Grigorios Dimitriadis

Designing lighter aircraft structures while satisfying strength, stability, and aeroelastic requirements remains a central challenge in preliminary aircraft design. Many early-stage aircraft structural optimization studies rely on beam, equivalent-plate, or semi-empirical structural representations, while many higher-fidelity studies focus on isolated components, particularly wings, rather than the simultaneous sizing of the wing, fuselage, and tail within a unified structural model. The present study addresses this gap by introducing a surrogate-assisted structural optimization framework for complete aircraft configurations. The framework is built around a fully parametric shell/beam-based finite element model of a representative long-range transport aircraft configuration with a high-aspect-ratio wing, coupled to aerodynamic loads computed using the subsonic source and doublet panel method. The proposed framework allows simultaneous sizing of the wing, fuselage, and tail under static strength, global buckling, and flutter constraints, while retaining a computational cost suitable for preliminary design studies. The optimization results are assessed through three independent surrogate-assisted optimization runs initialized from different sampling distributions and compared to a derivative-free optimization solver using the same computational budget. The independent runs identify feasible low-mass configurations with consistent convergence trends. The optimized design gives a full-aircraft modeled mass of 230.48 t, which remains below the reference maximum takeoff mass of 268 t and is consistent with the expected mass scale of representative long-range transport aircraft. Overall, the proposed framework provides a practical preliminary-design tool for exploring feasible full-aircraft structural layouts, identifying dominant load paths and active constraints, and generating candidate designs for subsequent higher-fidelity aeroelastic and certification-level refinement.

The steady and unsteady aerodynamic predictions of the Doublet Lattice Method (DLM) are routinely corrected for camber, twist, and transonic effects using steady reference data obtained from experiments or higher-fidelity calculations. In this work, a novel correction technique is presented for the unsteady compressible Source and Doublet Panel Method (SDPM) in order to correct for transonic effects. Although the SDPM is more computationally expensive than the DLM, it can predict complete steady and oscillatory pressure distributions around the entire surface of the exact geometry of a wing or body. Consequently, it does not require corrections for geometric properties such as camber and twist, only for transonic flow phenomena. The corrected predictions are compared to unsteady pressure distributions measured experimentally for two experimental test cases from the literature. It is shown that that the predictions of the SDPM correction method are accurate as long as the reference pressure data are of adequate quality. Finally, the SDPM with transonic correction is applied to the prediction of the flutter boundary of the AGARD 445.6 wing, and it is demonstrated that the proposed approach is capable of predicting the transonic dip.

The source and doublet panel method (SDPM) developed by Morino in the 1970s can model unsteady compressible ideal flow around wings and bodies. In this work, the SDPM is adapted to the calculation of aeroelastic solutions for wings. A second-order nonlinear version of Bernoulli’s equation is transformed to the frequency domain and written in terms of the generalized mode shapes and displacements. It is shown that the pressure component at the oscillating frequency is a linear function of the generalized displacements, velocities, and accelerations and can therefore be used to formulate a linear flutter problem. The proposed approach has several advantages over the usual doublet lattice method (DLM) approach: the exact geometry is modeled (including thickness, camber, and twist effects), the motion of the surface can be represented using all six degrees of freedom, the pressure calculation is of higher order, and the aerodynamic mass, damping, and stiffness terms are calculated explicitly. The complete procedure is validated using experimental data from the weakened AGARD 445.6 wing, a NACA 0012 rectangular wing with pitch and plunge degrees of freedom, and an experimental model of a T-tail, yielding flutter predictions that lie closer to the experimental observations than those obtained from the DLM, particularly in the case of the T-tail.

This work presents the development of closed-form expressions for the aerodynamic stability derivatives of wings and aircraft flying at subsonic compressible conditions using the unsteady subsonic 3D source and doublet panel method (SDPM). The stability derivatives are obtained directly from the frequency domain version of the nonlinear unsteady subsonic Bernoulli equation and can be calculated for negligible additional computational cost around the true geometry of the wing or aircraft for each frequency of interest. The methodology is validated by applying it to an experimental test case of a swept wing oscillating in yaw and sideslip in the wind tunnel, demonstrating that the values of the aerodynamic stability derivatives obtained from the SDPM are in good agreement with the experimental data at low and moderate angles of attack. Then, the methodology is applied to a blended-wing–body unmanned aerial vehicle configuration and the SDPM stability derivative predictions are compared to estimates obtained from both steady computational fluid dynamics calculations and the USAF DATCOM methodology. It is shown that the SDPM approach is accurate and can yield more information about the stability of the aircraft than the DATCOM, given that the latter was developed for conventional aircraft configurations.

Static aeroelastic deformations are nowadays considered as early as in the preliminary aircraft design stage, where low-fidelity linear aerodynamic modeling is favored because of its low computational cost. However, transonic flows are essentially nonlinear. The present work aims at assessing the impact of the aerodynamic level of fidelity used in preliminary aircraft design. Several fluid models ranging from the linear potential to the Navier–Stokes formulations were used to solve transonic flows for steady rigid aerodynamic and static aeroelastic computations on two benchmark wings: the Onera M6 and a generic airliner wing. The lift and moment loading distributions, as well as the bending and twisting deformations predicted by the different models, were examined, along with the computational costs of the various solutions. The results illustrate that a nonlinear method is required to reliably perform steady aerodynamic computations on rigid wings. For such computations, the best tradeoff between accuracy and computational cost is achieved by the full potential formulation. On the other hand, static aeroelastic computations are usually performed on optimized wings for which transonic effects are weak. In such cases, linear potential methods were found to yield sufficiently reliable results. If the linear method of choice is the doublet lattice approach, it must be corrected using a nonlinear solution.

A method is presented to model the incompressible, attached, unsteady lift and pitching moment acting on a thin three-dimensional wing in the time domain. The model is based on the combination of Wagner theory and lifting line theory through the unsteady Kutta–Joukowski theorem. The results are a set of closed-form linear ordinary differential equations that can be solved analytically or using a Runge–Kutta–Fehlberg algorithm. The method is validated against numerical predictions from an unsteady vortex lattice method for rectangular and tapered wings undergoing step or oscillatory changes in plunge or pitch. Further validation is demonstrated on an aeroelastic test case of a rigid rectangular finite wing with pitch and plunge degrees of freedom.