Search

Corner separation is known to limit the operability of aeronautical compressors. Dedicated control devices such as guide fins are envisioned to reduce its negative effects. This paper presents a methodology based on RANS (Reynolds-Averaged Navier-Stokes) computations enabling to select guide fins efficient for that purpose. This methodology is applied to a reference case of linear compressor cascade operating at low Mach number (∼0.11). A set of 17 parameters is used to define two design spaces of interest, from which guide fins are generated. From then, an automated process generates and merges an unstructured mesh built around each guide fin with a fixed, structured mesh of reference representing a single channel of the cascade. Finally, RANS results on the resulting hybrid mesh are obtained using the Computational Fluid Dynamics solver elsA. This set up has proven successful in evaluating automatically hundreds of guide fins of various shapes. Several geometries illustrate the diversity of the design space. A selection of guide fins is then evaluated experimentally. Evolutions of the losses downstream of the cascade are compared to their respective RANS predictions, and to the reference case without guide fin. These experimental results validate the implemented methodology and show encouraging results in terms of loss redistribution brought by the control device.

This paper presents the investigation of realistic cavity leakage flow for the von Karman Institute for Fluid Dynamics H25 axial compressor stage, equipped with a shroud cavity of average Reynolds number Rer = 2.2 * 106 and average aspect ratio G = 0.06. On top of overall performance measurements, time averaged data from unprecedented experimental dataset are used to characterize the stator hub shroud cavity flow field and its interaction with the power stream. The sensitivity of the stage performance and stability as well as the cavity flow topology is evaluated against injection conditions, compressor throttling, and compressor speed. Parametric steady RANS simulations of the stage under injection are used to support the findings and state on the relevance of such approach in the preliminary design phase of axial compressor components. As an example, a stage efficiency reduction of 0.97% for a leakage fraction of 0.56% at design point is retrieved. This performance drop is attributed to the effect of cavity flow on: the blockage ratio, the boundary layer skewness, and total temperature increase at stator row inlet. It is also presented that there is a mutual interaction between the cavity geometry and the cavity flow field organization. The injection homogenizes the cavity flow field and the associated pressure gradient. The various operating conditions presented also demonstrate that the cavity flow studied is, on average (i.e., time), sensitive to changes in compressor rotation speed and changes in stage loading.

Non-synchronous blade vibration (NSV) has received enormous attention in turbomachinery research, since it can be safety critical and its occurrence is less predictable than typical aeroelastic instabilities such as flutter. Typical for front stages of core compressors, NSV arises due to the propagation of one or multiple aerodynamic disturbances, traveling around the circumference and locking-in with structural blade vibration modes. Recent studies have shown, that modern low-speed fan architectures are also susceptible for this type of instability. With a specific focus on such aeroelastic instabilities, the open-test-case fan stage ECL5 has been designed by Ecole Centrale de Lyon. It is representative of near-future UHBR fan concepts with carbon-fibre composite blades. To prepare the experiments planned for 2022, unsteady full-annulus simulations have been carried out. Depending on rotation speed, the emergence of aerodynamic disturbances traveling around the circumference at a typical convective speed is observed, such as reported in studies on NSV on different configurations. These aerodynamic disturbances are identified as a possible source for non-synchronous vibration. At design conditions, an unstable aerodynamic behavior is observed, with part-span stall cells developing within the operating range predicted by RANS simulations. A study of intentional aerodynamic mistuning, in form of an alternating leading edge pattern is presented. At design speed, influence on steady performance and the onset of part-span stall is observed. At part-speed, the evolution of small-scale vortical disturbances, which can be responsible for the onset of NSV is affected significantly, providing a possible method to suppress NSV.

This paper presents the investigation of realistic cavity leakage flow for the von Karman Institute for Fluid Dynamics H25 axial compressor stage, equipped with a shroud cavity of average Reynolds number Rer = 2.2 * 106 and average aspect ratio G = 0.06. On top of overall performance measurements, time averaged data from unprecedented experimental dataset are used to characterize the stator hub shroud cavity flow field and its interaction with the power stream. The sensitivity of the stage performance and stability as well as the cavity flow topology is evaluated against injection conditions, compressor throttling and compressor speed. Parametric steady RANS simulations of the stage under injection are used to support the findings and state on the relevance of such approach in the preliminary design phase of axial compressor components. As an example, a stage efficiency reduction of 0.97% for a leakage fraction of 0.56% at design point is retrieved. This performance drop is attributed to the effect of cavity flow on: the blockage ratio, the boundary layer skewness, and total temperature increase at stator row inlet. It is also presented that there is a mutual interaction between the cavity geometry and the cavity flow field organization. The injection homogenizes the cavity flow field and the associated pressure gradient. The various operating conditions presented also demonstrate that the cavity flow studied is, on average (i.e. time), sensitive to changes in compressor rotation speed and changes in stage loading.

Manuscript accepted for Proceedings of ASME Turbo Expo 2022 Turbomachinery Technical Conference and Exposition. GT2022June 13-17, 2022, Rotterdam, The Netherlands GT2022-81961; A composite fan stage representative of a modern Ultra High Bypass Ratio (UHBR) architecture has been investigated experimentally on a novel test facility at Ecole Centrale de Lyon. For these experiments, a turbulence control screen (TCS), a hemisphere, consisting of honeycomb and wiremesh panels, was installed in front of the rig in order to ensure homogeneous, disturbance-free inflow conditions. This kind of screen was developed in the early 1970 years to investigate flight-noise on static test beds. Today, they are commonly used in aero-engine and fan tests, because they improve the accuracy of acoustic measurements and provide more reproducible performance tests. However, the influence of such screens has been investigated only from an acoustic point of view in the past. This study presents a back-to-back comparison of measurements with and without the TCS, which allows a characterization of the effects on aerodynamic and aeroelastic behavior of the fan stage. The investigation was carried out for multiple stable operating points on two different speed lines to obtain a global image of the influence at part and design speed. The utilization of steady and unsteady instrumentation in different axial and circumferential duct positions allows a detailed analysis of similarities and differences of the configuration without and with the TCS and provides insight into underlying physical mechanisms.

A composite fan stage representative of a modern UHBR architecture has been investigated experimentally on a novel test facility at Ecole Centrale de Lyon. These measurements show indications for strong overloading of the tip region resulting in extensive blockage of the blade passage. The performance of the fan is analyzed with extensive instrumentation including radial profiles upstream and downstream of the rotor. Unsteady pressure measurements help to interpret the flow structure in the tip region. The results are presented across a range of operating points on the design speedline. At the stability limit, the machine suffers from Non-Synchronous Vibrations which result from small scale aerodynamic disturbances propagating between the leading edges. A detailed analysis on the occurring waveforms is presented for two operating speeds. In order to further analyze the observed phenomena, a numerical study has been conducted using the RANS solver elsA. The results of steady calculations are discussed in comparison with the detailed experiments. Unsteady simulations near the stability limit accurately predict the aerodynamic disturbances observed during NSV. The obtained results are unusual for typical state-of-the-art transonic fans, as they show the same behavior as high-pressure compressor front stages, dominated by blockage caused by tip leakage flow. Even though flutter is not observed, the observed Non-Synchronous Vibration mechanism is a critical aeroelastic phenomenon which is of great interest for future designs of low speed fans.

As a compressor is throttled three-dimensional separations develop in the corners between the blades and annulus endwall. Surprisingly, little is understood about the unsteady topology of these separations. One of the problems with studying corner separations is that it is often difficult to understand whether a particular flow structure in the separation is inherent to the separation itself, or due to the response of the separation to changes in the inlet flow. In this paper a novel experimental approach is taken with the aim of isolating the corner separation from external influences. A cascade is designed with the specific aim of precisely controlling the inlet flow. Contrary to previous work, it is shown that the key saddle and focus pair, which describes the time-mean topology of the corner separation on the endwall, moves smoothly and continuously as the incidence of the flow is raised. This behavior is shown to be the result of the time-resolved topology of the flow field, which comprises numerous saddle and focus pairs which are produced stochastically in regions of high shear strain rate. Most importantly, the separation is shown to exhibit an extremely low frequency behavior, changing in topology over timescales which are approximately 80 times the convection time through the blade passage. The behavior is shown to be intrinsic to the separation and causes the separation, for periods, to completely disappear from the endwall. This underlying unsteady structure of the separation is shown to have implications for the ability of RANS-based design codes to be able to accurately predict corner separations.

In axial compressors, shrouded stator cavity flows are responsible for performance degradation due to their interaction with the power stream. The present paper aims at exploring the possibility of employing a single stage high pressure axial compressor as a test vehicle for cavity flows investigations. In a first step, the robustness of the adopted RANS approach is tested against experimental data on the closed-cavity baseline configuration (i.e. no downstream-to-upstream recirculation). In a second phase, the effect of different hub cavities layouts of different levels of realism is numerically investigated. The focus is set on the representativeness of a closed cavity configuration with injection. The cavity flow topology and impact on the overall performance are considered in the analysis. At its final extent, this paper provides numerical and experimental guidelines for the robust assessment of cavity flows topology and performance effects.

In axial compressors, shrouded stator cavity flows are responsible for performance degradation due to their interaction with the power stream. The present paper aims at exploring the possibility of employing a single stage high pressure axial compressor as a test vehicle for cavity flows investigations. In a first step, the robustness of the adopted RANS approach is tested against experimental data on the closed-cavity baseline configuration (i.e. no downstream-to-upstream recirculation). In a second phase, the effect of different hub cavities layouts of different levels of realism is numerically investigated. The focus is set on the representativeness of a closed cavity configuration with injection. The cavity flow topology and impact on the overall performance are considered in the analysis. At its final extent, this paper provides numerical and experimental guidelines for the robust assessment of cavity flows topology and performance effects.

A specific phenomenon that has been observed in many experimental studies on turbomachinery compressors and fans is discussed under the term “rotating instabilities”. It is associated to a local aerodynamic phenomenon, typically occurring in the tip region at highly loaded near stall conditions and often linked to blade vibrations. Even though the effect has been discussed over more than two decades, a very ambiguous interpretation still prevails. A particular problem is that certain signatures in measurement data are often considered to characterize the phenomenon despite possible misinterpretations. The present paper illustrates that a specific image of a pulsating disturbance that has been established in the 1990s needs to be reconsidered. At the example of a recent investigation on a composite fan the difficulties concerning sensor placement and post-processing techniques is discussed with a focus on spectral averaging, isolation of non-synchronous phenomena and multi-sensor cross-correlation methods.

The inner flow environment of turbomachinery presents strong three-dimensional, rotational, and unsteady characteristics. Consequently, a deep understanding of these flow phenomena will be the prerequisite to establish a state-of-the-art design system of turbomachinery. Currently the development of more accurate turbulence models and CFD tools is in urgent need for a high-quality database for validation, especially the advanced CFD tools, such as large eddy simulation (LES). Under this circumstance, this paper presents a detailed experimental investigation on the 3D unsteady flow field inside a laboratory-scale isolated-rotor with multiple advanced measurement techniques, including traditional aerodynamic probes, hotwire probes, unsteady endwall static pressure measurement, and stereo particle image velocimetry (SPIV). The inlet boundary layer profile is measured with both hotwire probe and aerodynamic probe. The steady and unsteady flow fields at the outlet of the rotor are measured with a mini five-hole probe and a single-slanted hotwire probe. The instantaneous flow field in the rotor tip region inside the passage is captured with SPIV, and then a statistical analysis of the spatial distribution of the instantaneous tip leakage vortex/flow is performed to understand its dynamic characteristics. Besides these, the uncertainty analysis of each measurement technique is described. This database is quite sufficient to validate the advanced numerical simulation with LES. The identification process of the tip leakage vortex core in the instantaneous frames obtained from SPIV is performed deliberately. It is concluded that the ensemble-averaged flow field could not represent the tip leakage vortex strength and the trajectory trace. The development of the tip leakage vortex could be clearly cataloged into three phases according to their statistical spatial distribution. The streamwise velocity loss induced by the tip leakage flow increases until the splitting process is weak and the turbulent mixing phase is dominant.

A novel test facility for transonic fans has been constructed and commissioned at École Centrale de Lyon (ECL) within the Project PHARE-2. The facility is instrumented with multi-physical measurement systems to deeply investigate aeroelastic and aeroacoustic phenomena. To enable long-term fundamental research, the composite material open-test-case rotor ECL5 is currently under development, which shall be established as a new reference case for method development. Within the present publication, the project objectives, the current rotor design, ECL5v2, and details on the test facility are presented. With the goal to establish future academic collaborations, the authors aim to initiate a discussion in the research community in an early stage of the test case development to receive constructive feedback on the design and research approach.

In the present paper we show an analysis of Non-Synchronous-Oscillations in the rear stage of an experimental high-speed 3.5 stage compressor. The machine CREATE is representative of the rear median-stages of a modern civil turbofan engine except for an increased tip clearance on middle and last stage. Using distributed unsteady pressure measurements a comprehensive dataset during transient throttling maneuvers is presented, including temporal and spatial mode decomposition and the derivation of wave propagation speeds. To support the aerodynamic characterization, the steady flow structure at near stall conditions has been investigated using Laser-Doppler-Anemometry and unsteady pressure probes. In repeated experiments the machine has encountered alternating circumferential modes, several hundred revolutions before rotating stall at design speed. It was found that acoustic resonance of an upstream cut-on mode between the compressor stages enforces the development of particular modes. These phenomena have not been observed on the same machine with regular tip clearance. In contrast to front-stage observations reported in literature, where leading edge separations cause strong excitations of the blade eigenmodes, initially weak modulations of the flow in the rotor tip region are observed to be amplified due to acoustic feedback. With rising amplitude lock-in to distinct circumferential mode orders develops. Apart from the final lock-in procedure, signatures of solely aerodynamic fluctuations, discussed in literature as ‘Rotating Instabilities’ or ‘Part Span Stall’, are observed intermittently at different speedlines.

This paper describes observed modal oscillations arising from a feedback mechanism between an acoustic resonance in the exit flow channel and aerodynamic and aeroelastic disturbances in a transonic fan stage. During tests, the fan suffered from rotating stall and surge which were preceded by low frequency pressure fluctuations. Through a range of aerodynamic and aeromechanical instrumentations, it was possible to determine a clear chain of cause and effect, whereby geometrical asymmetries trigger local instabilities and modal oscillations through an interaction with the system acoustics. To the authors knowledge, this is the first time that modal oscillations occurring before stall are attributed to multiphysical interactions, showing that acoustic characteristics of the system can influence the aerodynamic as well as the aeromechanical stability of fans. This bears implications for the stability assessment of fans and compressors because first, the stability margin may be affected by standing waves generated in bypass ducts or combustion chambers, and second, geometrical variations of the rotor blades which are believed to be beneficial for aeromechanical stability may lead to complex coupling phenomena.

A local loss model and an integral loss model are proposed to study the irreversible flow loss mechanism in a linear compressor cascade. The detached eddy simulation model based on the Menter shear stress transport turbulence model (SSTDES) was used to perform the high-fidelity simulations. The flow losses in the cascade with an incidence angle of 2&deg, 4&deg, and 7&deg, were analyzed. The contours of local loss coefficient can be explained well by the three-dimensional flow structures. The trend of flow loss varying with incidence angle predicted by integral loss is the same as that calculated by total pressure loss coefficient. The integral loss model was used to evaluate the irreversible loss generated in different regions and its varying trend with the flow condition. It as found that the boundary layer shear losses generated near the endwall, the pressure surface and the suction surface are almost identical for the three incidence angles. The secondary flow loss in the wake-flow and blade-passage regions changes dramatically with the flow condition due to the occurrence of corner stall. For this cascade, the secondary flow loss accounts for 26.1%, 48.3% and 64.3% of the total loss for the flow when the incidence angles are 2&deg, respectively. Lastly, the underlying reason for the variation of the secondary flow loss with the incidence angle is explained using the Lc iso-surface method.

This paper presents an evaluation of the multiple-frequency phase-lagged approach, which enables the performance of unsteady Reynolds-averaged Navier–Stokes simulations on multistage turbomachinery configurations using a time-marching method. The major advantage of this approach is to reduce the computational domain to one single blade passage per row. The first part of the paper presents the method and discusses the associated assumptions and limitations. The method is then evaluated on the axial compressor configuration “Compresseur de Recherche pour l’Etude des effets Aérodynamiques et TEchnologiques” investigated experimentally at Laboratory of Fluid Mechanics and Acoustics. The computational fluid dynamics results are analyzed and compared both with experimental data and with a reference multipassage computation based on a sliding mesh approach. These comparisons enable the highlighting of the interests of this approach but also the underlining of its limits. The multiple-frequency phase-lagged approach enables the simulation of unsteady effects on a multistage turbomachinery and access to unsteady information that would not be available with a mixing-plane approach. However, if the method is capable of capturing unsteady effects linked to the adjacent upstream and downstream blade rows passing frequency, it fails modeling clocking effects, i.e., the relative influence between rows N and N+2. Read More: http://arc.aiaa.org/doi/abs/10.2514/1.B35247

The loss-generating mechanism of a linear compressor cascade at the corner stall condition was numerically studied in this paper. The hybrid RANS/LES method was used to perform the high-fidelity simulations. By comparing the results captured by SSTDES, DDES, SAS models with the experimental data, the SSTDES model is proven to be more accurate in capturing the detailed flow structure of the corner stall than the other two models. Taking the turbulence dissipation term of SSTDES model into account, the volumetric entropy generation rate and a new dimensionless local loss coefficient are proposed and used to analyze the loss-generating mechanism in this work. It was found that the main flow loss generated in this cascade could be sorted as the wake flow loss, the profile loss, the secondary flow loss and the endwall loss according to their amounts. The corner separation significantly affects the secondary flow loss, wake flow loss and profile loss in the cascade passage. The mixing between the separated boundary layer flow and the main flow, the shear between a tornado vortex and the main flow are the main sources of the secondary flow loss. The wake flow loss is the largest loss source of the cascade, accounting for 41.8% of the total loss. There are two peaks of the wake flow loss along the spanwise direction near the corner stall region. This phenomenon is related to the appearance of large velocity gradient flows when the main flows and the corner separation flows mix together. The profile loss takes up 40.06 % of the total loss. The profile loss intensity in the corner region is lower than the mid blade span due to the interaction of the boundary layer on the suction side with the corner separation.

This paper considers the inherent unsteady behavior of the three dimensional separation in the corner region of a subsonic linear compressor cascade equipped of thirteen NACA 65-009 profile blades. Detailed experimental measurements were carried out at different sections in spanwise direction achieving, simultaneously, unsteady wall pressure signals on the surface of the blade and velocity fields by time-resolved PIV measurements. Two configurations of the cascade were investigated with an incidence of 4° and 7°, both at Re = 3.8 * 105 and Ma = 0.12 at the inlet of the facility. The intermittent switch between the two statistical preferred sizes of separation, large and almost suppressed, is called bimodal behaviour. The existence of such oscillation, reported at first in previous experimental and numerical works on the same test rig, is confirmed for both incidences. Additionally, the present PIV measurements provide, for the first time, time-resolved flow visualizations of the size switch of the separation with an extended field of view covering the entire blade section. The interaction of random large structures of the incoming boundary layer with the blade is found to be a predominant element that destabilizes the separation boundary. The recirculation region enlarges when these high vorticity perturbations blend with larger eddies situated in the aft part of the blade. Such massive separation persists until the blockage in the passage causes the breakdown of the largest structures in the aft part of the blade. The flow starts again to accelerate and the separation is almost suppressed. Finally, POD analysis is carried out to decompose flow modes and to contribute to the clarification of underlying cause-effect-relations, which predominate the dynamics of the present flow scenario.

Large-eddy simulation (LES) is compared with experiment and Reynolds-averaged Navier-Stokes (RANS), and LES is shown to be superior to RANS in reproducing corner separation in the LMFA-NACA65 linear compressor cascade, in terms of surface limiting streamlines, blade pressure coefficient, total pressure losses and blade suction side boundary layer profiles. However, LES is too expensive to conduct an influencing parameter study of the corner separation. RANS approach, despite over-predicting the corner separation, gives reasonable descriptions of the corner separated flow, and is thus selected to conduct a parametric study in this paper. Two kinds of influencing parameters on corner separation, numerical and physical parameters, are analyzed and discussed: second order spatial scheme is necessary for a RANS simulation; incidence angle and inflow boundary layer thickness are found to show the most significant influences on the corner separation among the parameters studied; unsteady RANS with the imposed inflow unsteadiness (inflow angle varying sinusoidally with fluctuating amplitude of 0.92°) does not show any non-linear effect on the corner separation.

A study of the acoustic resonance occurring and leading to surge in the 3½-stage high-speed compressor CREATE (rotation speed: 11543 r/min, rotor 1 tip speed: 313 m/s) is presented in this paper. It is based on high-frequency wall static pressure measurements performed all along the machine from peak efficiency to surge. In this paper, a model of acoustic resonance will be applied to the experimental results. The axial distribution of the amplitude of the acoustic modes is first analyzed. Then, a study of the reflections occurring in the machine will be done in order to understand the mechanisms responsible for the resonance.

In this work, unsteady numerical results and high-frequency measurements are investigated from nominal to loaded operating points with the objective to contribute to the understanding of pre-stall rotating disturbances. A 3.5 stages high speed axial multistage compressor is investigated on a 2 MW test rig in the laboratory of fluid mechanics (LMFA) at Ecole Centrale de Lyon, France. The compressor has been built by Snecma, and is representative of modern high-pressure rear blocks of a modern aircraft engine. The unsteady numerical results predict a rotating disturbance in the tip flow field of the rotor 2 at the loaded operating point. It causes a frequency which is not a multiple of the periodicity of the compressor, and is rotating at about 72% of the shaft speed. The mesh independency of this disturbance is ensured. The analysis of the circumferential and axial propagation of the disturbance reveals a rotating instability like phenomenon. Most characteristic is the very important periodic oscillation of the tip leakage vortex trajectory, leading to a modulation of the leakage flow in the neighboring tip gap. The influence of the neighboring blade rows is investigated by filtering their unsteady contribution by means of mixing planes up and/or downstream of the rotor 2. In either case, the rotating disturbance is found to be still present. There are no traces of this rotating disturbance in the high-frequency measurements investigated at a near surge operating point. A spike like stall inception and almost instantaneous surge inception is identified. The mis-prediction of the tip region flow field in the rotors 2 and 3 is believed to cause the mis-prediction of the pre-stall disturbance.

Recently bimodal phenomenon in corner separation has been found by Ma et al. (Experiments in Fluids, 2013, doi:10.1007/s00348-013-1546-y). Through detailed and accurate experimental results of the velocity flow field in a linear compressor cascade, they discovered two aperiodic modes exist in the corner separation of the compressor cascade. This phenomenon reflects the flow in corner separation is high intermittent, and large-scale coherent structures corresponding to two modes exist in the flow field of corner separation. However the generation mechanism of the bimodal phenomenon in corner separation is still unclear and thus needs to be studied further. In order to obtain instantaneous flow field with different unsteadiness and thus to analyse the mechanisms of bimodal phenomenon in corner separation, in this paper detached-eddy simulation (DES) is used to simulate the flow field in the linear compressor cascade where bimodal phenomenon has been found in previous experiment. DES in this paper successfully captures the bimodal phenomenon in the linear compressor cascade found in experiment, including the locations of bimodal points and the development of bimodal points along a line that normal to the blade suction side. We infer that the bimodal phenomenon in the corner separation is induced by the strong interaction between the following two facts. The first is the unsteady upstream flow nearby the leading edge whose angle and magnitude fluctuate simultaneously and significantly. The second is the high unsteady separation in the corner region.

This study takes place in the frame of a research project to better understand the flow that develops in multistage high-speed compressors. First, the paper presents the high-speed 3.5-stage compressor CREATE and the methods that are considered to increase the data reliability and the investigation capability with such realistic compressors. Second, flow simulations, achieved with three-dimensional unsteady Reynolds-averaged Navier-Stokes computations over the whole compressor spatial and temporal periodicities, are analyzed through a study of their sensitivity to some parameters such as the inlet conditions, the space and time discretization, and technological effects such as tip and hub clearances. Finally the paper focuses on the methodology used to compare the experimental and numerical results over the whole compressor periodicities. The local spatial and temporal flow structures are well estimated. The key is the advection of these structures, which interact with each other and produce a significant part of the flow fluctuations in the downstream stages. The paper presents the zone close to the casing, where losses and blockage are induced by the interaction between the tip leakage flow and the incoming wakes. This is where the numerical simulation has to be improved, in order to accurately predict the performance.

Numerical and experimental investigations were conducted in a transonic centrifugal compressor stage composed of a backswept splittered unshrouded impeller and a vaned diffuser. The present article focuses on the results obtained within the impeller, at an operating condition close to the surge of the compressor. The experimental results were obtained from a laser Doppler anemometry investigation. Unsteady numerical simulations of the compressor stage were performed using a three-dimensional Reynolds-averaged Navier—Stokes code with a phase-lagged technique, at both peak efficiency and close to surge operating conditions. A good agreement between the experiments and simulations were obtained, which justifies the use of the computational fluid dynamics results for the comparison of the flow field at both operating conditions (peak efficiency and near surge). Even if the change in flow field within the impeller from peak efficiency to near surge looked to be gradual, an overall rotation of the whole flow in the blade passages led to a non-homogeneous flow at the impeller exit in terms of angle and velocity level. Therefore, the vaned diffuser has to tolerate upstream flows, which are all the more distorted as the operating point moves towards surge.

The aim of this work is to present detailed unsteady pressure measurements of three-dimensional flow field in a NACA 65 linear compressor cascade. Chord-based Reynolds number of 382000 and incidence angle of 4 degrees were chosen as target configuration of the rig, which clearly presents the corner separation phenomenon at the juncture of the blade suction side and the end-wall. Concerning the experiments, a characterization of the mean and fluctuating component of wall static pressure on the surface of a specially developed blade is achieved at first. This fluctuating component is investigated utilizing nineteen high sensitivity condenser microphones plugged into blade cavities which have been carefully calibrated. Transfer functions obtained by calibration are exploited to reconstruct the time-dependent pressure signal and finally statistics, conditional ensemble averages, coherence and spectra analyses of fluctuations are presented in order to investigate the unsteady characteristics of the corner separation. High values of root mean square are individuated near the leading edge and in the separation region on the suction surface of the blade. Skewness and kurtosis show an intermittent behavior of the separation onset, which moves upstream and downstream on the suction surface. This intermittency of the separation line is probably linked with the existence of a bimodal behavior of the size of the corner separation. The analyses of coherence and conditional ensemble average between the signals at the leading edge and at the onset of the separation suggest a critical influence of angle and velocity of the incoming end-wall boundary layer on the positive pressure signatures of the shear layer, which characterize the inception of the separation.

The present study carried out a pure large eddy simulation (LES) on a NACA65 linear compressor cascade (chord-based Reynolds number: 382 000), at an incidence angle of 4°. In this configuration, a corner separation at the juncture of the blade suction side and the end-wall is clearly observed experimentally. For the simulation, 852 CPUs are used in parallel and particular attention is paid to the inflow conditions, coupling LES with the calculation of the inlet boundary layer and associated velocity fluctuations. The shear-improved Smagorinsky model, which has been proved to be competent for turbomachines, is used in this study. Numerical results are carefully compared with pressure measurements, two-dimensional and stereo particle image velocimetry. LES is found to be superior to Reynolds-averaged Navier–Stokes approach in predicting the three-dimensional separated flows in the compressor cascade since pressure coefficients and losses are in very good agreement with the experiment. LES has been proven to simulate also the unsteady behavior of the separation. Finally, the experimental and numerical analysis of the inflow angle reveals high fluctuations of incidence upstream the leading edge of the blade near the end-wall, mostly associated with the incoming wall boundary layer.

Aerodynamic instabilities such as stall and surge may occur in compressors, possibly leading to mechanical failures so their avoidance is crucial. A better understanding of those phenomena and an accurate prediction are necessary to improve both the performance and the safety. A surge event in a compressor threatens the mechanical integrity of the aircraft engine, and this remains true for a research compressor on a test rig. As a result, few experimental data on surge are available. Moreover, there are technological, restrictive constraints that exist on test rigs and limit severely the type of data obtainable experimentally. This partially explains why numerical simulation has become a usual, complementary and convenient tool to collect data in a compressor, as it does not disturb the flow nor does it encounter technological limits. Despite the inherent difficulties, an entire surge cycle has been simulated in a high-speed, high-pressure, multistage research compressor, using an implicit, time-accurate, 3D compressible unsteady Reynolds-averaged Navier–Stokes solver. First, the paper presents the main features of the surge cycle obtained, along with those from the experimental cycle, for a validation purpose. Four phases compose the surge cycle: surge inception, the reversed-flow phase, the recovery phase, and the repressurization of the compressor flow. All of them are described, and focus is put on surge inception and the reversed-flow phase, as they induce greater risk for the mechanical integrity of the machine.

When measurements are performed in high speed, small-scale compressors, the use of curved glass windows is required in order to avoid any mismatch between the measurement window and the casing. However, the glass curvature leads to optical distortions, which hinder acceptable measurements and can even prevent the acquisition of any data. Thus, an original optical assembly, which consists in inserting a simple and inexpensive corrective window between the frontal lens of the anemometer and the shroud window, is proposed. The way of determining the geometric characteristics and the position of this corrective window, which restores very acceptable foci, is presented in the paper. The reliability of this corrective optical assembly is highlighted by comparative measurements in a test case. Using such an optical setting, L2F measurements were realised along a section, downstream of the inlet guide vane (IGV) of a transonic compressor stage. The spatial resolution leads to a good description of the interaction of the wake with the oblique shock emanating from the leading edge of the rotor. A phenomenological study of the wake/shock interaction with a change of frame is realised using the streamwise equation of the transport of vorticity.

Pressure measurements using high frequency response sensors have been carried out on the 3rd rotor of the 3 1/2-stage high speed compressor CREATE (rotation speed: 11543 RPM, Rotor 1 tip speed: 313 m/s), over the complete characteristic line and during the surge transient. It seems that precursors to the instabilities occurring near surge are observable at stable operating points. Just before the surge, these precursors characterized as rotating disturbances transition smoothly to part span rotating stall cells, and bring the machine to surge as they grow and reach a critical size: this paper presents a detailed analysis of the phenomena of rotating stall and surge transient and shows that it is possible to develop an anti-surge active control system based on the early detection of the precursors.Copyright © 2012 by ASME

It is widely accepted that in a turbulent boundary layer (TBL) with adverse pressure gradient (APG) an outer peak usually appears in the profile of streamwise Reynolds stress. However, the effect of APG on this outer peak is not clearly understood. In this paper, the effect of APG is analysed using the numerical and experimental results in the literature. Because the effect of upstream flow is inherent in the TBL, we first analyse this effect in TBLs with zero pressure gradient on flat plates. Under the individual effect of upstream flow, an outer peak already appears in the profile of streamwise Reynolds stress when the TBL continues developing in the streamwise direction. The APG accelerates the appearance of the outer peak, instead of being a trigger.

This contribution addresses the possibility of exploiting the temperature dependency of piezoresistive sensors as a temperature measurement per se. This requires the characterization of the sensor, or the probe as a temperature probe, i.e. determination of the recovery factor between the sensor temperature and the flow temperature. This temperature calibration as well as the determination of the thermal response time was performed for two probe geometries: a high temperature flush-mounted and a low temperature sub-surface mounted single sensor total pressure probe, both with a probe head diameter of 2.5 mm. Two applications are reported. The first application was performed with the flush-mounted sensor probe in the high-speed 3 1/2-stage axial compressor CREATE tested in the 2 MW test rig of LMFA at École Centrale de Lyon, in France. The probes were traversed at each inter-row section up to temperatures of 180°C and an absolute pressure of 3 bar. The second application was performed with the sub-surface mounted sensor probe in the high-speed single stage R-4 compressor test rig of the von Karman Institute in Belgium. Both applications have shown results in extremely good agreement with simultaneous total temperature measurements with a Kiel type thermocouple probe. They also underline the necessity of a very accurate temperature calibration. Finally, considering the fact that a simultaneous temperature measurement can be obtained at the same location as the pressure measurement from the sensor, it is possible to derive entropy generation after a blade row, based on the average pressure and temperature quantities. This unveils another extremely interesting aspect of using the fast response probe technique in turbomachinery applications.

This paper describes the investigations performed to better understand unsteady flows that develop in a three-stage high-pressure compressor. More specifically, this study focuses on rotor-stator interactions and tip leakage flow effects on overall performance and aerodynamic stability. The investigation method is based on three-dimensional unsteady RANS simulations, considering the natural spatial periodicity of the compressor. Indeed, all information related to rotor-stator interactions can be computed. A comparison is first done with experimental measurements to outline the capacity of the numerical method to predict overall performance and unsteady flows. The results show that the simulation correctly estimates most flow features in the multistage compressor. Then numerical data obtained for three configurations of the same compressor are analyzed and compared. Configurations 1 and 2 consider two sets of tip clearance dimensions and a casing treatment based on a honeycomb design is applied for configuration 3. Detailed investigations of the flow at the same operating line show that the tip leakage flow is responsible for the loss of stability in the last stage. An increase by 30% of the tip clearance dimension dramatically reduces the stable operating range (by 40% with respect to the standard configuration). A modal analysis shows that the stall process in this case involves the perturbation of the flow in the last rotor by upstream stator wakes, leading to the development of a rotating instability. The control device designed and investigated in this study allows for reducing the sensitivity of the compressor to tip leakage flow by recovering the initial stable operating range.

International audience; Passive control devices based on casing treatments have already shown their capability to improve the flow stability in axial compressors. However, their optimization remains complex due to a partial understanding of the related physical mechanisms. In order to quantitatively assess the interaction between slots and the blade tip flow, the present paper develops a novel analysis methodology based on a control-volume approach located in the rotor tip region. This methodology may be used for analyzing the casing treatment based on both axi- and non-axisymmetric slots design. The second issue of the paper focuses on the application of the current approach to better understand the effects of axi- and non-axisymmetric grooves in three different axial compressors which differ by the flow regime (subsonic/transonic) and the smooth casing shape (cylindrical/concave). Numerical simulations are performed, and results of the current approach with and without casing treatments are compared.

Over the last decades, fast response aerodynamic probes have been recognized as a robust measurement technique to provide time-resolved flow field data in turbomachinery environments. Still, most of the existing probe designs are restricted to low temperature applications (< 120 °C) either because of sensor temperature range limitations or packaging issues. Measurements in turbomachines also require a small probe size often with a very high bandwidth which are conflictual constraints difficult to satisfy simultaneously. This contribution therefore presents the development of a novel miniature (O 2.5 mm) high temperature single sensor total pressure probe, designed for operation up to 250 °C with a very high bandwidth of 250 kHz. The probe main element is a 1.7 mm diameter commercial piezoresistive transducer placed in a Pitot type arrangement with a flush mounted sensor to provide the highest bandwidth. The details of the probe design are presented as well as the probe calibrations in pressure and in temperature. The effects of using a thermal compensation module or a sense resistor to monitor the temperature drift are described in the context of measurement uncertainty. The probes were characterized in terms of aerodynamic characteristics versus flow angle and Mach number. Shock tube tests have shown a dynamic response of the probe with sensor resonance frequencies well over 300 kHz, with a flat frequency range up to 250 kHz. Two probe prototypes were manufactured and first used in the 3 1/2-stage high speed axial compressor CREATE of the LMFA at Ecole Centrale de Lyon in France. The probes were traversed at each inter blade row plane up to temperatures of 180 °C and absolute pressure of 3 bar. The probe was able to resolve the high blade passing frequencies (∼16 kHz) and several harmonics including rotor-stator interaction frequencies up to 200 kHz. Besides the average total pressure distributions from the radial traverses, phase-locked averages and random unsteadiness are presented. The probe spatial and temporal resolutions are discussed in the context of those results.Copyright © 2011 by ASME

A modal analysis method of the rotor-stator interactions in multistage compressors has been developed by LMFA. This method, based on a double modal decomposition of the flow over space and time, has been applied to numerical and experimental results of the high-speed 3½-stage compressor CREATE based at LMFA, Lyon-France. It reveals the presence of a very strong rotor-stator interaction which completely drives the flow at casing behind all the rotors. This modal analysis method applied to an unsteady RANS simulation permits to calculate the energy of the rotor-stator interactions and to plot energetic meridian maps to explain experimental results and to analyze the interaction in the whole machine.

In applied research, a lack of understanding of corner stall, i.e. the three-dimensional (3D) separation in the juncture of the endwall and blade corner region, which has limited the efficiency and the stability of compressors. Both Reynolds-averaged Navier-Stokes (RANS) and large eddy simulation (LES) still need to be calibrated for turbomachinery applications. In the fundamental research of the turbulent boundary layer (TBL), there are a lot of findings of the effects of curvature and pressure gradients, which also play an important role in physics of corner stall. The purpose of this thesis is (i) to carry out an experiment in a cascade, (ii) to gain a database that could be used to calibrate both RANS and LES, and (iii) to give some basic explanations of corner stall through investigating the TBL on the suction side at the mid-span which is more complex than those in the basic investigations but simpler than those in a real engine. A detailed and accurate experiment of 3D flow field through a linear compressor cascade has been set up. Experimental data were acquired for a Reynolds number of 3.82×10 ^5 based on blade chord and inlet flow conditions. Measurements have been achieved by hot-wire anemometry, pressure taps on blade and endwall, five-hole pressure probe, oil visualization, 2D particle image velocimetry (PIV),and two-component laser Doppler anemometry (LDA). An original and complete database was thus obtained. The TBL on the suction side at mid-span was investigated. The wall-normal negative pressure gradient restrains the separation, on the contrary to its influence in the corner stall. The streamwise adverse pressure gradient can be responsible for the development of Reynolds stresses. The remarkable phenomenon at measurement stations near the trailing edge of blade is that an inflection point occurs in each profile of the mean streamwise velocity. At this inflection point, the magnitudes of the Reynolds stresses reach their maximum values, and the direction of energy diffusion also changes. The velocity field in the corner stall was presented. Bimodal histograms of velocity exist in the experiment. The bimodal points mainly appear in the region around the mean interface of separated flow and non-separated flow. At a bimodal point the local two velocity components are non-independent from each other, due to the aperiodic interplay of two basic modes in the flow field. Two modes were proposed to interpret the physics of bimodal behaviour.

Control devices based on casing treatments have already shown their capability to improve the flow stability in compressors. However their optimization remains complex due to a partial understanding of the related physical mechanisms. The present paper proposes to use a budget analysis of the Navier Stokes equations to support the understanding of such flow phenomena. Based on the original work of Shabbir and Adamczyk (2005), the strength of the present contribution is to generalize the flow analysis method to all Navier-Stokes equations, including unsteady terms. A high-pressure multistage compressor equipped with circumferential casing grooves is chosen to demonstrate the potential of this approach. Steady and unsteady Reynolds-Averaged Navier-Stokes (URANS) equations are solved with a structured multi-blocks solver. Results are then briefly compared to experimental data to validate the numerical method. The analysis of the unsteady axial momentum equation for configurations with and without casing treatment points out some of the mechanisms responsible for the stability improvement. The analysis also indicates that the flow unsteadiness generated by upstream stator wakes (stator/rotor interaction) reduces viscous efforts and increases convective forces, significantly modifying the compressor stability. Finally, the proposed post processing method shows very interesting results for the understanding of circumferential grooves and it should be also used for non-axisymmetric casing treatment configurations.Copyright © 2011 by ASME

Numerical and experimental investigations were conducted in a transonic centrifugal compressor stage composed of a backswept splittered unshrouded impeller and a vaned diffuser. Unsteady 3D simulations were performed with the code elsA that solves the turbulent averaged Navier-Stokes equations, at three operating points: choked flow, peak efficiency and near surge. Unsteady pressure measurements up to 150 kHz were carried out in the entry zone of the vaned diffuser (in the vaneless space and in the semi-vaneless space) when the compressor came into surge. These static pressure sensors were mounted on the shroud enwall. The paper focuses on the vaneless and semi-vaneless space where the surge originates. A detailed analysis of the flow pattern coming from the unsteady computations from choked flow towards surge led to identify the physical mechanisms involved in the surge inception. It is shown that, when approaching surge, the flow is destabilized by a severe modification of the shock system in the vaned diffuser inlet. The first perturbation is acquired from the transducer located just upstream of the shock foot (i.e. on the vane suction side surface), indicating a movement of the shock towards the vaneless space. This perturbation travels upstream and leads to the strongest short-wavelength perturbation acquired from the transducer located just upstream of the vane leading edge. This strongest short-wavelength perturbation which level may reach almost four times the mean exit pressure value triggers the full scale instability.Copyright © 2011 by ASME

Stator vanes which are found at different stages in axial compressors are subject to major aerodynamic load fluctuations, due to cyclic excitations induced by the wakes of the rotors located upstream and also due to the pressure waves generated by the blade rows located downstream. In a vibratory fatigue context, it is important to ensure the structural strength of these stator vanes. A new technology of monoblock sectorised stator vanes has been developed to provide an easier manufacturing process and lower costs; in return, cyclic symmetry properties, which make the prediction studies easier, are lost in this case. Besides, these kinds of monoblock sectors are characterized by low damping which increases the criticality of the blade modes localization. The division into sectors of stator vanes creates a high density of eigenmodes in some short frequency ranges; this phenomenon is accompanied by non-negligible modal participations of neighbouring eigenmodes. Moreover, stator vanes sectors have shown a strong sensitivity to mistuning, making their vibratory behaviour prediction more difficult by means of deterministic prediction methods. In this paper, a phenomenological approach is followed to quantify the mistuning. Thus, several sensitivity to mistuning numerical studies were performed. They allowed to have first results about the relative predictable errors that could be done, depending on the mistuning. Some engine tests were also performed with gauges-equipped stator vanes sectors.

This paper presents the unsteady data acquisition system used to measure the pressure field in high speed compressors. Details and electronic sketches are given for the conditioners developed in-house that have been used to amplify and to filter the pressure signal with the aim of acquiring data up to 150 kHz. A discussion of the experimental results carried out in a centrifugal compressor is proposed. Through different processing of the pressure signals and a comparison with URANS simulations, the excitation of the pressure transducers by the pressure waves generated by shock waves that occur between the impeller and the diffuser is highlighted. The levels of pressure fluctuations measured when entering into surge are also presented and reveal very repetitive behaviour of the flow instabilities.

Doppler global velocimetry (DGV) has already been shown to be an interesting technique capable of measuring the three components of velocity in a plane. A 1-component DGV system is currently under development at the LMFA, using a stabilized continuous wave (CW) argon ion laser for emission. The receiver features only one camera for both signal and reference images and incorporates a DEFI system to adjust the incident laser light frequency and its transmission coefficient on the iodine cell absorption line. A description of the whole system is presented and a validation with measurements of axial velocities at several positions in a round free jet is proposed.

International audience; The increase of the thrust/weight ratio of aircraft engines is extremely restricted by different 3-D flow loss mechanisms. One of them is the corner separation that can form at the junction between a blade suction side and a hub or shroud. In this paper, in order to further investigate the turbulent characteristics of corner separation, large-eddy simulation (LES) is conducted on a compressor cascade configuration using NACA65 blade profiles (chord based Reynolds number: 3.82 × 105), in comparison with the previous obtained experimental data. Using the shear-improved Smagorinsky model as subgrid-scale model, the LES gives a good description of the mean aerodynamics of the corner separation, especially for the blade surface static pressure coefficient and the total pressure losses. The turbulent dynamics is then analyzed in detail, in consideration of the turbulent structures, the one-point velocity spectra, and the turbulence anisotropy. Within the recirculation region, the energy appears to concentrate around the largest turbulent eddies, with fairly isotropic characteristics. Concerning the dynamics, an aperiodic shedding of hairpin vortices seems to induce an unsteadiness of the separation envelope.

The present study is concerned with wake-induced unsteady effects in axial-compressor blade rows. The goal is to exploit these effects in order to design high-lift blades without increasing the profile loss, as has been achieved for low-pressure turbine blades. In the first part of this paper, the experimental means and the computational fluid dynamics tools are described. The rig features a flat plate that can be subjected to different velocity distributions representative of the suction side of a real compressor blade. Cylindrical bars mounted on a moving system simulate the incoming wakes from the upstream blade row in the compressor. Results are presented for steady flow and for unsteady compressor-like conditions. In all cases, the separation bubble of the steady flow is suppressed by the turbulence that is induced in the boundary layer by the wakes at approximately 10% of the suction side. Its reappearance is then delayed by a region of stable laminar-like flow and low loss due to the so-called calming effect that follows the wake-induced patch. The paper describes these phenomena in detail for one particular pressure distribution. It is then shown that it should be possible to increase the lift by 35% while keeping the same level of loss as the initial conventional pressure distribution of the study.