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Aircraft wings bend while in flight. To see this for yourself, all you have to do is look out of the window during a flight. Beyond a certain limit, these wing deflections become known as Large Deformations. Professor Wolf Krüger of the DLR Institute of Aeroelasticity explains whether these deformations are dangerous, the conditions in which they occur and why modern wings are becoming increasingly elastic.

Flugzeugflügel verbiegen sich während des Fluges. Um sich selbst davon zu überzeugen, genügt während einer Flugreise ein Blick aus dem Fenster. Ab einer bestimmten Auslenkung sprechen die Wissenschaftlerinnen und Wissenschaftler von großen Deformationen. Ob das gefährlich ist, in welchen Grenzen sich diese Deformationen üblicherweise bewegen und warum moderne Flügel zunehmend elastischer werden, erklärt Prof. Dr. Wolf Krüger vom DLR-Institut für Aeroelastik.

Im Arbeitspaket 1.1 wurden zwei generische und zwei komplexere Testfälle hinsichtlich der Vorhersagequalität von Turbulenzmo-dellen im Kontext von Sekundärströmungseffekten untersucht. Es konnte erfolgreich demonstriert werden, dass sich höherwerti-ge Reynoldsspannungsmodelle für die Simulation von mehrstufigen Verdichtern eignen. Darüber hinaus konnte die Kombination aus Reynoldsspannungsmodell und Transitionsmodell für Verdichterströmungen verifiziert werden. Allerdings musste ebenso geschlussfolgert werden, dass Turbulenzmodelle im RANS-Kontext bei immer komplexer werdenden Testfällen und Strömungssi-tuationen aufgrund ihrer konzeptionellen Schwächen an ihre Vorhersagegrenze stoßen. Im Arbeitspaket 1.2 wurde an insgesamt sieben Testfällen aus dem akademischen und dem industriellen Kontext die Eignung des Programmsystems TRACE zur realitätsgetreuen Vorhersage nicht synchroner Schwingungsphänomene erprobt. Die Ergebnisse waren zufriedenstellend und, sofern Vergleichsdaten vorlagen, in guter Übereinstimmung zu Referenzwerten. Die Turbulenzmodel-lierung erwies sich durchweg als kritischste Komponente mit signifikantem Einfluss auf die untersuchten Strömungsphänomene.

This report presents comprehensive flutter calculations of the DLR-D150-VAMP configuration with respect to the influence of structural changes on flutter behaviour. The main focus thereby was on the influence of modifications concerning the position of the engines/nacelles. Furthermore the influence of gyroscopic and thrust vector effects as well as the influence of pylon stiffness regarding the flutter behaviour has been investigated. Finally, the influence of an integrated ultra high bypass ratio engine/nacelle (UHBR) instead of a classical standard CFM56 engine on the flutter stability has been investigated. The results are shown in comparison to the aeroelastic flight envelope for this assessed configuration, which is similar to an A320 aircraft.

The DLR project VicToria brings together disciplinary methods and tools of different fidelity for collaborative multidisciplinary design optimization (MDO) of long-range passenger aircraft configurations, necessitating the use of high-performance computing. Three different approaches are being followed to master complex interactions of disciplines and software aspects: an integrated aero-structural wing optimization based on high-fidelity methods, a multi-fidelity gradient-based approach capable of efficiently dealing with many design parameters and many load cases, and a many-discipline highly-parallel approach, which is a novel approach towards computationally demanding and collaboration intensive MDO. The XRF-1, an Airbus provided research aircraft configuration representing a typical long-range wide-body aircraft, is used as a common test case to demonstrate the different MDO strategies. Additional results are presented for the NASA Common Research Model (CRM) to show their flexibility. Parametric disciplinary models are used in terms of overall aircraft design synthesis, loads analysis, flutter, structural analysis and optimization, engine design, and aircraft performance. The different MDO strategies are shown to be effective in dealing with complex, real-world MDO problems in a highly collaborative, cross-institutional design environment, involving many disciplinary groups and experts and a mix of commercial and in-house design and analysis software.

This paper presents the cybermatrix protocol, a novel approach to multidisciplinary design optimization in the contex of many involved disciplinary experts and high use of high-performance computing resources. The approach is presented from its formal mathematical background to actual on-disk implementation of a running process. As the demonstration case, a twin-engine long-range transport aircraft is optimized. Two optimization processes are shown, an overall aircraft optimization (wing planform free) and a local parameter optimization (fixed planform), with five disciplines between them: overall aircraft wing planform design, aircraft synthesis and mission evaluation, aeroelastic wing airfoil design (RANS flow, trimmed static-aeroelastic state), maneuver loads evaluation and structural wing design (panel aerodynamics, global shell-element FEM model), and gust loads evaluation and selection (panel aerodynamics, dynamic FEM beam model). Optimized solutions are in line with those expected from previous studies under similar conditions, serving as a validation point for the new approach. Moreover, since the approach can deal flexibly with a higher-complexity loads subprocesses, such as configuration-specific selection of design load cases, also some interesting new effects are seen in optimized designs.

The unsteady boundary layer behavior of a supercritical laminar airfoil model, which undergoes limit cycle oscillations in pitch, is investigated by the application of hot-film anemometry. The data basis is a 2D flutter experiment under transonic flow conditions. The laminar airfoil model was elastically mounted with a single degree of freedom in pitch and performed self-excited limit cycle oscillations at a Mach number of$${\text {Ma}}=0.73$$Ma=0.73and a Reynolds number of$${\text {Re}}\approx 2\times 10^6$$Re≈2×106. An analysis of the hot-film signals on the basis of the quasi-wall shear stress is carried out, with which the boundary layer state for steady and unsteady flow is resolved. An algorithm is presented which allows an automated detection of the transition position, so that a correlation between airfoil motion and transition location movement can be quantified. A sudden movement of the boundary layer transition is observed at the upper and lower reversal points of the limit cycle oscillation, while in parts of the up- and downstroke of the laminar airfoil a shock-induced transition occurs. During the limit cycle oscillation, a delayed response of the boundary layer occurs, resulting in a significant phase lag between the movement of the boundary layer transition and the motion of the laminar airfoil.

This paper presents the cybermatrix protocol, a novel approach to multidisciplinary design optimization in the contex of multiple-fidelity disciplinary analyses, many involved disciplines and high use of high-performance computing resources. The approach is presented from its formal mathematical background to actual on-disk implementation of running processes. As the demonstration case, a twin-engine long-range transport aircraft is optimized. Four disciplines are employed: overall aircraft wing planform design, aerodynamic airfoil design using 3D RANS computations, structural wing design using global shell-element FEM model, and loads selection and evaluation process based on low fidelity aerodynamics.

In order to allow for a more economic and environmentally friendly aircraft operation and to fulfill the greener imperative demanded by today's society, fuel savings and cost reduction play a key role in the development of modern aircraft. Besides the efficiency of engines and aerodynamics, the aircraft weight has a major impact on fuel consumption. Reducing structural wing loads caused by atmospheric disturbances (such as gusts, turbulence and wake vortices) thereby became a main research interest of today's aircraft industry. Reducing such loads will allow the aircraft manufacturer building and certifying aircraft to smaller load envelopes, inherently reducing the airframe structural weight and thus reduce fuel, emissions and cost. In this contribution we discuss latest research and development trends in the field of modelling and control for load alleviation on modern, fly-by-wire aircraft. Covering a broad range of relevant topics, the paper is divided into four main subtopics as follows: (i) For analyzing loads on the aircraft in an early design stage with high fidelity simulation techniques, detailed models of the structure and the aerodynamics are generated. Here, we discuss the latest developments on how these complex aeroelastic models (generated by CFD methods, e.g.) are derived and can be approximated by lower order models for facilitating the subsequent controller design. (ii) Directly linked to the modelling in (i) is the feedback control design task. The use of available local measurements as well as virtual sensors for incoming gusts provides a profound basis for alleviating the associated structural loads. This subsection discusses the current trends in control development such as estimator design and robust control techniques to feed back the gust estimates. (iii) Also, in direct link to the modelling (i) but also to the lidar sensors (iv) feedforward control allows for an enhanced load alleviation performance than a pure feedback due to its anticipative behavior (control actions can be taken earlier). In contrast to the feedback approaches (ii), it requires additional forward sensing capabilities, such as lidar for the gust anticipation. Detailed models of these sensors are needed to allow for a dedicated feedforward control design: these lidar sensor models are investigated in great depth within a dedicated activity (iv) and in surrogate models – derived from them – are considered in the feedforward design tool. In this subtopic the surrogate models of the lidar sensors, the wind reconstruction algorithms as well as the feedforward load alleviation design workflow are presented. (iv) The feedforward control approaches in (iii) require measurements of the incoming gust, which shall be provided by advanced lidar measurement systems. The latest development directions for adapting the lidar technology to fulfil the needs of feedforward gust load alleviation applications are discussed in this subtopic. These progresses represent a key enabling technology for viable anticipation of incoming gusts and winds, paving the way for improving the gust load alleviation capabilities for future aircraft generations.

This paper presents an approach to multi-disciplinary optimization (MDO) of transport aircraft developed within the DLR project VicToria that attempts to strike a balance between two broad classes of MDO approaches: those arising from the formal optimization background, and those coming from the aircraft design background. It starts from the observation that any kind of numerical design process can be viewed as an approximation of a formal optimization process, where Jacobians of cost functions may be inexact and are often not explicitly computed. Based on that, a specific MDO problem representation and a highly parallel process assembly and execution protocol (the ``cybermatrix'' protocol) is defined, as well as one possible realization on high-performance computing (HPC) resources. The approach is applied to an optimization of a long-range transport aircraft, employing disciplinary subprocesses for high-fidelity aerodynamic design of wing airfoil shapes, structural sizing of lifting surfaces, and determination and evaluation of design loads.

The DLR project VicToria deals also with high fidelity-based MDO with various approaches. Aside from aerodynamic optimization, further sub-processes are part of the developed MDO processes. They belong to overall aircraft design synthesis, loads analysis, and structural sizing and optimization. The Paper lays out the mentioned MDO sub-processes in order show their contribution and capabilities for the selected MDO process, but also their complexity when dealing with a high fidelity based MDO approach.

The DLR project VicToria brings together design methods that necessite use of high-performance computing, for the multi-disciplinary optimization of a long-range transport aircraft. Three different approaches are followed: the integrated aerostructural wing optimization, the multi-fidelity gradient-based approach, and the many-discipline highly-parallel approach. A common test case is used: XRF1, an Airbus provided research configuration representing a typical long-range wide-body aircraft.

Das DLR entwickelt die "Plattform Windenergie", wovon die Errichtung mehrerer Windenergieanlagen wesentlicher Bestandteil ist. Eine diese Anlagen ist die als modularer Versuchsträger konzipierte Experimentalturbine. Sie soll der Validierung von Methoden und Modellen sowie die Erprobung von Technologien dienen. Ziel des Projekts war die Entwicklung eines entsprechenden Anlagendesigns. Das Ergebnis sind Dokumente, Daten und numerische Modelle, die diese einzigartige Experimentalturbine beschreiben. Das Design der pitchgesteuerten Anlage mit 500 kW Nennleistung wurde durch ein erfahrenes Ingenieurbüro von einem bewährten kommerziellen Entwurf abgeleitet. Seine Besonderheit liegt in der konsequenten Entwicklung für die Erforschung des Rotors der Zukunft. Daher umfasst das Design ein eigens entwickeltes sehr schlankes Rotorblatt, das die Vergleichbarkeit mit der Aerodynamik aktueller Multi-MW-Anlangen herstellt, hohe Schnelllaufzahlen ermöglicht und eine wechselbare Blattspitze besitzt. Die Anlagengröße war ebenso ein Designkriterium wie der Zugang zu allen Baugruppen. Ein besonders dynamisches Pitchsystem und eine Lidar-Plattform auf dem Gondeldach sind weitere Merkmale. Die Experimentalturbine ist durch die Kombination dieser Eigenschaften ein besonderer Versuchsträger für Experimente zur Untersuchung der Aerodynamik, Aeroakustik, Aeroelastik, Regelung und Meteorologie von Windenergieanlagen.

Reale Tests von Windenergieanlagen sind, aufgrund der großen Komponentenabmessungen und wegen der benötigten Extremlasten durch Wind oder Stromnetz, sehr aufwändig. Die Simulation ist ein notwendiges Hilfsmittel bei der Analyse von Windenergieanlagen. Die Güte dieser Analyse hängt stark von der Güte der verwendeten Simulationsmodelle und deren multidisziplinärer Interaktionsfähigkeit ab. Ziel des Projekts WindMUSE war, die langjährigen Erfahrungen beider Verbundpartner, DLR und Fraunhofer IWES, bezüglich hochgenauer multidisziplinärer Modellierung in einem Optimum zu bündeln. Das DLR brachte Methoden, Modelle und Werkzeuge aus der Luftfahrt bezüglich der Abbildung von Wind, atmosphärischer Grenzschicht, Rotoraerodynamik und -aeroelastik, Strukturdynamik und Systemtechnik mit ein und das Fraunhofer IWES brachte langjährige Erfahrungen aus der Entwicklung des Simulationsframeworks OneWind® ein. Entstanden ist eine flexible Simulationsumgebung mit einer werkzeugunabhängigen Beschreibungssprache für Windturbinen, einem Modellgenerator zur Mehrkörpersimulation, einer Komponentenbibliothek mit standardisiertem Functional Mock-up Interface (FMI), der Integrationsfähigkeit weiterer Simulationstools und mehreren transparenten Beispielworkflows. Die Modellbibliothek bildet neben dem flexiblen Turm und Rotor die Meteorologie inkl. konvektiver Grenzschicht, CFD-berechnete 360° Profilpolaren, Abwind- und Querströmeffekte sowie den kompletten Antriebstrang ab. Es wurde gezeigt, dass sich die WindMUSE-Umgebung eignet, eine Blattstrukturdimensionierung mit der Lastrechnung automatisiert zu koppeln, einen alternativen Regler zu testen und reale Hardware im Echtzeitbetrieb zu integrieren. Damit eignet sich die Umgebung auch als Vorbereitung für Realversuche auf der DLR Forschungsplattform Windenergie, mit deren Hilfe künftig die verwendeten Modelle gegenüber Messdaten validiert und z.B. bzgl. elektrischer Komponenten erweitert werden können.

The effect of transition from laminar to turbulent flow on the flutter behaviour is discussed on basis of classical and new wind-tunnel tests. Requirements are defined for numerical methods in order to predict the risk of transition effects on the flutter behavior of transport aircraft with laminar wings. Due to the lack of validation data and reliable transition modelling, the accuracy of flutter predictions for aircraft with laminar wing is expected to be lower than for conventional aircraft. Instead of managing late the risk in an engineering way with possible loss of performance, the authors propose to develop early enough adequate transition models and to prepare a wind-tunnel test at high Reynolds and Mach number conditions with a laminar wing in forced motion.

Aeroelastic effects strongly influence the design of an aircraft. To be able to assess those effects early on, reliable simulation models representing the global aeroelastic properties of a new design are required. At a conceptual or a pre-design stage, an intelligent parameterization concept allows for limited changes of the configuration while the simulation models are adapted accordingly. In the DLR project Integrated Green Aircraft, the goal was to investigate the impact of technologies for the reduction of fuel consumption on the aeroelastic properties of aircraft. One main aspect was the influence of laminar wing design on divergence, flutter and dynamic loads. As the reference aircraft in the project, the concept of a forward-swept wing aircraft with rear-mounted engines has been analysed. An aeroelastic model has been built up in the project. The model design procedure is based on the DLR in-house tool set MONA (ModGen/ NASTRAN). Focus of this design process is the generation of a parameterized structural model, representing the global dynamic properties of the elastic aircraft, but as detailed as reasonable to capture relevant local effects and to result in a feasible structural design. In the article, the aircraft design is presented. The modelling and sizing process for the structure is described. Results of the loads analysis as well as of the aeroelastic stability analyses are discussed.

Since the development of CRISP, a counter rotating integrated shrouded propfan, within a MTU-DLR cooperative project between 1985 and 2000, major advances have been made in various areas of engine development and fan technology. Therefore DLR launched a successive project CRISPmulti in 2010, redesigning the existing blades to increase the level of aerodynamic efficiency, lower the noise emissions and reduce the overall weight. The improvements are made possible by using the method of multidisciplinary optimization involving the DLR Institutes of Propulsion Technology, Aeroelastics and Structures and Design, together with the DLR facility Systemhaus Technik. This paper focuses on the structural mechanics and aeroelastic aspects in design and simulation for a counter rotating fan stage with blades made from carbon fiber reinforced PEEK. With the use of the thermoplastic resin PEEK, a new approach for the manufacturing of fan blades was used. Different from the so called “onion skin configuration” of CRISP the layers of the CRISPmulti blades are stacked up to a flat plate and are reshaped using a hot press. Afterwards the final blade geometry is obtained through a milling process. The material testing program to identify the important material parameters for this approach is specified and adjusted to suit the highly sophisticated simulation techniques. Based on the material properties, the structural mechanic approach for the calculation of the static and dynamic behavior is presented. The aeroelastic simulation validates the forced response and flutter behavior of the CFRP fan blades. The presented design strategy has been successfully completed and undergone a critical design review to prove mechanical and aeroelastic feasibility regarding a planned rig test.

The modal identification of large and dynamically complex structures often requires a multi-point excitation. Sine sweep excitation runs are applied when it is necessary to concentrate more energy on each line of the frequency spectrum. The conventional estimation of FRFs from multi-point excitation requires uncorrelated excitation signals. In case of multi-point (correlated) sine sweep excitation, several sweep runs with altered excitation force patterns have to be performed to estimate the FRFs. An alternative way, which offers several advantages, is to process each sine sweep run separately. The paper first describes the conventional method for FRF estimation in case of multi-point excitation, followed by two alternative methods applicable in case of correlated excitation signals. Both methods generate a virtual single-point excitation from a single run with multi-point excitation. In the first method, an arbitrary structural point is defined as a virtual driving point. This approach requires a correction of the modal masses obtained from modal analysis. The second method utilizes the equality of complex power to generate virtual FRFs along with a single virtual driving point. The computation of FRFs and the modal identification using virtual single-point excitation are explained. It is shown that the correct set of modal parameters can be identified. The application of the methods is elucidated by an illustrative analytical example. It could be shown that the separate evaluation of symmetric and anti-symmetric multi-point excitation runs yield obviously better and more reliable results compared to the conventional method. In addition, the modal analysis of the separate symmetric and anti-symmetric excitation runs is easier, since the stabilization diagrams are easier to interpret. The described methods were successfully applied during the Ground Vibration Tests on Airbus A380 and delivered excellent results. The methods are highly advantageous and may thus be established as a new standard procedure for testing aerospace structures.

Along with the growing importance of software, the quality requirements for software also increase constantly. Every day, we rely on the correct functionality of software we use. Therefore, errors in the software can have a direct impact on us. This also applies to the software developed at the German Aerospace Center, which is used for vibration testing of airplane structures. The software has to meet both, requirements for state-of-the-art functionality and a high quality standard, at the same time. Currently, no such quality assurance concept for the development of software exists. To fulfil the compulsory requirements, it is the purpose of this bachelor thesis to create a quality concept for future software developments. The main component of the concept is a tailored development process, which contains quality assurance measures. The process is largely automated and focusses on the processing of tasks resulting from user feedback. For the implementation of the development process, suitable software tools are selected and evaluated. Furthermore, the implementation is put into operation for a selected development project. In addition to the improved process, software development guidelines are supposed to guarantee compliance with the quality assurance measures. To ensure this, the guidelines have to be improved continuously and changes in the development process must be applied to the guidelines.

The application of composite material in load carrying structural components of an aircraft is rapidly gaining momentum. While part of the reason for this can certainly be attributed to an increasing confidence of designers in the new material as a result of growing experience, two other crucial points can be made. One, the continuous enhancements in the area of automated production technologies, which are an absolute necessity for ensuring consistent quality in a series production. Two, the progress in the development of computational methods to analyze and optimize composite structures in order to fully exploit their possible advantages over homogeneous materials. Nevertheless, it is still by virtue of challenges in their production as well as computational complexity, that full-fledged variable stiffness designs have not yet found their way in industrial large scale applications. Considering the complex path from the stiffness of a single laminate to the aeroelastic performance of an entire aircraft wing, it becomes clear that variable stiffness optimization is a non-trivial, laborious task. Not only does it require a large amount of design variables in order to achieve an adequate resolution, in addition the diverseness of responses impedes the problem definition. The research presented in this thesis aims at an advancement of the computational treatment, i.e. the development of a variable stiffness composite optimization framework, allowing for the consideration of static aeroelastic responses in the structural design of aircraft wings. Considering the different ways of optimizing composite structures, the strategy pursued in this thesis relates to a separation of the problem in three consecutive parts, the advantage being that each step can be handled with the most suitable optimization tools. The first part comprises an optimization based on laminate stiffnesses and is the main subject of this dissertation. It will be discussed in more detail below. The second part involves a stacking sequence optimization on the basis of the optimal stiffnesses derived in the first part. Part three deals with the optimal conversion of stacking sequences to fiber paths suitable for the chosen production technology. Parts two and three do not depend significantly on the physics of the problem. However, since it closely relates to the continuous optimization in part one, the stacking sequence optimization will also shortly be addressed. Part three is not dealt with in this thesis. The composite optimization framework consists of a successive convex subproblem iteration procedure, in which a gradient based optimizer consecutively solves a local approximation problem. Each response to be considered in the optimization, either as objective or as constraint, is approximated as a linear and/or reciprocal function of the laminate membrane and bending stiffness matrices A and D. Together with the laminate thicknesses h, they constitute the design variables in the optimization process. The distribution of design fields - each of which comprises its own set of A, D, h variables - within a structural entity like a wing skin, determines the variable stiffness resolution. Inside the optimization algorithm, stiffness matrices are parametrized by means of lamination parameters, resulting in a reduction in the amount of design variables on the one hand, and the constitution of a continuous, well-posed optimization problem on the other. The response sensitivities with respect to the design variables form an essential input in the assembly of response approximations. In this research, the finite element (FE) software Nastran is applied in order to generate sensitivities. Three major reasons account for this choice: one, the ability of specifying various types of responses, two, the time efficient implementation, and three, its prevalence in the aircraft industry. A Nastran FE model, suitable for the derivation of the required responses, is generated in a parametric model generation process. Aside from the structural FE representation, the model comprises a doublet lattice description for the computation of aeroelastic loads, and a mass model to incorporate non-structural masses like leading and trailing edge or fuel. Structural responses considered in the stiffness optimization are strength, buckling and mass. For strength, a failure criterion in lamination parameter space is adopted. Buckling is covered by a simply supported flat plate buckling model. Aside from the regular structural responses, the aeroelastic responses aileron effectiveness, divergence, and twist are also directly considered in the optimization process. While response values and sensitivities are an immediate result of Nastran, their approximations with respect to the design variables originate from a sensitivity convexification process, ensuring the approximation to incorporate as much reciprocal share as possible. The stiffness optimization fully relying upon the applied aeroelastic loads, a correction strategy by means of a higher order computational fluid dynamics (CFD) method is developed to enhance the doublet lattice aeroelastic loads. Eventually, the functionality of the stiffness optimization framework is verified by three applications, comprising different levels of complexity. Mostly wing skin weight serves as objective to be minimized, but also the maximization of aileron effectiveness for a prescribed weight is demonstrated. The possible advantages of unbalanced over balanced laminates are studied, as well as the influence of different sets of aeroelastic constraints on the achievable minimum wing skin weight. Finally, the modifications implied by an aero load correction are analyzed.

Due to increasing airport congestion the German research project ’Burgernahes Flugzeug’ (BNF, citizen friendly airplane) investigates the potential of aircraft concepts for a possible future air transportation system integrating small airports. Key technologies in aerodynamics for developing new generation quiet, efficient, short take-off and landing (QESTOL) aircraft are studied through the synergistic exploitationof Computational Fluid Dynamics (CFD) and new wind tunnel data. One subject of the conducted investigations is a ninebladed high power propeller mounted in tractor configuration on a wing that is equipped with an active gapless high-lift system. As the aerodynamic focus of the BNF project is on the interaction between the propeller slipstream and the high-lift flowfield, this paper is focussed on the experimental investigation of the propeller performance in comparison with simulation data.

The present paper deals with steady and unsteady aeroelastic simulation results obtained on the High Reynolds Number Aero-Structural Dynamics (HIRENASD) wind tunnel configuration [1], selected as common test case in the numerical aeroelasticity ONERA/DLR cooperation project “NLAS2”. In order to assess the aeroelastic prediction capabilities of the flow solvers elsA (ONERA) and TAU (DLR), 3 types of simulations have been realized: static coupling, harmonic forced motion and dynamic coupling. The set-up of the numerical simulations as well as comparisons with experimental data are presented for selected test cases.

This paper summarizes the plans for the first Aeroelastic Prediction Workshop. The workshop is designed to assess the state of the art of computational methods for predicting unsteady flow fields and aeroelastic response. The goals are to provide an impartial forum to evaluate the effectiveness of existing computer codes and modeling techniques, and to identify computational and experimental areas needing additional research and development. Three subject configurations have been chosen from existing wind tunnel data sets where there is pertinent experimental data available for comparison. For each case chosen, the wind tunnel testing was conducted using forced oscillation of the model at specified frequencies.

This paper addresses the challenges in ground vibration testing and in finite element model updating of superlight aircraft structures where the influence of the surrounding air on experimental modal data is no longer negligible. This so called apparent mass effect can lead to complex experimental mode shapes and to biased estimates for experimental damping ratios and experimental eigenfrequencies obtained from ground vibration test. It is essential that the effect of surrounding air on experimental modal data is properly understood to detect finite element model parameters that need to be adjusted while at the same time to avoid adjustment of unreasonable parameters in model updating.

Dynamic analysis of flexible aircraft typically involves the separation of rigid body and structural dynamics. This approach is justified, if an adequate distance between the frequencies of the elastic and the flight mechanic modes is present. For aircraft structures characterized by relatively low elastic frequencies (e.g. large passenger aircraft or sailplanes) the combined calculation of the coupled rigid body and structural dynamics becomes important and the setup of an integrated aeroelastic model of the aircraft is necessary. This article describes the derivation of the integrated aeroelastic model, composed of governing equations for the translational, the rotational, and the elastic motion. A modal approach is used for the calculation of the elastic deformations of the aircraft, therefore using unconstrained free-free vibration modes from a Finite-Element analysis. The aerodynamic forces are calculated by a CFD solver in Arbitrary Lagrangian Eulerian (ALE) formulation. The integration of all involved disciplines is finally done via a weak coupling approach applying a CSS (Conventional-Serial-Staggered) algorithm. The integrated model is intended to be used for the prediction of maneuver or gust loads.

Industrial flutter prediction for large transport aircraft requires numerical capabilities to predict unsteady aerodynamic forces induced by harmonic oscillations for the complete flight envelope including transonic and separated flow conditions, with shock boundary layer interaction and inverse shock motion. Adequate numerical methods have to model both flow physical and geometrical complexities in more detail than the standard DLM approach. The application of CFD methods results in a more or less time consuming task. On the other hand the aircraft development process, which leads to structural modifications with impact on the flutter behaviour, requires robust and accurate analyses for various stiffness and/or mass conditions in short iteration times. Thus the CFD calculation process should be decoupled from the structural analysis process. This can be achieved by application of aerodynamic reduced order models (ROM). A new method following this strategy is presented. It relies on a limited number of unsteady CFD computations forming the ROM data base, combined with an arbitrary number of Doublet Lattice computations. Thus compatibility with standard DLM based linear flutter prediction process is conserved. The quality of the new approach depends significantly on the choice and number of single CFD computations, as well as on mapping of these results to the DLM aerodynamic data structure. The validation of this approach is performed in a first step by comparing unsteady pressure distributions and generalised airloads for cases of the parameter space, which have not been included in the data base. In a second step the complete process validation requires a common aeroelastic reference test case of adequate complexity. A brief review of available windtunnel data for both unsteady aerodynamics and flutter outlines the shortcomings of these data, like missing measurements of a clear transonic dip significantly below Mach number one, and of inverse shock motions. A new common test configuration with a transonic dip flutter boundary in the Mach number range between 0.80 and 0.95 is proposed. The aircraft geometry from the Drag-Prediction Workshop 4 fulfils the above mentioned unsteady aerodynamic requirements. It is extended to a flutter model of a generic aircraft. The capability of this model is demonstrated by applying the above flutter process. An unsteady aerodynamic ROM is generated in the 3 dimensional parameter space of Mach number, reduced frequency and elastic mode shape. For selected points of this parameter space a sufficient number of unsteady RANS simulations is performed to display unsteady pressure distributions at Mach numbers between 0.6 and 0.90, and reduced frequencies up to 4. A constant lift coefficient of 0.50 has been chosen for all Mach numbers. DLR’s TAU code is applied for attached flow conditions in linear modus. The ROM is completed by performing this procedure for several so called synthetic modes, which are chosen properly to display all realistic structural modes (of the fixed aircraft geometry), without their detailed knowledge.

Within the European research project FUTURE a transonic compressor rotor has been evaluated with respect to its aeroelastic stability. Based on the identical computational mesh different numerical approaches of the DLR flow solver TRACE and the ONERA code elsA are compared in this paper: whereas a time-linearized method based on small time-harmonic perturbations of the steady flow field is chosen in TRACE, elsA uses a non-linear time-domain harmonic balance method in order to determine local aerodynamic work entries and the corresponding global damping values for all inter-blade phase angles (IBPA). Furthermore, nonlinear reference solutions for the critical IBPA range will be provided by both codes using their classical time-marching schemes. In this context the difference in terms of computational effort between the fast linearized or harmonic balance methods and the classical nonlinear techniques shall be highlighted as well.

CFD-CSMcoupled simulations of the 2001/2002 Aerostabil windtunnel experiments, making use of a beam finite element model and RANS computations have not agreed satisfyingly with the experimental data. A detailed analysis has led to the generation of a new shell FEM-model which improved the simulations enormously. The new model showed that the deviations can be explained by airfoil deformations of the wing. By using data from the 2011 FLIB experiment the FE-model has been updated and finally is used for the simulation of a complex steady flow setting.

In this paper the experimental setup and first results for a gust generator experiment will be presented. The experiment was carried out in the transonic wind-tunnel facility in Goettingen (DNW-TWG). The first goal of the experiment was to answer the question, whether it is possible to induce an unsteady flow field with a rigid moving profile as a gust generator. Furthermore the main objective of the experiment was the understanding of the physical phenomena of an impacting gust on an elastic structure. For that purpose a new set-up has been constructed and built, which allows the investigation of two subsequent airfoil-models in the flow of the wind tunnel. The first airfoil is active and should produce a generic gust. The second model located downstream is passive and responds to the unsteady load changes induced by the moving active profile. For a first investigation, amplitudes and frequencies were varied as well as the vertical position of both airfoils relative to each other. The experiment was conducted to deliver a comprehensive data base for validating the DLR in-house fluid-structure-interaction software “PyCSM” which has been developed for aeroelastic simulations in time domain. In order to support the experiment in the preparation phase, fluid-structure-interaction simulations in time domain have been carried out. These simulations mainly show the response amplitudes expected for the passive wing. For that purpose the DLR TAU Code for the calculation of the steady and unsteady aerodynamic loads and a finite element shell model have been used.

The propagation of parameter uncertainty in structural dynamics has become a feasible method to determine the probabilistic description of the vibration response of industrial scale �nite element models. Though methods for uncertainty propagation have been developed extensively, the quanti�cation of parameter uncertainty has been neglected in the past. But a correct assumption for the parameter variability is essential for the estimation of the uncertain vibration response. This paper shows how to identify model parameter means and covariance matrix from uncertain experimental modal test data. The common gradient based approach from deterministic computational model updating was extended by an equation that accounts for the stochastic part. In detail an inverse approach for the identi�cation of statistical parametric properties will be presented which will be applied on a numerical model of a replica of the GARTEUR SM-AG19 benchmark structure. The uncertain eigenfrequencies and mode shapes have been determined in an extensive experimental modal test campaign where the aircraft structure was tested repeatedly while it was 130 times dis- and reassembled in between each experimental modal analysis.

This paper describes the computational analysis of aerodynamic damping for prediction of flutter characteristics of a transonic fan stage that consists of a highly loaded rotor along with a tandem stator. Three dimensional, linearized Navier-Stokes flow solver TRACE is used to numerically analyse the flutter stability of the fan. The linear flow solver enables the modeling of a single blade passage to simulate the desired inter-blade phase angle. The unsteady aerodynamic load on a vibrating blade is obtained by solving the unsteady Navier-Stokes equations on a dynamically deforming grid and the energy exchange method is used to calculate the aerodynamic damping. The calculation of aerodynamic damping for the prediction of flutter characteristics of the fan rotor is carried out with and without considering the influence of the disk. The blade mode shapes from finite element modal analysis are obtained accordingly and the flutter calculations are carried out for three blade vibration modes at the design speed and at part speeds for all possible inter-blade phase angles. Two operating points, one on the working line and the other near stall are investigated at every rotational speed. Different aspects that affect the aerodynamic damping behaviour like part speed operation, variation in unsteady blade surface pressure fluctuation between operating points on the working line and at near stall and the corresponding variation in aerodynamic work, inter-blade phase angle etc., are described. This analysis primarily focuses on the variations in aerodynamic damping of the fan with and without the influence of the disk. In addition, influence and effect of shock wave on the aerodynamic damping is also discussed.

Within the DLR Institute of Aeroelasticity project iGREEN (integrated green aircraft) the aeroelastic behavior of a typical laminar airfoil under unsteady transonic flow conditions was to be investigated. For that a 2D airfoil model was build and tested in the transonic wind tunnel facility DNW-TWG in Goettingen. The model is a 2D-section with a CAST-10 profile, made of carbon reinforced fiber (CRF). It is equipped with 60 unsteady pressure sensors placed in one section and with hot-film sensors located on its upper and lower surface between 10% and 85% in steps of 5% in streamwise direction, 16 on each side. In order to get a smooth surface without disturbing the laminar flow, the hot-film foil is embedded in the model. In the steady case the upper surface of the model was observed with two infrared cameras. The model was installed in the institutes 2D test set-up for forced pitch oscillations. Under harmonic pitch oscillations the time dependent pressure distribution, global forces and the variation of the transition region were measured for three different flow conditions (subsonic, weak transonic and transonic with strong shock and possible separation). The results may lead to a physical understanding of the interaction of unsteady transition and shock position and act as validation data for new unsteady transition prediction models.

Aerospace manufacturers are faced with the challenge of designing systems and components that have to be safer, more reliable, affordable, offer improved passenger comfort and have less environmental impact than their competitors. In addition, they have to systematically reduce development times in order to get new products to market earlier. In the usual case, only a few prototypes are available for testing and experimental verification, in most cases, only towards the end of the development cycle. To shorten development times while still assuring design critical issues like structural integrity and safety, aircraft development teams are faced with several Ground Vibration Testing (GVT) challenges: (i) ensure identification of all critical modes and assess their non-linear behavior; (ii) manage challenging logistical requirements linked to boundary conditions, structure configurations, the typical usage of high number of sensors with multiple shakers and multi-shift testing teams; (iii) reduce test campaign duration to match development schedule and costs; (iv) deliver accurate, validated and traceable reference data in support of the mathematical model updating and of the flutter certification process. This paper will focus on some recent advances in systems and services that allow testing teams to realize an important testing and analysis time reduction without compromising the accuracy of the results. Emphasis will be put on system performance and openness of the selected industrial GVT platform that allows customization in terms of data acquisition and post-processing. Technical advances include efficient handling of very-high channel count (700) data, new capabilities in Normal Mode Testing, and the possibility to stream userdefined shaker excitation signals for dedicated experimental analysis of structures. The new developments will be illustrated by means of a recently conducted Research Ground Vibration Test on an A340-600 aircraft with ONERA, DLR and Airbus as project partners.

For the last 12 years, DLR and ONERA are involved in the realization of the Ground Vibration Testing (GVT) campaigns for AIRBUS civil aircrafts. Continuous developments are carried out in order to improve the test productivity without jeopardizing the quality of the experimental database deliveries, especially in the case of non-linear structural behaviors. Among the recent developments, DLR and ONERA are suggesting new excitation stimuli, such as multi-sine sweeps and low crest factor random signals. These signals have been tested during a research GVT campaign in March and April 2011, on an A340-600 aircraft. This paper is devoted to assessing the relevance of these new signals within an industrial context.

In the context of the flutter analysis of a modern transonic axial compressor two different CFD codes have been compared: based on the same three-dimensional steady solution, the timelinearized Navier-Stokes flow solver TRACE of the German Aerospace Center DLR as well as the industrial time-linearized Euler code Lin3D (developed at MTU Aero Engines) were used to assess the global aerodynamic damping of the rotor. No major differences could be observed for operating points on the working line. However, several operating points (at different shaft speeds) near the surge line were susceptible to flutter according to TRACE whereas Lin3D predicted aeroelastic stability. The capturing of shock movements as well as the modeling of the tip gap is identified to be responsible for these discrepancies. In this paper the sources of the different local excitation and damping mechanisms are highlighted. Moreover influence coefficient representations for the different approaches are compared in the complex plane.

The flow over an OA209 airfoil subjected to a sinusoidal pitching motion under dynamic stall conditions is investigated experimentally by means of time resolved particle image velocimetry (TR-PIV) and surface pressure measurements. Dynamic stall is distinguished by the formation and convection of large scale coherent structures and a delay in massive flow separation. A vortex detection scheme based on an identification function derived directly from the velocity fields is adopted to identify vortex cores. The combination of global time resolved imaging and an automated vortex identification algorithm allows for the investigation of the spatial and temporal evolution of vortical structures within a single oscillation. Furthermore, the mechanisms associated with the dynamic stall delay are considered.

An ongoing numerical investigation of unsteady shock/boundary layer interaction on a 2-d supercritical airfoil in transonic flow is presented. Initially, the finitevolume URANS solver DLR-TAU is used to simulate self-sustained periodic shock oscillations well known as shock buffet. Next, emphasis is put on the fixed-point stability of the steady flow field below the shock buffet onset. Therefore the flow is perturbed in time with small sinusoidal deflections of the airfoil geometry and random impulses. With increasing angle of attack the mean flow is shown to develop a damped aerodynamic resonance, that degenerates finally towards self-amplification. The occurrence of the aerodynamic resonance is closely related to the development of shock-induced separation, accompanied by quasi-steady inverse shock motion.

A method for the computation of real normal modes from complex eigenvectors is described in this paper. The method is based on the exact eigensystem equations, which commonly cannot be used due the problem of modal truncation. A solution to this problem is achieved by a reduction transformation. The theory of the method is described in detail. With the purpose to demonstrate the applicability, a simulated vibration system is employed. In addition, experimental data from modal identification tests on an aircraft are used. It can be shown that the method is able to transform identified complex eigenvectors into real normal modes with good reliability and accuracy.

Steigende Treibstoff- und Rohstoffpreise, Betriebs und Verwertungsvorschriften und an die Umweltverträglichkeit gekoppelte Steuern und Gebühren stellen immer neue Anforderungen an die Flugzeugbauer. Mit einem hochintegrierten aeroelastisch optimierten Flugzeugentwurf lassen sich neue Leistungs- und Sparpotenziale im Flugzeugbau erschließen – ein weiterer Forschungsschwerpunkt, bei dem der neue Versuchsträger ATRA zum Einsatz kommen kann.

Details of unsteady aerodynamics for a two dimensional NLR7301 airfoil with flap were numerically investigated by a Navier-Stokes code in the transonic regime. Unsteady aerodynamic coefficients such as lift and pitching moment show nonlinear aspects in a small amplitude range of forced excitation modes. The pressure distribution of the supersonic region on the upper surface is very sensitive to the incidence angle, which yields the nonlinearity of the unsteady aerodynamics in a small amplitude range of pitching motion. The behaviour of reattachment of the boundary layer behind the shock wave also affects the nonlinearity. Shock wave motions were also examined. Linearized lift and pitching moment coefficients for each amplitude were applied to the Eigenvalue analysis for the flutter boundary. Time domain flutter simulations were also carried out and the result showed that the amplitude of the limit cycle oscillation could be explained by the flutter boundary of the Eigenvalue analysis.

Grundlagen und Ziele des Fachgebiets Instationäre Aerodynamik werden vorgestellt. Der Schwerpunkt liegt dabei auf der Bereitstellung instationärer Luftkräfte für Flatteranalysen von Flugzeugen, die von kleinen Schwingungen der umströmten Struktur induziert werden. Die Entwicklung entsprechender Berechnungsverfahren in Göttingen, die sich, wie auch weltweit überall, lange Zeit eigenständig und unabhängig von den Entwicklungen der stationären Aerodynamik vollzog, wird beginnend mit den historischen Arbeiten von Birnbaum, Possio, Küssner dargestellt. Danach folgt die Darstellung der Entwicklung von numerischenWirbelleiter- und Panelverfahren nach dem Kriege, sowie der Computational Fluid Dynamics (CFD) Verfahren ab 1980, und hier speziell der Highlights am Institut für Aeroelastik. Aktuelle Arbeiten auf dem Gebiet rein strömungsinduzierter Instationaritäten in transsonischen und abgelösten Strömungen (Buffet) werden zusammen mit entsprechenden Göttinger Windkanalexperimenten präsentiert. Prof. Förschings Bedeutung als Förderer des Fortschritts bei der Erforschung der instationären Aerodynamik wird gewürdigt.

First a brief overview of C. Possio's short but outstanding and fruitful carreer is given. This is followed by an outline of the state of the art in flutter and unsteady aerodynamic research, and of the challenges and problems arising in aircraft development at that time, like high speed flight. His first publications until 1937 about gas dynamic and supersonic problems are reviewed. Main focus is put on his 1938 report about unsteady subsonic compressible 2D flow, which became famous and was named after him, because he was the first one who developed an unsteady compressible aerodynamic theory, which was so urgently needed in those years. The theory, based on Prandtl's acceleration potential is briefly outlined. Some discussions and comments in Germany and other countries at that time highlight the importance of this work for the scientific community. Early solutions of Possio's integral equation by himself and later ones by other researchers are presented, as well as the following approaches to extend the theory to 3 dimensional flows before the war, like Kuessner's theory, which was probably influenced by Possio. Finally C. Possio's later scientific contributions to windtunnel interference and to hydrodynamics are sketched. Summarising some developments of the 2nd half of the 20th century, which are based on Possio’s work, demonstrate, that C. Possio created a milestone for modern aircraft research during his very short carreer.

The reduction of flight loads has the potential to reduce structural mass in wing design and to improve passenger comfort. In the DLR project KonTeKst, an actively controlled wing was designed, manufactured and tested in a wind tunnel experiment. The motivation of the activity was to validate methods for the design of combined passive and active load alleviation techniques, including aeroelastic tailoring and active load control. In addition, the experimental activities provided an opportunity to improve and validate modal identification techniques. The wind tunnel model is a flexible composite wing of 1.6 meter (half-) span, with three flaps used for load control. The control design uses H2 optimal blending techniques for input and output. The control strategy is capable of accounting for failure of one flap without critically losing performance. The wind tunnel experiment was performed in a wind tunnel with a maximum speed of 50 m/s. The paper will give an overview of the numerical design activities for wing and load control, of wing design and manufacturing, including the selection and installation of sensors and actuators, and of the wind tunnel set up including the actuation of the wing, data acquisition and online analysis, as well as selected results.

Through the combination of passive and active load alleviation techniques, this paper presents the design, optimization, manufacturing, and update of a flexible composite wind tunnel model. In a first step, starting from the specification of an adequate wing and trailing edge flap geometry, passive, static aeroelastic stiffness optimizations for various objective functions have been performed. The second optimization step comprised a discretization of the continuous stiffness distributions, resulting in manufacturable stacking sequences. In order to determine which of the objective functions investigated in the passive structural optimization most efficiently complemented the projected active control schemes, the condensed modal finite element models were integrated in an aeroelastic model, involving a dedicated gust load alleviation controller. The most promising design was selected for manufacturing. The finite element representation could be updated to conform to the measured eigenfrequencies, based on the dynamic identification of the model. Eventually, a wind tunnel test campaign was conducted in November 2018 and results have been examined in separate reports.