Evaluation of Flow Field Approximations for Transonic Compressor Stages

1996 ◽  
Author(s):  
Daniel J. Dorney ◽  
Om P. Sharma
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Loosely Coupled ◽  
Transonic Compressor ◽  
Blade Row ◽  
Flow Through ◽  
And Performance ◽  
Computer Resources ◽  
Fully Coupled

The flow through gas-turbine compressors is often characterized by unsteady, transonic and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor-stator interacting Navier-Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses and performance quantities associated with axial compressor stages are presented. The methods studied include, 1) the unsteady fully-coupled blade row technique, 2) the steady coupled blade row method, 3) the steady single blade row technique, and 4) the loosely-coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.

Evaluation of Flow Field Approximations for Transonic Compressor Stages

1997 ◽  
Vol 119 (3) ◽  
pp. 445-451 ◽  
Author(s):  
D. J. Dorney ◽  
O. P. Sharma
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Loosely Coupled ◽  
Transonic Compressor ◽  
Blade Row ◽  
Flow Through ◽  
And Performance ◽  
Computer Resources ◽  
Fully Coupled

The flow through gas turbine compressors is often characterized by unsteady, transonic, and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor–stator interacting Navier–Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses, and performance quantities associated with axial compressor stages are presented. The methods studied include: (1) the unsteady fully coupled blade row technique, (2) the steady coupled blade row method, (3) the steady single blade row technique, and (4) the loosely coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.

A Numerical Investigation of the Interaction Between Rotor Bow Waves and Upstream Wake Generators in a Transonic Compressor Stage

2003 ◽  
Author(s):  
Gregory Bloch ◽  
James Loellbach ◽  
Chunill Hah
Keyword(s):  
Numerical Investigation ◽  
Rotor Blade ◽  
Similar Case ◽  
Axial Compressor ◽  
Navier Stokes ◽  
Transonic Compressor ◽  
Blade Row ◽  
Transonic Axial Compressor ◽  
Main Effects ◽  
Blade Row Interaction

A numerical investigation of unsteady blade row interaction in a transonic axial compressor was performed. The compressor consists of an upstream wake generator (WG) blade row followed by a rotor blade row. Blade row interaction consists of two main effects: the downstream influence on the rotor flowfield of wakes and unsteady vortices shed from the wake generator, and the upstream influence on the wake generator of the rotor bow shock waves. An unsteady, two-dimensional, Navier-Stokes simulation was performed at the 75% span location of the compressor. Results from the numerical simulation are compared to previously reported numerical results and to experimental measurements from a similar case.

Numerical Transonic Flow Field Predictions for NASA Compressor Rotor 37

1995 ◽  
Author(s):  
Peter Dalbert ◽  
Donald H. Wiss
Keyword(s):  
Flow Field ◽  
Axial Compressor ◽  
Flow Simulation ◽  
Turbulence Models ◽  
Navier Stokes ◽  
Nasa Lewis Research ◽  
Solution Strategies ◽  
Compressor Rotor ◽  
High Prediction ◽  
Computer Resources

Flow field calculations of the NASA transonic axial compressor Rotor 37 are presented. These were obtained by the two commercially available 3D Navier Stokes-codes BTOB3D and TASCflow using different turbulence models, i.e. Baldwin-Lomax and k-ε. Some of the results were submitted to the CFD code assessment exercise organized in 1994 by the Turbomachinery Committee of the ASME, where a number of “blind” CFD predictions were compared against previously unknown experimental data taken at the NASA Lewis Research Center. The objective of these calculations was to use the codes in the same way as they are generally used by experienced engineers for standard industrial design tasks. Thus, the effort involved in grid generation, flow simulation runs and postprocessing was subject to the usual limitations in computer resources as well as a stringent observation of cost-effectiveness (manpower and time available). With both codes, two sets of calculations were carried out: BTOB3D with two different tip clearances and TASCflow with a uniform and a spanwise varying outlet static pressure. Generally, the results of both codes show good agreement with respect to the measured overall performance characteristics and averaged spanwise distributions. In particular, the TASCflow solutions display high prediction accuracy in some local details of the flow field while in the BTOB3D code, boundary effects seem to mix out the flow significantly. The solution strategies employed as well as the reasons for certain discrepancies between computations and measurements are discussed.

scholarly journals Three-Dimensional Analysis of a Multistage Compressor Flow Field

1998 ◽  
Author(s):  
Wolfgang Elmendorf ◽  
Frank Mildner ◽  
Ralf Röper ◽  
Uwe Krüger ◽  
Michael Kluck
Keyword(s):  
Flow Field ◽  
Gas Turbine ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Flow Simulation ◽  
Test Facility ◽  
Navier Stokes ◽  
Air Supply ◽  
Blade Row ◽  
Cooling Air

A 3D-Navier-Stokes solver was used to analyse the complete flow field of the 15-stage axial compressor of Siemens model V84.3A advanced gas turbine. The paper presents the flow simulation including modelling of rotor tip clearances and bleeds for turbine cooling air supply. All computations were performed for coupled blade rows to account for the time averaged impact of interaction effects arising from adjacent airfoil rows. The evaluation of such two-blade-row calculations allows the update of the inlet boundary conditions for the following downstream two-blade-row combination. Successive computations from inlet guide vanes to exit stator thus yield the flow field of the whole compressor. The main objective is the analysis of the numerical results. Special attention is given to the front stage, stage matching, endwall flow effects, tip leakage and the cooling air extractions. The comparison to experimental data of the full load gas turbine test facility generally shows a good agreement. The results demonstrate the reliability and power of a modern CFD tool to perform advanced design studies, geometry modifications and calibration of fast 2D-Codes more efficiently and less expensively than performing any physical experiments.

Flow Field Unsteadiness in the Tip Region of a Transonic Compressor Rotor

1997 ◽  
Vol 119 (1) ◽  
pp. 122-128 ◽  
Author(s):  
S. L. Puterbaugh ◽  
W. W. Copenhaver
Keyword(s):  
Flow Field ◽  
High Performance ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Vortex Interaction ◽  
Tip Leakage Vortex ◽  
Tip Leakage ◽  
Transonic Compressor ◽  
Compressor Rotor ◽  
Operating Points

An experimental investigation concerning tip flow field unsteadiness was performed for a high-performance, state-of-the-art transonic compressor rotor. Casing-mounted high frequency response pressure transducers were used to indicate both the ensemble averaged and time varying flow structure present in the tip region of the rotor at four different operating points at design speed. The ensemble averaged information revealed the shock structure as it evolved from a dual shock system at open throttle to an attached shock at peak efficiency to a detached orientation at near stall. Steady three-dimensional Navier Stokes analysis reveals the dominant flow structures in the tip region in support of the ensemble averaged measurements. A tip leakage vortex is evident at all operating points as regions of low static pressure and appears in the same location as the vortex found in the numerical solution. An unsteadiness parameter was calculated to quantify the unsteadiness in the tip cascade plane. In general, regions of peak unsteadiness appear near shocks and in the area interpreted as the shock-tip leakage vortex interaction. Local peaks of unsteadiness appear in mid-passage downstream of the shock-vortex interaction. Flow field features not evident in the ensemble averaged data are examined via a Navier-Stokes solution obtained at the near stall operating point.

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.

scholarly journals Simplified Method for Flow Analysis and Performance Calculation Through an Axial Compressor

1984 ◽  
Author(s):  
Hubert Miton ◽  
Youssef Doumandji ◽  
Jacques Chauvin
Keyword(s):  
Equations Of Motion ◽  
Computing Time ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Computation Method ◽  
Duct Flow ◽  
Axial Compressors ◽  
Flow Through ◽  
And Performance ◽  
Performance Calculation

This paper describes a fast computation method of the flow through multistage axial compressors of the industrial type. The flow is assumed to be axisymmetric between the blade rows which are represented by actuator disks. Blade row losses and turning are calculated by means of correlations. The equations of motion are linearized with respect to the log of static pressure, whose variation along the radius is usually of limited extent for the type of machines for which the method has been developed. In each computing plane (i.e. between the blade rows) two flows are combined: a basic flow with constant pressure satisfying the mass flow requirements and a perturbation flow fulfilling the radial equilibrium condition. The results of a few sample calculations are given. They show a satisfactory agreement with a classical duct flow method although the computing time is reduced by a factor five. The method has also been coupled with a surge line prediction calculation.

Influencing Parameters of Tip Blowing Interacting With Rotor Tip Flow

2016 ◽  
Vol 139 (2) ◽  
Author(s):  
Cyril Guinet ◽  
André Inzenhofer ◽  
Volker Gümmer
Keyword(s):  
Geometric Model ◽  
Axial Compressor ◽  
Axial Flow ◽  
Navier Stokes ◽  
Axial Position ◽  
Design Parameters ◽  
Stall Margin ◽  
Casing Treatment ◽  
Transonic Compressor ◽  
Casing Treatments

The design space of axial-flow compressors is restricted by stability issues. Different axial-type casing treatments (CTs) have shown their ability to enhance compressor stability and to influence efficiency. Casing treatments have proven to be effective, but there still is need for more detailed investigations and gain of understanding for the underlying flow mechanism. Casing treatments are known to have a multitude of effects on the near-casing 3D flow field. For transonic compressor rotors, these are more complex, as super- and subsonic flow regions alternate while interacting with the casing treatment. To derive design rules, it is important to quantify the influence of the casing treatment on the different tip flow phenomena. Designing a casing treatment in a way that it antagonizes only the deteriorating secondary flow effects can be seen as a method to enhance stability while increasing efficiency. The numerical studies are carried out on a tip-critical rotor of a 1.5-stage transonic axial compressor. The examined recirculating tip blowing casing treatment (TBCT) consists of a recirculating channel with an air off-take above the rotor and an injection nozzle in front of the rotor. The design and functioning of the casing treatment are influenced by various parameters. A variation of the geometry of the tip blowing, more specifically the nozzle aspect ratio, the axial position, or the tangential orientation of the injection port, was carried out to identify key levers. The tip blowing casing treatment is defined as a parameterized geometric model and is automatically meshed. A sensitivity analysis of the respective design parameters of the tip blowing is carried out on a single rotor row. Their impact on overall efficiency and their ability to improve stall margin are evaluated. The study is carried out using unsteady Reynolds-averaged Navier–Stokes (URANS) simulations.

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.

Sign in / Sign up

Export Citation Format

Share Document