Helicopter blade pin fatigue life assessment method considering folding and tethering overload

Abstract

The application belongs to the technical field of structural strength design, and discloses a helicopter blade pin fatigue life evaluation method considering folding and tethering overload, and the steps are as follows: step one, evaluating the blade pin load and fatigue damage under the folding state of the blade; step two, evaluating the blade pin load and fatigue damage under the tethering overload; and step three, evaluating the blade pin fatigue life under the folding state superimposed with the tethering 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 evolution of helicopter use has necessitated the need for folding and tethering overloads. Blade pins, critical components of helicopter rotor systems, are crucial for flight safety. Evaluating the effects of folding and tethering overloads on blade pin fatigue life is crucial. Currently, helicopter blade pins are typically evaluated using safety fatigue design methods, considering both operating modes without and with fretting wear. However, mature technology is lacking to analyze the effects of folding and tethering overloads on blade pin fatigue life, and further research is needed. Summary of the Invention

[0003] Purpose of the Invention: This invention considers the effects 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. It also considers the effects of maximum vertical overload and blade pull-down force at the clamping end during mooring, and proposes an analysis method for the maximum extrusion load and maximum local stress of the blade pin. Furthermore, considering the effects of both fretting and non-fretting, a method for assessing fatigue damage and fatigue life of helicopter blade pins is proposed, providing assistance for the design of key helicopter rotor components under folding and mooring overloads.

[0004] Technical Solution

[0005] A fatigue life assessment method for helicopter blade pins considering folding and tethering overloads is proposed. The steps are 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 the folded state and moored overload.

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

[0010] 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;

[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. max1 After correction, the blade pin low-cycle equivalent dynamic stress is obtained;

[0013] According to the low cycle equivalent stress σ aeq1 of the blade pin, using the fatigue performance Stromyer model, combining the material performance parameters of the blade pin, the low cycle life cycle number N 1- of the blade pin in the folding without fretting wear and the low cycle life cycle number N 1+ of the blade pin with fretting wear are calculated.

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

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

[0016] When tethered, the helicopter rotor blade is constrained at both ends, one end is fixed on the butt joint surface through the blade pin, and the other end is clamped on the tool clamp,

[0017] According to the maximum vertical overload load F g during the tethering process and the tensile force F0 generated by the clamping end limiting the blade to be pulled down, the constraint bending moment M y ′ borne by the blade butt joint surface is calculated.

[0018] Since the clamping methods of the blades in the helicopter air transport are different, the constraint bending moment borne by the butt joint surface of each blade is calculated respectively, and the maximum constraint bending moment M y ′ ,max is obtained.

[0019] The maximum extrusion load F2 borne by the blade pin is calculated according to the maximum constraint bending moment M y ′ ,max ;

[0020] Considering the shear and bending deformation, the maximum local stress σ max2 borne by the blade pin in the tethering overload is calculated.

[0021] The low cycle life cycle number N 2- of the blade pin without fretting wear and the low cycle life cycle number N 2+ of the blade pin with fretting wear under the tethering overload are calculated.

[0022] The fatigue damage D2 caused by the blade pin in the tethering overload per flight hour without fretting wear and the fatigue damage D caused by the blade pin in the tethering overload per flight hour with fretting wear are calculated. 2+

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

[0024] Fatigue life of the blade pin T considering the folding and tethering overload is calculated based on Miner cumulative damage theory under the micro-oscillation abrasion mode. - ;

[0025] Fatigue life of the blade pin T considering the folding and tethering overload is calculated based on Miner cumulative damage theory under the micro-oscillation abrasion mode. + ;

[0026] Fatigue life of the blade pin T considering the folding and tethering overload is T - and T + .

[0027] Further, the maximum extrusion load F1 of the blade pin is as follows:

[0028]

[0029] In the formula, m is the mass of the blade; g = 9.8 m / s 2 ; M y,wind is the flap bending moment at the blade pin caused by the land wind load, which is determined according to the test results; F manual is the artificial control load, generally taken as 100 N; L is the distance from the blade tip to the blade root, and l1 is the distance from the blade gravity center to the blade root; h is the distance between the upper and lower ears;

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

[0031]

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

[0033] 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] Under the micro-oscillation abrasion mode, the fatigue damage D 1- of the blade pin caused by the blade folding per flight hour is as follows:

[0037]

[0038] Under the micro-oscillation abrasion mode, the fatigue damage D 1+ of the blade pin caused by the blade folding per flight hour is as follows:

[0039]

[0040] T flight The service life of the rotor blade pin is in flight hours, and n1 is the number of times of folding the blade during the service life.

[0041] Further, the constraint bending moment M y The formula is as follows:

[0042]

[0043] L2 is the distance from the center of gravity of the blade to the clamping tool of the blade tip in air transportation;

[0044] The maximum extrusion load F2 of the blade pin is as follows:

[0045]

[0046] The maximum local stress σ of the blade pin in the tethered overload is as follows: max2 The formula is as follows:

[0047]

[0048] The fatigue damage D of the blade pin in the tethered overload per flight hour in the micro-attrition abrasion mode is as follows: 2- The formula is as follows:

[0049]

[0050] The fatigue damage D of the blade pin in the tethered overload per flight hour in the micro-attrition abrasion mode is as follows: 2+ The formula is as follows:

[0051]

[0052] Further, in the micro-attrition abrasion mode, the fatigue life T of the helicopter blade pin considering folding and tethered overload is as follows: - The formula is as follows:

[0053]

[0054] In the micro-attrition abrasion mode, the fatigue life T of the helicopter blade pin considering folding and tethered overload is as follows: + The formula is as follows:

[0055]

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

[0057] The present application discloses a method for evaluating fatigue life of a helicopter blade pin under folding and tethering overloads.

[0058] 1) A method for representing the maximum extrusion load F1 of the blade pin is proposed, and the maximum local stress σ max1 of the blade pin under folding overloads is given.

[0059] 2) A method for representing the maximum extrusion load F2 of the blade pin is proposed, and the maximum local stress σ max2 of the blade pin under tethering overloads is given.

[0060] 3) A method for evaluating fatigue damage and life of the helicopter blade pin under folding and tethering overloads is proposed based on the Stromyer model and the Miner cumulative damage theory. BRIEF DESCRIPTION OF DRAWINGS

[0061] Figure 1 The folding and tethering overloads of the blade pin are shown.

[0062] Figure 2 The bearing of the blade pin is shown.

[0063] Figure 3 The technical flow chart of the method is shown.

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

[0065] Figure 1 L in the figure is the distance between the blade tip and the blade root, l1 is the distance between the blade center of gravity and the blade root, and l2 is the distance between the blade center of gravity and the blade tip clamping tool in the air.

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

[0067] A method for evaluating fatigue life of a helicopter blade pin under folding and tethering overloads is provided, and the specific steps are as follows.

[0068] Step one, blade folding pin load analysis and fatigue damage assessment

[0069] In the blade folding, considering the blade pin bearing blade weight, land wind load and the most dangerous bending moment caused by manual operation, the maximum extrusion load F1 of the blade pin can be expressed as:

[0070]

[0071] In the formula, m is the mass of the blade; g = 9.8 m / s 2 ; M y,wind is the flap bending moment at the blade pin caused by land wind load, which is determined according to the test results; F manual is the manual operation load, generally taken as 100 N; L is the distance from the blade tip to the blade root, and l1 is the distance from the blade center of gravity to the blade root (as shown in Figure 1 ); h is the distance between the upper and lower ears (as shown in Figure 2 ).

[0072] Considering shear and bending deformation, the maximum local stress σ max1 of the blade pin in the blade folding can be expressed as:

[0073]

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

[0075] According to formula (2), the low cycle equivalent dynamic stress of the blade pin in the blade folding is

[0076]

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

[0078] According to the low cycle equivalent dynamic stress σ aeq1 of the blade pin in the blade folding obtained by formula (3), using the fatigue performance Stromyer model, combined with the material performance parameters of the blade pin, considering the two failure modes of no fretting wear and fretting wear of the blade pin, the low cycle life cycle number N 1- of the blade pin in the blade folding without fretting wear and the low cycle life cycle number N 1+ with fretting wear can be calculated.

[0079] The service life of the rotor blade pin is T flight (unit: flight hours), and the number of times of blade folding required during the service life is n1. Based on the Miner cumulative damage theory, the fatigue damage D 1- caused by the blade folding of the blade pin per flight hour under the mode of no fretting wear is:

[0080]

[0081] Fatigue damage D caused by blade fold pin per flight hour in the micro-motion abrasion mode 1+ is:

[0082]

[0083] Step two, blade pin load analysis and fatigue damage evaluation in tethering overload

[0084] In the tethering process, the helicopter rotor blade is constrained at both ends, one end is fixed on the docking surface through the blade pin, and the other end is clamped on the tool clamp, as shown in Figure 1 . The maximum vertical overload load in the tethering process is F g , considering the influence of the tensile force F0 generated by the clamping end limiting the blade to be pulled down, then the constraint bending moment M y ′ borne by the blade docking surface in the tethering overload can be represented as:

[0085]

[0086] In the formula, l2 is the distance from the center of gravity of the blade to the clamping tool at the tip of the blade in flight (as shown in Figure 1 ).

[0087] The helicopter rotor has n blades, and each blade has a different clamping position in the tethering process, so the corresponding l1 and l2 are also different. As can be seen from formula (6), the bending moment M y ′ of the blade docking surface is a function of l1 and l2, and the maximum constraint bending moment M y ′ of the blade docking surface is calculated considering the different clamping methods of each blade in the helicopter flight ,max . On this basis, the maximum extrusion load F2 borne by the blade pin in the tethering overload can be represented 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 max2 borne by the blade pin in the tethering overload can be represented as:

[0091]

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

[0093] If the number of times of ground loop overload during the service life of the rotor blade pin is required to be n2, the fatigue damage D of the blade pin per flight hour in the ground loop overload in the mode without fretting wear is 2- :

[0094]

[0095] The fatigue damage D of the blade pin per flight hour in the ground loop overload in the mode with fretting wear is 2+ :

[0096]

[0097] Step three, fatigue life assessment of the helicopter blade pin considering folding and ground loop overload

[0098] According to formula (4) and formula (10), based on the Miner cumulative damage theory, the fatigue life T of the helicopter blade pin considering folding and ground loop overload in the mode without fretting wear (unit: flight hour) is - :

[0099]

[0100] According to formula (5) and formula (11), the fatigue life T of the helicopter blade pin considering folding and ground loop overload in the mode with fretting wear (unit: flight hour) is + :

[0101]

[0102] Considering the influence of different fretting wear modes on fatigue life, from formula (12) and formula (13), T = min(T - ,T + ) is the fatigue life of the helicopter blade pin considering folding and ground loop overload.

[0103] An example is suggested

[0104] A specific scene, numerical value, all the calculation process is walked through

[0105] In the blade folding, the flap bending moment M y,windis 1566Nm, the manual control load is 144N, L is 7.3m, l1 is 3.1m, h is 0.1m, then the maximum extrusion load F1 of the blade pin 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 1030 MPa, 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 The flight time is 6000 hours. During the life span, the number of blade folding times n1 is 1000 times. Based on Miner's cumulative damage theory, in the non-fretting wear mode, the fatigue damage D caused by the blade pin during blade folding per flight hour 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] In mooring overload, take the maximum vertical overload F during mooring g The force F0 generated by pulling the blade downward is 4000N, and the force F0 generated by pulling the blade downward is 166N. Assume that the helicopter rotor has 5 blades, and each blade is clamped at a different position during the tethering process. The distance l2 between the blade center of gravity and the clamping fixture of each blade tip during air transportation is shown in the table below. Then, in the tethered overload, calculate the restraint bending moment M on the mating surface of each blade. y ’The results are as follows.

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

[0112]

[0113]

[0114] Considering the different clamping methods of each blade during helicopter air transportation, the maximum restraint bending moment M of the blade interface is calculated.y ,max is 5998 Nm, on this basis, the maximum extrusion load F2 of the blade pin under the tethered overload is 45097 N.

[0115] Considering the shear and bending deformation, the maximum local stress σ max2 is 200 MPa, then the low cycle equivalent dynamic stress of the blade pin under the tethered overload is 101 MPa. The low cycle life cycle number N 2- of the blade pin under the tethered overload without fretting wear is calculated to be ∞ times, the low cycle life cycle number N 2+ of the blade pin under the tethered overload with fretting wear is calculated to be 551283 times.

[0116] Taking the number of tethered overloads that the rotor blade pin can perform during the service life of the rotor blade pin as n2 is 1000 times, then under the mode without fretting wear, the fatigue damage D 2- of the blade pin under the tethered overload per flight hour is 0; under the mode with fretting wear, the fatigue damage D 2+ of the blade pin under the tethered overload per flight hour is 3.0×10 -7 .

[0117] Considering the fatigue damage of the blade pin caused by folding and tethered overload, based on the Miner cumulative damage theory, under the mode without fretting wear, the fatigue life T - of the helicopter blade pin is 6000 flight hours; under the mode with fretting wear, the fatigue life T + of the helicopter blade pin is 5982 flight hours.

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

Claims

1. A method for evaluating fatigue life of a helicopter blade pin taking into account folding and tethering overload, characterized by: The steps are as follows: Step one: evaluate the blade pin load and fatigue damage under the folded state of the blade, the process is as follows: The maximum extrusion load of the blade pin is calculated according to the most dangerous bending moment of the blade pin caused by the weight of the blade, the land wind load and the manual operation F 1; Maximum local stresses on the blade pin are calculated according to shear and bending deformations σ max1 ; The maximum local stress is calculated by using the equal life model σ max1 The low cycle equivalent dynamic stress of the pin is obtained by correcting the maximum local stress ; Based on the low cycle equivalent stress of the blade pin , the low cycle life cycle number of the blade pin in folding without fretting corrosion is calculated by using the fatigue performance Stromyer model combined with the material performance parameters of the blade pin , and the low cycle life cycle number of the blade pin with fretting corrosion ; Fatigue damage per flight hour caused by the blade fold in the blade pin under the micro-oscillation wear mode D 1- Fatigue damage per flight hour caused by the blade fold in the blade pin under the micro-oscillation wear mode D 1+ Step two: evaluate the blade pin load and fatigue damage under the tethered overload, the process is as follows: When tethered, the helicopter rotor blade is constrained at both ends, one end is fixed on the butt joint surface through the blade pin, the other end is clamped on the tool clamp, According to the maximum vertical overload load during mooring F g And the clamping end limits the pull force generated by the down-draw of the blade F 0, the constraint bending moment at the blade butt surface is calculated ; The constraint bending moment of each blade is calculated respectively according to the clamping mode of each blade in the helicopter air transportation, and the maximum constraint bending moment is obtained ; According to the maximum constraint bending moment Calculating the maximum extrusion load on the blade pin F 2; Consideration of shear and bending deformation, calculation of the maximum local stress on the pin of the blade in tethered overload σ max2 ; Low cycle life cycle number of a blade pin under tethered overload without fretting wear and with fretting wear ; Fatigue damage per flight hour to the blade pin in tethered overload with no fretting corrosion mode D 2- Fatigue damage per flight hour to the blade pin in tethered overload with fretting corrosion mode D 2+ ; Step three: evaluate the fatigue life of the blade pin under the folded state superimposed with the tethered overload, the process is as follows: Fatigue life of helicopter blade pin is calculated based on Miner cumulative damage theory, considering folding and tether overload in micro-motion abrasion mode T - ; Based on Miner cumulative damage theory, the fatigue life of helicopter blade pin is calculated under the mode of fretting wear and considering the folding and tether overload T + ; The fatigue life of a helicopter blade pin considering folding and tethered overload is T - and T + the minimum value.

2. The method of claim 1, wherein: Maximum extrusion load on paddle pin F 1The formula is as follows: wherein m M = mass of the blade; g = 9.8 ; M = the flapwise bending moment at the blade pin due to land wind load, determined from test results; M = the manual handling load, typically 100 N; L M = the distance from the blade tip to the blade root; l M = the distance from the blade center of gravity to the blade root; h M = the distance between the upper and lower ears; Maximum local stress on the blade pin σ max1 The formula is as follows: wherein D and d Dp and Di are the outer diameter and inner diameter of the paddle pin, respectively. The low-cycle equivalent dynamic stress formula of the blade pin is as follows: In the formula, Y is the yield strength of the material; Fatigue damage per flight hour caused by the blade fold mid-blade pin in the absence of fretting corrosion mode D 1 - The formula is as follows: Fatigue damage per flight hour caused by the blade fold mid-blade pin in the fretting wear mode D 1+ formula is as follows: Tflight life of the rotor blade pin, in flight hours, n 1 is the number of times the blade can be folded during its life.

3. The method of claim 2, wherein: The constraint bending moment that the leaf pair interface bears The formula is as follows: l 2 is the distance from the center of gravity of the blade to the tip of the blade in the air; Maximum extrusion load on the blade pin F 2The formula is as follows: Maximum local stress on pin of centerboard in bridle overload σ max2 The formula is as follows: Fatigue damage per flight hour caused by the tethered overload mid-paddle pin in the no-fretting wear mode D 2- The formula is as follows: Fatigue damage per flight hour caused by the blade pin in the tethered overload mode of micro-oscillation abrasion D 2+ The formula is as follows: 。 4. The method of claim 3, wherein: Fatigue life of helicopter blade pin under no-fretting wear mode considering folding and tether overload T - The formula is as follows: 。 5. The method of claim 4, wherein: Fatigue life of helicopter blade pin under fretting and tethered overload in fretting mode T + The formula is as follows: 。

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