36 results found
Zoepke-Sonntag R, Hill G, Stapelfeldt S, 2023, Analysis of Flutter Mechanisms and Unsteady Aerodynamics of a Transonic Fan Blade Near Stall, Turbomachinery Technical Conference & Exposition
This paper investigates the onset of aeroelastic instability of a modern low-pressure ratio transonic fan using numerical simu- lations, validated against experimental measurements. Different domain configurations are compared in terms of fan characteris- tic and radial profiles. It is shown that flutter can be predicted over the investigated speed range using a reduced computational domain. The benefit in computational speed justifies simplifica- tions made and allows a more detailed unsteady analysis. During the unsteady analysis, it is shown that the transition from stable to aeroelastic unstable along a constant speed characteristic results from increased negative aerodynamic damping on the pressure side of the blade, accompanied by a drop of damping on the suc- tion side near the tip. The shock and the radial migration have a stabilizing effect on the blade, while the region downstream of the shock tends to destabilize the blade. The drop in stability occurs in multiple low nodal diameters, independent of acoustic propagation conditions. While previous work tried to identify unique mechanisms that drive flutter, this research concludes that the change from stable to unstable is subtle and that it is a combi- nation of many drivers. In addition, the tip region, especially the region downstream of the shock, tends to have a more significant impact than assumed, while the shock shows only small variations near the tip.
Merson J, Stapelfeldt S, 2023, Aerodynamic Forcing Analysis of Aerodynamically Mistuned Outlet Guide Vane Assemblies, Turbo Expo 2023: Turbomachinery Technical Conference and Exposition
Brandstetter C, Stapelfeldt S, 2022, Near-Stall Modelling of a Pitching Airfoil at High Incidence, Mach Number and Reduced Frequency, INTERNATIONAL JOURNAL OF TURBOMACHINERY PROPULSION AND POWER, Vol: 7
Rao A, Sureshkumar P, Stapelfeldt S, et al., 2022, UNSTEADY ANALYSIS OF AEROENGINE INTAKE DISTORTION MECHANISMS: VORTEX DYNAMICS IN CROSSWIND CONDITIONS, Turbo Expo 2022
Rao A, Sureshkumar P, Stapelfeldt S, et al., 2022, Unsteady analysis of aeroengine intake distortion mechanisms: vortex dynamics in crosswind conditions, Journal of Engineering for Gas Turbines and Power: Transactions of the ASME, Vol: 144, Pages: 1-11, ISSN: 0742-4795
The formation of a ground vortex and its ingestion into an aero-engine intake under crosswind conditions play a significant role in the aerodynamic excitation of the fan. Using steady and unsteady numerical simulations, an analysis of the dynamics of several distortion features is presented. For a simplified intake at high crosswind velocities, there is a substantial movement of the ingested ground vortex at the aerodynamic interface plane (AIP). The ingested ground vortex follows a specific trajectory while varying both in size and strength. The transition from a periodic to an aperiodic regime of the intake distortion at the AIP occurs as the crosswind velocity is increased. Circumferential mode decomposition shows that the largest amplitude of the distortion occurs at the first circumferential mode, and the amplitudes of the higher modes decrease monotonically. Furthermore, the amplitude of the spatial harmonics is time-dependent, which may be an influential feature at the time of assessing fan forced response.
Trafford A, Stapelfeldt S, Puente Rico R, 2022, A computational study of burner failure related low engine order forced response mechanisms in a high pressure turbine, International Symposium of Unsteady Aerodynamics, Aeroelastics and Aeroacoustics of Turbomachines (ISUAAAT)
Harris J, Lad B, Stapelfeldt S, 2022, Two-Dimensional Investigation of the Fundamentals of OGV Buffeting, INTERNATIONAL JOURNAL OF TURBOMACHINERY PROPULSION AND POWER, Vol: 7
Hill G, Gambel J, Schneider S, et al., 2022, Aeroelastic stability of combined plunge-pitch mode shapes in a linear compressor cascade, International Journal of Turbomachinery, Propulsion and Power, Vol: 7, Pages: 1-16, ISSN: 2504-186X
Modern aeroengine designs strive for peak specific fuel and thermal efficiency. To achieve these goals, engines have more highly loaded compressor stages, thinner aerofoils, and blended titanium integrated disks (blisks) to reduce weight. These configurations promote the occurrence of aeroelastic phenomena such as flutter. Two important parameters known to influence flutter stability are the reduced frequency and the ratio of plunge and pitch components in a combined flap mode shape. These are used as design criteria in the engine development process. However, the limit of these criteria is not fully understood. The following research aims to bridge the gap between semi-analytical models and modern compressors by systematically investigating the flutter stability of a linear compressor cascade. This paper introduces the plunge-to-pitch incidence ratio, which is defined as a function of reduced frequency and pitch axis setback for a first flap (1F) mode shape. Using numerical simulations, in addition to experimental validation, aerodynamic damping is computed for many modes to build stability maps. The results confirm the importance of these two parameters in compressor aeroelastic stability as well as demonstrate the significance of the plunge-to-pitch incidence ratio for predicting the flutter limit.
Unsteady aerodynamic phenomena appearing close to the stall boundary in axial fans and compressors have been the subject of extensive investigations. A particular phenomenon known as “rotating instabilities” typically occurs in the tip region at highly loaded conditions and is often linked to blade vibrations. During rig and engine testing, it is usually identified by the characteristic shape of its pressure spectra. In the 1990s, this shape has been explained as the result of a circumferentially propagating disturbance, which is unsteady in a frame of reference rotating with the disturbance. In this analysis, conclusions regarding its propagation speed and frequency have been made based on the analysis of spectral peaks, and this method is still often used to classify unsteady aerodynamic phenomena. However, in high subsonic and transonic machines, where aeroacoustic and aeroelastic phenomena interact with aerodynamic disturbances, the interpretation of measurements using spectra alone is challenging. The present article aims to demonstrate the difficulties and subtleties associated with the analysis of measurement signals, which need to be overcome to correctly interpret stall precursor signatures. At the example of a recent investigation on a composite fan, the consequences of sensor placement and postprocessing techniques are discussed with a focus on spectral averaging, isolation of non-synchronous phenomena, and multisensor cross correlation methods. It is seen that the interpretation of phenomena based solely on spectral peaks and their spacing can be misleading and that the characteristic features of a rotating instability spectrum do not require an unstable pulsating disturbance.
Stapelfeldt S, Brandstetter C, 2021, Suppression of nonsynchronous vibration through intentional aerodynamic and structural mistuning, Journal of Turbomachinery, Vol: 144, Pages: 1-8, ISSN: 0889-504X
Nonsynchronous vibrations (NSVs) arising near the stall boundary of compressors are a recurring and potentially safety-critical problem in modern axial compressors and fans. Recent research has improved predictive capabilities and physical understanding of NSV, but prevention measures are still lacking. This article addresses this by systematically studying the influence of aerodynamic and structural mistuning on NSV. This is achieved by incorporating mistuning effects in a validated linear model, in which individual blade modes are modeled as single-degree-of-freedom mass oscillators coupled by a convected aerodynamic disturbance term. The results demonstrate that both structural and aerodynamic mistuning are effective. While structural mistuning improves stability by preventing aero-structure lock-in, aerodynamic mistuning, which locally reduces the tip blockage and attenuates the aerodynamic disturbance causing NSV. In the latter case, the circumferentially averaged conditions are shown to be most influential, while the pattern plays a minor role. A combination of moderate aerodynamic and structural mistuning (1%) was also found to be effective. These findings are relevant for design decisions, demonstrating that small blade-to-blade variations can suppress NSV.
Brandstetter C, Stapelfeldt S, 2021, Analysis of a linear model for non-synchronous vibrations near stall, International Journal of Turbomachinery, Propulsion and Power, Vol: 6, Pages: 1-14, ISSN: 2504-186X
Non-synchronous vibrations arising near the stall boundary of compressors are a recurring and potentially safety-critical problem in modern aero-engines. Recent numerical and experimental investigations have shown that these vibrations are caused by the lock-in of circumferentially convected aerodynamic disturbances and structural vibration modes, and that it is possible to predict unstable vibration modes using coupled linear models. This paper aims to further investigate non-synchronous vibrations by casting a reduced model for NSV in the frequency domain and analysing stability for a range of parameters. It is shown how, and why, under certain conditions linear models are able to capture a phenomenon, which has traditionally been associated with aerodynamic non-linearities. The formulation clearly highlights the differences between convective non-synchronous vibrations and flutter and identifies the modifications necessary to make quantitative predictions.
Xu Z, Puente R, Stapelfeldt S, 2021, Validation of three-dimensional grid refinement for Lattice Boltzmann methods, AIAA Scitech 2021 Forum, Publisher: American Institute of Aeronautics and Astronautics, Pages: 1-15
The lattice Boltzmann method (LBM) is a method for modeling fluid dynamics which is based on discrete kinetic theory. It discretizes the continuous Boltzmann equation in velocity, spatial and temporal space and solves a transport equation for particle populations. To enable an efficient time-space discretization with discrete velocity sets, common LBM solvers employ uniform grids making them prohibitively expensive for the simulation of high Reynolds number flows. This paper describes the implementation of a vertex-centered grid refinement method with the BGK operator in the Open-source Lattice Boltzmann code, OpenLB. The method is validated for flow around a 3D circular cylinder with Reynolds number varying from 100 to 300 and the 3D lid-driven cavity with Reynolds number equals 3200. The results are in good agreement with experimental measurements and the reference solutions with uniform grid spacing. The simulation with grid refinement allows for remarkable computational cost savings for a given accuracy.
Mitchell SJ, Stapelfeldt S, Puente R, 2021, Evaluation of the single-population lattice Boltzmann method for one-dimensional compressible flows, AIAA Scitech 2021 Forum, Publisher: American Institute of Aeronautics and Astronautics, Pages: 1-16
In this work the stability and accuracy of the one-dimensional single-population Lattice Boltzmann Method (LBM) is studied in its application to compressible flow regimes. A linear stability analysis framework using the von Neumann method is developed, framed in terms of flow regime parameters. In order to validate this analysis and study the behavior of errors with respect to different simulation parameters, two academic test cases are presented. A shock tube configuration is used to assess the shock capturing capabilities. A high subsonic convergent-divergent nozzle is used to evaluate the error dependence on grid size. This test case requires the inclusion of an additional source term to the Boltzmann equation, and a methodology is proposed for the implementation of physically motivated source terms. The proposed numerical method is shown to be accurate, matching theoretical error trends. Through stability analysis, bounds for the simulation parameters resulting in a stable simulation are given, which can inform best practices in the application of LBM to these regimes.
Stapelfeldt S, Brandstetter C, 2020, Non-synchronous vibration in axial compressors: lock-in mechanism and semi-analytical model, Journal of Sound and Vibration, Vol: 488, Pages: 1-20, ISSN: 0022-460X
Non-Synchronous-Vibration (NSV) in high-speed turbomachinery compressors is an aeroelastic phenomenon which can have devastating consequences, including loss of rotor blades. Despite extensive research over the past two decades its underlying mechanisms are not yet understood. This paper aims to explain the physical mechanisms causing NSV in a modern transonic compressor rotor. Referring to previous experimental results and using validated computational fluid dynamics (CFD), a parametric study is performed in order to characterize the aerodynamic disturbance causing NSV, and to understand the lock-in mechanism between the fluid and the structure seen during NSV. The results show that the process is driven by aerodynamics in the tip region. Under highly throttled conditions, the tip leakage flow blocks the passage and causes the disturbance, which is characterised as a vorticity fluctuation, to propagate circumferentially in the leading edge plane. It is found that the propagation speed of the disturbance is determined by the mean flow conditions and only its phase is periodically modulated through interaction with oscillating blades. This is the mechanism facilitating lock-in. Based on these findings a semi-analytic model is developed and calibrated with the numerical results. The model is capable of simulating the lock-in process and correctly predicts unstable vibration modes. It can therefore be used to identify critical operating conditions and develop mitigation measures early in the design.
Moreno J, Dodds J, Stapelfeldt SC, et al., 2020, Deficiencies in the SA Turbulence model for the prediction of the stability boundary in highly loaded compressors, Journal of Turbomachinery, Vol: 142, Pages: 121012-1-121012-11, ISSN: 0889-504X
Reynolds-averaged Navier–Stokes (RANS) equations are employed for aerodynamic and aeroelastic modeling in axial compressors. Their solutions are highly dependent on the turbulence models for closure. The main objective of this work is to assess the widely used Spalart–Allmaras model suitability for high-speed compressor flows. For this purpose, an extensive investigation of the sources of uncertainties in a high-speed multi-stage compressor rig was carried out. The grid resolution near the casing end wall, which affects the tip leakage flow and casing boundary layer, was found to have a major effect on the stability limit prediction. Refinements in this region led to a stall margin loss prediction. It was found that this loss was exclusively due to the destruction term in the SA model.
Zhang W, Stapelfeldt S, Vahdati M, 2020, Influence of the inlet distortion on fan stall margin at different rotational speeds, AEROSPACE SCIENCE AND TECHNOLOGY, Vol: 98, ISSN: 1270-9638
Lu Y, Green J, Stapelfeldt SC, et al., 2019, Effect of Geometric Variability on Running Shape and Performance of a Transonic Fan, JOURNAL OF TURBOMACHINERY-TRANSACTIONS OF THE ASME, Vol: 141, ISSN: 0889-504X
Lu Y, Lad B, Green J, et al., 2019, Effect of Geometry Variability on Transonic Fan Blade Untwist, Publisher: MDPI
Lu Y, Lad B, Green J, et al., 2019, Effect of geometry variability on transonic fan blade untwist, International Journal of Turbomachinery, Propulsion and Power, Vol: 4
Due to manufacturing tolerance and deterioration during operation, fan blades in the same engine exhibit geometric variability. The absence of symmetry will inevitably exacerbate and contribute to the complexities of running geometry prediction as the blade variability is bound to be amplified by aerodynamic and centrifugal loading. In this study, we aim to address the fan blade untwist related phenomenon known as alternate passage divergence (APD). As the name suggests, APD manifests as alternating passage geometry (and hence alternating tip stagger pattern) when the fan stage is operating close to/at peak efficiency condition. APD can introduce adverse influence on fan performance, aeroacoustics behaviour, and high cycle fatigue characteristics of the blade. The main objective of the study is to identify the parameters contributing to the APD phenomenon. In this study, the APD behaviours of two transonic fan blade designs are compared.
Stapelfeldt S, Vahdati M, 2019, Improving the Flutter Margin of an Unstable Fan Blade, JOURNAL OF TURBOMACHINERY-TRANSACTIONS OF THE ASME, Vol: 141, ISSN: 0889-504X
Lu Y, Lad B, Vahdati M, et al., 2019, Nonsynchronous vibration associated with transonic fan blade untwist
Due to manufacturing tolerance and deterioration during operation, fan blades in the aero-engine exhibit geometric variability. This leads to asymmetry in the assembly which will be amplified in the running geometry by centrifugal and aerodynamic loads. This study investigates a phenomenon known as Alternative Passage Divergence (APD), where the blade untwist creates an alternating pattern in passage geometry and stagger angle around the circumference, resulting in two groups of blades. This phenomenon occurs close to, or at, peak efficiency conditions and can significantly reduce overall efficiency. This study focuses on a type of non-integral vibration which occurs during APD. After the formation of alternating tip stagger pattern, APDs unsteady effect can cause the blades from one group to switch to the other, creating a travelling wave pattern around the circumference.It was found from numerical assessment on a randomly mis-staggered assembly that real engines can potentially experience such travelling disturbance and suffer fatigue damage. An idealised case is used to capture the bulk behaviour from the more complex cases in real engines and to decipher the underlying mechanism of this travelling disturbance. The results indicate that the driving force originates from the interaction between passage shock displacement and the passage geometry.
Moreno J, Dodds J, Vahdati M, et al., 2019, DEFICIENCIES IN TURBULENCE MODELLING FOR THE PREDICTION OF THE STABILITY BOUNDARY IN HIGHLY LOADED COMPRESSORS, ASME Turbo Expo: Turbomachinery Technical Conference and Exposition, Publisher: AMER SOC MECHANICAL ENGINEERS
Lu Y, Vahdati M, Green J, et al., 2018, Effect of Geometry Variability on Fan Performance and Aeromechanical Characteristics, 15th International Symposium on Unsteady Aerodynamics, Aeroacoustics and Aeroelasticity of Turbomachines
Stapelfeldt SC, Vahdati M, 2018, On the importance of engine-representative models for fan flutter predictions, Journal of Turbomachinery, Vol: 140, Pages: 081005-1-081005-10, ISSN: 0889-504X
Discrepancies between rig tests and numerical predictions of the flutter boundary for fan blades are usually attributed to the deficiency of computational fluid dynamics (CFD) models for resolving flow at off-design conditions. However, as will be demonstrated in this paper, there are a number of other factors, which can influence the flutter stability of fan blades and lead to differences between measurements and numerical predictions. This research was initiated as a result of inconsistencies between the flutter predictions of two rig fan blades. The numerical results agreed well with rig test data in terms of flutter speed and nodal diameter (ND) for both fans. However, they predicted a significantly higher flutter margin for one of the fans, while measured flutter margins were similar for both blades. A new set of flutter computations including the whole low-pressure system was therefore performed. The new set of computations considered the effects of the acoustic liner and mistuning for both blades. The results of this work indicate that the previous discrepancies between CFD and tests were caused by, first, differences in the effectiveness of the acoustic liner in attenuating the pressure wave created by the blade vibration and second, differences in the level of unintentional mistuning of the two fan blades. In the second part of this research, the effects of blade mis-staggering and inlet temperature on aerodynamic damping were investigated. The data presented in this paper clearly show that manufacturing and environmental uncertainties can play an important role in the flutter stability of a fan blade. They demonstrate that aeroelastic similarity is not necessarily achieved if only aerodynamic properties and the traditional aeroelastic parameters, reduced frequency and mass ratio, are maintained. This emphasizes the importance of engine-representative models, in addition to accurate and validated CFD codes, for the reliable prediction of the flutter boundar
Stapelfeldt SC, Vahdati M, 2018, Improving the Flutter Margin of an Unstable Fan Blade, ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition
Franz D, Salles L, Stapelfeldt S, 2017, Analysis of a Turbine Bladed Disk With Structural and Aerodynamic Mistuning, ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition
Stapelfeldt S, Vahdati M, 2017, On the importance of engine-representative models for fan flutter predictions
This paper examines the factors which can result in discrepancies between rig tests and numerical predictions of the flutter boundary for fan blades. Differences are usually attributed to the deficiency of CFD models for resolving the flow at off-design conditions. This work was initiated as a result of inconsistencies between the flutter prediction of two rig fan blades, called here Fan F1 and Fan F2. The numerical results agreed well with the test data in terms of flutter speed and nodal diameter for both fans. However, they predicted a significantly higher flutter margin for F2 than for Fan F1, while rig tests showed that the two blades had similar flutter margins. A new set of flutter computations for both blades using the whole LP domain (intake, fan, OGV and ESS) was therefore performed. The new set of computations considered the effects of the acoustic liner and mistuning for both blades. The results of this work indicate that the previous discrepancies between CFD and tests were due to: 1. Differences in the effectiveness of the acoustic liner in attenuating the pressure wave created by the blade vibration as a result of differences in flutter frequencies between the two fan blades. 2. Differences in the level of unintentional mistuning of the two fan blades due to manufacturing tolerances. In the second part of this research, the effects of blade misstaggering and inlet temperature on aerodynamic damping were investigated. The data presented in this paper clearly show that manufacturing and environmental uncertainties can play an important role in the flutter stability of a fan blade. They demonstrate that aeroelastic similarity is not necessarily achieved if only aerodynamic properties and the traditional aeroelastic parameters, reduced frequency and mass ratio, are maintained. This emphasises the importance of engine-representative models, in addition to an accurate and validated CFD code, for the reliable prediction of the flutter boundary.
Brandstetter C, Holzinger F, Schiffer H, et al., 2016, Near Stall Behavior of a Transonic Compressor Rotor With Casing Treatment, ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition
Stapelfeldt SC, Parry AB, Vahdati M, 2016, Investigation of flutter mechanisms of a contra-rotating open rotor, Journal of Turbomachinery, Vol: 138, Pages: 1-10, ISSN: 0889-504X
The growing pressure to reduce fuel consumption and cut emissions has triggered renewed interest in contra-rotating open rotor (CROR) technologies. One of their potential issues is self-excited or forced vibration of the unducted, light-weight, highly swept blades. This paper presents a numerical study into the flutter behavior of a CROR rig at take-off conditions. The study presented in this paper aimed to validate the numerical approach and provide insights into the flutter mechanisms of the open rotor under investigation. For the initial validation, pressure profiles and thrust coefficients from steady-state mixing plane calculations were compared against rig measurements. A full domain unsteady analysis predicted front rotor instability at low advance ratios. Flutter occurred in the first torsional mode in 0 and 1 nodal diameter (ND) which agreed with experimental observations. Subsequent unsteady computations focused on the isolated front rotor and first torsional mode. The flow field and aerodynamic damping over a range of advance ratios were studied. It was found that minimum aerodynamic damping occurred at low advance ratios when the flow was highly three-dimensional on the suction side. A correlation between the quasi-steady loading on the blade and aeroelastic stability was made and related to the numerical results. The effects of variations in frequency were then investigated by linking local aerodynamic damping to the unsteady pressure on the blade surface.
Stapelfeldt SC, Di Mare L, 2015, Reduced passage method for multirow forced response computations, AIAA Journal: devoted to aerospace research and development, Vol: 53, Pages: 3049-3062, ISSN: 0001-1452
This paper presents a time-domain Fourier method for modeling steady and unsteady nonaxisymmetric flows in turbomachinery on a reduced computational domain. The method extends well-established single-passage multirow methods, which efficiently model periodic unsteadiness in single stages, to assemblies with stationary circumferential perturbations with periodicity different from the blade count. Such perturbations are caused, for example, by rotor–rotor/stator–stator interaction or geometric circumferential variations. The method is therefore suitable to study low-engine-order forcing problems, flow past nonuniform assemblies, and clocking problems. The proposed method solves the flow inside several discrete passages, located at different circumferential positions, using a time-accurate scheme. Boundary conditions at the azimuthal and interrow surfaces are approximated via time–space Fourier series and couple the individual passages. The reduced passage model is validated against the whole annulus solution for three test cases: 1) a row of outlet guide vanes with stagger pattern, 2) a high-pressure turbine stage with throat area variation in the stator, and 3) a 1.5-stage compressor. It is demonstrated that the time-domain Fourier method yields results equivalent to the whole annulus model but at a much reduced computational cost.
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