CN114082988A - Method for repairing aero-engine cold and hot end blade - Google Patents
Method for repairing aero-engine cold and hot end blade Download PDFInfo
- Publication number
- CN114082988A CN114082988A CN202111260755.1A CN202111260755A CN114082988A CN 114082988 A CN114082988 A CN 114082988A CN 202111260755 A CN202111260755 A CN 202111260755A CN 114082988 A CN114082988 A CN 114082988A
- Authority
- CN
- China
- Prior art keywords
- blade
- repairing
- points
- scanning
- edge
- Prior art date
- Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
- Pending
Links
- 238000000034 method Methods 0.000 title claims abstract description 42
- 238000004519 manufacturing process Methods 0.000 claims abstract description 13
- 238000013461 design Methods 0.000 claims abstract description 11
- 238000005457 optimization Methods 0.000 claims abstract description 10
- 238000004458 analytical method Methods 0.000 claims abstract description 9
- 238000004088 simulation Methods 0.000 claims abstract description 8
- 239000000463 material Substances 0.000 claims description 26
- 239000000843 powder Substances 0.000 claims description 22
- 230000008569 process Effects 0.000 claims description 15
- 229910052751 metal Inorganic materials 0.000 claims description 11
- 239000002184 metal Substances 0.000 claims description 11
- 238000006073 displacement reaction Methods 0.000 claims description 9
- 238000009826 distribution Methods 0.000 claims description 9
- 238000012546 transfer Methods 0.000 claims description 7
- 230000009467 reduction Effects 0.000 claims description 6
- 238000003892 spreading Methods 0.000 claims description 6
- 230000007480 spreading Effects 0.000 claims description 6
- 239000000758 substrate Substances 0.000 claims description 6
- 229910001069 Ti alloy Inorganic materials 0.000 claims description 3
- 239000003086 colorant Substances 0.000 claims description 3
- 238000013016 damping Methods 0.000 claims description 3
- 239000000155 melt Substances 0.000 claims description 3
- 238000012545 processing Methods 0.000 claims description 3
- 230000011218 segmentation Effects 0.000 claims description 3
- 230000005540 biological transmission Effects 0.000 claims description 2
- 239000011148 porous material Substances 0.000 claims description 2
- 238000007789 sealing Methods 0.000 claims description 2
- 238000005507 spraying Methods 0.000 claims description 2
- 229910000601 superalloy Inorganic materials 0.000 claims description 2
- 238000005516 engineering process Methods 0.000 abstract description 13
- 230000008439 repair process Effects 0.000 abstract description 12
- 238000004372 laser cladding Methods 0.000 abstract description 7
- 238000003466 welding Methods 0.000 abstract description 6
- 239000000654 additive Substances 0.000 abstract description 4
- 230000000996 additive effect Effects 0.000 abstract description 4
- 238000010923 batch production Methods 0.000 abstract description 2
- 238000010586 diagram Methods 0.000 description 6
- 230000000694 effects Effects 0.000 description 3
- 230000006872 improvement Effects 0.000 description 3
- 239000007787 solid Substances 0.000 description 3
- 238000005452 bending Methods 0.000 description 2
- 239000003795 chemical substances by application Substances 0.000 description 2
- 238000012423 maintenance Methods 0.000 description 2
- 238000002844 melting Methods 0.000 description 2
- 230000008018 melting Effects 0.000 description 2
- 238000007639 printing Methods 0.000 description 2
- 238000010146 3D printing Methods 0.000 description 1
- 125000000174 L-prolyl group Chemical group [H]N1C([H])([H])C([H])([H])C([H])([H])[C@@]1([H])C(*)=O 0.000 description 1
- 229910045601 alloy Inorganic materials 0.000 description 1
- 239000000956 alloy Substances 0.000 description 1
- 238000005266 casting Methods 0.000 description 1
- 230000008859 change Effects 0.000 description 1
- 239000011248 coating agent Substances 0.000 description 1
- 238000000576 coating method Methods 0.000 description 1
- 238000005242 forging Methods 0.000 description 1
- 239000000446 fuel Substances 0.000 description 1
- 238000000227 grinding Methods 0.000 description 1
- 238000003754 machining Methods 0.000 description 1
- 230000002035 prolonged effect Effects 0.000 description 1
- 238000011160 research Methods 0.000 description 1
- 238000012827 research and development Methods 0.000 description 1
- 239000013585 weight reducing agent Substances 0.000 description 1
Images
Classifications
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
- B22F10/00—Additive manufacturing of workpieces or articles from metallic powder
- B22F10/80—Data acquisition or data processing
- B22F10/85—Data acquisition or data processing for controlling or regulating additive manufacturing processes
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B22—CASTING; POWDER METALLURGY
- B22F—WORKING METALLIC POWDER; MANUFACTURE OF ARTICLES FROM METALLIC POWDER; MAKING METALLIC POWDER; APPARATUS OR DEVICES SPECIALLY ADAPTED FOR METALLIC POWDER
- B22F10/00—Additive manufacturing of workpieces or articles from metallic powder
- B22F10/20—Direct sintering or melting
- B22F10/28—Powder bed fusion, e.g. selective laser melting [SLM] or electron beam melting [EBM]
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B33—ADDITIVE MANUFACTURING TECHNOLOGY
- B33Y—ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3-D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3-D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING
- B33Y10/00—Processes of additive manufacturing
-
- B—PERFORMING OPERATIONS; TRANSPORTING
- B33—ADDITIVE MANUFACTURING TECHNOLOGY
- B33Y—ADDITIVE MANUFACTURING, i.e. MANUFACTURING OF THREE-DIMENSIONAL [3-D] OBJECTS BY ADDITIVE DEPOSITION, ADDITIVE AGGLOMERATION OR ADDITIVE LAYERING, e.g. BY 3-D PRINTING, STEREOLITHOGRAPHY OR SELECTIVE LASER SINTERING
- B33Y50/00—Data acquisition or data processing for additive manufacturing
- B33Y50/02—Data acquisition or data processing for additive manufacturing for controlling or regulating additive manufacturing processes
-
- G—PHYSICS
- G06—COMPUTING; CALCULATING OR COUNTING
- G06F—ELECTRIC DIGITAL DATA PROCESSING
- G06F30/00—Computer-aided design [CAD]
- G06F30/10—Geometric CAD
- G06F30/15—Vehicle, aircraft or watercraft design
-
- Y—GENERAL TAGGING OF NEW TECHNOLOGICAL DEVELOPMENTS; GENERAL TAGGING OF CROSS-SECTIONAL TECHNOLOGIES SPANNING OVER SEVERAL SECTIONS OF THE IPC; TECHNICAL SUBJECTS COVERED BY FORMER USPC CROSS-REFERENCE ART COLLECTIONS [XRACs] AND DIGESTS
- Y02—TECHNOLOGIES OR APPLICATIONS FOR MITIGATION OR ADAPTATION AGAINST CLIMATE CHANGE
- Y02P—CLIMATE CHANGE MITIGATION TECHNOLOGIES IN THE PRODUCTION OR PROCESSING OF GOODS
- Y02P10/00—Technologies related to metal processing
- Y02P10/25—Process efficiency
Landscapes
- Engineering & Computer Science (AREA)
- Chemical & Material Sciences (AREA)
- Materials Engineering (AREA)
- Physics & Mathematics (AREA)
- Manufacturing & Machinery (AREA)
- Geometry (AREA)
- Theoretical Computer Science (AREA)
- General Physics & Mathematics (AREA)
- Computational Mathematics (AREA)
- Aviation & Aerospace Engineering (AREA)
- Mathematical Analysis (AREA)
- Mathematical Optimization (AREA)
- Pure & Applied Mathematics (AREA)
- Computer Hardware Design (AREA)
- Evolutionary Computation (AREA)
- General Engineering & Computer Science (AREA)
- Automation & Control Theory (AREA)
- Plasma & Fusion (AREA)
- Turbine Rotor Nozzle Sealing (AREA)
Abstract
The invention discloses a method for repairing a blade at a cold end and a hot end of an aero-engine, which comprises the following steps: (1) 3D scanning the blade; (2) reverse modeling; (3) optimizing the structure; (4) simulation analysis; (5) 3D manufacturing; the technology of the invention utilizes the modern mature three-dimensional scanning technology and reverse modeling, and plays a role in repairing while reconstructing and restoring the original structural shape; the blade is designed by using four basic principles and innovative means, the internal optimization design is carried out on the blade, and the internal innovative design can improve the performance and the service life of the blade while the weight of the blade is reduced; finally, batch production is carried out by utilizing an advanced 3D additive manufacturing technology; the cost of the blade part reduced by the technology is far lower than that of the blade part directly purchased with a new part or laser cladding/welding repair. The method is not only suitable for the repair market, but also suitable for the restoration of other neighborhoods and other parts.
Description
Technical Field
The invention belongs to the technical field of engine blade repair, and particularly relates to a method for repairing a cold-hot end blade of an aircraft engine.
Background
The aero-engine is a short plate in China from design, research and development to maintenance, once engine parts are damaged, the aero-engine needs to be directly purchased from abroad, the period is long, and the price is high, so that the repair of the aero-engine parts becomes a necessary matter. The blade of the aircraft engine, which is the most easily damaged part, is a key medium for converting aviation fuel into flying power, the working temperature of the blade can reach more than 1000 ℃, the working condition is very severe, and the damage is an inevitable event.
At present, the method for repairing the engine blade mainly comprises laser cladding/welding and self-adaptive grinding, and the method has various problems: (1) the method is a mainstream repairing method for foreign engine maintenance enterprises, but is still not applied to the field of civil aviation in a mature and industrialized manner due to technical blockade and lack of domestic research on the aeroengine material process; (2) common ultrahigh-speed laser cladding equipment is expensive and is mostly imported equipment, the thickness of a repair coating applicable to the equipment is limited, for example, the thickness is 0.05-1mm, and the seriously damaged blade cannot be restored in a reducing manner; (3) only the body part can be repaired by using laser cladding/welding equipment, and the body part cannot be continuously produced in large batch; (4) the blade repaired by the laser cladding/welding equipment can keep the original performance of the part at most, and the performance of the repaired blade can not be improved greatly.
Disclosure of Invention
Aiming at the problem, the invention provides a method for repairing a cold and hot end blade of an aircraft engine, which is used for solving the problems in the prior art and adopts the technical scheme that:
a method for repairing a blade at a cold end and a hot end of an aircraft engine comprises the following steps: (1) 3D scanning the blade; (2) reverse modeling; (3) optimizing the structure; (4) simulation analysis; (5) 3D manufacturing;
the specific process of the step (1) comprises the following steps:
step (1-1): uniformly spraying a developer (white powder) on the surface of a blade part, sticking signal acquisition mark points on the surface of the part after the developer is completely dried, wherein the signal acquisition mark points are stuck and arranged to form V-shaped distribution, each visible surface is stuck with at least 3 signal acquisition mark points as far as possible, and the larger the size is, the more the signal acquisition mark points are;
step (1-2): and scanning the blades by using a blue light three-dimensional scanner and Vtop Studio software to generate point cloud data, and splicing.
The specific process of reverse modeling in the step (2) comprises the following operation steps:
step (2-1): importing data: importing the point cloud STL format model into Geomagic Design X software to perform reverse modeling operation;
step (2-2): dividing the field: classifying the surface patches into different fields by using an automatic segmentation option according to the curvature and the characteristics of the scanning data, and distinguishing the fields by using different colors;
step (2-3): creating a sketch: taking a plane passing through the blade tip or the blade root as an initial plane, intercepting a group of contour lines, drawing a sketch contour (containing 6 points and 6 lines) by using a spline curve, creating a group of parallel planes by taking the initial plane as a reference, and creating a sketch by adopting the mode;
step (2-4): creating a guideline: using a spline curve command in the 3D sketch and taking points with the same property in each group of curves as key points to create 6 guide lines;
step (2-5): creating an entity: using lofting commands, using each group of sketches as an outline, and using the 3D sketches as guide lines to create a model of the blade;
step (2-6): deviation comparison: selecting the body deviation on the right side of the software, and checking the deviation between the entity blade and the blade obtained by scanning;
step (2-7): drawing tenons and leaf tops: the sketch is drawn by drawing the section lines obtained by sketching with plane truncation. Drawing the whole tenon by using related characteristic commands; the blade top needs to use a corresponding cylindrical or conical surface to intercept the outline of the outer edge of the blade, so that the blade needs to be correspondingly extended, corresponding blade information is obtained after relevant data is consulted, and the blade is correspondingly intercepted after the radius of the circle and the interception surface are determined;
step (2-8): and (4) rounding the blade root and other positions.
Further, the structural optimization mode of the step (3) mainly comprises:
(3-1): unnecessary materials below the surface of the blade are removed through hollowing, and a sub-surface hole-shaped structure is arranged, so that the weight can be reduced;
(3-2): a flexible load transfer surface is adopted, so that the transfer area is increased, and the load can be reduced;
(3-3): by changing the stiffness distribution over the inner pore structure, the frequency can be adjusted;
(3-4): stress concentration can be reduced by improving and changing a load transmission path;
(3-5): dynamic friction is caused by adding a contact interface on the internal porous structure so as to increase the damping of the internal structure;
(3-6): the surface texture structure is arranged, so that the resistance reduction, noise reduction and sealing can be realized, and the external pneumatic performance is enhanced.
Further, the specific process of the simulation analysis in the step (4) is as follows:
(4-1): determining input parameters: the main parameters to be determined are: the number of stages of the blades, the position of a rotating shaft, the rotating speed and the assembling mode of the blades;
(4-2): determining a property of the material;
(4-3): and (3) carrying out grid division: dividing the model mesh by adopting an unstructured tetrahedral mesh, wherein the average mesh quality is 0.76, the root part of the blade is locally provided with meshes, the total mesh number is 99611, and the node number is 162684;
(4-4): setting boundary conditions and loads: firstly, adding a local cylindrical coordinate system according to the position of a rotating shaft, applying radial displacement constraint C to two side surfaces of a blade root, applying circumferential displacement constraint D to one side surface, applying axial displacement constraint B to the end surface of an extending tooth at the bottommost part of the blade root, and finally adding a centrifugal load;
(4-5): and (6) analyzing results.
Further, the operation process of the step (5) of 3D printing is as follows: before the laser beam starts scanning, a powder spreading device firstly pushes metal powder to a substrate of a forming cylinder, the laser beam selectively melts the powder on the substrate according to a filling scanning line of a current layer to process the current layer, then the forming cylinder descends by a distance of one layer thickness, a powder material cylinder ascends by a distance of a certain thickness, the powder spreading device spreads the metal powder on the processed current layer, equipment is adjusted to data of the profile of the next layer to process, and the layer-by-layer processing is carried out until the whole part is processed.
Further, in the step (2-4), the number of the key points is determined, and under the condition of meeting the precision, the fewer the number of the key points is, the better the key point recommendation value containing 6 points is: number of key points of suction side and pressure side: 5-10, and the number is required to be consistent, otherwise, the blade forming is influenced; the number of key points of the four curves of the air inlet edge and the air outlet edge is as follows: 4-6; number of contour lines: 4-8, determined by the size of the blade, the accuracy is met, and the generation rate and quality of the blade model are ensured.
Further, in the step (2), a point-line-plane-body is used as a theoretical basis of reverse modeling, and the following points are: the intersection point of 2 positions of the air inlet edge and the pressure edge of the suction edge, the intersection point of 2 positions of the air outlet edge and the pressure edge of the suction edge, and the 2-position boundary point of the air inlet edge and the air outlet edge for distinguishing the suction surface and the pressure surface are 6 positions in total; line: six curves are formed among the 6 points, wherein each curve is formed by fitting a plurality of key points; dough making: 6 faces are formed by several sets of closed curves: a suction surface, a pressure surface, an air inlet side suction surface, an air outlet side pressure surface and an air outlet side suction surface; body: the 6 surfaces and the 2 surfaces of the blade root and the blade tip are combined into a complete blade entity.
Preferably, the step (4-2) material is set to the IN718 material.
Further, the materials of 3D fabrication are mainly titanium alloys and superalloys.
Advantageous effects
The technology of the invention utilizes the modern mature three-dimensional scanning technology and reverse modeling, and plays a role in repairing while reconstructing and restoring the original structural shape; the blade is designed by using four basic principles and innovative means, the internal optimization design is carried out on the blade, and the internal innovative design can improve the performance and the service life of the blade while the weight of the blade is reduced; finally, batch production is carried out by utilizing an advanced 3D additive manufacturing technology; the cost of the blade part reduced by the technology is far lower than that of the blade part directly purchased with a new part or laser cladding/welding repair. The method is not only suitable for the repair market, but also suitable for the restoration of other neighborhoods and other parts.
Drawings
FIG. 1 is a schematic diagram of the process steps of the present invention;
FIG. 2 is a schematic diagram of a blade to be repaired and a signal acquisition punctuation thereon according to the present invention;
FIG. 3 is a schematic blade contour diagram of the present invention;
FIG. 4 is a schematic diagram of contour line creation in reverse modeling according to the present invention;
FIG. 5 is a schematic diagram of entity creation using lofting in reverse modeling according to the present invention;
FIG. 6 is a graph of deviation versus effect in reverse modeling according to the present invention;
FIG. 7 is a schematic view of a blade model after reverse modeling is completed according to the present invention;
FIG. 8 is a schematic view of an innovative structural optimization of the blade of the present invention;
FIG. 9 is a comparison of the model before and after repair of the blade according to the invention;
FIG. 10 is a schematic representation of meshing of a blade simulation analysis of the present invention;
FIG. 11 is a schematic diagram of boundary conditions and load settings for a blade simulation analysis of the present invention.
Detailed Description
The technical solutions in the embodiments of the present invention are clearly and completely described below, and it is obvious that the described embodiments are only a part of the embodiments of the present invention, and not all 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, a method for repairing a blade at a cold end and a hot end of an aircraft engine includes the following steps: (1) scanning a blade; (2) reverse modeling; (3) optimizing the structure; (4) simulation analysis; (5) 3D manufacturing;
the specific process is as follows:
1. leaf 3D scanning:
(1) the developing agent (white powder) is uniformly sprayed on the surface of the blade part, after the developing agent is completely dried, signal acquisition mark points are pasted on the surface of the part, the signal acquisition mark points are pasted and arranged to form V-shaped distribution, each visible surface is pasted with at least 3 signal acquisition mark points as far as possible, the larger the size is, the more the signal acquisition mark points are, as shown in figure 2.
(2) And scanning the blades by using a blue light three-dimensional scanner and Vtop Studio software to generate point cloud data, and splicing.
Reverse modeling: and the generation rate and quality of the blade model are ensured by adopting a point-line-plane-body as a theoretical basis of reverse modeling and a key point quantitative principle.
The steps of the invention adopt point-line-plane-body as the theoretical basis of reverse modeling, as shown in figure 3, (1) point: the intersection point of 2 positions of the air inlet edge and the pressure edge of the suction edge, the intersection point of 2 positions of the air outlet edge and the pressure edge of the suction edge, and the 2-position boundary point of the air inlet edge and the air outlet edge for distinguishing the suction surface and the pressure surface are 6 positions in total; (2) line: six curves are formed among the 6 points, wherein each curve is formed by fitting a plurality of key points; (3) dough making: 6 faces are formed by several sets of closed curves: a suction surface, a pressure surface, an air inlet side suction surface, an air outlet side pressure surface and an air outlet side suction surface; (4) body: the 6 surfaces and the 2 surfaces of the blade root and the blade tip are combined into a complete blade entity.
The method comprises the following specific steps:
(1) importing data: importing the point cloud STL format model into Geomagic Design X software to perform reverse modeling operation;
(2) dividing the field: classifying the surface patches into different fields by using an automatic segmentation option according to the curvature and the characteristics of the scanning data, and distinguishing the fields by using different colors;
(3) creating a sketch: taking the plane passing through the blade tip or the blade root as an initial plane, intercepting a group of contour lines, as shown in FIG. 4a, drawing a sketch contour (comprising 6 points and 6 lines) by using a spline curve, as shown in FIG. 4b, creating a group of parallel planes by taking the initial plane as a reference, and creating a sketch by adopting the mode, as shown in FIG. 4 c;
(4) creating a guideline: using spline curve commands in the 3D sketch, 6 guide lines are created with the same property points in each set of curves as key points. The selection principle of the number of the key points is as follows: under the condition of meeting the precision, the fewer the key points are, the better the key points are;
key points recommendation (containing 6 points): number of key points of suction side and pressure side: 5-10, and the number is required to be consistent or the blade forming is influenced; the number of key points of the four curves of the air inlet edge and the air outlet edge is as follows: 4-6; number of contour lines: 4-8, the size of the blade is determined, and the precision is met;
(5) creating an entity: using loft commands, outline each set of sketches, and create a model of the blade for the guideline, as shown in FIG. 5, with the 3D sketches;
(6) deviation comparison: selecting the body deviation on the right side of the software, and checking the deviation between the solid blade and the blade obtained by scanning, as shown in FIG. 6;
(7) drawing tenons and leaf tops: the sketch is drawn by drawing the section lines obtained by sketching with plane truncation. Drawing the whole tenon by using related characteristic commands; the blade top needs to use a corresponding cylindrical or conical surface to intercept the outline of the outer edge of the blade, so that the blade needs to be correspondingly extended, corresponding blade information is obtained after relevant data is consulted, and the blade is correspondingly intercepted after the radius of the circle and the interception surface are determined;
(8) and others: rounding off the blade root and other positions;
the final completed blade model is shown in fig. 7.
Structure optimization: the sub-surface porous structure and the surface texture structure can be used as an important means for structural optimization, so that the overall performance of the blade is improved, and the weight is reduced.
The basic principle is as follows: innovative structural optimization is performed on the blade, as shown in fig. 8, 1, the application of the subsurface porous structure can reduce weight (hollowing out, removing unnecessary material below the surface of the blade), reduce load (flexible load transfer surface, increasing transfer area), modulate frequency (changing rigidity distribution on the internal porous structure), reduce stress concentration (improving and changing load transfer path), increase internal structure damping (adding contact interface on the internal porous structure, causing dynamic friction), and save material and cost for manufacturing; 2. the application of the surface texture structure can reduce resistance, reduce noise, seal and enhance the external pneumatic performance.
A comparison graph of the leaves before and after repair is shown in FIG. 9, and a three-dimensional model of an original model is established through scanning and repair by taking the leaves as an example, and the internal structure of the model is subjected to weight reduction and innovative design on the basis of the three-dimensional model. The weight of the modified and upgraded three-dimensional model is reduced by 18.4% compared with that of the original entity three-dimensional model.
Simulation analysis:
(1) input parameter determination:
taking the above blade as an example, the main parameters to be determined are: the number of blade stages, the position of a rotating shaft, the rotating speed and the blade assembly mode. The rotating speed is inquired through relevant literature data, and the design rotating speed of FT8 is 12000 rpm. The vane stage number, the vane assembling mode and the position of the rotating shaft are mainly determined by a sectional view of an FT8 model.
(2) Material properties:
the material was designated as IN718, and both the original and modified models were analyzed and compared using the same material.
(3) Grid division:
and dividing the model mesh by adopting an unstructured tetrahedral mesh, wherein the average mesh quality is 0.76, and the root part of the blade is locally provided with the mesh. The total grid number is 99611, and the node number is 162684, as shown in fig. 10.
(4) Boundary conditions and loads:
firstly, a local cylindrical coordinate system is added according to the position of a rotating shaft, radial displacement constraint C is applied to two side faces of a blade root, circumferential displacement constraint D is applied to one side face, and axial displacement constraint B is applied to the end face of an extending tooth at the bottommost part of the blade root. Finally, the centrifugal load is added as shown in FIG. 11.
(5) And (4) analyzing results:
through comparing and analyzing the parameters of the original model and the improved model such as stress distribution, radial deformation distribution, maximum stress, maximum radial deformation, former ten-order modal distribution and the like, the weight of the three-dimensional model with the optimized structure is reduced by 18.4 percent compared with the original entity three-dimensional model. Position of center of mass
. The radial position of the center of mass of the blade changes the most. The maximum equivalent stress of the improved model is reduced by 4.6%, and the radial deformation is reduced by 14%. The main reasons are to reduce the blade mass and to position the center of mass closer to the center of rotation to reduce the effect of centrifugal bending moments.
Modal analysis shows that except for the improvement of the first-order bending vibration frequency by 6.6 percent, other second-order to tenth-order modal frequencies are reduced to different degrees and are far away from the rotating speed frequency.
In a word, after the model is optimized, the quality of the original model can be effectively reduced, and the maximum equivalent stress and radial deformation are reduced, so that the safety coefficient of parts is improved, and the service life of the parts is prolonged.
Manufacturing: 3D metal printing makes the blade, from 0 to 1, can change blade structure or material, also can be as the route of mass production.
Traditional blade machining is through casting and forging, mill material reduction again, and this technique utilizes present advanced metal 3D (selective laser melting SLM) printing apparatus additive manufacturing, no matter how complicated structure can become the real object, therefore also provides more infinite possibility for structural design optimization.
The principle is as follows: the basic principle of the selective laser melting technology is that a three-dimensional solid model of a part is designed on a computer by using three-dimensional modeling software such as Pro/e, UG and CATIA, then the three-dimensional model is sliced and layered by slicing software to obtain profile data of each section, a filling scanning path is generated by the profile data, and equipment controls laser beams to selectively melt metal powder materials of each layer according to the filling scanning lines and gradually stack the metal powder materials into the three-dimensional metal part.
Before the laser beam starts scanning, a powder spreading device firstly pushes metal powder to a substrate of a forming cylinder, the laser beam selectively melts the powder on the substrate according to a filling scanning line of a current layer to process the current layer, then the forming cylinder descends by a distance of one layer thickness, a powder material cylinder ascends by a distance of a certain thickness, the powder spreading device spreads the metal powder on the processed current layer, equipment is adjusted to data of the profile of the next layer to process, and the layer-by-layer processing is carried out until the whole part is processed.
At present, materials manufactured by 3D are mainly titanium alloy and high-temperature alloy, and new materials can be gradually applied.
The technical scheme of the invention has the technical advantages that:
1. the technology of the invention utilizes the modern mature three-dimensional scanning technology and reverse modeling, and plays a role in repairing while reconstructing and restoring the original structural shape;
2. the blade is designed by using four basic principles and innovative means, the structure of the blade is optimally designed, the weight of the blade is reduced, and the performance and the service life of the blade can be improved;
3. no matter how complex the structure is, parts are manufactured by using advanced 3D additive manufacturing technology, and a solid foundation is provided for continuous optimization and improvement, including material improvement, and 3D manufacturing can be carried out in batch;
4. the cost of the blade part reduced by the technology is far lower than that of the blade part directly purchased with a new part or laser cladding/welding repair.
5. The technical method is not only suitable for the repair market, but also suitable for the reduction of other neighborhoods and other parts.
Claims (9)
1. A method for repairing a blade at a cold end and a hot end of an aircraft engine is characterized by comprising the following steps: (1) 3D scanning the blade; (2) reverse modeling; (3) optimizing the structure; (4) simulation analysis; (5) 3D manufacturing;
the step (2) comprises the following steps:
step (2-1): importing data: importing the point cloud STL format model into Geomagic Design X software to perform reverse modeling operation;
step (2-2): dividing the field: classifying the surface patches into different fields by using an automatic segmentation option according to the curvature and the characteristics of the scanning data, and distinguishing the fields by using different colors;
step (2-3): creating a sketch: taking a plane passing through the blade tip or the blade root as an initial plane, intercepting a group of contour lines, drawing a sketch contour containing 6 points and 6 lines by using a sample curve, creating a group of parallel planes by taking the initial plane as a reference, and creating a sketch by adopting the mode;
step (2-4): creating a guideline: using a spline curve command in the 3D sketch and taking points with the same property in each group of curves as key points to create 6 guide lines;
step (2-5): creating an entity: using lofting commands, using each group of sketches as an outline, and using the 3D sketches as guide lines to create a model of the blade;
step (2-6): deviation comparison: selecting the body deviation on the right side of the software, and checking the deviation between the entity blade and the blade obtained by scanning;
step (2-7): drawing tenons and leaf tops: drawing a sketch by adopting a plane interception method and drawing a section line obtained by drawing the sketch, and drawing the whole tenon by using a related characteristic command; correspondingly extending the blades, looking up related data to obtain corresponding blade information, and correspondingly intercepting the blades after determining the radiuses of the circular and intercepting surfaces;
step (2-8): and (5) rounding off the blade root position by rounding off operation.
2. The method for repairing a cold-hot end blade of an aircraft engine as claimed in claim 1, wherein the specific process of the step (1) comprises:
step (1-1): uniformly spraying a developer on the surface of a blade part, pasting signal acquisition mark points on the surface of the part after the developer is completely dried, wherein the signal acquisition mark points are pasted and arranged in a V-shaped distribution, at least 3 signal acquisition mark points are pasted on each visible surface, and the larger the size is, the more the signal acquisition mark points are;
step (1-2): and scanning the blades by using a blue light three-dimensional scanner and Vtop Studio software to generate point cloud data, and splicing.
3. A method for repairing an aircraft engine hot and cold end blade according to claim 1, wherein the structural optimization of the step (3) comprises:
(3-1): unnecessary materials below the surface of the blade are removed through hollowing, and a sub-surface hole-shaped structure is arranged, so that the weight can be reduced;
(3-2): a flexible load transfer surface is adopted, so that the transfer area is increased, and the load can be reduced;
(3-3): by changing the stiffness distribution over the inner pore structure, the frequency can be adjusted;
(3-4): stress concentration can be reduced by improving and changing a load transmission path;
(3-5): dynamic friction is caused by adding a contact interface on the internal porous structure so as to increase the damping of the internal structure;
(3-6): the surface texture structure is arranged, so that the resistance reduction, noise reduction and sealing can be realized, and the external pneumatic performance is enhanced.
4. The method for repairing the cold and hot end blade of the aircraft engine as claimed in claim 1, wherein the specific process of the step (4) is as follows:
(4-1): determining input parameters: the parameters to be determined are: the number of stages of the blades, the position of a rotating shaft, the rotating speed and the assembling mode of the blades;
(4-2): determining a property of the material;
(4-3): and (3) carrying out grid division: dividing the model mesh by adopting an unstructured tetrahedral mesh, wherein the average mesh quality is 0.76, the root part of the blade is locally provided with meshes, the total mesh number is 99611, and the node number is 162684;
(4-4): setting boundary conditions and loads: firstly, adding a local cylindrical coordinate system according to the position of a rotating shaft, applying radial displacement constraint C to two side surfaces of a blade root, applying circumferential displacement constraint D to one side surface, applying axial displacement constraint B to the end surface of an extending tooth at the bottommost part of the blade root, and finally adding a centrifugal load;
(4-5): and (6) analyzing results.
5. The method for repairing a cold-hot end blade of an aircraft engine as claimed in claim 1, wherein the operation process of the step (5) is as follows: before the laser beam starts scanning, a powder spreading device firstly pushes metal powder to a substrate of a forming cylinder, the laser beam selectively melts the powder on the substrate according to a filling scanning line of a current layer to process the current layer, then the forming cylinder descends by a distance of one layer thickness, a powder material cylinder ascends by a distance of a certain thickness, the powder spreading device spreads the metal powder on the processed current layer, equipment is adjusted to data of the profile of the next layer to process, and the layer-by-layer processing is carried out until the whole part is processed.
6. The method for repairing the cold and hot end blade of the aircraft engine as claimed in claim 1, wherein the number of the key points in the step (2-4) is determined, the fewer the number of the key points is, the better the number of the key points is, and the recommended value of the number of the key points comprising 6 points is: number of key points of suction side and pressure side: 5-10, and the number is required to be consistent, otherwise, the blade forming is influenced; the number of key points of the four curves of the air inlet edge and the air outlet edge is as follows: 4-6; number of contour lines: 4-8, determined by the size of the blade, the accuracy is met, and the generation rate and quality of the blade model are ensured.
7. The method for repairing the cold and hot end blade of the aircraft engine as claimed in claim 1, wherein in the step (2), a point-line-plane-body is used as a theoretical basis for reverse modeling, and the point: the intersection point of 2 positions of the air inlet edge and the pressure edge of the suction edge, the intersection point of 2 positions of the air outlet edge and the pressure edge of the suction edge, and the 2-position boundary point of the air inlet edge and the air outlet edge for distinguishing the suction surface and the pressure surface are 6 positions in total; line: six curves are formed among the 6 points, wherein each curve is formed by fitting a plurality of key points; dough making: 6 faces are formed by several sets of closed curves: a suction surface, a pressure surface, an air inlet side suction surface, an air outlet side pressure surface and an air outlet side suction surface; body: the 6 surfaces and the 2 surfaces of the blade root and the blade tip are combined into a complete blade entity.
8. A method for repairing an aircraft engine hot and cold end blade according to claim 4, wherein the material of step (4-2) is IN718 material.
9. A method of repairing an aircraft engine hot and cold end blade according to claim 5, wherein the 3D fabricated materials are primarily titanium alloys and superalloys.
Priority Applications (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111260755.1A CN114082988A (en) | 2021-10-28 | 2021-10-28 | Method for repairing aero-engine cold and hot end blade |
Applications Claiming Priority (1)
| Application Number | Priority Date | Filing Date | Title |
|---|---|---|---|
| CN202111260755.1A CN114082988A (en) | 2021-10-28 | 2021-10-28 | Method for repairing aero-engine cold and hot end blade |
Publications (1)
| Publication Number | Publication Date |
|---|---|
| CN114082988A true CN114082988A (en) | 2022-02-25 |
Family
ID=80297998
Family Applications (1)
| Application Number | Title | Priority Date | Filing Date |
|---|---|---|---|
| CN202111260755.1A Pending CN114082988A (en) | 2021-10-28 | 2021-10-28 | Method for repairing aero-engine cold and hot end blade |
Country Status (1)
| Country | Link |
|---|---|
| CN (1) | CN114082988A (en) |
Cited By (3)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114686732A (en) * | 2022-04-19 | 2022-07-01 | 北航（四川）西部国际创新港科技有限公司 | High-temperature alloy repair material and preparation method thereof, additive remanufacturing method and remanufacturing service evaluation method of high-temperature alloy repair part |
| CN116174649A (en) * | 2023-02-28 | 2023-05-30 | 中国航发北京航空材料研究院 | Preparation method of air inlet side precision casting for maintaining high-temperature alloy guide vane |
| CN117301527A (en) * | 2023-10-30 | 2023-12-29 | 科路睿(天津)生物技术有限公司 | Single-path slicing method and printing method for four-axis 3D printing sheet model |
Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103093065A (en) * | 2013-02-20 | 2013-05-08 | 天津工业大学 | Three-dimensional model rebuilding method of aviation engine blade tip missing parts |
| CN103919631A (en) * | 2014-04-28 | 2014-07-16 | 四川大学 | Manufacturing method for jaw defect individual restoration |
| CN106599359A (en) * | 2016-11-08 | 2017-04-26 | 南方科技大学 | Design method of hollow leaf filled with spherical mesh structure, and engine |
| CN107818590A (en) * | 2017-09-11 | 2018-03-20 | 童和平 | Free form surface product reverse-engineering based on three-dimensional optical scanning models forming method |
| CN108145393A (en) * | 2017-11-29 | 2018-06-12 | 株洲丰发精工实业有限公司 | A kind of aero-engine compressor blade and its processing method |
| CN110781573A (en) * | 2019-08-13 | 2020-02-11 | 中国航发贵阳发动机设计研究所 | Reverse design method for rotor blade of aircraft engine |
| CN110853132A (en) * | 2019-10-09 | 2020-02-28 | 安徽徽博先临三维云打印技术有限公司 | Three-dimensional modeling method for complex curved surface part for 3D printing |
| CN111859577A (en) * | 2020-07-27 | 2020-10-30 | 江苏科技大学 | Direct layering method for laser additive repair |
| CN111962069A (en) * | 2020-09-02 | 2020-11-20 | 中国航发北京航空材料研究院 | Deformed high-temperature alloy and stainless steel gas compressor rotor blade tip repairing method and tool |
| CN112765732A (en) * | 2021-01-25 | 2021-05-07 | 沈阳工业大学 | Aviation blade topology optimization design method based on selective laser melting process |
| CN113409464A (en) * | 2021-06-15 | 2021-09-17 | 中国科学院近代物理研究所 | Method for reversely mapping electrode part of cyclotron |
-
2021
- 2021-10-28 CN CN202111260755.1A patent/CN114082988A/en active Pending
Patent Citations (11)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN103093065A (en) * | 2013-02-20 | 2013-05-08 | 天津工业大学 | Three-dimensional model rebuilding method of aviation engine blade tip missing parts |
| CN103919631A (en) * | 2014-04-28 | 2014-07-16 | 四川大学 | Manufacturing method for jaw defect individual restoration |
| CN106599359A (en) * | 2016-11-08 | 2017-04-26 | 南方科技大学 | Design method of hollow leaf filled with spherical mesh structure, and engine |
| CN107818590A (en) * | 2017-09-11 | 2018-03-20 | 童和平 | Free form surface product reverse-engineering based on three-dimensional optical scanning models forming method |
| CN108145393A (en) * | 2017-11-29 | 2018-06-12 | 株洲丰发精工实业有限公司 | A kind of aero-engine compressor blade and its processing method |
| CN110781573A (en) * | 2019-08-13 | 2020-02-11 | 中国航发贵阳发动机设计研究所 | Reverse design method for rotor blade of aircraft engine |
| CN110853132A (en) * | 2019-10-09 | 2020-02-28 | 安徽徽博先临三维云打印技术有限公司 | Three-dimensional modeling method for complex curved surface part for 3D printing |
| CN111859577A (en) * | 2020-07-27 | 2020-10-30 | 江苏科技大学 | Direct layering method for laser additive repair |
| CN111962069A (en) * | 2020-09-02 | 2020-11-20 | 中国航发北京航空材料研究院 | Deformed high-temperature alloy and stainless steel gas compressor rotor blade tip repairing method and tool |
| CN112765732A (en) * | 2021-01-25 | 2021-05-07 | 沈阳工业大学 | Aviation blade topology optimization design method based on selective laser melting process |
| CN113409464A (en) * | 2021-06-15 | 2021-09-17 | 中国科学院近代物理研究所 | Method for reversely mapping electrode part of cyclotron |
Cited By (6)
| Publication number | Priority date | Publication date | Assignee | Title |
|---|---|---|---|---|
| CN114686732A (en) * | 2022-04-19 | 2022-07-01 | 北航（四川）西部国际创新港科技有限公司 | High-temperature alloy repair material and preparation method thereof, additive remanufacturing method and remanufacturing service evaluation method of high-temperature alloy repair part |
| CN114686732B (en) * | 2022-04-19 | 2022-10-18 | 北航（四川）西部国际创新港科技有限公司 | High-temperature alloy repair material and preparation method thereof, and additive remanufacturing method and re-service evaluation method of high-temperature alloy repair part |
| CN116174649A (en) * | 2023-02-28 | 2023-05-30 | 中国航发北京航空材料研究院 | Preparation method of air inlet side precision casting for maintaining high-temperature alloy guide vane |
| CN116174649B (en) * | 2023-02-28 | 2024-04-16 | 中国航发北京航空材料研究院 | Preparation method of air inlet side precision casting for maintaining high-temperature alloy turbine guide vane |
| CN117301527A (en) * | 2023-10-30 | 2023-12-29 | 科路睿(天津)生物技术有限公司 | Single-path slicing method and printing method for four-axis 3D printing sheet model |
| CN117301527B (en) * | 2023-10-30 | 2024-04-05 | 科路睿(天津)生物技术有限公司 | Single-path slicing method and printing method for four-axis 3D printing sheet model |
Similar Documents
| Publication | Publication Date | Title |
|---|---|---|
| CN114082988A (en) | Method for repairing aero-engine cold and hot end blade | |
| US8881392B2 (en) | Method of repairing machined components such as turbomachine blades or blades of blisks | |
| US7797828B2 (en) | Adaptive machining and weld repair process | |
| CN109918701B (en) | Turbine disc crack propagation simulation method based on piecewise weight function | |
| CN108595887B (en) | Simulation calculation and prediction method for abrasion evolution of polycrystalline CBN (cubic boron nitride) abrasive particles | |
| CN105739440A (en) | Adaptive machining method of wide-chord hollow fan blade | |
| CN106777783B (en) | Method for predicting blade cracks of aircraft engine | |
| CN109128168A (en) | A kind of synchronous powder feeding system increasing material manufacturing Process Planning Method based on structure feature | |
| CN107900622B (en) | A kind of band coating casing processing method of thin-walled | |
| Vasilopoulos et al. | CAD-based aerodynamic optimization of a compressor stator using conventional and adjoint-driven approaches | |
| EP4257804A1 (en) | Mistuning analysis systems and methods for repairing inspected bladed rotors | |
| CN105302995B (en) | A kind of method of numerical simulation of optimum blade roll milling mold and Blank Design | |
| CN104915500B (en) | Powder injection forming turbine and its optimum structure design method | |
| CN102108883B (en) | Semi-rotation speed nuclear turbine final stage moving blade | |
| Salnikov et al. | A centrifugal compressor impeller: a multidisciplinary optimization to improve its mass, strength, and gas-dynamic characteristics | |
| CN111353249B (en) | Non-circular vent hole integrated design optimization method for turbine sealing disc | |
| CN114165477B (en) | Axial ultrasonic through-flow fan serial configuration and serial configuration optimization method | |
| CN110489925B (en) | Gas turbine fir-type blade root and blade groove modular design method | |
| CN114547786A (en) | Method for planning measuring points of bent and twisted thin-wall blade in consideration of machining deformation | |
| CN113468788A (en) | Deformation coordination design method for engine blisk | |
| CN106339543A (en) | Construction method for hydraulic turbine flow passage component morphology repairing model | |
| McGlumphy et al. | Numerical investigation of tandem airfoils for subsonic axial-flow compressor blades | |
| CN104942551A (en) | Inclined stator blade ring and manufacturing method thereof | |
| EP4257802A1 (en) | Structural repair analysis system and method for inspected bladed rotors | |
| Dang et al. | A model reconstruction strategy with geometric parameters adjustment function for repairing damaged blade |
Legal Events
| Date | Code | Title | Description |
|---|---|---|---|
| PB01 | Publication | ||
| PB01 | Publication | ||
| SE01 | Entry into force of request for substantive examination | ||
| SE01 | Entry into force of request for substantive examination | ||
| WD01 | Invention patent application deemed withdrawn after publication | ||
| WD01 | Invention patent application deemed withdrawn after publication |
Application publication date: 20220225 |