CN112182875B - Fatigue design method for test-oriented helicopter rotor wing metal piece - Google Patents
Fatigue design method for test-oriented helicopter rotor wing metal piece Download PDFInfo
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Abstract
The invention discloses a fatigue design method for a test-oriented helicopter rotor wing metal piece, which comprises the following steps: determining a proofing fatigue limit corresponding to a target life test, calculating the maximum loading load of the fatigue test carried out under different materials and friction and corrosion modes of the rotor wing metal part, and analyzing under the test load; according to the materials of the rotor moving parts in the inverse calculation fatigue test and the test proof fatigue limit under the abrasion mode state; taking stress ratio effect into consideration to carry out static load correction, calculating the equivalent dynamic load of each characteristic load proofing, and evaluating the fatigue life of the structure; and determining the optimized size of the dangerous part of the rotor wing moving part, guiding the design and meeting the requirement of the fatigue test assessment target. The method has the advantages that the strength, the rigidity and the boundary constraint of the test piece and the matching part under the characteristic test acceleration load can meet the requirements or not, and the early failure caused by unreasonable design of the test scheme is avoided.
Description
Technical Field
The invention belongs to the field of helicopter rotor strength design, and relates to a test-oriented method for evaluating the fatigue life of a helicopter rotor metal part and designing the key size of a structure.
Background
The strength design of the traditional rotor system is that the size is defined by reserving strength margin according to engineering experience on the premise that various failure modes meet the service life. The structural life of a helicopter rotor system part is basically checked by adopting a fatigue characteristic test, and the fatigue characteristic test has the problems of asynchronous life of each failure mode and great acceleration difference, so that the structural fatigue test fails. Such test results do not match the expected failure mode, resulting in wasted test cost and lead time.
Disclosure of Invention
The invention aims to provide a test-oriented fatigue design method for a helicopter rotor wing metal piece, which considers the fatigue failure conditions of different parts and different abrasion states of a part in the fatigue characteristic test process, takes the fatigue test assessment of each part as a target, adopts a test sampling fatigue limit to carry out strength design and guides the structure key size definition, and provides technical support for the development of a helicopter.
In order to realize the task, the invention adopts the following technical scheme:
a fatigue design method for a test-oriented helicopter rotor metal piece comprises the following steps:
establishing a fatigue performance characterization model of a helicopter rotor component based on an accumulated damage theory to obtain a relation between the fatigue limit and the service life of a rotor wing metal piece; determining fatigue limits of different material types and ablation states under the target service life based on the fatigue performance characterization model; carrying out fatigue test verification on the rotor wing metal piece, and determining fatigue test loading amplitudes corresponding to the target service life under different material types and the abrasion corrosion states; determining the maximum loading load of carrying out fatigue tests on different materials and under the condition of the fretting corrosion in the rotor wing metal piece based on the fatigue test loading amplitude, substituting the maximum loading load of each characteristic load into a structural strength analysis model of the metal piece and analyzing according to the test design requirements;
according to the reliability of the load spectrum data, taking a load correction coefficient to carry out fatigue limit calculation on the basis of the maximum load;
and selecting proper static load according to the rotor load spectrum, carrying out static load correction on the proper static load to obtain equivalent dynamic load, comparing the equivalent dynamic load with the safe fatigue limit for corresponding metal piece material design, and evaluating whether the key size of the metal piece meets the requirement.
Further, the determining fatigue limits for different material types and ablation states at a target life based on the fatigue performance characterization model comprises:
selecting the material type and the ablation state of the rotor wing part according to actual conditions, and respectively drawing fatigue life curves of different material types and the ablation state according to the fatigue performance characterization model and the load spectrum of the rotor wing part; and determining the fatigue limit of different material types and the friction corrosion state under the target service life through the fatigue life curve, namely determining the target service life fatigue limit of the component.
Further, the fatigue performance characterization model is represented as:
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material and are determined according to test data; t is a,i (i ═ 1,2, …, p) is a set of dynamic loads for characteristic load T in each flight regime; n is i (i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is ∞ The high cycle fatigue limit of the characteristic load T corresponding to the life L.
Further, the fatigue test loading amplitudes for the different material types and the fretting regime corresponding to the target life are determined by:
in the formula, C i,j And m i,j (i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the metal piece material corresponding to different materials and in the state of abrasion corrosion, and the test frequency is N 1 ，(T a ) i,j For different material types and for different states of erosion corresponding to a target lifetime L target Loading amplitude for fatigue testing.
Further, the proof fatigue limit of each material and the fretting corrosion state of the rotor component obtained by the fatigue test examination can be expressed as follows:
wherein eta is a load correction coefficient and takes a value of 1.0-1.2.
Further, the suitable static load is a maximum static load or a static load corresponding to a dynamic load state.
Further, the equivalent dynamic load T is obtained Sizing Load Safe fatigue limit S for corresponding metal part material design ∞ And comparing, and evaluating whether the critical dimension of the metal part meets the requirement, wherein the specific method comprises the following steps:
i) if S ∞ ≥T Sizing Load The flight load spectrum is shown at the safe fatigue limit S ∞ The cumulative damage caused by the lower layer is less than T Sizing Load Damage caused by the rotor, whereby the high cycle fatigue life of the rotor structure is greater than the target life L corresponding to the proof dynamic load target And can be examined through component tests;
ii) if S ∞ ＜T Sizing Load The flight load spectrum is shown at the safe fatigue limit S ∞ Cumulative damage caused by lower energy at T or above Sizing Load Damage caused by the lower part, the high cycle fatigue life of the rotor structure does not meet the target life L corresponding to the proofing dynamic load target Further optimization is required.
Further, the fatigue design method for the test-oriented helicopter rotor metal piece further comprises the following steps:
and optimally designing the metal piece which does not meet the requirements, acquiring corresponding local maximum stress according to the corresponding part of each equivalent load dynamic load of the metal piece, and determining the optimal size of the dangerous part of the rotor wing component by combining the dangerous section size of the metal piece obtained by measurement before design optimization under the specified load state, loading position, force transmission path and constraint mode.
Further, the optimal size of the rotor component hazard area is expressed as:
in the above formula, σ max, optimization Representing the fatigue maximum stress, σ, expected for the metal part performance max The method comprises the following steps that A, local maximum stress of parts corresponding to dynamic load of each equivalent load of the metal part is obtained through measurement before design optimization, and A is the size of a dangerous section of the metal part obtained through measurement before design optimization;
and guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula, so that the fatigue life target requirement can be met.
The invention has the following technical characteristics:
1. the method is based on the characteristic of fatigue life test and assessment of helicopter rotor system components, and fully considers the problem of inconsistent life acceleration conditions of various loads and different failure modes under the same test. The method takes the test as the guide, avoids the conditions of insufficient structure assessment caused by different loads and failure mode test service life acceleration difference, inaccurate boundary simulation under the condition of test acceleration load and the like, and improves the efficiency of strength design and design iteration and the success rate of the test by the method for testing the sampling fatigue limit, thereby saving the development period and the test cost.
2. The strength, the rigidity and the boundary constraint of the test piece and the matching part under the characteristic test acceleration load are pre-judged in advance, and early failure caused by unreasonable design of a test scheme is avoided.
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FIG. 1 is a schematic flow diagram of the process of the present invention.
Detailed Description
Referring to fig. 1, the invention provides a fatigue design method for a test-oriented helicopter rotor metal piece, which realizes fatigue life evaluation and structural design of the metal piece, and comprises the following steps:
step one, analyzing fatigue test loads of different metal piece materials in an abrasion corrosion state
Based on the accumulated damage theory, establishing a fatigue performance characterization model of the helicopter rotor component:
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material and are determined according to test data; t is a,i (i ═ 1,2, …, p) is a set of dynamic loads for characteristic load T in each flight regime; n is i (i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is a unit of ∞ The high cycle fatigue limit of the characteristic load T corresponding to the life L. Formula (1) establishes fatigue limit T of rotor wing metal piece ∞ With respect to the life L, a set of fatigue limits T is input ∞ The corresponding fatigue life L can be obtained.
Selecting the material type M of the rotor component according to actual conditions i (i-1, 2, …, k) and an ablation status of F j (j ═ 1,2, …, l), the fatigue life T under different material types, rub-out conditions can be plotted according to equation (1) and rotor component load spectrum, respectively ∞ -an L-curve; through fatigue life T ∞ L-curve, from which a target life L can be determined target Fatigue limit (T) for different material types and states of fretting ∞ ) i,j (i-1, 2, …, k; j-1, 2, …, l), is the target life fatigue limit for the component.
Carrying out fatigue test verification on the rotor wing metal piece, wherein the test frequency is N 1 The different material types and the state of the ablation correspond to the target life L target Fatigue test load amplitude (T) a ) i,j Can be determined by the following formula:
in the formula, C i,j And m i,j And (i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the corresponding metal piece material in different materials and in an abrasion state. According to the formula (2), gyroplanMaximum loading load T for carrying out fatigue test under different materials and friction and corrosion states in accessory a,max Can be expressed as:
T a,max ＝max[(T a ) 1,1 ,(T a ) 2,1 ,…,(T a ) c,1 ,(T a ) 1,2 ,…(T a ) c,1 ,…(T a ) c,d ] (3)
loading the maximum load T of each characteristic load a,max And substituting the structural strength analysis model of the metal piece into the model to analyze the test design requirements such as strength, rigidity, boundary conditions and the like. The structural strength analysis model is a known model.
Step two, calculating the fatigue limit of the test guide proofing under different material and friction and corrosion states
The formula (3) shows the maximum load T of the fatigue test under different rotor wing metal piece materials and wiping corrosion states a,max Considering the change of the structure adjustment to the load in the design process, and according to the reliability of the load spectrum data at T a,max On the basis, the load correction coefficient eta (1.0-1.2) is taken to carry out fatigue limit calculation.
Combining formula (2) and formula (3), fatigue test examination of each material of the rotor part and proof fatigue limit (Sizing Load) under the state of abrasion i,j Can be expressed as:
the fatigue limit Sizing Load calculated by the formula (4) is a proof fatigue Load corresponding to the target service life and guided by the test, and is used for strength design and guiding structure optimization.
Step three, fatigue life assessment and design optimization
Obtaining the test-oriented proofing load fatigue of different metal piece materials in the friction and corrosion state through a formula (4), selecting proper static load according to a rotor wing load spectrum, and carrying out static load correction on the proper static load to obtain an equivalent dynamic load T Sizing Load (ii) a The appropriate static load is generally the maximum static load or the corresponding dynamic load stateThe static load of (2); the method is determined according to actual needs.
Equivalent dynamic load T Sizing Load Safe fatigue limit S for corresponding metal part material design ∞ And comparing, and evaluating whether the critical dimension of the metal piece meets the requirement, wherein the specific method comprises the following steps:
i) if S ∞ ≥T Sizing Load The flight load spectrum is shown at the safe fatigue limit S ∞ The cumulative damage caused by Sizing Load Damage caused by the rotor, whereby the high cycle fatigue life of the rotor structure is greater than the target life L corresponding to the proof dynamic load target And can be examined through component tests;
ii) if S ∞ ＜T Sizing Load The flight load spectrum is shown at the safe fatigue limit S ∞ Cumulative damage caused by Sizing Load Damage caused by the lower part, the high cycle fatigue life of the rotor structure does not meet the target life L corresponding to the proofing dynamic load target Further optimization is needed, and the optimization design process specifically comprises the following steps:
dynamic load T for each equivalent load of the metal part Sizing Load Corresponding part, obtaining corresponding local maximum stress sigma max Under the appointed load state, loading position, power transmission route and restraint mode, combine the dangerous cross-section size A of metalwork that obtains through the measurement before the design optimization, can confirm the optimum size of rotor part danger position:
in the above formula, σ max, optimization A fatigue maximum stress representing a desired metal part performance; and (4) guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula (5), so that the fatigue life target requirement can be met.
Importantly, the design method taking the test as the guide avoids the condition that after the local part or part of the structure meets the requirement and is firstly damaged, the rest parts or parts are not examined, and saves the test cost and the development period.
The above embodiments are only used for illustrating the technical solutions of the present application, and not for limiting the same; although the present application has been described in detail with reference to the foregoing embodiments, it should be understood by those of ordinary skill in the art that: the technical solutions described in the foregoing embodiments may still be modified, or some technical features may be equally replaced; such modifications and substitutions do not substantially depart from the spirit and scope of the embodiments of the present application, and are intended to be included within the scope of the present application.
Claims (9)
1. A fatigue design method for a test-oriented helicopter rotor metal part is characterized by comprising the following steps:
establishing a fatigue performance characterization model of a helicopter rotor component based on an accumulated damage theory to obtain a relation between the fatigue limit and the service life of a rotor wing metal piece; determining fatigue limits of different material types and ablation states under the target service life based on the fatigue performance characterization model; carrying out fatigue test verification on the rotor wing metal piece, and determining fatigue test loading amplitudes corresponding to the target service life under different material types and the friction and corrosion states; determining the maximum loading load of carrying out fatigue tests on different materials and under the condition of the fretting corrosion in the rotor wing metal piece based on the fatigue test loading amplitude, substituting the maximum loading load of each characteristic load into a structural strength analysis model of the metal piece and analyzing according to the test design requirements;
according to the reliability of the load spectrum data, taking a load correction coefficient on the basis of the maximum loading load to carry out fatigue limit calculation;
and selecting proper static load according to the rotor load spectrum, carrying out static load correction on the proper static load to obtain equivalent dynamic load, comparing the equivalent dynamic load with the safe fatigue limit for corresponding metal piece material design, and evaluating whether the key size of the metal piece meets the requirement.
2. A test-oriented helicopter rotor metal part fatigue design method according to claim 1 wherein said determining fatigue limits for different material types and rub-out conditions at a target life based on said fatigue performance characterization model comprises:
selecting the material type and the ablation state of the rotor wing part according to actual conditions, and respectively drawing fatigue life curves of different material types and the ablation state according to the fatigue performance characterization model and the load spectrum of the rotor wing part; and determining the fatigue limit of different material types and the friction corrosion state under the target service life through the fatigue life curve, namely determining the target service life fatigue limit of the component.
3. A test-oriented helicopter rotor metal part fatigue design method according to claim 1, characterized in that said fatigue performance characterization model is represented as:
in the formula, C and m are the slope and intercept parameters of the fatigue curve of the metal piece material; t is a,i (i ═ 1,2, …, p) is the dynamic load of the characteristic load T in each flight condition; n is i (i ═ 1,2, …, p) is the frequency of hourly occurrences of the various flight states; l is the high cycle fatigue life of the helicopter structure; t is a unit of ∞ The high cycle fatigue limit of the characteristic load T corresponding to the life L.
4. A test-guided helicopter rotor metal hardware fatigue design method according to claim 1 wherein said fatigue test loading amplitudes for different material types and erosion states corresponding to a target life are determined by the following equation:
in the formula, C i,j And m i,j (i 1,2, …, c, j 1,2, …, d) are different materials, and the corresponding metal piece material fatigue under the state of abrasion corrosionSlope, intercept parameters of the curve, and the number of tests N 1 ，(T a ) i,j For different material types and for different states of erosion corresponding to a target lifetime L target Loading amplitude for fatigue testing.
5. A method of test-oriented helicopter rotor metal fatigue design according to claim 1, wherein the proof fatigue limit for each material and fretting condition of the rotor component assessed by a fatigue test is expressed as:
wherein, C i,j And m i,j (i is 1,2, …, c, j is 1,2, …, d) is the slope and intercept parameters of the fatigue curve of the metal piece material corresponding to different materials and in the state of abrasion corrosion, and the test frequency is N 1 ，T a,max The maximum load of the fatigue test under different rotor wing metal piece materials and the friction and corrosion states is shown, eta is a load correction coefficient, and the value is 1.0-1.2.
6. A trial-oriented helicopter rotor metal fatigue design method according to claim 1 wherein said suitable static load is the maximum static load or a static load corresponding to a dynamic loading condition.
7. The trial-oriented helicopter rotor metal part fatigue design method of claim 1, characterized by the equivalent dynamic load T SizingLoad Safe fatigue limit S for corresponding metal part material design ∞ And comparing, and evaluating whether the critical dimension of the metal piece meets the requirement, wherein the specific method comprises the following steps:
i) if S ∞ ≥T SizingLoad The flight load spectrum is shown at the safe fatigue limit S ∞ The cumulative damage caused by the lower layer is less than T SizingLoad Damage caused by the rotor structure, whereby the high cycle fatigue life of the rotor structure is greater than the target life corresponding to the dynamic load of the proofL target And can be examined through component tests;
ii) if S ∞ ＜T SizingLoad The flight load spectrum is shown at the safe fatigue limit S ∞ Cumulative damage caused by SizingLoad Damage caused by the lower rotor wing, the high cycle fatigue life of the rotor wing structure does not meet the target life L corresponding to the proofing dynamic load target Further optimization is required.
8. A method of trial-guided helicopter rotor metal part fatigue design according to claim 1, further comprising:
and optimally designing the metal piece which does not meet the requirements, acquiring corresponding local maximum stress according to the corresponding part of each equivalent load dynamic load of the metal piece, and determining the optimal size of the dangerous part of the rotor wing component by combining the dangerous section size of the metal piece obtained by measurement before design optimization under the specified load state, loading position, force transmission path and constraint mode.
9. A trial oriented helicopter rotor metal hardware fatigue design method according to claim 8, wherein the optimal dimensions for the rotor component hazard zone are expressed as:
in the above formula, σ max, optimization Maximum stress, σ, of fatigue representing the desired performance of the metal part max A is the dangerous section size of the metal piece obtained by measurement before design optimization for the local maximum stress of the corresponding part of each equivalent load dynamic load of the metal piece;
and guiding the design optimization of the dangerous section of the structure according to the calculation result of the formula, so that the fatigue life target requirement can be met.
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