Papers by Keyword: Fuselage

Abstract: In the case of catastrophic events, such as an emergency landing, the fuselage structure is demanded to absorb most of the impact energy preserving, at the same time, a survivable space for the passengers. Moreover, the increasing trend of using composites in the aerospace field is pushing the investigation on the passive safety capabilities of such structures in order to get compliance with regulations and crashworthiness requirements. This paper deals with the development of a numerical model, based on the explicit finite element (FE) method, aimed to investigate the energy absorption capability of a full-scale 95% composite made fuselage section of a civil aircraft. A vertical drop test, performed at the Italian Aerospace Research Centre (CIRA), carried out from a height of 14 feet so to achieve a ground contact velocity of 30 feet/s in according to the FAR/CS 25, has been used to assess the prediction capabilities of the developed FE method, allowing verifying the response under dynamic load condition and the energy absorption capabilities of the designed structure. An established finite element model could be used to define the reliable crashworthiness design strategy to improve the survival chance of the passengers in events such as the investigated one.

Abstract: Fatigue plays a significant role in the crack growth of the fuselage skin structures. In addition, the fuselage may suffer also from the corrosion damage, and the wear defects. The proper maintenance and scheduled test intervals can avoid the sudden skin failure. Therefore, the inspection interval has to be shortened. Nevertheless, the young machines may be also suffering from the unexpected skin rupture. The cracks are emanating from the rivets and the holes under cyclic loading. The stress concentration around the notch has an effective role under the effect of cyclic loading. The cracks propagate toward the high stressed area such as the notches or other crack locations. The propagation into a critical crack size is rather fast and causes a sudden aircraft fuselage cracking. Hence, the number of cycles to failure will be decreased dramatically. During the last decades, the fracture toughness, design, and the new alloying element have been enhanced. The previous fuselage failures show that the inspections against the cracking are recommended even after a few thousand of cycles. To prevent the crack extending, the crack arresting is recommended to use around the fuselage.

Abstract: This work includes design, fabrication and comparative vibration analysis of multi-layered filament wound composite cylinders of S-Glass epoxy and Basalt epoxy of 500mm length, 100mm diameter and 3mm thickness. The Vibration analyses were done separately on both the cylinders for various end conditions viz., free-free, fixed-free, and fixed-fixed. The natural frequencies of two materials are compared experimentally. Natural frequencies varies with end conditions, material properties, proportions of constituent materials present in the composites, geometrical parameters etc.The natural frequencies of all such parameters were determined using experimental setup which consists of a exciter, accelerometer and DAQ (DEWESOFT 7.0.5 software).

Abstract: The utilization of reconfigurable numerically controlled tooling is important for flexible fuselage assembly. An online programming and simulation system for flexible tooling is presented, which is composed of two modules. The online programming module generates NC code from an adjustment sequence and path, taking into consideration the process and device parameters. This paper details the second module, the visualized simulation module, which validates and optimizes the related adjustment process. This module covers three steps, those are, the creation of a virtual simulation environment based on the light-weight 3D model, the loading of the key feature points related to the adjustment and process information, and the adjustment simulation. The system can rapidly response to “one fixture with more models” and “one fixture with more states” deep flexibility demands for digital fuselage assembly, enable the flexible tooling to a have better capacity for independent adaptability. An example of flexible fuselage assembly process is used to illustrate the detail and advantages of the system. The three-step simulation strategy is proposed and detailed in this paper.

Abstract: The strength of a welded truss type fuselage of a light aircraft – named SAFAT 01 – is considered in this paper. The aircraft is a monoplane with high strut-braced wings configuration with flaps. The fuselage is of welded tubular steel fabric-covered construction. According to its production contract; the aircraft is fully produced and assembled in Sudan whereas the documentation is limited to technical side only with no information available about design procedures and calculations. This makes it difficult to further modify or upgrade the aircraft. The fuselage geometry has been modeled using a CAD program. The main dimensions have been obtained using 2-D drawings and the missing dimensions and data due to lack of documents were extracted experimentally from a built aircraft. Aerodynamic loads were determined using computational fluid dynamic program for the horizontal tail. Static structural analysis was conducted by finite element method (FEM) using fastened connection property between the tubes. The results were observed in three main parts; rear part which supports the vertical and horizontal stabilizers and the rear landing gear; the mid part which provides cantilever reaction for the wing and supports front landing gear, and the front part on which the engine is mounted. The Von Mises stresses, displacements and principal stresses were observed in the three parts and found acceptable except for a small region near the attachment between wing and fuselage. However, further experimental validation is needed. Presently, experimental and dynamic analyses are being conducted and the results will be published later.

Abstract: Experiments were conducted on thin delta wings to investigate, for subsonic flow, the effect of both privileged apex angle values and the wing-fuselage interactions on the aerodynamic characteristics, i.e. the distribution of the defect pressure on the extrados, the drag and the lift coefficients. For this purpose, several delta wing models of various apex angle (β = 75, 80 and 85°) were realized and tested without and with fuselages of cylindrical form, with diameters of 20 and 30 mm, downstream the apex and appropriately disposed on the extrados. The impact of the apex angle as well as the interaction on the defect pressure were specially considered along the apex vortices where the pressure defect is usually maximum. The above mentioned effects were investigated via the variations of the mean velocity in the wind tunnel and the incidence (attack) angle.

Abstract: This paper primarily presents the development and application of automation computational analysis techniques to determine the dynamic stress intensity factor for the damaged aircraft fuselage subjected to triangle blast load. A program based on automated procedure to simulate cracked fuselage is developed. It may create 3-dimention panel model using parameterization. The stress around the crack tips will be captured and the dynamic stress intensity factor can be obtained at every moment of the blast automatically. A typical curved panel model which consists of 7 frames and 8 stringers is calculated. The calculation results shown that the form of the dynamic SIF curve is similar to that of the triangle load curve while the peak point of dynamic SIF curve occurs a little later than that of the load curve due to the inertia effect. The longer the crack is, the more obvious the effect is. The peak SIF value of the crack under blast load is bigger than that under the static load for certain crack length. The longer the crack is, the bigger the difference between the dynamic peak SIF value and static SIF is. At the same time, the load time has effect on the dynamic SIF curve and its peak value. These results show good agreements with theoretical principles.

Abstract: Reducing structural weight is one of the major ways to improve aircraft performance. Lighter and/or stronger materials allow greater range and speed and may also contribute to reducing operational costs. Nowadays composite materials are widely used in “primary” structural components such as fuselage, for which contrasting requirements like lightness and structural strength are required, so particular attention is necessary during its design. In this paper a composite front bulkhead, subjected to ultimate pressure load, was examined. The front bulkhead is made of a composite skin, stiffened with seven vertical stiffeners linked through metallic fittings; the whole system is joined to the fuselage by rivets. A Finite Element model was established: the used elements were four nodes shells, simulating composite layers, and two nodes bar elements, simulating rivets; the structure was clamped and a pressure load was applied to the skin. A linear static stress analysis was performed to calculate strains in particular points in which strain gauges or rosettes are placed: the numerical results, compared with experimental ones, show a good degree of correlation. Stress calculations were performed in order to verify the front and rear bulkhead structural safety.

Abstract: This paper presents a practical approach for health monitoring of an aircraft fuselage using vibration measurements with piezoelectric transducers. For the test specimen, a fuselage element of an Airbus A320, the feasibility of health monitoring of stringers, frames and panels, is studied. The proposed structural health monitoring (SHM) system consists of three major components: vibration measurement, signal processing and damage diagnosis. By using applied piezoelectric patch actuators, flexural waves are excited which propagate along the monitored component. The used signal for the actuator is a broadband sine-sweep signal with an upper frequency of 100 kHz. The structural response is measured with piezoelectric patch transducers, which generate a charge proportional to the induced strain due to the vibration. These sensors are positioned on different locations around the exciting piezoelectric actuator. The sampled time signal is used to compute the complex frequency response functions between actuator and sensor. After measuring the vibration behaviour of the undamaged structure, the structure is artificially damaged by saw cuts to simulate cracks. After that the vibration behaviour is measured again in the same way. The measurements of the damaged and undamaged state are now evaluated with different mathematical algorithms, like root mean square value, norm, energy content, analysis of the phase and correlation coefficient. The aim is to find a method that needs very small computation effort and that even might be implemented by an analogue circuit. Finally the data are analysed to find appropriate damage metrics. It can be shown that some of the tested methods are a sensitive damage metric. With only a few actuators and sensors cracks in the outer panel with a length of 30-60 mm and in the stringer profile with a length of 10 mm can be found.

Abstract: In order to calculate the fracture parameters (Stress intensity factor) in a complicated 3- dimention aircraft model with damage in the aircraft panel, a new two steps global-local hierarchical analysis strategy is used. This paper primarily describes the development and application of advanced computational analysis techniques to determine stress intensity factors for the damaged panels based on the two steps hierarchical analysis strategy from global to 3-D local model, the bulging deformation of crack can be considered in the local model. A fracture parameter calculation programme based on automated global-local procedure to simulate cracked aircraft panel tests is developed for the hierarchical strategy. This programme may create models of two stages, transfer boundary conditions, calculate and obtain fracture parameter automatically. Finally, this paper presents some of the experimental data and the calculated fracture parameters are compared with the experimental results.