The Influence of Compressor Blade Row Interaction Modeling on Performance Estimates From Time-Accurate, Multistage, Navier–Stokes Simulations

2008 ◽  
Vol 130 (1) ◽  
Author(s):  
Dale Van Zante ◽  
Jenping Chen ◽  
Michael Hathaway ◽  
Randall Chriss
Keyword(s):  
Boundary Conditions ◽  
High Speed ◽  
Axial Compressor ◽  
Time History ◽  
Navier Stokes ◽  
Strong Interactions ◽  
Phase Lag ◽  
Single Passage ◽  
Blade Row ◽  
Performance Estimates

The time-accurate, multistage, Navier–Stokes, turbomachinery solver TURBO was used to calculate the aeroperformance of a 2 1∕2 stage, highly loaded, high-speed, axial compressor. The goals of the research project were to demonstrate completion times for multistage, time-accurate simulations that are consistent with inclusion in the design process and to assess the influence of differing approaches to modeling the effects of blade row interactions on aeroperformance estimates. Three different simulation setups were used to model blade row interactions: (1) single-passage per blade row with phase lag boundaries, (2) multiple passages per blade row with phase lag boundaries, and (3) a periodic sector (1∕2 annulus sector). The simulations used identical inlet and exit boundary conditions and identical meshes. To add more blade passages to the domain, the single-passage meshes were copied and rotated. This removed any issues of differing mesh topology or mesh density from the following results. The 1∕2 annulus simulation utilizing periodic boundary conditions required an order of magnitude fewer iterations to converge when all three simulations were converged to the same level as assessed by monitoring changes in overall adiabatic efficiency. When using phase lag boundary conditions, the necessity to converge the time history information requires more iterations to obtain the same convergence level. In addition to convergence differences, the three simulations gave different overall performance estimates where the 1∕2 annulus case was 1.0 point lower in adiabatic efficiency than the single-passage phase lag case. The interaction between blade rows in the same frame of reference sets up spatial variations of properties in the circumferential direction, which are stationary in that reference frame. The phase lag boundary condition formulation will not capture this effect because the blade rows are not moving relative to each other. Thus, for simulations of more than two blade rows and strong interactions, a periodic simulation is necessary to estimate the correct aeroperformance.

The Influence of Compressor Blade Row Interaction Modeling on Performance Estimates From Time-Accurate, Multi-Stage, Navier-Stokes Simulations

2005 ◽  
Author(s):  
Dale Van Zante ◽  
Jenping Chen ◽  
Michael Hathaway ◽  
Randall Chriss
Keyword(s):  
Boundary Conditions ◽  
High Speed ◽  
Time History ◽  
Navier Stokes ◽  
Strong Interactions ◽  
Phase Lag ◽  
Single Passage ◽  
Multi Stage ◽  
Blade Row ◽  
Performance Estimates

The time-accurate, multi-stage, Navier-Stokes, turbomachinery solver TURBO was used to calculate the aero performance of a 2 1/2 stage, highly-loaded, high-speed, axial compressor. The goals of the research project were to demonstrate completion times for multi-stage, time-accurate simulations that are consistent with inclusion in the design process, and to assess the influence of differing approaches to modeling the effects of blade row interactions on aero performance estimates. Three different simulation setups were used to model blade row interactions: 1.) single passage per blade row with phase lag boundaries, 2.) multiple passages per blade row with phase lag boundaries, and 3.) a periodic sector (1/2 annulus sector). The simulations used identical inlet and exit boundary conditions and identical meshes. To add more blade passages to the domain, the single passage meshes were copied and rotated. This removed any issues of differing mesh topology or mesh density from the following results. The 1/2 annulus simulation utilizing periodic boundary conditions required an order of magnitude less iterations to converge when all three simulations were converged to the same level as assessed by monitoring changes in overall adiabatic efficiency. When using phase lag boundary conditions the need to converge the time history information necessitates more iterations to obtain the same convergence level. In addition to convergence differences, the three simulations gave different overall performance estimates where the 1/2 annulus case was 1.0 point lower in adiabatic efficiency than the single passage phase lag case. The interaction between blade rows in the same frame of reference set up spatial variations of properties in the circumferential direction which are stationary in that reference frame. The phase lag boundary condition formulation will not capture this effect because the blade rows are not moving relative to each other. Thus for simulations of more than two blade rows and strong interactions, a periodic simulation is necessary to estimate the correct aero performance.

Full-Annulus Simulation of Nonsynchronous Blade Vibration Excitation of an Axial Compressor

2017 ◽  
Vol 140 (3) ◽  
Author(s):  
Daniel Espinal ◽  
Hong-Sik Im ◽  
Ge-Cheng Zha
Keyword(s):  
High Speed ◽  
Pressure Ratio ◽  
Excitation Frequency ◽  
Radial Profile ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Phase Lag ◽  
Blade Vibration ◽  
Weighted Essentially Nonoscillatory ◽  
Unsteady Fluid

A high speed 1–1/2 axial compressor stage is simulated in this paper using an unsteady Reynolds-averaged Navier–Stokes (URANS) solver for a full-annulus configuration to capture its nonsynchronous vibration (NSV) flow excitation with rigid blades. A third-order weighted essentially nonoscillatory scheme for the inviscid flux and a second-order central differencing for the viscous terms are used to resolve nonlinear unsteady fluid flows. A fully conservative rotor/stator sliding boundary condition (BC) is employed with multiple-processor capability for rotor/stator sliding interface that accurately captures unsteady wake propagation between the rotor and stator blades while conserving fluxes across the rotor/stator interfaces. The predicted dominant frequencies using the blade tip response signals are not harmonic to the engine order, which is the NSV excitation. The simulation is based on a rotor blade with a 1.1% tip-chord clearance. Comparison with the previous 1/7 annulus simulations show that the time-shifted phase-lag BCs used in the 1/7 annulus are accurate. For most of the blades, the NSV excitation frequency is 6.2% lower than the measurement in the rig test, although some blades displayed slightly different NSV excitation frequencies. The simulation confirms that the NSV is a full annulus phenomenon. The instability of the circumferential traveling vortices in the vicinity of the rotor tip due to the strong interaction of incoming flow is the main cause of the NSV excitation. This instability is present in all blades of the rotor annulus. For circumferentially averaged parameters like total pressure ratio, NSV is observed to have an effect on the radial profile, particularly at radial locations above 70% span. A design with a lower loading of the upper blade span and a higher loading of the midblade spans is recommended to mitigate or remove NSV.

scholarly journals Navier-Stokes and Potential Calculations of Axial Spacing Effect on Vortical and Potential Disturbances and Gust Response in an Axial Compressor

1995 ◽  
Author(s):  
Meng-Hsuan Chung ◽  
Andrew M. Wo
Keyword(s):  
Lower Order ◽  
Spacing Effect ◽  
Axial Compressor ◽  
Interaction Effects ◽  
Moving Frame ◽  
Navier Stokes ◽  
Blade Row ◽  
Vortical Disturbance ◽  
Blade Row Interaction ◽  
Gust Response

The effect of blade row axial spacing on vortical and potential disturbances and gust response is studied for a compressor stator/rotor configuration near design and at high loadings using 2D incompressible Navier-Stokes and potential codes, both written for multistage calculations. First, vortical and potential disturbances downstream of the isolated stator in the moving frame are defined; these disturbances exclude blade row interaction effects. Then, vortical and potential disturbances for the stator/rotor configuration are calculated for axial gaps of 10%, 20%, and 30% chord. Results show that the potential disturbance is uncoupled; the potential disturbance calculated from the isolated stator configuration is a good approximation for that from the stator/rotor configuration for all three axial gaps. The vortical disturbance depends strongly on blade row interactions. Low order modes of vortical disturbance are of substantial magnitude and decay much more slowly downstream than do those of potential disturbance. Vortical disturbance decays linearly with increasing mode except very close to the stator trailing edge. For a small axial gap, lower order modes of both vortical and potential disturbances must be included to determine the rotor gust response.

Evaluation of a Flow Measurement Probe Influence on the Flow Field in High Speed Axial Compressors

2021 ◽  
Author(s):  
Ryosuke Seki ◽  
Satoshi Yamashita ◽  
Ryosuke Mito
Keyword(s):  
Flow Field ◽  
Approximation Method ◽  
High Speed ◽  
Static Pressure ◽  
Axial Compressor ◽  
Pressure Change ◽  
Operating Point ◽  
Blade Row ◽  
Probe Insertion ◽  
Probe Geometry

Abstract The aerodynamic effects of a probe for stage performance evaluation in a high-speed axial compressor are investigated. Regarding the probe measurement accuracy and its aerodynamic effects, the upstream/downstream effects on the probe and probe insertion effects are studied by using an unsteady computational fluid dynamics (CFD) analysis and by verifying in two types of multistage high-speed axial compressor measurements. The probe traverse measurements were conducted at the stator inlet and outlet in each case to evaluate blade row performance quantitatively and its flow field. In the past study, the simple approximation method was carried out which considered only the interference of the probe effect based on the reduction of the mass flow by the probe blockage for the compressor performance, but it did not agree well with the measured results. In order to correctly and quantitatively grasp the mechanism of the flow field when the probe is inserted, the unsteady calculation including the probe geometry was carried out in the present study. Unsteady calculation was performed with a probe inserted completely between the rotor and stator of a 4-stage axial compressor. Since the probe blockage and potential flow field, which mean the pressure change region induced by the probe, change the operating point of the upstream rotor and increase the work of the rotor. Compared the measurement result with probe to a kiel probe setting in the stator leading edge, the total pressure was increased about 2,000Pa at the probe tip. In addition, the developed wake by the probe interferes with the downstream stator row and locally changes the static pressure at the stator exit. To evaluate the probe insertion effect, unsteady calculations with probe at three different immersion heights at the stator downstream in an 8-stage axial compressor are performed. The static pressure value of the probe tip was increased about 3,000Pa in the hub region compared to tip region, this increase corresponds to the measurement trend. On the other hand, the measured wall static pressure showed that there is no drastic change in the radial direction. In addition, when the probe is inserted from the tip to hub region in the measurement, the blockage induced by the probe was increased. As a result, operating point of the stator was locally changed, and the rise of static pressure of the stator increased when the stator incidence changed. These typical results show that unsteady simulations including probe geometry can accurately evaluate the aerodynamic effects of probes in the high-speed axial compressor. Therefore, since the probe will pinpointed and strong affects the practically local flow field in all rotor upstream passage and stator downstream, as for the probe measurement, it is important to pay attention to design the probe diameter, the distance from the blade row, and its relative position to the downstream stator. From the above investigations, a newly simple approximation method which includes the effect of the pressure change evaluation by the probe is proposed, and it is verified in the 4-stage compressor case as an example. In this method, the effects of the distance between the rotor trailing edge (T.E.) and the probe are considered by the theory of the incompressible two-dimensional potential flow. The probe blockage decreases the mass flow rate and changes the operating point of the compressor. The verification results conducted in real compressor indicate that the correct blockage approximation enables designer to estimate aerodynamic effects of the probe correctly.

Numerical Validation of Characteristic and Linearized Unsteady Boundary Conditions for Non-Integer Blade Ratios in a Non-Linear Navier Stokes Solver

2005 ◽  
Author(s):  
Thomas Mokulys ◽  
Stephen C. Dewhurst ◽  
Reza S. Abhari
Keyword(s):  
Boundary Conditions ◽  
New Method ◽  
Navier Stokes ◽  
Conservative Approach ◽  
Single Row ◽  
Numerical Validation ◽  
Single Passage ◽  
Non Linear ◽  
Improved Stability ◽  
Similar Accuracy

Unsteady single passage methods for turbomachinery applications are becoming a more and more mature method in predicting unsteady blade-to-blade aerodynamics in a more efficient and accurate way compared to multipassage computations with scaled airfoil geometries. The outlined work presents results of a fourier-based method for phase-lagged boundary conditions using a usual conservative approach at the interfaceplane and compares it to a new characteristic treatment as well as a linearized single row approach. The new method provides results of similar accuracy with improved stability of the algorithm and faster convergence times.

scholarly journals Numerical Optimization of a Stall Margin Enhancing Recirculation Channel for an Axial Compressor

Fluids ◽  
2019 ◽  
Vol 4 (2) ◽  
pp. 88
Author(s):  
Motoyuki Kawase ◽  
Aldo Rona
Keyword(s):  
High Speed ◽  
High Performance ◽  
Fluid Dynamic ◽  
Axial Compressor ◽  
Leading Edge ◽  
Navier Stokes ◽  
Test Case ◽  
Dynamic Simulations ◽  
Stall Margin ◽  
Casing Treatment

A proof of concept is provided by computational fluid dynamic simulations of a new recirculating type casing treatment. This treatment aims at extending the stable operating range of highly loaded axial compressors, so to improve the safety of sorties of high-speed, high-performance aircraft powered by high specific thrust engines. This casing treatment, featuring an axisymmetric recirculation channel, is evaluated on the NASA rotor 37 test case by steady and unsteady Reynolds Averaged Navier Stokes (RANS) simulations, using the realizable k-ε model. Flow blockage at the recirculation channel outlet was mitigated by chamfering the exit of the recirculation channel inner wall. The channel axial location from the rotor blade tip leading edge was optimized parametrically over the range −4.6% to 47.6% of the rotor tip axial chord c z . Locating the channel at 18.2% c z provided the best stall margin gain of approximately 5.5% compared to the untreated rotor. No rotor adiabatic efficiency was lost by the application of this casing treatment. The investigation into the flow structure with the recirculating channel gave a good insight into how the new casing treatment generates this benefit. The combination of stall margin gain at no rotor adiabatic efficiency loss makes this design attractive for applications to high-speed gas turbine engines.

scholarly journals The Computation of Adjacent Blade-Row Effects in a 1.5 Stage Axial Flow Turbine

1997 ◽  
Author(s):  
Rolf Emunds ◽  
Ian K. Jennions ◽  
Dieter Bohn ◽  
Jochen Gier
Keyword(s):  
Experimental Data ◽  
Numerical Simulation ◽  
Boundary Conditions ◽  
Axial Flow ◽  
The Other ◽  
Navier Stokes ◽  
Blade Row ◽  
Flow Through ◽  
The University

This paper deals with the numerical simulation of flow through a 1.5 stage axial flow turbine. The 3-row configuration has been experimentally investigated at the University of Aachen where measurements behind the first vane, the first stage and the full configuration were taken. These measurements allow single blade row computations, to the measured boundary conditions taken from complete engine experiments, or full multistage simulations. The results are openly available inside the framework of ERCOFTAC 1996. There are two separate but interrelated parts to the paper. Firstly, two significantly different Navier-Stokes codes are used to predict the flow around the first vane and the first rotor, both running in isolation. This is used to engender confidence in the code that is subsequently used to model the multiple bladerow tests, the other code is currently only suitable for a single blade row. Secondly, the 1.5 stage results are compared to the experimental data and promote discussion of surrounding blade row effects on multistage solutions.

Compressor Stability Enhancement Using Discrete Tip Injection

2000 ◽  
Vol 123 (1) ◽  
pp. 14-23 ◽  
Author(s):  
Kenneth L. Suder ◽  
Michael D. Hathaway ◽  
Scott A. Thorp ◽  
Anthony J. Strazisar ◽  
Michelle B. Bright
Keyword(s):  
Axial Velocity ◽  
High Speed ◽  
Axial Compressor ◽  
Flow Coefficient ◽  
Navier Stokes ◽  
Injection Velocity ◽  
Compressor Rotor ◽  
Annulus Flow ◽  
Tip Injection ◽  
Stability Enhancement

Mass injection upstream of the tip of a high-speed axial compressor rotor is a stability enhancement approach known to be effective in suppressing stall in tip-critical rotors. This process is examined in a transonic axial compressor rotor through experiments and time-averaged Navier-Stokes CFD simulations. Measurements and simulations for discrete injection are presented for a range of injection rates and distributions of injectors around the annulus. The simulations indicate that tip injection increases stability by unloading the rotor tip and that increasing injection velocity improves the effectiveness of tip injection. For the tested rotor, experimental results demonstrate that at 70 percent speed the stalling flow coefficient can be reduced by 30 percent using an injected massflow equivalent to 1 percent of the annulus flow. At design speed, the stalling flow coefficient was reduced by 6 percent using an injected massflow equivalent to 2 percent of the annulus flow. The experiments show that stability enhancement is related to the mass-averaged axial velocity at the tip. For a given injected massflow, the mass-averaged axial velocity at the tip is increased by injecting flow over discrete portions of the circumference as opposed to full-annular injection. The implications of these results on the design of recirculating casing treatments and other methods to enhance stability will be discussed.

The Simulation of Turbomachinery Blade Rows in Asymmetric Flow Using Actuator Disks

1997 ◽  
Vol 119 (4) ◽  
pp. 723-732 ◽  
Author(s):  
W. G. Joo ◽  
T. P. Hynes
Keyword(s):  
Supersonic Flow ◽  
Flow Field ◽  
Boundary Conditions ◽  
High Speed ◽  
Axial Position ◽  
Numerical Implementation ◽  
Asymmetric Flow ◽  
Actuator Disk ◽  
Blade Row ◽  
Flow Through

This paper describes the development of actuator disk models to simulate the asymmetric flow through high-speed low hub-to-tip ratio blade rows. The actuator disks represent boundaries between regions of the flow in which the flow field is solved by numerical computation. The appropriate boundary conditions and their numerical implementation are described, and particular attention is paid to the problem of simulating the effect of blade row blockage near choking conditions. Guidelines on choice of axial position of the disk are reported. In addition, semi-actuator disk models are briefly described and the limitations in the application of the model to supersonic flow are discussed.

Simulation of Non-Synchronous Blade Vibration of an Axial Compressor Using a Fully Coupled Fluid/Structure Interaction

2012 ◽  
Author(s):  
Hong-Sik Im ◽  
Ge-Cheng Zha
Keyword(s):  
Boundary Condition ◽  
Stokes Equations ◽  
Axial Compressor ◽  
Tip Vortex ◽  
Weno Scheme ◽  
Navier Stokes ◽  
Phase Lag ◽  
Tip Leakage Flow ◽  
Blade Vibration ◽  
Fully Coupled

In this paper non-synchronous vibration (NSV) of a GE axial compressor is simulated using a fully coupled fluid/strcuture interaction (FSI). Time accurate Navier-Stokes equations are solved with a system of 5 decoupled structure modal equations in a fully coupled manner. A 3rd order WENO scheme for the inviscid flux and a 2nd order central differencing for the viscous terms are used to resolve nonlinear interaction between vibrating blades and fluid flow. 1/7th annulus is used with a time shifted phase-lag (TSPL) boundary condition to reduce computational efforts. A fully conservative rotor/stator sliding boundary condition is employed to accurately capture unsteady wake propagation between the rotor and stator blades. The predicted dominant frequencies using the blade tip response signals are not harmonic to the engine order, which is the NSV. The blade vibration is torsionally coupled with highly oscillating blade pressure and is not damped out during the NSV. No resonance to the blade natural frequencies is found. The instability of tornado vortices in the vicinity of the rotor tip due to the strong interaction of incoming flow, tip vortex and tip leakage flow is the main cause of the NSV observed in this study.

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