A method for determining the whole-aircraft health index for aircraft structural health monitoring

Abstract

Currently, domestic and international aircraft structure PHM technologies mainly use macro prediction methods based on the aircraft center of gravity overload spectrum, local prediction methods based on the stress spectrum of key parts, and prediction methods based on damage monitoring sensors. These methods are too macroscopic and cannot reflect the actual situation of the structure, or only reflect the local situation of the structure and lack the overall health assessment of the aircraft. There is a lack of aircraft health status evaluation methods covering key parts, key components, and the entire aircraft. The present invention proposes a method for determining the whole-aircraft health index for aircraft structural health monitoring. It adopts a bottom-up comprehensive and top-down judgment approach, which not only covers the health assessment and life prediction of all key parts of the main load-bearing structure, but also solves the health assessment problem of the aircraft's global structure. It can be used in aircraft structural health monitoring to guide and optimize the inspection and maintenance plans of in-service aircraft, reduce the maintenance cost of the aircraft throughout its life cycle, and ensure the safety of aircraft use.

Description

Technical Field

[0001] The present invention belongs to the technical field of aircraft health management, and in particular relates to a method for determining a whole-aircraft health index for aircraft structural health monitoring. Background Art

[0002] Aircraft structural prognostic health management (PHM) technology is a key enabler for aircraft structural health assessment, remaining life prediction, and optimized aircraft structural maintenance plans. Currently, this technology is being vigorously developed and applied in both military and civilian aircraft. Based on the research and application of fighter aircraft both domestically and internationally, the main methods used in aircraft structural PHM include: macroscopic prediction methods based on the aircraft's center of gravity overload spectrum, local prediction methods based on stress spectra at key locations, and prediction methods based on damage monitoring sensors. These methods are either too macroscopic and fail to reflect the entire structure, or they only reflect local structural conditions and lack a comprehensive assessment of the aircraft's overall health. Each has its advantages and disadvantages, but none are comprehensive. A comprehensive method for measuring aircraft health from a comprehensive perspective has yet to be established. Summary of the Invention

[0003] Purpose of the present invention: This invention proposes a method for determining the whole-aircraft health index for aircraft structural health monitoring. This method adopts a bottom-up synthesis and top-down judgment approach, covering the health assessment and life prediction of all key parts of the main load-bearing structure, and providing a health evaluation of the aircraft's global structure. This method can be used in aircraft structural health monitoring to guide and optimize the inspection and maintenance plans of in-service aircraft, reduce maintenance costs, and ensure aircraft safety.

[0004] The technical solution of the present invention:

[0005] A method for determining an aircraft health index for aircraft structural health monitoring comprises the following steps:

[0006] Step 1: Obtain the flight parameter history of each aircraft in service, use machine learning methods such as clustering and artificial neural networks to identify maneuvers and extract corresponding flight parameters, such as center of gravity lateral overload, center of gravity heading overload, center of gravity normal overload, Mach number, pressure altitude, roll rate, pitch rate, yaw rate, angle of attack, sideslip angle, pitch angle, roll angle, etc.;

[0007] Maneuvers to be identified include: dives, jumps, somersaults, rolls, circling, bucket flips, etc.

[0008] Based on various flight parameter-load models established during the aircraft design phase, such as the flight parameter-wing bending moment model, the flight parameter-wing shear force model, the flight parameter-wing torque model, the flight parameter-vertical tail bending moment model, the flight parameter-vertical tail shear force model, the flight parameter-vertical tail torque model, etc., the corresponding flight parameters extracted above are substituted into the model to calculate the total load of aircraft components under various maneuvers, such as the wing bending moment, torque, shear force, vertical tail bending moment, torque, shear force, and rudder hinge moment.

[0009] Step 2: Substitute the component loads calculated in Step 1 into the load-stress equation established during the aircraft design phase, calculate the stress history of key parts of each key component of the aircraft, and obtain the stress spectrum of each key part of each fatigue critical component;

[0010] Key parts include: fracture key parts and durability key parts;

[0011] Critical fracture components are structures that affect flight safety and whose failure alone could result in aircraft damage, casualties of flight crew members, or unintentional external loadouts. Examples include the main frame connecting the wing and fuselage, main beams, landing gear support structures, and engine mounting support structures.

[0012] Durability-critical components are structures that do not affect flight safety, but whose failure alone may cause a reduction in aircraft performance or a significant increase in maintenance costs. For example, other main load-bearing frames and beams other than the wing-fuselage connection are considered durability-critical components.

[0013] Critical parts refer to several detailed parts with the highest stress levels in durability-critical components and fracture-critical components. A critical component may have multiple critical parts, such as the ear holes connecting the wing and body to the main load-bearing frame, high-stress fastening holes inside the frame, system openings, fillet areas, etc.; the life of critical parts is determined by multiple factors such as their stress spectrum, detailed characteristic parameters, surface roughness, etc. The part with the highest stress in the same critical component is not necessarily the shortest life.

[0014] Step 3: Based on the stress spectra of each key part of the aircraft, use the probabilistic strain fatigue life prediction model or the probabilistic fracture mechanics model to predict the probabilistic life distribution of each key part;

[0015] For fracture-critical components, the strain fatigue life prediction model and the fracture mechanics life prediction model are used. It is assumed that the model input parameters obey the normal distribution. Based on the parameter fitting results of multiple groups of test data, the characteristic values (mean, standard deviation) of the parameter probability distribution function are statistically obtained. Several parameter combinations are randomly selected from the parameter probability distribution function. The crack initiation life or crack propagation life of each key part under each parameter combination is predicted, and the probabilistic crack initiation life distribution and the probabilistic crack propagation life distribution are statistically obtained.

[0016] For durability key parts, the strain fatigue life prediction model is used and the same method as the fracture key parts is adopted to obtain the probabilistic crack initiation life distribution of each key part.

[0017] Step 4: Calculate the durability health index (DHI) and damage tolerance health index (DTHI) of each key part. The specific method is as follows:

[0018] Assuming that the crack initiation life or crack propagation life of key parts obeys the log-normal distribution, N is used to represent the crack initiation life or crack propagation life. The durability health index DHI is calculated based on the probabilistic crack initiation life distribution and refers to the probability that the crack initiation life is greater than a certain maintenance time T. The calculation method is shown in formula (1):

[0019]

[0020] The damage tolerance health index DTHI is calculated based on the probabilistic crack growth life, which refers to the probability that the crack growth life is greater than a certain inspection time t. The calculation method is shown in formula (2).

[0021]

[0022] Step 5: Calculate the durability health index Struc_DHI or damage tolerance health index Struc_DTHI of each key component (including fracture key components and durability key components);

[0023] The calculation method of the durability health index Struc_DHI of key parts is shown in formula (3):

[0024]

[0025] m is the number of key parts contained in a fracture key component or durability key component;

[0026] The calculation method of the damage tolerance health index Struc_DTHI of the fracture key component is shown in formula (4):

[0027]

[0028] n is the number of key parts contained in a fracture key component.

[0029] Step 6: Calculate the aircraft's structural durability health index Total_DHI and damage tolerance health index Total_DTHI;

[0030] The calculation method of the aircraft global structural durability health index Total_DHI is shown in formulas (5) to (8);

[0031] Total_DHI=a·Total_DHI1+b·Total_DHI2 (5)

[0032]

[0033]

[0034] Here, p represents the number of critical parts (including fracture critical parts and durability critical parts) on the aircraft. There are x parts with good repair economics and px parts with poor repair economics, respectively. These two types of critical parts have different impacts on the aircraft's health. With the same remaining service life, critical parts with poor repair economics should have a greater impact on the aircraft's health index than critical parts with good repair economics. To account for this difference in impact, a comprehensive weight distribution coefficient is assigned to critical parts with good repair economics and poor repair economics, respectively. The comprehensive weight distribution coefficients for critical parts with good repair economics and poor repair economics are a and b, respectively, as specified:

[0035] a+b=1 (b>a) (8)

[0036] Repair economics is determined through a comprehensive assessment of each key component's characteristics, taking into account assembly complexity, repair difficulty, inspectability, accessibility, and repair cycle. Components that are complex to assemble and difficult to disassemble, have poor accessibility, have long repair cycles, or are difficult to repair are considered poor repair economics. The main frame and main beam connecting the wing and fuselage generally fall into this category.

[0037] The calculation method of the damage tolerance health index Total_DTHI of the entire aircraft structure is shown in formulas (9) to (11).

[0038]

[0039] Struc_DTHI j =Struc_DTHI 2j-1 Struc_DTHI 2j (10)

[0040]

[0041] Where q represents the number of fracture critical parts in the aircraft body structure, and the damage tolerance health indexes of y groups of fracture critical parts (assuming each group contains two fracture critical parts) affect each other, while the damage tolerance health indexes of the remaining q-2y fracture critical parts are independent of each other and of the aforementioned y groups of critical parts; qy represents the number of independent fracture critical parts among the q fracture critical parts; and c represents the minimum value of the structural damage tolerance health index among the q-2y fracture critical parts without mutual influence and among the y groups of fracture critical parts with mutual influence.

[0042] Step 7: Perform a global health evaluation of the aircraft, including the Total_DHI evaluation criteria and the Total_DTHI evaluation criteria.

[0043] Total_DHI evaluation criteria include:

[0044] a) Within a certain future overhaul interval T, if Total_DHI ≥ 99.9%, the aircraft is considered to be in good health and the aircraft structure does not require maintenance;

[0045] b) Within a certain future overhaul interval T, if 95% ≤ Total_DHI < 99.9%, the aircraft is considered to be in good overall health. There is a certain probability that the damage accumulation and life consumption of individual structures have reached or will soon reach the critical value for economic repair. The corresponding structures need to be inspected according to the periodic inspection of the maintenance regulations. Based on the inspection results, a decision will be made to allow continued use to a specified time or to perform immediate repairs;

[0046] c) If Total_DHI < 95% within a certain future overhaul interval T, the aircraft is considered to be in poor overall health. There is a high probability that the damage accumulation and life consumption of some structures have reached or fallen below the critical value for economic repair. Immediate nondestructive inspection of the corresponding structures is required, and a decision on whether to repair them is made based on the inspection results and combined with design analysis.

[0047] If damage is found during non-destructive testing, the aircraft must be grounded immediately and a repair and reinforcement plan must be developed as soon as possible to complete the repair and restore the aircraft to normal use. If no damage is found during non-destructive testing, but the design analysis supports with a high degree of confidence that the part needs repair, preventive maintenance must also be arranged immediately. The aircraft's service life can only be extended if there is sufficient analysis to show that the structure will not cause a loss of structural repair economy within a certain period of time in the future (until the next maintenance time) and will not bring unacceptable failure risks that affect flight safety.

[0048] Total_DTHI evaluation criteria include:

[0049] Assume that from the current accumulated service life, the first inspection time point is t1, and the second inspection time point is t2;

[0050] a) If Total_DTHI ≥ 99.9% during the inspection interval from the current time to t2, the aircraft structure has very good damage tolerance characteristics and the risk of fracture failure during the expected interval is extremely low. At t1, the aircraft does not need to be scheduled for non-destructive inspection;

[0051] b) During the inspection interval from now to t2, if Total_DTHI ≥ 99.9% before t1, but gradually decreases to below 99.9% after t1, the aircraft maintains good damage tolerance characteristics before t1, but if it continues to be used after t1, there is a greater risk of partial structural failure. Therefore, when the aircraft is in normal use until t1, non-destructive inspections and preventive maintenance work as appropriate should be arranged;

[0052] c) If, during the inspection interval from now until t1, Total_DTHI gradually decreases and ultimately falls below 99.9% before reaching t1, this indicates that a portion of the aircraft structure presents a significant risk of failure before t1. Unless sufficient analysis and testing demonstrate that the structure can continue to be used until t1 without inspection or repair without an unacceptable risk of fracture, a nondestructive inspection (NDT) of the structure should be scheduled as soon as possible, and preventive maintenance should be performed as appropriate. If the NDT results indicate a good structural health, combined with a failure risk assessment, and provided the failure risk is acceptable, continued use until t1 with (preventive) maintenance may be permitted.

[0053] Beneficial effects of the present invention:

[0054] This invention establishes a method for determining the health index of the entire aircraft for aircraft structural health monitoring and proposes calculation methods for the durability health index and damage tolerance health index for evaluating the health status of the aircraft body structure. From a probabilistic perspective, the health index of key parts, critical components, and the entire aircraft structure is calculated step by step. The health status is graded by the health index, and inspection and maintenance recommendations corresponding to different levels are given. Each grade takes into account a high degree of life reliability, meeting the requirements of durability ensuring the economic efficiency of repairs and damage tolerance ensuring flight safety. This invention supplements and improves the macro-evaluation of the health status of the entire aircraft by aircraft structural health monitoring, which not only helps users understand the current health status of the aircraft body, but also helps optimize the inspection and maintenance intervals of the aircraft, thereby reducing the maintenance cost throughout its life cycle. BRIEF DESCRIPTION OF THE DRAWINGS

[0055] Figure 1 Flowchart of a method for determining the whole-aircraft health index for aircraft structural health monitoring

[0056] Figure 2 Schematic diagram showing that the probability distribution of crack initiation life or crack propagation life at key parts continuously moves toward the lower side with increasing service time;

[0057] Figure 3 This is a schematic diagram of the connotation of the key parts durability health index or damage tolerance health index. The shaded area represents the probability that the life span is greater than a certain time, that is, the key parts durability health index or damage tolerance health index;

[0058] Figure 4 A schematic diagram showing the changing trend of the durability health index or damage tolerance health index of key components with service time;

[0059] Figure 5 This is a schematic diagram showing the changing trend of the aircraft structure durability health index or damage tolerance health index with service time. Specific embodiments

[0060] The following will be combined with the drawings in the embodiments of the present invention to clearly and completely describe the technical solutions in the embodiments of the present invention. Obviously, the embodiments described are only part of the embodiments of the present invention, not all of the embodiments. Based on the embodiments of the present invention, all other embodiments obtained by ordinary technicians in this field without making creative work are within the scope of protection of the present invention. The overall technical process of the present invention is as follows: Figure 1 As shown in the figure, during the aircraft design phase, a mapping relationship between flight parameters and component loads (bending moment, torque, shear force, intersection loads, etc.) is established based on a load design database, known as the "flight parameter-load" equation. Finite element simulation is then used to obtain the structural stress distribution, identify key locations, and establish a mapping relationship between local stresses in these locations and component loads, known as the "load-stress" equation. After the aircraft enters service, the measured flight parameter histories are substituted into the "flight parameter-load" equation to obtain the measured component load histories. Substituting these into the "load-stress" equation yields the stress histories at these key locations, known as the stress spectrum. Using the stress spectrum as input, classical strain-based fatigue life prediction methods and linear elastic fracture mechanics theory are employed, with model parameters entered as probability distributions. The remaining life distribution of key locations corresponding to this spectrum is predicted. Based on the remaining life distribution of each key location, durability health indices and damage tolerance health indices are established for key locations, key components, and the entire aircraft, according to structural classification. This allows for a bottom-up assessment of the aircraft's health and a top-down decision-making process regarding the continued use, inspection, and maintenance of the aircraft structure.

[0061] A method for determining a whole-aircraft health index for aircraft structural health monitoring, one of the possible specific embodiments of which is as follows:

[0062] Assume that a certain aircraft has 15 critical parts, of which 10 are fracture-critical parts, including the fuselage, main load-bearing frames of the wings, main load-bearing beams, and landing gear support beams. Five are durability-critical parts, including door joints and control surface suspension joints. Nine of these parts are economically repairable, while six are economically unrepairable. Among the 10 fractured critical parts, the damage tolerance indexes of two groups (two parts each) affect each other, and the critical parts have one or more critical locations.

[0063] The first step is to use machine learning methods such as clustering and artificial neural networks to identify various maneuvers such as dives, jumps, and somersaults based on the flight parameter histories recorded by in-service aircraft, and extract their corresponding flight parameter histories, such as center of gravity lateral overload, center of gravity heading overload, center of gravity normal overload, Mach number, pressure altitude, roll rate, pitch rate, yaw rate, angle of attack, sideslip angle, pitch angle, roll angle, etc.

[0064] The extracted flight parameter histories are substituted into various flight parameter-load models established during the aircraft design phase to calculate the total load histories of aircraft components under various maneuvers, such as wing bending moment, torque, shear force, vertical tail bending moment, torque, shear force, rudder hinge moment, etc.

[0065] The second step is to substitute the component loads calculated in the first step into the 15 key load-stress equations established during the aircraft design phase, calculate the stress history of the key parts of the 15 key components, and compile the resulting stress spectrum;

[0066] The third step is to use the stress spectrum of the second step, apply the strain fatigue life prediction model or linear elastic fracture mechanics theory, introduce probability theory, and predict the probability life distribution of each key part. Figure 2 It can be seen that the probability distribution of the life of key parts gradually moves to the lower side as the service time increases;

[0067] For fracture-critical components, the strain fatigue life prediction model and the fracture mechanics life prediction model are used. It is assumed that the model input parameters obey the normal distribution. Based on the parameter fitting results of multiple groups of test data, the characteristic values (mean, standard deviation) of the probability distribution function of the parameters are statistically obtained. Several groups of parameter combinations are randomly selected from the parameter probability distribution. The crack initiation life or crack propagation life of each key part under each parameter combination is predicted, and the probabilistic crack initiation life distribution and the probabilistic crack propagation life distribution are obtained by statistics.

[0068] For durability key parts, the strain fatigue life prediction model is used and the same method as the fracture key parts is adopted to obtain the probabilistic crack initiation life distribution of each key part.

[0069] Step 4: Given a maintenance time of T = 2000 flight hours and an inspection time of t = 1000 flight hours, the durability health index DHI of each key part of the critical component and the damage tolerance health index DTHI of each key part of the fractured critical component are calculated using formulas (1) and (2) in step 4 of the technical solution.

[0070] like Figure 3 As shown, DHI and DTHI are the areas of the parts with lifetime greater than T or t in the probability lifetime distribution of the key parts.

[0071] In the fifth step, based on the DHI and DTHI of each key part calculated in the fourth step, the durability health index Struc_DHI of 15 key parts and the damage tolerance health index Struc_DTHI of 10 fractured key parts are calculated using the method of step five and formula (3) and formula (4), as shown in Table 1; the change trend of the durability health index of key parts with the increase of service time is shown in Figure 4 shown.

[0072]

[0073] In the sixth step, it is assumed that the key parts with good and poor repair economy are as shown in Table 1, and the repair economy weight coefficients a = 0.4, b = 0.6; the damage tolerance health index of the two groups of key parts that affect each other are shown in Table 1, and the method of step six and formula (5) to formula (11) are used to calculate the aircraft, the total aircraft structure durability health index Total_DHI and the total aircraft structure damage tolerance health index Total_DTHI, the results are shown in Table 1; the change trend of Total_DHI or Total_DTHI with service time is as follows Figure 5 shown.

[0074] The seventh step is to conduct a health assessment of the entire aircraft structure based on the calculation results of the sixth step, including the health status assessment of the durability of the entire aircraft structure and the health assessment of the damage tolerance of the entire aircraft structure.

[0075] Since Total_DHI = 0.8995, which is less than 0.95, when the estimated maintenance time T = 2000 flight hours, the aircraft's overall health is considered poor. There is a high probability that the damage accumulation and life consumption of some structures have reached or fallen below the critical value for economic repair. Immediate non-destructive inspection of the relevant structures is required, and a decision is made based on the inspection results. If damage is found during non-destructive inspection, the aircraft must be grounded immediately, and repair work must be carried out according to the repair plan for the key parts of the structure, or a new repair and reinforcement plan must be formulated as appropriate to complete the repair as soon as possible and restore the aircraft to normal service. If no damage is found during non-destructive inspection, but the design analysis has a high degree of confidence that the part needs repair, preventive maintenance must also be arranged immediately. The aircraft's service extension can only be allowed if sufficient analysis shows that the structure will not cause a loss of economic structural repair within the next 2000 flight hours and will not introduce an unacceptable failure risk that will affect flight safety.

[0076] Since Total_DTHI = 0.9508, less than 99.9% at the estimated inspection time t = 1000 flight hours, this indicates that some aircraft structures are at high risk of failure before reaching 1000 flight hours. Unless sufficient analysis and testing demonstrate that the corresponding structure can continue to be used for 1000 flight hours without inspection or repair without an unacceptable risk of fracture failure, non-destructive testing (NDT) should be scheduled as soon as possible and preventive maintenance should be carried out as appropriate. If the NDT results indicate a good structural health, combined with a failure risk assessment, and provided that the failure risk is acceptable, (preventive) maintenance can be approved until 1000 flight hours.

[0077] The above is merely one specific embodiment of the present invention, and the present invention is described in detail. Any unspecified portion represents conventional technology. However, the scope of protection of the present invention is not limited thereto. Any changes or substitutions that can be easily conceived by a person skilled in the art within the technical scope disclosed in the present invention should be included within the scope of protection of the present invention. The scope of protection of the present invention should be based on the scope of protection of the claims.

Claims

1. A method for determining an aircraft health index for aircraft structural health monitoring, characterized by: The method comprises the following steps: Step 1: Obtain the flight parameter history of each aircraft in service, use machine learning methods to identify maneuvers and extract the corresponding flight parameters; based on the various flight parameter-load models established during the aircraft design phase, substitute the extracted flight parameters into the corresponding flight parameter-load models to calculate the total load history of aircraft components under various maneuvers; Step 2: Substitute the total component load history calculated in Step 1 into the load-stress equation established during the aircraft design phase, calculate the stress history of key locations of each critical component of the aircraft, and obtain the stress spectrum of each key location of each fatigue critical component; Step 3: Based on the stress spectra of each key part of the aircraft, use the probabilistic strain fatigue life prediction model or the probabilistic fracture mechanics model to predict the probabilistic life distribution of each key part; Step 4: Calculate the durability health index DHI and damage tolerance health index DTHI of each key part; Step 5: Calculate the durability health index Struc_DHI and damage tolerance health index Struc_DTHI of each key component; Step 6: Calculate the aircraft's structural durability health index Total_DHI and damage tolerance health index Total_DTHI; Step 7: Perform a global health evaluation of the aircraft, including the Total_DHI evaluation criteria and the Total_DTHI evaluation criteria.

2. The method according to claim 1, wherein: In step 1, the maneuvers to be identified include: dive, jump, somersault, roll, circling, and half-pound flip; The flight parameters to be extracted include: lateral overload at the center of gravity, directional overload at the center of gravity, normal overload at the center of gravity, Mach number, pressure altitude, roll rate, pitch rate, yaw rate, angle of attack, sideslip angle, pitch angle, and roll angle; The total loads on aircraft components include: bending moment, torque, shear force at the root of the wing and tail surface, hinge moment of the rudder, and intersection load of the landing gear.

3. The method according to claim 2, wherein: In the step 2, the key parts include: fracture key parts and durability key parts; The key parts that broke included: wings, main frame connected to the fuselage, main beams, landing gear support structure, and engine mounting support structure; Durability key components include: other main load-bearing frames and beams except the wing-fuselage connection; Critical parts refer to the detailed parts in durability key parts and fracture key parts where the stress level is higher than a certain value.

4. The method according to claim 3, wherein: In the third step, for the fracture key parts, a strain fatigue life prediction model and a fracture mechanics life prediction model are used, assuming that the model input parameters obey the normal distribution, and based on the parameter fitting results of multiple groups of test data, the characteristic values of the parameter probability distribution function are statistically obtained. Several parameter combinations are randomly selected from the parameter probability distribution function, and the crack initiation life or crack propagation life of each key part under each parameter combination is predicted, and the probabilistic crack initiation life distribution and the probabilistic crack propagation life distribution are statistically obtained; For durability key parts, the strain fatigue life prediction model is used and the same method as the fracture key parts is adopted to obtain the probabilistic crack initiation life distribution of each key part.

5. The method according to claim 4, characterized in that: In step 4, it is assumed that the crack initiation life or crack propagation life of the key parts obeys a log-normal distribution, and N is used to represent the crack initiation life or crack propagation life. The durability health index DHI is calculated based on the probabilistic crack initiation life distribution and refers to the probability that the crack initiation life is greater than a certain maintenance time T. The calculation method is as follows: The damage tolerance health index DTHI is calculated based on the probabilistic crack growth life, which refers to the probability that the crack growth life is greater than a certain inspection time t. The calculation formula is as follows: 。 6. The method according to claim 5, characterized in that: In step 5, the durability health index Struc_DHI of the fracture key parts and the durability key parts is calculated as follows: m is the number of key parts contained in a fracture key component or durability key component; The calculation formula of the damage tolerance health index Struc_DTHI of the fracture key component is as follows: n is the number of key parts contained in a fracture key component.

7. The method according to claim 6, characterized in that: In step 6, the calculation formula of the aircraft global structural durability health index Total_DHI is as follows: Where p represents the number of critical parts of the aircraft, which includes fracture critical parts and durability critical parts. Among them, there are x parts with good repair economy and px parts with poor repair economy, respectively. The comprehensive weight distribution coefficients of critical parts with good repair economy and poor repair economy are a and b, respectively, and satisfy the following relationship: a+b=1, b>a.

8. The method according to claim 7, wherein: In step 6, the calculation formula of the damage tolerance health index Total_DTHI of the aircraft structure is as follows: Where q represents the number of fracture critical parts in the aircraft structure, and there are y groups of fracture critical parts, each group containing two fracture critical parts. The damage tolerance health indexes of the fracture critical parts affect each other, and the damage tolerance health indexes of the remaining q-2y fracture critical parts are independent of each other and of the y groups of critical parts. qy represents the number of independent fracture critical parts among the q fracture critical parts. c represents the minimum value of the structural damage tolerance health index among the q-2y fracture critical parts that do not affect each other and among the y groups of fracture critical parts that affect each other.

9. The method according to claim 8, characterized in that: In step 7, the Total_DHI evaluation criteria include: a) Within a certain future overhaul interval T, if Total_DHI ≥ 99.9%, the aircraft is considered to be in good health and the aircraft structure does not require maintenance; b) Within a certain future overhaul interval T, if 95% ≤ Total_DHI < 99.9%, the aircraft is considered to be in good overall health. The damage accumulation and life consumption of individual structures have a certain probability of reaching or about to reach the critical value for economic repair. The corresponding structures need to be inspected according to the periodic inspection of the maintenance regulations. Based on the inspection results, a decision will be made to allow continued use to a specified time or to perform immediate repairs. c) If Total_DHI < 95% within a certain future overhaul interval T, the aircraft is considered to be in poor overall health. There is a high probability that the damage accumulation and life consumption of some structures have reached or fallen below the critical value for economic repair. Immediate non-destructive inspection of the corresponding structures is required, and a decision on whether to repair them is made based on the inspection results and combined with design analysis.

10. The method according to claim 8, characterized in that: In step seven, the Total_DTHI evaluation criteria include: Assume that from the current accumulated service life, the first inspection time point is t1, and the second inspection time point is t2; a) If Total_DTHI ≥ 99.9% during the inspection interval from the current time to t2, the aircraft structure has very good damage tolerance characteristics and the risk of fracture failure during the expected interval is extremely low. At t1, no nondestructive inspection is scheduled for the aircraft. b) During the inspection interval from now to t2, if Total_DTHI is ≥ 99.9% before t1, but gradually decreases to below 99.9% after t1, the aircraft maintains good damage tolerance characteristics before t1, but there is a high risk of partial structural failure if it continues to be used after t1. Therefore, when the aircraft is in normal use until t1, non-destructive inspections and preventive maintenance work as appropriate should be arranged; c) If, during the inspection interval from the current time to t1, the Total_DTHI gradually decreases and ultimately falls below 99.9% before reaching t1, it indicates that a portion of the aircraft structure is at significant risk of failure before t1. Unless sufficient analysis and testing demonstrate that the structure can continue to be used until t1 without inspection or repair without incurring an unacceptable risk of fracture failure, nondestructive inspection of the structure and, if appropriate, preventive maintenance should be performed as soon as possible.

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