Interblade Phase Angle Research Articles

Spanwise Variations in Aerodynamic Damping of an Oscillating Annular Compressor Cascade

Aeroelastic instability in turbomachinery is one of the severe problems limiting its performance. Of particular interest is the phenomenon of stall flutter in compressors, which occurs when blades encounter large incidence flows. With the objective to understand the aerodynamics associated with stall flutter, the present experimental study is conducted on a subsonic annular compressor cascade. Five consecutive blades are oscillated, with the central blade instrumented for unsteady surface pressure measurements. The flow incidence is established using inlet guide vanes to provide three inflow profiles, representative of nominal, near-stall, and poststall incidences. With the blades oscillating at specific interblade phase angles and reduced frequencies, the stability of the cascade is determined globally and locally in terms of aerodynamic damping. The near-hub flow is distinctive due to the thickened boundary layer, resulting in an irregular variation of pressure with low magnitudes, eventually approaching negative stability. At larger incidences, due to the low-amplitude fluctuations in pressure at all spans, the overall stability is reduced. The present work also describes and rationalizes the nature of the dependence of blade unsteady pressure distribution on reduced frequency and incidence.

Numerical simulations of flutter mechanism for high-speed wide-chord transonic fan

Flutter is a self-excited aeroelastic instability phenomenon that can cause blade failure and lead to safety accidents. This paper investigates the first bending mode flutter behavior of a high-speed wide-chord transonic fan under different operating conditions. The aim is to reveal the relationship between the flow structure and blade flutter stability at different inter-blade phase angles (IBPA). Two methods are used to model the traveling wave propagation along the circumferential direction: the influence coefficient method (ICM) and the traveling wave method (TWM). The former is used to calculate the aerodynamic damping at different IBPA with fewer computational resources, and the latter is used to analyze the relationship between the flow and flutter stability. The results suggest that the most unstable behaviors occur with two nodal diameters (ND) at three representative conditions: near stall, peak efficiency, and choke. The aerodynamic damping distribution of the blade surface is concentrated primarily at the shock wave, separation, and large vibration amplitude regions, whose effects on the blade flutter stability are greatly impacted by the IBPA. When the blade IBPA is small and positive, the shock and separation areas in the suction side reduce the blade flutter stability. This shows that the IBPA strongly affects the unsteady pressure fluctuations, including the direction and shape of the pressure wave propagation, and whether the unsteady pressure can propagate downstream (cut-on) and result in different aerodynamic damping distributions on the blade surface.

Validation Studies of Linear Oscillating Compressor Cascade and Use of Influence Coefficient Method

Abstract Advanced predictions of blade flutter have been continually pursued. It is noted however that validation cases of unsteady CFD methods against experimental cases with detailed 3D unsteady pressures are still rather lacking. The main objectives of the present work are two-folds. First, validate and understand the characteristics of blade tip clearance, as well as a bubble-type flow separation for an unsteady CFD solver against a 3D oscillating cascade experiment. And second, examine the applicability of the influence coefficient method (ICM) as widely used in an oscillating linear cascade setup. In the first part, the capability of a widely used commercial solver (CFX) for unsteady flows induced by a 3D oscillating compressor cascade is examined. The present computations have shown consistently a destabilizing effect of increasing blade tip clearance, in agreement with the experiment. More remarkably, the computational analyses reveal a distinctive interplay between the inlet endwall boundary layer and the tip clearance in relation to the aerodynamic damping. Different inlet endwall boundary layer thicknesses are shown to lead to qualitatively different aeroelastic stability characteristics in relation to tip clearance. The aero-damping variation with the tip clearance under the influence of the inlet endwall boundary layer seems to correlate closely to a balancing act between the passage vortex and the tip leakage vortex. The tip clearance aeroelastic behavior seems also in line with a simple quasi-steady analysis. On the other hand, the mid-chord laminar bubble separation on suction surface, though with a clear signature in the local aero-damping, has negligible effects on the overall stability. The second part aims to examine computationally the applicability of the influence coefficient method in a linear cascade setup. The comparison between the cascade-based ICM data and a baseline “tuned cascade” shows that the differences in the sensitivity to the far-field treatment can be significant, depending on inter-blade phase angles. On the other hand, non-linearity effects closely relevant to the basic linear assumption of the ICM are shown to only have a small influence. The present results suggest that extra caution should be exercised when comparing a CFD-based tuned cascade model with a finite cascade-based ICM model, at conditions close to acoustic resonance. The resultant discrepancies may well arise from the inherently different far-field sensitivities between the two models, rather than those typical numerical and physical modeling aspects of interest (e.g., meshing, spatial and temporal discretization errors as well as turbulence modeling).

Efficient prediction of classical flutter stability of turbomachinery blade using the boundary element type numerical method

In this paper classical flutter phenomena in power turbine rotor is studied. A medium fidelity 2D flow solver based on boundary element method is developed for this purpose. The classical flutter parameters in turbomachinery cascades such as aerodynamic damping at different inter-blade phase angle are estimated using flutter stability analysis flow solver. The flow solver is developed using 2D unsteady potential flow based panel method. The unsteady aerodynamic loading on the vibrating cascade is estimated using a traveling-wave mode oscillation. The simulated aerodynamic damping and pressure coefficient using boundary element method are compared against both experimental data and the computational fluid dynamics model’s results at different flow conditions. The boundary element based method results demonstrate good agreement with experimental data. The boundary element based flow solver shows significant reduction in computational time compared to computational fluid dynamics model.

Effect of acoustic treatment on fan flutter stability

This paper presents an investigation of the effect of acoustically treated wall on fan flutter stability. The analysis is based on a three-dimensional analytical model which captures the aerodynamic coupling between oscillating annular rotor and finite-length liner in subsonic flow. With the response function of the entire system constructed as a whole, the perturbed duct field in modal description is determined by a simultaneous solution of the rotor and liner responses. Whether or not the aeroelastic instability occurs under a given wall impedance condition is evaluated through the energy method. Numerical experiments are conducted in the parametric space of inter-blade phase angle, liner distribution and acoustic impedance. The results show that under different conditions the acoustically treated wall can take diverse effects on the fan flutter stability. In the presented cases where the rotor stability is considerably affected, it is found that the aerodynamic interaction between the rotor and the lined section modify the overall distribution of the aerodynamic loadings on blade surface rather than change their magnitudes significantly. Moreover, the aeroelastic response of the oscillating blades turns out to be very susceptive to the reactance variation when the resistance is small, further indicating the important role of the reflections on the liner surface in such aerodynamic interaction. The analysis in general implies that the actual impact of a lined wall on the fan flutter stability depends on the aerodynamic excitation condition mutually determined by duct geometry, rotor state and configurations of liner.

NUMERICAL ANALYSIS OF THE TURBOMACHINE BLADE ROW AEROELASTIC OSCILLATIONS WITH TAKING INTO ACCOUNT THE DISC DEFORMATION

The unsteady phenomena caused by blades oscillations by action of forced forces are characterized with energy change between gas flow and oscillating blades and formulate the base of physical mechanism of self-excited oscillations that can or to attenuate (aerodamping), or to be displayed in stable form of autooscillations, or in unstable form of flutter, which can activate to the structure destruction. Therefore aeroelastic blades behaviour represents the important problem of reliability and safety of gas and steam turbines with high aerodynamic indicators and high loaded blades. One of approaches to increase the stable blades oscillations is detuning of natural forms bound to disc deformation. There presented the numerical analysis of effect of disc deformation on aeroelastic blades oscillations of turbomachine blade row. The disc deformation is characterized by disc nodal diameters number that defines the interblade phase angle of blades oscillations shift (IBPA), and impacts on unsteady aerodynamic loads and blades oscillations amplitudes. In paper there shown that decrease of IBPA causes to increase of aeroelastic stability that is to reduction of blades oscillations amplitudes. The proposed numerical method of coupled aeroelastic problem solution for threedimensional transonic ideal gas flow allows to predict aeroelastic behaviour of blades including the forced, self-excitation oscillations and autooscillations with purpose to increase the efficiency and reliability of turbomachines blades devices.

Aerodynamic Damping Prediction for Turbomachinery Based on Fluid-Structure Interaction with Modal Excitation

Aerodynamic damping predictions are critical when analyzing aeroelastic stability. A novel method has been developed to predict aerodynamic damping by employing two single time-domain simulations, specifically, one with the blade impulsed naturally in a vacuum and one with the blade impulsed in flow. The focus is on the aerodynamic damping prediction using modal excitation and the logarithmic decrement theory. The method is demonstrated by considering the first two bending modes with an inter-blade phase angle (IBPA) of 0° on a transonic compressor. The results show that the flutter boundary prediction is basically consistent with the experiment. The aerodynamic damping prediction with an IBPA of 180° is also performed, demonstrating that the method is suitable for different traveling wave mode representations. Furthermore, the influence of the amplitude of modal excitation and mechanical damping using the Rayleigh damping model for aerodynamic damping was also investigated by employing fluid-structure coupled simulations.

Detached-Eddy Simulation Applied to Aeroelastic Stability Analysis in a Last-Stage Steam Turbine Blade

Blade flutter in the last stage is an important design consideration for the manufacturers of steam turbines. Therefore, the accurate prediction method for blade flutter is critical. Since the majority of aerodynamic work contributing to flutter is done near the blade tip, resolving the tip leakage flow can increase the accuracy of flutter predictions. The previous research has shown that the induced vortices in the tip region can have a significant influence on the flow field near the tip. The structure of induced vortices due to the tip leakage vortex cannot be resolved by unsteady Reynolds-averaged Navier–Stokes (URANS) simulations because of the high dissipation in turbulence models. To the best of author’s knowledge, the influence of induced vortices on flutter characteristics has not been investigated. In this paper, the results of detached-eddy simulation (DES) and URANS flutter simulations of a realistic-scale last-stage steam turbine are presented, and the influence of induced vortices on the flutter stability has been investigated. Significant differences for the predicted aerodynamic work coefficient distribution on the blade surface, especially on the rear half of the blade suction side near the tip, are observed. At the least stable interblade phase angle (IBPA), the induced vortices show a destabilizing effect on the blade aeroelastic stability. The motion of induced vortices during blade oscillation is dependent on the blade amplitude, and hence, the aerodynamic damping is also dependent on the blade vibration amplitude. In conclusion, the induced vortices can influence the predicted flutter characteristics of the steam turbine test case.

Experimental Investigation of an Aerodynamically Mistuned Oscillating Compressor Cascade

In this paper, the effect of aerodynamic mistuning on stability of a compressor cascade is studied. The experiments have been carried out at a low-speed test facility of the Technische Universität Berlin. The test section contains a linear cascade with compressor blades that are forced to oscillate in sinusoidal pitching motion. The aerodynamic mistuning is realized by a blade-to-blade stagger angle variation, and three mistuning patterns have been investigated: one-blade mis-staggering, alternating mis-staggering, and random mis-staggering. Mis-staggering can have a stabilizing or destsabilizing effect, but depends strongly on the amount of detuning that alters the flow passage. For positive stagger angle variation for the one-blade and alternating mis-staggering, the trend of the damping curve was maintained, in the sense that the unstable interblade phase angles (IBPAs) remained unstable. For negative stagger angle variation, one IBPA shifted from stable to unstable. For the random pattern, only very moderate changes are observed. The cascade stability was not noticeably affected by the aerodynamic mistuning.

Influence of the Tip Clearance on the Aeroelastic Characteristics of a Last Stage Steam Turbine

In this paper, the tip clearance effects on the aeroelastic stability of a last-stage steam turbine model are investigated. Most of the unsteady aerodynamic work contributing to flutter of the long blades of the last-stage of a steam turbine is done near the tip of the blade. The flow in this region is transonic and sensitive to geometric parameters such as the tip clearance height. The KTH Steam Turbine Flutter Test Case was chosen as the test case, which is an open geometry with similar parameters to modern free-standing last-stage steam turbines. The energy method based on 3D URANS simulation was applied to investigate the flutter characteristics of the rotor blade with five tip gap height varying from 0–5% of the chord length. The numerical results show that the global aerodynamic damping for the least stable inter-blade phase angle (IBPA) increases with the tip gap height. Three physical mechanisms are found to cause this phenomenon. The primary cause of the variation in total aerodynamic damping is the interaction between tip clearance vortex and the trailing edge shock from the adjacent blade. Another mechanism is the acceleration of the flow near the aft side of the suction surface in the tip region due to the well-developed tip leakage vortex when the tip clearance height is greater than 2.5% of chord. This causes a stabilizing effect at the least stable IBPA. The third mechanism is the oscillation of the tip leakage vortex due to the blade vibration. This has a negative influence on the aeroelastic stability.

A Low-Speed Compressor Test Rig for Flutter Investigations

A test facility for aeroeolastic investigations has been set up at the chair of Aero Engines at the Technische Universität Berlin. The test rig provides data for tool and code validation, and is used for basic aeroelastic experiments. It is a low-speed wind tunnel, which allows free and controlled flutter testing. The test section contains a linear cascade with eleven compressor blades. Nine of them are elastically suspended. The paper presents a detailed description of the test facility results to evaluate the overall flow quality alongside an aeroelastic model to predict the flutter velocity and critical interblade phase angles (IBPAs). Furthermore, chordwise pressure distributions, measured with traveling wave (TW) mode experimental tests, are presented. These measurements have been carried out for a wide range of IBPAs and have been compared to numerical results. Hot-wire anemometry has been applied to examine the inlet flow for several Mach numbers and Reynolds numbers. The results show small turbulence intensities. The blade surface pressure distribution and the flow field of the blade's suction and pressure sides have been obtained by oil flow visualization.

Experimental Study of Aerodynamic Damping of an Annular Compressor Cascade With Large Mean Incidences

This study addresses flutter that can occur in compressors when operating at high relative incidence. Studies are performed on a subsonic annular compressor cascade containing a sector of blades that can be subjected to controlled torsional oscillation. Measurements taken on the centrally located blade are used to study the unsteady surface pressures developed. Three large mean incidences are considered to characterize the aeroelastic performance. Aerodynamic damping is calculated for each test condition and its variation due to interblade phase angle (IBPA), reduced frequency, and incidence is studied. The source of stability or instability is traced to the nature of unsteady pressures. When the incidence is below the static-stall limit, an increasing incidence is found to exhibit aeroelastically more stable performance, whereas beyond the limit, the stability worsens. For the latter, the amount of improvement in stability by increasing reduced frequency is less compared to the former and its variation with IBPA is not as regular owing to the associated large uncertainty. The nonlinearity effects were found to be relatively higher for this case, especially from the aft region of the suction surface. It is also found that the phase of the local fluctuating pressure and its location on the chord has a decisive influence on the aerodynamic damping and its trends with respect to various parameters are discussed. The results are expected to be useful in the assessing aerodynamic damping trends in relation to the pressure phase variations in specific regions along the chord.

Consequences of Borescope Blending Repairs on Modern High Pressure Compressor Blisk Aeroelasticity

Objective of this paper is to analyze the consequences of borescope blending repairs on the aeroelastic behavior of a modern high pressure compressor (HPC) blisk. To investigate the blending consequences in terms of aerodynamic damping and forcing changes, a generic blending of a rotor blade is modeled. Steady-state flow parameters like total pressure ratio, polytropic efficiency, and the loss coefficient are compared. Furthermore, aerodynamic damping is computed utilizing the aerodynamic influence coefficient (AIC) approach for both geometries. Results are confirmed by single passage flutter (SPF) simulations for specific interblade phase angles (IBPA) of interest. Finally, a unidirectional forced response analysis for the nominal and the blended rotor is conducted to determine the aerodynamic force exciting the blade motion. The frequency content as well as the forcing amplitudes is obtained from Fourier transformation of the forcing signal. As a result of the present analysis, the change of the blade vibration amplitude is computed.

Effect of Front and Rear Rotor Stages on Aeroelasticity in Multi-Stage Environment

An energy method based on the mixing-plane model and phase lagged boundary condition has been developed to estimate the flutter characteristics of rotor blades in multi-stage environment. The effects of front and rear rotor stages on the aerodynamic damping of the rotor blades have been investigated using this method. The results show that the mixing-plane model enables to consider the averaging effect of the other stages on the aeroelasticity of the checked rotor blades without having to perform the unsteady full annual multi-stage (FAMS) flow computations. Comparing with the isolated rotor blade, the upstream and downstream rotor stages have a significant influence on the aeroelasticity of the rotor blade with altering the intensity and location of the shock wave and separation flow region on suction surface. It is worth to point out that the neighbor rotor stages reduce the effect of the inter-blade phase angle (IBPA) on the aerodynamic damping. Moreover, the impact of the rear rotor stage on aerodynamic damping of the rotor blade is more remarkable than that of the front one. Compared to the measured data, the capability of this method used in the aeroelasticity assessment of a multi-stage turbomachine has been validated. Furthermore, the relationship between the aerodynamic damping and the motion of the shock wave has been revealed, which can assist the compressor design.

Dielectric Barrier Discharge Plasma Actuator for Load Alleviation and Instability Control in a Compressor Cascade

The aim of this study is to investigate the performance of dielectric barrier discharge plasma actuators (DBD-PAs) applied to control the aeroelastic response of a compressor cascade in subsonic flow conditions. Simulations involve a cascade composed by 7-blade, 3 of which show a not-synchronous pitching behaviour. Two different inter blade phase angle (IBPA) have been tested and compared. Actuators capabilities in terms of load alleviation and instability control have been evaluated through the measurement of mean value, standard deviation and hysteresis area of the airfoil response in terms of lift and moment coefficients.

A semi-analytical method for oblique gust - cascade interaction

A semi-analytical method is proposed for calculating the response of a linear cascade with spanwise mean flow subject to oblique gusts. It is developed based on the boundary value problems defined by Lloyd and Peake (AIAA paper 2008–2840). A gust strength parameter is introduced and the correct three-dimensional (3-D) response is obtained. The classic similarity rules are extended; the approach can be used to extend any 2-D methods to account for oblique gusts and 3-D mean flows. It is validated against analytical approximations for single-airfoil and cascade responses. The method is used to investigate the effect of gust angle αg on the unsteady lift and the sound field. It is found that as αg increases, the 2-D equivalent response varies slightly. However, the 3-D lift is amplified by factor 1/cosαg, and the spanwise phase variation increases. Cascade effects are also studied. The inter-blade phase angle (IBPA) is important even for very low solidity. As the solidity increases, the chordwise distribution of lift is no longer leading-edge dominant. Cascade effects are small only when the cascade blade count is lower than a limit. A statistics analysis reveals that Mach number is the most important parameter for determining this blade count limit, and frequency is the least important.

Flutter of long blades in a steam turbine

Cases of flutter appearing in the last stages of LP steam turbines have recently been reported. This paper presents a numerical analysis of flutter in the last stage of LP steam turbine blades using in-house, non-viscous and viscous codes as well as the ANSYS CFX viscous code. In the case of harmonic blade oscillations, instability regions were calculated using an aeroelastic damping coefficient for various interblade phase angles in five mode shapes. The distribution of this coefficient was also calculated along the blade length. It was found that flutter appears in the first mode shape. 3D viscous and non-viscous calculations showed that flutter appeared only in the first harmonic oscillation mode. In the other four modes, flutter did not appear. A negative aerodamping coefficient for the harmonic blade oscillations is not sufficient evidence of flutter. Conclusive evidence may be obtained using the two-way fluid–structure interaction procedure without assuming harmonic oscillation. In practice, the last stage LP rotor blades are mistuned and the pressure behind them is not uniform. Analysis of such conditions has to be done.

Numerical Assessment of Virtual Control Surfaces for Load Alleviation on Compressor Blades

Virtual control surfaces for the optimization of steady and unsteady airloads on a compressor cascade are assessed numerically. The effects of mechanical surfaces are realized with plasma actuators, located both on the pressure and on the suction side of the blade trailing edge. Suction side plasma actuation is thought to reproduce the effects of mechanical wing spoilers, whereas pressure side plasma actuation is meant to act as a mechanical Gurney flap. Indeed, actuators are operated to generate an induced velocity field that is opposite relative to the direction of the freestream velocity. As a consequence, controlled recirculating flow areas are generated, which modify the effective mean line shape, as well as the position of the Kutta condition application point—and in turn the developed airloads. Proper triggering of pressure/suction side actuation is found to be effective in altering the blade loading, with effects comparable to those of mechanical control surfaces. Traveling wave mode simulations show that significant reductions in the peaks of the blade pitching moment can be achieved on the whole spectrum of interblade phase angles. It is proved that virtual control surfaces can provide effective load alleviation on the cascade, with potential remarkable reduction of fatigue phenomena.

On the effect of moving blade grid on the flow field characteristics

The motivation of this paper is the continual development of the blades for the last stage of the steam turbine. The biggest problem is the slenderness of such blades and the extreme sensitivity to aeroelastic vibrations (flutter) caused by the instabilities present in the flow. This experimental research is dealing with the aeroelastic binding of the moving blades located in the blade grid with the flow field and vice versa. A parallelogram is used to ensure one order of freedom of the blade. The grid has five blades in total, three of them are driven by force control using three shakers. The deviation as well as force response is measured by strain gauges and dynamometers. The flow field statistical as well as dynamical characteristics are measured by optical method Particle Image Velocimetry. The grid is connected to the blow-down wind tunnel with velocity range up to 40 m/s. The aeroelastic binding is investigated in dependency on used actuation frequency and maximal amplitude (the intensity of force actuation) and on different Reynolds numbers. The flow field and the wake behind each individual blade are studied and the maximal interaction is examined for individual inter-blade phase angle of the grid.

Influence of the Tip Clearance on a Compressor Blade Aerodynamic Damping

The flutter CFD calculations conducted in industry generally use a single-row, single-passage approach with a minimum of detailing to reduce the computational time as much as possible. The question arises of what level of detailing is necessary for accurate aerodamping predictions. In this study, the influence of tip clearance size on the steady flowfield and aerodynamic damping is assessed for a high-pressure compressor second-stage rotor. The steady analyses are conducted on an isolated blade passage by imposing the experimental boundary conditions at the inlet and outlet of the passage. Two operating points are considered: 74 and 100% corrected speed. It is observed that, at the same pressure ratio, the mass flow decreases with a larger tip gap due to the tip clearance flow rolling into a vortex on the blade suction side. Then, unsteady analyses are conducted by enforcing the first torsion mode shape at the corresponding natural frequency. For all interblade phase angles, the same trend is observed: the damping first increases (stabilizing effect) with the tip clearance before reaching a maximum and dropping for the larger gaps. Such changes in damping would have a significant impact on both flutter and forced response predictions.