Multi-bladerow fan forced response predictions using an integrated three-dimensional time-domain aeroelasticity model

2000 ◽  
Vol 214 (12) ◽  
pp. 1467-1483 ◽  
Author(s):  
A I Sayma ◽  
M Vahdati ◽  
M Imregun
Keyword(s):  
Unsteady Flow ◽  
Stokes Equations ◽  
Pressure Ratio ◽  
Forced Response ◽  
Rotor Blades ◽  
Time Step ◽  
Blade Vibration ◽  
Inlet Distortion ◽  
Non Linear ◽  
Good Agreement

A non-linear integrated aeroelasticity system to predict the forced vibration response of aero-engine fans is presented in this paper. The computational fluid dynamics (CFD) solver, which uses Favre-averaged Navier-Stokes equations on unstructured grids of mixed elements, is coupled to a modal model of the structure so that the effects of blade flexibility can be accommodated. The structural motion due to unsteady fluid forces is computed at every time step and the flow mesh is moved to follow the structure so that the resulting flow unsteadiness is determined in a time-accurate fashion. Two fan forced response case studies are reported in detail. The first one deals with a high-pressure ratio fan, the excitation being due to the upstream variable-angle inlet guide vanes (VIGVs). The unsteady flow analysis with blade motion was conducted using a sector of three VIGVs and four rotor blades. The wake predictions were found to be in good agreement with the corresponding laser measurements. The flow was observed to be completely separated for high VIGV angles and the excitation encompassed several harmonics. The predicted rotor blade vibration levels were generally found to be within 30 per cent of the measured values. The forced response to upstream obstructions was studied in the next fan case study. Three whole bladerows, consisting of 11 struts, 33 VIGVs and 26 rotor blades, were modelled in full. The model also included a prescribed inlet distortion pattern so that the combined effects of stator wakes and inlet distortion on the response of the rotor blades could be studied. The unsteady flow calculations were conducted using a time-accurate non-linear viscous flow representation. Blade motion was also included. Such an undertaking required about 4.2 million grid points to include all three bladerows in a complete stage calculation. To reduce the computational effort, a number of smaller computations were conducted by considering the stator and rotor domains separately: the outflow solution of the stator domain was used as an inflow boundary condition to the rotor domain. The results indicated that such isolated bladerow computations were likely to under-predict the response levels because of neglecting rotor-stator interactions. A number of low engine order (LEO) harmonics were identified from an inspection of the unsteady forcing created by the inlet distortions. Good agreement was obtained for cases where experimental data were available.

Numerical Research on Forced Response of a Compressor Rotor Row Under Inlet Distortion

2012 ◽  
Author(s):  
Zhang Zhang ◽  
Anping Hou ◽  
Wei Tuo ◽  
Aiguo Xia ◽  
Sheng Zhou
Keyword(s):  
Dynamic Response ◽  
Total Pressure ◽  
Three Dimensional ◽  
Forced Response ◽  
Vibration Response ◽  
Rotor Blades ◽  
Response Level ◽  
Blade Vibration ◽  
Inlet Distortion ◽  
Fatigue Life Analysis

Under inlet total pressure distortion, forced response of compressor blades poses a threat to aircraft propulsion system. Research on blade dynamic response is premise and basis for high-cycle fatigue life analysis. Blades of a compressor first rotor row are studied with three dimensional numerical simulation in fluid-structure coupling methods. The inlet distortion’s influence on blade aeroelastic dynamic response and flow field characteristics are analyzed. The results demonstrate that circumferential and radial total pressure distortion should be considered together in the phenomenon of actual inlet distortion induced blade vibration response. At the condition of low angle of attacks, radial distortion intensity is weak, the relation between vibration response level of rotor blades and circumferential distortion intensity is proportional. With the angle of attack increases, the vibratory stress under aerodynamic forces grows sufficiently. The radial total pressure distortion near hub increases dynamic response severity of rotor blades.

A Resonance Tracking Algorithm for the Prediction of Turbine Forced Response With Friction Dampers

2000 ◽  
Author(s):  
C. Bréard ◽  
J. S. Green ◽  
M. Vahdati ◽  
M. Imregun
Keyword(s):  
Rotor Blade ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Forced Response ◽  
Rotor Blades ◽  
Test Case ◽  
Friction Damper ◽  
Blade Vibration ◽  
Frequency Shifts ◽  
Friction Dampers

This paper presents an iterative method for determining the resonant speed shift when non-linear friction dampers are included in turbine blade roots. Such a need arises when conducting response calculations for turbine blades where the unsteady aerodynamic excitation must be computed at the exact resonant speed of interest. The inclusion of friction dampers is known to raise the resonant frequencies by up to 20% from the standard assembly frequencies. The iterative procedure uses a viscous, time-accurate flow representation for determining the aerodynamic forcing, a look-up table for evaluating the aerodynamic boundary conditions at any speed, and a time-domain friction damping module for resonance tracking. The methodology was applied to an HP turbine rotor test case where the resonances of interest were due to the 1T and 2F blade modes under 40 engine-order excitation. The forced response computations were conducted using a multi-stage approach in order to avoid errors associated with “linking” single stage computations since the spacing between the two bladerows was relatively small. Three friction damper elements were used for each rotor blade. To improve the computational efficiency, the number of rotor blades was decreased by 2 to 90 in order to obtain a stator/rotor blade ratio of 4/9. However, the blade geometry was skewed in order to match the capacity (mass flow rate) of the components and the condition being analysed. Frequency shifts of 3.2% and 20.0% were predicted for the 1T/40EO and 2F/40EO resonances in about 3 iterations. The predicted frequency shifts and the dynamic behaviour of the friction dampers were found to be within the expected range. Furthermore, the measured and predicted blade vibration amplitudes showed a good agreement, indicating that the methodology can be applied to industrial problems.

scholarly journals A Numerical Model of the Onset of Stall Flutter in Cascades

1995 ◽  
Author(s):  
William S. Clark ◽  
Kenneth C. Hall
Keyword(s):  
Unsteady Flow ◽  
Stokes Equations ◽  
Fluid Dynamic ◽  
Computational Fluid Dynamic Model ◽  
Navier Stokes ◽  
Aeroelastic Stability ◽  
Blade Vibration ◽  
Navier Stokes Equations ◽  
The Mean ◽  
Stall Flutter

In this paper, we present a computational fluid dynamic model of the unsteady flow associated with the onset of stall flutter in turbomachinery cascades. The unsteady flow is modeled using the laminar Navier-Stokes equations. We assume that the unsteadiness in the flow is a small harmonic disturbance about the mean or steady flow. Therefore, the unsteady flow is governed by a small-disturbance form of the Navier-Stokes equations. These linear variable coefficient equations are discretized on a deforming computational grid and solved efficiently using a multiple-grid Lax-Wendroff scheme. A number of numerical examples are presented which demonstrate the destabilizing influence of viscosity on the aeroelastic stability of airfoils in cascade, especially for torsional modes of blade vibration.

Numerical Study on the Effect of a Volute on Surge Phenomena in a Centrifugal Compressor

2016 ◽  
Author(s):  
S. H. Jeon ◽  
D. H. Hwang ◽  
J. H. Park ◽  
C. H. Kim ◽  
J. H. Baek ◽  
...  
Keyword(s):  
Centrifugal Compressor ◽  
Stokes Equations ◽  
Flow Instability ◽  
Numerical Study ◽  
Pressure Ratio ◽  
Unstable Flow ◽  
Step Size ◽  
Time Step ◽  
Time Step Size ◽  
Blade Passing Frequency

Numerical investigation of the effect of the volute on stall flow phenomenon is presented by solving three-dimensional Reynolds-averaged compressible Navier-Stokes equations. Two different configurations of a centrifugal compressor were used to compare their performance: One is an original centrifugal compressor which is composed of impeller, splitter, vaned diffuser and a volute and the other is the one without a volute. Steady calculations were performed to predict aerodynamic performance in terms of the pressure ratio, efficiency and mass flow rate. The results show that the operating range of the compressor with a volute is narrower than that of the compressor without a volute. This can be interpreted that flow instability is strongly influenced by the tongue of a volute which is highly asymmetric. Unsteady calculations were also performed with a time-step size of 38μs corresponding to a pitch angle of 5 degrees at the given rotational speed. The flow characteristics for two configurations are analyzed and compared at various instantaneous times showing unsteady dynamic features. Based on the unsteady flow simulation, fast Fourier transform at several discrete points in semi-vaneless space was performed at peak efficiency and near surge point in order to illustrate the unstable flow physics in both configurations. It is found that the blade passing frequency is dominant, indicating that diffuser passages have a periodicity of 40 degrees due to the rotational blades. Besides blade passing frequency, there were several noticeable frequencies which affect the instability of the whole system. Those frequencies in both configurations are compared and analyzed in various aspects.

Effects of S-Shaped Intake on Aeromechanical Characteristics of a Transonic Fan

2020 ◽  
Author(s):  
Yun Zheng ◽  
Kang Xu ◽  
Hui Yang ◽  
Qingzhe Gao ◽  
Xiubo Jin
Keyword(s):  
Stokes Equations ◽  
Three Dimensional ◽  
Forced Response ◽  
Mode Shapes ◽  
Flow Distortion ◽  
Plane Model ◽  
Response Level ◽  
Blade Vibration ◽  
Structural Dynamic ◽  
Transonic Fan

Abstract S-shaped intakes are widely used in aero-engines of modern fighters because of the demand for reducing radar cross-section. Besides, boundary layer ingestion (BLI) configurations are proposed in civil engines recently due to the high propulsion efficiency and low fuel consumption. And S-shaped ducts are usually used as transition sections of diffusers in BLI intakes. Compared with normal straight intakes, it is inevitable to bring in the influence of inlet distortion and acoustic reflection for S-shaped intakes. Meanwhile, composite fan blades, shorter intakes and integrated blisks are common in engine designs. So, fan blades are prone to serious vibrations such as flutter and forced response, which may lead to high-cycle fatigue, and further cause structural failure. The aeromechanical characteristics of a transonic fan (NASA rotor67) in presence of a s-shaped intake are predicted by an in-house integrated time-domain aeroelasticity code. The three dimensional, time-accurate, unsteady Reynolds-Averaged Navier-Stokes equations are solved in fluid domain, and the structural dynamic equations of blade vibration are solved with a modal superimposition method. Mode shapes and natural frequencies of rotor blade are obtained with a commercial Finite Element code, and the Campbell diagram is presented. Full-annulus aeroelastic calculations are conducted to obtain the transient response and the aerodynamic damping of fan blades. Different techniques for interface between the intake and the rotor are used for comparison to demonstrate the influence of upstream interaction. A mixing-plane model is used at the interface to model the blade vibration without interactions with the distortion, while a sliding-plane model is used at the same condition to include the flow distortion and acoustic effects on the fan blade motion. S-shaped intakes with two different axial length are investigated for the forced response and flutter stability. This study indicates that the forced response level is attenuated due to the decrease of distortion level as the length increases, while the flutter stability is determined by the phase difference between the upstream and the reflective acoustic wave.

Investigation of an innovative non-free vortex aerodynamic procedure to design a single-stage transonic compressor

2017 ◽  
Vol 232 (11) ◽  
pp. 2132-2143 ◽  
Author(s):  
A Shahsavari ◽  
M Nili-Ahmadabadi
Keyword(s):  
Fluid Dynamics ◽  
Computational Fluid Dynamics ◽  
Stokes Equations ◽  
Design Method ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Rotor Blades ◽  
Meridional Plane ◽  
Transonic Compressor ◽  
Performance Map

This paper presents an innovative design method for a transonic compressor based on the radial equilibrium theory by means of increasing blade loading. Firstly, the rotor blade of a transonic compressor is redesigned based on the constant spanwise de-Haller number and diffusion. The design method leads to an unconventional increased axial velocity distribution in tip section, which originates from non-uniform enthalpy distribution assumption. A code is applied to extract the compressor meridional plane and blade-to-blade geometry containing rotor and stator in order to design the blade three-dimensional view. A structured grid is generated for the numerical domain of fluid. Finer grids are used for the regions near walls to capture the boundary layer effects and behavior. Reynolds-averaged Navier–Stokes equations are solved by finite volume method for rotating zones (rotor) and stationary zones (stator). The experimental data, available for the performance map of NASA Rotor67, is used to validate the results of the current simulations. Then, the capability of the design method is validated by computational fluid dynamics that is capable of predicting the performance map. The numerical results of the new geometry by representing 11% improvement in efficiency and 19% in total pressure ratio verify the new method advantages. The computational fluid dynamics results also show that the newly designed rotor blades due to a higher velocity in the tip section have a special capacity to increase the loading without any separation. The mass flow reduction is observed in the new geometry, which could be easily improved by changing stagger angle.

Forced Response in Axial Turbines Under the Influence of Partial Admission

2013 ◽  
Vol 135 (4) ◽  
Author(s):  
Jens Fridh ◽  
Björn Laumert ◽  
Torsten Fransson
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Forced Response ◽  
Rotor Blades ◽  
Response Analysis ◽  
Validation Data ◽  
Pressure Transducers ◽  
Control Stage ◽  
Engine Order ◽  
Partial Admission

High cycle fatigue (HCF) due to unforeseen excitation frequencies, underestimated force magnitudes, or a combination of both causes control-stage failures for steam turbine stakeholders. This paper provides an extended design criteria toolbox, as well as validation data, for control-stage design based on experimental data to reduce HCF incidents in partial-admission turbines. The upstream rotor in a two-stage air test turbine is instrumented with pressure transducers and strain gauges. Admission degrees extend from 28.6% to 100%, as one or two admission arcs are simulated by blocking segmental arcs immediately upstream of the first stator vanes with aerodynamically shaped filling blocks. Sweeps across a speed range of 50%–105% of design speed are performed at a constant turbine pressure ratio during simultaneous high-speed acquisition. A forced-response analysis is performed and results presented in Campbell diagrams. Partial admission creates a large number of low-engine-order forced responses because of the blockage, pumping, loading, and unloading processes. Combinations of the number of rotor blades and low-engine-order excitations are the principal sources of forced-response vibrations for the turbine studied here. Altering the stator and/or rotor pitches changes the excitation pattern. We observed that a relationship between the circumferential lengths of the admitted and nonadmitted arcs dictates the excitation forces and may serve as a design parameter.

Forced Response in Axial Turbines Under the Influence of Partial Admission

2012 ◽  
Author(s):  
Jens Fridh ◽  
Björn Laumert ◽  
Torsten Fransson
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Forced Response ◽  
Rotor Blades ◽  
Response Analysis ◽  
Validation Data ◽  
Pressure Transducers ◽  
Control Stage ◽  
Engine Order ◽  
Partial Admission

High cycle fatigue (HCF) due to unforeseen excitation frequencies or due to under predicted force magnitudes, or a combination of both causes control stage failures for steam turbine stakeholders. The objectives of this paper is to provide an extended design criteria toolbox and validation data for control stage design based on experimental data, with the aim to decrease HCF incidents for partial admission turbines. The upstream rotor in a two stage air test turbine is instrumented with pressure transducers and strain gauges. Admission degrees stretching from 28.6% to 100% as one or two admission arcs are simulated by blocking segmental arcs immediately upstream of first stator vanes by aerodynamically shaped filling blocks. Sweeps across a speed range from 50 to 105% of design speed are performed at constant turbine pressure ratio during simultaneous high speed acquisition. A forced response analysis is performed and results presented in Campbell diagrams. Partial admission creates a large number of low engine order forced responses because of the blockage, pumping, loading and unloading processes. Combinations of the number of rotor blades and low engine order excitations are the principal sources of forced response vibrations for the turbine studied herein. Altering the stator and/or rotor pitches will change the excitation pattern. A relation between the circumferential lengths of the admitted and non-admitted arcs that dictates the excitation forces is observed that may serve as a design parameter.

Forced Response Excitation Due to the Stator Vanes of Two and Three Compressor Stages Away

2021 ◽  
Author(s):  
Toshimasa Miura ◽  
Naoto Sakai ◽  
Naoki Kanazawa ◽  
Kentaro Nakayama
Keyword(s):  
Gas Turbines ◽  
Rotor Blade ◽  
High Efficiency ◽  
Potential Effect ◽  
Forced Response ◽  
Rotor Blades ◽  
Test Facility ◽  
Blade Vibration ◽  
Stator Vane ◽  
Aero Engines

Abstract State-of-the-art axial compressors of gas turbines employed in power generation plants and aero engines should have both high efficiency and small footprint. Thus, compressors are designed to have thin rotor blades and stator vanes with short axial distances. Recently, problems of high cycle fatigue (HCF) associated with forced response excitation have gradually increased as a result of these trends. Rotor blade fatigue can be caused not only by the wake and potential effect of the adjacent stator vane, but also by the stator vanes of two, three or four compressor stages away. Thus, accurate prediction and suppression methods are necessary in the design process. In this study, the problem of rotor blade vibration caused by the stator vanes of two and three compressor stages away is studied. In the first part of the study, one-way FSI simulation is carried out. To validate the accuracy of the simulation, experiments are also conducted using a gas turbine test facility. It is found that one-way FSI simulation can accurately predict the order of the vibration level. In the second part of the study, a method of controlling the blade vibration is investigated by optimizing the clocking of the stator vanes. It is confirmed that the vibration amplitude can be effectively suppressed without reducing the performance. Through this study, ways to evaluate and control the rotor blade vibration are validated.

Prediction of Low Engine Order Inlet Distortion Driven Resonance in a Low Aspect Ratio Fan

2000 ◽  
Author(s):  
J. G. Marshall ◽  
L. Xu ◽  
J. Denton ◽  
J. W. Chew
Keyword(s):  
Aspect Ratio ◽  
Three Dimensional ◽  
Coupled Model ◽  
Response Prediction ◽  
Euler Method ◽  
Forced Response ◽  
Blade Vibration ◽  
Inlet Distortion ◽  
Low Aspect Ratio ◽  
Engine Order

This paper presents a forced response prediction of 3 resonances in a low aspect ratio modern fan rotor and compares with other worker’s experimental data. The incoming disturbances are due to low engine-order inlet distortion from upstream screens. The resonances occur in the running range at 3 and 8 engine orders which cross low modes (flap, torsion and stripe) of the blade. The fan was tested with on-blade instrumentation at both on- and off-resonant conditions to establish the unsteady pressures due to known distortion patterns. The resulting steady and unsteady flow in the fan blade passages has been predicted by three methods, all three-dimensional. The first is a linearised unsteady Euler method; the second is a non-linear unsteady Navier-Stokes method; the third method uses a similar level of aerodynamic modelling as the second but also includes a coupled model of the structural dynamics. The predictions for the 3 methods are presented against the test data, and further insight into the problem is obtained through post-processing of the data. Predictions of the blade vibration response are also obtained. Overall the level of agreement between calculations and measurements is considered encouraging although further research is needed.

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