@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} } @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{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{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{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} } @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{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{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} }