CN115600413A - Basic reliability modeling method for aviation turbofan engine - Google Patents
Basic reliability modeling method for aviation turbofan engine Download PDFInfo
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Abstract
The application belongs to the field of design of aero-engines, and provides a basic reliability modeling method for an aero-turbofan engine.
Description
Technical Field
The application belongs to the field of aero-engine design, and particularly relates to a basic reliability modeling method for an aero turbofan engine.
Background
At present, the development of domestic engines obtains a series of achievements, but no matter which type of engine is not affected by the reliability problem, the reason is also that the reliability design system is difficult to establish due to extremely complex development process, complex structure, complex failure mechanism and large number of parts of the aero-engine, and the reliability design cannot effectively fall to the real place. To realize the quantitative design and evaluation of reliability, it is a necessary way to establish an effective basic reliability modeling method. However, no effective reliability modeling method is seen in the industry at present. At present, basic reliability modeling of various industries in China basically connects a part list into a reliability block diagram by using series-parallel connection models, is suitable for simple products and is not suitable for complex products of aero-engines.
As shown in fig. 1, basic reliability modeling mainly includes that parts lists are connected into a reliability block diagram by series-parallel connection models in a general manner, and problems in several aspects mainly exist in the application of an aero-engine, 1) the aero-engine has extremely large part number, is difficult to directly model and express by graphs, and a modeling result cannot guide application; 2) The aircraft engine development process is extremely complex, the structure level is complex, and the model cannot support the evaluation verification requirements of different levels; 3) The reliability test of a single part of the aircraft engine is high in verification cost and high in dependence on simulation verification evaluation, but the aircraft engine is complex in structure and various in failure mechanism, and the reliability design needs to be specific to the specific failure mechanism of a specific part, so that a basic reliability model established based on the part cannot meet the design requirement.
The reliability work of the conventional aeroengine is mainly characterized by qualitative analysis, wherein the FMECA of the qualitative analysis relates to basic reliability modeling, and the specific method mainly comprises the steps of selecting series-parallel block diagrams of part of main parts based on experience, wherein the modeling is not systematic, incomplete and normative, and the effectiveness of a reliability model is seriously influenced. Because the reliability design of the aircraft engine is essentially a system engineering, the design is more lean and refined than deterministic design, and the requirements on systematicness, integrity and precision are higher, the association between a reliability model and specific FMECA analysis and risk evaluation work is not large due to an irregular modeling method, and the quantitative estimation of risks cannot be realized.
Therefore, a basic reliability modeling method suitable for complex structures of aero-engines is urgently needed.
Disclosure of Invention
The application aims to provide a basic reliability modeling method for an aviation turbofan engine, and the basic reliability modeling method is used for solving the problem that basic reliability modeling cannot be performed on an aviation engine complex structure by adopting a series-parallel model in the prior art.
The technical scheme of the application is as follows: a basic reliability modeling method for an aviation turbofan engine comprises the following steps: classifying and classifying the general failure configuration of the aero-engine according to the structure level of the aero-engine, the classification list among all levels, the failure list of each level, the mapping relation among different levels and the mapping relation among different structure levels and failures to form a typical aero-engine failure configuration; the method comprises the steps of exporting EBOM lists of all the existing aircraft engines, carrying out data cleaning according to preset cleaning rules, carrying out typical part failure configuration, supplementing load information of each structure according to design requirements and usage requirements, and forming a structure design basic information requirement table of each structure; the method comprises the steps of exporting EBOM lists of all the existing aircraft engines, carrying out data cleaning according to preset cleaning rules to form typical part failure configurations, supplementing load information of each structure in the typical part failure configurations according to design requirements and usage requirements, and forming a structure design basic information requirement table of each structure; establishing a failure configuration primitive library according to a structural design basic information requirement table of each structure and a mapping relation of failure configurations of the aero-engine of a plurality of levels; then, EBOM lists of all the current novel aero-engines are derived regularly, after data cleaning is carried out according to current cleaning rules, data of the EBOM lists of the novel aero-engines are respectively placed into corresponding positions of the established failure configuration element library according to the current different-level mapping relation, and the failure configuration element library is continuously updated; establishing a mapping relation with a corresponding structure according to the basic design information of each structure in the failure configuration primitive library, and carrying out mapping marking on a design scheme and a failure configuration appointed level; and continuously updating the newly-imported EBOM list data mapping mark through the current mapping relation.
Preferably, an EBOM list is derived by adopting an EXCEL format to form a structural design basic information requirement table; and (3) putting the basic design information of the failure configuration primitive library into a corresponding Excel table, and carrying out mapping marking on the design scheme and the failure configuration convention level in the Excel table.
Preferably, the failure configuration element library comprises a component function library, a typical component and outsourcing component library, a failure mode library, a failure mechanism library and a danger site library.
Preferably, the failure configuration of the aircraft engine comprises a structure type convention layer and a failure type convention layer, wherein the structure type convention layer comprises a whole machine layer, a component functional layer, a component layer, a part and a small purchased component layer; the failure class commitment layer comprises a failure mode layer, a failure mechanism layer and a typical characteristic failure component layer.
Preferably, the specific establishment method of the mapping mark comprises the following steps:
selecting a required functional component from a pull-down menu according to the EBOM information;
the pull-down menu screens out a function list of the corresponding component according to the component name, preliminary prejudgment is carried out on the component function according to the EBOM information, a mapping relation between the component and the component function is established, and the component function is selected from the pull-down menu;
according to the EBOM information, selecting a required component layer, an assembly layer, a part layer and a small assembly layer from the menu, and carrying out structure level agreement;
the pull-down menu screens out a typical component list of the corresponding component according to the component name, and mapping between the component and the typical component is carried out from the pull-down menu according to the EBOM information;
selecting a mapping relation between the part or the outsourced external member and the typical part from a pull-down menu according to the EBOM information;
according to the EBOM information, the load and the material, the mapping between the failure modes of the part or the external cooperation external member and the typical part is selected from a pull-down menu, and a row is added to each failure mode;
according to the EBOM information, the load and the material, the mapping of the failure mechanisms of the part or the external cooperation external member and the typical part is selected from a pull-down menu, and one row is added to each failure mechanism;
screening and determining a dangerous component characteristic set of the typical part as a pull-down menu according to the name of the typical part, and selecting a corresponding dangerous part from the pull-down menu according to the typical part, the name and a failure mechanism;
recording the number of dangerous positions according to the structural drawing and the names of the dangerous positions;
and selecting a corresponding component function layer, a failure mode layer, a failure mechanism layer and a typical characteristic site layer from the pull-down menu to carry out mark mapping of the failure contract layer.
Preferably, the method for establishing the failure configuration element library comprises the following steps:
establishing a failure configuration element list;
obtaining an FMECA analysis result of any model, matching the FMECA analysis result with corresponding data of different levels of the established failure configuration primitive list, and putting the FMECA analysis data into the FMECA analysis result and replacing the corresponding data of the failure configuration primitive list.
Preferably, the method for establishing the failure configuration primitive list comprises the following steps:
defining the whole machine layer;
defining the component of the component layer;
performing functional classification of the components in the component layer;
classifying the components in the component layer;
classifying the parts and typical parts in the outsourcing small component layer;
classifying failure modes in the failure mode layer;
classifying failure mechanisms in the failure mechanism layer;
and classifying the realized positions in the typical characteristic failure part layer.
Preferably, the structural design basic information requirement table comprises indexes of parent component diagram numbers, child component diagram numbers, component names, quantities, material types, load environments, component lists and overall distribution.
Preferably, after the structural design is finished, carrying out technical state management and application on a basic reliability model of the complex structure of the aircraft engine, wherein the method for the technical state management and application comprises the steps of describing the basic reliability model of the engine in different states by adopting a ternary array, and the basic reliability model comprises M { model number, derivative serial number, batch number } -data-derivative-GENERAL DRAWING; the model description method is consistent with a standard channel, the serial number of the derived model is consistent with the suffix of the derived model name, the batch number is a technical verification machine, an engineering verification machine, a prototype machine, identification, design and finalization and outfield service, each matrix marks modeling completion time, designer name and drawing number according to, if the scheme name is changed in each stage and the derived model, the model is directly inherited from a basic reliability model, and an EBOM list is modified to complete basic reliability modeling.
According to the basic reliability modeling method for the aviation turbofan engine, all structures and failure characteristics of the aviation engine are sorted, classified and hierarchically divided, an EBOM list of the aviation engine is derived on the basis of the sorting and hierarchical division to fill basic information of failure configurations of the aviation engine, then a failure configuration primitive library is established to sort and summarize all the structures and failure configurations in the aviation engine, a mapping relation with a corresponding structure is established according to the failure configuration primitive library and by combining basic design information of each structure, a mapping mark of a design scheme and a failure configuration convention level is formed, the failure configuration convention level, the list primitive and the configuration establishing management method based on reliability transfer rules are achieved, and a fast, standard, complete and effective basic reliability model is achieved through ordered layer-by-layer mapping.
Drawings
In order to more clearly illustrate the technical solutions provided by the present application, the following briefly introduces the accompanying drawings. It is to be understood that the drawings described below are merely exemplary of some embodiments of the application.
FIG. 1 is a general basic reliability modeling diagram in the background art;
FIG. 2 is a schematic illustration of a failure configuration convention hierarchy for an aircraft engine complex structure according to the present application;
FIG. 3 is a schematic view of an aircraft engine failure configuration build scenario of the present application;
FIG. 4 is a diagram illustrating a basic information requirement table structure according to the present application;
FIG. 5 is a schematic diagram illustrating basic reliability modeling of a complex structure of an aircraft engine according to the present application;
FIG. 6 is a schematic diagram of the mechanism of the present application for mapping the design and failure configuration convention hierarchy.
Detailed Description
In order to make the implementation objects, technical solutions and advantages of the present application clearer, the technical solutions in the embodiments of the present application will be described in more detail below with reference to the drawings in the embodiments of the present application.
The patent firstly proposes a reliability modeling method based on failure configuration, establishes an agreed hierarchy and an element list of the failure configuration and a configuration establishment management method according to a reliability transmission rule, and establishes the failure configuration as shown in figure 2.
The method comprises the following steps:
step S100, failure configuration convention hierarchy division and classification
As shown in fig. 2-3, classifying and classifying the general failure configuration of the aircraft engine according to the structural hierarchy of the aircraft engine, the classification list between each hierarchy, the failure list between each hierarchy and the mapping relationship between different hierarchies and the mapping relationship between different structural hierarchies and failures to form a typical failure configuration of the aircraft engine;
preferably, the failure configuration of the aircraft engine comprises a structure type convention layer and a failure type convention layer, wherein the structure type convention layer comprises a whole machine layer, a component function layer, a component layer, a part and an outsourcing small component layer; the failure class commitment layer comprises a failure mode layer, a failure mechanism layer and a typical characteristic failure component layer.
Step S200, EBOM information collection
As shown in fig. 4, an EBOM list of all existing aircraft engines capable of reflecting the assembly level of the aircraft engine structure is derived by using the EXCEL format, and includes material information, a typical component failure configuration is formed, and load information of each structure is supplemented in the typical component failure configuration according to design requirements and usage requirements, so as to form a structure design basic information requirement table of each structure; the basic information requirement table mainly comprises allocated indexes, basic information requirements, parent component diagram numbers, child component diagram numbers, component names, quantities, material types, load environments, component lists and overall allocated indexes, and the specific form is shown in fig. 4.
Step S300, establishing and dynamically updating failure configuration primitive database
Establishing a failure configuration primitive library according to a structural design basic information requirement table of each structure and a mapping relation of failure configurations of the aero-engine of a plurality of levels; and then, regularly exporting the EBOM lists of all the current novel aircraft engines, after data cleaning is carried out according to the current cleaning rule, respectively putting all data of the EBOM lists of the novel aircraft engines into corresponding positions of the established failure configuration element library according to the current different-level mapping relation, and continuously updating the failure configuration element library.
Preferably, the method for establishing the failure configuration element library comprises the following steps:
s310, establishing a failure configuration primitive list;
the method specifically comprises the following steps:
s111, defining the complete machine of the complete machine layer;
the functional layer of the whole machine is one of the following functions: the system comprises a power supply module, an emergency oil discharge module, a mode conversion module, an engine control module, a communication module and a health management module.
S312, defining the component layer component;
the component layer is one of the following components: the engine comprises an air inlet casing, a fan, an intermediate casing, a high-pressure compressor, a combustion chamber, an afterburner, a high-pressure turbine, a low-pressure turbine, a turbine rear casing, an intermediate casing, an air inlet casing, an external structure, a nozzle, a transmission system, a lubricating oil system and a fuel oil and control system, wherein the lubricating oil system and the fuel oil and control system are not taken as main parts and are not further subdivided in the patent.
S313, classifying the functions of the components in the component layer;
the method comprises the following steps:
an air inlet casing: the gas rectification function is provided; the device has the functions of bearing and transferring load; the device has the function of adjusting the inlet airflow; the anti-icing function is achieved; providing the support function of the low pressure rotor bearing.
A fan: providing compressed air for engine operation; the anti-icing function is achieved; has the functions of bearing and transmitting; providing compressed air for the cabin, turbine cooling, etc.
An intermediary case: providing an internal culvert flow path and an external culvert flow path; has the function of internal air release; providing an engine main mounting joint; providing a fulcrum support function; providing access for mounting and wiring; has the function of bearing force.
A high-pressure compressor: providing high-pressure air required by the operation of an engine; the device has a force transmission function; the function of providing rotor support; a function of providing power transfer; has the adjustable function.
A combustion chamber: the device has a combustion function; providing a suitable temperature field; the function of providing a cooling air channel is provided; the device has the function of transmitting axial force and torque.
An afterburner: force application connection; the afterburner is used for stable combustion; force bearing installation; mounting a test sensing part; discharging residual fuel oil inside; emergency oil drainage; adjusting the mixing area of the inner culvert and the outer culvert; and (4) a heat insulation function.
A high-pressure turbine: the expansion device has the function of expanding and doing work by the high-pressure turbine; the function of transferring load is provided; has the functions of sealing and providing cold air; the expansion work of the low-pressure turbine is achieved; the function of transferring load is provided; possesses the function of obturating.
Turbine rear case: the function of providing rotor support; the function of transferring load is provided; forming a gas channel; providing a bleed air channel; the function of fixing the fulcrum bearing; the auxiliary installation function is provided.
External structure: the function of conveying the specified working medium required by the engine and the function of installing and fixing the engine as accessories are provided.
Outer culvert casing: constructing an outer culvert airflow channel; force transfer function; installation and maintenance functions; the device has the functions of fixing accessories, pipelines and supports.
Spraying a pipe: the high-temperature gas expansion and accelerated discharge device has the functions of high-temperature gas expansion and accelerated discharge, adjusting the cross-sectional area and injecting airflow in an engine compartment.
A transmission system: providing a power extraction function; a function of transmitting torque; providing a sealing function; providing installation and interface functions; providing a rotation speed measuring function; providing a pan interface function.
S314, classifying the components in the component layer;
the method comprises the following steps:
an air inlet casing: welding a combined key on the air inlet casing; a rectifier blade, an actuator cylinder assembly, a thin/rear adjustable blade; a cap cover, an oil inlet pipe assembly,
A fan: the device comprises fan rotor blades, a fan rotor assembly balance assembly, a two-stage and three-stage disc welding assembly, a two-stage and three-stage casing welding machine adding assembly, a two-stage and three-stage retainer ring assembly, a fan air-entraining pipe, a low-pressure compressor drum barrel, a disc drum assembly, a multifunctional shaft assembly, a lifting lug and an air-entraining pipe.
An intermediary case: intermediate casing assembly, intermediate casing welding assembly, shunt ring assembly, mounting joint ball seat and ball body
A high-pressure compressor: the high-pressure gas turbine engine comprises rotor blades, a high-pressure gas compressor rotor assembly, stator blades, a rotor disc drum assembly, a high-pressure gas compressor front casing assembly, a gas compressor wheel disc, a sealing disc, an upper half part and a lower half part of a front casing and an upper half part and a lower half part of an extension casing; the upper/lower half part of the rear casing; a high pressure compressor rotor assembly; an actuator cylinder;
a combustion chamber: a fuel manifold with nozzles; a head with a venturi; a nozzle housing; a headed outer wall assembly; a sleeve; nozzles (standard); an oil inlet pipe; a pre-diffuser; a pre-diffuser assembly; a fuel branch pipe assembly; a main nozzle; a rear housing assembly; an inner wall; an outer air inlet hood; the structure comprises a casing inner sleeve, a flow guide ring, an outer casing front section, a combustion chamber outer sleeve with a diffuser, a casing front section with a diffuser, an outer casing rear section, a rear sleeve wall, a rear mounting edge, a middle mounting edge, an outer casing front section, a combustion chamber outer sleeve with a diffuser and an outer support wall.
An afterburner: the device comprises a first-zone oil injection ring, a second-zone fuel main pipe, a third-zone fuel main pipe, a fourth-zone fuel main pipe, a V-zone fuel main pipe, a main stabilizer, a diffuser outer wall rear section, a mixer unit body, a mixer case, a barrel section, a stressing barrel shell, a bearing ring, an ion flame detector, an oil discharge port, an emergency oil discharge device, a heat shield front section and a heat shield rear section.
A high-pressure turbine: the turbine front sealing disc, the high-pressure turbine working blades, the high-pressure turbine rear shaft, the high-pressure turbine rotor balancing assembly, the working blade assembly, the guide blade assembly, the high-pressure turbine casing, the high-pressure turbine disc, the drum shaft, the front baffle and the air guide pipe.
A low-pressure turbine: the low-pressure turbine comprises low-pressure turbine guide blades, a low-pressure turbine rotor balance assembly, low-pressure turbine working blades, a low-pressure turbine disc, a low-pressure turbine shaft, a low-pressure secondary casing, a sealing ring and a sealing ring between discs.
Turbine rear case: the device comprises an inclined support plate bearing frame, a rear casing, a pull rod, an inclined support plate bearing frame, a bearing seat and an auxiliary mounting section.
External structure: external pipelines, brackets and clamp assemblies;
outer culvert casing: outer culvert casing body assembly, hole detection maintenance and mounting seat
Spraying a pipe: adjusting plate assembly and spray pipe hydraulic actuator cylinder
A transmission system: the booster pump comprises a booster pump intermediate gear shaft, a central driven gear shaft, a central driving bevel gear, a central driven bevel gear, a booster fuel pump gear shaft, a central driving bevel gear support, a booster fuel pump gear shaft, a fly-attached transmission rod, a driving spiral bevel gear, a driven spiral bevel gear, a shell welding assembly, a central gear case shell, a mounting seat assembly, a bearing, a sealing assembly, a front bearing case, a labyrinth sealing seat, an accessory case shell assembly, a bearing mounting seat assembly, a bearing, a quick-release ring, a lifting lug and a speed measuring gear shaft;
s315, classifying the parts and typical parts in the outsourcing small component layer;
the method comprises the following steps:
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S316, classifying failure modes in the failure mode layer;
including some typical failure modes, but not limited to the following:
and (3) fracture class: cracks, local fractures, overall fractures, coating chipping, coating peeling and the like;
wear type: local wear, overall wear, and the like;
and (4) deformation: dimensional changes, local deformations, global deformations, etc.;
erosion: ablation, corrosion, erosion, abrasion, etc.;
connection type: bonding, welding, loosening, peeling and the like;
and (3) seepage: air leakage, gas leakage, oil leakage, fuel leakage, etc.
S317, classifying failure mechanisms in the failure mechanism layer;
including but not limited to: static strength, low cycle fatigue, high cycle fatigue, creep endurance, corrosion, wear, aging, stress relaxation, abnormal loading, thermal deformation mismatch, chemical oxidation, material instability, electrical displacement.
And S318, classifying the realized positions in the typical characteristic failure part layer.
The method comprises the following steps: the device comprises a welding part, a blade front edge, a blade rear edge, a casing opening, a disc web, a vent hole, a tenon, a mortise, a honeycomb, a blade body, a corner, a boss, a comb tooth, a blade tip, a coating, a fillet, a blade air inlet and outlet edge, a mounting hole edge, a mounting edge root and a blade back switching sub-cavity.
The failure configuration of the aircraft engine formed by sorting the data comprises a complete machine layer, a component functional layer, an assembly layer, a part and outsourcing small assembly layer, a failure mode layer, a failure mechanism layer and a typical characteristic failure component layer.
S320, obtaining an FMECA analysis result of any model, matching the FMECA analysis result with corresponding data of different levels of the established failure configuration primitive list, putting the FMECA analysis data into the failure configuration primitive list, and replacing the corresponding data of the failure configuration primitive list, wherein the updating of all failure configuration elements such as typical components, assemblies, parts, failure modes, failure mechanisms, dangerous positions and the like is included, and therefore the real-time updating of the failure configuration primitive library is achieved.
Step S400, design plan and failure configuration contract level mapping mark
As shown in fig. 5-6, a mapping relation with a corresponding structure is established according to the basic design information of each structure in the failure configuration primitive library, and mapping marks of a design scheme and a failure configuration convention level are carried out; and continuously updating the newly-introduced EBOM list data mapping marks through the current mapping relation.
Putting the basic design information of the failure configuration primitive library into a corresponding Excel table, and carrying out mapping marking on a design scheme and a failure configuration convention level in the Excel table, wherein the information is shown in fig. 5:
preferably, the specific establishment method of the mapping mark is as follows:
s410, selecting a required functional component from a pull-down menu according to the EBOM information; the pull-down menu includes: the engine comprises an air inlet casing, a fan, an intermediate casing, a high-pressure compressor, a combustion chamber, an afterburner, a high-pressure turbine, a low-pressure turbine, a turbine rear casing, an intermediate casing, an air inlet casing, an external structure, an outer culvert casing, a spray pipe, a transmission system and a lubricating oil system. If it is imported per component, all components are automatically marked uniformly as component names corresponding to the column.
And S420, screening out a function list of the corresponding component according to the component name by the pull-down menu, preliminarily prejudging the component function according to the EBOM information, establishing a mapping relation between the component and the component function, and selecting the component function from the pull-down menu, wherein the pull-down menu specifically comprises the corresponding content in the S313.
S430, selecting a required component layer, an assembly layer, a part layer and a small assembly layer from the menu according to the EBOM information, and carrying out structure level agreement;
s440, screening a typical component list of the corresponding component by the pull-down menu according to the component name, and mapping the component and the typical component from the pull-down menu according to the EBOM information; the pull-down menu specifically includes the corresponding contents in S314.
S450, selecting a mapping relation between the part or the outsource external member and the typical part from the pull-down menu according to the EBOM information; the pull-down menu specifically includes the corresponding content in S315.
S460, according to the EBOM information, the load and the material, selecting the pull-down menu to map the failure modes of the part or the external cooperation external member and the typical part, and adding one row for each failure mode; the pull-down menu specifically includes the corresponding content in S316.
S470, according to the EBOM information, the load and the material, the mapping of the failure mechanisms of the parts or the outsourced external components and the typical parts is selected from a pull-down menu, and a row is added to each failure mechanism; the pull-down menu specifically includes the corresponding content in S317.
S480, screening and determining a dangerous component characteristic set of the typical part as a pull-down menu according to the name of the typical part, and selecting a corresponding dangerous part from the pull-down menu according to the typical part, the name and a failure mechanism; the pull-down menu specifically includes the corresponding content in S318.
S490, recording the number of dangerous positions according to the structural drawing and the names of the dangerous positions;
s4100, selecting corresponding component function layers, failure mode layers, failure mechanism layers and typical characteristic part layers from the pull-down menu to map the marks of the failure contract layers.
By the steps, mapping and marking of all typical parts, assemblies, parts, failure modes, failure mechanisms and dangerous parts of the aircraft engine are completed, and the relation between the structure and the reliability is clear and comprehensive.
Step S500, technical state management and application
And after the structural design is finished, carrying out technical state management and application of a basic reliability model of the complex structure of the aero-engine.
The method for managing and applying the technical state comprises the steps of describing basic reliability models of engines in different states by using a ternary array, wherein the basic reliability models comprise M { model number, derivative serial number and batch number } -DATE-derivative-GENERAL DRAWING, and the expression requirements of each item are as follows:
wherein, 1) the description method of the model is consistent with the standard channel;
2) The serial number of the derived model is consistent with the suffix named by the derived model, and the preferred suggestions are defined as D-A, D-B, D-C, D-D, \ 8230;
3) The batch number is a technical verifier, an engineering verifier, a prototype, identification, design and finalization, and outfield service, and is defined as C1.C2.C3.C4.C5.C6 once, and if a plurality of batches still exist in different stages, the batches are expressed by adopting extension forms of C1-01, C1-02 and the like;
4) Each matrix marks the modeling completion time, the name of a designer and the number of a drawing according to the modeling completion time;
5) And if the scheme naming is changed, each stage and derivative model directly inherits the basic reliability model, and an EBOM list is modified to perfect the basic reliability modeling.
After the basic reliability model is established, when the basic reliability model is used, the required data such as components, assemblies or parts can be found by searching the model number, the derivative serial number, the batch number or the name of a designer, so that the inheritance and the comparative analysis of technical states of different schemes are facilitated.
According to the method, all structures and failure characteristics of the aero-engine are sorted, classified and hierarchically divided, an EBOM list of the aero-engine is derived on the basis to fill basic information of failure configurations of the aero-engine, then a failure configuration element library is established to sort and summarize all structures and failure configurations in the aero-engine, a mapping relation with a corresponding structure is established according to the failure configuration element library and basic design information of each structure, a mapping mark of a design scheme and a failure configuration contract level is formed, the failure configuration contract level based on reliability transmission rules, the element list and configuration establishment management method are achieved, and a rapid, standard, complete and effective basic reliability model is achieved through ordered layer-by-layer mapping.
Has the following advantages:
1) The classification of thousands of parts is realized by mapping the hierarchical levels and the typical element list of the hierarchical levels, the number of elements of the reliability model is reduced, each element data is regularly maintained by a special database, the basic data of the built model can be conveniently obtained, the effectiveness of modeling application is improved, and the reliability model is not limited by the large number of the parts and incapability of displaying through the hierarchical display of the EXCEL;
2) The elements of the reliability modeling are arranged under the parts, a failure mode, a failure mechanism and a typical failure part level are arranged, the content of the reliability modeling is greatly expanded, and the model can be in one-to-one correspondence with the structural integrity design to realize the effective estimation of the structural reliability design;
3) The requirement for the reliability model in the development of the aircraft engine is deeply met, the basic reliability modeling appointed level of the aircraft engine is established according to the requirement and is divided into 8 levels of complete machine, part function, assembly, part, failure mode, failure mechanism and dangerous part, the design and evaluation requirements of each level in different stages can be met, and the requirement for multi-stage verification and evaluation of the engine is met.
4) The EBOM lists based on the aircraft engine are mapped one by one, the integrity of parts is ensured, in addition, the dependence on a designer is reduced by establishing typical libraries of all levels, a mapping flow, a template and a primary database are established, the dependence on the experience of the designer is reduced, and the method is more standard, complete and effective.
5) Because the development difficulty of the aero-engine is extremely high, the number of derived models of each type of engine is large, in addition, certain differences exist among the development stages, the development stages are complex, the technical states are multiple, but inheritance of different degrees exists, a matrix-based basic reliability model management method is provided, inheritance and comparative analysis of the technical states of different schemes are facilitated, and the modeling time is shortened. The method is suitable for the characteristics of multiple derivative models, multiple states, multiple development stages and the like of the aircraft engine, and is convenient for realizing the information management of huge modeling data.
The above description is only for the specific embodiments of the present application, but the scope of the present application is not limited thereto, and any changes or substitutions that can be easily conceived by those skilled in the art within the technical scope of the present application should be covered within the scope of the present application. Therefore, the protection scope of the present application shall be subject to the protection scope of the claims.
Claims (9)
1. A basic reliability modeling method for an aviation turbofan engine is characterized by comprising the following steps:
classifying and classifying the general failure configuration of the aero-engine according to the structure level of the aero-engine, the classification list among all levels, the failure list of each level, the mapping relation among different levels and the mapping relation among different structure levels and failures to form a typical aero-engine failure configuration;
the method comprises the steps of exporting EBOM lists of all the existing aircraft engines, carrying out data cleaning according to preset cleaning rules to form typical part failure configurations, supplementing load information of each structure in the typical part failure configurations according to design requirements and usage requirements, and forming a structure design basic information requirement table of each structure;
establishing a failure configuration primitive library according to a structural design basic information requirement table of each structure and a mapping relation of failure configurations of the aero-engine of a plurality of levels; then, regularly exporting EBOM lists of all the current novel aircraft engines, after data cleaning is carried out according to current cleaning rules, respectively putting all data of the EBOM lists of the novel aircraft engines into corresponding positions of the established failure configuration element library according to current different-level mapping relations, and continuously updating the failure configuration element library;
establishing a mapping relation with a corresponding structure according to the basic design information of each structure in the failure configuration primitive library, and carrying out mapping marking on a design scheme and a failure configuration appointed level; and continuously updating the newly-introduced EBOM list data mapping marks through the current mapping relation.
2. The aircraft turbofan engine basic reliability modeling method of claim 1 wherein: an EBOM list is derived by adopting an EXCEL format to form a structure design basic information requirement table; and (3) putting the basic design information of the failure configuration primitive library into a corresponding Excel table, and carrying out mapping marking on a design scheme and a failure configuration convention level in the Excel table.
3. The aircraft turbofan engine basic reliability modeling method of claim 1 wherein: the failure configuration element library comprises a component function library, a typical component and outsourcing component library, a failure mode library, a failure mechanism library and a dangerous part library.
4. The aircraft turbofan engine basic reliability modeling method of claim 1 wherein: the failure configuration of the aero-engine comprises a structure type convention layer and a failure type convention layer, wherein the structure type convention layer comprises a whole machine layer, a component function layer, a component layer, a part and an outsourcing small component layer; the failure class commitment layer comprises a failure mode layer, a failure mechanism layer and a typical characteristic failure component layer.
5. The aircraft turbofan engine basic reliability modeling method of claim 4 wherein the specific establishment method of the mapping indicia is:
selecting a required functional component from a pull-down menu according to the EBOM information;
the pull-down menu screens out a function list of the corresponding component according to the component name, preliminary prejudgment is carried out on the component function according to the EBOM information, a mapping relation between the component and the component function is established, and the component function is selected from the pull-down menu;
according to the EBOM information, selecting a required component layer, an assembly layer, a part layer and a small assembly layer from the menu to carry out structure level agreement;
the pull-down menu screens out a typical component list of the corresponding component according to the component name, and mapping between the component and the typical component is carried out from the pull-down menu according to the EBOM information;
selecting a mapping relation between the part or the outsourced external member and the typical part from a pull-down menu according to the EBOM information;
according to the EBOM information, the load and the material, the mapping between the failure modes of the part or the external cooperation external member and the typical part is selected from a pull-down menu, and a row is added to each failure mode;
according to the EBOM information, load and material, the mapping of failure mechanisms of the part or the external cooperation external member and the typical part is selected from a pull-down menu, and one row is added for each failure mechanism;
screening and determining a dangerous component characteristic set of the typical part as a pull-down menu according to the name of the typical part, and selecting a corresponding dangerous part from the pull-down menu according to the typical part, the name and a failure mechanism;
recording the number of dangerous positions according to the structural drawing and the names of the dangerous positions;
and selecting a corresponding component function layer, a failure mode layer, a failure mechanism layer and a typical characteristic site layer from the pull-down menu to carry out mark mapping of the failure contract layer.
6. The aircraft turbofan engine basic reliability modeling method of claim 4 wherein the method of building the failure configuration primitive library comprises:
establishing a failure configuration element list;
obtaining an FMECA analysis result of any model, matching the FMECA analysis result with corresponding data of different levels of the established failure configuration primitive list, and putting the FMECA analysis data in the corresponding data of the failure configuration primitive list and replacing the FMECA analysis data with the corresponding data of the failure configuration primitive list.
7. The aircraft turbofan engine basic reliability modeling method of claim 6 wherein the method of building the list of failure configuration primitives comprises:
defining the whole machine layer;
defining the component of the component layer;
performing functional classification of the components in the component layer;
classifying the components in the component layer;
classifying the parts and typical parts in the outsourcing small component layer;
classifying failure modes in the failure mode layer;
classifying failure mechanisms in the failure mechanism layer;
and classifying the realized positions in the typical characteristic failure part layer.
8. The aircraft turbofan engine basic reliability modeling method of claim 1 wherein: the structural design basic information requirement table comprises a parent component diagram number, a child component diagram number, a component name, a quantity, a material type, a load environment, a component list and overall distribution indexes.
9. The aircraft turbofan engine basic reliability modeling method of claim 1 wherein: after the structure design is finished, carrying out technical state management and application of a basic reliability model of a complex structure of the aircraft engine, wherein the method for the technical state management and application comprises the steps of describing the basic reliability model of the engine with different states by adopting a ternary array, wherein the basic reliability model comprises M { model number, derivative serial number and batch number } -data-design-GENERALDAWING; the model description method is consistent with a standard channel, the serial number of the derived model is consistent with a suffix named by the derived model, the batch number is a technical verifier, an engineering verifier, a prototype machine, identification, design and finalization and external field service, each matrix marks modeling completion time, designer name and figure number according to the model, and if the scheme naming is changed, each phase and the derived model directly inherits a basic reliability model and modifies an EBOM list to complete basic reliability modeling.
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