scholarly journals Development of a portable, focused‐beam ultrasonic scanner for the NDI of adhesively bonded aircraft fuselage skin structures

1995 ◽  
Vol 98 (5) ◽  
pp. 2875-2875
Author(s):  
Thadd C. Patton ◽  
David K. Hsu
Keyword(s):  
Adhesively Bonded ◽  
Aircraft Fuselage ◽  
Ultrasonic Scanner ◽  
Fuselage Skin ◽  
Focused Beam

FEM-BEM Modeling and Experimental Verification of the Vibro-Acoustic Behaviour of a Section of Aircraft Fuselage

2016 ◽  
Author(s):  
Viken N. Koukounian ◽  
Chris K. Mechefske
Keyword(s):  
Transmission Loss ◽  
Pressure Fluctuations ◽  
Experimental Methodology ◽  
Reverberation Chamber ◽  
Aircraft Fuselage ◽  
Ongoing Work ◽  
Damping Materials ◽  
Fuselage Skin ◽  
International Testing ◽  
Testing Standards

The aerodynamics of an aircraft in flight impose significant stresses upon the structure. Specifically, the mechanics of fluid flow are highly turbulent and, the layer around the aircraft, is referred to the turbulent boundary layer (TBL). The TBL incites a gradient of pressure fluctuations across the fuselage skin resulting in its vibration, and in turn, the generation of noise inside the passenger cabin. The investigation herein proposes a hybrid FEM-BEM modeling technique to predict the aforementioned vibro-acoustic response and an experimental methodology to verify the results (following ASTM and ANSI international testing standards). The described expectations required construction of an acoustic facility consisting of a reverberation chamber and a semi-anechoic room, the development of DAQ software using LabVIEW, an assembly of DAQ hardware using National Instruments products, and the post-processing of test data using Microsoft Excel. The principal quantity of interest is transmission loss (though insertion loss, absorption and other metrics are also calculated). Two panels (0.04in (40thou) and 0.09in (90thou) in thickness) were simulated and tested (0.01in = 1thou). The calculated error of the proposed methodology is within a maximum of 5dB, with an average of 1dB. Ongoing work is investigating complex constructions and the use of damping materials.

Case Studies of Aircraft Fuselage Cracking

2019 ◽  
Vol 33 ◽  
pp. 11-18 ◽  
Author(s):  
A.M. Al-Mukhtar
Keyword(s):  
Cyclic Loading ◽  
Corrosion Damage ◽  
Crack Size ◽  
Critical Crack ◽  
Aircraft Fuselage ◽  
Effective Role ◽  
Number Of Cycles ◽  
Fuselage Skin ◽  
Critical Crack Size ◽  
Cycles To Failure

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.

scholarly journals AFP mandrel development for composite aircraft fuselage skin

2014 ◽  
Vol 15 (1) ◽  
pp. 32-43 ◽  
Author(s):  
Deepak Kumar ◽  
Myung-Gyun Ko ◽  
Rene Roy ◽  
Jin-Hwe Kweon ◽  
Jin-Ho Choi ◽  
...  
Keyword(s):  
Aircraft Fuselage ◽  
Fuselage Skin

Structural health monitoring of an adhesively bonded CFRP aircraft fuselage by ultrasonic Lamb Waves

2020 ◽  
Vol 234 (13) ◽  
pp. 2000-2010
Author(s):  
J Weiland ◽  
DF Hesser ◽  
W Xiong ◽  
A Schiebahn ◽  
B Markert ◽  
...  
Keyword(s):  
Structural Health Monitoring ◽  
Health Monitoring ◽  
Adhesive Bonding ◽  
Lamb Waves ◽  
Reconstruction Algorithm ◽  
Aviation Industry ◽  
Adhesively Bonded ◽  
Structural Health ◽  
Aircraft Fuselage ◽  
Future Work

The aviation industry faces the challenge of offering aircraft that are lighter, more economical, and safer. One of the solutions is to increase the use of composites. For these materials, adhesive bonding has proven to be the appropriate joining technology. To check these adhesive joints, costly and time-consuming maintenance measures are carried out. An intelligent Structural Health Monitoring (SHM) system can extend these intervals and allow the use of a predictive maintenance system. This paper describes the method of Ultrasonic Lamb Waves for monitoring a adhesively bonded Carbon Fiber Reinforced Polymer (CFRP) aircraft fuselage. Prior to this, the production of a segment of a fuselage and the characterization of the materials (CFRP and adhesive) is shown. Afterwards the method of Ultrasonic Lamb Waves with the use of piezoelectric transducers and signal processing based on the Reconstruction Algorithm for Probabilistic Inspection of Damage (RAPID) algorithm are explained. At the end, the experimental evaluation of an undamaged and a damaged fuselage structure is done. The results have shown the possibility of RAPID algorithm for damage detection on adhesive. An outlook on future work is given.

Design and Analysis of Composite Window Frame for a Regional Aircraft

2020 ◽  
Vol 11 (04) ◽  
pp. 2050006
Author(s):  
João Afonso Gaspar Lopes ◽  
Omar Bacarreza ◽  
Zahra Sharif Khodaei
Keyword(s):  
Structural Integrity ◽  
Thermoplastic Composite ◽  
Design Criteria ◽  
Adhesively Bonded ◽  
Window Frame ◽  
Damage And Failure ◽  
Shear Loads ◽  
Failure Loads ◽  
Flight Loads ◽  
Fuselage Skin

This work presents the design and analysis of a thermoplastic composite window frame for integration into a regional aircraft. The main parameters which are investigated include buckling, damage and failure loads of a composite window frame subjected to shear loads repesentative of fuselage skin stress distribution due to flight loads. The attachment of such thermoplastic window frame to a thermoset fuselage skin was investigated including both adhesively bonded interface as well as riveting. Even though the bonded frame did meet the design criteria, its failure was very sudden, and the riveted assembly showed a considerably higher strength and structural integrity. The numerical simulation resulted in failure loads which matched very closely to experimental results.

Analytical-Experimental Evaluation of the Fatigue Resistance for Aircraft Fuselage Skin in the Case of Corrosion Damage

2021 ◽  
Vol 64 (2) ◽  
pp. 181-188
Author(s):  
V. S. Shapkin ◽  
A. V. Lapaev ◽  
K. A. Matveev ◽  
V. A. Gorshkov ◽  
A. A. Bogoyavlenskii
Keyword(s):  
Fatigue Resistance ◽  
Experimental Evaluation ◽  
Corrosion Damage ◽  
Aircraft Fuselage ◽  
Fuselage Skin

Fracture Analysis of Adhesively Bonded Cracked Panels

1978 ◽  
Vol 100 (1) ◽  
pp. 10-15 ◽  
Author(s):  
T. Swift
Keyword(s):  
Air Force ◽  
Complex Stress ◽  
Intensity Factors ◽  
Adhesively Bonded ◽  
Stiffened Panels ◽  
Aircraft Fuselage ◽  
Linear Elastic ◽  
Elastic Fracture ◽  
Shear Distortion ◽  
Using Data

An analysis procedure is presented for predicting crack tip stress intensity factors, adhesive stiffener stresses and strains, and the residual strength of adhesively bonded stiffened panels containing cracks equally spaced between stiffeners. The approach considers the cracked skin as a linear elastic fracture mechanics problem but accounts for the effects of plasticity on both adhesive and stiffener. The solution is based on displacement compatibility between the cracked sheet and the stiffener, accounting for shear distortion of the adhesive. Displacements in the cracked sheet are determined using the Westergaard complex stress function approach. The analysis was correlated with test using data from a cracked panel, simulating typical aircraft fuselage construction, previously tested during the Air Force Primary Adhesive Bonded Structures Technology (PABST) Program.

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