scholarly journals Comparison of Calculated and Experimental Cascade Performance for Controlled-Diffusion Compressor Stator Blading

1986 ◽  
Vol 108 (1) ◽  
pp. 42-50 ◽  
Author(s):  
N. L. Sanger ◽  
R. P. Shreeve
Keyword(s):  
Aspect Ratio ◽  
Flow Field ◽  
Boundary Layers ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Surface Boundary ◽  
Performance Parameters ◽  
Blade Surface ◽  
Controlled Diffusion ◽  
Inlet Angle

The midspan section of a previously reported controlled-diffusion compressor stator has been experimentally evaluated in cascade. Measurements were taken over a range of incidence angles for blade chord Reynolds numbers from 470,000 to 690,000. Blade chord length was 12.7 cm, aspect ratio was 2.0, and solidity was 1.67. Measurements included conventional cascade performance parameters as well as blade surface pressures. Computations were made for the inviscid flow field, surface boundary layers, and loss for several of the blade inlet angle conditions, and compared against corresponding data.

Axial-Flow Compressor Turning Angle and Loss by Inviscid-Viscous Interaction Blade-to-Blade Computation

1980 ◽  
Vol 102 (1) ◽  
pp. 28-34 ◽  
Author(s):  
E. C. Hansen ◽  
G. K. Serovy ◽  
P. M. Sockol
Keyword(s):  
Boundary Layers ◽  
Axial Flow ◽  
Inviscid Flow ◽  
Interactive System ◽  
Surface Boundary ◽  
Blade Surface ◽  
Axially Symmetric ◽  
Viscous Interaction ◽  
Interaction Scheme ◽  
Flow Compressor

A method for computation of the flow field around an arbitrary airfoil cascade on an axially symmetric blade-to-blade surface was developed which takes into account the development and separation of the blade surface boundary layers and mixing in the wake. The method predicts the overall fluid turning and total pressure loss in the context of an inviscid-viscous interaction scheme. The inviscid flow solution is obtained from a compressible flow matrix method. The viscous flow is obtained from a differential boundary layer method which calculates laminar, transitional and turbulent boundary layers. Provisions for the calculation of laminar and turbulent separation regions were added to the viscous scheme. The combined inviscid-viscous interaction scheme described yields results which are quantitatively consistent with experimental data. This suggests that the physical basis for the interactive system is correct and justifies continued exploration and use of the method.

Essential Ingredients for the Computation of Steady and Unsteady Blade Boundary Layers

1991 ◽  
Vol 113 (4) ◽  
pp. 608-616 ◽  
Author(s):  
H. M. Jang ◽  
J. A. Ekaterinaris ◽  
M. F. Platzer ◽  
T. Cebeci
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Stokes Equations ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Navier Stokes ◽  
Transitional Region ◽  
Low Reynolds Numbers ◽  
Flow Equations ◽  
Low Reynolds

Two methods are described for calculating pressure distributions and boundary layers on blades subjected to low Reynolds numbers and ramp-type motion. The first is based on an interactive scheme in which the inviscid flow is computed by a panel method and the boundary layer flow by an inverse method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier–Stokes equations with an embedded grid technique that permits accurate calculation of boundary layer flows. Studies for the Eppler-387 and NACA-0012 airfoils indicate that both methods can be used to calculate the behavior of unsteady blade boundary layers at low Reynolds numbers provided that the location of transition is computed with the en method and the transitional region is modeled properly.

Numerical Study of Winglet at Low Reynolds Numbers

2013 ◽  
Author(s):  
Sina Pooladsanj ◽  
Mehran Tadjfar
Keyword(s):  
Aspect Ratio ◽  
Flow Field ◽  
Vortex Flow ◽  
Numerical Study ◽  
Reynolds Numbers ◽  
Flow Zone ◽  
Low Reynolds Numbers ◽  
Structured Grid ◽  
Flow Field Characteristics ◽  
Wing Aspect Ratio

A numerical study has been performed to evaluate the aerodynamics coefficients of a winglet in the range of Reynolds numbers below 30,000. In this study some parameters on winglet design have been considered. The effect of winglet-tip airfoil thickness has been investigated on aerodynamics coefficients. In order to explore this effect, two different airfoils (NACA0002 and NACA0012) were employed at the winglet-tip. The influence of varying the winglet connection angle to the wing on aerodynamics coefficients and flow field characteristics in the vortex flow zone such as; circulation magnitude and vorticity magnitude in the vortex core have been studied. Six connection angles including 20°, 30°, 40°, 50°, 60° and 70° have been studied. Negative values of these angles have also been considered. In addition, the effect of changing wing aspect ratio on aerodynamics coefficients has been investigated. To solve the flow field around the studied geometry a fully structured grid was used which consists of 84 blocks.

scholarly journals The Use of Optimization Techniques to Design Controlled Diffusion Compressor Blading

1982 ◽  
Author(s):  
N. L. Sanger
Keyword(s):  
Optimization Procedure ◽  
Inviscid Flow ◽  
General Purpose ◽  
Optimization Techniques ◽  
Conjugate Directions ◽  
Surface Boundary ◽  
Controlled Diffusion ◽  
Constrained Problems ◽  
Blade Section ◽  
Stator Blade

A method is presented for automating compressor blade design using numerical optimization, and is applied to the design of a controlled diffusion stator blade row. A general purpose optimization procedure is employed, which is based on conjugate directions for locally unconstrained problems and on feasible directions for locally constrained problems. Coupled to the optimizer is an analysis package consisting of three analysis programs which calculate blade geometry, inviscid flow, and blade surface boundary layers. The optimization concepts are briefly discussed. Selection of design objective and constraints is described. The procedure for automating the design of a two-dimensional blade section is discussed, and design results are presented.

scholarly journals Essential Ingredients for the Computation of Steady and Unsteady Blade Boundary Layers

1990 ◽  
Author(s):  
H. M. Jang ◽  
M. F. Platzer ◽  
J. A. Ekaterinaris ◽  
T. Cebeci
Keyword(s):  
Boundary Layer ◽  
Boundary Layers ◽  
Stokes Equations ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Navier Stokes ◽  
Transitional Region ◽  
Low Reynolds Numbers ◽  
Flow Equations ◽  
Low Reynolds

Two methods are described for calculating pressure distributions and boundary layers on blades subjected to low Reynolds numbers and ramp–type motion. The first is based on an interactive scheme in which the inviscid flow is computed by a panel method and the boundary layer flow by an inverse method that makes use of the Hilbert integral to couple the solutions of the inviscid and viscous flow equations. The second method is based on the solution of the compressible Navier–Stokes equations with an embedded grid technique that permits accurate calculation of boundary layer flows. Studies for the Eppler and NACA–0012 airfoils indicate that both methods can be used to calculate the behavior of unsteady blade boundary layers at low Reynolds numbers provided that the location of transition is computed with the en–method and the transitional region is modelled properly.

Analysis of the Airfoil Stall With a Modification of Viscous-Inviscid Interaction Concept

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
E. L. Amromin
Keyword(s):  
Boundary Layers ◽  
Separation Bubble ◽  
Inviscid Flow ◽  
Reynolds Numbers ◽  
Laminar Separation ◽  
Laminar Boundary Layers ◽  
Integral Methods ◽  
Turbulent Transition ◽  
Turbulent Separation ◽  
Separation Zones

A modification of the viscous-inviscid interaction concept with the employment of coupled vortices around the airfoil wake is introduced for analyzing the airfoil stall. The analyzed flow includes the laminar boundary layers, laminar separation bubble, laminar-turbulent transition zone, turbulent boundary layers, turbulent separation zone, wake, and outer inviscid flow. Integral methods are employed for the boundary layers. The boundaries of separation zones are analyzed as free surfaces, however, their lengths and shapes depend on the Reynolds number. The described modification is validated by a comparison of the numerical results with the previously published experimental data for various airfoils and Reynolds numbers at low Mach numbers. This modification achieves a reasonably good agreement of the computed lift and moment coefficients with their measured values.

scholarly journals The Use of Optimization Techniques to Design-Controlled Diffusion Compressor Blading

1983 ◽  
Vol 105 (2) ◽  
pp. 256-264 ◽  
Author(s):  
N. L. Sanger
Keyword(s):  
Optimization Procedure ◽  
Inviscid Flow ◽  
General Purpose ◽  
Optimization Techniques ◽  
Conjugate Directions ◽  
Surface Boundary ◽  
Controlled Diffusion ◽  
Constrained Problems ◽  
Blade Section ◽  
Stator Blade

A method is presented for automating compressor blade design using numerical optimization and is applied to the design of a controlled diffusion stator blade row. A general purpose optimization procedure is employed, which is based on conjugate directions for locally unconstrained problems and on feasible directions for locally constrained problems. Coupled to the optimizer is an analysis package consisting of three analysis programs which calculate blade geometry, inviscid flow, and blade surface boundary layers. The optimization concepts are briefly discussed. Selection of design of a two-dimensional blade section is discussed, and design results are presented.

Studies of the flow field near a IMACA 4412 aerofoil at nearly maximum lift

1987 ◽  
Vol 91 (901) ◽  
pp. 29-44 ◽  
Author(s):  
R. C. Hastings ◽  
B. R. Williams
Keyword(s):  
Flow Field ◽  
Inviscid Flow ◽  
Surface Boundary ◽  
Reynolds Stresses ◽  
Surface Boundary Layer ◽  
Laser Anemometry ◽  
Pressure Distributions ◽  
Two Dimensional Flow ◽  
Integral Treatment ◽  
Momentum Integral

Summary Measurements made at a Mach number of 0.18 and a chord-based Reynolds number of 4·2 x 106 on a constant-chord model having a NACA 4412 aerofoil section are described and compared with the results of flow field calculations. Both the experimental arrangement and the difficulties initially experienced in achieving an adequate approximation to two-dimensional flow above the wing are briefly outlined. The measurements include static pressure distributions on the wing surface and on the wind-tunnel walls above and below the mid-span section of the wing. The main emphasis in the experiment was, however, on defining the development of the upper surface boundary layer through separation (at about 20% chord ahead of the trailing-edge) and on into the wake, making extensive use of laser anemometry to measure mean velocities. In addition, Reynolds stresses were measured in certain parts of the flow field by hot-wire anemometry. The flow field calculations are of the semi-inverse kind in which an inverse momentum-integral treatment of the shear flow, used to avoid difficulties at separation, is coupled to a direct solution of the inviscid flow problem. The main features of the method are outlined.

The Influence of Blade Sweep Technique in Linear Cascade Configuration

2011 ◽  
Author(s):  
Rosario Spataro ◽  
Gabriele D’Ippolito ◽  
Vincenzo Dossena
Keyword(s):  
Aspect Ratio ◽  
Flow Field ◽  
Surface Pressure ◽  
Turbine Blades ◽  
Blade Surface ◽  
Aerodynamic Optimization ◽  
Blow Down ◽  
Linear Cascade ◽  
Low Aspect Ratio ◽  
Design Techniques

One key issue in the advanced aerodynamic optimization of turbomachinery involves the application of 3D blade design techniques. The complex shape of the resulting blades is often a combination of simple techniques such as sweep. Such a blade arrangement is often imposed to the designer by structural constrains, space reduction needs, diameter optimization or spanwise blade loading control. This work aims to study the aerodynamic effects produced on turbine passages by blade sweep; with this term we refer to a configuration where the flow mainstream direction and the blade stacking axis are not orthogonal. A linear cascade of turbine blades, obtained by stacking the same profile with a sweep angle of 20 degrees, was investigated in a blow down facility at an isentropic downstream Mach number of 0.65. Standing the low aspect ratio of the cascade, the blade was built by simply shifting in axial direction the 2D profile originally used in the reference prismatic blade. The choice to build the swept blade keeping the same 2D section parallel to the incident flow was considered taking into account the blade low aspect ratio. Measurements were performed by means of blade surface pressure taps and five holes probe traversing downstream of the cascade; oil and dye flow visualizations were also performed to study the effects on the secondary vortices evolution inside of the passage. Moreover, a commercial CFD code was applied to provide information on the flow field all along the passage. The same profile was already extensively investigated both by measurements and CFD calculations [1, 2] in order to clarify the effects of blade lean and bowing. This additional paper gives a final contribution addressed to deeply understand the aerodynamic effects produced on turbine cascade flow field by the separate application of each one of the typical 3D design techniques. Results from both the experimental and computational investigations are presented and discussed in the paper where a phenomenological approach has been preferred. Measurements of the blade surface pressure distribution, performed at several blade heights, support the analysis of the pressure field inside of the passage which is mainly based on numerical results. In particular, the paper shows the influence of pressure contours shape on streamwise vorticity inside and downstream of the passage focusing the attention on secondary structures. The downstream vorticity field is then discussed together with the loss distribution in the same region to provide a more exhaustive description.

Compliant Coatings: A Decade of Progress

1996 ◽  
Vol 49 (10S) ◽  
pp. S147-S157 ◽  
Author(s):  
Mohamed Gad-el-Hak
Keyword(s):  
Boundary Layers ◽  
Computational Methods ◽  
Reynolds Numbers ◽  
Turbulent Boundary Layers ◽  
Surface Boundary ◽  
Rigid Surface ◽  
The Past ◽  
Marine Vehicles ◽  
Order Of Magnitude ◽  
Suppress Noise

This brief article reviews the important developments in the field of compliant coatings that took place in the past ten years. During this period progress in theoretical and computational methods somewhat outpaced that in experimental efforts. There is no doubt that compliant coatings can be designed to delay transition and to suppress noise on marine vehicles as well as other practical hydrodynamic devices. Transition Reynolds numbers that exceed by an order of magnitude those on rigid-surface boundary layers can be achieved. There is renewed evidence of favorable interactions of compliant coatings even for air flows and even for turbulent boundary layers, but more research is needed to confirm these latest results.

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