A Note on Helicopter Rotor-blade Fatigue-crack Propagation Rates Under Equivalent-lifetime Fatigue Loadings PDF Download

Author: John F. Ward
Publisher:
ISBN:
Category : Loads (Mechanics)
Languages : en
Pages : 14

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Author: John F. Ward
Publisher:
ISBN:
Category :
Languages : en
Pages : 1

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Book Description
Results are given for a brief investigation of the relative rates of fatigue-crack propagation obtained in helicopter-rotor-blade fatigue tests in which simplified, equivalent-total-lifetime, fatigue-test loadings at zero mean load are used to simulate a flight fatigue loading that includes a mean tension load. The conventional equivalent-lifetime loadings do not give equivalent rates of crack propagation. For typical rotor-blade loadings, which included large mean tension load, the general trend was toward greatly reduced rates of fatigue-crack propagation under the equivalent-lifetime loadings provided a nonconservative basis for establishing rotor-blade inspection intervals. (Author).

Author: Margery A. Dean
Publisher:
ISBN:
Category : Airframes
Languages : en
Pages : 38

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The life prediction of aerospace structures subjected to service loading conditions can be complex and expensive. Fortunately, many flight-by- flight loadings can be analyzed as constant amplitude loading if the variability in crack growth rate is low. A test program was designed such that the stress intensity factor (Kmax) was held constant by uniformly reducing the load as the fatigue crack grew. With a constant Kmax, the crack growth rate remains constant throughout the life of the test. The baseline fatigue crack growth rate variability was established for the 7075-T6 center-cracked panels under a flight-by-flight load history. The effect of introducing overloads to the baseline flight history on the crack growth rate variability was reported. The results indicate that the delay region for the applied 114% overloads was predicted by Irwin's plastic zone model and was equal to 0.0382 inches which was equal to 225 flights. The delay behavior was consistent for the several overload conditions that were studied. When the occurrence of overload increased, the fatigue crack growth decreased. The standard error of estimate remained relatively constant, suggesting that a block approach to life prediction is feasible for flight-by-flight load history.

Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 36

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Author: Harry T. Jensen
Publisher:
ISBN:
Category : Helicopters
Languages : en
Pages : 112

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Author: M. J. RICH
Publisher:
ISBN:
Category :
Languages : en
Pages : 9

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Author: K. B. Amer
Publisher:
ISBN:
Category : Aircraft
Languages : en
Pages : 19

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A technique is described whereby helicopter rotor blade service life can be substantiated by combining laboratory fatigue tests with cyclic loading applied during conventional ground endurance tests of rotor control and power-transmission systems. The method for applying cyclic loading involves the use of excitation panels, which consist of an array of stationary panels properly spaced and located close beneath the rotor to excite a natural flapwise bending mode. During the investigation, blade cyclic moments were excited that were ten per cent higher than the maximum cyclic bending moment measured in flight on the same rotor blade. The level of fatigue loading could be controlled by adjusting the vertical distance between the panels and rotor. The more rapid application of load (third-mode excitation at nine per revolution instead of the conventional one-per-revolution occurrence of peak inflight loads), coupled with sustained application of damage (100 per cent of the time for panels instead of approximately 16 per cent of the time for a normal flight spectrum), permits the excitation panels to impose damage approximately 56 times faster than experienced in flight. Tests are now being conducted to determine generalized design rules for application of excitation panels to rotors of any number of blades of any size.

Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 35

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The rate of crack propagation from an induced defect in a metal Bell Helicopter 540 main rotor blade was investigated. A controlled crack front was introduced into the top surface of the blade spar. Fatigue testing was conducted at maximum level flight loads and the crack growth was monitored. Experimental data was then comapred with analytical predictions to measure the ability to predict crack growth characteristics. (Author).

Author:
Publisher:
ISBN:
Category :
Languages : en
Pages : 114

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Author: M. Giglio
Publisher:
ISBN:
Category : Crack
Languages : en
Pages : 10

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A comparison between two types of helicopter fuselage panels with a traditional design of skin and stringers was carried out. Panels were made with the same geometrical shape and dimensions but with two different aluminium alloys for the skin: Conventional 2024 and 8090 Al-Li alloy. Two different crack locations were considered to be taken into account for the damage tolerant behavior of the two panel typologies: A crack into the skin between two stringers and a crack into the skin under the stringer after the removal of a complete section of the stringer. The crack growth was monitored during the fatigue load application, and a finite element (FE) model was utilized in order to calculate the fracture mechanics parameters. The experimental data (crack growth rate da/dN) and the FE results (stress intensity range ?KI) were checked with the crack propagation material data, with good agreement. According to these results, it is possible to compare the damage tolerant behaviors of the panels made by the two different materials, with the purpose of optimizing crack propagation behavior versus weight in this fundamental helicopter component.