@misc{KuehhornBeirowSchrapeetal., author = {K{\"u}hhorn, Arnold and Beirow, Bernd and Schrape, Sven and Golze, Mark and Kn{\"o}pke, Martin}, title = {Simulation fluidged{\"a}mpfter Strukturschwingungen mittels partitionierter Fluid-Struktur-Kopplung}, series = {Forum der Forschung}, volume = {9}, journal = {Forum der Forschung}, number = {18}, issn = {0947-6989}, pages = {79 -- 86}, language = {de} } @phdthesis{Schrape, author = {Schrape, Sven}, title = {Zur Simulation von Fluid-Struktur-Wechselwirkungen schwingender Verdichtergitter mittels kommerzieller Software}, publisher = {Shaker}, address = {Aachen}, isbn = {978-3-8440-1541-6}, pages = {X, 151}, abstract = {Die vorliegende Arbeit hat zum Ziel, dem Stand der Wissenschaft und Technik entsprechende uni- und bidirektional gekoppelte Berechnungsmethoden unter Verwendung einer partitionierten Kopplung kommerzieller FE- und CFD-Programme am Lehrstuhl Strukturmechanik und Fahrzeugschwingungen der Brandenburgischen Technischen Universit{\"a}t Cottbus zu etablieren. Dazu erfolgt bez{\"u}glich der Kopplungsverfahren eine {\"U}berpr{\"u}fung der Funktionalit{\"a}t am akademischen Beispiel einer querangestr{\"o}mten, elastischen Platte. Vor dem Hintergrund der intensiven Forschung des Lehrstuhls auf dem Gebiet der Strukturdynamik integraler Verdichterlaufr{\"a}der (Blisks) schließt sich die Validierung des eingesetzten Str{\"o}mungsl{\"o}sers hinsichtlich einer transsonischen, instation{\"a}ren Verdichterstr{\"o}mung innerhalb schwingender Schaufelgitter an. Letztlich wird das grundlegende aeroelastische Verhalten eines realen Hochdruckverdichterlaufrades anhand eines unverstimmten, zweidimensionalen Modells analysiert. Ein Vergleich der Methoden zur Berechnung aeroelastischer Parameter ist Bestandteil der Untersuchungen. Basierend auf bidirektional gekoppelten Ergebnissen wird abschließend eine Verifizierung von abgeleiteten {\"a}quivalenten aerodynamischen Elementen innerhalb eines mechanischen, unverstimmten Ersatzmodells vorgestellt.}, language = {de} } @inproceedings{MaywaldKuehhornSchrape, author = {Maywald, Thomas and K{\"u}hhorn, Arnold and Schrape, Sven}, title = {Experimental Validation of a Model Update Procedure Focusing on Small Geometric Deviations}, series = {ECCOMAS VII European Congress on Computational Methods in Applied Sciences and Engineering, Crete, Greece, June 5-10, 2016}, booktitle = {ECCOMAS VII European Congress on Computational Methods in Applied Sciences and Engineering, Crete, Greece, June 5-10, 2016}, abstract = {This contribution presents a model update procedure and its experimental validation using the example of a blade integrated disk rotor. This so called blisk is discretized using the finite element method. It is well known that numerical blisk models based on the ideal tuned design show major differences in structural dynamic behavior compared to the real rotor. In this context a modification of the mechanical simulation model should lead to a better accordance of numerical results and the real blisk characteristics. The described model update procedure utilizes data of an optical 3D measurement system. Using this data enables to identify geometric deviations between the ideal design and its real counterpart. Within the update procedure the originally tuned finite element mesh is modified in order to match the measured geometry of the real part. This is done by defining several morph regions. The outer surface nodes of these morph regions change their position along the surface normal vector until they meet the defined deviation constraint. Based on eigenvalue calculations employing free boundary conditions the sensitivity of structural dynamic behavior is shown with respect to small geometric changes. Finally computed eigenvalues and eigenvectors of the updated simulation model are compared with vibration measurement data. A laser Doppler vibrometer is used to detect the vibration responses of the impact excited structure. All experiments are carried out under technical vacuum conditions in order to minimize ambient air damping. In the context of an experimental modal analysis this low damping condition helps to identify more natural frequencies of the investigated structure. This leads to a much more efficient model validation.}, language = {en} } @misc{FranzKuehhornGierschetal., author = {Franz, Falco and K{\"u}hhorn, Arnold and Giersch, Thomas and Schrape, Sven and Figaschewsky, Felix}, title = {Influence of Inlet Distortions on the Forced Vibration of a High Pressure Compressor Rig}, series = {ASME 2020 Turbo Expo - Virtual Conference, September 2020}, journal = {ASME 2020 Turbo Expo - Virtual Conference, September 2020}, abstract = {The accurate prediction of blade vibrations is a key factor for the development of reliable turbomachines. This paper focusses on forced vibrations. The excitation frequency is an integer multiple of the rotor revolution frequency, which is commonly called engine order. Aerodynamic excitation of blades is created by stator wakes or the potential fields of downstream obstacles, which usually leads to high engine orders correlating to the number of vanes. Resonance crossings appear at higher frequencies corresponding to higher modes. Besides high engine orders, low engine orders not related to the number of vanes may exist. They can be caused by a disturbance of the perfect cyclic symmetry of the flow pattern due to geometry variations or inlet distortions. Inlet distortions result from installation effects, maneuvers or crosswind. Low engine orders affect fundamental modes at high engine speeds. High static loads due to centrifugal forces combined with dynamic excitation and low damping may lead to unacceptable high stresses. This paper aims at getting a better understanding of the simulative prediction of low engine order excitation with special focus on inlet distortions. Under investigation is a 4.5 stage research compressor rig, for which an extensive amount of test data is available. A three dimensional CFD-model of the compressor is used to compute the forcings generated by different distortion patterns. The first two stages are modeled as a full-annulus, which allows to fully resolve the spatial content of the inlet distortion patterns. The rotor 2 blisk is of special interest in this investigation. The propagation of the distortion after stage 2 with rotor 2 is not of interest, therefore the downstream stages are modeled as single passages in order to save computational time. The distortion patterns are the outcome of traversals of different screens with total pressure probes. During distortion measurements, the screens located in the inlet duct were rotated relative to the fixed instrumentation. The traversals in resonance of the first bending mode of rotor 2 with a low engine order four showed a dependency of the screen angle on the vibration amplitude. Acceleration and deceleration maneuvers through this resonance were conducted with screen angles set to those of smallest and highest response. Vibration amplitudes of the blisk rotor are measured by strain gauges and a blade tip timing system. Simulation results are compared against vibration measurements. Aerodynamic damping is calculated with the influence coefficient method. The effects of mistuning are included in the calculation of vibration amplitudes via a subset of nominal system modes model to give a meaningful comparison against real engine hardware. The mistuning distribution of the blisk was identified at rest for the fundamental bending mode. The presence of a 2nd excitation mechanism of unknown source explains the observed test data. This unknown source is not included in the CFD model. A direct comparison of simulation and measurement is still possible by leveraging the observed superposition effects of both excitation sources. The consequent approach is to identify and substract the forcing due to the unknown source, leaving only the delta forcing due to inlet distortions.}, language = {en} } @misc{GambittaKuehhornSchrape, author = {Gambitta, Marco and K{\"u}hhorn, Arnold and Schrape, Sven}, title = {Geometrical Variability Modelling of Axial Compressor Blisk Aerofoils and Evaluation of Impact on the Forced Response Problem}, series = {ASME 2020 Turbo Expo - Virtual Conference, September 2020}, journal = {ASME 2020 Turbo Expo - Virtual Conference, September 2020}, abstract = {The manufacturing process always produces onto the components a certain amount of geometrical uncertainty. This results inevitably in the introduction of a certain amount of variability within the manufactured parts. Even if the differences are small, all the resulting geometries will differ from each other. The present work focuses on the effect of the manufacturing geometrical variability on the high pressure compressor of a turbofan engine for civil aviation. The deviations of the geometry over the axial compressor blades are studied and modelled for the representation in the computational models. Such variability is of particular interest for the forced response problem, where small deviations of the geometry from the ideal nominal model can imply significant differences in the vibrational responses. The information regarding the geometrical mistuning is extracted from a set of manufactured components surface scans of a blade integrated disk (blisk) rotor. The measured geometries are analyzed over a large amount of set radial sections, defining a set of opportune parameters to represent the deviations from the nominal design. A spline fit of the parameters over the radial sections allows the creation of a set of variables describing the geometry. The dimension of the variables domain is reduced using the principal component analysis approach, this allows to obtain an optimal subset of geometrical modes as linear combination of the above mentioned parameters. The reconstruction of the modelled geometries is performed for the implementation in complex CFD and FEM solvers. This is done via the application of the modelled delta nominal-to-measure geometrical offset to the hot geometry of the desired test case. The generated model allows a stochastic representation of the variability, providing an optimal set of variables to represent it. Moreover the approach as defined allows to apply the modelled variability to different blades, e.g. different stators or rotors, utilizing the nominal geometry as input. The aeroelastic analyses considering geometry based mistuning is carried on a test-rig case, focusing on how such variability can affect the modal forcing generated on the blades. A validated CFD model is used to extract the force generated by the unsteady pressure field over the selected vibrational mode shapes of the rotor blades. The blade mode shapes are extracted form a FEM model of the whole blisk and the blades displacements are mapped over the CFD model nodes. The uncertainty quantification of the geometrical variability effect on the modal forcing is performed utilizing Monte Carlo methods. A reduced model for the CFD solution is employed, utilizing a single passage multi blade row which assumes a time-space periodicity solving the governing equations in the frequency domain. This allows for conducting an uncertainty quantification considering the large domain of the variables used to describe the geometries compared to the computational resources needed for the single solution. The unsteady modal forcing is studied as amplitude and phase shift for the different engine orders (frequencies arising from the engine working condition as higher harmonics of the shaft speed). In particular the scatter of the main engine orders forcing amplitudes for the manufactured blades can be compared with the nominal responses to predict the possible amplification due to the geometrical variability. Finally the results are compared to a larger computational model to assess the influence of multiple variable blades in the assembly.}, language = {en} } @inproceedings{HanschkeKuehhornSchrapeetal., author = {Hanschke, Benjamin and K{\"u}hhorn, Arnold and Schrape, Sven and Giersch, Thomas}, title = {Consequences of Borescope Blending Repairs on Modern HPC Blisk Aeroelasticity}, series = {Proceedings of ISROMAC 2017, Maui, Hawaii, December 16-21, 2017}, booktitle = {Proceedings of ISROMAC 2017, Maui, Hawaii, December 16-21, 2017}, pages = {8}, abstract = {Objective of this paper is to analyse the consequences of borescope blending repairs on the aeroelastic behaviour of a modern HPC blisk. To investigate the blending consequences in terms of aerodynamic damping and forcing changes, an exemplary blending of a rotor blade is modelled. Steady state flow parameters like total pressure ratio, polytropic efficiency and the loss coefficient are compared. Furthermore, aerodynamic damping is computed utilising the AIC approach for both geometries. Results are confirmed by SPF simulations for specific nodal diameters of interest. Finally, an unidirectional forced response analysis for the nominal and the blended rotor is conducted to determine the aerodynamic force exciting the blade motion. Fourier transformation of the forcing signal yields to the frequency content as well as the forcing amplitudes. As a result of the present analysis, the amplification of expected blade vibration amplitude is computed.}, language = {en} } @inproceedings{FigaschewskyKuehhornBeirowetal., author = {Figaschewsky, Felix and K{\"u}hhorn, Arnold and Beirow, Bernd and Giersch, Thomas and Schrape, Sven}, title = {Analysis of Mistuned Forced Response in an Axial High Pressure Compressor Rig With Focus on Tyler-Sofrin Modes}, series = {ISABE 2017, ISABE-2017-22614, Manchester, September 3.-8., 2017}, booktitle = {ISABE 2017, ISABE-2017-22614, Manchester, September 3.-8., 2017}, publisher = {ISABE}, pages = {21}, language = {en} } @inproceedings{BeirowKuehhornFigaschewskyetal., author = {Beirow, Bernd and K{\"u}hhorn, Arnold and Figaschewsky, Felix and H{\"o}nisch, Peter and Giersch, Thomas and Schrape, Sven}, title = {Model Update and Validation of a Mistuned High Pressure Compressor Blisk}, series = {Proceedings of ISABE 2017, ISABE-2017-22568, Manchester, September 3.-8., 2017}, booktitle = {Proceedings of ISABE 2017, ISABE-2017-22568, Manchester, September 3.-8., 2017}, publisher = {ISABE}, pages = {14}, language = {en} } @inproceedings{GierschFigaschewskyHoenischetal., author = {Giersch, Thomas and Figaschewsky, Felix and H{\"o}nisch, Peter and K{\"u}hhorn, Arnold and Schrape, Sven}, title = {Numerical Analysis and Validation of the Rotor Blade Vibration Response Induced by High Pressure Compressor Deep Surge}, series = {ASME Turbo Expo 2014: Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-26295}, booktitle = {ASME Turbo Expo 2014: Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-26295}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-4577-6}, doi = {10.1115/GT2014-26295}, pages = {12}, abstract = {The following paper presents a numerical analysis of a deep surge cycle of a 4.5 stage research compressor. The resulting unsteady loads are used to determine the response of two particular rotor blade rows that are then compared to strain gauge data from measurements. Within a deep surge cycle the compressor experiences a rapid change of the flow field from forward to reversed flow. This rapid breakdown is linked to a new mean blade load. Hence, the rapid change in blade loads are able to excite fundamental blade modes similar to an impulse load. The resulting vibration magnitudes might reach critical levels. This paper demonstrates two different approaches to evaluate the unsteady flow during a surge cycle. The first uses a three dimensional, time accurate finite volume solver for viscid compressible flows to calculate the transient surge cycle of the compressor. The compressor itself is represented by a multi-blade-row sector model. The second approach makes use of the same solver and compressor domain to determine steady state characteristics of the HPC in forward, stalled and reversed flow. Based on these characteristics an one dimensional finite volume solver for inviscid compressible flows was developed to determine the transient compressor behavior. The one dimensional solver represents the compressor by source terms that are linked to the previously determined steady state characteristics. Copyright © 2014 by Rolls-Royce Deutschland Ltd \& Co KG}, language = {en} } @misc{WeberKuehhornKlaukeetal., author = {Weber, Robby and K{\"u}hhorn, Arnold and Klauke, Thomas and Schrape, Sven}, title = {The Effect of Sand Erosion on a Compressor Blade and its Modal Properties}, series = {Proceedings of ASME Turbo Expo 2020, Turbomachinery Technical Conference and Exposition, GT2020, September 21-25, 2020, Virtual, Online}, journal = {Proceedings of ASME Turbo Expo 2020, Turbomachinery Technical Conference and Exposition, GT2020, September 21-25, 2020, Virtual, Online}, pages = {9}, abstract = {The wear and damage of High-Pressure Compressor (HPC) blades due to erosion or Foreign Object Damage (FOD) have a significant influence on HPC aerodynamic performance, vibration resistance against High-Cycle Fatigue (HCF) and thus component lifetime. The changes in airfoil geometry reduce the overall engine efficiency. Furthermore extended off-wing engine maintenances due to blade failures are increasing the cost of ownership. The safe operation of every engine within a reduced number of shop visits requires a reliable prediction of future deterioration. This enables the optimization of services and off-wing time. One contribution to this is a better understanding of the component's dynamics and based on this providing an improved wear modeling to reliably predict the remaining lifetime and the decreased efficiency. This contribution determines the material removal of HPC blades due to sand erosion. Originally, this stage was built as a blisk (Blade Integrated Disk). After sand erosion test completion, the blisk was cut into segments containing one airfoil only. First, the material removal is determined for ten blades of one exemplary rotor. A blue light fringe projector is employed to identify the geometrical differences between the eroded blades and the nominal design. Second, realistic finite element models are generated to enable comparable modal analyses of eroded blades. This procedure suffers from unavoidable and mostly random imperfections due to the manufacturing process, which significantly affects the blade surface before the erosion test can be conducted. Therefore, an already published approach is implemented in the third step to predict the blade surface after erosion based on nominal blade design. The investigation is completed by comparing measured and predicted surfaces. Finally, the aforementioned tool is employed to predict the locations and intensities of the material losses and the accompanying change in modal properties of this compressor blade concerning operational time.}, language = {de} } @misc{GambittaBeirowSchrape, author = {Gambitta, Marco and Beirow, Bernd and Schrape, Sven}, title = {Modelling Method for Aeroelastic Low Engine Order Excitation Originating from Upstream Vanes' Geometrical Variability}, series = {Preprints : the multidisciplinary preprint platform}, volume = {2023}, journal = {Preprints : the multidisciplinary preprint platform}, issn = {2310-287X}, doi = {10.20944/preprints202311.0493.v1}, abstract = {The manufacturing geometrical variability in axial compressors is a stochastic source of uncertainty, implying that the real geometry differs from the nominal design. This causes the real geometry to lose the ideal axial symmetry. Considering the aerofoils of a stator vane, the geometrical variability affects the flow traversing it. This impacts the downstream rotor, especially when considering the aeroelastic excitation forces. Optical surface scans coupled with a parametrization method allow for acquiring the information relative to the real aerofoils geometries. The measured data are included in a multi-passage and multi-stage CFD setup to represent the mistuned flow. In particular, low excitation harmonics on the rotor vane are introduced due to the geometrical deviations of the upstream stator. The introduced low engine orders as well as their amplitude depend on the stator geometries and their order. A method is proposed to represent the phenomena in a reduced CFD domain, limiting the size and number of solutions required to probabilistically describe the rotor excitation forces. The resulting rotor excitation forces are reconstructed as a superposition of disturbances due to individual stator aerofoils geometries. This indicates that the problem is linear in in the combination of disturbances from single passages.}, language = {en} } @misc{GambittaBeirowSchrape, author = {Gambitta, Marco and Beirow, Bernd and Schrape, Sven}, title = {Modelling Method for Aeroelastic Low Engine Order Excitation Originating from Upstream Vanes' Geometrical Variability}, series = {International Journal of Turbomachinery Propulsion and Power}, volume = {9}, journal = {International Journal of Turbomachinery Propulsion and Power}, number = {2}, issn = {2504-186X}, doi = {10.3390/ijtpp9020012}, abstract = {The manufacturing geometrical variability in axial compressors is a stochastic source of uncertainty, implying that the real geometry differs from the nominal design. This causes the real geometry to lose the ideal axial symmetry. Considering the aerofoils of a stator vane, the geometrical variability affects the flow traversing it. This impacts the downstream rotor, especially when considering the aeroelastic excitation forces. Optical surface scans coupled with a parametrisation method allow for acquiring the information relative to the real aerofoils geometries. The measured data are included in a multi-passage and multi-stage CFD setup to represent the mistuned flow. In particular, low excitation harmonics on the rotor vane are introduced due to the geometrical deviations of the upstream stator. The introduced low engine orders, as well as their amplitude, depend on the stator geometries and their order. A method is proposed to represent the phenomena in a reduced CFD domain, limiting the size and number of solutions required to probabilistically describe the rotor excitation forces. The resulting rotor excitation forces are reconstructed as a superposition of disturbances due to individual stator aerofoils geometries. This indicates that the problem is linear in the combination of disturbances from single passages.}, language = {en} } @misc{GambittaKuehhornBeirowetal., author = {Gambitta, Marco and K{\"u}hhorn, Arnold and Beirow, Bernd and Schrape, Sven}, title = {Stator Blades Manufacturing Geometrical Variability in Axial Compressors and Impact on the Aeroelastic Excitation Forces}, series = {Journal of Turbomachinery}, volume = {144}, journal = {Journal of Turbomachinery}, issn = {1528-8900}, doi = {10.1115/1.4052602}, pages = {10}, abstract = {The manufacturing geometrical variability is a source of uncertainty, which cannot be avoided in the realization of machinery components. Deviations of a part 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. In particular, the amplitude of the forces acting as a source of excitation on the vibrating blades is studied. The geometrical variability of the upstream stator is investigated as input uncertainty. The variability is modeled starting from a series of optical surface scans. A stochastic model is created to represent the measured manufacturing geometrical deviations from the nominal model. A data reduction methodology is proposed in order to represent the uncertainty with a minimal set of variables. The manufacturing geometrical variability model allows to represent the input uncertainty and probabilistically evaluate its impact on the aeroelastic problem. An uncertainty quantification is performed in order to evaluate the resulting variability on the modal forcing acting on the vibrating rotor blades. Of particular interest is the possible rise of low engine orders due to the mistuned flow field along the annulus. A reconstruction algorithm allows the representation of the variability during one rotor revolution. The uncertainty on low harmonics of the modal rotor forcing can be therefore identified and quantified.}, language = {en} } @misc{HanschkeKuehhornSchrapeetal., author = {Hanschke, Benjamin and K{\"u}hhorn, Arnold and Schrape, Sven and Giersch, Thomas}, title = {Consequences of Borescope Blending Repairs on Modern HPC Blisk Aeroelasticity}, series = {Journal of Turbomachinery}, volume = {141}, journal = {Journal of Turbomachinery}, number = {2}, issn = {1528-8900}, doi = {10.1115/1.4041672}, pages = {7}, abstract = {Objective of this paper is to analyze the consequences of borescope blending repairs on the aeroelastic behavior of a modern high pressure compressor (HPC) blisk. To investigate the blending consequences in terms of aerodynamic damping and forcing changes, a generic blending of a rotor blade is modeled. Steady-state flow parameters like total pressure ratio, polytropic efficiency, and the loss coefficient are compared. Furthermore, aerodynamic damping is computed utilizing the aerodynamic influence coefficient (AIC) approach for both geometries. Results are confirmed by single passage flutter (SPF) simulations for specific interblade phase angles (IBPA) of interest. Finally, a unidirectional forced response analysis for the nominal and the blended rotor is conducted to determine the aerodynamic force exciting the blade motion. The frequency content as well as the forcing amplitudes is obtained from Fourier transformation of the forcing signal. As a result of the present analysis, the change of the blade vibration amplitude is computed.}, language = {en} } @misc{GambittaBeirowSchrape, author = {Gambitta, Marco and Beirow, Bernd and Schrape, Sven}, title = {A Digital Twin of Compressor Blisk Manufacturing Geometrical Variability for the Aeroelastic Uncertainty Quantification of the Aerodynamic Damping}, series = {Turbo Expo 2022 : Rotterdam Ahoy Convention Centre, Rotterdam, The Netherlands, Conference and Exhibition: June 13 - 17, 2022}, journal = {Turbo Expo 2022 : Rotterdam Ahoy Convention Centre, Rotterdam, The Netherlands, Conference and Exhibition: June 13 - 17, 2022}, abstract = {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.}, language = {en} } @misc{GambittaKuehhornBeirowetal., author = {Gambitta, Marco and K{\"u}hhorn, Arnold and Beirow, Bernd and Schrape, Sven}, title = {Stator Blades Manufacturing Geometrical Variability in Axial Compressors and Impact on the Aeroelastic Excitation Forces}, series = {Proceedings of ASME Turbo Expo 2021, ASME Paper Number: GT2021-59642}, journal = {Proceedings of ASME Turbo Expo 2021, ASME Paper Number: GT2021-59642}, abstract = {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., G{\"u}mmer 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}, language = {en} } @misc{FigaschewskyKuehhornBeirowetal., author = {Figaschewsky, Felix and K{\"u}hhorn, Arnold and Beirow, Bernd and Giersch, Thomas and Schrape, Sven}, title = {Analysis of mistuned forced response in an axial high-pressure compressor rig with focus on Tyler-Sofrin modes}, series = {The Aeronautical Journal}, journal = {The Aeronautical Journal}, number = {123}, issn = {2059-6464}, doi = {10.1017/aer.2018.163}, pages = {356 -- 377}, abstract = {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.}, language = {en} } @misc{BeirowKuehhornFigaschewskyetal., author = {Beirow, Bernd and K{\"u}hhorn, Arnold and Figaschewsky, Felix and H{\"o}nisch, Peter and Giersch, Thomas and Schrape, Sven}, title = {Model update and validation of a mistuned high-pressure compressor blisk}, series = {The Aeronautical Journal}, volume = {123}, journal = {The Aeronautical Journal}, number = {1260}, issn = {2059-6464}, doi = {10.1017/aer.2018.149}, pages = {230 -- 247}, abstract = {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.}, language = {en} } @misc{MaywaldHeinrichKuehhornetal., author = {Maywald, Thomas and Heinrich, Christoph Rocky and K{\"u}hhorn, Arnold and Schrape, Sven and Backhaus, Thomas}, title = {Prediction of Geometrically Induced Localization Effects Using a Subset of Nominal System Modes}, series = {ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition June 17-21, 2019 Phoenix, Arizona, USA}, journal = {ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition June 17-21, 2019 Phoenix, Arizona, USA}, isbn = {978-0-7918-5869-1}, doi = {10.1115/GT2019-90884}, pages = {9}, abstract = {It is widely known that the vibration characteristics of blade integrated discs can dramatically change in the presence of manufacturing tolerances and wear. In this context, an increasing number of publications discuss the influence of the geometrical variability of blades on phenomena like frequency splitting and mode localization. This contribution is investigating the validity of a stiffness modified reduced order model for predicting the modal parameters of a geometrically mistuned compressor stage. In detail, the natural frequencies and mode shapes, as well as the corresponding mistuning patterns, are experimentally determined for an exemplary rotor. Furthermore, a blue light fringe projector is used to identify the geometrical differences between the actual rotor and the nominal blisk design. With the help of these digitization results, a realistic finite element model of the whole compressor stage is generated. Beyond that, a reduced order model is implemented based on the nominal design intention. Finally, the numerical predictions of the geometrically updated finite element model and the stiffness modified reduced order model are compared to the vibration measurement results. The investigation is completed by pointing out the benefits and limitations of the SNM-approach in the context of geometrically induced mistuning effects.}, language = {en} } @misc{FigaschewskyKuehhornBeirowetal., author = {Figaschewsky, Felix and K{\"u}hhorn, Arnold and Beirow, Bernd and Giersch, Thomas and Schrape, Sven and Nipkau, Jens}, title = {An inverse approach to identify tuned aerodynamic damping, system frequencies and mistuning - Part 3: Application to engine data}, series = {ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition, June 17-21, 2019, Phoenix, Arizona, USA}, journal = {ASME Turbo Expo 2019: Turbomachinery Technical Conference and Exposition, June 17-21, 2019, Phoenix, Arizona, USA}, isbn = {978-0-7918-5868-4}, doi = {10.1115/GT2019-91337}, pages = {13}, abstract = {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.}, language = {en} } @misc{MaywaldBackhausSchrapeetal., author = {Maywald, Thomas and Backhaus, Thomas and Schrape, Sven and K{\"u}hhorn, Arnold}, title = {Geometric Model Update of Blisks and its Experimental Validation for a Wide Frequency Range}, series = {ASME Turbo Expo 2017, GT2017-63446, June 26-30, 2017, Charlotte, NC, USA, Volume 7A}, journal = {ASME Turbo Expo 2017, GT2017-63446, June 26-30, 2017, Charlotte, NC, USA, Volume 7A}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-5092-3}, doi = {10.1115/GT2017-63446}, pages = {9}, abstract = {The contribution discusses a model update procedure and its experimental validation in the context of blisk mistuning. Object of investigation is an industrial test blisk of an axial compressor which is milled from solid using a state of the art 5-axis milling machine. First, the blisk geometry is digitized by a blue light fringe projector. Digitization is largely automated using an industrial robot cell in order to guarantee high repeatability of the measurement results. Additionally, frequency mistuning patterns are identified based on vibration measurements. Here, the system excitation is realized by a modal impact hammer. The blade response is detected using a laser scanning vibrometer. Furthermore, all blades except the currently excited one are detuned with additional masses. Applying these masses allows to identify a blade dominated natural frequency for each blade and every mode of interest. Finally, these blade dominated frequencies are summarized to mode specific mistuning patterns. The key part of the contribution presents a model update approach which is focused on small geometric deviations between real engine parts and idealized simulation models. Within this update procedure the nodal coordinates of an initially tuned finite element blisk model were modified in order to match the geometry of the real part measured by blue light fringe projection. All essential pre- and post-processing steps of the mesh morphing procedure are described and illustrated. It could be proven that locally remaining geometric deviations between updated finite element model and the optical measurement results are below 5 μm. For the purpose of validation blade dominated natural frequencies of the updated finite element blisk model are calculated for each sector up to a frequency of 17 kHz. Finally, the numerically predicted mistuning patterns are compared against the experimentally identified counterparts. At this point a very good agreement between experimentally identified and numerically predicted mistuning patterns can be proven across several mode families. Even mistuning patterns of higher modes at about 17 kHz are well predicted by the geometrically mistuned finite element model. Within the last section of the paper, possible uncertainties of the presented model update procedure are analyzed. As a part of the study the digitization of the investigated blisk has been repeated for ten times. These measurement results serve as input for the model update procedure described before. In the context of this investigation ten independent geometrical mistuned simulation models are created and the corresponding mistuning patterns are calculated. Copyright © 2017 by Rolls-Royce Deutschland Ltd \& Co KG}, language = {en} } @misc{BackhausMaywaldSchrapeetal., author = {Backhaus, Thomas and Maywald, Thomas and Schrape, Sven and Voigt, Matthias and Mailach, Roland}, title = {A Parametrization Describing Blisk Airfoil Variations Referring to Modal Analysis}, series = {ASME Turbo Expo 2017, GT2017-64243, June 26-30, 2017, Charlotte, NC, USA, Volume 7A}, journal = {ASME Turbo Expo 2017, GT2017-64243, June 26-30, 2017, Charlotte, NC, USA, Volume 7A}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-5092-3}, doi = {10.1115/GT2017-64243}, abstract = {This paper will present a way to capture the geometric blade by blade variations of a milled from solid blisk as well as the manufacturing scatter. Within this idea it is an essential task to digitize the relevant airfoil surface as good as possible to create a valid surface mesh as the base of the upcoming evaluation tasks. Since those huge surface meshes are not easy to handle and are even worse in getting quantified and easy interpretable results, it should be aimed for an easily accessible way of presenting the geometric variation. The presented idea uses a section based airfoil parametrization that is based on an extended NACA-airfoil structure to ensure the capturing of all occurring characteristic geometry variations. This Paper will show how this adapted parametrization method is suitable to outline all the geometric blade by blade variation and even more, refer those airfoil design parameters to modal analysis results such as the natural frequencies of the main mode shapes. This way, the dependencies between the modal and airfoil parameters will be proven.}, language = {en} }