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.

Linearized Solutions for the Supersonic Flow Through Turbomachinery Blade Rows (Using Actuator Disk Theory)

1980 ◽  
Vol 102 (3) ◽  
pp. 330-337
Author(s):  
J. H. Horlock ◽  
C. F. Grainger
Keyword(s):  
Supersonic Flow ◽  
Method Of Characteristics ◽  
Three Dimensional ◽  
Actuator Disk ◽  
Blade Row ◽  
Flow Through ◽  
Twisted Blade ◽  
Disk Method

An actuator disk method is developed for calculating the flow through the blade rows of a turbomachine in which the velocity relative to the blading may be supersonic. The method is compared with calculations of the fully supersonic flow through a twisted blade row using a three-dimensional method of characteristics.

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 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.

scholarly journals A Numerical Simulation of the Inviscid Flow Through a Counterrotating Propeller

1986 ◽  
Vol 108 (2) ◽  
pp. 187-193 ◽  
Author(s):  
M. L. Celestina ◽  
R. A. Mulac ◽  
J. J. Adamczyk
Keyword(s):  
Numerical Simulation ◽  
Flow Field ◽  
High Speed ◽  
Inviscid Flow ◽  
Solution Procedure ◽  
Numerical Results ◽  
Governing Equations ◽  
Flow Through ◽  
Coding Strategy

This paper presents the results of a numerical simulation of the time-averaged inviscid flow field through the blade rows of a multiblade row turboprop configuration. The governing equations are outlined along with a discussion of the solution procedure and coding strategy. Numerical results obtained from a simulation of the flow field through a modern high-speed turboprop will be shown.

scholarly journals Supersonic Stall Flutter of High-Speed Fans

1982 ◽  
Vol 104 (3) ◽  
pp. 675-682 ◽  
Author(s):  
J. J. Adamczyk ◽  
W. Stevans ◽  
R. Jutras
Keyword(s):  
Dynamic Response ◽  
Rotor Blade ◽  
High Speed ◽  
Aerodynamic Force ◽  
Rotor System ◽  
Flexural Mode ◽  
Disk Model ◽  
Actuator Disk ◽  
Blade Row ◽  
Flutter Boundary

An analytical model is developed for predicting the onset of supersonic stall bending flutter in axial-flow compressors. The analysis is based on a modified two-dimensional, compressible, unsteady actuator disk theory. It is applied to a rotor blade row by considering a cascade of airfoils whose geometry and dynamic response coincide with those of a rotor blade element at 85 percent of the span height (measured from the hub). The rotor blades are assumed to be unshrouded (i.e., free standing) and to vibrate in their first flexural mode. The effects of shock waves and flow separation are included in the model through quasisteady, empirical, rotor total-pressure-loss and deviation-angle correlations. The actuator disk model predicts the unsteady aerodynamic force acting on the cascade blading as a function of the steady flow field entering the cascade and the geometry and dynamic response of the cascade. Calculations show that the present model predicts the existence of a bending flutter mode at supersonic inlet Mach numbers. This flutter mode is suppressed by increasing the reduced frequency of the system or by reducing the steady-state aerodynamic loading on the cascade. The validity of the model for predicting flutter is demonstrated by correlating the measured flutter boundary of a high-speed fan stage with its predicted boundary. This correlation uses a level of damping for the blade row (i.e., the log decrement of the rotor system) that is estimated from the experimental flutter data. The predicted flutter boundary is shown to be in good agreement with the measured boundary. These results show that the model can be used to estimate the relative stability between operating points of a given rotor system. If, in addition, a measure of the mechanical damping of the rotor system is available, the model can also be used to estimate the absolute stability at an operating point.

Effects of combustor geometry on the combustion process of an RBCC combustor in high-speed ejector mode

2019 ◽  
Vol 33 (27) ◽  
pp. 1950330
Author(s):  
Taiyu Wang ◽  
Zhenguo Wang ◽  
Zun Cai ◽  
Jian Chen ◽  
Mingbo Sun ◽  
...  
Keyword(s):  
Supersonic Flow ◽  
Flow Field ◽  
Heat Release ◽  
Reaction Mechanisms ◽  
Total Pressure ◽  
High Speed ◽  
Combustion Process ◽  
Numerical Approach ◽  
Combined Cycle ◽  
Chemical Reaction Mechanisms

The combustion characteristics of high-speed ejector mode in a 2-dimensional strut-based RBCC (rocket-based combined cycle) combustor had been investigated numerically in a Mach 2.5 supersonic flow. The numerical approach had been validated by comparing numerical results with available experimental data. Besides, three different hydrogen-air chemical reaction mechanisms had also been compared. The effect of the combustor geometry on the combustion process was then discussed by analyzing the heat release distribution and flow field. It was found that the wall configuration, closeout angle of the converging location and converging ratio all have significant influences on the heat release distribution and flow field structures. It is demonstrated that a converging–diverging wall configuration is beneficial for the combustion process with significant heat release increase compared to the other wall configurations. In addition, the closeout angle of the converging location is also closely related to the combustion performance, and there exists an optimized closeout angle in a specific combustor geometry. It is also revealed that the major heat release region moves upstream obviously with increase in the converging ratio, leading to an enhanced combustion process. However, the converging ratio is still to be optimized to keep a balance between heat release increase and total pressure loss of the supersonic flow.

The Measured and Visualized Behavior of Rotating Stall in an Axial-Flow Compressor and in a Two-Dimensional Cascade

1959 ◽  
Vol 81 (1) ◽  
pp. 24-34 ◽  
Author(s):  
Gino Sovran
Keyword(s):  
Flow Field ◽  
Axial Flow ◽  
Rotating Stall ◽  
Operating Conditions ◽  
Axial Position ◽  
Compressor Cascade ◽  
Axial Flow Compressor ◽  
Blade Row ◽  
The Absolute ◽  
Flow Compressor

The technique of smoke-flow visualization has been used to show clearly the action of the flow field upstream of a single rotor during rotating stall. The flow processes on the blades and in the blade passages of a stationary compressor cascade also have been observed using the same technique. An audio method of detecting rotating-stall patterns has been developed and has indicated that some compressor operating conditions at which no periodic flow disturbance was previously thought to occur actually contained rotating-stall patterns whose number of stalled regions changed very quickly from one value to another. The absolute speed of rotation of a stall zone was reduced to zero and its over-all shape clearly outlined by means of smoke visualization. The qualitative results of all these investigations have given a good physical picture of the rotating-stall phenomenon. Severe reverse flows were found to exist during rotating stall and these caused the formation of reversed flow regions extending upstream of a blade row. Furthermore, the distortions of the flow field in the vicinity of a blade row were found to be of such large magnitude that it does not seem likely that they can be described adequately by any linearized theory. Quantitative investigations were made into the effects of guide-vane turning and axial position on the rotating-stall characteristics of an axial-flow-compressor rotor. The absolute direction of the fluid entering such a configuration was found to affect nearly all facets of rotating-stall behavior.

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.

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.

The Effects of Unsteadiness and Compressibility on the Interaction Between Hub Leakage and Mainstream Flows in High-Pressure Turbines

2013 ◽  
Vol 135 (6) ◽  
Author(s):  
Ivan Popović ◽  
Howard P. Hodson ◽  
Erik Janke ◽  
Torsten Wolf
Keyword(s):  
Flow Field ◽  
High Speed ◽  
Leakage Flow ◽  
Average Flow ◽  
Linear Cascade ◽  
Low Speed ◽  
Present Test ◽  
Blade Row ◽  
Time Average ◽  
Unsteady Effects

This paper investigates the effects of compressibility and unsteadiness due to the relative blade row motion and their importance in the interaction between hub leakage (purge) and mainstream flows. First, the challenges associated with the blade redesign for low-speed testing are described. The effects of Mach number are then addressed by analyzing the experiments in the low-speed linear cascade equipped with the secondary airflow system and computations performed on the low- and high-speed blade profiles. These results indicate that the compressibility does not significantly affect the interaction between the leakage and mainstream flows despite a number of compromises made during the design of the low-speed blade. This was due to the fact that the leakage–mainstream interaction takes place upstream of the blade throat where the local Mach numbers are still relatively low. The analysis is then extended to the equivalent full-stage unsteady computations. The periodic unsteadiness resulting from the relative motion of the upstream vanes appreciably affected the way in which the leakage flow is injected and the rotor flow field in general. However, the time-average flow field was still found to be dominated by the rotor blade's potential field. For the present test arrangement, the unsteady effects were not very detrimental and caused less than a 10% increase in the losses due to the leakage injection relative to the steady calculations.

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