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This paper aims to present the results of calculations that checked how the longerons and frames arrangement affects the stiffness of a conventional structure. The paper focuses only on first stage of research – analysis of small displacement. Main goal was to compare different structures under static loads. These results are also compared with the results obtained for a geodetic structure fuselage model of the same dimensions subjected to the same internal and external loads.

The finite element method analysis was carried out for a section of the fuselage with a diameter of 6.3 m and a length equal to 10 m. A conventional and lattice structure – known as geodetic – was used.

Finite element analyses of the fuselage model with conventional and geodetic structures showed that with comparable stiffness, the weight of the geodetic fuselage is almost 20 per cent lower than that of the conventional one.

This analysis is limited to small displacements, as the linear version of finite element method was used. Research and articles planned for the future will focus on nonlinear finite element method (FEM) analysis such as buckling, structure stability and limit cycles.

The increasing maturity of composite structures manufacturing technology offers great opportunities for aircraft designers. The use of carbon fibers with advanced resin systems and application of the geodetic fuselage concept gives the opportunity to obtain advanced structures with excellent mechanical properties and low weight.

This paper presents very efficient method of assessing and comparison of the stiffness and weight of geodetic and conventional fuselage structure. Geodetic fuselage design in combination with advanced composite materials yields an additional fuselage weight reduction of approximately 10 per cent. The additional weight reduction is achieved by reducing the number of rivets needed for joining the elements. A fuselage with a geodetic structure compared to the classic fuselage with the same outer diameter has a larger inner diameter, which gives a larger usable space in the cabin. The approach applied in this paper consisting in analyzing of main parameters of geodetic structure (hoop ribs, helical ribs and angle between the helical ribs) on fuselage stiffness and weight is original.

The purpose of this paper is to present the result of calculations that were performed to estimate the structural weight of the passenger aircraft using novel technological solution. Mass penalty resulting from the installation of the fuselage boundary layer ingestion device was needed in the CENTRELINE project to be able to estimate the real benefits of the applied technology.

This paper focusses on the finite element analysis (FEA) of the fuselage and wing primary load-carrying structures. Masses obtained in these analyses were used as an input for the total structural mass calculation based on semi-empirical equations.

Combining FEA with semi-empirical equations makes it possible to estimate the mass of structures at an early technology readiness level and gives the possibility of obtaining more accurate results than those obtained using only empirical formulas. The applied methodology allows estimating the mass in case of using unusual structural solutions, which are not covered by formulas available in the literature.

Accurate structural mass estimation is possible at an earlier design stage of the project based on the presented methodology, which allows for easier and less costly changes in designed aircrafts.

The presented methodology is an original method of mass estimation based on a two-track approach. The analytical formulas available in the literature have worked well for aeroplanes of conventional design, but thanks to the connection with FEA presented in this paper, it is possible to estimate the structure mass of aeroplanes using unconventional technological solutions.

The purpose of this paper is to present the result of simulations that were performed to assess the uncontrolled motion of the passenger aircraft following its wing tip was suddenly cut. Such a simulation can help to understand the tendencies of aircraft behaviour after wing tip cut, assess how fast this phenomenon is going on and estimate the values of angles of attack, sideslip and pitch angle basing on given aerodynamic characteristics. Also, answer the question if pilot can counteract high deviations from flight path initially planned during the final phase of approach to landing.

Simulation is based on the full non-linear equations of motion derived from generalised equations of change of momentum and moment of momentum of rigid body. Dynamic equations of motion in the so-called normal mode are solved in the so-called stability frame of reference.

It was found that asymmetric rolling moment must be compensated by essential increase of pitching moment. Moreover, it appeared that aircraft goes into high angles of attack and high pitch angle and, therefore, for reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs.

For accurate simulation, it is strongly recommended to perform to perform first the wind tunnel testing in the range of +20o ÷ 120o and use it in flight simulation.

The presented methodology is an original for numerical simulation of flight trajectory during the final phase of approach to landing in a hazardous state of flight. For reliable simulation, the available aerodynamic characteristics must include angles of attack till 90 degrees when total flow separation occurs, whereas usually maximum angles of attack used in wind tunnel experiments for passenger aircraft are not higher than 25 degrees. The influence of limited range of experimental data on results of simulation is another value which can be adopted in the future investigations of hazardous states of flight.

Aircraft structure mass estimation is a very important issue in aerospace. Multiple methods of different fidelity are available, which give results with varying accuracy. Sometimes these methods are giving a high discrepancy of estimated mass compared to the real mass of the structure. The discrepancy is especially noticeable in the case of small aircraft with a composite structure. Their mass properties highly depend not only on the material but also on technology and the human factor. Moreover, methods of mass estimation for unmanned aerial vehicle (UAV) platforms are even less established and examined. The purpose of this paper is to present and discuss various methods of mass estimation.

The paper presents different procedures of mass estimation for small UAVs with a composite structure. Beginning from the simplest one, where mass is estimated basing on a single equation and finishing with a mass estimation based on finite element method model and three-dimensional computer-aided design model. The results from all methods are compared with the airworthy aircraft and conclusions are discussed.

Mass of flying aircraft was estimated with different methods and compared. It revealed levels of accuracy of the investigated methods. Moreover, the influence on structure mass of human factor, glueing and painting is underlined.

Mass of the structure is a key factor in aerospace, which influences the performance of the aircraft. Thorough knowledge about the accuracy of the mass estimation methods and possible sources of discrepancies in mass analyses provides an essential tool for designers, which can be used with confidence and allows for the development of new cutting-edge constructions.

There are very few comparisons of mass estimation methods with an actual mass of manufactured and functional airframes. Additionally, mass estimation inaccuracies based on technological issues are presented, which is seldom done.

The study presents test results concerning the impact of high temperature and different cooling conditions on the mechanical properties of quenched and self-tempered reinforcing steel. The purpose of this paper is to clarify the extent of the history of the material’s temperature development profile, the course and the intensity of fire exposure and how cooling conditions determines its properties.

Each specimen series was heated to the temperatures of T = 200°C, 400°C, 600°C, 700°C, 800°C and 1,000 °C. The specimens were either slowly cooled down or subjected to rapid cooling with water quenching, which can be encountered during a firefighting operation. Additionally, stress–strain relationships, microhardness and structural observations were also performed.

The results of the presented experiments have shown that the steel bars previously heated in fire conditions were very sensitive to the cooling intensity. The test results from the steel specimens – that were heated and quenched with water – demonstrate an increase in tensile strength but a significant reduction in material plasticity.

The presented piece of work provides a contribution for fire safety engineering giving insight into the fire behaviour of reinforcing steel under fire conditions and subjected to rapid or slow cooling. This study has shown the threats arising from thermally induced changes in steel microstructure because of high-temperature exposure. It should also be noted that structure changes may have a local character and refer to steel rebars that are exposed because of fire spalling of concrete cover.