 and the tool (cylinder).
The working process of the invention is as follows: firstly, the cylinder 5 and the base are respectively fixed on the working platform 55, and the upper base 341 and the lower base 342 of the measuring mechanism 3 are positioned and finely adjusted through the waist circular hole 343, so as to realize the coaxiality of the measuring mechanism 3 and the tool (cylinder). Then, the upper cross beam 37 is turned upwards, the rear shaft portion 23 of the high-pressure compressor penetrates through the upright column 35 and is positioned in a matched mode through a precise bolt, the rear shaft portion 23 of the high-pressure compressor is coaxial with the cylinder, then a rear sealing disc of the compressor and the front shaft portion 21 of the high-pressure turbine are stacked on the rear shaft portion 23 of the high-pressure compressor respectively, the bolt holes are aligned, the aviation bolt 4 is placed in the bolt holes correctly, the elastic piece of the bolt head is guaranteed to be clamped in place, the elastic piece is clamped and welded or bonded to the end of the bolt head, and the bolt is prevented from falling off under the condition that the nut is not screwed through clamping with the rear shaft portion 23 of the high-pressure compressor. After the installation is finished, each aviation bolt 4 is measured in sequence.
Example 3
An assembly quantity measuring method for simulating an assembly process of an aircraft engine rotor is shown in fig. 4 and 5, and comprises the following steps:
1. the shaft is inserted into the hole by a temperature difference method, the rear shaft of the high-pressure compressor, the rear sealing disc 22 of the compressor and the front shaft of the high-pressure turbine form a test component 2, and then the rear shaft of the high-pressure compressor, the rear sealing disc 22 of the compressor and the front shaft of the high-pressure turbine are tightly pressed by process bolts, wherein the number of the process bolts is multiple, and the multiple process bolts are uniformly distributed along the hole; after the test assembly 2 returns to the room temperature, measuring the thickness of the joint part of the test assembly 2, wherein each bolt hole corresponds to a measuring point which is positioned beside the bolt hole; measuring the thickness of each bonding part at each measuring point, wherein the thickness of each bonding part obtained at each measuring point is measured once, and the thickness of the bonding part obtained by the measurement is taken as the thickness of the first group of bonding parts;
2. disassembling the process bolts, sequentially measuring the thickness of the joint of each measuring point after all the process bolts are disassembled, and taking the thickness of the joint obtained by the measurement as the thickness of the second group of joints;
3. sequentially applying bolt pre-tightening to the test assembly 2 according to a tightening process scheme, and measuring the thickness of the sequential joint part at a measuring point corresponding to each pre-tightening bolt when each pre-tightening bolt is tightened; sequentially assembling all the pre-tightening bolts, and obtaining the thickness of a third group of combined parts; the tightening process was pre-designed before the start of the test.
4. After all the pre-tightening bolts are tightened, the thicknesses of the joint parts of all the measuring points are sequentially obtained, and the thickness of the fourth group of joint parts is obtained;
and (4) each joint thickness obtained in the step 1-4 comprises a respective measuring point identifier and a thickness value of the measuring point. For example, the first hole location corresponds to a first measurement point C1 having a joint thickness of Xmm.
In the step 1, the assembly by the temperature difference method is to heat the rear sealing disc 22 of the compressor, and after the hole of the rear sealing disc 22 of the compressor expands, the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are/is installed in the rear sealing disc 22 of the compressor.
Or, in the step 1, the assembly by the temperature difference method is to cool the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine, and then the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are installed in the rear sealing disc 22 of the compressor.
The method of this example can be implemented using the test system of example 1 or example 2.
Example 4
An assembly quantity measuring method for simulating an assembly process of an aircraft engine rotor is shown in fig. 4 and 5, and comprises the following steps:
1. the shaft is inserted into the hole by a temperature difference method, the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine form a test component 2, then the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine are tightly pressed by process bolts, the process bolts are multiple and are uniformly distributed along the hole, and after the test component 2 is recovered to the room temperature; disassembling the process bolt;
2. sequentially applying bolt pre-tightening to the test assembly 2 according to a tightening process scheme, and measuring the original length of each bolt before pre-tightening; after pre-tightening, measuring the pre-tightened length of the bolt once; the elongation of each bolt was calculated. In this scheme, the extension of using the bolt represents the bolt pretightning force. After the elongation of all bolts is measured, the distribution condition of the bolt pretightening force of the joint part of the test assembly 2 can be reflected.
Preferably, in the scheme of measuring the elongation of the bolts, the joint degree of the joint surface of the front shaft part 21 of the high-pressure turbine and the rear sealing disc 22 of the compressor and the joint degree of the joint surface of the rear sealing disc 22 of the compressor and the rear shaft part 23 of the high-pressure compressor are also measured.
The method of this example can be implemented using the test system of example 1 or example 2.
The invention shown and described herein may be practiced in the absence of any element or elements, limitation or limitations, which is specifically disclosed herein. The terms and expressions which have been employed are used as terms of description and not of limitation, and there is no intention in the use of such terms and expressions of excluding any equivalents of the features shown and described or portions thereof, and it is recognized that various modifications are possible within the scope of the invention. It should therefore be understood that although the present invention has been specifically disclosed by various embodiments and optional features, modification and variation of the concepts herein described may be resorted to by those skilled in the art, and that such modifications and variations are considered to be within the scope of this invention as defined by the appended claims.
The contents of the articles, patents, patent applications, and all other documents and electronically available information described or cited herein are hereby incorporated by reference in their entirety to the same extent as if each individual publication was specifically and individually indicated to be incorporated by reference. Applicants reserve the right to incorporate into this application any and all materials and information from any such articles, patents, patent applications, or other documents.
Claims (3)
1. An assembly quantity measuring method for an aircraft engine rotor simulation assembly process is characterized by comprising the following steps: the method comprises the following steps of forming a test assembly by a rear shaft of the high-pressure compressor, a rear sealing disc of the compressor and a front shaft of the high-pressure turbine by adopting a temperature difference method, wherein a joint part of the test assembly is used as a measurement object in the measurement method, and a measurement point is positioned beside a bolt hole of the joint part; representing the assembly quantity change of the assembly process by the thickness change of the joint part; the measuring method comprises the following steps:
1) the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are/is arranged in a hole of the rear sealing disc by a temperature difference method, the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine form a test assembly, then the rear shaft of the high-pressure compressor, the rear sealing disc of the compressor and the front shaft of the high-pressure turbine are tightly pressed by process bolts, the number of the process bolts is multiple, and the multiple process bolts are uniformly distributed along the hole; after the test assembly recovers the room temperature, measuring the thickness of the joint part of the test assembly, wherein each bolt hole corresponds to one measuring point which is positioned beside the bolt hole; measuring the thickness of each bonding part at each measuring point, wherein the thickness of each bonding part obtained at each measuring point is measured once, and the thickness of the bonding part obtained by the measurement is taken as the thickness of the first group of bonding parts;
2) the process bolts are disassembled, after all the process bolts are disassembled, the thickness of the joint part of each measuring point is measured in sequence, and the thickness of the joint part obtained through the measurement is used as the thickness of the second group of joint parts;
3) sequentially applying bolt pre-tightening to the test assembly according to a tightening process scheme, and sequentially measuring the thickness of the joint at a measuring point corresponding to each pre-tightening bolt when each pre-tightening bolt is tightened; sequentially assembling all the pre-tightening bolts, and obtaining the thickness of a third group of combined parts;
4) after all the pre-tightening bolts are tightened, the thicknesses of the joint parts of all the measuring points are sequentially obtained, and the thickness of the fourth group of joint parts is obtained;
and (4) each joint thickness obtained in the step 1-4 comprises a respective measuring point identifier and a thickness value of the measuring point.
2. The assembly quantity measuring method for simulating an assembly process of an aircraft engine rotor as claimed in claim 1, wherein: in the step 1, the assembly by the temperature difference method is to heat the rear sealing disc of the gas compressor, and after the hole of the rear sealing disc of the gas compressor expands, the rear shaft of the high-pressure gas compressor and/or the front shaft of the high-pressure turbine are/is installed in the rear sealing disc of the gas compressor.
3. The assembly quantity measuring method for simulating an assembly process of an aircraft engine rotor as claimed in claim 1, wherein: in the step 1, the assembly by the temperature difference method is to cool the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine, and then the rear shaft of the high-pressure compressor and/or the front shaft of the high-pressure turbine are arranged in a rear sealing disc of the compressor.
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Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101799354A (en) * | 2010-02-09 | 2010-08-11 | 西安交通大学 | Method for testing assembly performance of rotor of aircraft engine |
| CN103292959A (en) * | 2012-03-05 | 2013-09-11 | 沈阳黎明航空发动机(集团)有限责任公司 | Method for analyzing and detecting parameter of dynamic balance simulation rotor of engine rotor component |
| CN103791816A (en) * | 2014-02-14 | 2014-05-14 | 哈尔滨工业大学 | Aircraft engine rotor assembly method and device based on concentricity and perpendicularity measurement |
| CN105241598A (en) * | 2015-11-04 | 2016-01-13 | 中国航空动力机械研究所 | Method and system for measuring pre-tightening force of engine rotor |
| CN106289120A (en) * | 2016-08-11 | 2017-01-04 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of aero-engine sigmatron is tested with adjustable measuring basis device |
| CN106625398A (en) * | 2017-01-03 | 2017-05-10 | 哈尔滨工业大学 | Assembling method and device of aero-engine rotor |
Family Cites Families (1)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| GB0822859D0 (en) * | 2008-12-16 | 2009-01-21 | Rolls Royce Plc | Timing analysis |
-
2020
- 2020-08-03 CN CN202010766675.2A patent/CN112254966B/en active Active
Patent Citations (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN101799354A (en) * | 2010-02-09 | 2010-08-11 | 西安交通大学 | Method for testing assembly performance of rotor of aircraft engine |
| CN103292959A (en) * | 2012-03-05 | 2013-09-11 | 沈阳黎明航空发动机(集团)有限责任公司 | Method for analyzing and detecting parameter of dynamic balance simulation rotor of engine rotor component |
| CN103791816A (en) * | 2014-02-14 | 2014-05-14 | 哈尔滨工业大学 | Aircraft engine rotor assembly method and device based on concentricity and perpendicularity measurement |
| CN105241598A (en) * | 2015-11-04 | 2016-01-13 | 中国航空动力机械研究所 | Method and system for measuring pre-tightening force of engine rotor |
| CN106289120A (en) * | 2016-08-11 | 2017-01-04 | 中国航空工业集团公司沈阳发动机设计研究所 | A kind of aero-engine sigmatron is tested with adjustable measuring basis device |
| CN106625398A (en) * | 2017-01-03 | 2017-05-10 | 哈尔滨工业大学 | Assembling method and device of aero-engine rotor |
Non-Patent Citations (3)
| Title |
|---|
| 叠加优化技术在高压涡轮转子维修装配中的应用;王洪明;《长沙航空职业技术学院学报》;20190630;第19卷(第2期);摘要 * |
| 涡扇发动机先进装配工艺与装备;孙贵青等;《航空制造技术》;20171115(第22期);73页左栏倒数第2段、74页右栏倒数第1段-75页右栏第1段，图7 * |
| 燃气轮机拉杆转子拉伸装配过程的变形分析;袁奇等;《中国动力工程学会透平专业委员会2011年学术研讨会论文集》;20111021;第1-2页，图1 * |
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