A structural sizing methodology for the wing-fuselage of the Flying-V

Abstract

The demand for sustainability in aviation is at a historical peak. Efforts from industry and institutions have made the conventional tube-and-wing aircraft greatly efficient, to the degree that a performance ceiling is being approached. Unconventional aircraft configurations promise to reduce fuel burn further, at the cost of design complexity and risk. One aircraft configuration which has the potential to allow for this increase in performance is the Flying-V concept. This research aims to contribute to a better understanding of the structural design methods of the Flying-V.

Several studies into the viability and performance of the Flying-V have been performed, showing promising results. However, these results mainly pertain to the aerodynamic and flight performance fields. Preliminary structural analyses have shown a significant reduction in structural mass. However, these structural analyses have not regarded the wing-fuselage enough as a novel structure, applying analyses which work well conventional wings or fuselages, but not on the oval fuselage structure. More specifically, the hoop stresses in the skins and frames were not assessed beyond the membrane response. Furthermore, a clear opportunity was identified in the inclusion of a fatigue analysis, beyond using the material fatigue limit.

The wing-fuselage structure was identified to be dominant in terms of weight. Furthermore, there is a knowledge gap in the understanding of the behaviour of the unconventional wing-fuselage structure. Therefore, the research goal was to determine an improved structural analysis and sizing methodology for the wing-fuselage. For the sizing procedure, failure modes in three categories were selected: static material failure, fatigue and stability. With the selected failure modes, the stresses were determined.

An extensive load case study was performed, resulting in a weight and load distribution over the aircraft for a large number of different load cases and different weights. Due to the favourable weight distribution and open cabin structure, the cross-section stresses were considered to be equally important as the longitudinal stresses in terms of failure. Megson's (2013) boom method was applied to determine the stresses in longitudinal direction, while an enriched 2-dimensional FEM model was used to determine the stresses in the cross-sectional plane. The first important insights were obtained in the stress analysis stage: the hoop stresses vary greatly over the constant arc sections, due to significant bending. The stress analysis was validated using a 3-dimensional ABAQUS model of a fuselage section.

With the obtained stress calculation approach, the failure modes were determined for each category. Static material failure was calculated in limit and ultimate load conditions for the yield and ultimate static material strength, respectively. The fatigue performance was assessed by calculating the crack initiation life and crack propagation life (including residual strength). A fatigue load spectrum was determined from the standardised load spectrum by Jonge et al. (1973). Finally, multiple stability-related failure phenomena were assessed. The well-known column and plate buckling phenomena were evaluated. Global buckling and panel buckling were assessed by using a smeared property plate buckling method. The buckling behaviour under combined loading was approximated with an interaction relation.

With the obtained failure index distributions, a sizing methodology was determined. The sizing methodology was based on the identification of the driving variable(s) for each failure mode using stress and failure distributions from a generic geometry. As the sizing and analysis loop proved to require a substantial runtime, the sizing procedure was reduced in scope. The critical load cases were selected, and an early material selection was performed, based on the failure index distributions of a solution using mean material properties.
The wing-fuselage structure was sized for two selected concepts, leading to two similarly performing designs. The designs showed large frames and cross-beams compared to conventional fuselage frames. This was the case as high stiffness was necessary in the components, due to high bending and compressive loads, primarily induced by pressurisation. The frames show a large potential for reusing the same frame design along the constant cabin, potentially leading to significant economical gains. Secondly, the floor and ceiling needed a far higher amount of stringers than the pressure shells, as the high compressive loads lead to an instability-prone structure. A structural mass estimation was performed, for the two concepts. The mass of the original and NLES concept were obtained to be 65.4e3 kg and 67.2e3 kg, respectively. Additionally, a mesh convergence analysis and input sensitivity analysis was performed.