Method for determining a durability design allowable stress for an aircraft metal structure

Abstract

This invention belongs to the field of aircraft structural performance evaluation technology, specifically involving a method for determining the allowable stress for durability design of aircraft metal structures. First, the baseline local allowable stress is calculated under a given material and load spectrum using fatigue analysis. Considering actual structural factors, an equivalent stress concentration factor is introduced to correct the baseline local allowable stress. Furthermore, a stress correction factor is introduced based on the correction of theoretical analysis results using fatigue tests. Considering reliability requirements, a fatigue uncertainty factor is also introduced, thereby establishing a set of analogous calculation formulas to determine the allowable stress for durability design. The calculation formula is: where, [σ max,0 [K] is the reference local allowable stress. te C is the equivalent stress concentration factor, SCF is the stress correction factor, FSF is the fatigue uncertainty factor, and C is the stress concentration factor. load This is the load conversion factor. This invention simplifies the fatigue test matrix while ensuring the accuracy of theoretical analysis; it can quickly determine whether the structural fatigue margin meets the requirements, thus enabling durability analysis to cover the entire load-bearing structure of the aircraft.

Description

Technical Field

[0001] This invention belongs to the field of aircraft structural performance evaluation technology, specifically relating to the field of aircraft structural durability analysis, and particularly to a method for determining the allowable stress for durability design of aircraft metal structures. Background Technology

[0002] In the process of aircraft design and development, in order to quickly and quantitatively assess whether the durability of the entire load-bearing structure meets the design requirements, a commonly used assessment method is the durability design allowable stress method. This method provides the durability allowable stress of various materials and details of the aircraft structure under the design load spectrum and design life index. By comparing the allowable stress of the structural details with the service stress (or working stress), it is determined whether the structural fatigue margin (or durability margin) meets the requirements, thereby achieving durability analysis covering the entire load-bearing structure of the aircraft.

[0003] This method is for metallic structures, and the formula for calculating fatigue margin (FMS) is as follows:

[0004]

[0005] In the formula, [σ ref ] represents the allowable stress, σ ref The value indicates the working stress, and the subscript ref indicates the reference stress or nominal stress.

[0006] The allowable stress for durability is the stress value that a target structure can withstand under expected service conditions when its durability life, obtained through durability analysis or testing, just meets the design service life requirement. It depends on multiple factors, including the design load spectrum, design life, structural details, materials, and structural reliability requirements. Among these, the most important influencing factors are materials, load spectrum, and design life index.

[0007] The allowable stress for durability can be theoretically calculated based on fatigue analysis methods (stress fatigue method or strain fatigue method), or obtained based on durability tests under random spectrum. Given the load spectrum sequence, material property parameters, detailed characteristics, and dimensions, the stress level corresponding to the maximum load in the spectrum is calculated based on the design life index and reliability requirements. This stress is then converted to the limiting load or ultimate load as needed to obtain the allowable stress. The allowable stress can correspond to either the limiting load or the ultimate load; the two are converted using the load uncertainty factor (also known as the safety factor). The allowable stress value under the ultimate load is equal to the allowable stress value under the limiting load multiplied by the load uncertainty factor. The allowable stress in this invention refers to the limiting load.

[0008] Since numerous factors influence the allowable stress for durability design, relying solely on fatigue tests to obtain the allowable stress would result in an excessively large test matrix; conversely, relying solely on theoretical analysis would compromise the accuracy of the calculations. Against this backdrop, this invention proposes a method for determining the allowable stress for durability design of metal structures, primarily based on fatigue analysis supplemented by experimental corrections. Summary of the Invention

[0009] The technical problem solved by this invention is: a method for determining the allowable stress for durability design of metal structures, which mainly uses fatigue analysis and is supplemented by experimental correction. It is used to determine the allowable stress for durability of various materials and details of aircraft structures under design load spectrum and design life index, so as to conduct rapid durability assessment by judging the positive or negative value of fatigue margin.

[0010] Technical Solution: A method for determining the allowable stress for durability design of aircraft metal structures. First, the baseline local allowable stress under given material and load spectrum is calculated using fatigue analysis. Considering actual structural factors, an equivalent stress concentration factor is introduced to correct the baseline local allowable stress. Then, a stress correction factor is introduced based on the correction of theoretical analysis results by fatigue tests. Finally, a fatigue uncertainty factor is introduced to consider reliability requirements, thereby establishing a set of analogous calculation formulas to determine the allowable stress for durability design.

[0011] The formula for calculating the allowable stress in durability design is as follows:

[0012]

[0013] Among them, [σ max,0 [K] is the reference local allowable stress. te C is the equivalent stress concentration factor, SCF is the stress correction factor, FSF is the fatigue uncertainty factor, and C is the stress concentration factor. load This is the load conversion factor.

[0014] Furthermore, for the target structure, the materials involved, fatigue load spectrum type, detail type, and design life are determined;

[0015] Using fatigue analysis methods, based on the aforementioned material and load spectrum, the theoretical local elastic stress corresponding to the fatigue life equaling the design life is calculated, and the reference local allowable stress [σ] is determined. max,0 ];

[0016] For the aforementioned detail type, the theoretical stress concentration factor K is calculated based on the actual structural dimensions. t Considering the effects of surface treatment, surface roughness, cold work strengthening, and fastener installation factors, K is introduced. t Correction factor, used to calculate the equivalent stress concentration factor Kte;

[0017] For the aforementioned detail types and materials, based on fatigue test data of similar detail simulation test specimens under the same material and similar load spectrum, the test stress is compared with the theoretical stress derived from the fatigue analysis method, the stress correction factor SCF is calculated, and statistical aggregation is performed according to material and detail type;

[0018] For the aforementioned load spectrum type, the load reduction factor C is determined based on the ratio of the limiting load to the maximum load in the spectrum. load ;

[0019] For the target structure, the fatigue uncertainty factor (FSF) is determined based on factors such as fatigue life dispersion coefficient, structural importance, material property differences, and load spectrum differences.

[0020] Furthermore, the calculation method for the reference local allowable stress is as follows:

[0021] 4) Given the material and load spectrum sequence, select several local elastic stress levels σ max The theoretical fatigue life N is calculated using fatigue analysis methods, where σ max There are at least three stress levels: low, medium, and high, to ensure that N is sufficiently dispersed and located between 2000FH and 20000FH, where FH represents flight hours.

[0022] 5) Based on several groups (σ) max For sample points N, the local elastic stress-fatigue life curve is fitted using a power function formula, i.e., σ max The -N curve is shown below:

[0023]

[0024] In the formula, m and C are constants that are related to the fatigue analysis method, material, and load spectrum.

[0025] 6) According to σ max -N curve, inversely calculate the reference local allowable stress [σ] corresponding to the design life N0. max,0 The calculation formula is as follows, where [σ] max,0 [Corresponds to the maximum load condition in the fatigue spectrum;]

[0026] [σ max,0 ]=(C / N0) 1 / m .

[0027] Furthermore, the theoretical stress concentration factor is calculated as follows: for different detail types, consult the curves or calculation formulas in the stress concentration factor handbook, or obtain the ratio of local elastic stress to nominal stress through detailed finite element analysis; the equivalent stress concentration factor is calculated as follows:

[0028] K te =Kt / ΠC i

[0029] Among them, C i The Kt correction coefficient, representing various influencing factors, is strongly correlated with the material type and needs to be obtained through experiments or by using empirical values.

[0030] C st This indicates a surface treatment correction factor, such as chromic anodizing.

[0031] C cw This indicates the correction factor for cold work strengthening, such as whether cold extrusion strengthening is used;

[0032] C sr This represents a surface roughness correction factor, such as Ra = 1.6, 3.2, 6.4;

[0033] C fi This indicates the fastener installation correction factor, such as for rivets, pull studs, interference fit high lock, clearance fit high lock, etc.

[0034] C se This represents the size effect correction factor.

[0035] Furthermore, C cw When the cold work strengthening correction factor is obtained through experiments, firstly, test specimens with and without cold work strengthening are designed. Then, nominal stress-life curves are constructed through fatigue tests under multiple stress levels. Finally, the corresponding stress is calculated based on the target life. The ratio of the two stresses is the correction factor, as shown in the following formula.

[0036] C cw =S cw / S0.

[0037] Furthermore, the formula for calculating the stress correction factor SCF is as follows:

[0038]

[0039] Among them, [σ max '] refers to the theoretical local elastic stress corresponding to the median life of the test under the test load spectrum and the test specimen material, which is back-calculated using fatigue analysis methods. The fatigue analysis method must be the same as the method used to determine the reference local allowable stress, and N0 is replaced with the median life of the test. It refers to the crack initiation life of detectable engineering cracks; K te 'Refers to K of the test specimen' te ;σ ref 'Refers to the nominal stress of the test.'

[0040] SCF reflects the degree of matching between fatigue analysis and testing. The closer the SCF is to 1.0, the better the matching is. If SCF > 1.0, the fatigue analysis life is longer than the test life, and the analysis results are more dangerous.

[0041] This coefficient is mainly related to the detail type, material, and load spectrum type. It needs to be calculated and statistically / merged based on a large number of fatigue tests of typical and simulated parts under random spectrum to verify and correct the theoretical analysis results, thereby improving the accuracy and reliability of the theoretical calculation results.

[0042] Furthermore, the load reduction factor C load The calculation formula is:

[0043] C load =F lim / F spec

[0044] Among them, F lim Indicates the limiting load, F spec This represents the maximum load in the spectrum. This coefficient is mainly related to the type of load spectrum.

[0045] Furthermore, the formula for calculating the fatigue uncertainty factor FSF is as follows:

[0046] FSF = Πβ i

[0047] β0 represents the baseline dispersion coefficient, which is derived by inversely from the fatigue life dispersion coefficient. The calculation formula is as follows:

[0048] β0=LSF 1 / m

[0049] Where LSF represents the fatigue life dispersion factor, and m represents the local elastic stress-life curve under a given material and load spectrum, i.e., σ max -N curve parameters.

[0050] β r This represents the importance coefficient of a component, which is related to the component's impact on aircraft safety and economy. Critical components have the highest importance, followed by important components, and general components have the lowest. For example, critical components = 1.0, important components = 0.9, and general components = 0.8.

[0051] β m This represents the coefficient of variation in material properties. When fatigue performance parameters for a current material grade are unavailable, and only similar materials can be used (e.g., borrowing performance parameters from materials of the same grade but with different heat treatments or different dimensions), β... m ≥1.0;

[0052] β sβ represents the load spectrum difference coefficient, used when there is a significant difference between the load spectrum used in the analysis and the actual load spectrum. s ≥1.0.

[0053] Furthermore, load spectrum type refers to component load spectrum that is directly related to the stress spectrum of structural details, including wing bending moment spectrum, fuselage longitudinal bending moment spectrum, control surface hinge moment spectrum, landing gear load spectrum, etc.; detail types include several categories such as fastening holes, hollow holes and notches, fillets, lugs, etc.

[0054] Furthermore, the fatigue analysis methods refer to stress fatigue method and strain fatigue method; among which, stress fatigue method includes nominal stress method and stress severity coefficient method.

[0055] Beneficial technical effects: This invention proposes a method for determining the allowable stress for durability design of metal structures. It mainly uses fatigue analysis and is supplemented by experimental correction. By analogy with the calculation formula, it simplifies the fatigue test matrix and ensures the accuracy of theoretical analysis. By comparing the allowable stress and the working stress, it can quickly determine whether the fatigue margin of the structure meets the requirements, thereby achieving durability analysis covering the entire load-bearing structure of the machine. Attached Figure Description

[0056] To more clearly illustrate the technical solutions of the embodiments of the present invention, the drawings used in the embodiments of the present invention will be briefly introduced below. Obviously, the drawings described below are only some embodiments of the present invention. For those skilled in the art, other drawings can be obtained based on these drawings without creative effort.

[0057] Figure 1 This is the technical process of the present invention;

[0058] Figure 2 The nominal stress-life curve;

[0059] Figure 3 This is the wing bending moment spectrum sequence in the example (with the ordinate axis normalized).

[0060] Figure 4 This is the local elastic stress-fatigue life curve (σmax-N curve) in the example. Detailed Implementation

[0061] To make the objectives, technical solutions, and advantages of the embodiments of the present invention clearer, the technical solutions of the embodiments of the present invention will be clearly and completely described below with reference to the accompanying drawings. Obviously, the described embodiments are only some embodiments of the present invention, not all embodiments. Based on the embodiments of the present invention, all other embodiments obtained by those skilled in the art without creative effort are within the scope of protection of the present invention.

[0062] The features and illustrative embodiments of various aspects of the present invention will now be described in detail. Numerous specific details are set forth in the following detailed description to provide a thorough understanding of the invention. However, it will be apparent to those skilled in the art that the invention may be practiced without requiring some of these specific details. The following description of embodiments is merely intended to provide a better understanding of the invention by illustrating examples of the invention. The invention is by no means limited to any specific setups and methods set forth below, but covers any improvements, substitutions, and modifications to structures, methods, and devices without departing from the spirit of the invention. Well-known structures and techniques are not shown in the drawings and the following description to avoid unnecessarily obscuring the invention.

[0063] It should be noted that, unless otherwise specified, the embodiments of the present invention and the features thereof can be combined with each other, and the various embodiments can be referenced and cited in each other. The present invention will now be described in detail with reference to the accompanying drawings and embodiments.

[0064] See Figure 1 The present invention first calculates the baseline local allowable stress under a given material and load spectrum using fatigue analysis. Considering the geometric characteristics, surface treatment, surface roughness, cold work strengthening, and other factors of the actual structure, an equivalent stress concentration factor is introduced to correct the baseline local allowable stress. Furthermore, a stress correction factor is introduced based on the correction of theoretical analysis results by fatigue tests. In addition, considering reliability requirements, a fatigue uncertainty factor is introduced, thereby establishing a set of analogous calculation formulas to determine the allowable stress for durability design.

[0065] The formula for calculating the allowable stress in durability design is as follows:

[0066]

[0067] Among them, [σ max,0 [K] is the reference local allowable stress. te C is the equivalent stress concentration factor, SCF is the stress correction factor, FSF is the fatigue uncertainty factor, and C is the stress concentration factor. load This is the load conversion factor.

[0068] The calculation steps are as follows:

[0069] S1: For the target structure, determine the material grade, fatigue load spectrum type, detail type, and design life involved;

[0070] In step S1, the load spectrum type refers to the component load spectrum that is directly related to the structural detail stress spectrum, including the wing bending moment spectrum, the fuselage longitudinal bending moment spectrum, the control surface hinge moment spectrum, the landing gear load spectrum, etc.; the detail types include several categories such as fastening holes, hollow holes and notches, fillets, lugs, etc.

[0071] S2: Using fatigue analysis methods, based on the material and load spectrum described in S1, calculate the theoretical local elastic stress corresponding to a fatigue life equal to the design life, and determine the reference local allowable stress [σ]. max,0 ];

[0072] In step S2, the fatigue analysis method refers to the stress fatigue method (including the nominal stress method and the stress severity coefficient method) and the strain fatigue method (also known as the local stress-strain method).

[0073] In step S2, the method for calculating the reference local allowable stress is as follows:

[0074] 1) Given the material and load spectrum sequence, select several (at least 3) local elastic stress levels σ max The theoretical fatigue life N is calculated using fatigue analysis methods, where σ max There are at least three stress levels: low, medium, and high, to ensure that N is sufficiently dispersed and located between 2000FH and 20000FH, where FH represents flight hours.

[0075] 2) Based on several groups (σ) max For sample points (N), the local elastic stress-fatigue life curve (σ) is fitted using a power function formula. max -N curve), as follows:

[0076]

[0077] In the formula, m and C are constants that are related to the fatigue analysis method, material, and load spectrum.

[0078] 3) According to σ max -N curve, inversely calculate the reference local allowable stress [σ] corresponding to the design life N0. max,0 The calculation formula is as follows, where [σ] max,0 This corresponds to the maximum load condition in the fatigue spectrum.

[0079] [σ max,0 ]=(C / N0) 1 / m

[0080] S3: For the detail type described in S1, calculate the theoretical stress concentration factor K based on the actual structural dimensions. t Considering the effects of surface treatment, surface roughness, cold work strengthening, and fastener installation, a Kt correction factor is introduced to calculate the equivalent stress concentration factor K. te ;

[0081] In step S3, the theoretical stress concentration factor is calculated as follows: for different detail types, consult the curves or formulas in the stress concentration factor handbook, or obtain the ratio of local elastic stress to nominal stress (or reference stress) through detailed finite element analysis.

[0082] In step S3, the equivalent stress concentration factor is calculated as follows:

[0083] K te =K t / ΠC i

[0084] Among them, C i The Kt correction coefficient represents various influencing factors and is strongly correlated with the material type. It needs to be obtained through experiments. Currently, certain empirical values ​​and formulas have been accumulated in engineering, which can be obtained by consulting relevant manuals.

[0085] C st This indicates a surface treatment correction factor, such as chromic anodizing.

[0086] C cw This indicates the correction factor for cold work strengthening, such as whether cold extrusion strengthening is used;

[0087] C sr This represents a surface roughness correction factor, such as Ra = 1.6, 3.2, 6.4;

[0088] C fi This indicates the fastener installation correction factor, such as for rivets, pull studs, interference fit high lock, clearance fit high lock, etc.

[0089] C se This represents the size effect correction factor.

[0090] Taking the cold work strengthening correction factor as an example, the method for obtaining this factor through experiments is as follows; the methods for obtaining other correction factors are similar. First, design test specimens with and without cold work strengthening. Construct nominal stress-life curves through fatigue tests at multiple stress levels. Then, calculate the corresponding stress based on the target life, such as... Figure 2 As shown, the ratio of the two stresses is the correction factor, and the formula is as follows.

[0091] C cw =S cw / S0

[0092] S4: For the detail types and materials mentioned in S1, based on the fatigue test data of similar detail simulation test pieces under the same material and similar load spectrum, compare the test stress with the theoretical stress derived from the fatigue analysis method, calculate the stress correction factor SCF, and perform statistical merging according to material and detail type;

[0093] In step S4, the formula for calculating the stress correction factor SCF is:

[0094]

[0095] Among them, [σ max '] refers to the theoretical local elastic stress corresponding to the median life of the test under the test load spectrum and test specimen material by back-calculation using fatigue analysis methods. The calculation method is described in step S2. The fatigue analysis method must be the same as the method used to determine the reference local allowable stress, and N0 should be replaced with the median life of the test, which refers to the crack initiation life of detectable engineering cracks; K te 'Refers to K of the test specimen' te ;σ ref 'Refers to the nominal stress of the test.'

[0096] SCF reflects the degree of matching between fatigue analysis and testing. The closer the SCF is to 1.0, the better the matching is. If SCF > 1.0, the fatigue analysis life is longer than the test life, and the analysis results are more dangerous.

[0097] This coefficient is mainly related to the detail type, material, and load spectrum type. It needs to be calculated and statistically / merged based on a large number of fatigue tests of typical and simulated parts under random spectrum to verify and correct the theoretical analysis results, thereby improving the accuracy and reliability of the theoretical calculation results.

[0098] S5: For the load spectrum type described in S1, determine the load reduction factor C based on the ratio of the limiting load to the maximum load in the spectrum. load ;

[0099] In step S5, the load reduction factor C load The calculation formula is:

[0100] C load =F lim / F spec

[0101] Among them, F lim Indicates the limiting load, F spec This represents the maximum load in the spectrum. This coefficient is mainly related to the type of load spectrum. Note that static load conditions with extremely low probability of occurrence, such as failure conditions, should be excluded when determining the limiting load.

[0102] S6: For the target structure described in S1, determine the fatigue uncertainty factor FSF based on factors such as fatigue life dispersion coefficient, structural importance, material property differences, and load spectrum differences;

[0103] In step S6, the fatigue uncertainty factor FSF is calculated using the following formula:

[0104] FSF = Πβi

[0105] β0 represents the baseline dispersion coefficient, which is derived by inversely from the fatigue life dispersion coefficient. The calculation formula is as follows:

[0106] β0=LSF 1 / m

[0107] Where LSF represents the fatigue life dispersion factor, and m represents the local elastic stress-life curve (σ) under a given material and load spectrum. max -N curve parameters.

[0108] β r This represents the importance coefficient of a component, which is related to the component's impact on aircraft safety and economy. Critical components have the highest importance, followed by important components, and general components have the lowest. For example, critical components = 1.0, important components = 0.9, and general components = 0.8.

[0109] β m This represents the coefficient of variation in material properties. When fatigue performance parameters for a current material grade are unavailable, and only similar materials can be used (e.g., borrowing performance parameters from materials of the same grade but with different heat treatments or different dimensions), β... m ≥1.0;

[0110] β s β represents the load spectrum difference coefficient, used when there is a significant difference between the load spectrum used in the analysis and the actual load spectrum. s ≥1.0.

[0111] S7: Substitute the parameters determined in steps S2 to S6 into the following calculation formula to determine the allowable stress [σ] for durability design of the actual structural details. ref ].

[0112]

[0113] It should be noted that the above calculation of the damage tolerance design allowable stress [σ] ref In the process of determining each parameter, there is no strict requirement for the order of execution; the calculation process can be performed in parallel during the computer program's operation.

[0114] Example 1: Implemented using the design described above, the technical process of the present invention is as follows: Figure 1 Taking a certain fuselage reinforcement frame of a certain type of aircraft as an example, the process of determining the allowable stress for durability design is explained.

[0115] 1) Determine the material grade, fatigue load spectrum type, detail type, and design life involved in this box;

[0116] The frame is made of 7050-T7451 thick plate and is mainly affected by the wing bending moment spectrum (spectral sequence as follows). Figure 2 The key components include four categories: fastening holes, hollow holes, rounded corners, and lugs, with a design life of 8000 FH.

[0117] 2) Using the local stress-strain method, calculate the theoretical local elastic stress corresponding to a fatigue life of 8000FH for the 7050-T7451 thick plate under the wing bending moment spectrum, and determine [σ]. max,0 ];

[0118] In the local stress-strain method, the Neuber method and the Morrow mean stress correction method are used to calculate four sets of (σ) max Sample points (N) were used, and the local elastic stress-fatigue life curve (σ) was fitted using a power function formula. max -N curve), such as Figure 3 As shown, m = 5.807, C = 9.467E+19. Then, the allowable local stress corresponding to 8000FH is calculated, [σ max,0 =586 MPa.

[0119] 3) For the four types of details, based on the actual structural dimensions, the theoretical Kt is calculated by consulting the stress concentration factor handbook. Then, considering various Kt correction factors and using values ​​accumulated in engineering, the equivalent Kt is further calculated. te ;

[0120] Equivalent K te The calculation process is as follows:

[0121] Table 1 Equivalent Kte for various details of 7050-T7451

[0122] Fastening hole empty hole Rounded corners ear <![CDATA[Theory K t > 3.2 2.5 2.0 2.0 <![CDATA[Surface treatment C st > 0.8 0.8 0.8 0.8 <![CDATA[Cold work strengthening C cw > 1.0 / / 1.20 <![CDATA[Surface roughness C sr > 1.0 1.0 1.0 1.0 <![CDATA[Fastener installation C fi > 1.20 / / / <![CDATA[Size effect C se > 0.98 0.85 0.98 0.85 <![CDATA[Equivalent K te > 3.40 3.68 2.55 2.45

[0123] Note: Kt corresponds to the extrusion stress of the ear hole, and Kt for other parts corresponds to the stress of the gross section.

[0124] 4) Based on existing fatigue test data of various simulated parts under similar load spectra, calculate the stress correction factor SCF and perform statistical aggregation according to material and detail type;

[0125] The SCF merging results are as follows:

[0126] Table 2 SCF details for various aspects of 7050-T7451

[0127] Fastening hole empty hole Rounded corners ear SCF 0.9 1.1 1.1 0.9

[0128] 5) For the wing bending moment spectrum, determine the load reduction factor C based on the ratio of the limiting load to the maximum load in the spectrum. load =1.25;

[0129] 6) Determine the fatigue uncertainty factor (FSF) based on factors related to reliability indicators such as fatigue life dispersion factor and structural importance;

[0130] This structure is a critical fatigue component with a life dispersion factor of 4.0. Therefore, β0 = 4^(1 / 5.807) = 1.27, β r =1.0. Since the fatigue performance parameters of the material are fixed, β m =1.0; the load spectrum sequence is also determined, β s =1.0. Therefore, FSF = β0 * β r *β m *β s =1.27.

[0131] 7) Substitute the parameters determined in the above steps into the allowable stress calculation formula to determine the allowable stress [σ] for durability design of the actual structural details. ref ].

[0132]

[0133] The allowable durability stress (corresponding to the limiting load) of typical structural details of 7050-T7451 thick plate under the wing bending moment spectrum is listed in the table below.

[0134] Table 3 7050-T7451-Aircraft Wing Bending Moment Spectrum-Allowable Durability Stress (MPa)

[0135] Fastening hole empty hole Rounded corners ear <![CDATA[[σ ref ]]]> 188 143 206 262

[0136] Note: The allowable stress of the lug corresponds to the compressive stress of the lug hole, and other allowable stresses correspond to the stress of the gross section.

[0137] Through the above description of the embodiments, those skilled in the art can clearly understand that each embodiment can be implemented by means of software plus necessary hardware platforms, and of course, it can be implemented directly by hardware. Based on this understanding, the above technical solutions, in essence, or the part that contributes to the prior art, can be embodied in software form. This computer software can be stored in a computer-readable storage medium, such as ROM / RAM, magnetic disk, optical disk, etc. The readable storage medium can store programs, instructions, etc., causing a computer device (such as a personal computer, server, or network device, etc.) to execute the methods described in the various embodiments.

[0138] Those skilled in the art will recognize that the various program (functional) units and execution steps described in conjunction with the embodiments disclosed herein can be implemented in electronic hardware, computer software, or a combination of both. To clearly illustrate the interchangeability of hardware and software, whether a unit or step of the above technical solution is executed in hardware or software may depend on the specific application and design constraints of the technical solution. Those skilled in the art can use different methods to implement the various embodiments for each specific application.

[0139] Finally, it should be noted that the above embodiments are only used to illustrate the technical solutions of the present invention, but the protection scope of the present invention is not limited thereto. Any person skilled in the art can easily conceive of various equivalent modifications or substitutions within the technical scope disclosed in the present invention, and these modifications or substitutions should be covered within the protection scope of the present invention.

Claims

1. A method for determining the allowable stress for durability design of aircraft metal structures, characterized in that, First, the reference local allowable stress under given material and load spectrum is calculated by fatigue analysis. Then, considering actual structural factors, an equivalent stress concentration factor is introduced to correct the reference local allowable stress. Furthermore, a stress correction factor is introduced based on the correction of theoretical analysis results by fatigue tests; and a fatigue uncertainty factor is also introduced to take into account reliability requirements, thereby establishing a set of analog calculation formulas to determine the allowable stress for durability design; Durability design allowable stress The calculation formula is as follows: Among them, [σ max,0 [K] is the reference local allowable stress. te C is the equivalent stress concentration factor, SCF is the stress correction factor, FSF is the fatigue uncertainty factor, and C is the stress concentration factor. load This is the load conversion factor; For the target structure, determine the materials involved, fatigue load spectrum type, detail type, and design life; Using fatigue analysis methods, based on the aforementioned material and load spectrum, the theoretical local elastic stress corresponding to the fatigue life equaling the design life is calculated, and the reference local allowable stress [σ] is determined. max,0 ]; For the aforementioned detail type, the theoretical stress concentration factor K is calculated based on the actual structural dimensions. t Considering the effects of surface treatment, surface roughness, cold work strengthening, and fastener installation factors, K is introduced. t Correction factor, used to calculate the equivalent stress concentration factor Kte; For the aforementioned detail types and materials, based on fatigue test data of similar detail simulation test specimens under the same material and similar load spectrum, the test stress is compared with the theoretical stress derived from the fatigue analysis method, the stress correction factor SCF is calculated, and statistical aggregation is performed according to material and detail type; For the aforementioned load spectrum type, the load reduction factor C is determined based on the ratio of the limiting load to the maximum load in the spectrum. load ; For the target structure, the fatigue uncertainty factor (FSF) is determined based on factors such as fatigue life dispersion coefficient, structural importance, material property differences, and load spectrum differences.

2. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, The calculation method for the reference local allowable stress is as follows: 1) Given the material and load spectrum sequence, select several local elastic stress levels σ max The theoretical fatigue life N is calculated using fatigue analysis methods, where σ max There are at least three stress levels: low, medium, and high, to ensure that N is sufficiently dispersed and located between 2000FH and 20000FH, where FH represents flight hours. 2) Based on several groups (σ) max For sample points N, the local elastic stress-fatigue life curve is fitted using a power function formula, i.e., σ max The -N curve is shown below: In the formula, m and C are constants that are related to the fatigue analysis method, material, and load spectrum. 3) According to σ max -N curve, inversely calculate the reference local allowable stress [σ] corresponding to the design life N0. max,0 The calculation formula is as follows, where [σ] max,0 [Corresponds to the maximum load condition in the fatigue spectrum;] 。 3. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, The theoretical stress concentration factor is calculated as follows: for different detail types, consult the curves or calculation formulas in the stress concentration factor handbook, or obtain the ratio of local elastic stress to nominal stress through detailed finite element analysis; equivalent stress concentration factor. The calculation method is as follows: Among them, C i The Kt correction coefficient, representing various influencing factors, is strongly correlated with the material type and needs to be obtained through experiments or by using empirical values. C st Indicates the surface treatment correction factor; C cw This indicates the correction factor for cold working intensification; C sr This represents the surface roughness correction factor; C fi This indicates the fastener installation correction factor; C se This represents the size effect correction factor.

4. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 3, characterized in that, C cw When the cold work strengthening correction factor is obtained through experiments, firstly, test specimens with and without cold work strengthening are designed. Then, nominal stress-life curves are constructed through fatigue tests under multiple stress levels. Finally, the corresponding stress is calculated based on the target life. The ratio of the two stresses is the correction factor, as shown in the following formula. C cw = S cw / S0。 5. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, The formula for calculating the stress correction factor SCF is: Among them, [σ max '] refers to the theoretical local elastic stress corresponding to the median life of the test under the test load spectrum and the test specimen material, which is back-calculated using fatigue analysis methods. The fatigue analysis method must be the same as the method used to determine the reference local allowable stress, and N0 is replaced with the median life of the test. It refers to the crack initiation life of detectable engineering cracks; K te 'Refers to K of the test specimen' te ;σ ref 'Refers to the nominal stress of the test.' 6. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, Load conversion factor C load The calculation formula is: Among them, F lim Indicates the limiting load, F spec This represents the maximum load in the spectrum; this coefficient is mainly related to the type of load spectrum.

7. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, The formula for calculating the fatigue uncertainty factor (FSF) is: β0 represents the baseline dispersion coefficient, which is derived by inversely from the fatigue life dispersion coefficient. The calculation formula is as follows: Where LSF represents the fatigue life dispersion factor, and m represents the local elastic stress-life curve under a given material and load spectrum, i.e., σ max -N curve parameters; β r The importance coefficient represents the component's impact on aircraft safety and economy. Critical components have the highest importance, followed by important components, and general components have the lowest importance. β m This represents the coefficient of variation in material properties. Since fatigue performance parameters for the current material grade are currently unavailable, we can only borrow from similar materials. (β) m ≥1.0; β s β represents the load spectrum difference coefficient, used when there is a significant difference between the load spectrum used in the analysis and the actual load spectrum. s ≥1.

0.

8. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 1, characterized in that, Load spectrum type refers to the component load spectrum that is directly related to the stress spectrum of structural details, including wing bending moment spectrum, fuselage longitudinal bending moment spectrum, control surface hinge moment spectrum, and landing gear load spectrum; detail types include fastening holes, holes and notches, fillets, and lugs.

9. The method for determining the allowable stress for durability design of aircraft metal structures as described in claim 8, characterized in that, The fatigue analysis methods mentioned refer to stress fatigue method and strain fatigue method; among them, stress fatigue method includes nominal stress method and stress severity coefficient method.

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