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Journal articles on the topic 'Aircraft panel'

Author: Grafiati

Published: 4 June 2021

Last updated: 1 February 2022

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1

Chenxi, L. I., H. U. Ying, and H. E. Liyan. "Exploration and optimization on the usage of micro-perforated panels as trim panels in commercial aircrafts." Noise Control Engineering Journal 68, no. 1 (2020): 87–100. http://dx.doi.org/10.3397/1/37687.

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Micro-perforated panels (MPPs), as an alternative to porous materials for sound absorption, have been commonly used in electronic industries and aircraft engines but are barely used in aircraft cabins. The effect of MPPs on the sound insulation and absorption properties of aircraft cabin panels has been investigated in this article. Theoretical modeling has been conducted on an aircraft cabin panel structure with a trim panel replaced by an MPP trim panel, using the transfer matrix method and the classic MPP theory. It is indicated by the theoretical results that, although the sound transmissi

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2

Zhang, Li, Kai Xiang Li, and Xia Sheng Sun. "Study of Piezoelectric Vibration Damping Based on the SSDI Technology for Aircraft Panels." Applied Mechanics and Materials 422 (September 2013): 105–12. http://dx.doi.org/10.4028/www.scientific.net/amm.422.105.

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In this paper, the typical aircraft panel is excited by the noise induced in traveling wave tube, the vibratory phenomenon of the typical aircraft panel is researched in detail. The piezoelectric vibration of the aircraft panels is damped by Synchronized Switch Damping on Inductor technology (SSDI technology). The acceleration parameters of the structure are controlled and the effect of structural damping is achieved.

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3

Anonymous. "Panel to allocate aircraft time." Eos, Transactions American Geophysical Union 69, no. 51 (1988): 1652. http://dx.doi.org/10.1029/eo069i051p01652-06.

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4

Sanchez-Carmona, Alejandro, and Cristina Cuerno-Rejado. "Composite stiffened panel sizing for conceptual tail design." Aircraft Engineering and Aerospace Technology 90, no. 8 (2018): 1272–81. http://dx.doi.org/10.1108/aeat-05-2017-0129.

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Purpose A conceptual design method for composite material stiffened panels used in aircraft tail structures and unmanned aircraft has been developed to bear compression and shear loads. Design/methodology/approach The method is based on classical laminated theory to fulfil the requirement of building a fast design tool, necessary for this preliminary stage. The design criterion is local and global buckling happen at the same time. In addition, it is considered that the panel does not fail due to crippling, stiffeners column buckling or other manufacturing restrictions. The final geometry is de

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5

TUDOSIE, Alexandru Nicolae. "CONTROL LAWS FOR AN AIRCRAFT SUPERSONIC INLET WITH MOBILE PANEL." SCIENTIFIC RESEARCH AND EDUCATION IN THE AIR FORCE 19, no. 1 (2017): 231–42. http://dx.doi.org/10.19062/2247-3173.2017.19.1.26.

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Malinowski, Pawel, Tomasz Wandowski, and Wieslaw Ostachowicz. "Guided Waves for Aircraft Panel Monitoring." Key Engineering Materials 558 (June 2013): 107–15. http://dx.doi.org/10.4028/www.scientific.net/kem.558.107.

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The reported research concerns experimental investigation toward the monitoring of an aircraft panel. Guided wave propagation phenomena were used to obtain information about the state of the monitored structure. A curved aluminium panel with rivets was investigated. Piezoelectric transducer was used to excite guided waves in chosen structural element. The generated signal was amplified before applying it to the transducer in order to ensure measurable amplitude of excited guided waves. Measurement of the wave field was realized using laser scanning vibrometer that registered the velocity respo

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7

Cervi, B. "Knockout punch [aircraft body panel production]." Engineering & Technology 3, no. 11 (2008): 70–71. http://dx.doi.org/10.1049/et:20081110.

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8

Li, Xiao Ge, Kai Fu Zhang, and Yuan Li. "The Modeling of Multi-State Deformation in Aircraft Panel Automated Riveting System." Applied Mechanics and Materials 224 (November 2012): 123–27. http://dx.doi.org/10.4028/www.scientific.net/amm.224.123.

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Automatic drilling and riveting process, in order to improve the quality of drilling and riveting aircraft panel deformation are needed for compensation. Due to panels in automatic drilling and riveting process with more posture, therefore, solving panel deformation of plates under more posture is important. This paper establishes aircraft panel in the automatic drilling riveting process multiple posture deformation model. Firstly, simplify the manner of support, locate and clamping between the panel and the splints, supportive of integral panel at each location are calculated using three-mome

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Lu, Cong, Deng-Sheng Huo, and Zi-Yue Wang. "Assembly variation analysis of the aircraft panel in multi-stage assembly process with N-2-1 locating scheme." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 233, no. 19-20 (2019): 6754–73. http://dx.doi.org/10.1177/0954406219869040.

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The flexible aircraft panel parts are easy to deform during the assembly process, the deformation can directly affect the assembly accuracy of the aircraft panel, and further affect the assembly quality of the aircraft fuselage. In order to evaluate the assembly accuracy of the aircraft panel in multi-stage assembly process with N-2-1 locating scheme, this paper proposes an assembly variation analyzing approach, considering the effect of the contact force between the panel skin and the fixture components on the deformation of the panel in the assembly process. The mathematical models are estab

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10

Ju, Jin San, Xiao Chuan You, Xiu Gen Jiang, and Jin Zhao Zhuang. "Fracture Analysis for Damaged Aircraft Fuselage Using Substructure Method." Advanced Materials Research 33-37 (March 2008): 29–34. http://dx.doi.org/10.4028/www.scientific.net/amr.33-37.29.

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This paper primarily describes the development and application of substructure computational analysis techniques to determine stress intensity factors for the damaged panels subjected to fatigue internal pressure. A program based on substructure analysis technique has been developed for the fracture analysis of curved aircraft panels containing cracks. This program may create whole model which consists of substructure superelements and obtain fracture parameter of the crack by expanding results in superelement automatically. For instance, a typical test curved panel model consists of 7 frames

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El-Salamony, Mostafa, and Mohamed A. Aziz. "Solar Panel Effect on Low-Speed Airfoil Aerodynamic Performance." Unmanned Systems 09, no. 04 (2021): 333–47. http://dx.doi.org/10.1142/s2301385021500175.

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Recently, a great interest in search of alternate means of power for the traditional fuel for aircraft propulsion is raised so as to decrease gas emissions and reduce operating costs. For the small and micro unmanned aerial vehicles or small transportation aircraft, there are many challenges in the direction of constructing an electric or solar powered airplane whose wings may possibly be sheltered with photo voltaic PV solar panels to harvest sun’s energy for propulsion. Greatest remarkably success solar powered aircraft has attracted the attention of researchers other than UAV and small airc

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Jiang, Junxia, Chen Bian, Yunbo Bi, and Yinglin Ke. "A new type of inner-side working head for automatic drilling and riveting system." Assembly Automation 39, no. 1 (2019): 154–64. http://dx.doi.org/10.1108/aa-09-2017-107.

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PurposeThe purpose of this paper is to design, analyze and optimize a new type of inner-side working head for automatic horizontal dual-machine cooperative drilling and riveting system. The inner-side working head is the key component of automatic drilling and riveting system, and it is a challenge to design an inner-side working head which must be stiffness and stable with a compact structure to realize its functions.Design/methodology/approachAccording to the assembly structure features of large aircraft panels and riveting process requirements, a new type of inner-side working head is desig

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Chen, Xiao Ning, Jin Long Zhao, Yun Sheng Zhang, and Bin Zhang. "Simulation and Test for the Lightning Damage of the Glass Fiber Composites." Advanced Materials Research 1003 (July 2014): 78–84. http://dx.doi.org/10.4028/www.scientific.net/amr.1003.78.

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Theoretical deducing, simulated lightning test and finite element simulation are used to research the mechanism and state of lightning damage of the aircraft composites sandwich panels. It provides the basis for the design of the aircraft lightning protection. The three-dimensional finite element model of the composites panel is constructed through the thermal electrical-mechanical multi-Physics coupling field. According to the structure and the role process, the lightning effect of the aircraft composites is analysed to study the damage mechanism and the possible state of the composites panel

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14

Li, Xi-Ning, Xiao-Gang Dang, Bao-Qiang Xie, and Yu-Long Hu. "Flexible tooling design technology for aircraft fuselage panel component pre-assembly." Assembly Automation 35, no. 2 (2015): 166–71. http://dx.doi.org/10.1108/aa-06-2014-055.

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Purpose – The purpose of this paper is to develop digital flexible pre-assembly tooling system for fuselage panels. Design/methodology/approach – First, the paper analyzes the technological characteristics of fuselage panels and then determines the pre-assembly object. Second, the pre-assembly positioning method and assembly process are researched. Third, the panel components pre-assembly flexible tooling scheme is constructed. Finally, the pre-assembly flexible tooling system is designed and manufactured. Findings – This study shows the novel solution results in significantly smaller tooling

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Haftirman, K. S. Basaruddin, A. R. Syayuthi, and M. S. Fartini. "Failure Behavior of Aircraft Sandwich Panels under Bending Load." Applied Mechanics and Materials 754-755 (April 2015): 844–48. http://dx.doi.org/10.4028/www.scientific.net/amm.754-755.844.

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Failure behavior of aircraft sandwich panels under bending load has been investigated in this study. In this study focused in effect of support span length under bending load. Three-points bending test was performed to the specimens with various span length 125 mm, 80 mm, 70 mm, and 55 mm. Standard test method and dimensions were adhering to the ASTM C393. Deflection and energy absorption of the sandwich panels have been characterized by the variation span lengths.It was found that the deflection and the energy absorption of the sandwich panels were strongly influenced by the length of support

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16

Smoot, Jessie W. "Aircraft interior panel noise dampening support brackets." Journal of the Acoustical Society of America 95, no. 4 (1994): 2300. http://dx.doi.org/10.1121/1.408628.

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17

Abramowitz, Allan, and Thor I. Eklund. "Model Tests of Aircraft Interior Panel Flammability." Journal of Fire Sciences 3, no. 2 (1985): 129–40. http://dx.doi.org/10.1177/073490418500300205.

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18

TUDOSIE, Alexandru Nicolae, and Mádálina Luciana PĂUNESCU. "AUTOMATIC CONTROL SYSTEM FOR AN AIRCRAFT PLAN SUPERSONIC INLET WITH MOBILE PANEL." SCIENTIFIC RESEARCH AND EDUCATION IN THE AIR FORCE 19, no. 1 (2017): 243–52. http://dx.doi.org/10.19062/2247-3173.2017.19.1.27.

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19

Arunkumar, MP, Jeyaraj Pitchaimani, KV Gangadharan, and MC Lenin Babu. "Sound transmission loss characteristics of sandwich aircraft panels: Influence of nature of core." Journal of Sandwich Structures & Materials 19, no. 1 (2016): 26–48. http://dx.doi.org/10.1177/1099636216652580.

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Sandwich panel which has a design involving acoustic comfort is always denser and larger in size than the design involving mechanical strength. The respective short come can be solved by exploring the impact of core geometry on sound transmission characteristics of sandwich panels. In this aspect, the present work focuses on the study of influence of core geometry on sound transmission characteristics of sandwich panels which are commonly used as aircraft structures. Numerical investigation has been carried out based on a 2D model with equivalent elastic properties. The present study has found

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Viscardi, Massimo, Maurizio Arena, Valerio Porpora, Giuliano Di Paola, and Edoardo Aubry. "Feasibility investigation of a smart thermoacoustic configuration for general aviation aircrafts." MATEC Web of Conferences 233 (2018): 00012. http://dx.doi.org/10.1051/matecconf/201823300012.

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One of the most disturbing noise for passengers is related to the engine fan. This noise is mainly tonal and in the low-frequency range. Reducing its impact for the comfort of passengers is an important and challenging issue in aeronautics. One suitable way to reduce cabin noise is to use efficient thermal acoustic insulation blankets between the interior trim panel and the exterior shell. General Aviation Aircrafts, are generally equipped with a cabin poorly efficient both in terms of thermal and acoustic insulation; this due mainly to the not availability of the most innovative solutions or

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Kalinowski, Miłosz. "Aero-Structural Optimization of Joined-Wing Aircraft." Transactions on Aerospace Research 2017, no. 4 (2017): 48–63. http://dx.doi.org/10.2478/tar-2017-0028.

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Abstract Joined-wing aircraft due to its energy characteristics is a suitable configuration for electric aircraft when designed properly. However, because of the specific for this aircraft phenomenons (e.g. static indeterminacy of structure, aerodynamic interference of lifting surfaces) it demands more complicated methods to model its behavior than a traditional aircraft configurations. For these reasons the aero-structural optimization process is proposed for joined-wing aircrafts that is suitable for preliminary design. The process is a global search, modular algorithm based on automatic geo

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Ramly, Ramzyzan, Wahyu Kuntjoro, and Mohd Kamil Abd Rahman. "Embedded FBG Sensor in Aircraft Smart Composite Materials for Structural Monitoring." Applied Mechanics and Materials 393 (September 2013): 311–16. http://dx.doi.org/10.4028/www.scientific.net/amm.393.311.

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This paper describes the use of embedded Fiber Brag Grating (FBG) sensor in the honeycomb core carbon fiber sandwich panel in smart composite materials for the application of monitoring the structural integrity of an aircraft. A part of vertical stabilizer was selected and reproduced using carbon fiber honeycomb core sandwich panels. The sandwich panel was fabricated in accordance to the generic sandwich structure and aviation industry standards, including the materials and also the method of construction. Using a carbon fiber from Hexcel as the face-sheet, Nomex honeycomb as the core, the san

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23

Mohd Aris, K. D., F. Mustapha, S. M. Sapuan, and Dayang Laila Majid. "Experimental Validation on Time Base Analysis of Various Aircraft CFRP Panel Conditions for Structural Health Monitoring." Key Engineering Materials 594-595 (December 2013): 935–39. http://dx.doi.org/10.4028/www.scientific.net/kem.594-595.935.

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This paper evaluates the feasibility and effectiveness, within controlled conditions of an active pitch catch sensing PZT sensors on two panels and an aircraft structure made from carbon fiber reinforced plastic (CFRP) pre-impregnated materials. Once cured, the exhibits were subjected to partial and full penetration damages. Two PZT sensors each acting as an actuator and receiver were placed across the investigated region at 100mm apart. Three conditions were set on each panel for each undamaged, damaged and repaired area. Fifty readings were carried out on each panel for each condition. Featu

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Wood, O. J., C. A. Featherston, D. Kennedy, Mark J. Eaton, and Rhys Pullin. "Optimised Vibration Energy Harvesting for Aerospace Applications." Key Engineering Materials 518 (July 2012): 246–60. http://dx.doi.org/10.4028/www.scientific.net/kem.518.246.

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Accurate knowledge regarding the ongoing condition of an aircraft’s structural condition together with future life predictions enable optimal use of material, hence reducing mass, cost and environmental effects. Previous work by the authors has demonstrated the potential for using energy harvested from vibrating aircraft panels to power a self contained health monitoring system based on the use of wireless sensor nodes for an aircraft structure. However the system proposed was far from optimal. Research is being undertaken to investigate the various factors affecting the power output of such a

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Trendafilova, Irina, Emil Manoach, Matthew P. Cartmell, Wiesław M. Ostachowicz, and Arkadiusz Zak. "An Investigation on Damage Detection in Aircrafts Panels Using Nonlinear Time Series Analysis." Key Engineering Materials 347 (September 2007): 213–18. http://dx.doi.org/10.4028/www.scientific.net/kem.347.213.

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This study investigates a possibility for representing, interpreting and visualising the vibration response of aircraft panels using time domain measurements. The aircraft panels are modelled as thin orthotropic plates and their vibration response is simulated using FE modelling. The vibration response of a thin aluminium panel is simulated using FE modelling. The first ten resonant frequencies are estimated for the FE model and for the dynamically tested panel. They were found to show somewhat low sensitivity to damage. Then the simulated vibration response of the panel is transformed and exp

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Katrňák, Tomáš, and Jaroslav Juračka. "Detailed Topometry FEM Optimization of Wing Structural Panel." Applied Mechanics and Materials 821 (January 2016): 357–63. http://dx.doi.org/10.4028/www.scientific.net/amm.821.357.

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The detailed topometry optimization of the critical part of an aircraft wing is presented in this article. The integral lower wing structural panel of aircraft in the Commuter category of the CS-23 regulation standard is selected for optimization. The first case demonstrates significant weight savings using modern Finite Element (FE) optimization methods for determined structural constraints. A practical aircraft operation and additional regulation requirements affect optimization constraints in the second case. This detailed optimization also consists of FE model validation, stress analyses a

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Goetzendorf-Grabowski, Tomasz, and Jacek Mieloszyk. "Common computational model for coupling panel method with finite element method." Aircraft Engineering and Aerospace Technology 89, no. 5 (2017): 654–62. http://dx.doi.org/10.1108/aeat-01-2017-0044.

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Purpose Conceptual and preliminary aircraft concepts are getting mature earlier in the design process, than ever before. To achieve that advanced level of maturity, multiple multidisciplinary analyses have to be done, often with usage of numerical optimization algorithms. This calls for right tools that can handle such a demanding task. Often the toughest part of a modern design is handling an aircraft’s computational models used for different analysis. Transferring geometry and loads from one program to another, or modifying internal structure, takes time and is not productive. Authors define

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Ju, Jin San, Xiu Gen Jiang, and Xiang Rong Fu. "Global-Local Hierarchical Analysis Techniques for Damaged Aircraft Fuselage Considering Bulging Deformation of Crack." Key Engineering Materials 324-325 (November 2006): 871–74. http://dx.doi.org/10.4028/www.scientific.net/kem.324-325.871.

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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

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Bi, Yunbo, Weimiao Yan, and Yinglin Ke. "Optimal placement of measurement points on large aircraft fuselage panels in digital assembly." Proceedings of the Institution of Mechanical Engineers, Part B: Journal of Engineering Manufacture 231, no. 1 (2016): 73–84. http://dx.doi.org/10.1177/0954405414564808.

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A large aircraft fuselage panel is commonly composed of a variety of thin-walled components. Most of these components are large, thin and compliant, and they are also prone to some flexible deformation during assembly and remain deformed after assembly. Besides, many different fabrication and assembly manners are adopted in order to guarantee the complicated assembly relationships between each component. The above characteristics often cause large aircraft fuselage panels to exhibit low stiffness and weak strength, thereby inducing deformation during assembly. Since the posture of a large airc

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Kubit, Andrzej, Tomasz Trzepiecinski, Łukasz Święch, Koen Faes, and Jan Slota. "Experimental and Numerical Investigations of Thin-Walled Stringer-Stiffened Panels Welded with RFSSW Technology under Uniaxial Compression." Materials 12, no. 11 (2019): 1785. http://dx.doi.org/10.3390/ma12111785.

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Many aircraft structures are thin walled and stringer-stiffened, and therefore, prone to a loss of stability. This paper deals with accurate and validated stability analysis of the model of aircraft skin under compressive loading. Both experimental and numerical analyzes are conducted. Two different methods of joining panel elements are considered. In the first case, the panel is fabricated using rivets. In the second variant, the refill friction stir spot welding technique is used. Both types of panels are loaded in axial compression in a uniaxial tensile testing machine. The geometrically an

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Osadchy, N. V., and V. T. Shepel. "STATIC TESTS OF FULL-SIZE ACOUSTIC PANELS OF AIRCRAFT GAS-TURBINE ENGINES." Kontrol'. Diagnostika, no. 266 (August 2020): 18–23. http://dx.doi.org/10.14489/td.2020.08.pp.018-023.

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The article suggests the installation and methods of static tests for the limit working and fracture loads of acoustic panels of aircraft gas-turbine engines (GTE). It is not possible to perform the full scope of static tests on the engine as it is technically impossible to obtain the airflow pressures required by airworthiness standards. The required pressure in the proposed installation is created by the punch, which is the same shape as the panel to be tested. Flexible rubber sheets are used to equalize contact pressure on the panel surface. The similarity in the distribution of pressures o

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Osadchy, N. V., and V. T. Shepel. "STATIC TESTS OF FULL-SIZE ACOUSTIC PANELS OF AIRCRAFT GAS-TURBINE ENGINES." Kontrol'. Diagnostika, no. 266 (August 2020): 18–23. http://dx.doi.org/10.14489/td.2020.08.pp.018-023.

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The article suggests the installation and methods of static tests for the limit working and fracture loads of acoustic panels of aircraft gas-turbine engines (GTE). It is not possible to perform the full scope of static tests on the engine as it is technically impossible to obtain the airflow pressures required by airworthiness standards. The required pressure in the proposed installation is created by the punch, which is the same shape as the panel to be tested. Flexible rubber sheets are used to equalize contact pressure on the panel surface. The similarity in the distribution of pressures o

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33

Zhang, Xian Jie, Jun Biao Wang, Yong Jun Wang, and Ming Jie Qiao. "Prediction of Shot Peen Forming Parameters of Integral Aircraft Wing Panels." Materials Science Forum 532-533 (December 2006): 937–40. http://dx.doi.org/10.4028/www.scientific.net/msf.532-533.937.

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In this paper, algorithms are presented for predicting peen forming parameters of integral aircraft wing panels with complex airfoil shapes. The peen forming deformation is divided into stretching deformation and bending deformation. The stretching deformation is assumed to result from the tensile strain within the plane panel, and the bending deformation corresponds to the difference of maximum and minimum curvature of the airfoil surface. The distribution of the forming tensile strain within the panel is obtained by optimal mapping of the airfoil surface in the sense that the stretching defo

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Uday Deepika. A , K. Veeranjaneyulu, Uday Deepika A. ,. K. Veeranjaneyulu. "Buckling Analysis of Stiffened Panel for Aircraft Fuselage." International Journal of Mechanical and Production Engineering Research and Development 8, no. 1 (2018): 1299–308. http://dx.doi.org/10.24247/ijmperdfeb2018150.

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Dehm, S., and D. Wurzel. "Fast, in-situ repair of aircraft panel components." Journal of Aircraft 26, no. 5 (1989): 476–81. http://dx.doi.org/10.2514/3.45788.

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York, C. B., and F. W. Williams. "Aircraft wing panel buckling analysis: efficiency by approximations." Computers & Structures 68, no. 6 (1998): 665–76. http://dx.doi.org/10.1016/s0045-7949(98)00050-9.

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Fangfang, Yu, Du Baorui, Ren Wenjie, Zheng Guolei, and Chu Hongzhen. "Slicing Recognition of Aircraft Integral Panel Generalized Pocket." Chinese Journal of Aeronautics 21, no. 6 (2008): 585–92. http://dx.doi.org/10.1016/s1000-9361(08)60178-8.

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Liu, Chuang, Jun Biao Wang, and Xian Jie Zhang. "Digital Manufacturing System of Aircraft Wing Integral Panel." Advanced Materials Research 97-101 (March 2010): 2732–35. http://dx.doi.org/10.4028/www.scientific.net/amr.97-101.2732.

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The function model of digital manufacturing system of aircraft wing integral panel is proposed using IDEF0. The digital integral panel manufacturing subsystems including process design, manufacturing model definition, tooling design and manufacturing, process information management, fabrication and inspection are described. In this system, the integral panel is formed by shot peen forming process, and the corresponding approach to solve some key manufacturing problems is described. The part manufacturing model is presented to represent the intermediate part in the fabrication process. The info

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Wang, Yiwei, Christian Gogu, Nicolas Binaud, Christian Bes, Raphael T. Haftka, and Nam-Ho Kim. "Predictive airframe maintenance strategies using model-based prognostics." Proceedings of the Institution of Mechanical Engineers, Part O: Journal of Risk and Reliability 232, no. 6 (2018): 690–709. http://dx.doi.org/10.1177/1748006x18757084.

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Aircraft panel maintenance is typically based on scheduled inspections during which the panel damage size is compared to a repair threshold value, set to ensure a desirable reliability for the entire fleet. This policy is very conservative since it does not consider that damage size evolution can be very different on different panels, due to material variability and other factors. With the progress of sensor technology, data acquisition and storage techniques, and data processing algorithms, structural health monitoring systems are increasingly being considered by the aviation industry. Aiming

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Hai, Gong, Yi Bin, Wu Yunxin, Liao Zhiqi, Liu Yaoqiong, and Du Fei. "Integral Aircraft Wing Panels with Penetration Cracks: The Influence of Structural Parameters on the Stress Intensity Factor." Applied Sciences 10, no. 12 (2020): 4142. http://dx.doi.org/10.3390/app10124142.

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The finite element model of integral wing panels with central penetration cracks under bending load was established, and the crack propagation process of the aircraft panel was simulated. The stress intensity factor (SIF) of the crack tip during crack propagation under varying conditions of crack length and panel structural parameters was determined. The effects of the panel structure parameters and crack size on the crack tip SIF were obtained. The regression analysis of the finite simulation element results has been performed and a regression model of SIF at the crack tip of the integral pan

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Zhou, Junjie, and Shengnan Wang. "Initial Fatigue Quality Assessment for Aircraft Wing Panel Fastener Hole." Xibei Gongye Daxue Xuebao/Journal of Northwestern Polytechnical University 36, no. 1 (2018): 91–95. http://dx.doi.org/10.1051/jnwpu/20183610091.

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An effective initial fatigue quality assessment method is presented in order to verify aircraft wing panel fastener hole whether to satisfy the design requirements. Firstly, after finishing fatigue test of bolted specimens and fatigue fracture interpretation, the time to crack initiation distributions under 3 stress levels are obtained and then a general equivalent initial flaw size distribution is established. Secondly, a method of fatigue life prediction with 95% reliability is proposed. Finally, the initial fatigue quality of aircraft wing panel fastener hole is evaluated based on the econo

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Liu, C. G., F. X. Tan, and X. J. Li. "Multi-Point Positioning Tooling Technology for Aircraft Manufacturing." Materials Science Forum 628-629 (August 2009): 517–22. http://dx.doi.org/10.4028/www.scientific.net/msf.628-629.517.

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Multi-point positioning tooling is a flexible fixture used in assembly of contoured panels in aircraft manufacturing. It uses many numerical controlled punch elements to configure a flexible tooling system which shape contour can be adjusted according CAD data. The paper is focused on analysis of the operations to be done on the formed panel when they are clamped and supported on the MPPT. From the results of finite element analysis, it can be concluded that the interval between punches is the key factor which affects accuracy of MPPT. And also sheet thickness, material and geometric shape hav

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Sahwee, Zulhilmy, Halil Husain, Muhd Khairulzaman Abdul Kadir, Mohamad Fiteri Razali, Mohamad Zikri Zainol, and Shahliza Azreen Sarmin. "Automatic Monitoring of Photovoltaic Cells Performance on Solar Aircraft." Applied Mechanics and Materials 225 (November 2012): 356–60. http://dx.doi.org/10.4028/www.scientific.net/amm.225.356.

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The development towards the usage of green energy sources is currently a major concern in today’s lifestyle due to the limited supply of fossil-based energy. Many have turned to alternative energy resources that are less polluting to the environment. Replacement of the vehicle’s fuel-based engine to an electric-based engine by using alternative or hybrid energy is one of the possible paths that many studies are conducted. In the aviation field, harvesting solar energy by using solar panel to charge aircraft on-board batteries could reduce the dependency on fuel. As with any other energy conver

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Ju, Jin San, Xiu Gen Jiang, and Xiang Rong Fu. "Fracture Analysis for Stiffened Aircraft Fuselage Using Substructure Method." Key Engineering Materials 385-387 (July 2008): 837–40. http://dx.doi.org/10.4028/www.scientific.net/kem.385-387.837.

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This paper primarily describes the development and application of substructure computational analysis techniques in two-step hirarchical strategy to determine stress intensity factors for the stiffened damaged panels subjected to fatigue internal pressure. A program based on substructure analysis technique and global-local hierarchical strategy has been developed for the fracture analysis of curved aircraft panels containing cracks. This program may create superelements in global and local models, and obtain fracture parameter of the crack in local model by expanding results in superelements a

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Li, Shu Lin, and Man Yi Hou. "Finite-Element Simulation and Analysis on High Velocity Impact Damage of Aircraft Panel Structure." Key Engineering Materials 353-358 (September 2007): 1033–36. http://dx.doi.org/10.4028/www.scientific.net/kem.353-358.1033.

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The finite-element simulation models of the projectile and the discrete rod impacting to the aircraft panel structure in high velocity are established according to some experiment projects. Based on dynamic finite-element Program, the forming of impact damage in the panel structure is simulated. Through comparing the simulation results of damage pattern and size in the panel to the experiment results, the reliability of the material models and equations of state and contact algorithm used in the simulations is testified. Take the simulation of projectile vertically impacting to the panel as ex

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Hou, Man Yi, Shu Lin Li, Yang Yi Jiang, and Shou An Li. "Simulation Study on Impact Damage of Aircraft Panel Structure under Stress." Key Engineering Materials 324-325 (November 2006): 391–94. http://dx.doi.org/10.4028/www.scientific.net/kem.324-325.391.

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Based on ANSYS/LS-DYNA code, the finite-element models were created to simulate the response of the aircraft panel structure impacted by high velocity projectile. The models proved to be effective through the comparison between the results of simulation and relative experiments. Then the impact process was simulated respectively considering the states of various types of stress in the panel. Through analyzing the simulation results, the influence of various stress states in the panel on impact response and damage mechanism was summed up. The conclusions indicated that the stress and particular

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Schoř, Pavel, Martin Kouřil, and Vladimír Daněk. "SIMULATION OF A MANEUVERING AIRCRAFT USING A PANEL METHOD." Acta Polytechnica 61, no. 2 (2021): 378–90. http://dx.doi.org/10.14311/ap.2021.61.0378.

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We present a method for numerical simulations of a maneuvering aircraft, which uses a first-order unsteady panel method as the only source of aerodynamic forces and moments. By using the proposed method, it is possible to simulate a motion of an aircraft, while the only required inputs are geometry and inertia characteristics, which significantly reduces the time required to start the simulation. We validated the method by a comparison of recordings of flight parameters (position, velocities, accelerations) from an actual aerobatic flight of a glider and the results obtained from the simulatio

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Croitoru, Emilian Ionut, and Gheorghe Oancea. "Impact Analysis of an Oxygen Mask Locking Panel of Aircraft Using Finite Element Modelling." Applied Mechanics and Materials 657 (October 2014): 735–39. http://dx.doi.org/10.4028/www.scientific.net/amm.657.735.

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This paper presents a method of finite element modelling used for the impact analysis of a composite panel. In this research, the composite panel consists of an oxygen mask locking panel of an aircraft. This panel is loaded with one concentrated abuse loading and three uniform distributed abuse loading cases and the stress variation within the composite panel for each load case is determined. In order to assess the impact analysis on the oxygen mask panel of the aircraft, a finite element model is created using Patran as the main application for pre/post-processing and Nastran as the main proc

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Ratcliffe, Stanley. "Safe Vertical Separation of Aircraft." Journal of Navigation 44, no. 3 (1991): 386–91. http://dx.doi.org/10.1017/s0373463300010213.

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A recent paper by Cox, ten Have and Forrester is planned to be the first of a series describing the work of the European Vertical Studies Sub Group (VSSG) set up under the ICAO panel for review of the general concept of separation. The present paper has no connection with the VSSG work.

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Conlan-Smith, Cian, Néstor Ramos-García, Ole Sigmund, and Casper Schousboe Andreasen. "Aerodynamic Shape Optimization of Aircraft Wings Using Panel Methods." AIAA Journal 58, no. 9 (2020): 3765–76. http://dx.doi.org/10.2514/1.j058979.

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