Failure and fatigue life estimate of a pre-stressed Aircraft seat support tube (original) (raw)
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STRESS ANALYSIS AND FATIGUE DAMAGE ESTIMATION FOR SEAT SUPPORT ATTACHMENT
Aircraft structure is the most obvious example where structural efficiency results in light weight and high operating stresses. Airframe experiences variable loading in service. The paper presents the response of fuselage seat attachment due to cabin pressure and fluctuating load on seat attachment. The stresses developed in the fuselage at the altitude during actual flight condition were simulated by using the FEA software MSC.PATRAN & MSC.NASTRAN. The response of structure for the hoop stress and longitudinal stress developed in the fuselage due to cabin pressurization is studied by using finite element analysis technique. Linear static stress analysis is carried out for the identification of the fatigue critical location. The global and local analysis is carried out. Miner's rule will be used for fatigue damage calculation with the help of the S-N diagram of the respective material used in the structure.
Fatigue Life Extimation of a Military Aircraft Structure subjected to Random Loads
Procedia Structural Integrity, 2018
During their operation, modern aircraft engine components are subjected to increasingly demanding operating conditions, especially the high pressure turbine (HPT) blades. Such conditions cause these parts to undergo different types of time-dependent degradation, one of which is creep. A model using the finite element method (FEM) was developed, in order to be able to predict the creep behaviour of HPT blades. Flight data records (FDR) for a specific aircraft, provided by a commercial aviation company, were used to obtain thermal and mechanical data for three different flight cycles. In order to create the 3D model needed for the FEM analysis, a HPT blade scrap was scanned, and its chemical composition and material properties were obtained. The data that was gathered was fed into the FEM model and different simulations were run, first with a simplified 3D rectangular block shape, in order to better establish the model, and then with the real 3D mesh obtained from the blade scrap. The overall expected behaviour in terms of displacement was observed, in particular at the trailing edge of the blade. Therefore such a model can be useful in the goal of predicting turbine blade life, given a set of FDR data.
FATIGUE LIFE ESTIMATION OF REAR FUSELAGE STRUCTURE OF AN AIRCRAFT
Integrity of the airframe structure is achieved through rigorous design calculations, stress analysis and structural testing. Finite element method (FEM) is widely used for stress analysis of structural components. Each component in the airframe becomes critical based on the load distribution, which in-turn depends on the attitude of the aircraft during flight. Fuselage and wing are the two major components in the airframe structure. The current study includes a portion of the fuselage structure. Empennage is the rear portion of the aircraft, which consists of rear fuselage, Horizontal tail and vertical tail. The air loads acting on the HT also get transferred to rear portion of the fuselage. First step in ensuring the safety of the structure is the identification of critical locations for crack initiation. This can be achieved through detailed stress analysis of the airframe In this project one of the major stress concentration areas in the fuselage is considered for the analysis. Rear fuselage portion with a cargo door cutout region will be analysed. The structure considered for the stress analysis consists of skin, bulkheads and longerons, which are connected to each other through rivets. Aerodynamic load acting on the aircraft components is a distributed load. Depending on the mass distribution of the fuselage structure the inertia forces will vary along the length of the fuselage. The inertia force distribution makes the fuselage to bend about wing axis. During upward bending, bottom portion of the fuselage will experience tensile stress. A cutout region in the tensile stress field will experience high stress due to concentration effect. These high stress regions will be probable fatigue crack initiation locations in the current work, fatigue damage calculation will be carried out to estimate the fatigue life of the structure under the fluctuating loads experienced during flight. Miner's rule will be adopted for fatigue damage calculation.
Finite Element Stress Analysis of Airplane Seat
European Mechanical Science , 2021
Finite element method (FEM) is frequently used in the seat industry, as well as in the aircraft seat industry, which is a sub-branch of it, especially in the last 10-15 years. Developments in finite element (FE) analysis have enabled safer and cheaper designs to be created in the seat industry. The accuracy of the finite element analysis performed while using this method is extremely important. For this reason, in creating the finite element model, some important parameters must be selected and processed correctly for the model to give the correct result. These parameters can be listed as element size, time scale, analysis type, and material model. The verification of the Finite element analysis (FEA) results is usually done using experimental methods. It is known that in the finite element analysis results almost equivalent to experimental results are obtained when the aforementioned parameters are modeled correctly. This study aims to perform static stress analysis and topology optimization of an airplane seat using the FEM. The static stresses and displacements created at the seat are calculated under simulated loading conditions. Thanks to the topology optimization study, the weight of the airplane seat is minimized by a 30% without sacrificing seat safety. A comparison of static stresses obtained from the FE and analytical models indicates a reasonable correlation, demonstrating confidence in our FE analysis.
Design and characterization of a shock and vibration mitigation seat system
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Extensive research has been conducted into the development of pneumatic seatbladder systems for shock and vibration mitigation for use in current U.S military vehicle envelopes. This research expands on the previous work through an elaborate experimental characterization of four prototype air bladder seat cushion systems. The experimental characterization conducted included shock testing, continuous vibration, and internal dynamic pressure measured during the shock event. The shock testing was conducted both at the Army Research Lab as well at UNLV. The shock testing conducted at UNLV was performed on a drop tower designed and constructed during the time of this research. The scope of the testing was extended beyond the U.S military's requirements to include random continuous vibrations which can cause physical harm to the occupant over extended durations. The primary considerations are to increase the survivability of crewmembers exposed to mine blasts and mitigation of the vib...
Fatigue life analysis using random vibration data
International Conference on Engineering Technologies (ICENTE'17), 2017
In this study, fatigue analysis of a missile body structure during captive carriage at the underwing of a fixed wing aircraft platform is performed. The F-16 jet aircraft was determined as a fixed wing aircraft platform. As the external store, a data measurement store (DMS) was designed to measure loads during five different sorties. Acceleration and strain data were collected during the test. Accelerometers were used to generate power spectral densities (PSD) for the vibration tests and fatigue analyses. Strain gages were placed to the most critical locations for fatigue calculations
Civil transport aircraft is used for ferrying passengers from one place to another. Aircraft is a highly complex flying structure. Generally transport aircraft undergoes nominal maneuring flights. During the flight when the maximum lift is generated, the wings of the aircraft will undergo highest bending moment. The bending moment will be maximum at the root of the wing which caused highest stress at this location. Wings are attached to the fuselage structure through wing-fuselage attachment brackets. The bending moment and shear loads from the wing are transferred to the fuselage through the attachment joints. In this project bending load transfer joint is considered for the analysis. First one needs to ensure the static load carrying capability of the wing-fuselage attachment bracket. Stress analysis will be carried out for the given geometry of the wing-fuselage attachment bracket. Finite element method is used for the stress analysis. In the current project, an attempt will be made to predict the fatigue life of wing-fuselage attachment bracket in a transport airframe. In a metallic structure fatigue manifests itself in the form of a crack which propagates. Fatigue cracks will appear at the location of high tensile stress locations. These locations are invariably of high stress concentration. Fatigue life calculation will be carried out for typical service loading condition using constant amplitude S-N data for various stress ratios and local stress history at stress concentration. In this paper for modeling CATIA V5 software is used and for analysis tool MSC/ PATRAN and MSC/ NASTRAN 2010.
Fatigue Life Analysis On A Missile Body Exposed To Flight Loads
ANKARA INTERNATIONAL AEROSPACE CONFERENCE, 2017
In this study, fatigue analysis of a missile body structure during captive carriage at the underwing of a fixed wing aircraft platform is performed. The F-16 jet aircraft was determined as a fixed wing aircraft platform. As the external store, a data measurement store (DMS) was designed to measure loads during captive carriage for five different sorties. Acceleration and strain data were collected during the test. Accelerometers were used to generate power spectral densities (PSD) for the vibration tests and fatigue analyses. Strain gages were placed to the most critical locations for fatigue calculations. In data processing and analyses
Fatigue life estimate of landing Gear's leg using modal analysis
The International Journal of Multiphysics, 2014
The present work concerns solving Noise, Vibration and Harshness (NVH) and fatigue based on Power Spectrum Density (PSD) analysis of a landing gear's leg for an Un-Manned Aerial Vehicle (UAV). This analysis includes random vibration and high-cycle fatigue analysis in a random vibration environment. In this analysis, the cumulative damage ratio is computed using material S-N (Stress-Number of cycles) fatigue curve. Dirlik method is used for the analysis of lifetime as it is proven to provide accurate results for large number of applications, both in automotive and aerospace industry. It is also compared to other methods that have been developed in LS-DYNA ® as well. The input acceleration PSD data are provided through measurements. The obtained analysis results shows that although the landing gear design is safe according to dynamic and static load, its service life is about 3037 hours due to random vibration effect.