Effects of Non-Axisymmetric Volute on Rotating Stall in the Vaneless Diffuser of a Centrifugal Compressor

Rotating stall is an important unstable flow phenomenon that leads to performance degradation and limits the stability boundary in centrifugal compressors. The volute is one of the sources to induce the non-axisymmetric flow in a centrifugal compressor, which has an important effect on the performance of compressors. However, the influence of volute on rotating stall is not clear. Therefore, the effects of volute on rotating stall by experimental and numerical simulation have been explored in this paper. It’s shown that one rotating stall cell generates in a specific location and disappears in another specific location of the vaneless diffuser as a result of the distorted flow field caused by the volute. Also, the cells cannot stably rotate in a whole circle. The frequency related to rotating stall captured in the experiment is 43.9% of the impeller passing frequency (IPF), while it is 44.7% of IPF captured by three-dimensional unsteady numerical simulation, which proves the accuracy of the numerical method in this study. The numerical simulation further reveals that the stall cell initialized in a specific location can be split into several cells during the evolution process. The reason for this is that the blockage in the vaneless diffuser induced by rotating stall is weakened by the mainstream from the impeller exit to make one initialized cell disperse into several ones. The volute has an important influence on the generation and evolution process of the rotating stall cells of compressors. By optimizing volute geometry to reduce the distortion of the flow field, it is expected that rotating stall can be weakened or suppressed, which is helpful to widen the operating range of centrifugal compressors.

Assessment of the Loss Map of a Centrifugal Compressor’s External Volute

A volute’s loss coefficient is highly sensitive to Mach number, circumferential velocity and flow rate at volute inlet. In case of a backswept impeller, these parameters are coupled to each other. An increased flowrate leads to a steeper absolute flow angle at impeller exit and hence to a decrease of circumferential velocity. The absolute Mach number is also altered. Therefore, in order to investigate the effects of flowrate and flow angle separately, one would have to vary the diffuser width together with the flowrate, keeping the flow angle constant. This corresponds to coupling the volute with aerodynamically similar impellers, designed for higher and lower flowrates. Since this is elaborate, there is no adequate study available in open literature, assessing a volute’s global loss map. In this work, a new numerical approach for the prediction of a volute’s representative loss map is presented: The volute is calculated by means of steady CFD as a standalone component. The inlet boundary conditions are carefully selected by means of 1D and applied together with different diffuser widths. This allows for separate investigation of the impacts of flow angle, flow rate and Mach number. Validation against full stage CFD confirms the applicability of the standalone model. The results exhibit that minimum losses do not necessarily occur at the theoretical matching point but either when the volute is smaller or bigger, depending on the inlet flow angle. Investigations of the loss mechanisms at different operating conditions provide useful guidelines for volute design. Finally, the validity of these study’s findings for volutes with different geometrical features is examined by comparison with experimental data as well as with fullstage CFD.

Effect of the Condition Number of the Inverse Fourier Transform Matrix on the Solution Behavior of the Time Spectral Equation System

The time spectral method is a very popular reduced order frequency method for analyzing unsteady flow due to its advantage of being easily extended from an existing steady flow solver. Condition number of the inverse Fourier transform matrix used in the method can affect the solution convergence and stability of the time spectral equation system. This paper aims at evaluating the effect of the condition number of the inverse Fourier transform matrix on the solution stability and convergence of the time spectral method from two aspects. The first aspect is to assess the impact of condition number using a matrix stability analysis based upon the time spectral form of the scalar advection equation. The relationship between the maximum allowable Courant number and the condition number will be derived. Different time instant groups which lead to the same condition number are also considered. Three numerical discretization schemes are provided for the stability analysis. The second aspect is to assess the impact of condition number for real life applications. Two case studies will be provided: one is a flutter case, NASA rotor 67, and the other is a blade row interaction case, NASA stage 35. A series of numerical analyses will be performed for each case using different time instant groups corresponding to different condition numbers. The conclusion drawn from the two real life case studies will corroborate the relationship derived from the matrix stability analysis.

Low Rank Education of Cascade Loss Sensitivity to Unsteady Parameters by Proper Orthogonal Decomposition

In the present work, Proper Orthogonal Decomposition (POD) has been applied to a large dataset describing the profile losses of Low Pressure Turbine (LPT) cascades, thus allowing: i) the identification of the most influencing parameters that affect the loss generation; ii) the identification of the minimum number of requested conditions useful to educate a model with a reduced number of data. The dataset is constituted by the total pressure loss coefficient distributions in the pitchwise direction. The experiments have been conducted varying the flow Reynolds number, the reduced frequency and the flow coefficient. Two cascades are considered: the first for tuning the procedure and identifying the number of really requested tests, and the second for the verification of the proposed model. They are characterized by the same axial chord but different pitch-to-chord ratio and different flow angles, hence two Zweifel numbers. The POD mode distributions indicate the spatial region where losses occur, the POD eigenvectors provide how such losses vary for the different design conditions and the POD eigenvalues provide the rank of the approximation. Since the POD space shows an optimal basis describing the overall process with a low rank representation (LRR), a smooth kernel is educated by means of Least-Squares method (LSM) on the POD eigenvectors. Particularly, only a subset of data (equal to the rank of the problem) has been used to generate the POD modes and related coefficients. Thanks to the LRR of the problem in the POD space, predictors are low order polynomials of the independent variables (Re, f+ and ϕ). It will be shown that the smooth kernel adequately estimates the loss distribution in points that do not participate to the education. Additionally, keeping the same steps for the education of the kernel on another cascade, loss distribution and magnitude are still well captured. Thus, analysis show that the rank of the problem is much lower than the tested conditions, and consequently a reduced number of tests are really necessary. This could be useful to reduce the number of hi-fidelity simulations or detailed experiments in the future, thus further contributing to optimize LPT blades.

Experimental Investigation of Surge Phenomena in a Transonic Centrifugal Compressor With Inlet Distortion

Inlet distortion has influence on the aerodynamic performance of turbomachinery such as compressors, turbines and fans. On turbochargers, bent pipes are installed around the compressor due to the spatial limitations in the engine room of the vehicle. As the result, the compressor is operated with the distorted inflow. In the low flow rate operation, the distorted inflow also affects the flow instability like stall and surge. Especially, the operation range on the low flow rate side is defined based on the flow rate where surge occurs, so it is important to investigate the effect of the distorted inflow on surge. In this study, the effect of the inlet distortion to surge phenomena has been investigated by the experiments with a transonic centrifugal compressor. A bent pipe has been installed at the upstream of the compressor to generate a distorted flow. Experiments have been also conducted under the condition that a straight pipe was installed upstream of the compressor, and unsteady measurements with high response pressure sensors and an I-type hot wire probe have been carried out to each experiments. In addition, Fast Fourier transform (FFT) and Wavelet transform have been applied to the unsteady measurement results obtained from each experiment.

Impact of Multi-Row Aerodynamic Interaction on the Forced Response Behaviour of an Embedded Compressor Rotor

This paper focuses on the impact of multi-row interaction on the forced response behavior of an embedded compressor rotor at higher order modes. The authors in previous papers have discussed about the multi-row influence at the torsional mode resonant crossing and this paper extends the study to higher order modes. The paper talks about both the steady and unsteady influence of having additional rows in the configuration. It makes use of the time transformation (TT) method available in CFX to reduce the number of passages required in each row. Since the number of vanes from both the stators and the inlet guide vanes (IGV) is the same, the excitations from upstream rows and the potential field influence of the downstream row all contribute to the forcing, which is quantified both in terms of modal force and individual blade response. This paper describes the multi-row influence on the chordwise bending modes at both the peak efficiency (PE) and the high loading (HL) operating condition. To ascertain this influence, a 3-row case with just the two neighboring stators (S1, R2, S2 a 4-row case with the downstream rotor as well (S1, R2, S2, R3) and a 5-row with the upstream IGV were considered. While the 3-row case helped to determine the influence of neighboring stators on the forcing, the 4-row case provided the influence of the downstream rotor on the forced response behavior. Since the number of IGV vanes was the same as the neighboring stators the nature of interference between the stator and IGV wakes was determined as well. The 4-row case helped investigate physical wave reflections off a downstream rotating row, which had a significant influence on the modal force. The final section of the paper focuses on the mistuning response, which essentially couples frequency variations with the structural and aerodynamic aspects to predict individual blade responses, which are compared to experimental data. A mistuning analysis was carried out with the frequency mistuning present in the experimental facility Some of the key conclusions from this investigation are: 1) The interference of the IGV with the downstream stator (S1) is destructive at peak efficiency and constructive at high loading in line with the previous observation at torsional modes; 2) Physical wave reflections are constructive at all operating conditions at higher order modes unlike torsional modes where it was destructive; 3) The 3-row case gives the most accurate prediction in terms of average blade response and the 5-row case in terms of maximum blade response. Hence one of the significant findings is that, the aeromechanical behavior can be ascertained to a great deal of accuracy using just 3-rows at higher order modes crossings.

Time-Resolved Measurements of the Unsteady Boundary Layer in an Annular Low-Pressure Turbine Configuration With Perturbed Inlet

In this study the unsteady behavior of the boundary layers developing on a LPT stator profile and their effect on secondary flow patterns in a 1.5-stage turbine configuration are investigated under the influence of periodic inflow perturbations. The experimental setup previously employed to analyze the unsteady secondary flow in the stator wake has been enhanced by hotfilm sensor arrays placed on the stator profiles at different blade height positions to provide time-resolved data from within the passage. The turbine inflow is perturbed by periodically passing circular bars and a modified T106-profile has been considered for the blading. The modified profile, labeled as T106RUB, was developed for matching the transition and separation characteristics of the original T106 profile at low flow speeds, thus facilitating measurements to be taken in a large-scale test rig with its improved accessibility. The transition phenomena occurring in the profile boundary layers are investigated under both unperturbed and periodically perturbed inflow by means of spectral analysis, the semi-quantitative characterization of the wall-stress system and an evaluation of the statistic quantities. In particular, the periodic changes of the suction side boundary layer flow region towards the trailing edge are studied in detail. Furthermore, time-resolved hot-film measurements at different blade height positions facilitate a detailed comparison of the quasi two-dimensional mid-span profile flow and the near end wall profile flow which is subject to influence of secondary flow structures. These information are employed to assess to which extent the additional turbulence originating from the wakes affects the blade boundary layers and thus the secondary flow structures. Furthermore, the role of the perturbation frequency on the coupled system of boundary layers and secondary flow structures is evaluated.

Large Eddy Simulation of Wake-Shear Layer Interactions Over a Multi-Element Aerofoil

Modern commercial airliners use multi-element aerofoils to enhance take-off and landing performance. Further, multielement aerofoil configurations have been shown to improve the aerodynamic characteristics of wind turbines. In the present study, high resolution Large Eddy Simulation (LES) is used to explore the low Reynolds Number (Re = 0.832 × 104) aerodynamics of a 30P30N multi-element aerofoil at an angle of attack, α = 4°. In the present simulation, wake shed from a leading edge element or slat is found to interact with the separated shear layer developing over the suction surface of the main wing. High receptivity of shear layer via amplification of free-stream turbulence leads to rollup and breakdown, forming a large separation bubble. A transient growth of fluctuations is observed in the first half of the separation bubble, where levels of turbulence becomes maximum near the reattachment and then decay depicting saturation of turbulence. Results of the present LES are found to be in close agreement with the experiment depicting high vortical activity in the outer layer. Some features of the flow field here are similar to those occur due to interactions of passing wake and the separated boundary layer on the suction surface of high lift low pressure turbine blades.

Loss Predictions in the High-Pressure Film-Cooled Turbine Blade Cascade T120D by Mean of Wall-Resolved Large Eddy Simulation

Film cooling is commonly used to protect turbine vanes and blades from the hot gases produced in the combustion chamber. The design and optimization of these systems can however only be achieved if a precise prediction of the fluid mechanics and film efficiency is guaranteed at a level where induced losses are fully mastered. Such a prerequisite induces at the numerical level to be able to identify and assess losses. In this context, the present study addresses loss assessment in a wall-resolved Large Eddy Simulation (LES) of the film-cooled high-pressure turbine blade cascade T120D from the European project AITEB II. The objectives are twofolds: (1) to evaluate the capacity of LES to predict adiabatic film cooling effectiveness in a mastered academic case; and (2) to investigate loss generation mechanisms in a fully anisothermal configuration. When it comes to LES predictions of T120D, the flow structure around the blade and the coolant jet organization are coherent with literature findings. Satisfactory agreements are furthermore retrieved for the pressure load prediction as well as the adiabatic film effectiveness if compared to the experiment. Loss generation is then investigated illustrating the fact that aerodynamics losses dominate mixing losses which are mainly located in the coolant film. This is in line with the temperature difference between the hot and coolant flows that is low for this experimental condition. Distinct contributions can however be made available by studying the local loss generation maps by means of Second Law Analysis if recast in the specific context of anisothermal flows when simulated by LES.

Criteria for the Stability Limit Prediction of High Speed Centrifugal Compressors With Vaneless Diffuser: Part I — Flow Structure Analysis

High-speed centrifugal compressor requirements include a wide operating range between choking and stall especially for turbocharging applications. The prediction of the stability limit at different speeds is still challenging. In literature, several studies have been published on the phenomena that trigger the compressor instability. However, a comprehensive analysis of criteria that can be used in the first steps of centrifugal compressors design to predict the stability limit is still missing. In previous work the authors have already presented a criterion, so called “Stability Parameter”, to predict the surge line of centrifugal compressors based on a simplified CFD approach that does not require excessive computational resources and that can be efficiently used in the preliminary design phases. The above methodology has demonstrated its accuracy for centrifugal compressors with vaned diffuser, but a lower accuracy has been detected for vaneless diffusers. Before proceeding to identify additional criteria focused on compressors with vaneless diffuser, an in-depth fluid dynamics analysis has been necessary. This analysis has been also carried out through fully 3D unsteady simulations to allow identifying the real phenomena linked to the trigger of the instability of centrifugal compressors. It has been found how these phenomena are strongly related to the rotational speed, in particular have been shown the key role of the volute at high rotational speed.