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This study develops a rapid algorithm coupled with the finite element method to predict the fatigue crack propagation process and select the enhancement factor for the equivalent random load spectrum of accelerated fatigue tests. The proposed algorithm is validated by several fatigue tests of an aluminum alloy under the accelerated random load spectra. In the validation process, two kinds of panels with different geometries and sizes are used to calculate the stress intensity factor, critical crack length, and crack propagation life. The simulated and experimental findings indicate that when the aluminum alloy is in a low plasticity state, the crack propagation life exhibits a linear relationship with the acceleration factor. When the aluminum alloy is in a high plasticity state, this study proposes an empirical formula to calculate the equivalent stress intensity factor and crack propagation life. The normalized empirical formula is independent of the geometry and size of different samples, although the fracture processes are different in the two kinds of panels used in our study. Overall, the numerical method proposed in this paper can be applied to predict the fatigue crack propagation life for the random spectrum of large samples based on the results of the simulated accelerated crack propagation process and the accelerated fatigue tests of small samples to reduce the cost and time of the testing. [ABSTRACT FROM AUTHOR]

This study investigates the thermal stress and bolt load distribution in a hybrid panel structure of an aircraft mechanical joint under extreme temperatures. The hybrid panel structure comprises two aluminum alloy splices, six T-shaped composite stringers, and two composite skins, secured together with 96 bolts. This study analyzed the strain induced by thermal stress on composite materials and metals within the structure across temperatures, employing temperature environment tests ranging from room temperature to −54 °C, alongside a carrying capacity test at −54 °C. Furthermore, a three-dimensional simulation model of the panel structure was developed, incorporating considerations for contact, metal elastoplasticity, and the progressive damage failure of composite materials. This model facilitated the determination of thermal stress and bolt load distribution patterns. The results indicate a strong consistency between the finite element analysis outcomes and the experimental data. Temperature variations exacerbate the uneven distribution of bolt loads, concentrating the load near the edges of the hybrid structure while diminishing it in the center. The bolt load distribution parallel to the mechanical load direction forms an "M" shape, with a maximum load magnitude of approximately 31 kN. Perpendicular to the mechanical load, the bolt load undergoes significant changes, especially at the edges, reaching a maximum of about 20 kN, which warrants attention. The bolt-load distribution of the structure with the increase in mechanical load at −54 °C tends to be consistent with that at room temperature. [ABSTRACT FROM AUTHOR]

Experimental and simulation methods were used to predict the fatigue life of the connecting part of the aircraft beam structure. First of all,based on the statistical analysis,the marker load was chosen. The experimental results show that the marker load method can quantitatively record the initiation and propagation of cracks. Meanwhile this paper established a finite element model of the three-dimensional connection part of aircraft beam structure,with soft-soft contact and pre-stressing force modal. The errors between the results of the finite element model and the static experimental results are within 10%. Finally,with the help of the local stress-strain method and the software FEMFAT,the fatigue life of the test specimens was successfully predicted. The two results were compared to each other. It is find out that the crack formation time generated by the local stressstrain method is radical,while the FEMFAT software can obtain a conservative crack formation time which is consistent with the test results.

The evaluation of structural crashworthiness is important for the development of civil aircraft. In this study, cargo luggage was considered in the vertical drop simulation of an aircraft mid-fuselage section. First, quasi-static compression tests on three types of luggage made of different materials were conducted. Then, a finite element model (FEM) of the cargo luggage was developed with a crushable foam material in LS-PrePost4.5 software. The finite element analysis results of the luggage made of different materials were consistent with the corresponding compression experiments. Secondly, the FEM of the mid-fuselage section was established without cargo luggage. The simulated displacement and acceleration of the fuselage section were consistent with the test results. Finally, the influences of cargo luggage on the fuselage vertical drop response were studied with the FEM considering both the fuselage structure and cargo luggage. Compared with the responses of the fuselage without cargo luggage, the cargo luggage could reduce the deformation of the cargo bay and maintain the integrity of the passenger living space in the cabin at the vertical velocity of 6 m/s, even though the initial kinetic energy was higher. [ABSTRACT FROM AUTHOR]

3D woven composites have been widely used in aerospace field. It is of great engineering value to study the damage mechanism of C/C woven composites with openings under tension and compression. A mosaic finite element model of C/C woven composite with opening-hole is established, and the progressive damage analysis under tensile and compressive loads is performed to study the mesoscopic damage model of opening plate. In the finite element modeling, a mesoscopic finite element model is established based on the mesostructure of the material in the perforated area of the open-hole plate, and the rest is equivalent to a uniform model. The failure criterion based on meso-mechanics is used to predict the ultimate load and strength of open-hole plates under tensile and compressive loads, and the damage process is studied. Based on the simulation, the damage initiation occurred in many areas of the warp yarn due to the stress concentration in the contact area between the binder and the warp yarn. And the damage area of the open-hole plate exhibited an ‘X’ shape expansion both under tensile and compressive loads.

In this research, the design of a robust curved-line path-following control system for fixed-wing unmanned aerial vehicles (FWUAVs) affected by uncertainties on the latitude plane is studied. This is undertaken to enhance closed-loop system robustness under unknown uncertainties and derive the control surface deflection angle directly used to control FWUAVs, which has rarely been studied in previous works. The system is formed through the mass center position control (MCPC) and yaw angle control (YAC) subsystems. In the MCPC, the desired yaw angle, which is treated as the reference signal for the YAC subsystem, is calculated analytically using path-following errors, current flow angles, and the yaw angle. In the YAC, a disturbance estimator is designed to estimate uncertainties such as nonlinearities, couplings, time variations, model parameter perturbations, and unmodeled dynamics. Predictive functional controllers are designed to target nominal systems in the absence of uncertainties, such that the estimations of the uncertainties can be incorporated through feedback for closed-loop system robustness enhancement. The simulation results show that higher path-following precision and stronger robustness for the FWUAVs based on the proposed approach can be achieved using only rough model parameters compared with the conventional nonlinear dynamic inversion, which requires detailed model information. [ABSTRACT FROM AUTHOR]

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To evaluate the multi-element damages that occurred during the aircraft services, a random failure model was established based on element fatigue characteristics, and the average behavior of widespread fatigue damages were calculated based on element failure life to determine the time for aircraft inspection and repair. The validity of the model was verified by a five-frame fuselage section example, and the influences of simulation times and distribution model on the repair time of multi-element damage structures were studied. According to the fatigue test results of the seven-frame fuselage section, multi-element damage assessment was performed, the structural inspection and repair time were given. The results show that the evaluation results are in good agreement with the calculation example data, and normal regression may better simulate structural failure characteristics. [ABSTRACT FROM AUTHOR]

The failure types and ultimate loads for eight carbon-epoxy laminate specimens with a central circular hole subjected to tensile load were tested experimentally and simulated using two different progressive failure analysis (PFA) methodologies. The first model used a lamina level modeling based on the Hashin criterion and the Camanho stiffness degradation theory to predict the damage of the fiber and matrix. The second model implemented a micromechanical analysis technique coined the generalized method of cells (GMC), where the 3D Tsai–Hill failure criterion was used to govern matrix failure, and the fiber failure was dictated by the maximum stress criterion. The progressive failure methodology was implemented using the UMAT subroutine within the ABAQUS/implicit solver. Results of load versus displacement and failure types from the two different models were compared against experimental data for the open hole laminates subjected to tensile displacement load. The results obtained from the numerical simulation and experiments showed good agreement. Failure paths and accurate damage contours for the tested specimens were also predicted. [ABSTRACT FROM AUTHOR]

In order to enhance the safety of the catapult launch of the carrier-based aircraft, the catapult launch multi- body dynamic model is established aiming at the problem of off - center catapult launch. The whole catapult process including four stages which are buffering, tensioning, releasing and taxiing is taken into consideration and the body dynamics of the off-center catapult during each stage is analyzed. The catapult launch dynamic differences between the conditions only considering taxiing and that considering four stages are compared, and the effects of the different initial off - center distances considering four stages on the attitude, landing gear load and acceleration of the carrier - based aircraft during catapult launch are discussed. The results show that only considering taxiing may underestimate the dynamics of the carrier-based aircraft substantially. When taking four stages into consideration, the initial off-center distance has small influence on the aircraft dynamic characteristics during buffering and tensioning but has larger influence on that during releasing and taxiing. The increase of the off-center distance will cause the enhancement of the aircraft rolling and yawing, which may lead to the load difference between the left and right landing gears and the increase of the aircraft lateral acceleration. The establishment and simulation of the catapult launch multi - body dynamic model founded on buffering, tensioning, releasing and taxiing provide reference for the carrier-based aircraft design and analysis of the catapult launch dynamics. [ABSTRACT FROM AUTHOR]