Experimental Investigation of the Aeroelastic Stability of an Annular Compressor Cascade at Reverse Flow Conditions

Compressor surge events are unsafe operating regimes yielding highly unsteady flow fields in which complex aeroelastic phenomena occur. If the blade flutter and forced response behaviour (i.e. aeroelastic stability) can be predicted reliably for normal flow conditions, its assessment at severe-off design conditions remains a critical task for compressor development programs. During the flow reversal sequence of a surge cycle, combined aerodynamic phenomena occur which make the accurate prediction of the unsteady forces acting on the blades difficult to assess. The main objective of this investigation is to increase the physical understanding of the unsteady phenomena present during the reverse flow sequence of a typical deep surge cycle. The analysis of the blade surface unsteady pressure distribution enables the identification of the main physical mechanisms present during such extreme flow operating conditions, as well as the evaluation of their contribution on the blade global aerodynamic stability. The approach adopted consists in performing aeroelastic investigations on an annular compressor cascade at established reverse flow conditions. The investigations are carried out at EPFL, in the annular test facility for non-rotating cascades. The cascade is forced to vibrate in a torsional travelling wave mode (controlled vibration). With an upstream swirled flow corresponding to real axial turbomachine conditions, a constant flow can be set in the test section. The steady-state operating conditions are measured upstream and downstream of the test section, using 5-hole aerodynamic probes. Several cascade blades are equipped with pressure taps at 50% span in order to acquire the steady-state and unsteady blade surface pressure distributions. Static pressure taps are also inserted in the casing wall of the test section to assess the steady-state flow field characteristics in the blade tip area. The inlet flow operating conditions are varied in order to determine their influence on the blade unsteady aerodynamic forces. This study presents the measurement results and analyses in details the aerodynamic response of a cascade subjected to controlled vibrations at reverse flow conditions. The data analysis is oriented towards both physical and practical approaches. In particular, the following features are addressed: Identification of the main unsteady physical mechanisms influencing the unsteady aerodynamic forces acting on a blade oscillating and subjected to reverse flow conditions. Determination of the influence of the inlet flow condition variations on these unsteady mechanisms. Evaluation of the contribution of each unsteady phenomenon to the blade aerodynamic stability (in terms of stabilizing or destabilizing impact). Assessment of the key parameters to control in order to minimize flutter risks in case that reverse flow conditions should occur. The data analysis reveals that during the reverse flow sequence of the surge cycle, blade interaction mechanisms play a major role in the blade aerodynamic stability. The global aerodynamic damping coefficient highlights this feature and indicates that aerodynamic instability exists for some operating conditions. A second unsteady phenomenon was detected, generated by the uncommon steady-state flow field characteristics present at reverse flow operating conditions. Within this frame, the presence of a large recirculation zone on the blade suction side was identified, influencing the blade aerodynamic stability. From a more general point of view, this study constitutes a step forward to the understanding of the blade loading processes occurring during a typical deep surge sequence. Results highlight the impact of the steady-state and unsteady phenomena on the blade loading level at reverse flow conditions. For one test case, the measured data was compared with numerical results, performed in parallel to the measurements. The results indicate that even though the agreement is reasonable, the correct prediction of the aerodynamic damping curve re- quires the consideration of a complex blade interaction mechanism. Within this frame, since not many experiments exist at reverse flow conditions, these experimental results are also a precious data source for CFD validation. They enable the improvement of the prediction/simulation accuracy of the compressor performance at off-design flow conditions

AEROELASTIC INVESTIGATION OF AN ANNULAR TRANSONIC COMPRESSOR CASCADE: NUMERICAL SENSITIVITY STUDY FOR VALIDATION PURPOSES

The accuracy of flutter or forced response analyses of turbomachinery blade assemblies strongly depends on the correct prediction of the unsteady aerodynamic loads acting on the vibrating blades. This paper presents the aeroelastic numerical results of an annular transonic compressor cascade subjected to harmonic oscillation conditions. The measurements associated were performed in an annular test facility for non-rotating cascades. The aim of this investigation is to get a deeper understanding of the specific characteristics of this test facility as well as improving the flutter prediction procedure and accuracy. For a subsonic and a transonic flow condition, the steady-state blade surface pressure distributions were predicted with two mesh configurations and results were compared to the experimental results. The first configuration omits the geometrical complexity of the experimental model and only models the blade passage. The second mesh configuration includes the cascade’s detailed geometry and cavities. The presence of leakage flows arisen due to the cascade’s slits and cavities are identified and their impact on the main flow field is analyzed and discussed. For the flutter computations, two mesh resolutions were investigated. The global damping predicted with a fine and a coarse mesh was compared, as well as the local pressure amplitudes and phases predicted with both configurations. Results show that even though similar global damping curves are predicted with both mesh resolution, for some IBPAs, local differences exist on the pressure amplitudes and phases. This highlights that only comparing the global damping coefficient, is not sufficient for code validation

NUMERICAL INVESTIGATION OF AN ANNULAR TRANSONIC COMPRESSOR CASCADE INCLUDING SECONDARY FLOWS

Within the frame of the European Project FUTURE, an annular transonic compressor cascade dedicated to aeroelastic experiments was investigated at Ecole Polytechnique Fédérale de Lausanne. As a first part of the experimental investigations and for different inlet flow conditions, the static pressure distribution on the airfoil surface was determined at three different blade spans (20%, 50% and 90%) by static pressure taps located on both blade’s pressure and suction side along the blade chord. Parallel to the measurements, DLR performed numerical flutter simulations to predict the blade aerodynamic stability. For the computational fluid dynamics procedure, the associated steady-state flow field was computed using two different numerical setups. The first one considers the blade channel only and omits the geometrical complexity of the experimental model. The second mesh corresponds to the real test model and includes the cavity under the blade hub, thus enabling the modeling of the secondary flows. For a subsonic and a transonic flow operating condition, the steady-state CFD computations performed with both meshes are compared to the blade surface static pressure distributions measured. The results show that including the cavity in the numerical model is required to determine the steady-state aerodynamic forces active on the blades with reasonable accuracy

NUMERICAL AND EXPERIMENTAL INVESTIGATIONS OF AN ANNULAR TRANSONIC COMPRESSOR CASCADE CONSIDERING LEAKAGE FLOWS

The accuracy of flutter or forced response analyses of turbomachinery blade assemblies strongly depends on the correct prediction of the unsteady aerodynamic loads acting on the vibrating blades. For unsteady linearized CFD solvers, the quality of the steady-state flow solution constitutes the basis for efficient and accurate CFD unsteady computations. This paper presents the steady-state numerical and experimental results of an annular transonic compressor cascade dedicated to aeroelastic investigations. The measurements were performed in an annular test facility for non-rotating cascades. For both a subsonic and a transonic flow condition, the steady-state flow field measured was simulated and the steady-state static pressure distributions measured on the blade surface were compared to the computational predictions. Especially for the transonic flow condition, results show that the nonlinear effect induced by the presence of a shock in the blade channel requires a detailed computational model taking into account the geometrical features of the experiment. The simulations were performed with two different meshing setups. For a first setup omitting the geometrical complexity of the experimental model and only including the blade passage, results highlight that the steady-state blade loading is not predicted correctly. With a second computational setup including the cascade’s detailed geometry, the relevant physics of the experiment are captured and the prediction accuracy is improved. The presence of leakage flows that arise due to several cascade’s slits and cavities was identified and their impact on the main flow field is discussed. For both flow regimes, the computational setup, the boundary conditions imposed as well as the characterization of the physics associated to the experiment are analyzed to define the optimal level of modeling required to ensure a robust mean flow solution to the initialization of the unsteady CFD procedure

Aerodynamic Damping Predictions During Compressor Surge: A Numerical Comparison Between a Half and Full Transient Approach

The prediction of the aerodynamic damping during compressor surge is a challenging task, because the flow is continuously evolving along the four surge cycle phases: pressurization (PR), flow-breakdown (FB), reversed flow (RF), and regeneration (RG), and complex flow conditions such as shocks and separations occur. Damping predictions with current existing methods typically consist of two steps. In the first step, a modified numerical model is used to simulate transient surge cycles. In the second step, damping analyses are performed for multiple timesteps along the surge cycle phases, which are then assumed as quasi-steady. The damping simulation can be performed using nonlinear or linear approaches. If shocks or separations occur, the latter yields inaccuracies in the flow and thus in the damping predictions. A new approach was developed to take into account and improve these inaccuracies. This new method includes the damping prediction within the transient surge simulation. Thus, all surge cycle phases and the continuously evolving flow conditions are considered, and nonlinear simulations are performed to account for shocks and separations. The results of this new method are presented and compared to the former method

Unsupervised spectral pattern recognition by Self Organizing Maps (SOM) analysis for site effects estimation: a review

A reliable determination of the unsteady aerodynamic loads acting on the blades is essential to predict the aeroelastic stability of vibrating compressor cascades with accuracy. At transonic flow conditions, the vibration of the shock may change the blade aeroelastic behavior. Numerical tools still have difficulties to capture the physics associated to this effect. In order to increase the prediction’s accuracy, high quality experimental data at high spatial resolution is therefore required to enable the calibration and validation of these tools. Within the frame of the European project FUTURE, experimental aeroelastic investigations were performed on a transonic compressor cascade in the Non-Rotating Annular Test Facility at EPFL. Associated to the measurements, the numerical flutter prediction procedure was applied. This paper focuses on the experimental results. The experimental database gained during the project is presented and aims at helping the aeroelastic community to develop and improve their flutter prediction capabilities. The test model consists of twenty prismatic blades. Each blade of the cascade assembly was mounted on an elastic spring element enabling harmonic bending vibrations in the twenty possible cascade’s travelling wave modes. Large efforts were made to improve the measuring techniques and to provide high quality data at relatively high spatial resolution. For various sub- and transonic flow conditions, steady-state and unsteady blade surface pressure distributions were measured to evaluate the local contributions to the blade stability in terms of local aerodynamic work. The blade global aerodynamic stability is determined applying an integration of all unsteady pressure signals measured over the airfoil

Applicability of Harmonic Balance for the Determination of Blade Stability within the Prescribed Motion Approach

To predict blade aerodynamic damping during the design phase, unsteady linearized CFD methods are commonly used as they offer a reasonable accuracy at acceptable computational costs. However, for moderate blade oscillation amplitudes, nonlinear aerodynamic effects may appear, yielding eventually an evolution into a stable, limit cycle oscillation (LCO). In the perspective of raising performance and safety, identifying such scenarios might open new development possibilities. Therefore, a valuable alternative to expensive CFD time domain methods consists in applying the nonlinear frequency domain harmonic balance (HB) approach to determine the aerodynamic response. An appropriate number of higher harmonics have to be retained depending on the severity of the aerodynamic nonlinearity under consideration. This number can be identified using either a convergence study with an increasing number of harmonics, or a direct comparison with time-domain simulations. For weak to moderate aerodynamic nonlinearities, this work proposes a guideline to determine the number of harmonics without additional, comparative simulations. First, the HB convergence properties are derived using the wellknown Duffing oscillator. Next, the method is applied to a compressor rotor blade subjected to a prescribed harmonic motion for conditions with and without aerodynamic nonlinearities