Supersonic Wave/Blade-Row Interactions Establish Boundary Conditions for Unsteady Inlet Flows

2002 ◽  
Vol 18 (5) ◽  
pp. 1126-1128 ◽  
Author(s):  
Miklos Sajben
Keyword(s):  
Boundary Conditions ◽  
Blade Row

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.

Effect of Bleed Flows on Flutter and Forced Response of Core Compressors

2006 ◽  
Author(s):  
Luca di Mare ◽  
George Simpson ◽  
Bernhard Mueck ◽  
Abdulnaser I. Sayma
Keyword(s):  
Boundary Conditions ◽  
Flow Pattern ◽  
Forced Response ◽  
Accurate Prediction ◽  
Aeroelastic Response ◽  
Axial Compressors ◽  
Blade Row

This paper presents a methodology for the modeling of flutter and forced response in axial compressors while taking into account the effect of bleed off-takes. Usually, aeroelasticity analyses are performed assuming smooth solid end walls. This type of analysis has two main shortcomings. Firstly, it does not account for the change in the aerodynamic speed of the stages downstream of the bleed off-take, so that aeroelasticity analyses are not performed at the correct aerodynamic conditions. Secondly, bleed off-takes influence the flow pattern particularly in the stages around or close to the bleed off-take, thus leading to possibility of obtaining the wrong aeroelastic response. Another objective of this paper is to present a methodology for the accurate prediction of the flow in a compressor with bleed off-take, by both including the geometry of the bleed off-take and performing the calculations on the entire compressor, thus eliminating errors resulting from prescribing boundary conditions at inter-blade row boundaries. It is concluded that bleed off-takes could influence significantly the aeroelastic response of the blades.

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.

scholarly journals A Simplified Method for 3-D Potential Flow in Turbomachinery Using Vortex Sheet Boundary Conditions

1987 ◽  
Author(s):  
H. Jiang ◽  
R. Cai ◽  
Y. Zhu
Keyword(s):  
Boundary Conditions ◽  
Potential Flow ◽  
Three Dimensional ◽  
Vortex Sheet ◽  
Inviscid Flow ◽  
Three Dimensional Flow ◽  
Blade Row ◽  
Simplified Method ◽  
Outlet Flow ◽  
Two Sides

Within the framework of inviscid flow theory, the character of three-dimensional flow in turbomachinery blade row is discussed. One of the important differences between 3-D and 2-D flow in turbomachinery is the discontinuity of velocity at the two sides of trailing edge and across downstream boundary. The inconsistency of the traditional periodicity conditions for downstream boundary and of the axisymmetric assumption for outlet flow with the three-dimensionality of turbomachinery flow is discussed also. For 3-D potential flow, the vortex sheet boundary conditions (VSBC) for downstream boundary and a fully 3-D condition for outlet flow are presented. A simplified method is developed by implementation of VSBC on a fixed vortex boundary in order to predict the fully 3-D flow in blade passage as well as downstream of blade row. In the present investigation two calculations are carried out. In one calculation the traditional boundary conditions are imposed while in another one the VSBC are used to demonstrate the capability of the newly develped boundary conditions. The agreement between some calculated results and the theoretical analysis is very well.

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.

scholarly journals Computation of 3D Flow Phenomena in Axial Flow Compressor Blade Rows

1991 ◽  
Author(s):  
M. Janssen ◽  
H.-J. Dohmen ◽  
K. G. Grahl
Keyword(s):  
Flow Field ◽  
Boundary Conditions ◽  
Axial Flow ◽  
Compressor Blade ◽  
Numerical Results ◽  
Curvilinear Coordinates ◽  
Simple Method ◽  
Flow Development ◽  
Blade Row ◽  
Flow Effects

The main subject of the present publication is the comparison of results achieved with a 3D-partially parabolic calculation procedure and experimental data for the three dimensional flow in stationary and rotating blade rows of axial flow compressors. To set up the numerical solution procedure, the Navier-Stokes Equations are written in finite difference form by applying the control-volume approach. The turbulent flow effects are taken into account by using the standard k—ε model for the calculation of the turbulent viscosity. For precisely introducing the boundary conditions for arbitrary geometries, the differential equations are transformed to a body-fitted coordinate system by a very simple method. To construct the physical mesh, the nonorthogonal curvilinear coordinates are taken as solutions of a suitable elliptic boundary value problem. The abilities of the developed computer program are shown by comparing experimental and numerical results for three applications. The first, most simple case deals with the flow development in an isolated, stationary blade row of cylindrical blades and uniform boundary conditions upstream of the blade row. A more complex flow is regarded by calculating the flow field through highly turned, custom tailored airfoils working in a multistage environment. The flow conditions upstream of the flow domain under consideration show a well developed end wall boundary layer at the hub, leading to a strongly skewed inflow due to the superimposed tangential velocity component of the rotor upstream. The third application regards the flow development in a rotating axial compressor blade row in which the complexity of the flow field increases by flow effects that are due to centrifugal and Coriolis forces. The comparisons between experimental and numerical results show good agreements for all applications.

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

1999 ◽  
Vol 121 (1) ◽  
pp. 1-10 ◽  
Author(s):  
R. Emunds ◽  
I. K. Jennions ◽  
D. Bohn ◽  
J. 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 three-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. First, 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 blade-row tests; the other code is currently only suitable for a single blade row. Second, the 1.5-stage results are compared to the experimental data and promote discussion of surrounding blade row effects on multistage solutions.

Use of the Magnetostatic Analogue In Electromagnetic Lens Engineering

1970 ◽  
Vol 28 ◽  
pp. 360-361
Author(s):  
John W. Coleman
Keyword(s):  
Boundary Conditions ◽  
High Performance ◽  
Force Fields ◽  
Optical Design ◽  
Direct Conversion ◽  
Design Engineering ◽  
Design Data ◽  
Scalar Potentials ◽  
Geometric Elements ◽  
At Will

In the design engineering of high performance electromagnetic lenses, the direct conversion of electron optical design data into drawings for reliable hardware is oftentimes difficult, especially in terms of how to mount parts to each other, how to tolerance dimensions, and how to specify finishes. An answer to this is in the use of magnetostatic analytics, corresponding to boundary conditions for the optical design. With such models, the magnetostatic force on a test pole along the axis may be examined, and in this way one may obtain priority listings for holding dimensions, relieving stresses, etc..The development of magnetostatic models most easily proceeds from the derivation of scalar potentials of separate geometric elements. These potentials can then be conbined at will because of the superposition characteristic of conservative force fields.

Sign in / Sign up

Export Citation Format

Share Document