A method for determining a damage tolerance design allowable stress for an aircraft metallic structure

Abstract

This invention belongs to the field of aircraft structural performance evaluation technology, specifically relating to a method for determining the allowable stress for damage tolerance design of aircraft metal structures. It calculates the baseline allowable stress for structural details under given material and load spectra, and introduces a stress correction coefficient based on crack propagation tests to correct theoretical analysis results. Considering the actual structure, a strengthening correction coefficient is introduced; considering reliability requirements, a fatigue uncertainty coefficient is introduced, thereby establishing a set of analogous calculation formulas to determine the allowable stress for damage tolerance design. The calculation formula is: where, [σ ref,0 [ ] represents the reference allowable stress, SCF is the stress correction factor, and C cw To strengthen the correction factor, FSF is the fatigue uncertainty factor, C load This represents the load conversion factor. The method proposed in this invention simplifies the crack propagation test matrix while ensuring the accuracy of theoretical analysis; by comparing allowable stress and working stress, it quickly determines whether the structural damage tolerance margin meets the requirements, thereby achieving damage tolerance analysis covering all critical fracture components of the entire machine.

Description

Technical Field

[0001] This invention belongs to the field of aircraft structural performance evaluation technology, specifically relating to aircraft structural damage tolerance analysis, and particularly a method for determining the allowable stress for damage tolerance 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 damage tolerance of critical fracture components (or flight safety structures) meets the design requirements, a commonly used assessment method is the damage tolerance design allowable stress method. This method provides the damage tolerance allowable stress for various materials and details of the aircraft structure under the design load spectrum and specified service life requirements. By comparing the allowable stress of structural details with the service stress (or working stress), it is determined whether the structural damage tolerance margin meets the requirements, thereby achieving damage tolerance analysis covering all critical fracture components of the entire aircraft.

[0003] This method is for metal structures, and the formula for calculating the damage tolerance margin 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] Damage tolerance allowable stress is the allowable stress value at which the crack propagation life of a target structure under expected service conditions just meets the specified service life, obtained through damage tolerance analysis or testing. It depends on multiple factors, including the design load spectrum, the specified service life, the type of structural details, the materials, and the structural reliability requirements. Among these, the most important influencing factors are the materials, the load spectrum, and the specified service life.

[0007] The specified service life is mainly related to inspectability. For non-inspectable structures, the specified service life is the design life. For inspectable structures, the specified service life is the inspection interval. For base-level (overhaul) inspectable structures, the inspection interval is generally 1 / 4 of the design life. See GJB67.6A-2008 "Structural Strength Specification for Military Aircraft Part 6: Repeated Loads, Durability and Damage Tolerance" for details.

[0008] The type of structural detail determines the initial crack assumption, including the crack configuration and initial crack size. For details such as fastening holes, voids, and notches, the crack configuration is typically a corner crack or a through crack; for rounded corner details, the crack configuration is typically a corner crack or a through crack; for lug details, the crack configuration can be simplified to a corner crack or a through crack. For initial crack size, refer to the damage tolerance section of GJB67.6A-2008.

[0009] The allowable stress for damage tolerance can be theoretically calculated based on crack propagation analysis or obtained from crack propagation 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 specified service life requirements. This stress is then converted to the limiting load or ultimate load as needed to obtain the allowable stress.

[0010] The allowable stress for damage tolerance 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. In this invention, the allowable stress refers to the limiting load.

[0011] Since numerous factors influence the allowable stress for damage tolerance, relying solely on fatigue tests to obtain the allowable stress would result in an excessively large experimental 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 damage tolerance design of metal structures, primarily based on crack propagation analysis and supplemented by experimental correction. Summary of the Invention

[0012] The technical problem solved by this invention is as follows: This invention proposes a method for determining the allowable stress for damage tolerance design of aircraft metal structures. It mainly uses crack propagation analysis and is supplemented by experimental correction. It is used to determine the allowable stress for damage tolerance of various materials and details of aircraft structures under the design load spectrum and specified service life requirements, so as to conduct rapid damage tolerance assessment by judging the positive or negative value of the damage tolerance margin.

[0013] Technical solution: A method for determining the allowable stress for damage tolerance design of aircraft metal structures. The method first calculates the baseline allowable stress for structural details under given material and load spectrum, and introduces a stress correction coefficient based on the correction of theoretical analysis results by crack propagation test; considering the actual structure, a strengthening correction coefficient is introduced; considering the reliability requirements, a fatigue uncertainty coefficient is introduced, thereby establishing a set of analog calculation formulas to determine the allowable stress for damage tolerance design.

[0014] The formula for calculating the allowable stress for damage tolerance design is as follows:

[0015]

[0016] Among them, [σ ref,0 [ ] represents the reference allowable stress, SCF is the stress correction factor, and C cw To strengthen the correction factor, FSF is the fatigue uncertainty factor, C load This is the load conversion factor.

[0017] Furthermore, the process for determining each parameter in the formula for calculating the allowable stress for damage tolerance design is as follows:

[0018] For the target structure, determine the material grades, fatigue load spectrum types, detail types, and specified service life.

[0019] Select the crack configuration according to the described detailed type, and use the crack propagation analysis method. Based on the described material and load spectrum, calculate the theoretical nominal stress corresponding to the specified service life under the detailed geometry, and determine the reference allowable stress [σ]. ref,0 ];

[0020] Considering that the actual structure employs gain measures, a strengthening correction factor C is introduced. cw This coefficient needs to be obtained through experiments or by taking values ​​based on accumulated engineering experience;

[0021] For the aforementioned detail types and materials, based on crack propagation 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 crack propagation analysis method, the stress correction factor SCF is calculated, and statistical aggregation is performed according to material and detail type;

[0022] 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 ;

[0023] For the target structure, the fatigue uncertainty factor FSF is determined based on factors such as the lifetime dispersion factor, material property differences, and load spectrum differences.

[0024] Furthermore, the calculation process for the reference allowable stress is as follows:

[0025] 4) Given the material, load spectrum sequence, and crack configuration, select several nominal stress levels σ ref Using crack propagation analysis, the crack propagation life N from the initial crack size to the critical crack size is calculated, where σ ref 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.

[0026] 5) Based on several groups (σ) ref For sample points (N), the nominal stress-crack propagation life curve is fitted using a power function formula to obtain σ. ref The -N curve is shown below:

[0027]

[0028] In the formula, m and C are constants that are related to the crack propagation analysis method, material, load spectrum and crack configuration.

[0029] 6) According to σref The -N curve is used to calculate the theoretical nominal stress corresponding to the specified service life N0, which is the reference allowable stress [σ]. ref,0 The calculation formula is as follows, where [σ] ref,0 [Corresponds to the maximum load condition in the fatigue spectrum;]

[0030] [σ ref,0 ]=(C / N0) 1 / m .

[0031] Furthermore, the strengthening correction coefficient C was obtained through experiments. cw The method is as follows: First, design work-free and work-strength test specimens. Construct nominal stress-life curves through crack propagation tests at multiple stress levels. Then, calculate the stress based on the target life. The ratio of the two stresses is the strengthening correction factor C. cw The formula is as follows.

[0032] C cw =S cw / S0.

[0033] Furthermore, the formula for calculating the stress correction factor SCF is as follows:

[0034]

[0035] Among them, [σ ref '] refers to the theoretical nominal stress corresponding to the median life of the experiment, which is calculated by back-calculating the crack propagation life under the test load spectrum and the test specimen material. The crack propagation analysis method must be the same as the method used to determine the reference allowable stress, and N0 is replaced with the median life of the experiment, which refers to the crack propagation life from the initial crack size to the critical size; σ ref 'Refers to the nominal stress of the test.'

[0036] Furthermore, the load reduction factor C load The calculation formula is:

[0037] C load =F lim / F spec

[0038] Among them, F lim Indicates the limiting load, F spec This indicates the maximum load in the spectrum; this coefficient is related to the type of load spectrum.

[0039] Furthermore, the formula for calculating the fatigue uncertainty factor FSF is as follows:

[0040] FSF = Πβ i

[0041] β0 represents the baseline dispersion coefficient, which is derived by inversely from the crack propagation life dispersion coefficient. The calculation formula is as follows:

[0042] β0=LSF 1 / m

[0043] Where LSF represents the crack propagation life dispersion factor, and m represents the nominal stress-crack propagation life curve parameters under given material, load spectrum and crack configuration;

[0044] β m This represents the material property difference coefficient. Since crack propagation performance parameters for the current material grade are currently unavailable, we can only borrow from similar materials. (β) m ≥1.0;

[0045] 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.

[0046] Furthermore, the crack configuration depends on the detail type. For fastening holes, hollow holes, and notches, the crack configuration is a hole corner crack or a through crack; for rounded corners, the crack configuration is a flat plate edge corner crack or a through crack; for lug-type details, the crack configuration is a lug hole edge corner crack or a through crack.

[0047] Furthermore, the crack propagation analysis method includes the classical linear elastic fracture mechanics model and the probabilistic fracture mechanics model, and adopts a combination of crack propagation rate formula and load interaction model.

[0048] Furthermore, the crack propagation rate formula includes the Paris formula and the Walker formula, and the load interaction model includes the Willenborg model and the Closure model.

[0049] Beneficial technical effects: This invention proposes a method for determining the allowable stress for damage tolerance design of aircraft metal structures. It mainly uses crack propagation analysis and is supplemented by experimental correction. By analogy with the calculation formula, it simplifies the crack propagation test matrix and ensures the accuracy of theoretical analysis. By comparing the allowable stress and the working stress, it can quickly determine whether the structural damage tolerance margin meets the requirements, thereby achieving damage tolerance analysis covering all critical fracture components of the aircraft. Attached Figure Description

[0050] 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.

[0051] Figure 1 The technical process of this invention;

[0052] Figure 2 The nominal stress-life curve;

[0053] Figure 3 This is the wing bending moment spectrum sequence in the example (with the ordinate axis normalized).

[0054] Figure 4 These are three different crack configurations in the example.

[0055] Figure 5 The nominal stress-fatigue life curve (σ) of the corner crack of the fastening hole in the example. ref -N curve);

[0056] Figure 6 The nominal stress-fatigue life curve (σ) of the corner crack at the edge of the flat plate in the example. ref -N curve);

[0057] Figure 7 The nominal stress-fatigue life curve (σ) of the corner crack of the ear hole in the example. ref -N curve). Detailed Implementation

[0058] 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.

[0059] 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.

[0060] 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.

[0061] See Figure 1 The overall design concept of this invention is to first calculate the baseline allowable stress of the structural details under a given material and load spectrum using crack propagation analysis, and then introduce a stress correction coefficient based on the correction of the theoretical analysis results by crack propagation test; considering that the actual structure adopts measures such as cold working strengthening, a strengthening correction coefficient is introduced; in addition, considering the reliability requirements, a fatigue uncertainty coefficient is introduced, thereby establishing a set of analog calculation formulas to determine the damage tolerance design allowable stress.

[0062] The formula for calculating the allowable stress in damage tolerance design is as follows:

[0063]

[0064] Among them, [σ ref,0 [ ] represents the reference allowable stress, SCF is the stress correction factor, and C cw To strengthen the correction factor, FSF is the fatigue uncertainty factor, C load This is the load conversion factor.

[0065] The following is the calculation method for each parameter in the above formula during the specific implementation of this invention:

[0066] S1: For the target structure, determine the material grade, fatigue load spectrum type, detail type, and specified service life;

[0067] In S1, the load spectrum type refers to the component load spectrum directly related to the structural detail stress spectrum, including the wing bending moment spectrum, fuselage longitudinal bending moment spectrum, control surface hinge moment spectrum, landing gear load spectrum, etc.; the structural detail type includes several categories such as fastening holes, voids and notches, fillets, lugs, etc.; the specified service life is related to inspectability, see the damage tolerance chapter of GJB67.6A-2008.

[0068] S2: Select the crack configuration according to the detailed type described in S1, and use the crack propagation analysis method. Based on the material and load spectrum described in S1, calculate the theoretical nominal stress corresponding to the specified service life under the detailed geometry, and determine the reference allowable stress [σ]. ref,0 ];

[0069] In S2, crack propagation analysis methods include classical linear elastic fracture mechanics models and probabilistic fracture mechanics models. A combination of crack propagation rate formulas and load interaction models is often used. Commonly used crack propagation rate formulas include the Paris formula and the Walker formula, and commonly used load interaction models include the Willenborg model and the Closure model.

[0070] In S2, the crack configuration depends on the detail type. For fastening holes, hollow holes, and notches, the crack configuration is a hole corner crack or a through crack; for rounded corners, the crack configuration is a flat plate edge corner crack or a through crack; for lug-type details, the crack configuration is a lug hole edge corner crack or a through crack.

[0071] In S2, the method for calculating the reference allowable stress is as follows:

[0072] 1) Given the material, load spectrum sequence, and crack configuration, select several (at least 3) nominal stress levels σ. ref Using crack propagation analysis, the crack propagation life N from the initial crack size to the critical crack size is calculated, where σ ref 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.

[0073] 2) Based on several groups (σ) ref For sample points (N), the nominal stress-crack propagation life curve (σ) is fitted using a power function formula. ref -N curve), as follows:

[0074]

[0075] In the formula, m and C are constants that are related to the crack propagation analysis method, material, load spectrum, and crack configuration.

[0076] 3) According to σ ref The -N curve is used to calculate the theoretical nominal stress corresponding to the specified service life N0, which is the reference allowable stress [σ]. ref,0 The calculation formula is as follows, where [σ] ref,0 This corresponds to the maximum load condition in the fatigue spectrum.

[0077] [σ ref,0 ]=(C / N0) 1 / m

[0078] S3: Considering that the actual structure adopts gain measures such as cold working strengthening, a strengthening correction factor C is introduced. cw This coefficient needs to be obtained through experiments, or it can be obtained from accumulated engineering experience.

[0079] In S3, the method for obtaining this coefficient through experimentation is as follows: First, design work-hardened and non-work-hardened test specimens. Then, construct nominal stress-life curves through crack propagation tests at multiple stress levels. Finally, calculate the stress based on the target life, as follows: Figure 2 The ratio of the two stresses is the correction factor, and the formula is as follows.

[0080] C cw =S cw / S0

[0081] S4: For the detail type and material mentioned in S1, based on the crack propagation 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 crack propagation analysis method, calculate the stress correction factor SCF, and perform statistical merging according to material and detail type;

[0082] In S4, the formula for calculating the stress correction factor SCF is:

[0083]

[0084] Among them, [σ ref '] refers to the theoretical nominal stress corresponding to the median life of the experiment, which is derived from the crack propagation analysis method under the test load spectrum and the test specimen material. The calculation method is described in S2. The crack propagation analysis method must be the same as the method used to determine the reference allowable stress, and N0 should be replaced with the median life of the experiment, which refers to the crack propagation life from the initial crack size to the critical size; σ ref 'Refers to the nominal stress of the test.'

[0085] SCF reflects the degree of matching between crack propagation analysis and experiment. The closer the SCF is to 1.0, the better the match between the two. If SCF>1.0, the crack propagation analysis life is longer than the experimental life, and the analysis results are more dangerous.

[0086] 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 ;

[0087] In S5, the load reduction factor C load The calculation formula is:

[0088] C load =F lim / F spec

[0089] Among them, F lim Indicates the limiting load, F specThis 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.

[0090] S6: For the target structure described in S1, determine the fatigue uncertainty factor FSF based on factors such as the lifetime dispersion factor, material property differences, and load spectrum differences;

[0091] In S6, the fatigue uncertainty factor FSF is calculated using the following formula:

[0092] FSF = Πβ i

[0093] β0 represents the baseline dispersion coefficient, which is derived by inversely from the crack propagation life dispersion coefficient. The calculation formula is as follows:

[0094] β0=LSF 1 / m

[0095] Where LSF represents the crack propagation life dispersion factor, typically taken as 2.0, and m represents the nominal stress-crack propagation life curve (σ) under given material, load spectrum, and crack configuration. ref -N curve parameters.

[0096] β m This represents the material property difference coefficient. Since crack propagation performance parameters for the current material grade are currently unavailable, we can only borrow parameters from similar materials, such as materials of the same grade but with different heat treatments or different dimensions. (β) m ≥1.0;

[0097] 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.

[0098] S7: Substitute the parameters determined in S2 to S6 into the following calculation formula to determine the allowable stress [σ] for damage tolerance design of the actual structural details. ref ].

[0099]

[0100] It should be noted that the above calculation of 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.

[0101] Example 1: Implemented using the design described above, the technical process of this 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 damage tolerance design is explained.

[0102] 1) This frame is a TC4-DT die forging, mainly affected by the wing bending moment spectrum (spectral sequence as follows). Figure 2 The key parts involve the flange fastening holes and rounded corners; the frame is an uninspectable structure, and its service life is the design life, 8000FH.

[0103] 2) For flanged fastening holes, the crack configuration should be selected as a corner crack, such as... Figure 3 (a) With geometric dimensions W = 40 mm, T = 6 mm, and D = 4.76 mm, and an initial crack size of 1.27 mm, the Walker crack propagation rate formula and the improved Willenborg hysteresis model were used to calculate the theoretical nominal stress corresponding to a crack propagation life of 8000 FH under tensile load, and the reference allowable stress was determined.

[0104] [σ ref,0 =278 MPa, the process is as follows Figure 4 As shown, the nominal stress-crack propagation life curve is included.

[0105] 3) For rounded corner details, the crack configuration should be selected as a crack at the edge of a flat plate, such as... Figure 3 (b) With geometric dimensions W = 40 mm and T = 6 mm, and an initial crack size of 3.175 mm, the Walker crack propagation rate formula and the improved Willenborg hysteresis model were used to calculate the theoretical nominal stress corresponding to a crack propagation life of 8000 FH under tensile load, and the reference allowable stress [σ] was determined. ref,0 =227 MPa, the process is as follows Figure 5 As shown, the nominal stress-crack propagation life curve is included.

[0106] 4) For ear plate details, the crack configuration is selected from the corners of the ear plate holes, such as... Figure 3 (c) With geometric dimensions W = 150 mm, T = 24 mm, and D = 54 mm, and an initial crack size of 1.27 mm, the Walker crack propagation rate formula and the improved Willenborg hysteresis model were used to calculate the theoretical nominal stress corresponding to a crack propagation life of 8000 FH under tensile load, and to determine the reference allowable stress [σ]. ref,0 =336 MPa, the process is as follows Figure 6 As shown, the nominal stress-crack propagation life curve is included.

[0107] 5) The flange fastening holes are reinforced by cold extrusion. Based on previously accumulated test data, C cw Take 1.15; no reinforcement measures were used for the fillet, C cw Take 1.0; the ear hole adopts a press-fit bushing reinforcement measure. Based on the accumulated test data, C cw Take 1.25;

[0108] 6) Based on the crack propagation test data of existing fastening hole and fillet test specimens under similar load spectra, calculate the stress correction factor SCF. For TC4-DT, the SCF of fastening hole is 0.9, the SCF of fillet is 1.1, and the SCF of lug is 1.0.

[0109] 7) 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;

[0110] 8) Determine the fatigue uncertainty factor FSF based on factors such as the lifetime dispersion factor, material property differences, and load spectrum differences;

[0111] If the crack propagation life dispersion coefficient is 2.0, then the fastening hole β0 = 2^(1 / 2.833) = 1.277, the fillet β0 = 2^(1 / 2.887) = 1.271, and the lug hole β0 = 2^(1 / 3.129) = 1.248.

[0112] The fatigue performance parameters of the material are fixed, β m =1.0; the load spectrum sequence is also determined, β s =1.0.

[0113] Therefore, the fastening hole FSF = β0 * β m *β s =1.277, fillet FSF = β0 * β m *β s =1.271, Ear hole FSF = β0*β m *β s =1.248.

[0114] 9) Substitute the parameters determined above into the allowable stress calculation formula to determine the damage tolerance design allowable stress [σ] for the actual structural details. ref ].

[0115]

[0116] The damage tolerance allowable stress (corresponding limiting load) for typical structural details of TC4-DT die forging under the wing bending moment spectrum is listed in the table below.

[0117] Table 1 TC4-DT-Flying Wing Bending Moment Spectrum - Damage Tolerance Allowable Stress (MPa)

[0118] Fastening hole Rounded corners ear <![CDATA[[σ ref ]]]> 348 203 421

[0119] 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.

[0120] 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.

[0121] 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.

[0122] 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 damage tolerance design of aircraft metal structures, characterized in that, The method first calculates the baseline allowable stress for structural details under given material and load spectrum, and introduces a stress correction factor based on the correction of theoretical analysis results by crack propagation test. Considering the actual structure, a strengthening correction factor is introduced; Considering the reliability requirements, fatigue uncertainty coefficients are introduced, and a set of analogous calculation formulas are established to determine the allowable stress for damage tolerance design. The damage tolerance design allowable stress The calculation formula is: Among them, [σ ref,0 [ ] represents the reference allowable stress, SCF is the stress correction factor, and C cw To strengthen the correction factor, FSF is the fatigue uncertainty factor, C load This is the load conversion factor; The process for determining each parameter in the formula for calculating the allowable stress for damage tolerance design is as follows: For the target structure, determine the material grades, fatigue load spectrum types, detail types, and specified service life. Select the crack configuration according to the described detailed type, and use the crack propagation analysis method. Based on the described material and load spectrum, calculate the theoretical nominal stress corresponding to the specified service life under the detailed geometry, and determine the reference allowable stress [σ]. ref,0 ]; Considering that the actual structure employs gain measures, a strengthening correction factor C is introduced. cw This coefficient needs to be obtained through experiments or by taking values ​​based on accumulated engineering experience; For the aforementioned detail types and materials, based on crack propagation 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 crack propagation 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 lifetime dispersion coefficient, material property differences, and load spectrum differences.

2. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 1, characterized in that, The calculation process for the reference allowable stress is as follows: 1) Given the material, load spectrum sequence, and crack configuration, select several nominal stress levels σ ref Using crack propagation analysis, the crack propagation life N from the initial crack size to the critical crack size is calculated, where σ ref 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) ref For sample points (N), the nominal stress-crack propagation life curve is fitted using a power function formula to obtain σ. ref The -N curve is shown below: In the formula, m and C are constants that are related to the crack propagation analysis method, material, load spectrum and crack configuration. 3) According to σ ref The -N curve is used to calculate the theoretical nominal stress corresponding to the specified service life N0, which is the reference allowable stress [σ]. ref,0 The calculation formula is as follows, where [σ] ref,0 [Corresponds to the maximum load condition in the fatigue spectrum;] 。 3. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 2, characterized in that, The enhancement correction coefficient C was obtained through experiments. cw The method is as follows: First, design work-free and work-strength test specimens. Construct nominal stress-life curves through crack propagation tests at multiple stress levels. Then, calculate the stress based on the target life. The ratio of the two stresses is the strengthening correction factor C. cw The formula is as follows; C cw = S cw / S0。 4. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 2, characterized in that, The formula for calculating the stress correction factor SCF is: Among them, [σ ref '] refers to the theoretical nominal stress corresponding to the median life of the experiment, which is calculated by back-calculating the crack propagation life under the test load spectrum and the test specimen material. The crack propagation analysis method must be the same as the method used to determine the reference allowable stress, and N0 is replaced with the median life of the experiment, which refers to the crack propagation life from the initial crack size to the critical size; σ ref 'Refers to the nominal stress of the test.' 5. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 2, characterized in that, Load conversion factor C load The calculation formula is: Among them, F lim Indicates the limiting load, F spec This indicates the maximum load in the spectrum; this coefficient is related to the type of load spectrum.

6. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 2, 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 crack propagation life dispersion coefficient. The calculation formula is as follows: Where LSF represents the crack propagation life dispersion factor, and m represents the nominal stress-crack propagation life curve parameters under given material, load spectrum and crack configuration; β m This represents the material property difference coefficient. Since crack propagation 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.

7. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 2, characterized in that, The crack configuration depends on the detail type. For fastening holes, hollow holes, and notches, the crack configuration is a hole corner crack or a through crack; for rounded corners, the crack configuration is a flat plate edge corner crack or a through crack; for lug-type details, the crack configuration is a lug hole edge corner crack or a through crack.

8. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 7, characterized in that, The crack propagation analysis method includes the classical linear elastic fracture mechanics model and the probabilistic fracture mechanics model, and adopts a combination of crack propagation rate formula and load interaction model.

9. The method for determining the allowable stress for damage tolerance design of aircraft metal structures as described in claim 8, characterized in that, The crack propagation rate formulas include the Paris formula and the Walker formula, and the load interaction models include the Willenborg model and the Closure model.

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