# ECL5/CATANA: Transition from Non-Synchronous Vibration to Rotating Stall at transonic speed

Alexandra P Schneider, Anne-Lise Fiquet, Nathalie Grosjean, Benoit Paoletti, Xavier Ottavy, Christoph Brandstetter

► To cite this version:

Alexandra P Schneider, Anne-Lise Fiquet, Nathalie Grosjean, Benoit Paoletti, Xavier Ottavy, et al. ECL5/CATANA: Transition from Non-Synchronous Vibration to Rotating Stall at transonic speed. European Turbomachinery Conference - ETC16, Mar 2025, Hannover, Germany. hal-04963573

HAL Id: hal-04963573

https://hal.science/hal-04963573v1

Submitted on 24 Feb 2025

HAL is a multi-disciplinary open access
archive for the deposit and dissemination of sci-
entific research documents, whether they are pub-
lished or not. The documents may come from
teaching and research institutions in France or
abroad, or from public or private research centers.

L'archive ouverte pluridisciplinaire HAL, est
destinée au dépôt et à la diffusion de documents
scientifiques de niveau recherche, publiés ou non,
émanant des établissements d'enseignement et de
recherche français ou étrangers, des laboratoires
publics ou privés.

<figure><img src="image_1.png" /></figure>

Distributed under a Creative Commons Attribution 4.0 International License
# ECL5/CATANA: Transition from Non-Synchronous Vibration to Rotating Stall at transonic speed

Alexandra P. Schneider*¹, Anne-Lise Fiquet¹, Nathalie Grosjean¹, Benoit Paoletti¹, Xavier Ottavy¹, and Christoph Brandstetter¹

¹Ecole Centrale de Lyon, CNRS, Universite Claude Bernard Lyon 1, INSA Lyon, LMFA, UMR5509, 69130, Ecully, France

## Abstract

Non-Synchronous Vibration (NSV), flutter or rotating stall can cause severe blade vibrations and limit the operating range of compressors and fans. To enhance the understanding of these phenomena, this paper investigates the corresponding mechanisms in modern composite Ultra-High-Bypass-Ratio (UHBR) fans based on the ECL5/CATANA test campaign. Extensive steady and unsteady instrumentation such as Stereo-PIV, fast-response pressure probes and rotor strain gauges are used to derive aerodynamic and structural characteristics of the rotor at throttled operating conditions. The study focuses on the analysis of the transition region from transonic to subsonic speeds where two distinct phenomena are observed. At transonic design speed, rotating stall is encountered, while NSV is observed at 90% speed. At the intermediate 95% speedline a peculiar behavior involving a single stalled blade was observed. The results emphasize that rotating stall and NSV exhibit different wave characteristics: Rotating stall comprises lower wave numbers and higher propagation speeds around 78% rotor speed, while small scale disturbances propagate at 57% rotor speed and lock in with blade eigenmodes, causing NSV. Both phenomena are observed in a narrow range of operation, and even simultaneously at specific conditions. The presented results contribute to the understanding of different types of operating range limiting phenomena in modern UHBR fans and serve as a basis for the validation of numerical simulations.

**Keywords:** Rotating stall, Non-Synchronous Vibration, Blade-to-blade variation, Stereo Particle Image Velocimetry, UHBR fan.

## 1 Introduction

Understanding the behavior of fans and compressors close to their stability limit remains an essential topic in turbomachinery research. Phenomena such as flutter [1, 2], Non-Synchronous Vibration (NSV) [3, 4] and rotating stall [5-7] can cause severe blade vibrations and limit the operating range. To ensure safe and efficient operation, the underlying mechanisms must be understood. Thus, the main objectives of the present study are:

1. to enhance understanding of relevant phenomena at high subsonic and transonic conditions.

2. to quantify disturbance characteristics at different rotation speeds to provide a benchmark for numerical simulations.

Within this context, an extensive experimental campaign was conducted with the open-test case fan stage ECL5/CATANA with the goal to characterize mechanisms that limit the operating range of a modern composite Ultra-High-Bypass-Ratio (UHBR) fan. During these experiments, two different phenomena were observed: Rotating stall occurs at transonic design speed conditions, NSV is observed at all subsonic part speedlines.

Rotating stall is a purely aerodynamic phenomenon that is characterized by the circumferential propagation of regions of separated flow, typically referred to as stall cells. Regarding its onset, two different types of precursors have been reported in literature. The first one is a short-length scale disturbance, called spike, that rapidly grows in amplitude forming a stall

cell within a few rotor revolutions [5-7]. The second type is a disturbance with larger length scale and often referred to as modal oscillation [8, 9]. Depending on the size and the number of stall cells, the onset of rotating stall is usually accompanied by a significant reduction of performance and high amplitude blade vibration [10-12].

NSV, in contrast, is an aeroelastic mechanism that involves the interaction of small-scale disturbances with a structural blade eigenmode. These features are described as small vorticity disturbances, resembling those associated to a spike, that occur at highly throttled but aerodynamically stable operating conditions and convect in the leading edge plane from one blade to another with a distinct speed. They can lock-in with a structural blade eigenmode when their frequency in the rotating frame of reference is coincident with an eigenfrequency of the blades [3, 4, 13-16].

To differentiate both phenomena in experiments, extensive instrumentation in the stationary and rotating frame of reference is required. The angular frequency of an aerodynamic disturbance in the stationary frame is given by

$$ \omega_a^S = N_a \Omega_a^S \quad (1) $$

and can be measured using unsteady pressure probes in the rotor casing. The respective propagation speed $\Omega_a^S$ can be derived based on cross-correlations of probes in different circumferential positions (see [17]) and used to calculate the aerodynamic wave number $N_a$, i.e. the number of disturbances around the circumference. For rotating stall in single stage fans and compressors the cell count is typically reported between one and three, large multistage compressors may de-

<footnote>*Corresponding author: alexandra.schneider@ec-lyon.fr</footnote>
velop higher cell counts. Wall pressure spectra show char-acteristic peaks at the angular frequency of propagation (≈ EO0.5 - EO0.8) and its harmonics according to Eq. 1. In con-trast, NSV is usually related to smaller disturbances, with a wave number slightly below the blade count [18–21]. Corre-sponding spectra show a rake of frequencies around half the blade passing frequency that is often, but misleadingly, re-ferred to as rotating instabilities in literature [13, 22–24].

The interaction of propagating disturbances with the rotor blades can be analyzed using blade-mounted strain gauges in the rotating frame of reference. The angular frequency of an aerodynamic disturbance $\omega_a^S$ in the rotating frame (again nor-malized with the shaft speed $\Omega_r$) is given by

$$ \frac{\omega_a^R}{\Omega_r} = \frac{\omega_a^S}{\Omega_r} - N_a \quad (2) $$

and lock-in with a structural mode can occur when this fre-quency is coincident with a blade’s eigenfrequency $\omega_v^R$, i.e. when

$$ \omega_v^R = \omega_a^R \quad (3) $$

Equation 3 can only be fulfilled if the inter-blade phase angle of the aerodynamic wave $\sigma_a$ is coherent with the structural inter-blade phase angle $\sigma_v$, i.e. the difference in vibration phase between one blade and its neighbors

$$ 2\pi \frac{N_v}{N_b} = \sigma_v = \sigma_a = 2\pi \frac{N_a}{N_b} \rightarrow N_v \equiv N_a \ (\text{mod } N_b) \quad (4) $$

This yields that aerodynamic wave $N_a$ and structural nodal di-ameter $N_v$ are identical if the wave number is below half the number of blades $N_b$, otherwise aliased to $N_v = N_b - N_a$ [4].

Based on this methodology, the following study presents an in-depth analysis of the behavior of the ECL5 fan stage close to the border of the operating range. The focus is laid on the transition between a purely aerodynamic phenomenon at transonic design and aeroelastic phenomenon at subsonic part speed conditions. It will be shown that the fan experiences strong circumferential variations of the flow field within the investigated speed range that affect these phenomena. The aerodynamic analysis based on wall pressure data is supported by Stereo Particle Image Velocimetry (SPIV) measurements for selected operating points to highlight the significance of aerodynamic asymmetries, i.e. spatial deviations of the flow field from the cyclic-symmetry of an ideal rotor assembly.

The data presented herein complements previous publica-tions on the open-test-case ECL5 and serves as extended vali-dation for simulations at highly throttled operating conditions.

## 2 Open-test-case ECL5

The investigated open-test-case fan stage ECL5 was designed by the Turbomachinery Department of the Ecole Centrale de Lyon and is representative of modern low-speed engine ar-chitectures. The stage consists of 16 carbon-fiber rotor blades and 31 aluminum outlet guide vanes (OGV). The aerodynamic design point targets a standard mass flow rate of 36.0 kg/s at 11.000 rpm with a total pressure of 1.35 and an isentropic ef-ficiency of 92 %. At this condition, the rotor is running tran-

sonic with a relative tip Mach number of approximately 1.

The open-test-case was investigated within the European CleanSky-2 project CATANA<fnref n="1" /> with the goal to identify and characterize the multi-physical interactions between aerody-namics, acoustics, and structural dynamics in the context of future Ultra-High Bypass Ratio (UHBR) engines, with a spe-cial focus on NSV.

A total of 48 blades were manufactured and inspected us-ing a 3D scanner and frequency response based on ping tests. This process determined individual static moments, mass, and eigenfrequencies enabling a thorough selection to build dif-ferent rotor configurations. The blades for the reference con-figuration (CATREF) were chosen with the goal to assemble a structurally tuned rotor. The respective aerodynamic perfor-mance and aeroelastic stability are detailed in [25] and [26].

To quantify the sensitivity of observed aerodynamic and aeroelastic phenomena to intentional structural mistuning (i.e. increased variation of eigenfrequencies of mode 1, 2 and 3), the fan configuration CATMIS, was experimentally tested. The study presented in [27] provides a detailed comparison of both rotor configurations at highly throttled operating condi-tions at 80 % speed, highlighting a phenomenologically sim-ilar aeroelastic behavior regarding observed vibration modes and corresponding nodal diameters. Measured differences in vibration amplitudes could be attributed to (unintended) local deviations in the aerodynamic field at the blade tip caused by manufacturing inaccuracies (correlations of tip geometry and flow field variations will be discussed in section 3.1 and 3.2) which were found to be more important than the introduced structural mistuning.

In this study, the mistuned configuration is used to inves-tigate the transition of operating range limiting mechanisms between high subsonic and transonic conditions. The paper comprises a short introduction of the test setup and employed methods, followed by a detailed decomposition of aerody-namic and aeroelastic characteristics at highly throttled con-ditions. The reader is referred to previous publications for details on permanent instrumentation and test case design.

### 2.1 Test facility and instrumentation

The experiments with the open-test-case were conducted on the facility ECL-B3/Phare-2 at Ecole Centrale de Lyon (ECL). As schematically shown in Fig. 1, the facility is operated in an open cycle and driven by an electric motor with a maximum power of 3MW, as detailed in [28]. The core section is located in an anechoic chamber and a turbulence control screen (TCS) is installed in front of the intake to ensure homogeneous in-flow conditions [29].

A schematic illustration of the instrumentation of the core section is provided in Fig. 2, focusing on the instrumentation in the rotor section. A commercial capacitive tip timing/tip clearance system provided by MTU [30] is used to derive blade individual tip clearances and stagger angles during ex-periments. A multitude of unsteady wall pressure transducers (WPT) in different axial and circumferential positions allows for the analysis of the flow field at the blade tip as well as to derive the characteristics of propagating disturbances (wave

<footnote n="1">¹Composite Aeroelastics and Aeroacoustics, catana.ec-lyon.fr</footnote>
<figure><img src="image_2.png" /><figcaption>Figure 1: Test facility ECL-B3/Phare-2</figcaption></figure>

number and propagation speed). During machine operation, in particular for investigations of the highly throttled operating conditions, unsteady pressure signals are online-monitored to detect the onset of rotating stall. Blade-mounted strain gauge signals are used simultaneously to derive blade vibration amplitudes and frequencies to ensure structural integrity.

The light-sheet for the PIV measurements is introduced via a probe slightly upstream of the leading edge plane and optical access for the cameras is provided by a casing window in the rotor section as presented in Fig. 2(b). In the following, the PIV setup will be described in more detail, further information on all other measurement systems can be found in [28].

<figure><img src="image_3.png" /><figcaption>Figure 2: Facility core section: (a) schematic illustration of instrumentation; (b) PIV light-sheet position</figcaption></figure>

## 2.2 Stereo-PIV setup

Particle Image Velocimetry (PIV) is an optical measurement technique that is used to derive velocity fields. The flow is enriched with petroleum oil fog (seeding particles), illuminated in one plane using a light-sheet optic and an Nd:YAG double-pulse laser (Litron Bernoulli-PIV 200-15, wave length 532 nm) and recorded with a double-frame camera (LaVision Imager sCMOS, resolution 2560x2160 px). Image cross-correlation methods are used to convert particle displacements between the first and the second frame into two dimensional velocity fields based on the time delay between both images. Using a second double-frame camera with a different angle of view with respect to the illuminated plane enables the calculation of the third velocity component perpendicular to the measurement plane using stereo-reconstruction [31].

The experimental setup used for the Stereo-PIV measurements is shown in Fig. 2 providing a light-sheet position at approximately 93% channel height. The angle between both cameras is approximately 35°, resulting in higher uncertainties of the velocity component normal to the light-sheet plane [31]. To align the focal plane of the cameras with the illuminated plane, Scheimpflug adapters (LaVision) are utilized. The cameras are calibrated using a three-dimensional calibration target and refined after image acquisition using a self-calibration procedure with recorded particle images to compensate potential misalignment between the light sheet and the calibration target as described by [32].

Seeding particles are introduced in the facility chamber. VapORIZATION of petroleum based smoke oil provides particles with a diameter of 200 to 300 nm.

The system enables the acquisition of images with a fixed phase angle with respect to the rotor. This is required to derive statistically reliable velocity fields by recording several images at the same angular position (here 100 images have been recorded per phase angle).

## 2.3 Stereo-PIV measurement procedure and post-processing

The setup described above is used to record two double-frames (one for each camera) of the illuminated particles in the light-sheet plane using the acquisition parameters summarized in Table 1.

<table><caption>Table 1: PIV recording parameters</caption><thead><tr><td>parameter</td><td>value</td><td>unit</td></tr></thead><tbody><tr><td>frequency</td><td>≈ 14</td><td>Hz</td></tr><tr><td>time delay</td><td>2-2.5</td><td>μs</td></tr><tr><td>image scale</td><td>≈ 24</td><td>px/mm</td></tr><tr><td>laser pulse energy</td><td>40-60</td><td>mJ</td></tr><tr><td>max. repetition rate</td><td>14</td><td>Hz</td></tr><tr><td>particle image diameter</td><td>3-6</td><td>px</td></tr><tr><td>camera opening</td><td>F4-F5.6</td><td>-</td></tr></tbody></table>

### Image pre-processing, cross-correlation and post-processing of velocity fields

To improve the quality of the recorded raw images, digital image processing methods are applied to correct camera and
machine vibrations and to reduce reflections and noise [16].

After the pre-treatment, image cross-correlations are performed for each camera individually. To achieve a robust velocity detection for complex three-dimensional flows, a multi-grid procedure with decreasing interrogation window sizes is chosen, yielding a final resolution of 32x32 px corresponding to approximately 1.34x1.34 mm.

Based on this procedure, one two-dimensional velocity field is obtained for each camera that corresponds to the projection of the actual three-dimensional field to the respective measurement planes. These velocity fields are used for a stereo-reconstruction as described in [31], which allows to calculate both in-plane velocity components u and v as well as the third velocity component w perpendicular to the light-sheet plane. Subsequently, this velocity field is transformed from the coordinate system, that is aligned with the calibration target, to the machine coordinate system with the components $c_{ax}$, $c_{θ}$ and $c_{r}$ in axial, circumferential and radial direction respectively. The flow in the rotor relative frame of reference can be analyzed by subtracting the rotational speed $U_{rot} = Ω_{r} \cdot r$ from the circumferential velocity component $c_{θ}$. The relative in-plane velocity is then given by

$$c_{rel} = \sqrt{c_{ax}^2 + (c_{\theta} - U_{rot})^2} \quad (5)$$

To account for changes of atmospheric conditions, velocities are standardized with respect to measured total pressure and temperature in the facility chamber.

## Phase angle superposition

As schematically illustrated in Fig. 3, the field of view of a single recording with a fixed angular position of the light sheet with respect to the rotor is limited and multiple phase angles (i.e. angular positions with respect to the once-per revolution trigger) are required to cover a full blade passage. After transformation to the machine coordinate system, data from different phase angles can be superimposed. To simplify visualization, data is projected to the x,rθ-plane at constant radius. This projection results in a direct superposition of values from different channel heights as visible in Fig. 3(b) showing the variation of the radius r within one phase angle and arising steps between adjacent angular measurement positions. They can cause discontinuities in the presented velocity contours between overlapping phase angles due to the presence of strong gradients in the flow field.

## Measurement uncertainty

The measurement conditions in the facility are far from ideal due to reflections, vibrations and seeding deposition on the casing window. As a result, the uncertainty of image correlation is estimated to be around 4% of the measured velocity components based on the study with a comparable setup in [33]. In regions of increased secondary flows and mixing, particle segregation and reduced raw image quality can lead to locally higher uncertainties of up to 10% within individual frames, in particular at throttled operating conditions.

Further parameters that affect the overall accuracy of PIV measurements are (a) the spatial measurement uncertainty, i.e. the position of measured flow features such as the compression shock with respect to the blade, caused by particle inertia

<figure><img src="image_4.png" /><figcaption>Figure 3: Phase angle superposition: (a) schematic illustration; (b) contour of radial positions of projected phase angles</figcaption></figure>

(see [34]) and (b) the transformation to the machine coordinate system which causes inaccuracies as a result of variations of the rotor blade with the respect to the laser pulse (due to trigger inaccuracies) or slight deviations of the light-sheet position (due to machine vibration).

Even if described errors of measured instantaneous velocity fields are quite high compared to other instrumentation systems, the acquisition of 100 individual images per recording and the application of an outlier detection yields reliable averaged velocity fields, i.e. all PIV measurements presented below show converged statistical quantities<fnref n="2" />. Since the focus of the presented study is the qualitative comparison of the tip flow structure at different operating points, related uncertainties are considered as acceptable.

## 2.4 Machine performance

The experimental results presented in this study were obtained with the mistuned rotor CATMIS. Figure 4 shows the measured machine total pressure ratio as a function of massflow rate for different rotational speeds, black lines correspond to performance measurements while gray crosses were obtained during operability measurements with less intrusive instrumentation (details can be found in [25, 28]).

Within this study, the operating points highlighted by colored symbols at 90, 95 and 100% speed (denoted as N90, N95 and N100) will be investigated in detail. For these operating points, uncertainties of measured standard massflow and total pressure ratio are below 0.2 and 0.02% respectively as shown in [28]. To ensure comparable throttle settings across speedlines, operating conditions at equal flow coefficient $Φ = c_{ax}/U_{rot}$ are investigated, ranging from 0.506 to 0.451. As reported in [26] for the reference rotor configuration, significant differences regarding observed phenomena are observed in this region. While convective NSV is encountered at subsonic part speed conditions up to 90% speed leading to high blade vibration amplitudes, the onset of rotating stall without any vibratory precursors is observed at 100%

<footnote n="2">²The average velocity is considered converged when the cumulative sliding average is within ±1 m/s range (approx. 0.5% with respect to the lowest velocity) with respect to the average of all images, for the standard deviation a range of ±0.5 m/s is considered. In a large part of the measurement area, 40 images are sufficient to achieve convergence, 100 images are only required in regions of high unsteadiness</footnote>
<figure><img src="image_5.png" /><figcaption>Figure 4: Machine total pressure ratio for mistuned rotor configuration (CATMIS)</figcaption></figure>

speed, where the rotor is running transonic. This behavior is visible in Fig. 5 showing the evolution of average vibration amplitude of mode 1, 2 and 3 as a function of flow coefficient $Φ$ for the mistuned rotor configuration. In this figure, the vibration amplitude of each mode is scaled with respect to the maximum allowed strain at the strain gauge position, i.e. in percent scope limit. This scaling requires an accurate FEM model of the fan, a calibration of each strain gauge and a correction to account for the influence of FFT evaluation parameters as discussed in detail in [35, 36].

Figure 5 shows that at 90% speed vibration amplitudes rise continuously from $Φ = 0.52$, while they are negligible at design speed until a sharp rise due to the onset of rotating stall at $Φ = 0.489$. The amplitude evolution of N95 is comparable to the N90 case, but with increased amplitudes of mode 1 and 2 and decreased amplitudes of mode 3. This is attributed to the fact that rotating stall is observed prior to the occurrence of NSV signatures at this speedline, as will be discussed in detail below. To better understand the transition between subsonic and transonic operation and the related changes in observed phenomena, a comprehensive analysis of the operating points highlighted in Fig. 4 will be presented in section 3.2.

In a first step, wall pressure data which is usually used to derive the flow characteristics at the blade tip, is compared to SPIV measurements in section 3.1 for two different operating points at design speed. This allows to validate whether the footprint of the flow captured at the casing provides an accurate representation of the complex three-dimensional flow at the rotor tip. At the same time, SPIV measurements at highly throttled operating conditions did not yield satisfying results due to excessive condensation and oil deposition on the casing window in the area of interest and cannot be used for the analysis of all operating points highlighted in Fig. 4. Thus, the cross-validation provided in section 3.1 aims to prove that wall pressure measurements provide an adequate description of the aerodynamic field and are suitable for an in-depth analysis of blade tip aerodynamics. Particular focus is laid on the reso-

<figure><img src="image_6.png" /><figcaption>Figure 5: Evolution of average vibration amplitude for mode 1, 2 and 3 with flow coefficient $Φ$ from strain gauge data (scaled to modal scope limit, see [26]) for CATMIS configuration. Contours indicate mode shapes by means of normalized displacement amplitudes</figcaption></figure>

lution of aerodynamic asymmetries which are caused by geometrical blade-to-blade variations due to manufacturing inaccuracies (these geometrical variations are hence unintended and not correlated to the structural mistuning of the rotor). They can affect the behavior at throttled operating conditions as will be discussed in section 3.2.

# 3 Results

In the first part of this section, wall pressure measurements will be compared to SPIV measurements at design speed. The second part focuses on the investigation of the operating points marked in Fig. 4.

## 3.1 Cross-validation of measurement systems

For the validation of wall pressure data, SPIV measurements consisting of 96 phase angles were performed at several operating points to cover the whole rotor and investigate local flow variations. Two of those 360° measurements that were conducted at peak efficiency (PE) and peak pressure (PP) operating conditions at design speed are presented in Fig. 6 side-by-side with wall pressure measurements. On the left-hand side of each operating point, the ensemble averaged wall pressure field of 200 rotor revolutions $p_s^{\text{ens}}$ is compared to the relative in-plane velocity $c_{\text{rel}}$ (average of 100 images per phase angle). The right-hand side compares the local deviation of the respective quantity of one passage $\Delta(\cdot)^{1\text{pass.}}$ from the average
of all 16 blade passages to emphasize local (spatial) deviations of the flow field from the cyclic symmetry that would be observed for an ideal rotor. This quantity is referred to as aerodynamic asymmetry in the following.

<figure><img src="image_7.png" /><figcaption>Figure 6: Comparison of ensemble averaged wall pressure, rel. in-plane velocity and their respective passage-wise deviations from the all blade average at peak efficiency (PE, $\Phi = 0.631$) and peak pressure (PP, $\Phi = 0.521$) at 100% speed</figcaption></figure>

Previous studies presented in [25, 28] indicated based on wall pressure data that the highest aerodynamic asymmetry occurs at design speed for all ECL5 rotor configurations, reaching maximum deviations close to peak efficiency operating conditions. The operating points shown in Fig. 6 are hence considered suitable for the desired validation.

At peak efficiency conditions, the contours of relative in-

plane velocity $c_{rel}$ show a strong acceleration of the flow within the supersonic expansion region on the blade's suction sides before the flow is abruptly decelerated through the compression shock at approximately mid-chord. At this operating point, the shocks are already far detached from the leading edge of the trailing blade and significant blade-to-blade variations of the flow field become apparent. For some blades, for example 5 and 8, the acceleration zone is larger, resulting in higher relative velocities, while other blades, for example blades 6 and 9, show reduced velocities. The corresponding wall pressure measurements consistently provide low static pressures for the first and higher static pressures for the latter in the respective regions.

The good agreement of PIV and wall pressure measurements is further confirmed by the contours of aerodynamic asymmetry. For both measurements systems, highest deviations from the all-blade average are observed in the sector from blade 4 to blade 9 and mainly attributed to variations of the shock position and the maximum velocity in the first acceleration zone close to the leading edge. For the other blades in the sector from blade 10 to 14, variations are low.

This is also visible in Fig. 7, showing the shock distance (measured from the leading edge of the trailing blade) for each blade based on the measurements presented in Fig. 6. While strong variations of around ±5% chord are present within the blade package 4-9, variations are below ±1% chord between blade 10 and 14. The quantitative comparison of the PIV and wall pressure data in Fig. 7 further confirms the good agreement of both instrumentation systems despite different measurement locations (at the casing wall vs. at 93% channel height).

<figure><img src="image_8.png" /><figcaption>Figure 7: Comparison of shock distance from leading edge of trailing blade measured with WPTs and SPIV at peak efficiency at 100% speed</figcaption></figure>

SPIV measurements provide a detailed image of the flow structure in the blade tip region and allow to derive characteristic velocities at a given channel height and to analyze correlations with the blade tip geometry. As an example, Fig. 8 shows for peak efficiency conditions at design speed that tip clearance variations amount to approximately ±0.1 mm and ±0.05 mm at leading edge and mid chord, i.e. 10% of the respective nominal clearances<fnref n="3" />. The presented blade-to-blade patterns were found during repeated measurements on different operating days within a range of ±0.01 mm and hence considered as significant.

<footnote n="3">³with respect to mid-chord blade height these deviations correspond to approx. 0.1% and 0.05% respectively</footnote>
<figure><img src="image_9.png" /><figcaption>Figure 8: Tip clearance variation (leading edge, LE and mid chord, MC) at peak efficiency at 100 % speed</figcaption></figure>

For peak efficiency conditions at design speed, Fig. 9 indicates that the maximum velocity in the acceleration zone close to the leading edge as well as the maximum velocity directly upstream of the shock are correlated with individual tip clearances near the leading edge. For blades with high LE clearance, lower relative velocities are observed upstream of the shock which is attributed to increased tip leakage flow and hence blockage in the blade tip region. For the location of the compression shock as presented in Fig. 7, no significant correlation with stagger and tip-clearance could be identified and the subject remains in the focus of further studies. But a complex dependency with the local passage shape and tip blockage is assumed.

<figure><img src="image_10.png" /><figcaption>Figure 9: Correlation of velocity in LE acceleration zone and upstream of shock with leading edge (LE) tip clearance at peak efficiency at 100 % speed from SPIV measurements</figcaption></figure>

The study presented in [25] using the reference rotor indicated that circumferential flow variations decrease towards higher and lower massflow rates along the design speedline with respect to peak efficiency. This observation is confirmed by the PIV measurement at peak pressure conditions presented in Fig. 6 that provides a mostly uniform flow field. This is particularly noticeable in the deviation plot of relative in-plane velocity, which provides even lower asymmetry than the respective wall pressure contour. As for peak efficiency, main deviations are caused by variations of the shock position between neighboring blades and are consistently observed in wall pressure and PIV measurements. The comparison in Fig. 6 thus shows that wall pressure measurements yield a good representation of the flow field at slightly lower channel heights and can be effectively used for the investigation of

flow topology and aerodynamic asymmetry.

The presented results emphasize again, that the unintentional variation of tip clearance causes a systematic and stable variation of the flow field in the tip region. At peak efficiency this effect is most pronounced and tightly correlated with the leading edge clearance.

## 3.2 Investigation of throttled operating points

In this section, the operating points highlighted in Fig. 4 corresponding to different flow coefficients are analyzed in detail to characterize occurring phenomena and differences between the speedlines N90, N95 and N100.

### 3.2.1 $\Phi = 0.506$

At a flow coefficient of $\Phi = 0.506$, the flow field at 100 % speed is, similar to the flow field at peak pressure in Fig. 6, mostly uniform and fluctuations are comparably low.

At 95 % speed, in contrast, the wall pressure measurements indicate a significantly altered flow topology for blade 5 compared to its neighboring blades as visible in Fig. 10. PIV measurements of the respective blade passages revealed that no compression shock is present on the suction side of blade 5 in contrast to the surrounding blades. The analysis of the corresponding temporal fluctuations shown on the right-hand side of the figure illustrates that the absence of the shock is attributed to a flow separation on the respective blade which results in a significantly increased unsteadiness along the suction side. The separation is fixed to blade 5 and temporally stable, hence it is observed at every rotor revolution and different circumferential positions.

<figure><img src="image_11.png" /><figcaption>Figure 10: Relative velocity, ensemble averaged static wall pressure and related temporal fluctuations at $\Phi = 0.506$ at 95 % speed</figcaption></figure>

This observation could be confirmed throughout repeated measurements and can be explained with the stagger angle distribution shown in Fig. 11. As for tip clearance, the stagger
pattern is observed throughout repeated measurements within a range of ±0.04° and hence significant. Figure 11 highlights a stagger angle of blade 5 that is approximately 0.3° smaller than the average value, leading to increased incidence and thus flow separation on the blade's suction side. Despite this flow separation of a single blade, the fan operates stable and no significant performance drop can be noticed in Fig. 4.

<figure><img src="image_12.png" /><figcaption>Figure 11: Blade-to-blade variation of stagger angle at $\Phi = 0.506$ at 95% speed</figcaption></figure>

This separation of blade 5 is neither observed at 100% speed, nor at 90% speed. As visible from Fig. 5, vibration amplitudes are negligible for all speedlines at a flow coefficient $\Phi = 0.506$, corresponding wall pressure spectra do not show any distinct peaks and are omitted for brevity.

### 3.2.2 $\Phi = 0.489$

At a flow coefficient of $\Phi = 0.489$, the flow separation of blade 5 (I) is still present at 95% speed as presented by the ensemble averaged wall pressure field in Fig. 12(a). In contrast, the 90% speedline provides uniform flow conditions.

Another difference becomes apparent from Fig. 12(b). At 90% speed, zones of increased unsteadiness show a tip leakage flow (TLF) that is nearly perpendicular to the main flow and highest fluctuations are observed in the region where the TLF impinges on the pressure side of the trailing blade (II). At 95 and 100% speed the TLF is less inclined and highest unsteadiness is found in the region of the shock and in the zone where the shock interacts with the TLF (III).

To investigate these differences in more detail, Fig. 12(c) presents an average wall pressure spectrum of all probes in the leading edge plane. To emphasize non-synchronous pressure fluctuations an ensemble average has been subtracted before FFT calculation [17].

For N90, a broadband hump around EO10 indicates the presence of convective disturbances, often confusingly referred to as rotating instabilities in literature [13, 15, 24], that travel in the leading edge plane from one blade to another. However, at this operating point no coherent wave patterns are formed and no discrete peaks between EO6 and EO15 are visible. The absence of coherent waves is confirmed by the spectrum of propagation speed presented in Fig. 12(e), which is derived based on cross-correlations of unsteady pressure probes in different circumferential positions as described in detail in [17] and normalized with the maximum value of all presented measurements. The graph shows a broad peak around 57% rotor speed. Even if not organized around the circumference, the aerodynamic disturbances cause blade vibrations of different modes as visible in the average strain gauge spectrum in Fig. 12(d). Those observations are clear indicators for the onset of convective NSV.

<figure><img src="image_13.png" /><figcaption>Figure 12: Comparison of operating points at $\Phi = 0.489$ at 90% and 95% speed: (a) ensemble averaged wall pressure; (b) temporal fluctuations; (c), wall pressure spectra (LE plane); (d) strain gauge spectra (e) propagation speed spectra</figcaption></figure>

The four peaks below EO5 in the pressure spectrum of N90 indicate disturbances with an increased length-scale, i.e. reduced cell-count around the circumference. However, their amplitude and coherence is low at this throttle degree.

At 95% speed, no distinct peaks are observed in the wall pressure spectrum and vibration amplitudes for all three modes are slightly lower compared to the measurement at 90% speed. The most coherent propagation speed is observed
at 85% rotor speed indicating the presence of stall precursors as will be shown below. No NSV signatures are observed at this operating point.

At N100, a flow coefficient of $Φ = 0.489$ corresponds to the operating point just before the onset of rotating stall. The contour of temporal fluctuations indicates a TLF that is less inclined than at N90. Rotating stall is eventually initiated when spill forward occurs on one blade, leading to the formation of a spike that rapidly develops into a large cell of separated flow. However, just before the onset, no distinct peaks are visible in the wall pressure spectrum and vibration amplitudes of all modes are very low. Only the spectrum of propagation speeds indicates the presence of some stall precursors that propagate with a speed around 73% rotor speed.

### 3.2.3 $Φ = 0.461$

At a flow coefficient of $Φ = 0.461$, no data for the design speedline can be shown, since the fan stage always encountered an abrupt onset of rotating stall at $Φ = 0.489$ throughout repeated measurements. One to two stall cells are observed.

For 95% speed, the contour of temporal fluctuations in Fig. 13(a) reveals a further inclined TLF compared to the operating point at $Φ = 0.489$, and increasing unsteadiness close to the leading edge on the blade's pressure sides. However, highest unsteadiness is still observed in the region of the compression shock. These fluctuations are the result of the presence of rotating stall cells, which is clearly visible in the wall pressure spectrum. The dominant peak at EO2.25 = 3.075 together with the vibration peak at EO0.75 = |2.25 - 3| in Fig. 13(c) indicates the presence of three stall cells that propagate with speeds around $0.75Ω_r$ as visible in Fig. 13(d) according to Eq. 1 and 2. Analogously, the peaks at EO1.44 = 2.077 and EO3.03 = 4.076 correspond to stall cell patterns with 2 and 4 cells respectively, causing (low) off-resonant vibration peaks at EO0.56 = 1.44 - 2 and EO0.93 = 3.03 - 4 (not marked but visible next to the peak at EO0.75). The simultaneous presence of these peaks shows that the stall pattern is not stable in time but changes between spatial mode orders of 2, 3 and 4, with a dominant order of 3. Despite the presence of rotating stall, vibration amplitudes are still low compared to those which are observed in stall at 100% speed. Moreover, no significant performance drop is observed in Fig. 4. Thus, measurements could be continued, i.e. the machine further throttled, in contrast to the measurement at design speed where the emergency throttle was activated as soon as rotating stall was encountered to avoid safety critical blade vibration.

For the 90% speedline, Fig. 13 shows fully developed convective NSV characteristics. The contour of temporal fluctuations (a) yields increased unsteadiness, in particular close to the leading edge of the blade's pressure sides which has been related to convected radial vortex disturbances in the leading edge plane in previous work [15, 16]. The respective wall pressure spectrum shows several distinct peaks around EO8, corresponding to aerodynamic wave numbers between $N_a=11$ and 15. High vibration amplitudes are observed especially for mode 3 at EO5.09 due to a lock-in of an aerodynamic wave number $N_a = 12$ (corresponding pressure peak at EO6.91 = 12 - 5.09 according to Eq. 2) with an structural

<figure><img src="image_14.png" /><figcaption>Figure 13: Comparison of operating points at $Φ = 0.461$ at 90 and 95% speed: (a) temporal fluctuations; (b) wall pressure spectra (LE plane); (c) strain gauge spectra (d) propagation speed spectra</figcaption></figure>

nodal diameter of $N_v = 4 = N_b - N_a$, see Eq. 4. The coherence of propagating disturbances is underlined by Fig. 13(d) showing a single clear peak at a propagation speed of 57% rotor speed, which agrees well with the propagation speed observed at different speedlines for this rotor (see [26, 28]).

### 3.2.4 $Φ = 0.451$

At a flow coefficient of $Φ = 0.451$, the flow conditions change significantly for 90 and 95% speed compared to $Φ = 0.461$. For both speedlines, unsteady pressures presented in Fig. 14(a) are increased, especially in the region close to the
leading edge on the blade's pressure sides at N95 (1).

<figure><img src="image_15.png" /><figcaption>Figure 14: Comparison of operating points at $\Phi = 0.451$ at 90 and 95% speed: (a) temporal fluctuations; (b) wall pressure spectra (LE plane); (c) strain gauge spectra (d) propagation speed spectra</figcaption></figure>

This observation is supported by the spectra in Fig. 14(b-d). For 90% speed, the distinct peaks around EO8 are still present and result from aerodynamic disturbances that coherently propagate with a speed of approximately 58% rotor speed. However, additional peaks at EO1.31 = 2.066 and EO2.07 = 3.067 can be noticed in the wall pressure spectrum, indicating the presence of two to three stall cells that propagate slower compared to those at 95% speed at $\Phi = 0.461$ (but still faster compared to the small-scale disturbances of higher wave numbers). Nevertheless, the vibration

spectrum shows no traces of off-resonant excitation and no related peak is visible in the propagation speed spectrum. Mode 3 vibration amplitudes reach a plateau as visible from Fig. 5 despite the persistent lock-in of $N_a = 12 = 6.95/0.58$ and $N_v = 4$ (resulting vibration peak at EO5.06 = $|6.95 - 12|$). A slight reduction of the flow coefficient $\Phi$ even results in decreasing mode 3 vibration, while mode 1 amplitudes further increase. It is hence considered that superimposed large-scale disturbances in the form of part-span rotating stall cells reduce the coherence of the small scale disturbances, leading to reduced vibration of mode 3.

At 95% speed, the pressure peaks around EO2 corresponding to two to four propagating stall cells are still highest (and even higher than at $\Phi = 0.461$), with a dominant peak at EO1.37 = 2.076 (i.e. 2 cells) that causes off-resonant vibration at EO0.63 = $|1.37 - 2|$. However, a noticeable broadband hump around EO8 with several peaks emerges. This indicates, consistently with the increased unsteadiness on the blade's pressure sides in the leading edge region and the peak at EO0.58 in the propagation speed spectrum, the presence of coherently propagating small-scale disturbances and the onset of convective NSV alongside the presence of a multi-cell stall. Off-resonant excitation due to the stall cells is significant at this condition and leads to critical vibration amplitudes. The propagation speed spectrum shows a wide range of speeds, with a tendency towards higher values for longer wavelengths reaching 78% of rotor speed in the stationary frame.

## 4 Discussion & Conclusion

The findings emphasize a drastic change of occurring phenomena with the transition from transonic to subsonic operation:

1. At 100% speed, the rotor experiences rotating stall with one to two cells and spike-type inception. The onset of rotating stall is accompanied by a strong performance drop and safety critical vibration amplitudes that necessitate immediate measurement abortion.

2. At 95% speed, significant aerodynamic asymmetry in the form of a single separated blade is observed, long before rotating stall or NSV occurs. This blade-boundary layer separation is attributed to an unintentional misstagger of 0.3° and does not propagate in the rotor system.

At lower flow coefficients, a multi-cell (2-4 cells) rotating stall establishes that leads to moderate vibration amplitudes. The associated performance degradation is negligible. Further reducing the flow coefficient causes continuously increasing vibration amplitudes until safety limits are reached and the emergency bleed is opened. These vibrations are off-resonant with blade-eigenmodes and considered as non-synchronous forced-response.

3. At 90% speed, small-scale convective disturbances form coherent aerodynamic waves, one of which locks-in with the third blade eigenmode resulting in non-synchronous vibration amplitudes that continuously increase with throttle degree until measurements are stopped at a comparably low flow coefficient.
The presented study highlights that rotating stall and NSV have different wave characteristics as schematically illustrated in Fig. 15. Rotating stall is observed with wave numbers up to 4 that propagate with a speed of approximately 76-78 % rotor speed. As a result, frequencies in the rotating frame of reference are low and cause mainly off-resonant vibrations. Lock-in with eigenmodes is not observed.

In contrast, small-scale disturbances related to NSV are convected at a mostly constant speed of 57 % rotor speed and form patterns with higher wave numbers between 10 and 15. Thus, their frequency in the rotating frame is comparably high and can lock-in with a blade eigenmode. The intermediate range of wave-numbers between 5 and 9 that is in between NSV and rotating stall is not observed for the investigated fan.

Another difference is observed in Fig. 15 regarding the axial location of disturbances. The superimposed contours show an accumulation of unsteady pressure minima at $\Phi = 0.451$ at 90 % and 95 % speed to identify the path of aerodynamic disturbances. The left plot illustrates the higher wave number disturbances convecting slightly donwstream of the leading edges, whereas larger disturbances with higher wavelength propagate further upstream in the right-hand case.

These findings imply that prediction of NSV and rotating stall can only be successful if those characteristics are correctly captured, also considering that both types of disturbances might occur superimposed to each other.

<figure><img src="image_16.png" /><figcaption>Figure 15: Schematic illustration of wave characteristics of (a) NSV and (b) rotating stall with superimposed contours of accumulation of unsteady pressure minima at $\Phi = 0.451$ at 90 % (right) and 95 % (left) speed</figcaption></figure>

The results moreover emphasize that a significant asymmetry

of the aerodynamic field resulting from geometric non-uniformity in the blade tip region is present which will affect the local propagation characteristics of aerodynamic disturbances as was shown in [37]. Thus, numerical simulations that aim to investigate or predict these phenomena must consider non-periodic full-annulus simulations to capture aerodynamic asymmetry.

The instrumentation used has been effective in investigating the observed phenomena. The synchronized wall pressure arrays and bladewise vibration sensors enable the identification of modal interactions. Phase-locked PIV helps examine local aerodynamic asymmetry and allows correlation with geometric variations of tip-clearance and stagger, as measured by capacitive tip-timing. The data provided within this study forms the basis for the validation of numerical simulations at highly throttled operating conditions and for different phenomena.

## Acknowledgements

The results presented in this paper are the outcome of contributions from a large research group, and the authors gratefully acknowledge the excellent collaboration and support over the past five years. We extend special thanks to Sébastien Goguey and Cédric Desbois of LMFA and Kévin Billon and Claude Gibert of LTDS for their support and contributions to the experiments.

The open-test case ECL5 was developed in close collaboration with Safran Aircraft Engines. We are grateful for the continuous collaboration and financial support of SAFRAN since the beginning of this project and specifically for the present measurement campaign, for which the MARLYSA test module was provided by SAFRAN. For the tip-timing measurements we received extensive support of MTU Aero Engines.

The presented research was financed through the European Union's Clean Sky 2 Joint Undertaking (JU) under grant agreement N864719, CATANA. The JU receives support from the European Union's Horizon 2020 research and innovation program and the Clean Sky 2 JU members other than the Union. The paper reflects only the author's view and the JU is not responsible for any use that may be made of the information it contains. Assessment of the test facility was enabled through financial supports of Agence Nationale de la Recherche (ANR, Project d'EquipEx PHARE) and Conseil pour la Recherche Aeronautique Civile (CORAC - Programme CUMIN). Buildings and infrastructure were supported by ECL, instrumentation supported by Institut Carnot (INGENIERIE@LYON - Project MERIT) and SAFRAN Aircraft Engines.

## Nomenclature

<table><tr><td>$\dot{m}_{\text{std}}$</td><td>standard massflow rate [kg/s]</td></tr><tr><td>$c$</td><td>velocity [m/s]</td></tr><tr><td>$N$</td><td>wave number</td></tr><tr><td>$t$</td><td>tip clearance [mm]</td></tr><tr><td>$\lambda$</td><td>wave length [m]</td></tr></table>
Ω propagation speed [rad/s]

ω angular frequency [rad/s]

Ωr rotor speed [rad/s]

Φ flow coefficient

Πt total pressure ratio

σ standard deviation

σa, σv aerodyn. / structural inter-blade phase angle

a aerodynamic

rel relative

s static

t total

v vibration

R rot. frame of ref.

S stat. frame of ref.

ECL Ecole Centrale de Lyon

EO Engine Order

NSV Non-Synchronous Vibration

OGV Outlet Guide Vane

PE Peak Efficiency

PP Peak Pressure

RS Rotating Stall

SPIV Stereo Particle Image Velocimetry

TCS Turbulence Control Screen

UHBR Ultra-High-Bypass-Ratio

WPT Wall Pressure Transducer

## References

[1] K. B. Lee, M. Wilson, and M. Vahdati. "Numerical study on aeroelastic instability for a low-speed fan". In: *J. Turbomach*. 139.7 (2016). DOI: 10.1115/1.4035569.

[2] M. Vahdati and N. A. Cumpsty. "Aeroelastic instability in transonic fans". In: *J. Turbomach*. 138.2 (2016). DOI: 10.1115/1.4031225.

[3] R. E. Kielb, J. W. Barter, J. P. Thomas, and K. C. Hall. "Blade Excitation by Aerodynamic Instabilities: A Compressor Blade Study". In: *Proc. ASME Turbo. 2003*, pp. 399–406. DOI: 10.1115/GT2003-38634.

[4] S. Stapelfeldt and C. Brandstetter. "Non-synchronous vibration in axial compressors: Lock-in mechanism and semi-analytical model". In: *J. Sound and Vibration* 488 (2020), p. 115649. DOI: 10.1016/j.jsv.2020.115649.

[5] M. Inoue, M. Kuroumaru, T. Tanino, and M. Furukawa. "Propagation of multiple short-length-scale stall cells in an axial compressor rotor". In: *J. Turbomach*. 122.1 (2000). DOI: 10.1115/1.555426.

[6] A. Young, I. Day, and G. Pullan. "Stall warning by blade pressure signature analysis". In: *J. Turbomach*. 135.1 (2013). DOI: 10.1115/1.4006426.

[7] G. Pullan, A. Young, I. J. Day, E. M. Greitzer, and Z. S. Spakovsky. "Origins and structure of spike-type rotating stall". In: *J. Turbomach*. 137.5 (2015), p. 51007. DOI: 10.1115/1.4028494.

[8] T. R. Camp and I. J. Day. "A study of spike and modal stall phenomena in a low-speed axial compressor". In: *Proc. ASME Turbo*. 1997. DOI: 10.1115/97-GT-526.

[9] F. K. Moore. "A theory of rotating stall of multistage axial compressors: Part II - Finite disturbances". In: *J. Eng. Gas Turbines Power* 106.2 (1984). DOI: 10.1115/1.3239566.

[10] E. M. Greitzer. "Review: Axial compressor stall phenomena". In: *J. Fluids Engineering* 102.2 (1980). DOI: 10.1115/1.3240634.

[11] M. Inoue, M. Kuroumaru, S. Yoshida, T. Minami, K. Yamada, and M. Furukawa. "Effect of tip clearance on stall evolution process in a low-speed axial compressor stage". In: *Proc. ASME Turbo*. 2004. DOI: 10.1115/GT2004-53354.

[12] I. J. Day. "Stall, surge, and 75 years of research". In: *J. Turbomach*. 138.1 (2016). DOI: 10.1115/1.4031473.

[13] M. Baumgartner, F. Kameier, and J. Hourmouziadis. "Non-engine order blade vibration in a high pressure compressor". In: *12th Int. Symp. On Airbreathing Engines*. 1995. DOI: https://hal.science/hal-01353829.

[14] J. Thomassin, H. D. Vo, and N. W. Mureithi. "The tip clearance flow resonance behind axial compressor nonsynchronous vibration". In: *J. Turbomach*. 133.4 (2011). DOI: 10.1115/1.4001368.

[15] H. D. Vo. "Role of tip clearance flow in rotating instabilities and nonsynchronous vibration". In: *J. Propulsion and Power* 26.3 (2010). DOI: 10.2514/1.26709.

[16] C. Brandstetter, M. Juengst, and H.-P. Schiffer. "Measurements of Radial Vortices, Spill Forward, and Vortex Breakdown in a Transonic Compressor". In: *J. Turbomach*. 140.6 (2018), p. 061004. DOI: 10.1115/1.4039053.

[17] C. Brandstetter, X. Ottavy, B. Paoletti, and S. Stapelfeldt. "Interpretation of stall precursor signatures". In: *J. Turbomach*. 143.12 (2021), p. 13. DOI: 10.1115/1.4051709.
[18] F. Holzinger, F. Wartzek, M. Jüngst, H.-P. Schiffer, and S. Leichtfuss. "Self-excited blade vibration experimentally investigated in transonic compressors: Rotating Instabilties and flutter". In: *J. Turbomach*. 138.4 (2016). DOI: 10.1115/1.4032163.

[19] D. Franke, M. Jüngst, D. Möller, H.-P. Schiffer, and T. Giersch. "Influence of pre-swirl, rotor speed and blade count on aeroelastic coupling mechanisms during stall inception of a transonic compressor". In: *Proc. ASME Turbo*. 2020. DOI: 10.1115/GT2020-15255.

[20] V. B. C. de Almeida, E. Tüzüner, M. Eck, and D. Peitsch. "Numerical characterization of prestall disturbances in a compressor stage". In: *J. Turbomach*. 146.8 (2024). DOI: 10.1115/1.4064993.

[21] V. Hernley, A. Jemcov, J. Kang, M. Montgomery, and S. C. Morris. "Relationship between casing pressure and non-synchronous vibration in an axial compressor". In: *Int. J. Turbomach., Propulsion and Power* 9.2 (2024). DOI: 10.3390/ijtpp9020014.

[22] J. März, C. Hah, and W. Neise. "An experimental and numerical investigation into the mechanisms of rotating instability". In: *J. Turbomach*. 124.3 (2002). DOI: 10.1115/1.1460915.

[23] F. Kameier and W. Neise. "Rotating blade flow instability as a source of noise in axial turbomachines". In: *J. Sound and Vibration* 203.5 (1997). DOI: 10.1006/jsvi.1997.0902.

[24] R. Mailach, I. Lehmann, and K. Vogeler. "Rotating instabilities in an axial compressor originating from the fluctuating blade tip vortex". In: *J. Turbomach*. 123.3 (2001), pp. 453–460. DOI: 10.1115/1.1370160.

[25] A. Schneider, A.-L. Fiquet, B. Paoletti, X. Ottavy, and C. Brandstetter. "Experiments on tuned UHBR open-test-case fan ECL5/CATANA: Performance and aerodynamics". In: *J. Turbomach*. 146.8 (2024). DOI: 10.1115/1.4064231.

[26] A.-L. Fiquet, A. Schneider, B. Paoletti, X. Ottavy, and C. Brandstetter. "Experiments on Tuned UHBR Open-Test-Case Fan ECL5/CATANA: Stability Limit". In: *J. Eng. Gas Turbines Power* 146.5 (2023). DOI: 10.1115/1.4063717.

[27] C. Brandstetter, A. P. Schneider, A.-L. Fiquet, B. Paoletti, and X. Ottavy. "Experiments on structurally mis-tuned UHBR open-test-case fan ECL5/CATANA". In: *Proc. ASME Turbo*. 2024. DOI: 10.1115/GT2024-129163.

[28] A. P. Schneider. "Aerodynamic and aeroelastic investigation of a composite fan for ultra-high-bypass-ratio aircraft engines". In: *PhD thesis, Ecole Centrale de Lyon*. 2024. DOI: https://theses.fr/2024ECDL0018.

[29] A. Schneider, B. Paoletti, X. Ottavy, and C. Brandstetter. "Influence of a turbulence control screen on the aerodynamic and aeroelastic behavior of a UHBR fan". In: *Proc. ASME Turbo Expo 2022, GT2022-81961*. 2022. DOI: https://doi.org/10.1115/GT2022-81961.

[30] M. Zielinski and G. Ziller. "Noncontact Vibration Measurements on Compressor Rotor Blades". In: *Meas. Sci. Technol*. 11.7 (2000), pp. 847–856. DOI: 10.1088/0957-0233/11/7/301.

[31] M. Raffel, C. E. Willert, F. Scarano, C. J. Kähler, S. T. Wereley, and J. Kompenhans. *Particle Image Velocimetry: A practical guide*. Spinger, 2018. ISBN: 978-3-319-68852-7.

[32] B. Wieneke. "Stereo-PIV using self-calibration on particle images". In: *J. Experiments in Fluids* 39.1 (2005). DOI: 10.1007/s00348-005-0962-z.

[33] M. Voges, C. Willert, R. Mönig, and H.-P. Schiffer. "The effect of a bend-slot casing treatment on the blade tip flow field of a transonic compressor rotor". In: *Proc. ASME Turbo*. 2013. DOI: 10.1115/GT2013-94006.

[34] C. Brandstetter. "Aerodynamische Stabilisierung transsonischer Axialverdichter: Eine experimentelle Untersuchung der Blattspitzenströmung unter dem Einfluss von Gehäusestrukturierungen". In: *PhD thesis, Technische Universität Darmstadt*. 2015.

[35] C. Brandstetter, A. P. Schneider, A.-L. Fiquet, B. Paoletti, K. Billon, and X. Ottavy. "ECL5/CATANA: Comparative analysis of advanced blade vibration measurement techniques". In: *Proc. ETC16, accepted manuscript*. 2025.

[36] A. Schneider, B. Paoletti, X. Ottavy, and C. Brandstetter. "Quantification of blade vibration amplitude in turbomachinery". In: *Int. J. Turbomach., Propulsion and Power* 9.1 (2024). DOI: 10.3390/ijtpp9010010.

[37] P. Tharreau, M. Hardy-Falch, S. Stapelfeldt, and C. Brandstetter. "Non-synchronous vibration: Characterisation of the aerodynamic disturbance and its dependency on local tip clearance". In: *Proc. ASME Turbo*. 2024. DOI: 10.1115/GT2024-129186.