scholarly journals Analysis of Fan Stage Conceptual Design Attributes for Boundary Layer Ingestion

2017 ◽  
Vol 139 (7) ◽  
Author(s):  
D. K. Hall ◽  
E. M. Greitzer ◽  
C. S. Tan
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Design Flow ◽  
Surface Geometry ◽  
Local Flow ◽  
Performance Loss ◽  
Inlet Distortion ◽  
Design Attributes ◽  
Stagnation Enthalpy

This paper describes a new conceptual framework for three-dimensional turbomachinery flow analysis and its use to assess fan stage attributes for mitigating adverse effects of inlet distortion due to boundary layer ingestion (BLI). A nonaxisymmetric throughflow analysis has been developed to define fan flow with inlet distortion. The turbomachinery is modeled using momentum and energy source distributions that are determined as a function of local flow conditions and specified blade camber surface geometry. Comparison with higher-fidelity computational and experimental results shows the analysis captures the principal flow redistribution and distortion transfer effects associated with BLI. Distortion response is assessed for a range of (i) design flow and stagnation enthalpy rise coefficients, (ii) rotor spanwise work profiles, (iii) rotor–stator spacings, and (iv) nonaxisymmetric stator geometries. Of the approaches examined, nonaxisymmetric stator geometry and increased stage flow and stagnation enthalpy rise coefficients provide the greatest reductions in rotor flow nonuniformity, and may offer the most potential for mitigating performance loss due to BLI inlet distortion.

scholarly journals Analysis of Fan Stage Design Attributes for Boundary Layer Ingestion

2016 ◽  
Author(s):  
D. K. Hall ◽  
E. M. Greitzer ◽  
C. S. Tan
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Surface Geometry ◽  
Flow Distortion ◽  
Local Flow ◽  
Stage Design ◽  
Flow Angle ◽  
Inlet Distortion ◽  
Design Attributes

This paper describes a new conceptual framework for three-dimensional turbomachinery flow analysis and its use to assess fan stage attributes for mitigating adverse effects of inlet distortion due to boundary layer ingestion (BLI). A non-axisymmetric throughflow method has been developed to describe the fan flow field with inlet distortion. In this the turbomachinery is modeled using momentum and energy source distributions that are determined as a function of local flow conditions and a specified blade camber surface geometry. Comparison with higher-fidelity computational and experimental results shows that the method captures the principal flow redistribution and distortion transfer effects associated with BLI. Distortion response is assessed for a range of (i) rotor spanwise work profiles, (ii) rotor-stator spacings, and (iii) non-axisymmetric stator geometries. For the parameters examined, changes in axisymmetric design result in trades between rotor and stator distortions, or between different radial sections of a given blade row with marginal overall gain. Of the approaches examined, non-axisymmetric stator exit flow angle distributions were found to provide the greatest reduction in rotor flow distortion and thus may offer the most potential for mitigating decreases in performance due to BLI inlet distortion.

Generation of steady longitudinal vortices in hypersonic boundary layer

2013 ◽  
Vol 729 ◽  
pp. 702-731 ◽  
Author(s):  
A. I. Ruban ◽  
M. A. Kravtsova
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Viscosity Coefficient ◽  
Momentum Equation ◽  
The Body ◽  
Nonlinear Behaviour ◽  
Hypersonic Boundary Layer ◽  
Stagnation Temperature ◽  
Characteristic Distance

AbstractIn this paper we study the three-dimensional perturbations produced in a hypersonic boundary layer by a small wall roughness. The flow analysis is performed under the assumption that the Reynolds number, $R{e}_{0} = {\rho }_{\infty } {V}_{\infty } L/ {\mu }_{0} $, and Mach number, ${M}_{\infty } = {V}_{\infty } / {a}_{\infty } $, are large, but the hypersonic interaction parameter, $\chi = { M}_{\infty }^{2} R{ e}_{0}^{- 1/ 2} $, is small. Here ${V}_{\infty } $, ${\rho }_{\infty } $ and ${a}_{\infty } $ are the flow velocity, gas density and speed of sound in the free stream, ${\mu }_{0} $ is the dynamic viscosity coefficient at the ‘stagnation temperature’, and $L$ is the characteristic distance the boundary layer develops along the body surface before encountering a roughness. We choose the longitudinal and spanwise dimensions of the roughness to be $O({\chi }^{3/ 4} )$ quantities. In this case the flow field around the roughness may be described in the framework of the hypersonic viscous–inviscid interaction theory, also known as the triple-deck model. Our main interest in this paper is the nonlinear behaviour of the perturbations. We study these by means of numerical solution of the triple-deck equations, for which purpose a modification of the ‘skewed shear’ technique suggested by Smith (United Technologies Research Center Tech. Rep. 83-46, 1983) has been used. The technique requires global iterations to adjust the viscous and inviscid parts of the flow. Convergence of such iterations is known to be a major problem in viscous–inviscid calculations. In order to achieve improved stability of the method, both the momentum equation for the viscous part of the flow, and the equations describing the interaction with the flow outside the boundary layer, are treated implicitly in this study. The calculations confirm the fact that in this sort of flow the perturbations are capable of propagating upstream in the boundary layer, resulting in a perturbation field which surrounds the roughness on all sides. We found that the perturbations decay rather fast with the distance from the roughness everywhere except in the wake behind the roughness. We found that if the height of the roughness is small, then the perturbations also decay in the wake, though much more slowly than outside the wake. However, if the roughness height exceeds some critical value, then two symmetric counter-rotating vortices form in the wake. They appear to support themselves and grow as the distance from the roughness increases.

Effects of Inlet Distortion on the Flow Field in a Transonic Compressor Rotor

1996 ◽  
Author(s):  
Chunill Hah ◽  
Douglas C. Rabe ◽  
Thomas J. Sullivan ◽  
Aspi R. Wadia
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Viscous Flow ◽  
Total Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Aerodynamic Loss ◽  
Transonic Compressor ◽  
Inlet Distortion ◽  
Compressor Rotor

The effects of circumferential distortions in inlet total pressure on the flow field in a low-aspect-ratio, high-speed, high-pressure-ratio, transonic compressor rotor are investigated in this paper. The flow field was studied experimentally and numerically with and without inlet total pressure distortion. Total pressure distortion was created by screens mounted upstream from the rotor inlet. Circumferential distortions of 8 periods per revolution were investigated at two different rotor speeds. The unsteady blade surface pressures were measured with miniature pressure transducers mounted in the blade. The flow fields with and without inlet total pressure distortion were analyzed numerically by solving steady and unsteady forms of the Reynolds-averaged Navier-Stokes equations. Steady three-dimensional viscous flow calculations were performed for the flow without inlet distortion while unsteady three-dimensional viscous flow calculations were used for the flow with inlet distortion. For the time-accurate calculation, circumferential and radial variations of the inlet total pressure were used as a time-dependent inflow boundary condition. A second-order implicit scheme was used for the time integration. The experimental measurements and the numerical analysis are highly complementary for this study because of the extreme complexity of the flow field. The current investigation shows that inlet flow distortions travel through the rotor blade passage and are convected into the following stator. At a high rotor speed where the flow is transonic, the passage shock was found to oscillate by as much as 20% of the blade chord, and very strong interactions between the unsteady passage shock and the blade boundary layer were observed. This interaction increases the effective blockage of the passage, resulting in an increased aerodynamic loss and a reduced stall margin. The strong interaction between the passage shock and the blade boundary layer increases the peak aerodynamic loss by about one percent.

Three-Dimensional Calculation of Wall Boundary Layer Flows in Turbomachines

1987 ◽  
Author(s):  
W. L. Lindsay ◽  
H. B. Carrick ◽  
J. H. Horlock
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Cross Flow ◽  
Flow Profile ◽  
Boundary Layer Flows ◽  
Wall Boundary Layer ◽  
Boundary Layer Development ◽  
Flow Through ◽  
The Cross

An integral method of calculating the three-dimensional turbulent boundary layer development through the blade rows of turbomachines is described. It is based on the solution of simultaneous equations for (i) & (ii) the growth of streamwise and cross-flow momentum thicknesses; (iii) entrainment; (iv) the wall shear stress; (v) the position of maximum cross-flow. The velocity profile of the streamwise boundary layer is assumed to be that described by Coles. The cross-flow profile is assumed to be the simple form suggested by Johnston, but modified by the effect of bounding blade surfaces, which restrict the cross-flow. The momentum equations include expressions for “force-defect” terms which are also based on secondary flow analysis. Calculations of the flow through a set of guide vanes of low deflection show good agreement with experimental results; however, attempts to calculate flows of higher deflection are found to be less successful.

scholarly journals Inner Workings of Aerodynamic Sweep

1997 ◽  
Author(s):  
A. R. Wadia ◽  
P. N. Szucs ◽  
D. W. Crall
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Axial Flow ◽  
Tip Clearance ◽  
Surface Boundary ◽  
Surface Boundary Layer ◽  
Stall Margin ◽  
Design Speed ◽  
The Impact

The recent trend in using aerodynamic sweep to improve the performance of transonic blading has been one of the more significant technological evolutions for compression components in turbomachinery. This paper reports on the experimental and analytical assessment of the pay-off derived from both aft and forward sweep technology with respect to aerodynamic performance and stability. The single stage experimental investigation includes two aft-swept rotors with varying degree and type of aerodynamic sweep and one swept forward rotor. On a back-to-back test basis, the results are compared with an unswept rotor with excellent performance and adequate stall margin. Although designed to satisfy identical design speed requirements as the unswept rotor, the experimental results reveal significant variations in efficiency and stall margin with the swept rotors. At design speed, all the swept rotors demonstrated a peak stage efficiency level that was equal to that of the unswept rotor. However, the forward-swept rotor achieved the highest rotor-alone peak efficiency. At the same time, the forward-swept rotor demonstrated a significant improvement in stall margin relative to the already satisfactory level achieved by the unswept rotor. Increasing the level of aft sweep adversely affected the stall margin. A three-dimensional viscous flow analysis was used to assist in the interpretation of the data. The reduced shock/boundary layer interaction, resulting from reduced axial flow diffusion and less accumulation of centrifuged blade surface boundary layer at the up, was identified as the prime contributor to the enhanced performance with forward sweep. The impact of tip clearance on the performance and stability for one of the aft-swept rotors was also assessed.

Aerodynamic Comparison Between Circumferential and Wing-Embedded Inlet Distortion for an Ultra-High Bypass Ratio Fan Stage

2019 ◽  
Author(s):  
Daniel Giesecke ◽  
Jens Friedrichs
Keyword(s):  
Boundary Layer ◽  
Pressure Ratio ◽  
Aircraft Design ◽  
Low Noise ◽  
Research Centre ◽  
Aircraft Wing ◽  
Local Flow ◽  
Inlet Distortion ◽  
Distortion Coefficient ◽  
Bypass Ratio

Abstract Future aircraft design concepts often show a somewhat wing embedded ultra-high bypass ratio engine. The aircraft concept of the Coordinated Research Centre 880 (CRC880) is a single-aisle configuration with engines partly integrated over the aircraft wing. The aircraft is designed to take off and land on regional airfields with low noise and fuel emissions to address the guidelines set by the ACARE. As a result of the engine installation, the inlet induces a non-axisymmetric boundary layer ingestion into the fan stage. In experimental setups, inlet distortion has often been seen as a 60-degree circumferential inlet stagnation pressure distortion. However, the fan stage inlet flow of the prescribed engine installation of the CRC880 differs to a great extent from a 60-degree sector. In this paper, an aerodynamic comparison between a realistic inflow situation and a similar 60-degree inlet distortion for the same ultra-high bypass ratio fan stage is given. The realistic inflow situation is a result of the flow moving over the aircraft wing suction side and entering the nacelle. As non-axisymmetric inlet geometry remains the same for both cases, therefore, only the total pressure boundary condition at nacelle inlet was changed between both cases. Hence, full annulus simulations are required. Both inlet distortion cases are equivalent by matching average 60-degree distortion coefficient. This study points out that the method, by using averaged 60-degree segment values, excludes specific inflow characteristics. For the same averaged 60-degree distortion coefficient, the local distortion of the embedded case is up to four times larger at rotor tip compared to the segmental approach. For constant mass flow, fan pressure ratio and isentropic efficiency drop by more than five and eight percent respectively. At peak efficiency operating condition, the decrease is even more significant with more than nine percent in stage efficiency. For future embedded aircraft engine configurations, the fan designer has to take into account the non-axisymmetric local flow changes. Hence, the 60-degree segment method is not sufficient when investigating experimental boundary layer ingesting fans and therefore, further method developments are necessary.

The Use of Fluid-Solid Cell Transformation to Model Volcanic Ash Deposition Within a Gas Turbine Hot Component

2018 ◽  
Author(s):  
Jonathon Connolly ◽  
Peter Forsyth ◽  
Matthew McGilvray ◽  
David Gillespie
Keyword(s):  
Boundary Layer ◽  
Volcanic Ash ◽  
Computational Models ◽  
Cell Transformation ◽  
Fluid Dynamic ◽  
Surface Geometry ◽  
Validation Data ◽  
Local Flow ◽  
Mesh Morphing ◽  
Order Of Magnitude

Creating robust empirical and computational models of the process of deposition of salts, dust, sand and volcanic ash has gained increased importance over the last two decades as civil aircraft flights in regions with particulate laden atmospheres have increased. This is associated with increased costs of maintenance to engine suppliers in a market where there is pressure from carriers to continue to fly. Thus, knowledge of the build-up of particulates within the engine over long or multiple deposition events is required in addition to predicting its onset. In this paper deposition in idealised geometries typical of internal cooling passages is examined. The fluid phase is modelled using the commercial flow solver FLUENT and a simple RANS approach. The discrete phase was then solved using Lagrangian particle tracking and a continuous random walk model using one-way coupling. Following identification of deposition fluxes, the local surface of the solid domain was modified using a bespoke cell transformation process. Particular care was taken to distribute deposited mass appropriately to surface cells to avoid large discontinuities at the boundaries. The model was implemented using user-defined functions. The functionality of the technique, is demonstrated through application two impingement cooling geometries for which experimental validation data are available. Here the solution was highly sensitive to the changing target surface geometry as deposition advanced temporally. Fair agreement was found with the experimental data of Burwash et al.[1] though the level of accretion found was an order of magnitude too high, highlighting the need to combine this approach with accurate stick-bounce and shedding models. Significant changes in deposition locations were observed as the deposition site grew in size. Comparison to a second validation case, by Clum et al [2], was used to test further the effect of deposition on the local flow field. Again, good qualitative agreement was obtained. The procedure is shown to create believable deposits of volcanic ash for all cases tested, without many of the typical problems encountered with mesh morphing — overlapping volumes and indeterminate boundary layer resolution. For the commercial computational fluid dynamic (CFD) code used, the process of identifying cells which are to be modified and their neighbours is, disappointingly, an order of magnitude slower than using a mesh morphing strategy. The procedure does, however maintain high, known, resolution throughout the thermal boundary layer, will allow the redistribution of particles to take into account features such as the fusing of neighbouring accretions and the breakaway of deposits from the surface as they grow.

Aerodynamic Performance of Large Centrifugal Compressors

1982 ◽  
Vol 104 (4) ◽  
pp. 796-804 ◽  
Author(s):  
Fumikata Kano ◽  
Noriyuki Tazawa ◽  
Yoshiteru Fukao
Keyword(s):  
Boundary Layer ◽  
Flow Analysis ◽  
Three Dimensional ◽  
Design Procedure ◽  
Suction Side ◽  
Aerodynamic Performance ◽  
Boundary Layer Theory ◽  
Air Separation ◽  
Three Dimensional Flow ◽  
Dimensional Measurements

The aerodynamic performance of impellers and diffusers of the large centrifugal compressor were studied. A performance design procedure based on the quasi-three-dimensional flow analysis which is combined with the boundary layer theory was developed. The conditions of the boundary layer at the impeller exit and at the diffuser vane throat were calculated, and the three-dimensional measurements were carried out. This result shows that the low momentum flow is accumulated at the corner of the shroud and the blade suction side of the impeller. These results were applied to the development of a large four-stage isothermal compressor which handles the air for an air separation apparatus. This was tested in the field and showed an isothermal efficiency of 76 percent.

scholarly journals How smooth is a dolphin? The ridged skin of odontocetes

2019 ◽  
Vol 15 (7) ◽  
pp. 20190103 ◽  
Author(s):  
Dylan K. Wainwright ◽  
Frank E. Fish ◽  
Sam Ingersoll ◽  
Terrie M. Williams ◽  
Judy St Leger ◽  
...  
Keyword(s):  
Boundary Layer ◽  
Drag Reduction ◽  
Three Dimensional ◽  
Morphological Data ◽  
Boundary Layer Flows ◽  
Surface Geometry ◽  
Average Roughness ◽  
Fluid Resistance ◽  
Layer Flow

Dolphin skin has long been an inspiration for research on drag reduction mechanisms due to the presence of skin ridges that could reduce fluid resistance. We gathered in vivo three-dimensional surface data on the skin from five species of odontocetes to quantitatively examine skin texture, including the presence and size of ridges. We used these data to calculate k + values, which relate surface geometry to changes in boundary layer flow. Our results showed that while ridge size differs among species, odontocete skin was surprisingly smooth compared to the skin of other swimmers (average roughness = 5.3 µm). In addition, the presence of ridges was variable among individuals of the same species. We predict that odontocete skin ridges do not alter boundary layer flows at cruising swimming speeds. By combining k + values and morphological data, our work provides evidence that skin ridges are unlikely to be an adaptation for drag reduction and that odontocete skin is exceptionally smooth compared to other pelagic swimmers.

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