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.

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.

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.

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.

Unsteady Three-Dimensional Analysis of Inlet Distortion in a Transonic Fan

2006 ◽  
Author(s):  
Peng Sun ◽  
Guotal Feng
Keyword(s):  
Shock Wave ◽  
Total Pressure ◽  
Large Scale ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Inlet Distortion ◽  
Flow Loss ◽  
Large Scale Vortex ◽  
Total Temperature ◽  
Transonic Fan

A time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan was carried out using "Fluent-parallel" in a parallel supercomputer. The numerical simulation focused on a transonic fan with inlet square wave total pressure distortion and the analysis of result consisted of three aspects. The first was about inlet parameters redistribution and outlet total temperature distortion induced by inlet total pressure distortion. The pattern and causation of flow loss caused by pressure distortion in rotor were analyzed secondly. It was found that the influence of distortion was different at different radial positions. In hub area, transportation-loss and mixing-loss were the main loss patterns. Distortion not only complicated them but enhanced them. Especially in stator, inlet total pressure distortion induced large-scale vortex, which produced backflow and increased the loss. While in casing area, distortion changed the format of shock wave and increased the shock loss. Finally, the format of shock wave and the hysteresis of rotor to distortion were analyzed in detail.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

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.

Flow Field Analysis of Inlet Distortion in a Transonic Fan With Straight and Bowed Stator Blade

2007 ◽  
Author(s):  
Peng Sun ◽  
Jingjun Zhong ◽  
Guotai Feng
Keyword(s):  
Flow Field ◽  
Total Pressure ◽  
Three Dimensional ◽  
Field Analysis ◽  
Navier Stokes ◽  
Fan Performance ◽  
Inlet Distortion ◽  
Flow Loss ◽  
Stator Blade ◽  
Transonic Fan

The performance and stability of a fan in clean and distorted inlet flow can be improved through the use of bowed stator blades. Measurements between the blade rows in transonic and supersonic flow are too complex to provide any useful insights, so 3D flow simulations are required. In this paper, a time-accurate three-dimensional Navier-Stokes solver of the unsteady flow field in a transonic fan is carried out using “Fluent-parallel” in a parallel supercomputer. Two sets of simulations are performed. The first simulation focuses on a better understanding of inlet total pressure distortion effects on a transonic fan. The second set of numerical simulation aims at studying the improvements of fan performance made by bowed stator blades. Three aspects are contained in this paper. The first is about the distortion effects on characteristics of the fan stage with straight stator. The effects of bowed stator on fan performance with inlet distortion are demonstrated secondly. One hand bowed stator increases the loss in rotor. On the other hand, it reduces the flow loss in stator. Finally, the patterns of flow loss caused by total pressure distortion with straight/bowed stator are compared. The scale of vortex in stator induced by inlet total pressure distortion is weakened by bowed blades, which decreases the stator loss.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1998 ◽  
Vol 120 (2) ◽  
pp. 233-246 ◽  
Author(s):  
C. Hah ◽  
D. C. Rabe ◽  
T. J. Sullivan ◽  
A. R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of eight periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier–Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20 percent of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

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 Flutter and Inlet Distortion Driven Response of a Transonic Fan Rotor Using Phase-Lagged Boundary Conditions

2001 ◽  
Author(s):  
H. D. Li ◽  
L. He
Keyword(s):  
Computing Time ◽  
Three Dimensional ◽  
Correction Method ◽  
Forced Response ◽  
Navier Stokes ◽  
Design Environment ◽  
Inlet Distortion ◽  
Single Passage ◽  
Shape Correction ◽  
Transonic Fan

Prediction of blade forced response and flutter is of great importance to turbomachinery designers. However, calculations of unsteady turbomachinery flows using conventional time-domain methods typically would lead to the use of multi-passage/whole-annulus domains due to the required direct periodic condition. This makes numerical computations extremely time-consuming and is one of the major difficulties for nonlinear unsteady calculations to be applied in a blading design environment. A single-passage approach to three-dimensional unsteady Navier-Stokes calculations using the Fourier-series based Shape-Correction method has been developed, and been applied to analyze inlet distortion driven response and flutter of a transonic fan rotor (NASA Rotor-67). The key feature is that the Shape-Correction method enables a single-passage solution to unsteady flows in blade rows under influences of multiple disturbances with arbitrary inter-blade phase angles. The results show that the single-passage solution can capture deterministic unsteadiness as well as time-averaged flows in good agreement with conventional multi-passage solutions, while the corresponding computing time can be reduced dramatically.

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