IFASD2024 Paper Submission & Registration
17 – 21 June 2024, The Hague, The Netherlands





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18:00   Poster session & drinks
18:00
1 mins
Deep Reinforcement Learning for Aeroservoelastic Control of Unmanned Multi-Body Aircraft with Geometric Nonlinearities
Zhuolin Ying, Ying Bi, Zijian Zhu, Chen Zhu, Xiaoping Ma
Abstract: The research delves into the unique challenges posed by high-altitude long-endurance (HALE) aircraft, particularly focusing on their extreme flexibility. The study explores the substantial deformation experienced by the lightweight and highly flexible wings of HALE aircraft during flight. This deformation not only alters aerodynamic calculation but also has a pronounced impact on the stiffness characteristics of the wings. To address issues of poor wind resistance during takeoff and landing, the paper introduces an innovative solution – an unmanned multi-body aircraft (MBA) design featuring wingtip docking. However, the introduction of multiple wingtip dockings brings about inevitable geometric nonlinearity issues, significantly influencing the aeroservoelastic design of the aircraft. Consequently, there is an urgent need for aeroservoelastic analysis for unmanned multi-body aircraft that takes geometric nonlinearity into account. The research methodology involves the establishment of a nonlinear finite element model tailored to flexible unmanned multi-body aircraft. The minimal state method is employed for rational aerodynamic approximation, allowing a comprehensive exploration of the coupling issues arising from geometric nonlinearity and aeroelasticity in the time domain. Additionally, an adaptive control law, incorporating principles of machine learning, is designed to actively suppress flutter. The findings of the research reveal that flexible unmanned multi-body aircraft considering geometric nonlinearity may exhibit complex behaviors such as limit cycle oscillations and chaotic phenomena under specific initial conditions. Importantly, the proposed adaptive learning control law emerges as a highly effective measure in reducing the critical flutter speed, offering robust support to enhance the overall flight stability of the aircraft. This comprehensive exploration contributes valuable insights into the intricate dynamics of flexible unmanned multi-body aircraft.
18:01
1 mins
Nonlinear Aeroelasticity Characteristics and Dynamics Response Analysis of Unmanned Multi-Body Aircraft
Zijian Zhu, Ying Bi, Chen Zhu, Zhuolin Ying, Jian Zhang
Abstract: Unmanned multi-body aircraft (MBA) represents an innovative aircraft configuration, wherein multiple aircrafts are interconnected at the wingtip. It has substantial applications in increasing aspect ratio, optimizing aeroelasticity characteristics, and mitigating takeoff and landing challenges. However, existing studies on MBA predominantly focus on the rigid-body level, overlooking the geometrical nonlinearities stemming from elastic deformation and the concentrated nonlinearities occurring at connections during flight. This paper centers around the configuration of MBA, utilizing a substructure method and modal reduction technique to establish a reduced-order nonlinear structural model. Additionally, describing function method is employed to formulate its nonlinear aerodynamic force model. The Lagrangian equation is then utilized to couple the nonlinear structural equation with the aerodynamic equation, resulting in an explicit dynamic expression for the high-dimensional aeroelastic system encompassing geometric and concentrated nonlinearities. Numerical simulation is used to Conduct parameter analysis and mechanism exploration based on the established aeroelastic system equations, which reveals correlations and impact mechanisms of various nonlinear factors under the MBA configuration, particularly affecting dynamic phenomena such as limit cycles and chaos. The findings suggest that both concentrated nonlinearities and geometrical nonlinearities significantly contribute to the system's bifurcations, limit cycles, and other dynamic responses under the MBA configuration. Specifically, when the number of joined aircraft is limited, concentrated nonlinearity assumes a dominant role, whereas with an increased number of joined aircraft, geometric nonlinearity takes precedence. Furthermore, adjustments to the number of connected aircrafts and model parameters at the connection point unveil the substantial influence of both types of nonlinear factors on the overall flutter characteristics of the system. This study delves into the vibrational characteristics and nonlinear dynamic behavior of unmanned MBA, providing a reference analysis that contributes to the development of this configuration.
18:02
1 mins
Fluid-structure interaction of supersonic membrane structures
Zexuan Yang, Faying Zhang, Chao Yang, Bing Feng NG, Zhigang Wu
Abstract: In order to enable wing morphing (e.g. change in camber or folds) without incurring additional weight to the aircraft, lightweight flexible materials such as membrane is needed. At present, the research on fluid-structure coupling of membranes is mainly focused on airbags and parachutes. However, there is a lack of research on supersonic membrane structures. In this study, a rectangular membrane of length 0.6m and width 0.4m is investigated. A finite element modeling method using shell elements was proposed to study membranes deformation. Establishing a shell model with a very small thickness can approximate a membrane structure due to its low bending stiffness. The method is more effective in solving deformation than using membrane elements directly. Examples are provided to verify the effectiveness of the proposed methodology. Moving a step further, in order to capture aeroelastic effects of the membrane in supersonic flow, a fluid-structure coupling framework was established. It involves an aerodynamic module (involving the piston theory or commercial CFD solvers), a structural module (through ABAQUS) and the coupling is performed through an in-house code developed in Python. Two aerodynamic force solvers are used to analyze the static deformation and dynamic response of the rectangular membrane, and the results are consistent. Subsequently, A parametric analysis is performed on the dynamic pressure and Mach number for their effects on membrane structures. A theoretical analysis is also conducted on a specific phenomenon where the membrane structure vibrates randomly at small angles of attack, but deformation amplitude converges to a common value at large angles of attack. Finally, considering the fact that membrane structure is relaxed, a model with initial relaxation was designed. The results prove that the initial relaxation has a great influence on the deformation, and the influence of various shape relaxations on the results are discussed. Research highlights are as follows: 1. A finite element modeling method using shell elements to simulate membrane structures was proposed. 2. A general framework for fluid-structure coupling of flexible structures in supersonic flow was established, and the aerodynamic force can be solved by piston theory or Fluent. 3. The dynamic response of membrane structure has two forms, the vibration is random at small AoA, and the deformation amplitude tends to converge to a common value at large AoA. 4. A membrane structure with initial relaxation was designed, and the effect of the initial relaxation size and form of deformation were analyzed.
18:03
1 mins
Structural optimization of a high-aspect-ratio wing with post-flutter constraint
Zhiying Chen, Yang Meng, Zhiqiang Wan, Changchuan Xie, Chao Yang
Abstract: Highly flexible wings of large aspect ratio are commonly found in high-altitude long-endurance aircraft, large deformations of such wings can lead to structural and aerodynamic nonlinearity, which includes geometric nonlinearity and stall. The relation between structural nonlinear force and stall aerodynamics affects the post-flutter behaviour of the wing. Although there has been a significant amount of work on post-flutter analyses, post-flutter constraints, which are critical for preventing undesirable subcritical limit cycle oscillations, have been rarely considered in design optimization. In this paper, the structural optimization of a clamped high-aspect-ratio highly flexible wing with post-flutter constraint is presented. Firstly, a nonlinear unsteady aeroelastic model is built by coupling nonlinear structural and aerodynamic model. Strain-based beam theory is used for efficient structural modelling, and strip theory combining the ONERA dynamic stall model is used for aerodynamic modelling. Then, the equilibrium position for a specific angle of attack is solved, and the flutter speed is determined by performing an eigenvalue analysis of the linear equation at the equilibrium position. The nonlinear perturbation solution near the flutter speed is calculated using the method of multiple scales, from which a scalar variable can be extracted to dictate the characteristic of the limit cycle. Finally, several design variables which controls the stiffness and inertia characteristics are defined, and the gradient-based optimization is carried out under flutter, post-flutter and other constrains. The results shows that the current methods employed can accurately and efficiently calculate the nonlinear characteristics of the system, meeting the requirements for design optimization. And through optimization, the limit cycle can be transformed from subcritical to supercritical, which ensures the effectiveness of the flutter constraint.
18:04
1 mins
A state-space rigid-elastic coupling aeroelastic model with geometrical accurate boundary condition
Nongyue Gao, Changchuan Xie, Chao An, Chao Yang
Abstract: Flying-wing aircraft has become an ideal choice for advanced aircraft design due to its excellent aerodynamic characteristic and strong loading capacity. However, this kind of aircraft usually has poor flight stability and controllability because of the small pitch inertia. In addition, with the widespread use of composite materials on aircraft, the stiffness of the aircraft is gradually decreasing. The minimum structural elastic natural frequency of the aircraft is becoming lower and closer to the maximum frequency of rigid-body flight modes. These factors lead to the occurrence of rigid-elastic coupling aeroelastic instability phenomenon and the flutter speed may become much lower compared to the one without considering rigid body flight modes, which can be a crucial safety disadvantage of applying flying-wing layout and need appropriate analysis model urgently. Based on the flight dynamic model, flight kinematic model, the state-space unsteady vortex lattice method (UVLM), and the modal-superposition structural dynamic model, a rigid-elastic coupling aeroelastic model can be built, which is fully coupled at each time step and can be written in state space form. To obtain a more accurate aerodynamic response, the UVLM has been improved with geometrical accurate boundary condition, which can consider the impact such as structural deformation, rudder deflection, and the change of unsteady wake through integral boundary changing and normal vectors deflection. According to the coupling model, the analytical solution of the model can be derived. This can be a credible reference when computing the time-domain response. A flying-wing model was built for the rigid-elastic coupling aeroelastic analysis. The results of rigid-elastic flutter analysis were proved to be consistent with commercial software MSC. Nastran and ZAERO. In addition, a significant decrease in critical speed after considering the rigid-body flight mode was obtained, which revealed the unignorable coupling effect between flight mode and structural elasticity. Those results show that the state-space rigid-elastic coupling aeroelastic model with geometrical accurate boundary condition can provide a kind of powerful and reliable analysis method for flying-wing aircraft aeroelastic design to improve the flight safety.
18:05
1 mins
A CFD/CTD/CSD based aero-thermo-elastic framework for full-vehicle scale analysis
Liang Ma, Zhiqiang Wan, Xiaozhe Wang, Keyu Li, Chang Li, Longfei He
Abstract: Due to the significant multidisciplinary coupling mechanism inherent in hypersonic flight mission, unnecessary and wasteful trade-off in vehicle performance will be cost if complex load distribution and aerodynamic heating effect are neglected at the early stage of design. This paper establishes an CFD/CTD/CSD based aero-thermo-elastic framework for analysis of the full-vehicle scale. The loose coupling strategy is chosen in this framework to reveal the specific efforts of each disciplinary, and the RBF-TFI method is introduced for mesh deformation. This study is carried on the rudders assembled on a hypersonic missile, with the high-fidelity aerodynamic data of the full-vehicle model extracted by CFD and only the component deformation of the rudders extracted by FEM. This paper demonstrates the aero-thermo-elastic effects of those factors concealed by engineering algorithms, with the influence mechanism revealed from the results discussed.
18:06
1 mins
Aeroelastic optimization of commercial aircraft considering high-precision aerodynamic performance
Keyu Li, Chao Yang, Xiaozhe Wang, Zhiqiang Wan, Liang Ma, Chang Li
Abstract: Traditional design methods tend to introduce the consideration of aeroelasticity effects on the aircraft late in the design process, leading to increased mechanical mass, reduced aerodynamic performance. Therefore, it is crucial to fully consider the effects of aeroelasticity during the preliminary design stage. The performance of the wing can be effectively improved by using the aeroelastic tailoring. The static aeroelastic analysis in this approach obtains aerodynamic forces by solving the linearized aerodynamic potential flow theory. However, it may not accurately calculate the drag of the wing. The use of high-precision aerodynamic calculation methods presents a problem of time-consuming calculation, which is difficult to apply in tailoring design. This paper proposes an aeroelastic optimization method that considers high-precision aerodynamic. The Euler equations are solved to obtain high-precision aerodynamic forces, and a viscous correction method is employed to improve the accuracy of drag results. To tackle the issue of low efficiency, aerodynamic force prediction is achieved through a Kriging surrogate model based on wing torsion angle and deflection. Next, the stiffness of the box section of a commercial aircraft wing is optimized. The optimization problem focuses on minimizing the structural mass of the wing by designing the thickness of the wing skin and web layup, as well as other variables. To obtain the global optimal solution, a genetic sensitivity hybrid algorithm is utilized. Aeroelastic constraints such as strength, deformation, and aileron efficiency are considered, and high-precision aerodynamic constraints are introduced to obtain the cruise profile with the optimal lift-to-drag ratio. The results indicate that the surrogate model has an error rate of only 0.32%, enabling efficient prediction of aerodynamic forces. Furthermore, the final wing configuration of the tailoring design reduces mass by 20 kg compared to the initial configuration while satisfying aeroelastic constraints, resulting in efficient aeroelastic optimization.
18:07
1 mins
Testing and Analysis of a Simplified Nonlinear Horizontal Stabilator of an Aircraft
Nichoas Stathopoulos, Viresh Wickramasinghe, Devon Downes
Abstract: The horizontal stabilator (H-Stab) of a fighter aircraft develops significant freeplay as the control surface mechanism ages. Modal tests on a simplified model of a H-Stab structure have been performed to investigate the nonlinear dynamic behavior of the spindle and bushing interface related to varied levels of wear. Extensive testing on this simplified model successfully revealed the critical changes in its dynamic properties. The H-Stab structure was represented by a rectangular aluminum block (labeled exciter plate), attached to a steel rod as a spindle fixed to a mast. A set of bushings with varying diameters was used at the inner edge of the plate to simulate different levels of freeplay, which was controlled in the order of 0.034° to be comparable with MIL-A-8870 standard. Steel bushings were fabricated that represent three freeplay cases, the tight fit, nominal tolerance, and double tolerance. The modal parameters of the first three modes of the structure were identified in three test configurations, and each configuration involved testing using two excitation loadings, namely burst random and swept sinusoid input. Each load type included several input levels. Comparing modal frequencies and damping ratios revealed that the spindle and bushing interface demonstrated nonlinear dynamic behavior. Relative to the tight fitting bushing, the nominal tolerance case experienced a frequency reduction of between 4 to 17 % in the first three modes, while the damping ratio, moved from 0.99% to 1.56% for the first mode. The double tolerance case exhibited frequency reductions of between 1 to 17%, while the damping ratio increased from the original value of 0.99% to 1.72%. In summary, with the spindle freeplay within the range of MIL-A-8870 standard, the modal frequencies of the H-Stab major modes decreased with increased spindle freeplay and increased input load levels, while the damping ratio increased with increased spindle freeplay and increased input load levels. These results indicate that tracking the modal parameters of the H-Stab during the aircraft’s service may be a promising approach in determining when the structure has undergone excessive wear.
18:08
1 mins
Reduced order modeling analysis for flexible wing using a co-rotational beam element
Yali Shao, Changchuan Xie, Chao An, Duoyao Zhang, Yuhui Zhang
Abstract: The objective of this paper is to develop a reduced order modelling (ROM) method suitable for aeroelastic analysis with high efficiency and sufficient fidelity. The method is applied for solving the static aeroelastic problems of highly flexible wing containing geometric nonlinearities. The structural ROM method is based on equations derived from the Galerkin approach to solve the geometric nonlinear dynamics in a weak form, in which the explicit calculation of nonlinear stiffness is not practical. Based on dynamic response data samples, nonlinear stiffness coefficients in structural dynamics equation are identified based on the fast Fourier transform (FFT) and the harmonic balance nonlinearity identification technique. And dynamic response data samples can be gained by the commercially finite element (FE) software, however, for aeroelastic analysis, the computational cost usually takes a lot. In this paper, we adopt the co-rotational (CR) finite element to shorten the simulation time and better deal with different load cases. The total and updated Lagrangian formulations for geometrically nonlinear analysis are used in most commercially available FE software. Their accuracy is enough for the engineering application, but are numerically quite demanding and may exhibit convergence issues. However, the CR finite element is computationally more efficient for the case of large displacements and/or large rotations in a geometrically nonlinear analysis, as the rigid part is purged from the total motion. More specifically, the main idea of CR theory is to decompose the motion of the element into rigid body and pure deformational parts, through the use of co-rotational frame which continuously rotates and translates with the element. In this paper, the CR beam element is adopted for the structural simulation of wing. The fidelity of the method (ROM with CR) is verified against ROM with MSC Nastran (a widely-used multidisciplinary structural analysis solver), on a high aspect ratio wing. Compared to solutions figured out by ROM with MSC Nastran, the result of ROM with CR shows high efficiency and adequate accuracy. The final paper will include detailed simulated results of ROM with CR and comparison with MSC Nastran.
18:09
1 mins
Research on nonlinear substructure method and aeroelastic analysis of large flexible aircraft
Rui Zhao, Chao An, Changchuan Xie
Abstract: Large flexible aircraft represented by high-altitude long-endurance UAVs and flying wing UAVs have broad application prospects in many fields. Due to specific requirements for flight performance, such aircraft has large structural flexibility. Under the aerodynamic loads, large flexible aircraft will undergo large elastic deformation, which will bring geometric nonlinear aeroelastic problems. Tradtional linear aeroelastic analysis methods cannot meet the aeroelastic analysis requirements of large flexible aircraft. In this paper, we use the nonlinear substructure method to establish a reduced-order model of a large flexible aircraft and calculate the aeroelastic response. Meanwhile, discussion and variation of calculation results will be performed. Aiming at the large flexible aircraft components, the natural structural mode is used as the basis to establish a nonlinear structural reduced-order model, and regression analysis method is used to solve the nonlinear stiffness coefficients. On this basis, combined with the surface aerodynamics and three-dimensional surface spline interpolation methods, the geometric nonlinear aeroelastic analysis of the wing components is carried out. Compared with the aeroelastic analysis results based on the nonlinear finite element method, the deviation of the analysis results is small. The nonlinear structural reduced-order model can balance the solution efficiency, accuracy and structural application range. Aiming at the whole aircraft structure of large flexible aircraft, aircraft is separated into nonlinear wing components and linear fuselage components, and a hybrid interface modal synthesis method considering geometric non-linear factors in the domain is established based on substructure method. Combined with the nonlinear structural reduced order model of the wing, a low-order structure model of the whole aircraft is formed. On this basis, the static aeroelastic trim and rigid/elastic coupling stability analysis of large flexible aircraft are carried out. In this paper, the structural modeling method and geometric nonlinear aeroelastic analysis of the complex engineering model of large flexible aircraft are carried out, which verifies the applicability of the structural modeling method and aeroelastic analysis established in the complex engineering model, and shows its great advantages in computational efficiency. Key words:Large flexible aircraft, Aeroelasticity, Geometric nonlinearity, Reduced order modeling, Nonlinear substructure method, Modal synthesis method
18:10
1 mins
Hypersonic flutter analysis based on three-dimensional local piston theory
Gaozhan Wang, Changchuan Xie, Chenyu Liu, Chao An
Abstract: Hypersonic vehicles have become a research hotspot these years. This kind of aircraft often has a slender body layout using a thin-walled structure and lightweight materials, which causes special and significant aeroelastic flutter problems. In this research, a hypersonic flutter analysis method based on three-dimensional local piston theory(3D-LPT) is developed and verified with wind tunnel test results. Using a local coordinate system of wall surface, local piston theory can be extended to 3D. Combined with Euler-CFD calculation, a high-precision three-dimensional discrete hypersonic unsteady aerodynamic force model is established in the body axis system Oxyz: where is the outward normal unit vector of a wall surface element and denotes the displacements of the element. , and are the projections of surface local flow velocity in the x, y, and z directions. and are the local density and sound speed, respectively. These local flow parameters are determined by CFD calculation. The generalized aerodynamic force can be expressed explicitly in the time domain as a combination of generalized aerodynamic influence coefficient (AIC) matrix A and generalized coordinate q in the small-disturbance range around a specific flow state. Therefore, an aerodynamic-structure tight coupling aeroelastic flutter equation can be established in the state-space form: where and are the generalized mass matrix and generalized stiffness matrix, respectively. Based on this state-space form flutter equation, Lyapunov's first method can be used to analyze the flutter stability of the aeroelastic system. As shown in the figure below, a supersonic fin model with wind tunnel test results is used to verify the accuracy of the proposed method. Comparing the numerical results calculated by the 3D-LPT-based method with wind tunnel test results, the numerical errors of flutter speeds at all AOA cases are less than 11%. This 3D-LPT-based hypersonic flutter analysis method can analyze 3D complex objects while suitable for conditions with high AOA or wide Mach number range. Combined with CFD calculation, aerodynamic nonlinear effects can be considered. Based on the aerodynamic-structure tight coupling model, the proposed method has high accuracy and efficiency. By comparing with wind tunnel data, the effectiveness of the proposed method is proved.
18:11
1 mins
Gust response analysis of supersonic aircraft based on three-dimensional piston theory
Chen Song, Changchuan Xie, Chenyu Liu, Yang Meng
Abstract: Though the influence of atmosphere disturbance (gust) on low-speed aircraft has been studied a lot since the disintegration of the Helios Prototype, the gust load on supersonic vehicles still needs further discussion, along with efficient analysis methods. On the other hand, with the development of hypersonic glide vehicles, the assembly of warhead and booster usually has a lower basic frequency than supersonic fighters and therefore, is more sensitive to atmosphere disturbance, which makes it necessary to analyze the gust response of supersonic aircraft. In this paper, first-order piston theory is applied in the local frame of 3-D aero mesh to calculate the unsteady aerodynamic force caused by elastic vibration and gust, which could be written into matrix form using modal coordinates and interface interpolation method, where and are aerodynamic damping and stiffness respectively. Integrate the aerodynamic model into the structural dynamics, and the aeroelastic dynamic equation could be formed, with the notation : where are the generalized mass, damping, and stiffness matrixes of the structure. In most research, the ODE is solved by numerical methods like Ruuge-Kutta methods though semi-analytical solution exists: , with . To justify the proposed method and semi-analytical solution, the dynamic response of a simple wing model under 1-cos gust is calculated by both the proposed method and ZAERO at Ma 3.0: (a) Fig. (a): Diagram of the wing model (b): Acceleration of monitor point (b) In conclusion, the gust response analysis based on three-dimensional piston theory has great consistency with commercial software, yet the proposed method could be applied to high-resolution 3-D mesh. The analytical solution gives almost identical results to Runge-Kutta with 1/20 of the time spent, which makes it practical to analyze complex objects.
18:12
1 mins
Influence of nonlinear aerodynamic effects on high aspect ratio aircraft model
Álvaro Antonio García Quesada, Pedro José González Ramirez, Guilherme Chaves Barbosa, Gerrit Sybe Stavorinus, Flávio José Silvestre
Abstract: ABSTRACT This paper studies the influence of stall effects and follower forces on High Aspect Ratio (HAR) aircraft model for different load cases and aircraft flexibility levels. The Modeling and Simulation Group Toolbox created by the Chair of Flight Mechanics, Flight Control and Aeroelasticity of Technical University Berlin (ModSiG-FMRA) is used for modeling and analyzing the aerodynamic nonlinearities on the TU-Flex aircraft model. TU-Flex is a scaled flight demonstrator embodying a HAR transport aircraft configuration with modular construction, allowing different flexibility levels (for instance, in this paper a flexible and very flexible set). The ModSiG-FMRA framework uses mean axes formulation, modal superposition for structural dynamics, therefore considered linear, nonlinear flight mechanics, and quasi-steady or unsteady strip theory for incremental aerodynamics due to elastic deformations. The aerodynamic formulation permits to incorporate stall effects and/or follower forces. The follower forces effect does not show an influence on the load cases simulations of the TU-Flex flight envelope while considering geometrically linear structure. On the other side, stall effects influence the TU-Flex dynamic behaviour by the coupling rigid body and structure, not captured by linear aerodynamics formulations. This coupling mechanism is emphasised by increasing the flexibility level.
18:13
1 mins
Aeroelastic State space modeling for flight flutter test simulation using expanded mode shapes from Ground Vibration Test data
Hemalatha E, Gourav Kumar Dutta, Mohamed Nishad K, Pavan Kumar Dasari
Abstract: During the design stage of an Aircraft, the global finite element model (GFEM) is used to predict modal characteristics. Prior to first flight, Ground Vibration Test (GVT) is conducted and the experimental modal data is used to validate and update the finite element model which is then used for aeroelastic analysis. However, experimental mode shape vectors exist only at a limited number of Degrees of Freedom (DoF) corresponding to the measurement sensor locations. Expansion of experimental vectors is often needed for better visualization of modeshapes/deformation shapes as well as correlation purposes. The data from the reduced order experimental model when expanded to the GFEM gives a modeshape matrix which is more realistic. This can be used to replace the computed modal matrix for further flutter or aeroelastic response analysis. Of the many techniques for model order reduction and expansion, the System Equivalent Reduction and Expansion Process (SEREP) is adopted for the present study. Using this, mode shapes obtained from ground vibration test (GVT) of a typical Fighter aircraft configuration are expanded back to the Global Finite element model (GFEM) of the full aircraft. These are then used to carry out flutter computations using MSc/Nastran Aeroelasticity module. Next, the Nastran generated unsteady aerodynamic loads are transformed to Laplace domain using Matrix polynomial approach (MPA) for creation of an aeroelastic state space model. The state space representation offers a greater degree of flexibility to model other system dynamics such as actuator transfer functions. The state space model is verified by comparing flutter results with those obtained from Nastran. A flight flutter test simulation is done using this aeroelastic state space model by applying excitation input through the control surface actuators and extracting output responses at wing and fin tips and control surfaces. The simulation confirms the adequacy of the excitation magnitudes and frequency bands prior to actual flight flutter tests.
18:14
1 mins
Aerodynamic and aeroelastic analysis of propeller-wing system with a unified UVLM framework
Ruijie Niu, Changchuan Xie, Zhitao Zhang, Chao An
Abstract: The aerodynamic and aeroelastic analysis on the propeller-wing tractor configurations has gained a lot of focus as the rapid development of electric distributed propulsion vehicles. Complex aerodynamic and structural couplings of propeller-wing system must be modelled properly and solved together as an aeroelastic problem, especially for flexible wings. Low to medium fidelity approaches remain popular due to high precision methods (CFD) are still expensive for aeroelastic problems. The Unsteady Vortex Lattice Method (UVLM) can relatively guarantee accuracy with less effort and has the ability to consider aerodynamic interference between wing and propellers. In this paper, a unified framework is presented to analyse the aerodynamic and aeroelastic of propeller-wing system with two paralleling propellers, providing the basis for further studies of distributed propulsion wing. The wing, propellers and their wakes are modelled meanwhile based on UVLM to capture the mutual aerodynamic interference without empirical formulations. Vortex strengths can be determined by boundary condition: (1) Numeric comparisons with the RANS (Reynolds-averaged Navier-Stokes) approach is made both on the two isolated propellers and the proposed system. As for structural analysis, propellers are considered as rigid bodies while wing is modelled by linear finite element model. The aerodynamic and structural models are loose coupled to calculate the aeroelastic response in time-domain. Examples are also included to illustrate the usefulness of the proposed method. In the future, we will investigate the effect of relative parameters on the stability of the propeller-wing system, such as numbers, locations and rotational directions (same or opposite) of propellers and so on. Gaining a deeper insight into the aerodynamic performance, vibration, stability, control and noise of the coupled system is also a challenging task.
18:15
1 mins
Measurement and Control System Design for Propeller-Wing System Experiments
Wei Wang, Zhitao Zhang, Changchuan Xie, Chao Yang, Kunhui Huang
Abstract: An integrated measurement and control system suitable for ground and wind tunnel testing of a distributed propeller-wing system was designed. The real-time measurement, recording and display of several indicators including the propeller aerodynamic load, wing root load, propeller rotation speed, wing structural deformation and wing surface pressure during the experiment were obtained by using a six-component balance, an optical sensor, a data acquisition device and graphic display. Computer software was also utilized to process the collected test data, which were then compared with the numerical simulation results to verify the effectiveness and reliability of the measurement and control system. Results of high-precision computational fluid dynamics (CFD)/computational structural dynamics (CSD) aeroelastic simulations and Nastran SOL144 calculations used in this study are collected. In terms of vertical displacement of the wing tip and wing deformation distribution tested during the experiment. It is observed that the simulation results obtained in this study are in consistency with the results of wind tunnel tests using commercial software, proving the effectiveness of the experiment design.
18:16
1 mins
On the preparation of a flexible real-time aircraft model for a flight simulator
Gefferson Silva, Bernd Boche, Hannes Wilke, Flavio Silvestre
Abstract: Before flight test campaigns are pursued, handling qualities analyses of new aircraft designs are typically performed by means of a pilot-in-the-loop full-flight simulator. This also includes the forthcoming generation of greener aircraft, for which environmental damages resulting from aircraft emissions is mitigated. However, this type of air-vehicle requires special treatment in flight dynamics modeling and simulation, due to the interactive coupling between rigid-body dynamics and aeroelasticity. The extended model simulations result in an expressive increase in computational costs, which makes real-time simulations of flexible aircraft a current challenge. Therefore, the work at hand discusses an approach to reduce the computational consumption required for simulating slightly flexible aircraft models, in order to be embedded into a flight simulator. The mean-axes formulation is used to represent the coupled relations between flight mechanics and flexibility effects. The incremental forces and moments expressions are then modified algebraic-wise to obtain an equivalent but computationally efficient formulation to those expressions.
18:17
1 mins
Nonlinear aeroelastic response of strut-braced high aspect ratio wings
Fahed Mohd, Mohammadreza Amoozgar, Stewart McWilliams
Abstract: High Aspect ratio Wings (HARW) are nowadays being used in the commercial aerospace industry due to their high lift-to-drag ratios and longer flight ranges. This is because aerodynamic-induced drag is reduced by increasing the aspect ratio of the wing, thereby resulting in more economical flight operations [1]. However, owing to these aerodynamic advantages, there are structural design constraints such as higher stress concentration on the wing root, and higher structural flexibility which causes the wing to be more prone to larger deflections, which in turn affect the overall aeroelastic behaviour [1]. To eliminate these downsides, Strut-Braced High Aspect Ratio Wings (SB-HARW) were proposed which may be proved to be more advantageous than the traditional wings [2]. Nonetheless, large deflections of SB-HARW can have structural nonlinearities associated with them, which can significantly alter the aeroelastic stability and response [3]. In, addition to accurately capturing their structural dynamics, their aeroelastic behaviours such as flutter instability onset, post-flutter response, etc. must be considered as well [4]. The studies related to assessing the post-flutter nonlinear aeroelastic behaviour of SB-HARW are very limited thus hindering its practical implementation in the commercial aviation industry. Therefore, this paper aims to study the nonlinear aeroelastic response of SB-HARW in the post-instability region considering various wing and strut nonlinearities. The SB-HARW is modelled as a cantilever nonlinear beam while the strut is modelled using a nonlinear spring connected between the wing and fuselage. The final paper will present the methodologies used for the analysis in more detail alongside with the verifications of the developed aeroelastic model. The objective will be to investigate the effect of strut spanwise location, chordwise offset, number of struts and strut stiffness nonlinearity on the flutter instability and post-instability behaviour of a wing with cubic nonlinearity. References [1] Afonso, Frederico, et al. "A review on non-linear aeroelasticity of high aspect-ratio wings." Progress in Aerospace Sciences 89 (2017): 40-57. [2] Sohst, Martin, et al. "Optimization and comparison of strut-braced and high aspect ratio wing aircraft configurations including flutter analysis with geometric non-linearities." Aerospace Science and Technology 124 (2022): 107531. [3] Patil, M. J., & Hodges, D. H. (2004). On the importance of aerodynamic and structural geometrical nonlinearities in aeroelastic behavior of high-aspect-ratio wings. Journal of Fluids and Structures, 19(7), 905-915. [4] Liu, Y., Xie, C., Yang, C., & Cheng, J. (2016). Gust response analysis and wind tunnel test for a high-aspect ratio wing. Chinese Journal of Aeronautics, 29(1), 91-103.
18:18
1 mins
Experimental And Numerical Activities For Aeroelastic Analysis Of An Aluminum Wing
Davide Mastrodicasa, Silvia Vettori, Massimiliano Chillemi, Alessandro Laurini, Aleli Sosa Chavez, Emilio Di Lorenzo, Karl Janssens
Abstract: In the aerospace field, comprehensive monitoring of systems structural behavior is imperative for enhancing structural safety, optimizing maintenance protocols, and predicting remaining useful life. New significant challenges have been raised in recent years by market innovations such as the use of composite materials and high aspect ratio geometries. These solutions generate light-weight components featuring larger deformations with increasing fluid loading. To deal with these features, this work pursues the long-term goal of establishing a validated procedure that combines the use of different methodologies for Fluid-Structure Interaction (FSI) problems. A multi-physics environment, including experimental and numerical analyses, is constructed to allow for FSI analysis. The case study presented in this work concerns wind tunnel testing of an aluminium wing featuring a NACA 0018 profile. An extensive test campaign has been conducted on such specimen in the wind tunnel of the University of Twente. During the measurements, the wing has been tested in a clamped-free configuration under different flow velocities and angles of attack (AoA). The unit under test has been instrumented with different measurement systems such as: i) strain gauges for measuring the structural deformation of the wing under wind loading, ii) pressure ports for gaining knowledge of the pressure value around the wing airfoil, iii) speckle pattern for Digital Image Correlation (DIC). Two 2 MPx high-speed cameras running at 800 fps have been used to measure the full-field displacements over the structure during wind excitation. The retrieved displacements have been compared to virtual sensing estimated displacements for validation purposes. A Finite Element Model (FEM) has been built in Simcenter 3D for numerical identification of the wing dynamic properties. The FEM has been validated and updated to match the experimental modal parameters identified via Experimental Modal Analysis (EMA) during an impact test conducted prior to wind tunnel testing. A Computational Fluid Dynamics (CFD) analysis has been conducted in Simcenter STAR CCM+ and validated via experimental pressure values. This work further combines the validated structural and aerodynamic models into a comprehensive FSI framework for aeroelastic analysis.
18:19
1 mins
Mathematical method for prediction aeroelastic phenomena and multidisciplinary optimization lifting surfaces of flight vehicle at preliminary stage design
Oleh Havaza, Vitalii Sukhov
Abstract: Every year, the air transportation market increases the requirements for aircraft performance in order to obtain greater profits. Satisfaction of this requirement is the main purpose for aircraft manufacturer. One of the ways to achieve this purpose is improving the design process by developing and implementing new approaches and tools for modelling different phenomena (in considered case – Aeroelastic phenomena) inherent to flight vehicles. The purpose of this report is to present a modern mathematical method that allows modelling static and dynamic Aeroelastic Phenomena by a symbolic operation, in contrast to the numerical methods that are widely used today. The theory of this method based on science analogies approach which allow to present interaction (considering 6 DOF) between aeroelastic forces in analytical formulation, using modern Computer Algebra System tools and as results exclude iteration calculations which inherent inversing and eigenvalue extraction of large dimension matrixes. Approach of this method based on reduced order modelling principle and main idea is presenting lifting surface (wing, blade, etc.) as a principal scheme which look as parallel, serial and star connection of three types of elements: aerodynamical, elastic and inertial. Each element describes by respective matrixes of parameters with maximum dimension of 6x6 for 6 DOFs (3 translations + 3 rotations). Analysis is performing by transformation scheme (connection nodes condensation) and finding equivalent matrix of “Aeroelasticity” using recurrent equations for each type of connection (analogical with electrical circuit). Described method was validate by modelling: Wing load distribution, Divergence, Effectiveness of control surface and Flutter. Finally, gotten Math model describes dependences between design variables and aeroelastic characteristic in explicit symbolic form which allow performs wide fast parametric investigation at preliminary stage design. Additional this model will be useful for structure optimization process using analytical methods and for machine learning and will allow expand using artificial intelligence in aerospace structure design.
18:20
1 mins
Aeroelastic characteristics of plane-symmetry vehicle main wing trailing edge rudder at high angle of attack
Guo Li, Gongde Li, Yingyou Hou, Junlin Shen, Pu Xing, Chen Ji
Abstract: The plane-symmetry shuttle vehicle experiences a long period of supersonic high angle of attack flight stage during its reentry. It is hard for the aerospace-vehicles structure to sustain long-time supersonic flight at high angle of attack. Therefore, the aeroelastic response characteristics of the trailing edge rudder surface of the typical plane-symmetric shuttle configuration require in-depth analysis. In this paper, the analysis and calculation of the trailing-edge rudder of at the high angle of attack is carried out in high speed. The numerical CFD/CSD coupling method is employed in the analysis. The calculation results show that the symmetrical and antisymmetric deflection modes of the trailing-edge rudder are unstable, and the aerodynamic damping identified by the ARMA method is less than 0, which indicates the aeroelastic instability. Therefore, the dynamic stability of the structure during the reentry period needs to be considered in the design of the plane-symmetric aircraft.


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