@misc{BeirowMaywaldFigaschewskyetal., author = {Beirow, Bernd and Maywald, Thomas and Figaschewsky, Felix and K{\"u}hhorn, Arnold and Heinrich, Christoph Rocky and Giersch, Thomas}, title = {Simplified Determination of Aerodynamic Damping for Bladed Rotors, Part 1: Experimental Validation at Rest}, series = {ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13-17, 2016}, journal = {ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13-17, 2016}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-4984-2}, doi = {10.1115/GT2016-56535}, abstract = {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.}, language = {en} } @inproceedings{GierschBeirowPopigetal., author = {Giersch, Thomas and Beirow, Bernd and Popig, Frederik and K{\"u}hhorn, Arnold}, title = {FSI-based forced response analyses of a mistuned high pressure compressor blisk}, series = {10th International Conference on Vibrations in Rotating Machinery, 11-13 September 2012, IMechE London, UK}, booktitle = {10th International Conference on Vibrations in Rotating Machinery, 11-13 September 2012, IMechE London, UK}, publisher = {Woodhead Publ.}, address = {Cambridge, UK}, isbn = {978-0-85709-452-0}, language = {en} } @inproceedings{GierschHoenischBeirowetal., author = {Giersch, Thomas and H{\"o}nisch, Peter and Beirow, Bernd and K{\"u}hhorn, Arnold}, title = {Forced Response Analyses of Mistuned Radial Inflow Turbines}, series = {Proceedings of the ASME Turbo Expo 2012 : presented at the 2012 ASME Turbo Expo, June 11 - 15, 2012, Copenhagen, Denmark, Vol. 7, part B}, booktitle = {Proceedings of the ASME Turbo Expo 2012 : presented at the 2012 ASME Turbo Expo, June 11 - 15, 2012, Copenhagen, Denmark, Vol. 7, part B}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-4473-1}, pages = {1559 -- 1570}, language = {en} } @misc{GierschHoenischBeirowetal., author = {Giersch, Thomas and H{\"o}nisch, Peter and Beirow, Bernd and K{\"u}hhorn, Arnold}, title = {Forced Response Analyses of Mistuned Radial Inflow Turbines}, series = {Journal of Turbomachinery}, volume = {135}, journal = {Journal of Turbomachinery}, number = {3}, issn = {1528-8900}, doi = {10.1115/1.4007512}, pages = {031034-1 -- 031034-9}, abstract = {Radial turbine wheels designed as blade integrated disks (blisk) are widely used in various industrial applications. However, related to the introduction of exhaust gas turbochargers in the field of small and medium sized engines, a sustainable demand for radial turbine wheels has come along. Despite those blisks being state of the art, a number of fundamental problems, mainly referring to fluid-structure-interaction and, therefore, to the vibration behavior, have been reported. Aiming to achieve an enhanced understanding of fluid-structure-interaction in radial turbine wheels, a numerical method, able to predict forced responses of mistuned blisks due to aerodynamic excitation, is presented. In a first step, the unsteady aerodynamic forcing is determined by modeling the spiral casing, the stator vanes, and the rotor blades of the entire turbine stage. In a second step, the aerodynamic damping induced by blade vibration is computed using a harmonic balance technique. The structure itself is represented by a reduced order model being extended by aerodynamic damping effects and aerodynamic forcings. Mistuning is introduced by adjusting the modal stiffness matrix based on results of blade by blade measurements that have been performed at rest. In order to verify the numerical method, the results are compared with strain-gauge data obtained during rig-tests. As a result, a measured low engine order excitation was found by modeling the spiral casing. Furthermore, a localization phenomenon due to frequency mistuning could be proven. The predicted amplitudes are close to the measured data.}, language = {en} } @misc{BeirowGierschKuehhornetal., author = {Beirow, Bernd and Giersch, Thomas and K{\"u}hhorn, Arnold and Nipkau, Jens}, title = {Forced Response Analysis of a Mistuned Compressor Blisk}, series = {Journal of Engineering for Gas Turbines and Power}, volume = {136}, journal = {Journal of Engineering for Gas Turbines and Power}, number = {6}, issn = {1528-8919}, doi = {10.1115/1.4026537}, pages = {13}, abstract = {The forced response of an E3E-type high pressure compressor (HPC) blisk front rotor is analyzed with regard to varying mistuning and the consideration of the fluid-structure interaction (FSI). For that purpose, a reduced order model is used in which the disk remains unchanged and mechanical properties of the blades, namely stiffness and damping, are adjusted to measured as well as intentional blade frequency mistuning distributions. The aerodynamic influence coefficient technique is employed to model the aeroelastics. Depending on the blade mode, the exciting engine order, and aerodynamic influences, it is sought for the worst mistuning distributions with respect to the maximum blade displacement based on optimization analyses. Genetic algorithms using blade-alone frequencies as design variables are applied. The validity of the Whitehead limit is assessed in this context. In particular, the question is addressed if and how far aeroelastic effects, mainly caused by aerodynamic damping, combined with mistuning can even cause a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the interblade phase angle is the main driver for a possible response attenuation considering the fundamental as well as a higher blade mode. Furthermore, the differences to the blisk vibration response without a consideration of the flow and an increase of the disk's stiffness are discussed. Closing, the influence of pure damping mistuning is analyzed again using optimization.}, language = {en} } @misc{BeirowGierschKuehhornetal., author = {Beirow, Bernd and Giersch, Thomas and K{\"u}hhorn, Arnold and Nipkau, Jens}, title = {Optimization-Aided Forced Response Analysis of a Mistuned Compressor Blisk}, series = {Journal of Engineering for Gas Turbines and Power}, volume = {137}, journal = {Journal of Engineering for Gas Turbines and Power}, number = {1}, issn = {1528-8919}, doi = {10.1115/1.4028095}, pages = {012504-1 -- 012504-10}, abstract = {The forced response of the first rotor of an engine 3E (technology program) (E3E)-type high pressure compressor (HPC) blisk is analyzed with regard to varying mistuning, varying engine order (EO) excitations and the consideration of aero-elastic effects. For that purpose, subset of nominal system modes (SNM)-based reduced order models are used in which the disk remains unchanged while the Young's modulus of each blade is used to define experimentally adjusted as well as intentional mistuning patterns. The aerodynamic influence coefficient (AIC) technique is employed to model aero-elastic interactions. Furthermore, based on optimization analyses and depending on the exciting EO and aerodynamic influences it is searched for the worst as well as the best mistuning distributions with respect to the maximum blade displacement. Genetic algorithms using blade stiffness variations as vector of design variables and the maximum blade displacement as objective function are applied. An allowed limit of the blades' Young's modulus standard deviation is formulated as secondary condition. In particular, the question is addressed if and how far the aero-elastic impact, mainly causing aerodynamic damping, combined with mistuning can even yield a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the interblade phase angle is the main driver for a possible response attenuation considering the fundamental blade mode. The results of the optimization analyses are compared to the forced response due to real, experimentally determined frequency mistuning as well as intentional mistuning.}, language = {en} } @inproceedings{StelldingerGierschFigaschewskyetal., author = {Stelldinger, Marco and Giersch, Thomas and Figaschewsky, Felix and K{\"u}hhorn, Arnold}, title = {A Semi-Unstructured Turbomachinery Meshing Library With Focus on Modeling of Specific Geometrical Features}, 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 = {Computational Fluid Dynamics is widely used for the analysis and the design of turbomachinery blade rows. A well established method is the application of semi-unstructured meshes, that uses a combination of structured meshes in the radial direction and unstructured meshes in the axial as well as the tangential direction. This takes advantage of the approximately two dimensional flow field through the blade rows, whereby a fine radial discretization, excepting the near wall region, is not necessary. Otherwise, it is possible to discretize particular regions, e.g. the leading and trailing edge regions, in the axial and tangential direction without generating unnecessary nodes in the far field. The meshing approach is based on the projection of a two dimensional unstructured mesh defined at a reference surface. Once, the two dimensional mesh is generated the projection is achieved by transfinite interpolation from the reference surface to further radial surfaces using a structured mesh. Due to the modeling of geometrical features, especially fillets, advanced methods for the generation of structured meshes and mesh smoothing algorithms are required. The paper presents two different approaches for the generation of an appropriate structured mesh. The first is based on the solution of elliptic partial differential equations. The second approach is based on the split of the domain into fourteen appropriately arranged blocks. Furthermore, two smoothing methods for two dimensional unstructured meshes, a constrained Laplace smoothing and an optimization based approach, are presented. Regarding a more realistic representation of the geometry, methods for the modeling of cavities, variable clearance sizes and fillets are presented. Finally, a comparison of the smoothing techniques applied to a rotor passage is presented and the influence of chosen geometrical features on the flow solution is evaluated.}, language = {en} } @phdthesis{Giersch, author = {Giersch, Thomas}, title = {Numerical Models for the Vibration Response of High Pressure Compressor Rotors with Validation for Forced Response and Surge}, publisher = {Mensch \& Buch}, address = {Berlin}, isbn = {978-3-86387-930-3}, pages = {XX, 136}, abstract = {The following dissertation contributes to the numerical modelling of the vibration response of integral real high pressure compressor rotor blades. It aims to validate the chosen modelling techniques with available measurement data and to demonstrate its capability for industrial applications. In detail the present dissertation investigates two different scenarios of blade excitation. Within the first scenario the predicted blade response of an integral bladed rotor is compared to measurements with focus on the blade mistuning problem. The second scenario deals with vibration response of compressor rotor blades due to surge of a transonic high pressure compressor. For the modelling purpose of the aeroelastic interaction a loose coupling technique is selected with a separate structural and flow model. The chosen structural model is based on the modal reduction technique published by Yang and Griffin, called subset of nominal system modes. The high accuracy of the model when reducing finite element models is discussed. It is shown, that the reduction algorithm is limited to mistuned modeshapes that can be expressed by a superposition of the tuned modeshapes of the rotor. The 3D Finite Volume Code AU3D, developed at Imperial College London, is used to model the steady and unsteady flow field. The flow solver is utilised to derive the external and motion induced aerodynamic forces for both investigated scenarios. To reduce the numerical effort an additional 1D flow solver is developed that allows the surge frequency and impulse loads for the compressor to be computed. The presented comparisons between measured and predicted vibration responses for the integral resonance passing are in a high agreement. The remaining deviations are within the measurement accuracy. In addition, it is demonstrated that the applied model can also be used for model identification purposes from measurement data. It is found that the investigated surge can qualitatively be well explained by the impulse loads that are generated due to the fast change of the aerodynamic loads. The predicted vibration levels are quantitatively and qualitatively in good agreement to the measured data. It is shown that further analysis is required to understand the considerable scatter in vibration response of successive surge cycles.}, 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} } @inproceedings{FigaschewskyKuehhornGiersch, author = {Figaschewsky, Felix and K{\"u}hhorn, Arnold and Giersch, Thomas}, title = {A Finite Element Based Least Square Fit for the Assessment of Integral and Non-Integral Vibrations With Blade Tip Timing}, series = {Proceedings of ISROMAC 2017, Maui, Hawaii, December 16-21, 2017}, booktitle = {Proceedings of ISROMAC 2017, Maui, Hawaii, December 16-21, 2017}, pages = {9}, abstract = {This paper aims at improving the robustness and accuracy of the least square fit technique utilized in blade tip timing (BTT) measurements of blade vibrations by proposing two modifications. The first proposal is to replace the lines of the original least square problem by differences of consecutive lines. Thereby, the static deflection as well as the circumferential blade positioning error cancels out and the robustness is improved by removing these uncertainties inherently. The second proposal is to replace the fit of piecewise constant vibration amplitudes within the chosen block length by a linear or cubic spline in the frequency (integral) or time (non-integral) regime. This does not only suppress overshoots due to distorted acceleration or deceleration manoeuvres but also allows for a "coarser analysis grid" (i.e. larger block length) without loosing amplitude accuracy. Thereby it smooths out random errors more efficiently and increases the orthogonality of the relevant EO or vibration frequency to unwanted signal components.}, 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{FigaschewskyGierschKuehhorn, author = {Figaschewsky, Felix and Giersch, Thomas and K{\"u}hhorn, Arnold}, title = {Forced Response Prediction of an Axial Turbine Rotor With Regard to Aerodynamically Mistuned Excitation}, series = {ASME Turbo Expo 2014: Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics, D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-25896}, booktitle = {ASME Turbo Expo 2014: Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics, D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-25896}, publisher = {ASME}, isbn = {978-0-7918-4577-6}, doi = {10.1115/GT2014-25896}, pages = {13}, abstract = {The design of both efficient and reliable turbomachinery blades demands a detailed knowledge of static and dynamic forces during operation. This paper aims to contribute to the proper identification of dynamic excitation mechanisms acting on an axial turbine rotor, particularly with regard to deviations of the NGV's nominal geometry due to the use of variable vanes or tolerances in manufacturing. As variations of the NGV's geometry disturb the perfectly periodic pattern of the downstream flow features, other spectral components than those correlated with the number of stator vanes are possible to appear. These frequency components may lead to low engine order excitation of fundamental blade modes at high engine speeds. Under these operating conditions the rotor is already highly loaded with centrifugal forces and additional dynamic excitation may cause unacceptable stresses. Thus aerodynamic mistuning might be a limiting criterion for the design of a highly loaded turbine rotor. Within this paper 2 dimensional CFD-models are used to investigate both, the determination of the wake of a geometric mistuned stator guide vane and the influence of the resulting excitation on the adjacent rotor stage due to aerodynamically mistuned flow. In order to generate a mistuned NGV geometry, variations of pitch and stagger angle are taken into account and a mesh morpher is used to produce computational domains of the mistuned geometry on the basis of a nominal mesh. Additionally a simplified reconstruction process based on a set of CFD computations will be introduced, being able to reproduce the spectral components of the mistuned wake by specifying a certain geometric mistuning distribution. The prediction of the resulting modal forces is carried out in time domain and approaches with lower fidelity are investigated with respect to their capability of reproducing the key features of an aerodynamically mistuned excitation mechanism.}, language = {en} } @inproceedings{BeirowKuehhornGierschetal., author = {Beirow, Bernd and K{\"u}hhorn, Arnold and Giersch, Thomas and Nipkau, Jens}, title = {Optimization-Aided Forced Response Analysis of a Mistuned Compressor Blisk}, series = {ASME Turbo Expo 2014, Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics, D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-25915}, booktitle = {ASME Turbo Expo 2014, Turbine Technical Conference and Exposition, Volume 7B: Structures and Dynamics, D{\"u}sseldorf, Germany, June 16-20, 2014, Paper GT2014-25915}, publisher = {ASME}, isbn = {978-0-7918-4577-6}, doi = {10.1115/GT2014-25915}, pages = {11}, abstract = {The forced response of the first rotor of an E3E-type high pressure compressor blisk is analyzed with regard to varying mistuning, varying engine order excitations and the consideration of aeroelastic effects. For that purpose, SNM-based reduced order models are used in which the disk remains unchanged while the Young's modulus of each blade is used to define experimentally adjusted as well as intentional mistuning patterns. The aerodynamic influence coefficient technique is employed to model aeroelastic interactions. Furthermore, based on optimization analyses and depending on the exciting EO and aerodynamic influences it is searched for the worst as well as the best mistuning distributions with respect to the maximum blade displacement. Genetic algorithms using blade stiffness variations as vector of design variables and the maximum blade displacement as objective function are applied. An allowed limit of the blades' Young's modulus standard deviation is formulated as secondary condition. In particular, the question is addressed if and how far the aeroelastic impact, mainly causing aerodynamic damping, combined with mistuning can even yield a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the inter-blade phase angle is the main driver for a possible response attenuation considering the fundamental blade mode. The results of the optimization analyses are compared to the forced response due to real, experimentally determined frequency mistuning as well as intentional mistuning. Copyright © 2014 by Rolls-Royce Deutschland Ltd \& Co KG}, 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} } @inproceedings{GierschKuehhornFigaschewsky, author = {Giersch, Thomas and K{\"u}hhorn, Arnold and Figaschewsky, Felix}, title = {Probabilistic Analysis of Low Engine Order Excitation Due to Geometric Perturbations of Upstream Nozzle Guide Vanes : ISABE-2015-20165}, series = {Conference Proceedings from the 22nd International Symposium on Air Breathing Engines, October 25-30, 2015, Phoenix, Arizona}, booktitle = {Conference Proceedings from the 22nd International Symposium on Air Breathing Engines, October 25-30, 2015, Phoenix, Arizona}, publisher = {ISABE}, pages = {1 -- 9}, language = {en} } @inproceedings{BeirowKuehhornGierschetal., author = {Beirow, Bernd and K{\"u}hhorn, Arnold and Giersch, Thomas and Nipkau, Jens}, title = {Forced Response Analysis of a Mistuned Compressor Blisk}, series = {ASME Turbo Expo 2013: Turbine Technical Conference and Exposition, San Antonio, Texas, USA, June 3-7, 2013, Vol. 7B, Structures and Dynamics, Paper GT2013-94142}, booktitle = {ASME Turbo Expo 2013: Turbine Technical Conference and Exposition, San Antonio, Texas, USA, June 3-7, 2013, Vol. 7B, Structures and Dynamics, Paper GT2013-94142}, publisher = {ASME}, address = {New York}, isbn = {978-0-7918-5527-0}, doi = {10.1115/GT2013-94142}, pages = {10}, abstract = {The forced response of an E3E-type HPC-blisk front rotor is analyzed with regard to varying mistuning and the consideration of the fluid-structure interaction (FSI). For that purpose, a reduced order model is used in which the disk remains unchanged and mechanical properties of the blades namely stiffness and damping are adjusted to measured as well as intentional blade frequency mistuning distributions. The aerodynamic influence coefficient technique is employed to model the aeroelastics. Depending on the blade mode, the exciting engine order and aerodynamic influences it is sought for the worst mistuning distributions with respect to the maximum blade displacement based on optimization analyses. Genetic algorithms using blade alone frequencies as design variables are applied. The validity of the Whitehead-limit is assessed in this context. In particular, the question is addressed if and how far aeroelastic effects, mainly caused by aerodynamic damping, combined with mistuning can even cause a reduction of the forced response compared to the ideally tuned blisk. It is shown that the strong dependence of the aerodynamic damping on the inter-blade phase angle is the main driver for a possible response attenuation considering the fundamental as well as a higher blade mode. Furthermore, the differences to the blisk vibration response without a consideration of the flow and an increase of the disk's stiffness are discussed. Closing, the influence of pure damping mistuning is analyzed again using optimization.}, language = {en} } @misc{YangBeirowGiersch, author = {Yang, Jingjie and Beirow, Bernd and Giersch, Thomas}, title = {Simulation and Investigation of an Intentionally Mistuned Blisk Rotor in a High Pressure Compressor}, series = {ASME 2022 Turbomachinery Technical Conference \& Exposition (GT2022)}, journal = {ASME 2022 Turbomachinery Technical Conference \& Exposition (GT2022)}, abstract = {In modern aircraft engines, blade integrated disk (blisk) is widely implemented. While blisk rotor design brings numerous advantages including weight reduction, aerodynamic efficiency improvement, and manufacturing simplification, its low mechanical damping due to the absence of friction between disk and blades makes the rotor more susceptible to vibration. Given that damage to blisk rotor sometimes requires the whole assembly to be replaced, effort has been made to alleviate the unexpected vibration amplitude within operating range, among which intentional mistuning is regarded as one of the commonly used technique. Mistuning refers to blade-to-blade deviation of mechanical properties, which is inevitable in practice due to manufacturing tolerances or wear. Through the application of intentional mistuning, it is expected that the amplitude of synchronous or nonsynchronous vibration (NSV) will be reduced without severely losing aerodynamic performance. In this paper, the effect of intentional mistuning has been investigated for the blisk rotor of a 1.5-stage transonic research compressor at Technical University of Darmstadt. According to the previous test campaign, the baseline rotor has shown its susceptibility to NSV due to first torsion mode in the near stall region. The rotor was then intentionally mistuned. Subsequent tests have proven a successful suppression of flutter problem. In order to have a comprehensive understanding of the effect of the applied mistuning pattern, simulations are performed using a FVM based CFD solver to produce comparable results as shown in the test campaign. In the simulation, mistuned systems are modelled in comparison with the nominal tuned reference. Geometrical disturbance and frequency disturbance are introduced to the tuned model first separately and then simultaneously. In this way, contribution of aerodynamic and structural mistuning to the suppression of NSV is identified based on the CFD results. Later, system eigenvalues of the mistuned aeromechanical model are determined by making use of the blade individual response in time domain. The obtained results are compared with mistuned eigenvalues calculated by a reduced order model (ROM), which utilizes the idea of subset of nominal modes (SNM). This makes it possible to demonstrate the feasibility of using SNM to carry out stability analysis when designing mistuning pattern for vibration of NSV type. It also allows a compare between the linear structural model of the SNM and the non-linear aeromechanic model of the CFD solver on capturing the non-linear nature of the flow, especially in the context of NSV.}, language = {en} } @inproceedings{SpringmannKuehhornRaueretal., author = {Springmann, Marcel and K{\"u}hhorn, Arnold and Rauer, Georg and Giersch, Thomas}, title = {Constitutive Modelling of Plastic and Creep Behavior of the Nickel Base Superalloy ALLVAC® 718PLUS® under Heat Treatment Conditions}, 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{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{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{FigaschewskyBeirowKuehhornetal., author = {Figaschewsky, Felix and Beirow, Bernd and K{\"u}hhorn, Arnold and Nipkau, Jens and Giersch, Thomas and Powers, Bronwyn}, title = {Design and Analysis of an Intentional Mistuning Experiment Reducing Flutter Susceptibility and Minimizing Forced Response of a Jet Engine Fan}, series = {ASME Turbo Expo 2017, GT2017-64621, June 26-30, 2017, Charlotte, NC, USA, Volume 7B}, journal = {ASME Turbo Expo 2017, GT2017-64621, June 26-30, 2017, Charlotte, NC, USA, Volume 7B}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-5093-0}, doi = {10.1115/GT2017-64621}, pages = {13}, abstract = {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}, language = {en} } @misc{KrauseStelldingerHanschkeetal., author = {Krause, Christoph and Stelldinger, Marco and Hanschke, Benjamin and K{\"u}hhorn, Arnold and Giersch, Thomas}, title = {Asynchronous Response Analysis of Non-Contact Vibration Measurements on Compressor Rotor Blades}, series = {ASME Turbo Expo 2017, GT2017-63200, June 26-30, 2017, Charlotte, NC, USA, Volume 7B}, journal = {ASME Turbo Expo 2017, GT2017-63200, June 26-30, 2017, Charlotte, NC, USA, Volume 7B}, publisher = {ASME}, address = {New York, NY}, isbn = {978-0-7918-5093-0}, doi = {10.1115/GT2017-63200}, abstract = {Although the research in non-intrusive techniques for the measurement of vibration have made major progress since the beginning in the 1960's, they are still mainly used as additional tool to the common strain gauges. Therefore, there is still a great deal of interest in the improvement of such non-contact vibration measurement techniques, to replace the intrusive ones with alternative techniques. One possibility to monitor all blades at once is blade tip-timing. The probes for a blade tip-timing measurement system are mounted circumferentially in the engine casing to log the passing times of the rotor blades. These logged time data will be compared with theoretically calculated passing times. The deviation between measured and calculated passing times can be transformed to blade displacement values. In recent years, several methods to analyse the acquired vibration data have been developed and improved. They are directed to evaluate synchronous and asynchronous blade vibration events. This paper focuses on the identification of asynchronous vibrations on rotor blades using blade tip-timing. Taking the data from all probes into account gives an opportunity to determine the vibration of each single blade. Due to the usage of a research test rig, all measurement data could be acquired in simulated real case operation scenarios. Analysis data were evaluated with a developed post processing routine based on a Fourier transformation algorithm coupled with a least square fitting procedure. Since compressor surge represents one of the most critical non synchronous events during compressor operation, in this paper a special interest is paid to the analysis of compressor surges. Vibration frequencies revealed during surge investigation will be compared with simultaneously measured strain gauge data to ensure the reliability of blade tip-timing measurement and analysis. To explain the results in more detail, the possibility of a blade damaged triggered shift of the blade characteristic frequency is shown. The most promising result of the analysis is the close correlation between the identified vibration frequencies of compressor surge events and the afterwards determined frequency mistuning and crack distributions. Blade damage becomes visible through increasing deviation between characteristic frequencies of different blades as result of multiple surge events. In addition, with the comparison of mean frequency records over each single surge among each other it is possible to restrict the blade damage time. Subsequently, the possibility to develop a process routine to predict blade damage during compressor test series could arise.}, language = {en} } @misc{FigaschewskyKuehhornBeirowetal., author = {Figaschewsky, Felix and K{\"u}hhorn, Arnold and Beirow, Bernd and Giersch, Thomas and Nipkau, Jens and Meinl, Ferdinand}, title = {Simplified Estimation of Aerodynamic Damping for Bladed Rotors, Part 2: Experimental Validation During operation}, series = {ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13-17, 2016}, journal = {ASME Turbo Expo 2016, Turbomachinery Technical Conference and Exposition, Volume 7B, Structures and Dynamics, Seoul, South Korea, June 13-17, 2016}, publisher = {ASME}, address = {New York, NY [u.a.]}, isbn = {978-0-7918-4984-2}, doi = {10.1115/GT2016-56458}, abstract = {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}, language = {en} }