TY - JOUR A1 - Grosch, A. A1 - Beirow, Bernd T1 - Ausgewählte bauphysikalische Probleme von Wohngebäuden in Stahl-Leichtbauweise Y1 - 2001 ER - TY - GEN A1 - Kober, Markus A1 - Beirow, Bernd A1 - Singh, Kai Navtej T1 - Towards the Isogeometric Aero-Engine T2 - Proceedings of 16th German LS-DYNA Forum, 11.-12. Oktober 2022, Bamberg Y1 - 2022 UR - https://www.dynamore.it/en/training/conferences/upcoming/16th-german-ls-dyna-forum-2022/preliminary-agenda#tag-1 SN - 978-3-9816215-8-7 ER - TY - GEN A1 - Gambitta, Marco A1 - Beirow, Bernd A1 - Schrape, Sven T1 - A Digital Twin of Compressor Blisk Manufacturing Geometrical Variability for the Aeroelastic Uncertainty Quantification of the Aerodynamic Damping T2 - Turbo Expo 2022 : Rotterdam Ahoy Convention Centre, Rotterdam, The Netherlands, Conference and Exhibition: June 13 – 17, 2022 N2 - This study is centered on the aeroelastic problem for axial compressors blisk airfoils in presence of geometrical uncertainties. The combined problem of structural dynamics and unsteady aerodynamics is of interest for these machines due to the stress induced by the blades vibration. In this field, deviations from the nominal cyclic symmetry (in geometry, material or fluid properties) are generally referred to as mistuning. In particular, the geometrical mistuning is addressed resulting from the manufacturing process of blisk airfoils. The impact of these uncertainties on the aeroelastic problem is evaluated, focusing on the aerodynamic damping. The analysis of the manufacturing geometrical variability is approached in a probabilistic manner. A model representing the uncertainty is created starting from a dataset of optical surface scans. The measured geometries are parameterized in order to numerically describe the differences from the nominal geometry with a set of variables. The creation of a mean geometry of the measured blades allows to simplify the description of the uncertainty, which can be then modelled describing the distributions of geometrical deviations over the blade height. In order to create a stochastic model for the geometrical uncertainty, a data reduction method is implemented in the model. This aims to describe the variability within a minimum required accuracy while using a minimal set of variables. For this purpose, an Autoencoder is used to define a compressed representation of the dataset of interest. The method is based on the training of a Neural-Network, which tries to represent the identity function for the given data while forcing a variables reduction in the intermediate layers. A regularization method for the reduced variables is also introduced in order to avoid correlations and normalize the distributions. The computation of the aerodynamic damping is performed using a CFD solver. A steady-state representation of the investigated axial compressor rig is validated using available experimental data. The unsteady computations are done for one configuration at one shaft speed, which is representative of two relevant crossings in the Campbell diagram for the studied blisk. This indicates resonance conditions for two vibrational mode shapes of the component. The Aerodynamic Influence Coefficients (AIC) method is used to calculate the aerodynamic damping curve for the two vibrational mode shapes of interest. This allows to obtain the damping values over the different inter-blade phase angles with one single solution per mode shape, while reducing the domain to a sub-assembly of the investigated blisk. The Uncertainty Quantification (UQ) uses the implemented geometrical variability model and the defined solution method for the calculation of the aerodynamic damping. To describe the input uncertainty (manufacturing geometrical variability) the space of the variables resulting from the Autoencoder data reduction is used. A sampling is generated, representing with each sample a set of three mistuned blades. For each sample, the three resulting blade surfaces are inserted in the AIC setup, representing the vibrating blade as well as the relative direct upstream and downstream blades. This allows to evaluate the uncertainty on the amplitude and phase of the influence coefficients relative to the three blades and finally on the aerodynamic damping curve. The data reduction provided by the Autoencoder proved to be very efficient, especially if compared to linear methods as the principal components analysis. This allowed to include in the UQ multi-passage variations for a better representation of a real geometry. The output uncertainty on the aerodynamic damping could therefore be evaluated taking these effects in consideration. The results can be combined in an aeroelastic reduced order model with the mistuning of the mechanical properties of the component to represent the mistuned blades vibrations. KW - Digital Twin KW - Compressor Blisk KW - Aerodynamic Damping Y1 - 2022 UR - https://asme-turboexpo.secure-platform.com/a/solicitations/167/sessiongallery/9707/application/82935 ER - TY - CHAP A1 - Beirow, Bernd A1 - Golze, Mark A1 - Popig, Frederik ED - Beran, Jaroslav ED - Bílek, Martin ED - Václavík, Miroslav ED - Žabka, Petr T1 - Vibration Reduction of a Steam Turbine Wheel by Means of Intentional Mistuning T2 - Advances in Mechanism Design III N2 - A last stage steam turbine wheel is analyzed with the objective to alleviate the flutter susceptibility by employing intentional mistuning (IM). In particular, the operation at nominal speed under part-load conditions may cause unfavorable flow conditions facilitating flow separation. In consequence, negative aerodynamic damping ratios occur for the first bending mode family in some circumstances. Employing intended alternate mistuning of adequate magnitude has proved to be a promising measure to stabilize rotors in terms of avoiding self-excited vibration phenomena. From the manufacturing point of view, this two-blade design is advantageous as well and hence, chosen here as a first measure to attenuate flutter susceptibility. Two prototypes of bladed disks series have been made, which are exhibiting small but unavoidable deviations from the design intention due to manufacturing. The real blade alone frequencies have been identified within foregoing experimental investigations. Numerical modal analyses carried out for the prototypes as manufactured finally reveal that there is an additional positive contribution of random mistuning in terms of further enhancing the least aerodynamic damping ratio. Another promising and robust IM pattern is found by using generic algorithms to optimize the least aerodynamic damping ratio yielding stable conditions at any time as well. Moreover, it shows that IM combined with random mistuning also mitigates the maximum forced response at part-speed conditions. KW - steam turbine KW - vibration KW - intentional mistuning Y1 - 2022 SN - 978-3-030-83593-4 SN - 978-3-030-83596-5 U6 - https://doi.org/10.1007/978-3-030-83594-1_8 SN - 2211-0984 SN - 2211-0992 SP - 73 EP - 82 PB - Springer International Publishing CY - Heidelberg ER - TY - CHAP A1 - Kühhorn, Arnold A1 - Beirow, Bernd A1 - Parchem, Roland A1 - Klauke, Thomas T1 - Schaufelschwingungen bei realen Verdichter-Integralrädern (BLISK) T2 - Deutscher Luft- und Raumfahrtkongress 2006, Braunschweig, 06. bis 09. November 2006, Bd. 2 KW - Schaufelschwingungen KW - Blisk Y1 - 2006 SP - 1199 EP - 1208 PB - Dt. Ges. für Luft- und Raumfahrt CY - Bonn ER - TY - GEN A1 - Gambitta, Marco A1 - Kühhorn, Arnold A1 - Beirow, Bernd A1 - Schrape, Sven T1 - Stator Blades Manufacturing Geometrical Variability in Axial Compressors and Impact on the Aeroelastic Excitation Forces T2 - Proceedings of ASME Turbo Expo 2021, ASME Paper Number: GT2021-59642 N2 - The manufacturing geometrical variability is a source of uncertainty, which cannot be avoided in the realization of a machinery. Deviations of a component geometry from its nominal design are inevitably present due to the manufacturing process. In the case of the aeroelastic forced response problem within axial compressors, these uncertainties may affect the vibration characteristics. For this reason, the impact of geometrical uncertainties due to the manufacturing process onto the modal forcing of axial compressor blades is investigated in this study. The research focuses on the vibrational behavior of an axial compressor rotor blisk (blade-integrated disk) and in particular the amplitude of the forces acting as source of excitation on the vibrating blades (modal forcing). Within this context, the geometry of the upstream stator plays an important role as in general the main harmonics of the rotor excitation forces are produced by its wake. Therefore, small variations of the upstream stators geometries, such as the ones caused by the manufacturing process, may affect the resulting forcing. In particular, the geometrical variability of the upstream stator implies that the hypothesis of a cyclic-symmetrical flow is no longer valid. This may cause the introduction of lower harmonic components in the modal forces, generally referred to as Low Engine Orders (LEO). The geometrical variability is modelled starting from a series of optical surface scans. A set of optical measurements of manufactured stator blades originating from the same nominal design constitutes the baseline dataset on which the geometrical model is built. The measured blades as well as the relative nominal geometry are parametrized to describe the individual blades surfaces. The parameterization is accomplished by slicing the surfaces in radial sections and describing each of these with a set of NACA-like parameters [1]. The measured geometrical deviations from the nominal model can therefore be described as an offset of such parameters. A reduced representation of the variables representing the input uncertainty (noise variables) is obtained via Principal Components Analysis. Afterwards a sampling on the reduced noise variables domain can be done to represent the modelled uncertainty and perform an Uncertainty Quantification (UQ) on the relative quantities of interest, in this case the modal forcing. The computation of the modal forcing is done through a CFD solver, computing the unsteady flow field around the rotor blades. The domain considered in this case is a 1.5 stage of the axial compressor, including the rotor and the up- and down-stream stators. The solutions are initialized from a validated steady state solution of the considered compressor rig. The time-dependent pressure field calculated on the rotor blades is projected onto the relative vibrational mode shapes of interests (from structural modal analyzes). The resulting forces are analyzed by means of their spectrum, evaluating the amplitudes for the present engine orders (higher harmonics of the shaft mechanical speed). The UQ uses Monte Carlo methods to evaluate the impact of the geometrical variability onto the modal forcing. The modelled uncertainty on the geometries is introduced into the CFD solver to compute the deviations on the quantities of interest. A reconstruction of the forces acting on the rotor during one revolution is obtained. This allows to evaluate the uncertainty on the present engine orders as well as the possible rise of LEO for the rotor blades in presence of a mistuned upstream stator. [1]: Lange A., Vogeler K., Gümmer V., Schrapp H. and Clemen C. (2009). “Introduction of a Parameter Based Compressor Blade Model for Considering Measured Geometry Uncertainties in Numerical Simulation.” Proceedings of ASME Turbo Expo. GT2009-59937 Y1 - 2021 UR - https://asme-turboexpo.secure-platform.com/a/solicitations/137/sessiongallery/6794/application/59642 ER - TY - GEN A1 - Maywald, Thomas A1 - Beirow, Bernd A1 - Kühhorn, Arnold T1 - Mistuning und Dämpfung von Radialturbinenrädern T2 - MTZ - Motortechnische Zeitschrift N2 - Moderne Verbrennungskraftmaschinen müssen ein stetig wachsendes Anforderungsprofil in Bezug auf Wirtschaftlichkeit, Leistung und Umweltfreundlichkeit erfüllen. In diesem Zusammenhang hat die Turboaufladung von Verbrennungsmotoren an Bedeutung gewonnen. Bei Turboladern kleiner und mittlerer Baugröße, deren Turbinen einen Durchmesser zwischen 30 und 250 mm aufweisen, kommen vornehmlich gegossene Laufräder zum Einsatz. Am Institut für Verkehrstechnik der Brandenburgischen Technischen Universität Cottbus-Senftenberg wurde im Rahmen eines FVV-Forschungsvorhabens der Einfluss charakteristischer Betriebsgrößen eines Turboladers auf das strukturdynamische Verhalten solcher Radialturbinenräder untersucht. Y1 - 2015 U6 - https://doi.org/10.1007/s35146-015-0043-7 SN - 2192-8843 VL - 76 IS - 06 SP - 68 EP - 75 ER - TY - GEN A1 - Beirow, Bernd A1 - Kühhorn, Arnold A1 - Weber, Robby A1 - Popig, Frederik T1 - Vibration Analyses of an Axial Turbine Wheel With Intentional Mistuning T2 - Turbo Expo 2020, Virtual Conference, Virtual Conference and Exhibition, Online, September 21 – 25, 2020 N2 - The last stage bladed disk of a steam turbine is analyzed with respect to both flutter susceptibility and limitation of forced response. Due to the lack of variable stator vanes unfavorable flow conditions may occur which can lead to flow separation in some circumstances. Consequently, there is the risk of flutter in principle, particularly at nominal speed under part load conditions. For this reason, intentional mistuning is employed by the manufacturer with the objective to prevent any self-excited vibrations. A first step in this direction is done by choosing alternate mistuning, which keeps the manufactural efforts in limits since only two different blade designs are allowed. In this sense, two different series of blades have been made. However, it is well known that small deviations from the design intention are unavoidable due to the manufacturing procedure, which could be proved by bonk tests carried out earlier. The influence of these additional but unwanted deviations is considered in numerical simulations. Moreover, the strong dependence of blade frequencies on the speed is taken into account since it significantly attenuates the blade to blade frequency difference in this particular case. Within an academic study the turbine wheel is modelled as blade integrated disk in order to demonstrate fundamental effects of intentional mistuning on flutter susceptibility and forced response. For that purpose, reduced order models are built up by using the subset of nominal system mode approach introduced by Yang and Griffin [1], which conveniently allows for taking into account both differing mistuning patterns and the impact of aeroelastic interaction. Focusing on the first flap mode it could be shown that a mitigation of flutter susceptibility is achieved by prescribing alternate mistuning, which indeed affects an increase of originally small aerodynamic damping ratios. Nevertheless, the occurrence of negative damping ratios could not be completely precluded at part load conditions. That is why optimization studies are conducted based on genetic algorithms with the objective function of maximizing the lowest aerodynamic damping ratios. Again only two different blade designs are admitted. Finally, mistuning patterns could be identified causing a tremendous increase of aerodynamic damping ratios. The robustness of the solutions found could be proved by superimposing additional random mistuning. Another study is focused on the impact of mistuning strength. Further analyses are addressing the forced response at part speed conditions, where different resonance crossings are becoming apparent in the Campbell plot. An increase of the forced response compared to the tuned counterpart is partly unpreventable because of unfavorable aerodynamic damping curves. Independently, the maximum forced response has to be limited also in case of applying large intentional mistuning. [1] Yang, M. T., Griffin, J. H., „A Reduced-Order model of Mistuning Using a Subset of Nominal System Modes“. J Eng Gas Turb Power, 123, pp. 893-900 (2001). Y1 - 2020 UR - https://asme-turboexpo.secure-platform.com/a/solicitations/105/sessiongallery/5325/application/45830 ER - TY - GEN A1 - Figaschewsky, Felix A1 - Kühhorn, Arnold A1 - Beirow, Bernd A1 - Giersch, Thomas A1 - Schrape, Sven T1 - Analysis of mistuned forced response in an axial high-pressure compressor rig with focus on Tyler–Sofrin modes T2 - The Aeronautical Journal N2 - This paper aims at contributing to a better understanding of the effect of Tyler–Sofrin Modes (TSMs) on forced vibration responses by analysing a 4.5-stage research axial compressor rig. The first part starts with a brief review of the involved physical mechanisms and necessary prerequisites for the generation of TSMs in multistage engines. This review is supported by unsteady CFD simulations of a quasi 2D section of the studied engine. It is shown that the amplitude increasing effect due to mistuning can be further amplified by the presence of TSMs. Furthermore, the sensitivity with respect to the structural coupling of the blades and the damping as well as the shape of the expected envelope is analysed. The second part deals with the Rotor 2 blisk of the research compressor rig. The resonance of a higher blade mode with the engine order of the upstream stator is studied in two different flow conditions realised by different variable stator vane (VSV) schedules which allows to separate the influence of TSMs from the impact of mistuning. A subset of nominal system modes representation of the rotor is used to describe its mistuned vibration behaviour, and unsteady CFD simulations are used to characterise the present strength of the TSMs in the particular operating conditions. Measured maximum amplitude vs blade pattern and frequency response functions are compared against the predictions of the aeromechanical models in order to assess the strength of the TSMs as well as its influence on vibration levels. Y1 - 2019 U6 - https://doi.org/10.1017/aer.2018.163 SN - 2059-6464 IS - 123 SP - 356 EP - 377 ER - TY - GEN A1 - Beirow, Bernd A1 - Kühhorn, Arnold A1 - Figaschewsky, Felix A1 - Hönisch, Peter A1 - Giersch, Thomas A1 - Schrape, Sven T1 - Model update and validation of a mistuned high-pressure compressor blisk T2 - The Aeronautical Journal N2 - In order to prepare an advanced 4-stage high-pressure compressor rig test campaign, details regarding both accomplishment and analysis of preliminary experiments are provided in this paper. The superior objective of the research project is to contribute to a reliable but simultaneously less conservative design of future high pressure blade integrated disks (blisk). It is planned to achieve trend-setting advances based on a close combination of both numerical and experimental analyses. The analyses are focused on the second rotor of this research compressor, which is the only one being manufactured as blisk. The comprehensive test program is addressing both surge and forced response analyses e.g. caused by low engine order excitation. Among others the interaction of aeroelastics and blade mistuning is demanding attention in this regard. That is why structural models are needed, allowing for an accurate forced response prediction close to reality. Furthermore, these models are required to support the assessment of blade tip timing (BTT) data gathered in the rig tests and strain gauge (s/g) data as well. To gain the maximum information regarding the correlation between BTT data, s/g-data and pressure gauge data, every blade of the second stage rotor (28 blades) is applied with s/g. However, it is well known that s/g on blades can contribute additional mistuning that had to be considered upon updating structural models. Due to the relevance of mistuning, efforts are made for its accurate experimental determination. Blade-by-blade impact tests according to a patented approach are used for this purpose. From the research point of view, it is most interesting to determine both the effect s/g-instrumentation and assembling the compressor stages on blade frequency mistuning. That is why experimental mistuning tests carried out immediately after manufacturing the blisk are repeated twice, namely, after s/g instrumentation and after assembling. To complete the pre-test program, the pure mechanical damping and modal damping ratios dependent on the ambient pressure are experimentally determined inside a pressure vessel. Subsequently the mistuning data gained before is used for updating subset of nominal system mode (SNM) models. Aerodynamic influence coefficients (AICs) are implemented to take aeroelastic interaction into account for forced response analyses. Within a comparison of different models, it is shown for the fundamental flap mode (1F) that the s/g instrumentation significantly affects the forced response, whereas the impact of assembling the compressor plays a minor role. KW - Blisk KW - Mistuning KW - Aerodynamic damping Y1 - 2019 U6 - https://doi.org/10.1017/aer.2018.149 SN - 2059-6464 SN - 0001-9240 VL - 123 IS - 1260 SP - 230 EP - 247 ER - TY - GEN A1 - Henke, Anna-Sophia A1 - Noack, Martin A1 - Geyer, Thomas A1 - Heinrich, Christoph Rocky A1 - Beirow, Bernd A1 - Sarradj, Ennes A1 - Kühhorn, Arnold T1 - Calculation of the Modal Behavior of Structured Sheet Metal T2 - International Journal of Lightweight Materials and Manufacture Y1 - 2019 U6 - https://doi.org/10.1016/j.ijlmm.2019.01.004 SN - 2588-8404 ER - TY - CHAP A1 - Beirow, Bernd A1 - Kühhorn, Arnold A1 - Golze, Mark T1 - Experimental and Numerical Investigations of High pressure Compressor Blades Vibration Behavior Considering Mistuning KW - vibration Y1 - 2005 ER - TY - GEN A1 - Figaschewsky, Felix A1 - Kühhorn, Arnold A1 - Beirow, Bernd A1 - Giersch, Thomas A1 - Schrape, Sven A1 - Nipkau, Jens T1 - An inverse approach to identify tuned aerodynamic damping, system frequencies and mistuning – Part 3: Application to engine data T2 - ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition, June 17–21, 2019, Phoenix, Arizona, USA N2 - A novel approach for the identification of tuned aerodynamic damping, system frequencies, forcing and mistuning has been introduced in the first part of this paper. It is based on the forced response equations of motion for a blade dominated mode family. A least squares formulation allows to identify the system’s parameters directly from measured frequency response functions (FRFs) of all blades recorded during a sweep through a resonance. The second part has dealt with its modification and application to experimental modal analyses of blisks at rest. This 3rd part aims at presenting the application of the approach to blade tip timing (BTT) data acquired in rig tests. Therefore, blisk rotors of two different engines are studied: a single stage fan rig and a 4.5 stage high pressure compressor (HPC) rig. The rig test campaign of the fan blisk included also an intentional mistuning experiment that allows to study the performance of the identification approach for a similar rotor with two different mistuning levels. It is demonstrated that the approach can identify aerodynamic damping curves, system frequencies, mistuning pattern and forced travelling wave modes (TWMs) from state of the art BTT data monitored during rig or engine tests. All derived mistuning patterns could be verified with reference measurements at standstill. The derived aerodynamic damping curves and system frequencies show a reasonable agreement with simulations. For the HPC case a multitude of excited TWMs could be identified which also lines up with previous simulations. KW - Damping KW - Engines KW - Blades KW - Engineering simulation KW - Rotors KW - Compressors KW - Modal analysis Y1 - 2019 UR - https://asmedigitalcollection.asme.org/GT/proceedings-abstract/GT2019/58684/V07AT36A014/1067111 SN - 978-0-7918-5868-4 U6 - https://doi.org/10.1115/GT2019-91337 ER - TY - GEN A1 - Beirow, Bernd A1 - Kühhorn, Arnold A1 - Figaschewsky, Felix A1 - Bornholm, Alfons T1 - Vibration analysis of a mistuned axial turbine blisk T2 - ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition, June 17–21, 2019, Phoenix, Arizona, USA N2 - An axial turbine blisk for turbocharger applications is analyzed with respect to the effect of intentional mistuning on the forced response. Originally, the intentional mistuning pattern has been designed by employing a genetic algorithm optimization in order to reduce the forced response caused by low engine order excitation (LEO) of the fundamental flap mode. The solution found has been implemented in a prototype of that blisk. For the purpose of comparison, a second reference blisk has been manufactured without intentional mistuning. The actual mistuning distributions of the blisks have been identified by employing blade-by-blade impact testing. Alternatively, a new inverse approach has been employed, which is based on a least squares formulation and benefits from less experimental effort. Based on the information gained by the aforementioned testing procedures, subset of nominal systems (SNM)-models have been updated, which allow for considering the aeroelastic coupling by means of aerodynamic influence coefficients (AIC). Despite of small but unavoidable deviations from the design intention it could be proved within numerical simulations that the intended 70 per cent reduction of the maximum forced response is nevertheless achieved. In addition, the paper is addressing the effect of the aforementioned intentional mistuning pattern on a higher mode, which is relevant for the durability as well. Hence, new SNM-models have to be updated in order to calculate the forced response due to EO-excitation caused by the nozzle guide vane. Although the original mistuning pattern has been optimized solely for reducing the forced response of the fundamental flap mode, it hardly affects the higher mode forced response in a negative manner. KW - Turbines KW - Vibration analysis KW - Blades KW - Excitation KW - Computer simulation KW - Design Y1 - 2019 UR - https://asmedigitalcollection.asme.org/GT/GT2019/volume/58691 SN - 978-0-7918-5869-1 U6 - https://doi.org/10.1115/GT2019-92047 ER - TY - GEN A1 - Weber, Robby A1 - Kühhorn, Arnold A1 - Beirow, Bernd T1 - Mistuning and Damping of Turbine and Compressor Impellers T2 - MTZ worldwide N2 - Turbocharging is known to be a well-established technology for an engine's efficiency and power output by forcing extra compressed air into the combustion chamber. The centrifugal loads, necessary flow deflections, unsteady pressure fluctuations, and structural temperature gradients put a high strain on rotating components. Additionally, those components are prone to high-cycle fatigue. The Chair of Structural Mechanics and Vehicle Vibrational Technology at the BTU Cottbus-Senftenberg investigated the impact of manufacturing tolerances on the vibrational behavior of several turbine and compressor impellers. Finally, it is shown that intentional mistuning can lead to significantly lower stresses. KW - Mistuning KW - Damping KW - Compressor Impellers Y1 - 2019 U6 - https://doi.org/10.1007/s38313-019-0090-4 SN - 2192-9114 VL - 80 IS - 9 SP - 72 EP - 77 ER - TY - JOUR A1 - Beirow, Bernd T1 - Secure Power for Jet Engines KW - Jet Engines Y1 - 2006 ER - TY - JOUR A1 - Osterrieder, Peter A1 - Beirow, Bernd T1 - Advanced dynamic methods for structural evaluation of TV towers, Proseedings of the Centre of Structural Integrity Y1 - 2003 ER - TY - GEN A1 - Figaschewsky, Felix A1 - Beirow, Bernd A1 - Kühhorn, Arnold A1 - Nipkau, Jens A1 - Giersch, Thomas A1 - Powers, Bronwyn T1 - Design and Analysis of an Intentional Mistuning Experiment Reducing Flutter Susceptibility and Minimizing Forced Response of a Jet Engine Fan T2 - ASME Turbo Expo 2017, GT2017-64621, June 26-30, 2017, Charlotte, NC, USA, Volume 7B N2 - Recent demands for a reduction of specific fuel consumption of jet engines have been opposed by increasing propulsive efficiency with higher bypass ratios and increased engine sizes. At the same time the challenge for the engine development is to design safe and efficient fan blades of high aspect ratios. Since the fan is the very first rotor stage, it experiences significant distortions in the incoming flow depending on the operating conditions. Flow distortions do not only lead to a performance and stall margin loss but also to remarkable low engine order (LEO) excitation responsible for forced vibrations of fundamental modes. Additionally, fans of jet engines typically suffer from stall flutter, which can be additionally amplified by reflections of acoustic pressure waves at the intake. Stall flutter appears before approaching the stall line on the fan’s characteristic and limits its stable operating range. Despite the fact that this “flutter bite” usually affects only a very narrow speed range, it reduces the overall margin of safe operation significantly. With increasing aspect ratios of ultra-high bypass ratio jet engines the flutter susceptibility will probably increase further and emphasizes the importance of considering aeromechanical analyses early in the design phase of future fans. This paper aims at proving that intentional mistuning is able to remove the flutter bite of modern jet engine fans without raising issues due to heavily increased forced vibrations induced by LEO excitation. Whereas intentional mistuning is an established technology in mitigating flutter, it is also known to amplify the forced response. However, recent investigations considering aeroelastic coupling revealed that under specific circumstances mistuning can also reduce the forced response due to engine order excitation. In order to allow a direct comparison and to limit costs as well as effort at the same time, the intentional mistuning is introduced in a non-destructive way by applying heavy paint to the blades. Its impact on the blade’s natural frequencies is estimated via finite element models with an additional paint layer. In parallel, this procedure is experimentally verified with painted fan blades in the laboratory. A validated SNM (subset of nominal system modes) representation of the fan is used as a computational model to characterize its mistuned vibration behavior. Its validation is done by comparing mistuned mode shape envelopes and frequencies of an experimental modal analysis at rest with those obtained by the updated computational model. In order to find a mistuning pattern minimizing the forced response of mode 1 and 2 at the same time and satisfying stability and imbalance constraints, a multi-objective optimization has been carried out. Finally, the beneficial properties of the optimized mistuning pattern are verified in a rig test of the painted rotor. Copyright © 2017 by Rolls-Royce Deutschland Ltd & Co KG KW - Flutter (Aerodynamics) KW - Desin KW - Jet engines Y1 - 2017 SN - 978-0-7918-5093-0 U6 - https://doi.org/10.1115/GT2017-64621 PB - ASME CY - New York, NY ER - TY - GEN A1 - Beirow, Bernd A1 - Figaschewsky, Felix A1 - Kühhorn, Arnold A1 - Bornholm, Alfons T1 - Modal Analyses of an Axial Turbine Blisk With Intentional Mistuning T2 - ASME Turbo Expo 2017, GT2017-63193, June 26-30, 2017, Charlotte, NC, USA, Volume 7B N2 - The potential of intentional mistuning to reduce the maximum forced response is analyzed within the development of an axial turbine blisk for ship diesel engine turbocharger applications. The basic idea of the approach is to provide an increased aerodynamic damping level for particular engine order excitations and mode shapes without any significant distortions of the aerodynamic performance. The mistuning pattern intended to yield a mitigation of the forced response is derived from an optimization study applying genetic algorithms. Two blisk prototypes have been manufactured a first one with and another one without employing intentional mistuning. Hence, the differences regarding the real mistuning and other modal properties can be experimentally determined and evaluated as well. In addition, the experimental data basis allows for updating structural models which are well suited to compute the forced response under operational conditions. In this way, the real benefit achieved with the application of intentional mistuning is demonstrated. Copyright © 2017 by ASME KW - Turbines KW - Modal Analysis Y1 - 2017 SN - 978-0-7918-5093-0 U6 - https://doi.org/10.1115/GT2017-63193 PB - ASME CY - New York, NY ER - TY - GEN A1 - Figaschewsky, Felix A1 - Kühhorn, Arnold A1 - Beirow, Bernd A1 - Giersch, Thomas A1 - Nipkau, Jens A1 - Meinl, Ferdinand T1 - Simplified Estimation of Aerodynamic Damping for Bladed Rotors, Part 2: Experimental Validation During operation T2 - ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13–17, 2016 N2 - Due to increasing requirements of future engine projects, much effort has been spent on the design of more efficient turbomachinery blades in the recent years. Besides aerodynamic efficiency constraints, these designs need to meet structural criteria ensuring that they are safe and robust with respect to High Cycle Fatigue (HCF). The estimation of the resonant vibration amplitude is done based on the aerodynamic force and the overall damping level. Since, for many applications the contribution of mechanical damping is often rather low compared to the aerodynamic counterpart, the determination of the aerodynamic damping is vital for the estimation of the forced vibration response. This second part is meant to contribute to a simplified computation of the aerodynamic damping during operation by making additional assumptions: The investigated mode family shall not suffer from flutter, has a high reduced frequency and the influence of adjacent blades is negligible. Under these circumstances a simplified approach can be introduced that allows for the computation of the mean value of the aerodynamic damping based on a steady state CFD solution of the regarded stage. It is well known, that the aerodynamic damping of a blade mode family depends on the inter blade phase angle (IBPA) and its direction of propagation, which is not covered by the simplified approach. For higher modes the difference between the minimum and maximum damping is often low and the mean value is a good approximation, whereas for fundamental modes there is often a significant difference. However, it is shown that considering a mistuned vibration response of the rotor, the expected value of the mistuned damping exhibits the mean value of IBPA-dependent aerodynamic damping. CFD simulations of an oscillating airfoil indicate a certain validity range of the simplified approach based on a modified reduced frequency and inlet Mach number, which allows to determine for which industrial applications the approach is most suitable. Finally, this range of validity is verified with experimentally determined overall damping values from strain gauge measurements during operation for 2 different industrial applications, an axial compressor stage of a jet engine and a radial turbine stage of a turbocharger. Copyright © 2016 by Rolls-Royce Deutschland Ltd & Co KG Y1 - 2016 SN - 978-0-7918-4984-2 U6 - https://doi.org/10.1115/GT2016-56458 N1 - Paper No. GT2016-56458 PB - ASME CY - New York, NY [u.a.] ER - TY - GEN A1 - Beirow, Bernd A1 - Maywald, Thomas A1 - Figaschewsky, Felix A1 - Kühhorn, Arnold A1 - Heinrich, Christoph Rocky A1 - Giersch, Thomas T1 - Simplified Determination of Aerodynamic Damping for Bladed Rotors, Part 1: Experimental Validation at Rest T2 - ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13–17, 2016 N2 - Considering both a radial turbine rotor of a turbocharger and an axial compressor test blisk at rest, aerodynamic damping characteristics are experimentally and numerically analyzed. Linear dependencies of modal damping ratios on the ambient pressure or the acoustic impedance, respectively, could be shown within experiments carried out inside a pressure chamber. The impact of the ambient air clearly dominates the modal damping ratios compared to the minor contribution of the structure. Assuming that acoustic emission can be regarded as main source of aerodynamic damping a simplified approach for its determination is introduced which only depends on natural frequency, mode shape and acoustic impedance. It is shown that a satisfying match between experiment and computation is achieved for those cases which are dedicated to sufficiently small ratios between wave lengths of acoustic emissions and blade distances. Y1 - 2016 SN - 978-0-7918-4984-2 U6 - https://doi.org/10.1115/GT2016-56535 N1 - Paper No. GT2016-56535 PB - ASME CY - New York, NY ER - TY - GEN A1 - Sasakaros, Marios A1 - Schafferus, Markus A1 - Wirsum, Manfred A1 - Zobel, Arthur A1 - Vogt, Damian A1 - Nakos, Alex A1 - Beirow, Bernd T1 - Experimental Investigation of Synchronous Flow Induced Blade Vibrations on a Radial Turbine - Part 2: Influence of Different Inlet Guide Vane Configurations T2 - Proceedings of ASME Turbo Expo 2023, Boston, Massachusetts, June 26-30, 2023, Volume 11, A. Aerodynamics excitation and damping, bearing and seal dynamics N2 - The occurrence of blade vibrations in radial turbines leads to limit cycle oscillations, which in time increase the risk of component failure due to high cycle fatigue. In this context, the precise determination of the resonance operating points and the estimation of the vibration magnitudes are essential for an accurate assessment of the service life of the turbocharger components. In radial turbines forced blade vibrations are primarily flow induced. These vibrations are produced by the non-uniform flow field in the circumferential direction which acts on the blades as a cyclic pressure fluctuation. Previous studies have identified the inlet guide vane (IGV) and the spiral turbine housing as the primary sources of the non-uniform flow field. In the present study a thorough experimental investigation of the synchronous blade vibrations of a radial turbine is performed. First, the blade vibration modes were measured experimentally and calculated numerically for the determination of the speed ranges that need to be examined. Subsequently, the vibrations were captured with two redundant measurement systems during real turbocharger operation. Strain gauges were applied on certain blades while eight optical sensors were distributed on the circumference of the turbine shroud for the measurement of the blades tips deflection through a commercial tip-timing system. In the first part, the blade vibrations caused by the “nominal” IGV are presented. Part 2 analyses the changes of the blade vibrations due to the application of two different IGVs. The first IGV has the same number of vanes as the “nominal” IGV. Nevertheless, it generates additional low engine order excitations by intentionally varying the distance between the vanes. Next, an IGV with a higher number of vanes is employed for the excitation at higher frequencies and thus of higher blade modes. Contrary to expectations, certain synchronous vibrations can be measured in the experiments of all IGVs. These cannot be attributed to the spiral turbine casing. KW - synchronous blade vibrations KW - flow induced vibrations KW - mistuning Y1 - 2023 SN - 978-0-7918-8705-9 U6 - https://doi.org/10.1115/GT2023-102243 N1 - Projekt Nr. BTU: 35057002 FVV/AiF-Projekt PB - ASME CY - New York ER -