Helicopter blade pin fatigue life evaluation method considering folding and mooring overload

Abstract

The invention belongs to the technical field of structural strength design, and discloses a helicopter blade pin fatigue life evaluation method considering folding and mooring overload, and the method comprises the following steps: 1, evaluating the blade pin load and fatigue damage of a blade in a folding state; 2, evaluating the load and fatigue damage of the blade pin under mooring overload; and 3, evaluating the fatigue life of the blade pin under the folding state superposition mooring overload.

Description

Technical Field

[0001] The invention belongs to the technical field of structural strength design, and in particular relates to a method for evaluating the fatigue life of a helicopter blade pin taking into account folding and mooring overloads. Background Art

[0002] The development of helicopter use has put forward the task requirements of folding and tethering overload. As a key component of the helicopter rotor system, the blade pin is very important for the flight safety of the helicopter to evaluate the impact of folding and tethering overload on the fatigue life of the blade pin. For helicopter blade pins, the working modes without fretting and with fretting are often considered, and the life assessment is carried out using the safety fatigue design method. There is no mature technology to analyze the impact of folding and tethering overload processes on the fatigue life of the blade pin, which needs further research. Summary of the invention

[0003] Purpose of the invention: This invention considers the influence of blade deadweight, land wind load and manual operation during blade folding, and proposes an analysis method for the maximum extrusion load and maximum local stress of the blade pin; considers the influence of the maximum vertical overload and the blade pull under the clamping end during the mooring process, and proposes an analysis method for the maximum extrusion load and maximum local stress of the blade pin. On this basis, considering the influence of fretting abrasion and fretting abrasion, an evaluation method for fatigue damage and fatigue life of helicopter blade pins is proposed, which provides assistance for the design of key components of helicopter rotors under folding and mooring overloads.

[0004] Technical Solution

[0005] A method for fatigue life assessment of helicopter blade pins considering folding and tethered overloads is as follows:

[0006] Step 1: Evaluate the blade pin load and fatigue damage in the folded blade state;

[0007] Step 2: Evaluate blade pin load and fatigue damage under moored overload;

[0008] Step 3: Evaluate the blade pin fatigue life under folded state superimposed on moored overload.

[0009] Further, in step one, the process is as follows:

[0010] The maximum extrusion load F of the blade pin is calculated based on the most dangerous bending moment caused by the blade's deadweight, land wind load and manual operation. 1 ;

[0011] Calculate the maximum local stress σ on the blade pin based on shear and bending deformations max1 ;

[0012] The maximum local stress σ is calculated using the equal life model. max1After correction, the low-cycle equivalent dynamic stress of the blade pin is obtained;

[0013] According to the low-cycle equivalent dynamic stress σ of the blade pin aeq1 , the fatigue performance Stromyer model is used, combined with the material performance parameters of the blade pin, to calculate the low cycle life cycle number N of the blade pin without fretting wear in folding 1- , and low cycle life cycle number N with fretting 1+ ;

[0014] Based on Miner's cumulative damage theory, the fatigue damage D caused by the blade pin per flight hour in the blade folding mode without fretting is calculated. 1- The fatigue damage D caused by the blade pin per flight hour in the blade folding mode with fretting wear 1+ .

[0015] Further, in step 2, the process is as follows:

[0016] When tethered, the helicopter rotor blades are constrained at both ends, one end is fixed to the docking surface by a blade pin, and the other end is clamped on the fixture.

[0017] According to the maximum vertical overload F during mooring g The tension F generated by the clamping end restricting the blade from pulling down 0 , calculate the restraint bending moment M on the blade interface y ′;

[0018] Since the clamping methods of each blade in helicopter air transportation are different, the restraint bending moment of each blade at the joint surface is calculated separately and the maximum restraint bending moment M is obtained. y ' ,max ;

[0019] According to the maximum restraint bending moment M y ' ,max Calculate the maximum extrusion load F on the blade pin 2 ;

[0020] Considering shear and bending deformation, calculate the maximum local stress σ on the blade pin during mooring overload max2 ;

[0021] Calculate the number of cycles N of low-cycle life without fretting erosion of the blade pin under moored overload 2- , and low cycle life cycle number N with fretting 2+ ;

[0022] Calculate the fatigue damage D caused by the blade pin per flight hour in the tethered overload mode without fretting wear 2 The fatigue damage D caused by the blade pin per flight hour in the blade tethered overload mode with fretting wear2+ .

[0023] Further, in step three, the process is as follows:

[0024] Based on Miner's cumulative damage theory, the fatigue life T of helicopter blade pins considering folding and tethered overloads is calculated in the non-fretting wear mode. - ;

[0025] Based on Miner's cumulative damage theory, the fatigue life T of helicopter blade pins considering folding and tethered overloads is calculated in the fretting wear mode. + ;

[0026] The fatigue life of helicopter blade pin considering folding and tethered overload is T - and T + The minimum value in .

[0027] Furthermore, the maximum extrusion load F 1 The formula is as follows:

[0028]

[0029] Where m is the mass of the blade; g = 9.8 m / s 2 ;M y,wind is the flapping bending moment at the blade pin caused by land wind load, determined according to the test results; F manual is the manual control load, generally 100N; L is the distance between the blade tip and the blade root, l 1 is the distance between the blade's center of gravity and the blade's root; h is the distance between the upper and lower lugs;

[0030] The maximum local stress σ borne by the blade pin max1 The formula is as follows:

[0031]

[0032] Where D and d are the outer diameter and inner diameter of the blade pin respectively;

[0033] The formula for the low-cycle equivalent dynamic stress of the blade pin is as follows:

[0034]

[0035] In the formula, R 0.2 is the yield strength of the material;

[0036] Fatigue damage D caused by blade pin per flight hour in blade folding without fretting abrasion mode 1- The formula is as follows:

[0037]

[0038] Fatigue damage D caused by blade pin per flight hour in blade folding with fretting mode 1+ The formula is as follows:

[0039]

[0040] T flight is the service life of the rotor blade pin in flight hours, n 1 is the number of times the paddle can be folded during its service life.

[0041] Furthermore, the restrained bending moment M on the blade interface is y The formula is as follows:

[0042]

[0043] l 2 The distance between the blade's center of gravity and the blade's tip clamping fixture during air transport;

[0044] Maximum extrusion load F on the blade pin 2 The formula is as follows:

[0045]

[0046] The maximum local stress σ on the blade pin during mooring overload max2 The formula is as follows:

[0047]

[0048] Fatigue damage D caused by the blade pin per flight hour in the tethered overload mode without fretting wear 2- The formula is as follows:

[0049]

[0050] Fatigue damage of blade pin per flight hour in tethered overload with fretting mode D 2+ The formula is as follows:

[0051]

[0052] Furthermore, in the non-fretting wear mode, the fatigue life T of the helicopter blade pin considering the folding and tethered overload is - The formula is as follows:

[0053]

[0054] Fatigue life T of helicopter blade pin considering folding and tethered overloads in fretting mode + The formula is as follows:

[0055]

[0056] In summary, the beneficial effects of the present invention are as follows:

[0057] Considering the influence of blade deadweight, land wind load and manual operation during blade folding, an analysis method for the maximum extrusion load and maximum local stress of the blade pin is proposed. Considering the influence of the maximum vertical overload and the blade pull under the clamping end during the mooring process, an analysis method for the maximum extrusion load and maximum local stress of the blade pin is proposed. Considering the influence of fretting abrasion without and with fretting abrasion, an evaluation method for fatigue damage and fatigue life of helicopter blade pins is proposed, which provides assistance for the design of key components of helicopter rotors under folding and mooring overloads.

[0058] 1) Considering the influence of blade deadweight, land wind load and manual operation during blade folding, the maximum extrusion load F that the blade pin can bear is proposed. 1 The maximum local stress σ borne by the blade pin during blade folding is given by max1 Analytical methods.

[0059] 2) Considering the influence of the maximum vertical overload and the downward pull of the blade at the clamping end during the mooring process, the maximum extrusion load F that the blade pin bears is proposed. 2 The maximum local stress σ borne by the blade pin during mooring overload is given by max2 Analytical methods.

[0060] 3) The equal life model is used for load correction. According to the Stromyer model of fatigue performance and Miner cumulative damage theory, an evaluation method for fatigue damage and life of helicopter blade pins with and without fretting corrosion and considering folding and tethered overloads is proposed, which provides assistance for the design of key components of helicopter rotors under folding and tethered overloads. BRIEF DESCRIPTION OF THE DRAWINGS

[0061] Figure 1 Indicates blade folding and moored overload.

[0062] Figure 2 This is an illustration of the blade pin load.

[0063] Figure 3 The technical flow chart of the method described in the present invention.

[0064] The symbols in the figure are explained as follows:

[0065] Figure 1 L in the formula is the distance between the blade tip and the blade root. 1 is the distance between the blade center of gravity and the blade root, l 2 It is the distance between the blade's center of gravity and the blade tip's clamping fixture during air transport.

[0066] Figure 2 The h in the figure is the distance between the upper and lower ears of the blade pin connecting the ears. DETAILED DESCRIPTION

[0067] A method for evaluating fatigue life of helicopter blade pins considering folding and tethered overloads is provided, and the specific steps of the method are as follows.

[0068] Step 1: Blade pin load analysis and fatigue damage assessment during blade folding

[0069] In the blade folding process, considering that the blade pin bears the most dangerous bending moment caused by the blade's deadweight, land wind load and manual operation, the maximum extrusion load F of the blade pin is 1 It can be expressed as:

[0070]

[0071] Where m is the mass of the blade; g = 9.8 m / s 2 ;M y,wind is the flapping bending moment at the blade pin caused by land wind load, determined according to the test results; F manual is the manual control load, generally 100N; L is the distance between the blade tip and the blade root, l 1 is the distance from the blade's center of gravity to the blade's root (e.g. Figure 1 As shown); h is the distance between the upper and lower ears (as shown Figure 2 shown).

[0072] Considering shear and bending deformation, the maximum local stress σ borne by the blade pin during blade folding is max1 It can be expressed as:

[0073]

[0074] Where D and d are the outer diameter and inner diameter of the blade pin, respectively.

[0075] According to formula (2), the equal life model is used for correction, and the low-cycle equivalent dynamic stress of the blade pin during blade folding is:

[0076]

[0077] In the formula, R 0.2 is the yield strength of the material.

[0078] According to formula (3), the low-cycle equivalent dynamic stress σ of the blade pin during blade folding is obtained aeq1 , using the fatigue performance Stromyer model, combined with the material performance parameters of the blade pin, considering the blade pin without fretting and with fretting two failure modes for analysis, the low cycle life cycle number N of the blade pin without fretting can be calculated. 1- , and low cycle life cycle number N with fretting 1+ .

[0079] The service life of the rotor blade pin is T flight (Unit: flight hours), the number of times the blades can be folded during the service life is required to be n 1 Based on Miner's cumulative damage theory, in the non-fretting abrasion mode, the fatigue damage D caused by the blade pin per flight hour in the blade folding process is 1- for:

[0080]

[0081] Fatigue damage D caused by blade pin per flight hour in blade folding with fretting mode 1+ for:

[0082]

[0083] Step 2: Blade pin load analysis and fatigue damage assessment during moored overload

[0084] During the mooring process, the helicopter rotor blades are constrained at both ends, one end is fixed to the docking surface by a blade pin, and the other end is clamped on the fixture, such as Figure 1 The maximum vertical overload during the mooring process is F g , considering the tension F generated by the clamping end restricting the blade from pulling down 0 In the moored overload, the restraint bending moment M on the blade interface is y ′ can be expressed as:

[0085]

[0086] In the formula, l 2 The distance between the blade center of gravity and the blade tip clamping fixture during air transport (e.g. Figure 1 shown).

[0087] The helicopter rotor has n blades in total. Each blade has a different clamping position during the tethering process, so the corresponding l 1 and l 2 It can be seen from formula (6) that the bending moment M at the blade interface is y ′ is l 1 and l 2 The maximum restraint bending moment M of the blade joint surface is calculated by considering the different clamping methods of each blade in helicopter air transportation. y ' ,max On this basis, the maximum extrusion load F borne by the blade pin during mooring overload is 2 It can be expressed as:

[0088] M y ' ，max =max(M y ' ，1,M y ' ，2 ,M y ' ，3 ,…,M y ' ，n ) (7)

[0089]

[0090] Considering shear and bending deformation, the maximum local stress σ borne by the blade pin during mooring overload is max2 It can be expressed as:

[0091]

[0092] According to formula (3), the low-cycle equivalent dynamic stress of the blade pin in the moored overload is calculated. Similarly, the low-cycle life cycle number N of the blade pin without fretting wear in the moored overload can be calculated: 2- , and low cycle life cycle number N with fretting 2+ .

[0093] The number of moored overloads required during the service life of the rotor blade pin is n. 2 , then the fatigue damage D caused by the blade pin per flight hour in the moored overload mode without fretting wear is 2- for:

[0094]

[0095] Fatigue damage D caused by blade pin per flight hour in the tethered overload mode with fretting 2+ for:

[0096]

[0097] Step 3: Fatigue life assessment of helicopter blade pins considering folding and tethered overloads

[0098] According to equations (4) and (10), based on Miner cumulative damage theory, the fatigue life T of helicopter blade pin considering folding and tethered overload in the non-fretting wear mode is - (in flight hours) are:

[0099]

[0100] According to equations (5) and (11), the fatigue life T of helicopter blade pin considering folding and tethered overload in the fretting wear mode is + (in flight hours) are:

[0101]

[0102] Considering the influence of different fretting wear modes on fatigue life, according to equations (12) and (13), we take T = min (T - ,T + ) is the fatigue life of helicopter blade pin considering folding and tethered overloads.

[0103] Suggest an example

[0104] Use a specific scenario and value to go through all the calculation processes

[0105] When the blade is folded, the flapping bending moment M at the blade pin caused by the land wind load is taken y,wind is 1566Nm, the manual operating load is 144N, L is 7.3m, l 1 is 3.1m, h is 0.1m, then the maximum extrusion load F of the blade pin is 1 It is 40318N.

[0106] Taking the outer diameter D of the blade pin as 0.04m and the inner diameter d as 0.03m, the maximum local stress σ borne by the blade pin during blade folding is max1 It is 178MPa.

[0107] Take the yield strength R of the material 0.2 is 1030MPa, then the low-cycle equivalent dynamic stress σ of the blade pin during blade folding is aeq1 It is 90MPa.

[0108] Using the fatigue performance Stromyer model, the safe fatigue limit of the blade pin without fretting wear mode is taken as 170MPa, and the low cycle life cycle number N of the blade pin without fretting wear in folding is calculated. 1- =∞ times; the safe fatigue limit of the blade pin with fretting wear mode is taken as 56MPa, and the low cycle life cycle number N with fretting wear is calculated 1+ It is 845575 times.

[0109] Take the service life of the rotor blade pin, T flight 6000 flight hours, blades can be folded n times during life 1 is 1000 times. Based on Miner's cumulative damage theory, in the non-fretting abrasion mode, the fatigue damage D caused by the blade pin per flight hour in the blade folding is 1- =0; In the fretting abrasion mode, the fatigue damage D caused by the blade pin per flight hour in the blade folding 1+ 2.0×10 -7 .

[0110] During mooring overload, take the maximum vertical overload F during mooring g The pulling force F generated by the blade pulling down is 4000N. 0The rotor of a helicopter has five blades. Each blade is clamped at a different position during the tethering process. The distance between the blade center of gravity and the clamping fixture of each blade tip during air transport is l 2 See the table below. In the moored overload, calculate the restraint bending moment M on the butt joint of each blade. y ’The results are as follows.

[0111] Table 1 Calculation of the constraint bending moment of the blade interface during mooring

[0112]

[0113]

[0114] Considering the different clamping methods of each blade in helicopter air transportation, the maximum restraint bending moment M of the blade joint surface is calculated. y ' ,max is 5998Nm. On this basis, the maximum extrusion load F borne by the blade pin in the mooring overload is 2 It is 45097N.

[0115] Considering shear and bending deformation, the maximum local stress σ borne by the blade pin during mooring overload is max2 If the low-cycle equivalent dynamic stress of the blade pin in moored overload is 200 MPa, the low-cycle life cycle number N of the blade pin without fretting erosion in moored overload is calculated as 2- =∞ times, low cycle life cycle number N with fretting erosion 2+ It is 551283 times.

[0116] Take the number of moored overloads that can be performed during the service life of the rotor blade pin as n 2 If the number of times is 1000, then in the non-fretting abrasion mode, the fatigue damage D caused by the blade pin per flight hour in the tethered overload is 2- =0; In the fretting abrasion mode, the fatigue damage D caused by the blade pin per flight hour in the tethered overload 2+ 3.0×10 -7 .

[0117] Considering the fatigue damage of the blade pin caused by folding and tethering overload, based on Miner cumulative damage theory, the fatigue life T of the helicopter blade pin in the non-fretting wear mode is - 6000 flight hours; Under the fretting wear mode, the fatigue life of helicopter blade pin is T + That's 5,982 flight hours.

[0118] Considering the influence of different fretting abrasion modes, the fatigue life T of helicopter blade pin under folding and tethered overload is 5982 flight hours.

Claims

1. A method for evaluating the fatigue life of helicopter blade pins considering folding and tethering overloads, characterized in that: Here are the steps: Step 1: Evaluate the blade pin load and fatigue damage in the folded blade state; Step 2: Evaluate blade pin load and fatigue damage under moored overload; Step 3: Evaluate the blade pin fatigue life under folded state superimposed on moored overload.

2. The method according to claim 1, characterized in that: In step one, the process is as follows: The maximum extrusion load F1 of the blade pin is calculated based on the most dangerous bending moment caused by the blade's deadweight, land wind load and manual operation. Calculate the maximum local stress σ on the blade pin based on shear and bending deformations max1 ; The maximum local stress σ is calculated using the equal life model. max1 After correction, the low-cycle equivalent dynamic stress of the blade pin is obtained; According to the low-cycle equivalent dynamic stress σ of the blade pin aeq1 , the fatigue performance Stromyer model is used, combined with the material performance parameters of the blade pin, to calculate the low cycle life cycle number N of the blade pin without fretting wear in folding 1- , and low cycle life cycle number N with fretting 1+ ; Based on Miner's cumulative damage theory, the fatigue damage D caused by the blade pin per flight hour in the blade folding mode without fretting is calculated. 1- The fatigue damage D caused by the blade pin per flight hour in the blade folding mode with fretting wear 1+ .

3. The method according to claim 2, characterized in that: In step 2, the process is as follows: When tethered, the helicopter rotor blades are constrained at both ends, one end is fixed to the docking surface by a blade pin, and the other end is clamped on the fixture. According to the maximum vertical overload F during mooring g The restraint bending moment M on the blade joint surface is calculated by the tension F0 generated by the clamping end to limit the blade from pulling down. y ′; Since the clamping methods of each blade in helicopter air transportation are different, the restraint bending moment of each blade at the joint surface is calculated separately and the maximum restraint bending moment M is obtained. y ' ,max ; According to the maximum restraint bending moment M y ' ,max Calculate the maximum extrusion load F2 borne by the blade pin; Considering shear and bending deformation, calculate the maximum local stress σ on the blade pin during mooring overload max2 ; Calculate the number of cycles N of low-cycle life without fretting erosion of the blade pin under moored overload 2- , and low cycle life cycle number N with fretting 2+ ; Calculate the fatigue damage D2 caused by the blade pin per flight hour in the tethered overload mode without fretting and the fatigue damage D2 caused by the blade pin per flight hour in the tethered overload mode with fretting. 2+ .

4. The method according to claim 3, characterized in that: In step three, the process is as follows: Based on Miner's cumulative damage theory, the fatigue life T of helicopter blade pins considering folding and tethered overloads is calculated in the non-fretting wear mode. - ; Based on Miner's cumulative damage theory, the fatigue life T of helicopter blade pins considering folding and tethered overloads is calculated in the fretting wear mode. + ; The fatigue life of helicopter blade pin considering folding and tethered overload is T - and T + The minimum value in .

5. The method according to claim 4, characterized in that: The formula for the maximum extrusion load F1 of the blade pin is as follows: Where m is the mass of the blade; g = 9.8 m / s 2 ;M y,wind is the flapping bending moment at the blade pin caused by land wind load, determined according to the test results; F manual is the artificial control load, generally 100N; L is the distance from the blade tip to the blade root, l1 is the distance from the blade center of gravity to the blade root; h is the distance between the upper and lower ears; The maximum local stress σ borne by the blade pin max1 The formula is as follows: Where D and d are the outer diameter and inner diameter of the blade pin respectively; The formula for the low-cycle equivalent dynamic stress of the blade pin is as follows: In the formula, R 0.2 is the yield strength of the material; Fatigue damage D caused by blade pin per flight hour in blade folding without fretting abrasion mode 1- The formula is as follows: Fatigue damage D caused by blade pin per flight hour in blade folding with fretting mode 1+ The formula is as follows: T flight is the service life of the rotor blade pin in flight hours, and n1 is the number of times the blade can be folded during its service life.

6. The method according to claim 5, characterized in that: The restrained bending moment M on the blade joint surface y The formula is as follows: l2 is the distance between the blade center of gravity and the blade tip clamping fixture during air transport; The formula for the maximum extrusion load F2 borne by the blade pin is as follows: The maximum local stress σ on the blade pin during mooring overload max2 The formula is as follows: Fatigue damage D caused by the blade pin per flight hour in the tethered overload mode without fretting wear 2- The formula is as follows: Fatigue damage of blade pin per flight hour in tethered overload with fretting mode D 2+ The formula is as follows:

7. The method according to claim 6, characterized in that: Fatigue life T of helicopter blade pin considering folding and tethered overloads in the non-fretting wear mode - The formula is as follows:

8. The method according to claim 7, characterized in that: Fatigue life T of helicopter blade pin considering folding and tethered overloads in fretting mode + The formula is as follows:

Images