Full Aircraft Model Research Articles

Enhanced Unsteady Vortex Lattice Aerodynamics for Nonlinear Flexible Aircraft Dynamic Simulation

Several modeling extensions and numerical improvements to the unsteady vortex-lattice method are discussed for the simulation of very flexible aircraft dynamics. In particular, fuselage aerodynamics are included by means of linear source panels, polar corrections are incorporated on the aerodynamics under potentially large deformations, and a new wake discretization scheme is derived to accelerate the computations. The theory behind each approach is detailed and successfully verified on representative test cases. The resulting modeling environment is then exemplified in a study of the flight dynamics of a flexible aircraft demonstrator model. Results indicate that, for large wing deformations, a noticeable effect of fuselage aerodynamic interference appears on the wing aeroelastic behavior. The polar corrections improve the drag estimations and allow capturing lift decrements due to stall onset at high induced angles of attack, which then affects the gust loads. The wake discretization scheme enables a reduction of the computational time for dynamic simulations on a full aircraft model by about 20% while keeping the arising error negligibly small.

Aeroacoustic testing on a full aircraft model at high Reynolds numbers in the European Transonic Windtunnel

This paper presents an end-to-end approach for the assessment of pressurized and cryogenic wind tunnel measurements of an EMBRAER scaled full model close to real-world Reynolds numbers. The choice of microphones, measurement parameters, the design of the array, and the selection of flow parameters are discussed. Different wind tunnel conditions are proposed which allow separating the influence of the Reynolds number from the Mach number, as well as the influence of slotted and closed test sections. The paper provides three-dimensional beamforming results with CLEAN-SC deconvolution, the selection of regions of interest, and the corresponding source spectra. The results suggest that slotted test sections have little influence on the beamforming results compared to closed test sections and that the Reynolds number has a profound, non-linear impact on the aeroacoustic emission that lessens with increasing Reynolds number. Further, sources show a non-linear Mach number dependency at constant Reynolds number but are self-similar in the observed Mach number range. The findings suggest that it is possible to study real-world phenomena on small-scale full models at real-world Reynolds numbers, which enable further investigations in the future such as the directivity of sources.

Wind tunnel test of gust load alleviation for a large-scale full aircraft model

Wind tunnel test is an important way to test the performance of Gust Load Alleviation (GLA). At present, some component-level wind tunnel tests have been carried out in big aviation countries, but there is a lack of full aircraft model GLA tests. In this study, a set of large-scale GLA test system in low-speed wind tunnel is developed, which includes a gust generator, a five-degree-of-freedom suspension system, a full elastic aircraft model with control system, and gust load measuring devices. Two control schemes based on closed-loop feedback control and open-loop feed forward control are respectively designed and verified by the full-model GLA tests in the wind tunnel. The experimental results show that the designed gust generator can generate a stable wideband, wide-area gust wind field; the suspension support system can sustain static and dynamic stability during wind tunnel test, and enables the model moving in the horizontal, vertical, pitching, rolling and yawing directions. The results show that the closed-loop feedback control can obtain good control performance of reducing the peak values of elastic vibration response induced by gust excitation, but has little effect on suppressing the rigid body motion excited by low-frequency gust, while the open-loop feed forward control presents superior performance in alleviating the high-frequency elastic vibration as well as the low-frequency rigid body motion with more than 40% overall reduction rate.

Analysis of the application of fuselage skin reinforcements with beam element representations in flexible full aircraft models for ditching simulations

The aim of this paper is the application of beam element representations for structural skin reinforcements in flexible full aircraft FE models used in ditching simulations. To verify this approach, it was initially analyzed on flexible reinforced bottom-aircraft panels under guided ditching conditions, considering also structural mesh size variations and partly corresponding fluid mesh densities. For this analysis two different numerical methods were used for comparisons, the coupled Finite Element-Smoothed Particle Hydrodynamics and the Arbitrary Lagrangian Eulerian methods. For the generation of the full aircraft model a multidisciplinary process chain approach and a standardized data format description are used. The beam element representations are considered for the modelling of skin reinforcement as well as other structures like cabin and cargo floor structures. By this approach, first time feasible full aircraft ditching simulations and the subsequent analysis of both global kinematics and the local fuselage structural response could be achieved.

Data-driven Bayesian inference of turbulence model closure coefficients incorporating epistemic uncertainty

We introduce a framework for statistical inference of the closure coefficients using machine learning methods. The objective of this framework is to quantify the epistemic uncertainty associated with the closure model by using experimental data via Bayesian statistics. The framework is tailored towards cases for which a limited amount of experimental data is available. It consists of two components. First, by treating all latent variables (non-observed variables) in the model as stochastic variables, all sources of uncertainty of the probabilistic closure model are quantified by a fully Bayesian approach. The probabilistic model is defined to consist of the closure coefficients as parameters and other parameters incorporating noise. Then, the uncertainty associated with the closure coefficients is extracted from the overall uncertainty by considering the noise being zero. The overall uncertainty is rigorously evaluated by using Markov-Chain Monte Carlo sampling assisted by surrogate models. We apply the framework to the Spalart–Allmars one-equation turbulence model. Two test cases are considered, including an industrially relevant full aircraft model at transonic flow conditions, the Airbus XRF1. Eventually, we demonstrate that epistemic uncertainties in the closure coefficients result into uncertainties in flow quantities of interest which are prominent around, and downstream, of the shock occurring over the XRF1 wing. This data-driven approach could help to enhance the predictive capabilities of computational fluid dynamics (CFD) in terms of reliable turbulence modeling at extremes of the flight envelope if measured data is available, which is important in the context of robust design and towards virtual aircraft certification. The plentiful amount of information about the uncertainties could also assist when it comes to estimating the influence of the measured data on the inferred model coefficients. Finally, the developed framework is flexible and can be applied to different test cases and to various turbulence models.

Noise prediction research of a scaled turboprop aircraft

Aerodynamic noise level has become an important performance index of civil aircraft, and it is drawing more and more attention. Most airframe noise research based on CFD method is aimed at aircraft individual parts at present, while lack of noise prediction for the complex high fidelity full aircraft model. Due to the interaction between aircraft parts, noise prediction between single part and the actual configuration are very different in the aspect of calculation conditions, noise generation and propagation. Only by using more realistic model can the noise be accurately predicted. In this paper, high-resolution mesh and high-precision hybrid RANS/LES method, combined with the FW-H acoustic analogy method, are employed to predict the noise of a turboprop's high-fidelity 1/6 scale model of landing configure. The improved delayed detached eddy simulation (IDDES) method was used to simulate the flow in the near-field sound source region, and the sound source characteristics of the whole turboprop are obtained. In addition to the two important noise sources-flap side edge and wing tip, numerical simulation also found other two important noise sources resulting from the interaction between parts-interaction between nacelle wake and flap, and the complex flow between flap inner side and fuselage. Results of far-field noise show that in the longitudinal symmetry plane of the fuselage, the noise mainly propagates to the lower front and over back direction, and the dipole characteristics are very obvious. In the plane perpendicular to the incoming flow, the noise in the transverse direction is weaker.

Barrier Lyapunov function and disturbance observer based omnidirectional robust gust response stabilization for multi-control-effectors aircraft

Purpose The purpose of this paper is to find an omnidirectional robust gust response stabilization (GRS) scheme with anti-disturbance and state-limited features. Design/methodology/approach Disturbance observer and barrier Lyapunov techniques, which can, respectively, estimate the lumped disturbances of the dynamic system in real-time and ensure the middle states within some prescribed ranges according to some flight safety indexes. Findings In the existing literature, almost all of the GRS controllers are either only for the longitudinal dynamics or only for the latitudinal dynamics. Few studies have considered the gust response alleviation problem with omnidirectional wind disturbance and full aircraft model. Originality/value This paper proposes a fresh scheme to deal with a more holistic GRS problem; the disturbance observer based (DOB) barrier Lyapunov backstepping longitudinal controller has been put forward; DOB nonlinear dynamic inversion to handle the multi-input-multi-output lateral dynamics; and to closely connect the two loops of the latitudinal dynamics, a manipulating variable conversion method is proposed.

Mathematical Model and Vibration Analysis of Aircraft with Active Landing Gear System using Linear Quadratic Regulator Technique

This paper deals with the study and comparison of passive and active landing gear system of the aircraft and dynamic responses due to runway irregularities while the aircraft is taxying. The dynamic load and vibration caused by the unevenness of runway will result in airframe fatigue, discomfort of passengers and the reduction of the pilot’s ability to control the aircraft. One of the objectives of this paper are to obtain a mathematical model for the passive and active landing gears for full aircraft model. The main porpuse of current paper is to design Linear Quadratic Regulator (LQR) for active landing gear system that chooses damping and stiffness performance of suspension system as control object. Sometimes conventional feedback controller may not perform well because of the variation in process dynamics due to nonlinear actuator in active control system, change in environmental conditions and variation in the character of the disturbances. To overcome the above problem, this research designs a controller for a second order system based on Linear Quadratic Regulator. The performance of active system is compared with the passive landing gear system by numerical simulation. The results of current paper in compared with the previous work mentioned in reference, demonstrate 37.04% improvement in body acceleration, 20% in fuselage displacement and 13.8% in the shock strut travel. The active landing gear system is able to increase the ride comfort and good track holding by reducing the fuselage acceleration and displacement and load induced to airframe caused by runway excitation.

THE WING-BODY AEROELASTIC ANALYSES USING THE INVERSE DESIGN METHOD

Flutter phenomenon is one of the most dangerous problems in aeroelasticity. When it occurs, the aircraft structure can fail in a few second. In recent aeroelastic research, computational fluid dynamics (CFD) techniques become important means to predict the aeroelastic unstable responses accurately. Among various flow equations like Navier-Stokes, Euler, full potential and so forth, the transonic small disturbance (TSD) theory is widely recognized as one of the most efficient theories. However, the small disturbance assumption limits the applicable range of the TSD theory to the thin wings. For a missile which usually has small aspect ratio wings, the influence of body aerodynamics on the wing surface may be significant. Thus, the flutter stability including the body effect should be verified. In this research an inverse design method is used to complement the aerodynamic deficiency derived from the fuselage. MGM (modified Garabedian-McFadden) inverse design method is used to optimize the aerodynamic field of a full aircraft model. Furthermore, the present TSD aeroelastic analyses do not require the grid regeneration process. The MGM inverse design method converges faster than other conventional aerodynamic theories. Consequently, the inverse designed aeroelastic analyses show that the flutter stability has been lowered by the body effect.