Aerospace Science and Technology

Published by Elsevier
Print ISSN: 1270-9638
In the Summer of 1995, the Confederation of European Aerospace Societies (CEAS), comprising the national Aerospace Societies of France (AAAF), Germany (DGLR), Italy (AIDAA), the Netherlands (NVvL), Spain (AIAE) and the United Kingdom (RAeS) formed the CEAS-Aeroacoustics Specialists' Committee (ASC).This Committee is to serve and support the scientific and industrial aeroacoustics community in Europe. Here “Acoustics” is to encompass all aerospace acoustics and related aeras.Each year the Committee will highlight some of the research and development activities in Europe. This is the report on the 1996 highlights.
The paper deals with the analysis of the effective stiffness of stiffened metallic panels under combined compression and shear stress as used, e.g. in aircraft fuselages. An important criterion for sizing and certification of aircraft fuselages is the local and global buckling behaviour. For optimisation of stiffened metallic panels it is necessary to know the buckling and postbuckling behaviour as accurately as possible. Due to the fact that nonlinear FE analyses of a whole aircraft fuselage are too time consuming, a very fast quasi nonlinear FE analysis with a coarse mesh under consideration of semi-empirical methods for the effective skin-stiffness are used. At this point the effective stiffness method derived in this paper is used. Admittedly previous semi-empirical methods like the effective width method [J. Wiedemann, Leichtbau 1: Elemente, second ed., 1996; M.C.Y. Niu, Airframe Stress Analysis and Sizing, second ed., Commilit. Press Ltd., 1999] (only for pure compression load) or the method of Kuhn [P. Kuhn, J.P. Peterson, L.R. Levin, A summary of diagonal tension part I – methods of analysis, Technical Note 2661, NACA, 1952] (only for pure shear load) have disadvantages for the general combined compression and shear load case. This is improved in the current method.
Planar laser-induced fluorescence (PLIF) has been used to measure time-resolved spatial distributions of the fuel, the OH radical, and the temperature field in a jet engine model combustor segment. For temperature measurements, a two-line PLIF scheme was used: two different spectral lines of the OH radical, which served as indicator molecule, were excited successively within a short time delay using nanosecond pulses from two UV laser systems operating on different wavelengths. The ratio of the two fluorescence signals depends on the temperature; this forms the basis of the temperature measurement. To our knowledge, this temperature mapping technique has been applied for the first time in the high pressure combustion of kerosene. The fluorescence signal resulting from excitation by one of the two lasers is proportional to the OH density and provides thus information on the OH radical distribution and on flame structure. The same technique can be utilized to excite fluorescence from the fuel, thus providing qualitative information on kerosene distributions. These measurements yield information on flame structure, heat release and mixing properties, which can serve as design aids, as well as for CFD code validation purposes. (C) 2000 Editions scientifiques et medicales Elsevier SAS.
A two-dimensional numerical study was performed to investigate the acoustic effects of micro-tab device on airframe noise reduction. As the noise generated by leading-edge slat and trailing-edge flap rise with their increased deflection angles, it is possible to mitigate such high-lift noise by using reduced settings without sacrificing the aerodynamic performance during approach. In this paper, micro-tab device attached to the pressure side of the flap surface is envisioned as a mean to achieve this goal. Hybrid method involving Computational Fluid Dynamics and acoustic analogy was used to predict the far-field noise spectrum. Results illustrate that the micro-tab device with reduced deflection angles of the high-lift settings provides lower noise signature at far-field positions, comparing to the baseline configuration, while the aerodynamic performance is maintained. In addition, two parametric studies which investigated the effects of micro-tab location and micro-tab height on acoustic spectra were also included.
The present paper aims to describe the potential of the adjoint technique for aerodynamic shape optimization. After a brief description of the continuous adjoint formulation and the aerodynamic optimization process developed at the DLR, specific requirements for an optimization framework combined with the adjoint technique are introduced. The drag reduction at constant lift and pitching moment for the RAE2822 airfoil in transonic flow is then presented as validation case. An extension to multi-point optimization demonstrates the capability of the framework to solve more complex problems. Finally, the wing-body optimization of a supersonic commercial aircraft confirms the flexibility of the framework and the efficiency of the adjoint technique.
A domain decomposition method which combines a vortex method and a finite difference method is presented to solve the Navier-Stokes equations in velocity-vorticity form. The lagrangian vortex method is used in the flow region where convective effects are dominant whereas the finite difference method using velocity-vorticity in this paper is used in the flow region where the convective effects may be dominant. Unsteady flow around an isolated NACA 0012 is studied. The comparison between numerical and experimental data show that the method is consistent for simulating two dimensional flows. Flow interaction between an airfoil and a circular cylinder has been analyzed showing the abilities of the method in multibodies flow simulation. The velocity-vorticity formulation can be extended easily in three dimensional flow simulation.
A Reynolds-averaged Navier–Stokes solver, a laminar boundary-layer code and different transition prediction methods for the prediction of Tollmien–Schlichting and cross flow instabilities were coupled in order to perform Reynolds-averaged Navier–Stokes computations of three-dimensional, finite wings with automatic laminar-turbulent transition prediction. The results from computations based on two database methods and a local, linear stability code together with the eN-method are compared for a three-dimensional wing configuration.
Over the last years, the discontinuous Galerkin method (DGM) has demonstrated its excellence in accurate, high-order numerical simulations for a wide range of applications in computational physics. However, the development of practical, computationally efficient flow solvers for industrial applications is still in the focus of active research. This paper deals with solving the Navier-Stokes equations describing the motion of three-dimensional, viscous compressible fluids. We present details of the PADGE code under development at the German Aerospace Center (DLR) that is aimed at large-scale applications in aerospace engineering. The discussion covers several advanced aspects like the solution of the Reynolds-averaged Navier-Stokes and k-ω turbulence model equations, a curved boundary representation, anisotropic mesh adaptation for reducing output error and techniques for solving the nonlinear algebraic equations. The performance of the solver is assessed for a set of test cases.
The numerical simulation of the flow around a 65° delta wing configuration with rounded leading edges is presented. For the numerical simulation the DLR TAU-Code is used which is based on an unstructured hybrid mesh approach. Within this paper several numerical results are shown, solving the steady RANS equations by different turbulence models. The simulations are carried out within the RTO/AVT 113 working group focusing on experimental and numerical research on delta wing configurations with rounded leading edges. Within this paper the focus is related to the flow topology depending on the angle of attack as well as on Reynolds number effects. Finally the results will be compared and verified by experimental data.
A numerical investigation of the transonic steady-state aerodynamics and of the two-degree-of-freedom bending/torsion flutter characteristics of the NLR 7301 section is carried out using a time-domain method. An unsteady, two-dimensional, compressible, thin-layer Navier-Stokes flow-solver is coupled with a two-degree-of-freedom structural model. Fully turbulent flows are computed with algebraic or one-equation turbulence models. Furthermore, natural transition is modeled with a transition model. Computations of the steady transonic aerodynamic characteristics show good agreement with Schewe's experiment after a simplified accounting for wind-tunnel interference effects is used. The aeroelastic computations predict limit-cycle flutter in agreement with the experiment. The computed flutter frequency agrees closely with the experiment but the computed flutter amplitudes are an order of magnitude larger than the measured ones. This discrepancy is likely due to the omission of the full wind-tunnel interference effects in the computations. (C) 2001 Editions scientifiques et medicales Elsevier SAS.
Typical turboprop noise spectra exhibit a series of characteristic peaks which are directly related to the product of propeller rpm and number of propeller blades. These blade passage frequencies contribute significantly to the overall sound pressure level both outside and inside the aircraft. Their contribution to cabin noise is usually reduced by appropriately adjusted mass dampers. However, since the engine rpm varies for different flight stages, any fixed eigenfrequency absorber will merely be a sub-optimal compromise. The Tunable Vibration Absorber (TVA) introduced in this article has a variable resonant frequency which enables an adaptation to different flight phases providing largely improved performance. Frequency tuning is achieved through a piezo-electric stack actuator, which applies a pressure force to a pair of leaf springs thus reducing their effective bending stiffness. Among the main advantages of this particular approach are a static control signal and low power consumption. To enable a light-weight construction the components which generate the pressure loading are incorporated into the oscillating mass. The TVA allows to cover a wide frequency range using only a single device. Additionally, it features damping control capability and optional active multi-mode operation. Structural-acoustic simulations have indicated a noise reduction potential of approximately 10 dB. This article gives a short overview of different tuneable vibration absorber concepts, lines out the theoretical background of the proposed approach, discusses the general components layout and describes the experimental verification of a prototype TVA for the Airbus A400M.
This paper addresses the design and application of active interfaces and semi-active vibration absorbers that can be used to reduce annoying vibration or noise levels. Currently, these devices are mainly used in automotive applications but they can easily be adapted to meet the requirements of aerospace applications. The design process and the numerical simulations that are necessary to assess the performance of such active interfaces are described. This includes the modeling of the mechanical structure, of the components of the active interface, and of the controller. The simulation results are compared with experimental measurement results, which agree very well. An application example is presented as well. Finally, the design, application, and effect of a semi-active vibration absorber are described.
A tunable diode laser spectrometer has been used to probe the ONERA-F4 arc-driven high-enthalpy wind tunnel. By fast wave number tuning of the laser beam passing through the flow, free-stream measurements from NO or H2O absorption lines can be made at high repetition rates (1 kHz) in the infrared domain near 5 μm. Temperature and density measurements are obtained from the absorption line shapes, and the flow velocity is deduced from the Doppler shift at the line centre position. This line-of-sight technique gives averaged values over the absorption path through the flow. Herein we report results concerning medium and high enthalpy runs, where velocities inferred from the Doppler shifts range between 2500 and 4500 m/s. The results are also compared to calculations from heat flux probes measurements and flow models.RésuméLa spectroscopie d'absorption par diode laser infrarouge a été utilisée sur les espèces moléculaires NO et H2O pour calibrer certains paramètres de l'écoulement libre de F4 (soufflerie hypersonique à arc bref ONERA/Le Fauga-Mauzac). La mesure de la vitesse de l'écoulement est déduite du décalage Doppler des positions des raies d'absorption. La mesure de la température cinétique de l'écoulement ainsi que celle des concentrations de NO et de H2O ont aussi été effectuées à partir du profil et de l'intensité de ces raies d'absorption. Ces mesures ont été réalisées avec un taux de répétition de 1 kHz. De par sa nature de technique d'absorption, la spectrométrie diode laser donne des valeurs moyennées sur le segment de trajet du faisceau traversant l'écoulement. Nous présentons ici des résultats des mesures effectuées pour des conditions de fonctionnement de la soufflerie à moyenne et à haute enthalpie où la vitesse de l'écoulement déduite du décalage Doppler est entre 2500 et 4500 m/s. Ces résultats sont aussi comparés à ceux obtenus par des sondes de flux thermiques ainsi qu'aux modélisations de l'écoulement.
There are huge potential applications of 3-D braided composite in aerospace engineering because of the non-delamination feature of the composite under impact loading. This paper presents the analysis of energy absorption features of 3-D braided composite under compression with different strain rates. The 3-D 4-step rectangular braided composite coupons were tested on a material tester MTS 810.23 and a split Hopkinson pressure bar (SHPB) apparatus to obtain out-of-plane and in-plane compression stress vs. strain curves at quasi-static and high strain rate state. The failure modes and energy absorption features of the 3-D braided composite under different strain rates were analyzed both in time domain and frequency domain. The energy absorbed by the 3-D braided composite increases with the strain rate. From fast Fourier transform (FFT) analysis of compression stress vs. time histories, the power of energy absorption of the 3-D braided composite increases with strain rate and mostly concentrate on the high frequency region. While for quasi-static compression, the power distributes in very narrow frequency region and also is less than that in high strain rates. This feature corresponds to the different damage and energy absorption mechanisms of the 3-D braided composite under quasi-static and high strain rate compression.
The flexible high-aspect-ratio wings of high-altitude long-endurance unmanned aerial vehicles experience large geometrical deformations. The nonlinear aeroelastic analysis for such a wing is carried out by using a loosely-coupled CFD/CSD method, which treats the fluid and structure as two separate modules and updates the CFD and CSD variables separately with a transfer of variables at the fluid–structure interface. In the loosely-coupled method employed here, an unsteady Euler solver and a nonlinear CSD solver are joined together by a fully three-dimensional integral constant volume tetrahedron (CVT) interfacing technique. The computation process is very time-consuming when the computed incremental displacements of every aerodynamic node on the wing surface are fully added to the previously computed deformed wing configuration. For a maximum deflection of approximately 8.7 mm (0.54% of semi-span length), the coupled computation scheme takes 20 coupling iteration steps with about 72 hours to converge on an HP XW6400 workstation with a 3 GHz Xeon5160 CPU. To reduce the computational expense of this loosely-coupled method, the golden section technique with an empirical parameter is introduced to speed up convergence. The deflections relaxed by using this technique are assimilated, and the wing bends up monotonically to its static equilibrium position with a maximum deflection 44.9 mm. The convergence history shows that, this accelerated algorithm takes just 6 coupling iteration steps with about 24 hours to monotonically converge to its static equilibrium position, although the maximum deflection 44.9 mm is 5 times larger than the maximum deflection 8.7 mm of the above test case with aeroelastic deflections fully assimilated. So, it is employed in the following nonlinear fluid–structure interaction (FSI) computations. After this nonlinear aeroelastic system has reached its static equilibrium position, the aerodynamic loads on wing surface are extracted and then applied onto the linear wing structure to calculate its deformation. In present paper, if the geometric nonlinear effects are taken into account for wing deflection calculation, the wing structure model is named as “nonlinear wing structure”; otherwise, the wing structure model is named as “linear wing structure”. The role of geometric nonlinearity on aeroelastic deformation is analyzed by comparing the deformations of linear and nonlinear wing structures. It is shown that, geometric nonlinearity plays an important role for large static aeroelastic deformation and should be accounted for in aeroelastic analyses for such high-aspect-ratio flexible wings.
Some flexible appendages of spacecrafts, such as solar panels, are cantilever plate structures. Thus, vibration problem is unavoidable when there is slewing maneuver or external disturbance excitation. Vibration of such cantilever plate structures includes coupled bending and torsional motion. Furthermore, the low amplitude vibration near the equilibrium point is very difficult to be quickly suppressed due to nonlinear factors of the hardware in the system, which is harmful to stability and attitude control accuracy. To solve these problems, acceleration sensor-based modal identification and active vibration control methods are presented for the first two bending and the first two torsional modes vibration of the cantilever plate. Optimal placements of three acceleration sensors and PZT patches actuators are performed to decouple the bending and torsional vibration of such cantilever plate for sensing and actuating, and identifications are achieved by experiments. A nonlinear control method is presented to suppress both high and low amplitude vibrations of flexible smart cantilever plate significantly. Experimental comparison researches are conducted by using acceleration proportional feedback and the presented nonlinear control algorithms. The experimental results demonstrate that the presented acceleration sensor-based methods can suppress the vibration of cantilever plate effectively.
In this paper we describe the implementation and validation of arbitrarily moving reference frames in the block-structured CFD-code EURANUS. We also present results from calculations on two applications involving accelerating missiles with generic configurations. It is shown that acceleration affects wave drag significantly. Also, it is shown that strake-generated vortices move significantly in turns. These results clearly show the necessity of including the acceleration effects in the calculations.
Ongoing and future geopotential space missions are equipped with one or more accelerometers. In order to use these observations, the accelerometer measurements have to be calibrated before processing them. In this paper they are introduced in the GPS based precise orbit determination, by replacing the non-gravitational force models for atmospheric drag and solar radiation pressure. Empirical accelerations are still estimated to account for deficiencies in the applied conservative force models. The in-orbit calibrated accelerometer observations are used to validate the accelerations determined by force modeling. In along-track direction they show the best agreement. During days of high solar activity the benefit of using accelerometer observations is clearly visible. The observations during these days show high frequency fluctuations which the modeled and empirical accelerations cannot follow. A long period of GRACE (second half of 2003) and of CHAMP (2004) data is processed. This results for GRACE in a mean orbit fit of a few centimeters with respect to high-quality JPL reference orbits, showing a slightly better consistency compared to the case when using force models, which is also supported by SLR residual analysis. The daily calibration factors determined with this technique show a small variation. When not including empirical accelerations in the estimation procedure, the calibration factors in radial and cross-track direction show a bigger spread, with an orbit fit below the decimeter level.
This paper deals with a micromechanical damage model, based on a constitutive theory for brittle materials weakened by microcracks. The model is implemented in the DYNA3D three-dimensional explicit finite element code. The phenomenological study shows the importance of taking micromechanical effects into account to model macroscopic failure of the material. The constitutive model relates damage to microscopic parameters (size of microcracks, cracks density etc.) and takes loading-induced anisotropy damage into account by correlating microcrack growth to preferential orientations. The unilateral character (behaviour difference between tension and compression) is treated by the microcrack growth criterion. The progressive reduction in material stiffness due to the presence of microcracks is modelled using Margolin's effective modulus expressions, and the material is pulverised if the microcrack density exceeds a critical value. Determination of the energy dissipated by damage is proposed. The constitutive model applied to SiC/SiAlYON ceramics is validated by a comparaison of the results between a Hopkinson's Bar Test and numerical simulation. Comparing the macroscopic brittle model results with the damage model results shows the ability of the second to predict microcrack effects on the dynamic failure behaviour of ceramics.
The Automated Transfer Vehicle (ATV) is a European spacecraft intended to service the International Space Station (ISS). It is designed to perform automated phasing, approach, rendezvous and docking to the ISS, then departure and deorbitation manoeuvres. Such an automated rendezvous mission towards a manned facility raises severe performance and safety constraints for the vehicle, which are declined towards the on-board Navigation in terms of availability, accuracy and failure tolerance: the ATV shall be operational after any first failure and safe conditions shall be reached after a second failure. The whole ATV navigation system has been designed to fulfil these very stringent requirements. Based on fully redundant hardware, the navigation algorithms present optimal estimators and multi-layers Failure, Detection, Isolation and Recovery (FDIR) capabilities to ensure the continuity of the state vector in case of failure. Several functions provide state vectors estimations and health reports, according to the flight phase: the attitude and drift estimation function provides the vehicle absolute attitude and angular rate during the whole flight; for the far rendezvous, position and velocity relative to the ISS are estimated by the relative GPS navigation and a dedicated relative navigation with Videometer applies in close rendezvous, in the final approach. All these autonomous navigation functions offer nevertheless a high level of monitoring and control to the ATV Control Centre operators. Thus the ATV innovative navigation chain provides the high level of performance, robustness and autonomy required by modem spacecrafts involved in human programs, today in Earth orbit but also for future space exploration missions.
The multi-site damage problem of aging aircraft has raised the issue of obtaining accurate stress intensity factors of interacting collinear cracks under external load. The simple approximate technique proposed by Kachanov (1987) [8] is specialized in this work to arbitrary length collinear cracks with arbitrary spacing under far field tension. This approach yields simple formulas that can be used to quickly and accurately assess the mode I stress intensity factors of any number of collinear cracks in an infinite body under far field tension considering interactions between all cracks. The advantage of this approach over others is that it is simple, has good accuracy, and is very efficient. Kachanov's method is further enhanced here such that a repeating subpanel of cracks of arbitrary number, size, and spacing can be modeled implicitly with resulting efficiency gains. The effect of interaction cracks can be used to build up stress intensity factor solutions for more complex geometries. Several numerical examples and timing studies are used to demonstrate the efficacy of this approach.
Forcing mechanisms in aerodynamic flow actuation in Mach 3 supersonic flow provided by a direct-current surface discharge are investigated experimentally and computationally. High-speed flow actuation is achieved by creating a near-wall ionization region produced by striking a discharge between two bare round electrodes which are flush mounted on a ceramic actuator plate. Flow actuation which is signatured by the occurrence of a weak oblique shock originating from the ionization region is observed with a lower power (∼10's W) diffuse discharge which exhibits a volumetric ionization region above the cathode. This actuation effect is achieved when the cathode is placed upstream of the anode while no actuation is detected for the opposite case albeit with the same power input. Gas heating effect on flow actuation is evaluated by estimating gas (rotational) temperature using optical emission spectroscopy and numerical simulation. Significant gas heating as high as 400 K (free stream temperature is about 110 K) is observed near the cathode and numerical simulation confirms that gas heating is effective in supersonic flow actuation. However, the fact that the peak magnitude and spatial profile of gas temperature is similar for both cathode-upstream and cathode-downstream cases implies similar gas heating effect on supersonic flow actuation even though the actuation effect is different. Electrohydrodynamic (EHD) effect on flow actuation is emphasized by experimental results whose effect has been ignored in supersonic flow actuation and is supported by phenomenological modeling of electrostatic force. Analytic estimate of electrostatic forcing and corresponding computational result propose the potential role of EHD effect on supersonic flow actuation explaining the absence of flow actuation in the cathode downstream case.
A numerical approach based on the solution of convected wave equations in the frequency domain is applied to the prediction of fan noise radiation from realistic engine configurations at realistic operating conditions. Fully three-dimensional computations based on the acoustic velocity potential at Helmholtz numbers up to 30 are carried-out in order to simulate the sound transmission through a scarfed inlet in the presence of a spliced liner. A verification of the method capability in featuring the modal scattering in the presence of rigid splices is carried out by comparing present results with analytical and numerical results available in the literature. In order to avoid the treatment of the vortex sheet shed from the edge of the bypass duct, a wave model for the acoustic pressure based on the Lilley's third-order wave operator is used for the aft noise radiation from the exhaust. Again, a fully three-dimensional computation is carried-out for a Helmholtz number equal to 30 and results are compared to analytical solutions available in the literature for an idealized flow configuration. Finally, the effects of a more realistic mean flow are investigated by evaluating the refractive effects due to a mixing-layer velocity profile.
This paper presents the results of an experimental program that was conducted to examine the effect of aspect-ratio on the flow development and the related acoustic properties of jets issuing from elliptic-slots. The aspect-ratios investigated are 2:1, 3:1 and 4:1. Fully expanded cases show a decrease in shear layer growth with increasing aspect-ratio resulting in a downstream shift in axis-switching location. Further the small aspect-ratio jet shows a dominance of large-scale structures close to slot exit relative to 3:1 and 4:1 jets. In the underexpanded condition the shock-structure development shows appreciable changes with aspect-ratio. At the end of first cell, oblique shock cross-over point in 2:1 jet is gradually replaced by a normal shock in 3:1 and 4:1 jets. The frequency content of jet noise also shows a corresponding increase in amplitude of fundamental screech frequency with increase in aspect-ratio. Introduction of notches along the minor-axis sides further modifies the shock development process which significantly alters the screech amplitudes for notched 3:1 and 4:1 jets indicating the weakening of the feedback process.
A porous solid, saturated with fluid, may be described as an elastic-viscoelastic and acoustic-viscoacoustic medium. The transport of vibroacoustic energy is carried both through the sound pressure waves propagating through the fluid in the pores, and through the elastic stress waves, carried through the solid frame of the material. For most porous materials, used to reduce sound and vibration, these waves are coupled to each other, i.e. they simultaneously propagate in both the fluid and the solid frame but with different strengths. A characteristic of this coupled wave propagation, is that the vibroacoustic energy is dissipated and converted into heat as the wave travels through the material. Clearly for a given situation, the balance between energy dissipated through vibration of the solid frame and changes in the acoustic pressure varies with the topological arrangement, choice of material properties, interfacial conditions, etc. This paper illustrates the influence such a balancing has on the performance of a multi-layer sound proofing arrangement applicable for an aircraft interior.
The acoustic post-processing of unsteady aerodynamic jet simulations using Kirchhoff or Ffowes Williams and Hawkings (FW-H) surface integral methods is investigated from the theoretical and practical points of view. This analysis is carried out for a supersonic hot jet, starting from the flow fields provided by an unsteady aerodynamic simulation whose characteristics are recalled. Both acoustic integral methods are first compared from a theoretical point of view. In particular, the role of the 'entropic' sources on the control surface is pointed out. Various acoustic calculations are then carried out for this hot jet to assess the influence of the type of method and of the position, extent and nature (open or closed) of the control surfaces on the noise predictions. This parametric study shows that the Kirchhoff method using density as input data is poorly suited for acoustic predictions of hot jets. The Kirchhoff method using pressure as input data provides results similar to those obtained with the FW-H surface integral but the latter appears to be more reliable. The study also shows that the use of a downstream closed control surface does not present a real interest, whatever the acoustic method is.
A block-structured adaptive mesh refinement (AMR) method was applied to the computational problem of acoustic radiation from an aeroengine intake. The aim is to improve the computational and storage efficiency in aeroengine noise prediction through reduction of computational cells. A parallel implementation of the adaptive mesh refinement algorithm was achieved using message passing interface. It combined a range of 2nd- and 4th-order spatial stencils, a 4th-order low-dissipation and low-dispersion Runge–Kutta scheme for time integration and several different interpolation methods. Both the parallel AMR algorithms and numerical issues were introduced briefly in this work. To solve the problem of acoustic radiation from an aeroengine intake, the code was extended to support body-fitted grid structures. The problem of acoustic radiation was solved with linearised Euler equations. The AMR results were compared with the previous results computed on a uniformly fine mesh to demonstrate the accuracy and the efficiency of the current AMR strategy. As the computational load of the whole adaptively refined mesh has to be balanced between nodes on-line, the parallel performance of the existing code deteriorates along with the increase of processors due to the expensive inter-nodes memory communication costs. The potential solution was suggested in the end.
This paper describes a new experimental approach to acoustic liner characterization in the presence of a grazing flow. The traditional methods of measurement use microphones to determine liner impedance. The in situ method in particular requires the simultaneous use of two microphones. The first is mounted flush with the surface of the liner grazed by the flow and the second is flush-mounted to the rear face of the liner. However, this method is invasive and assumes the reaction of the liner to be independent of the incidence of the waves (locally-reacting liner). The approach suggested here is radically different since Laser Doppler Velocimetry (LDV) is used to measure the acoustic perturbation of velocity, or acoustic velocity. This latter allows us to determine the acoustic displacement, which is the key parameter in Galbrun's linear theory for assessing the perturbation of pressure and the field of active intensity. The wall impedance and the propagation paths of acoustic energy in the presence of the liner may be deduced without any assumption and non-invasively. This approach was applied for characterizing a resistive liner in a test bench specially designed for aeroacoustic measurements, with a 2D LDV system. The flow was turbulent and the measured nominal Mach number was 0.13. The impedance and field of active intensity were then obtained. A comparison was carried out between the new approach and the in situ method using microphones. According to previous theoretical works in the literature and the presented test results, one has to be cautious about the definition of the impedance when performing in-flow acoustic measurements.
Nonlinear acoustic damping induced by a half-wave resonator is investigated numerically by adopting nonlinear analysis. Governing equations describing nonlinear acoustic fields in a chamber are solved simultaneously. The two acoustic properties of damping factor and insertion loss are adopted to quantify acoustic damping capacity of the resonator. As the amplitude of pressure disturbances increases, the baseline damping capacity of the chamber without the resonator increases gradually and then, rapidly. The optimal length found in a linear range is still valid in a nonlinear range. But, pure acoustic-damping effect induced by fine tuning of the resonator is degraded rather by nonlinear acoustics. The effect can be clearly quantified by the insertion loss, not the damping factor. From the acoustic fields in an acoustic tube with a single resonator, the insertion losses are calculated with two adjustable parameters of the resonator length and sound pressure level. It is validated even by the insertion-loss approach that the resonator functions as a half-wave resonator in a linear range. From nonlinear responses of the resonator, it is found that the damping capacity of the resonator is degraded and becomes nearly identical irrespective of the resonator length when high-amplitude acoustic oscillation is excited.
Helicopter flow computation is still a challenging task these days. The flow is instationary and transonic in a complex, time varying geometry and thus very difficult to simulate accurately. Besides that, reliable results can be achieved only if the dynamics of the blade are taken into account as well. Even in hover, but more so in forward flight the aerodynamic problem is tightly coupled to the structural response of the blades to aerodynamic forces. Rigid body motions occur as well as elastic deformations like bending and especially twist. In addition to the aerodynamic performance of a helicopter, acoustical effects become an important issue more and more. Consequently aeroacoustic data has to be generated as well as aerodynamic and aeroelastic. Realistic helicopter rotor simulations are therefore a multi disciplinary problem, where computational fluid dynamics admittedly plays the key role.
The helicopter project CHANCE contains, among other developments, the quasi-steady approximation to modelling rotors with actuator discs. This reduces the cost of an unsteady simulation down to a stationary one. In testing existing approaches in the literature, the source term implementation proved to perform best, especially in forward flight: source terms located on the disc bottom side impart impulse and energy to the fluid. These are obtained from a loose coupling between two DLR codes: the flow solver FLOWer and the rotor code S4. The latter provides a rotor map, a radial and azimuthal force distribution, to the former converting it to the actuator disc map (source terms). Low velocities are accounted for using preconditioning. A more flexible Chimera approach can be used. The actuator disc feature has been developed in a parallel framework for shorter turn-around times.
This study focusses on the numerical simulation of supersonic flows around a missile with lattice wings. In order to reduce the computational cost for such a configuration, the actuator disc concept is chosen to model the effects of grid fins. This approach is coupled with a Navier-Stokes solver in order to predict the forces and moments on a complete vehicle. The method consists in replacing the lattice controls by artificial boundary conditions where the forces involved by the grid fins are applied. These forces are interpolated from an experimental database. Numerical simulations are performed for laminar and turbulent flows for several Mach numbers and angles of attack. The results are compared to experimental data in order to validate the method. The comparisons reveal some discrepancies which are mainly due to the turbulence effects and to the database from which the forces resulting from the lattice wings are interpolated. The method allows the prediction of the forces applied on the vehicle and therefore an estimate of its aerodynamic performances. The computations validate the approach and show its potential as a tool for vehicle design.
In this paper a novel approach is developed for optimization of piezoelectric actuators in vibration suppression. A scaled model of a vertical tail of F/A-18 is developed in which piezoelectric actuators are bounded to the surface. The frequency response function (FRF) of the system is then recorded and maximization of the FRF peaks is considered as the objective function of the optimization algorithm to enhance the actuator authority on the mode, which assigns the optimal placement of the pair of piezoelectric actuators on the smart fin. Six multi-layer perceptron neural networks are employed to perform surface fitting to the discrete data generated by the finite element method (FEM). Invasive weed optimization (IWO), a novel numerical stochastic optimization algorithm, is then employed to maximize the FRF peak which in due reduces the vibration of the smart fin. Results indicated an accurate surface fitting for the FRF peak data as well as the optimal placement of the piezoelectric actuators for vibration suppression.
Active flutter velocity enhancement scheme is presented for lifting surfaces, employing Linear Quadratic Gaussian based multi-input multi-output controller with multilayered piezoelectric actuators. To numerically test the developed concept, a composite plate wing, surface bonded with eight piezoelectric bender actuators and sensors has been considered. A modal flutter control model is formulated in state-space domain using coupled piezoelectric finite element procedures along with unsteady aerodynamics and optimal control theory. The bending – torsion flutter instability has been actively postponed from 44.13 to 55.5 m/s using the energy imparted by the multilayered piezoelectric actuators. As the power requirement by these actuators is comparatively very low with respect to stack actuators, they can be employed in an integrated form to develop active lifting surfaces for real time applications.
This paper presents inviscid computations of an isolated rotor in hover using a specific mesh adaptation technique. The proposed adaptation method is simple, efficient and dedicated to this type of application. The user provides a short and fine grid around each blade; an overset cylindrical grid that discretizes the remaining part of the computational domain is then automatically generated and periodically adapted to improve the capture of the wake. This technique can also be applied to a portion of a multi-bladed rotor by using periodic boundary conditions. Transfers between blade and cylindrical grid are completed through the Chimera method. Results show that, with about 600000 points, the tip vortex can be followed for approximately 380 degrees of age. Also, this paper proposes a deep insight into different refinement indicators for this problem.
Launcher trajectory optimization is a complex task, especially when considering the specific problems arising in the study of reusable launch vehicles. Part of the difficulty comes from the different characteristics of the trajectory arcs which make up the vehicle's mission (constraints and controls may not be the same). Another difficulty is the necessity, in some cases, of a global optimization between ascent and re-entry phases (branching optimisation). Finally, optimization tools devoted to this task should be polyvalent and robust, as the studies of reusable launch vehicles usually cover many different concepts, and also many different trajectory cases (such as abort scenarios). The purpose of this paper is to present different approaches used in France by CNES and ONERA to solve optimal control problems in the context of launcher trajectory optimization. These approaches, which are powerful implementations of classical optimization methods, were designed to cover the needs for both expendable and reusable launchers trajectory calculation. The first optimization tool presented is OPTAX, which uses an indirect shooting method. The second and third tools presented are CNES's ORAGE and ONERA's FLOP/OLGA, which use two different variants of the gradient method. The paper describes the equations and methodology behind these tools, and also presents their advantages and drawbacks.
This paper concerns the flutter, post-flutter and adaptive control of a non-linear 2-D wing-flap system operating in supersonic/hypersonic flight speed regimes. An output feedback control law is implemented and its performance toward suppressing flutter and limit cycle oscillations (LCOs) as well as reducing the vibrational level in the subcritical flight speed range is demonstrated. This control law is applicable to minimum phase systems and we provide conditions for stability of the zero dynamics. The control objective is to design a control strategy to stabilize the pitch angle while adaptively compensating for uncertainties in all the aeroelastic model parameters. It is shown that all the states of the closed-loop system are asymptotically stable.
Un radar est un capteur spatio-temporel dont les performances dépendent beaucoup des traitements appliqués aux signaux qu'il délivre. Le but de cet article est, après avoir rappelé le principe de ces traitements, de livrer une réflexion sur les différentes possibilités offertes par le filtrage adaptatif en matière de réjection des signaux qui perturbent la détection des cibles. Ce type de filtrage, généralement appliqué dans le domaine angulaire, peut en effet être appliqué dans d'autres domaines avec des propriétés particulières que nous nous proposons de mettre en évidence. Ces propriétés seront illustrées par des résultats de simulation.
A backstepping control design procedure for uncertain nonlinear flight control system expressible in parameter-strict feedback form is presented in this paper. The proposed backstepping procedure, in association with sliding model control technique, exploits the possibility of avoiding, under certain suitable assumptions, the overparameterization problem existing in the classical backstepping process. In particular, a sliding-model-based integral filter is introduced to facilitate the development of the derivation of the virtual inputs, thus reducing the computational load with regard to the standard backstepping procedure. Moreover, in simulations, the control parameters in the resulted controller are optimally tuned using a genetic algorithm so as to show the full potential of the proposed control system.
This paper treats the question of control of nonlinear aeroelastic responses of a prototypical wing section with structural nonlinearity using leading- and trailing-edge control surfaces. It is assumed that all the aerodynamic, structural and inertia parameters are unknown to the designer. As such the limitation of a recent control design reported in the literature, which requires complete knowledge of aerodynamic derivatives and inertia parameters, is removed. An adaptive controller and a neural control system are designed for the trajectory control of the plunge displacement and pitch angle. For the derivation of the adaptive control law, a linearly parameterized model is used but the neural controller is designed by treating the stiffening-type structural nonlinearity as an unstructured function (not parameterizable). It is shown that the adaptive and neural controllers accomplish trajectory control in the closed-loop system. Simulation results are presented which show that these controllers are effective in regulating the nonlinear responses to the origin in the state space in spite of large model uncertainties. Moreover unlike the model with a single trailing-edge surface, two control surfaces provide flexibility in shaping both the plunge and pitch responses.
Applying adaptronics to helicopters has a high potential to significantly suppress noise, reduce vibration and increase the overall aerodynamic efficiency. This paper presents recent investigations on a very promising specific concept described as Adaptive Blade Twist (ABT). This concept allows to directly control the twist of the helicopter blades by smart adaptive elements and through this to positively influence the main rotor area which is the primary source for helicopter noise and vibration. Since the interaction of non-stationary helicopter aerodynamics and elastomechanical structural characteristics of the helicopter blades causes flight envelope limitations, vibration and noise, a good comprehension of the aerodynamics is essential for the development of structural solutions to effectively influence the local airflow conditions and finally develop the structural concept. With respect to these considerations, the ABT concept will be presented. This concept bases on the actively controlled tension-torsion-coupling of the structure. For this, an actuator is integrated within a helicopter blade that is made of anisotropic material based on fiber composites. Driving the actuator results in a local twist of the blade tip, in such a way that the blade can be considered as a torsional actuator. Influencing the blade twist distribution finally results in a higher aerodynamic efficiency. The paper starts at giving a review on conventional concepts and potential adaptive solutions for shape control (2),(3),(9),(13)-(15). Hereafter, some calculations of the adaptive twist control concept are presented. These are based on a representive model in which the active part of the rotor blade is simplified with a thin-walled rectangular beam, that is structurally equivalent to a model rotor blade of the Bo 105 with a scaling factor 2,54. The calculations are performed using an expanded Wlassow Theory. The results are valid for static and dynamic conditions. For the dynamic condition excessive deformations near the blade resonance frequency shall be utilised. Therefore, the actuated blade section has to be properly designed for this preconditions. This has been demonstrated and verified in experiments (7) which will not be discussed in this paper. For experimental investigations on the ABT concept the skin of the outer part of the model rotor blade was manufactured of fibre composite material using the above mentioned tension-torsion-coupling effect with an additional uncoupling layer between skin and spar. The experimental results have shown that near to the resonance frequency dynamic forces of 550+_550 N are required for a deformation of 3 degrees at the blade tip.
This paper describes an approach for augmenting a linear controller with a neural network based adaptive element in output feedback setting. The approach is applicable to non-affine nonlinear systems with parametric uncertainty and unmodeled dynamics. Stability of the adaptive control system is ensured using Lyapunov direct method. A numerical example of guided munitions illustrates the efficacy of the approach.
The military typically operates in demanding, dynamic, semi-structured and large-scale environments. This reality makes it difficult to detect, track, recognize/classify, and response to all entities within the volume of interest, thus increasing the risk of late (or non-) response to the ones that pose actual threat. A key challenge facing the military operators, in these contexts, is the focus of attention and effort, that is, how to make the most effective use of the available but scarce sensing and processing resources to gather the most relevant information from the environment and fuse it in the most efficient way. Adaptive Data Fusion and Sensor Management can aid this information gathering and fusion processes by automatically allocating, controlling, and coordinating the sensing and the processing resources to meet mission requirements. This paper presents results of a project initiated by Defence R&D Canada – Valcartier that aims at defining, developing, and demonstrating adaptive data fusion and sensor management concepts for distributed military surveillance operations.
A demonstrator representing a lightweight engine mounting for turboprop aircraft has been developed which allows for active suppression of disturbing vibrations. It consists of an advanced CFRP truss structure based upon high-load CFRP struts which enable an optimal integration of piezo-electric stack actuators. The dynamic behaviour of the truss structure was analyzed by a 3D-simulation and the optimal position of active struts was evaluated. The experimental results of static and dynamic tests showed very good agreement with the analytical predictions. First investigations of the active vibration suppression by means of a simple feedback controller exhibited an excellent functionality of the adaptive structure and a promising potential to future improvements.
Especially in the case of large transonic transport aircraft, flight conditions change considerably during a typical mission. This is accounted for by multiple but fixed design points which compromise, however, the aircraft performance. Employing adaptive wing technology where the wing geometry, or other means of flow control, adjust the flow development to the changing freestream and load conditions allows to fully explore the flow potential at each point of the flight envelope. Various means of flow control by geometric adaptation and by direct boundary layer control have been investigated within the German national program ADIF and the EU-project EUROSHOCK II and their potential explored. Here, corresponding results are presented and discussed indicating the applicability and benefits of the adaptive control methods considered. It is also demonstrated that generally flow control must be adaptive to work in real aeronautical conditions since these conditions change within the mission flight envelope.
The paper presents a new approach to the quantification of simulation fidelity based on an analysis of pilot guidance strategy. The manoeuvre guidance portrait is conceived as the solution to a low-order equivalent system and to properly allow for pilot adaptation to changing cues and task demands, the model parameters are allowed to vary. Thus the concept of the Adaptive Pilot Model (APM) is proposed and developed. The theoretical foundation to the concept is developed using the familiar spatial variables in flight control, such as distance and speed. Motion is then transformed into temporal variables and drawing on the theory of τ(t)-coupling from visual flow theory (τ(t) is the instantaneous time to contact) the APM model is transformed into a much simpler algebraic relationship when the pilot maintains constant during a deceleration. If we make assumptions about the separation of guidance and stabilisation control strategy, pilot guidance feedback gains are then closely related to the frequency and damping of the APM structure. Results are presented from the analysis of simulation trials with pilots flying an acceleration-deceleration manoeuvre that show strong correlation with the τ(t)-based guidance strategy. The interpretation of the theory in terms of simulation fidelity criteria is discussed.
In this paper, the design and evaluation of a helicopter trajectory tracking controller are presented. The control algorithm is implemented using the feedback linearization technique and the two time-scale separation architecture. In addition, an on-line adaptive architecture that employs a Sigma-Pi neural network, which is simple in its structure so that it is easily applicable to on-line adaptation, compensating the model inversion error caused by the deficiency of full knowledge of helicopter dynamics is applied to augment the attitude control system. Trajectory tracking performance of the control system is evaluated using a generic helicopter model simulation program. It is shown that the on-line neural network in an adaptive control architecture is very effective in dealing with the performance degradation problem of the trajectory tracking control caused by insufficient information of dynamics.
Applying adaptronics to helicopters has a high potential to significantly suppress noise, reduce vibration, and increase the overall aerodynamic efficiency. This paper presents recent investigations on a very promising specific concept described as Adaptive Blade Twist (ABT). This concept allows us to directly control the twist of the helicopter blades by smart adaptive elements. This influences positively the main rotor area which is the primary source for helicopter noise and vibration. Since the interaction of non-stationary helicopter aerodynamics and elastomechanical structural characteristics of the helicopter blades causes flight envelope limitations, vibration and noise, a good comprehension of the aerodynamics is essential for the development of structural solutions to effectively influence the local airflow conditions and finally develop the structural concept. With respect to these considerations, the ABT concept will be presented.
A theoretical analysis of on-line autonomous intelligent adaptive tracking controller based on emotional learning model in mammalians brain (BELBIC) for aerospace launch vehicle is presented. The control algorithm is provided with some sensory inputs and reward signal, subsequently it autonomously seeks the proper control signal to be executed by actuators, thus eliminating tracking error without pre-knowledge of the plant dynamics. The algorithm is very robust and fast in adaptation with dynamical change in the plant, due to its on-line learning ability. Development and application of this algorithm for an aerospace launch vehicle during atmospheric flight in an experimental setting is presented to illustrate the performance of the control algorithm.
The German Remote Sensing Data Center (DFD) of the German Aerospace Center (DLR) has been operating a ground segment for High Resolution Picture Transmission (HRPT) data acquisition, archiving, and distribution since the early 1980s. The station's visibility covers all of Europe. DFD started with the generation of thematic level-3 AVHRR value-added products consisting of Multichannel Sea Surface Temperatures (MCSST) and Normalized Difference Vegetation Indices (NDVI) in March 1993 [8]. Additionally, calibrated and registered 5-channel image subsets in two areas have been generated for supporting user-specific applications since 1994 [8]. The status of the current level-3 product generation chain as well as corresponding processing algorithms are presented. Perspectives are introduced to improve the existing products in terms of channel 1 and 2 radiometric optimization by implementing an atmospheric correction scheme, as well as to correct the solar channels for anisotropic reflectance with respect to different surfaces. As AVHRR data proved to be one of the major sources to derive global information on different land-oriented parameters, special emphasis is given in this paper on methods to extract land cover, the fraction of Absorbed Photosynthetic Active Radiation (fAPAR), and Leaf Area Index (LAI) with respect to operational use. Furthermore, different algorithms were discussed to derive Land Surface Temperatures (LST) by estimating surface emissivity based on NDVI time synthesis. First results over Germany are shown, problems addressed, and outlines for operational usage are given.
Top-cited authors
Wei Huang
  • National University of Defense Technology
Wang Honglun
  • Beihang University (BUAA)
Abdelouahed Tounsi
  • University of Sidi-Bel-Abbes
Mohammed Sid Ahmed Houari
  • University Mustapha Stambouli of Mascara
Haibin Duan
  • Beihang University (BUAA)