Viscous-Flow Two-Dimensional Analysis Including Secondary Flow Effects

2000 ◽  
Vol 123 (3) ◽  
pp. 558-567 ◽  
Author(s):  
Reinhard Mo¨nig ◽  
Frank Mildner ◽  
Ralf Ro¨per
Keyword(s):  
Secondary Flow ◽  
Boundary Layers ◽  
Gas Turbines ◽  
Axial Flow ◽  
Tip Clearance ◽  
New Method ◽  
Navier Stokes ◽  
Flow Effects ◽  
End Wall ◽  
Axial Flow Compressors

During the last few decades extremely powerful Quasi-three-dimensional (3D) codes and fully 3D Navier-Stokes solvers have been developed and successfully utilized in the design process and optimization of multistage axial-flow compressors. However, most of these methods proved to be difficult in handling and extremely time consuming. Due to these disadvantages, the primary stage design and stage matching as well as the off-design analysis is nowadays still based on fast 2D methods incorporating loss-, deviation- and end wall modeling. Only the detailed 3D optimization is normally performed by means of advanced 3D methods. In this paper a fast and efficient 2D calculation method is presented, which already in the initial design phase of multistage axial flow compressors, considers the influence of hub leakage flows, tip clearance effects, and other end wall flow phenomena. The method is generally based on the fundamental approach by Howard and Gallimore (1992). In order to allow a more accurate prediction of skewed and nondeveloped boundary layers in turbomachines, an improved theoretical approach was implemented. Particularly the splitting of the boundary layers into an axial and tangential component proved to be necessary in order to account for the change between rotating and stationary end walls. Additionally, a new approach is used for the prediction of the viscous end wall zones including hub leakage effects and strongly skewed boundary layers. As a result, empirical correlations for secondary flow effects are no longer required. The results of the improved method are compared with conventional 2D results including 3D loss- and deviation-models, with experimental data of a three-stage research compressor of the Institute for Jet Propulsion and Turbomachinery of the Technical University of Aachen and with 3D Navier-Stokes solutions of the V84.3A compressor and of a multistage Siemens research compressor. The results obtained using the new method show a remarkable improvement in comparison with conventional 2D methods. Due to the high quality and the extremely short computation time, the new method allows an overall viscous design of multistage compressors for heavy duty gas turbines and aeroengine applications.

Viscous-Flow 2D-Analysis Including Secondary Flow Effects

2000 ◽  
Author(s):  
Reinhard Mönig ◽  
Frank Mildner ◽  
Ralf Röper
Keyword(s):  
Secondary Flow ◽  
Boundary Layers ◽  
Gas Turbines ◽  
Axial Flow ◽  
Tip Clearance ◽  
New Method ◽  
Navier Stokes ◽  
Flow Effects ◽  
End Wall ◽  
Axial Flow Compressors

During the last few decades extremely powerful Quasi-3D codes and fully 3D Navier-Stokes solvers have been developed and successfully utilized in the design process and optimization of multistage axial-flow compressors. However, most of these methods proved to be difficult in handling and extremely time consuming. Due to these disadvantages, the primary stage design and stage matching as well as the off-design analysis is nowadays still based on fast 2D methods incorporating loss-, deviation- and end wall modeling. Only the detailed 3D optimization is normally performed by means of advanced 3D methods. In this paper a fast and efficient 2D calculation method is presented, which already in the initial design phase of multistage axial flow compressors considers the influence of hub leakage flows, tip clearance effects and other end wall flow phenomena. The method is generally based on the fundamental approach by Howard and Gallimore (1992). In order to allow a more accurate prediction of skewed and non-developed boundary layers in turbomachines an improved theoretical approach was implemented. Particularly the splitting of the boundary layers into an axial and tangential component proved to be necessary in order to account for the change between rotating and stationary end walls. Additionally, a new approach is used for the prediction of the viscous end wall zones including hub leakage effects and strongly skewed boundary layers. As a result, empirical correlations for secondary flow effects are no longer required. The results of the improved method are compared with conventional 2D-results including 3D loss- and deviation-models, with, experimental data of a 3-stage research compressor of the Institute for Jet Propulsion and Turbomachinery of the Technical University of Aachen and with 3D Navier-Stokes solutions of the V84.3A compressor and of a multi-stage Siemens research compressor. The results obtained using the new method show a remarkable improvement in comparison with conventional 2D-methods. Due to the high quality and the extremely short computation time the new method allows an overall viscous design of multistage compressors for heavy duty gas turbines and aeroengine applications.

scholarly journals An Investigation of Spanwise Mixing in Multistage Axial Flow Compressors

1992 ◽  
Author(s):  
S.-M. Li ◽  
M.-Z. Chen
Keyword(s):  
Secondary Flow ◽  
Turbulent Diffusion ◽  
Flow Model ◽  
Three Dimensional ◽  
Axial Flow ◽  
Equation System ◽  
Meridional Flow ◽  
Flow Effects ◽  
Annular Cascade ◽  
Axial Flow Compressors

An equation system has been deduced for meridional throughflow fields of multistage axial flow compressors, presenting different kinds of spanwise mixing effects of the fields in a unified form. The spanwise mixing in compressors is caused by three kinds of effects, molecular motion, turbulent diffusion, and circumferential non–uniformities, the last of which includes secondary flow effects and others. This equation system thus unifies the two models for spanwise mixing analyses by Adkins & Smith (1981) and Gallimore & Cumpsty (1986). The turbulent diffusion in the two–dimensional (2–D) meridional fields is determined by complex three–dimensional (3–D) shear flows in compressors, rather than the 2–D shearing alone, so a turbulence model for 2–D meridional flow calculations is proposed on the basis of a simplified 3–D shearing structure in compressors. The circumferentially non–uniform correlation terms in the equation system have been modeled on the basis of Adkins and Smith (1981) secondary flow model and the experimental data for annular cascade wakes. The results obtained agree well with the experiments for five compressors. The results also show some improvement over the previous theories.

Spanwise Mixing in Axial-Flow Turbomachines

1982 ◽  
Vol 104 (1) ◽  
pp. 97-110 ◽  
Author(s):  
G. G. Adkins ◽  
L. H. Smith
Keyword(s):  
Secondary Flow ◽  
Boundary Layers ◽  
Axial Flow ◽  
Separated Flow ◽  
Secondary Flows ◽  
Main Stream ◽  
Flow Measurements ◽  
Mixing Coefficient ◽  
Blade Row ◽  
End Wall

Flow measurements taken in multistage axial-flow turbomachines suggest that substantial spanwise mixing of flow properties often occurs. In addition, measured blade row turnings often show considerable deviation from two-dimensional cascade theory, particularly in the end-wall regions. An approximate method is presented with which both of these effects can be included in design through-flow calculations. The method is based on inviscid, small-perturbation secondary flow theory. Frictional effects are not directly included but secondary flows caused by annulus wall and blade boundary layers are included in an approximate way. The secondary flow model includes effects of 1) main-stream nonfree-vortex flow, 2) end-wall boundary layers, 3) blade end clearances, 4) blade end shrouding, and 5) blade boundary layer and wake centrifugation. The spanwise mixing phenomenon is modeled as a diffusion process, where the mixing coefficient is related to the calculated spanwise secondary velocities. Empirical adjustments are employed to account for the dissipation of the secondary velocities and interactions with downstream blade rows. The induced blade row overturnings are related to the calculated cross-passage secondary velocities. The nature of the assumptions employed restricts the method to design-point-type applications for which losses are relatively small and significant regions of separated flow are not present.

End-Wall Film Cooling Through Fan-Shaped Holes With Different Area Ratios

2006 ◽  
Vol 129 (2) ◽  
pp. 212-220 ◽  
Author(s):  
Giovanna Barigozzi ◽  
Giuseppe Franchini ◽  
Antonio Perdichizzi
Keyword(s):  
Secondary Flow ◽  
Mass Flow ◽  
Film Cooling ◽  
Three Dimensional ◽  
Wide Band ◽  
Lateral Spreading ◽  
Area Ratio ◽  
Adiabatic Effectiveness ◽  
Flow Effects ◽  
End Wall

The present paper reports on the aerothermal performance of a nozzle vane cascade, with film-cooled end walls. The coolant is injected through four rows of cylindrical holes with conical expanded exits. Two end-wall geometries with different area ratios have been compared. Tests have been carried out at low speed (M=0.2), with coolant to mainstream mass flow ratio varied in the range 0.5–2.5%. Secondary flow assessment has been performed through three-dimensional (3D) aerodynamic measurements, by means of a miniaturized five-hole probe. Adiabatic effectiveness distributions have been determined by using the wide-band thermochromic liquid crystals technique. For both configurations and for all the blowing conditions, the coolant share among the four rows has been determined. The aerothermal performances of the cooled vane have been analyzed on the basis of secondary flow effects and laterally averaged effectiveness distributions; this analysis was carried out for different coolant mass flow ratios. It was found that the smaller area ratio provides better results in terms of 3D losses and secondary flow effects; the reason is that the higher momentum of the coolant flow is going to better reduce the secondary flow development. The increase of the fan-shaped hole area ratio gives rise to a better coolant lateral spreading, but appreciable improvements of the adiabatic effectiveness were detected only in some regions and for large injection rates.

Predictions of Endwall Losses and Secondary Flows in Axial Flow Turbine Cascades

1987 ◽  
Vol 109 (2) ◽  
pp. 229-236 ◽  
Author(s):  
O. P. Sharma ◽  
T. L. Butler
Keyword(s):  
Flow Visualization ◽  
Secondary Flow ◽  
Axial Flow ◽  
Secondary Flows ◽  
Integral Boundary ◽  
Realistic Interpretation ◽  
Proposed Model ◽  
Flow Theories ◽  
Semi Empirical ◽  
End Wall

This paper describes the development of a semi-empirical model for estimating end-wall losses. The model has been developed from improved understanding of complex endwall secondary flows, acquired through review of flow visualization and pressure loss data for axial flow turbomachine cascades. The flow visualization data together with detailed measurements of viscous flow development through cascades have permitted more realistic interpretation of the classical secondary flow theories for axial turbomachine cascades. The re-interpreted secondary flow theories together with integral boundary layer concepts are used to formulate a calculation procedure for predicting losses due to the endwall secondary flows. The proposed model is evaluated against data from published literature and improved agreement between the data and predictions is demonstrated.

A Method for the Calculation of the Tip Clearance Flow Effects in Axial Flow Compressors: Part II — Calculation Procedure

1993 ◽  
Author(s):  
I. K. Nikolos ◽  
D. I. Douvikas ◽  
K. D. Papailiou
Keyword(s):  
Axial Flow ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Vorticity Distribution ◽  
Clearance Flow ◽  
Calculation Results ◽  
Flow Effects ◽  
Set Up ◽  
Induced Velocity ◽  
Theoretical Procedure

An algorithm was set up for the implementation of the tip clearance models, described in Part I, in a secondary flow calculation method. A complete theoretical procedure was, thus, developed, which calculates the circumferentially averaged flow quantities and their radial variation due to the tip clearance effects. The calculation takes place in successive planes, where a Poisson equation is solved in order to provide the kinematic field. The self induced velocity is used for the positioning of the leakage vortex and a diffusion model is adopted for the vorticity distribution. The calculated pressure deficit due to the vortex presence is used, through an iterative procedure, in order to modify the pressure difference in the tip region. The method of implementation and the corresponding algorithm are described in this part of the paper and calculation results are compared to experimental ones for cascades and single rotors. The agreement between theory and experiment is good.

Reduction of Secondary Flow Effects in a Linear Cascade by Use of an Air Suction From the Endwall

1996 ◽  
Author(s):  
K. Funazaki ◽  
T. Endo ◽  
T. Tanuma
Keyword(s):  
Secondary Flow ◽  
Flow Rate ◽  
Pressure Measurement ◽  
Gas Turbines ◽  
Stagnation Pressure ◽  
Working Fluid ◽  
Steam Turbines ◽  
Linear Cascade ◽  
Oil Flow ◽  
Flow Effects

The objective of this study is to reduce secondary flow effects in a linear cascade by sucking the working fluid from the endwall. It is widely known that the secondary flow developed in a cascade has a significant impact on the cascade loss or blade erosion in steam turbines. Therefore, a number of studies have been made on the physics of the secondary flow and several devices to control the secondary flow, such as a fence, have been examined. In this study, considering the application to nozzles in gas turbines or steam turbines, the air suction approach is investigated for reducing the secondary flow effects. A suction slit is provided on the lower endwall of the cascade and a flow rate of the sucked air is controlled by adjusting the exit pressure of the slit. The effects of the suction upon the flow nearby the endwall and the secondary flow are observed through several flow visualizing techniques, for example an oil flow method or a tuft method. Furthermore, velocity and stagnation pressure measurement are conducted by a five-hole pressure tube. This clearly demonstrates the vorticity and loss profiles downstream of the cascade with and without the endwall suction.

Leakage Effects in the Rotor Tip-Clearance Region of a Multistage Axial Compressor: Part 2 — Numerical Modelling

1998 ◽  
Author(s):  
E. S. Politis ◽  
K. C. Giannakoglou ◽  
K. D. Papailiou
Keyword(s):  
Axial Compressor ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Computational Domain ◽  
Loose Coupling ◽  
Tip Clearance Flow ◽  
Clearance Flow ◽  
Flow Effects ◽  
Multistage Axial Compressor ◽  
Asme Paper

Innovative measurements of tip-clearance flow for the 3rd stage rotor embedded in a four stage Low-Speed Research Compressor are presented in the companion ASME paper. Here, in Part 2, the rotor flow is numerically simulated through a Navier-Stokes solver implementing the k-ε turbulence model. The 3rd stage rows are considered as discrete parts of the same computational domain and the flow in each one of them is treated as steady in the corresponding system of reference. An iterative, though loose, coupling between the rotor exit and the stator inlet is established by artificially increasing the inter-row distance. To model tip-clearance flow effects with sufficient accuracy, a two-block grid system per row is used. Comparisons with measurements published in Part 1 for the average flow quantities at the exit of both rows are presented. Row patterns close to the rotor tip-clearance region are illustrated.

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