scholarly journals Nonaxisymmetric Stator Design for Boundary Layer Ingesting Fans

2019 ◽  
Vol 141 (7) ◽  
Author(s):  
Ewan J. Gunn ◽  
Cesare A. Hall
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Surface ◽  
Inlet Distortion ◽  
Design Changes ◽  
Final Design ◽  
Computational Fluid Dynamics Cfd ◽  
And Diffusion ◽  
Stator Design ◽  
Additional Loss

In a boundary layer ingesting (BLI) fan system, the inlet flow field is highly nonuniform. In this environment, an axisymmetric stator design suffers from a nonuniform distribution of hub separations, increased wake thicknesses, and casing losses. These additional loss sources can be reduced using a nonaxisymmetric design that is tuned to the radial and circumferential flow variations at exit from the rotor. In this paper, a nonaxisymmetric design approach is described for the stator of a low-speed BLI fan. First, sectional design changes are applied at each radial and circumferential location. Next, this approach is combined with the application of nonaxisymmetric lean. The designs were tested computationally using full-annulus unsteady computational fluid dynamics (CFD) of the complete fan stage with a representative inlet distortion. The final design has also been manufactured and tested experimentally. The results show that a 2D sectional approach can be applied nonaxisymmetrically to reduce incidence and diffusion factor at each location. This leads to reduced loss, particularly at the casing and midspan, but it does not eliminate the hub separations that are present within highly distorted regions of the annulus. These are relieved by nonaxisymmetric lean where the pressure surface is inclined toward the hub. For the final design, the loss in the stator blades operating with BLI was measured to be 10% lower than that for the original stator design operating with undistorted inflow. Overall, the results demonstrate that the nonaxisymmetric design has the potential to eliminate any additional loss in a BLI fan stator caused by the nonuniform ingested flow field.

scholarly journals Non-Axisymmetric Stator Design for Boundary Layer Ingesting Fans

2017 ◽  
Author(s):  
E. J. Gunn ◽  
C. A. Hall
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Diffusion Factor ◽  
Pressure Surface ◽  
Inlet Distortion ◽  
Design Changes ◽  
Final Design ◽  
And Diffusion ◽  
Stator Design ◽  
Additional Loss

In a Boundary Layer Ingesting (BLI) fan system the inlet flow field is highly non-uniform. In this environment, an axisymmetric stator design suffers from a non-uniform distribution of hub separations, increased wake thicknesses and casing losses. These additional loss sources can be reduced using a non-axisymmetric design that is tuned to the radial and circumferential flow variations at exit from the rotor. In this paper a non-axisymmetric design approach is described for the stator of a low-speed BLI fan. Firstly sectional design changes are applied at each radial and circumferential location. Next, this approach is combined with the application of non-axisymmetric lean. The designs were tested computationally using full-annulus unsteady CFD of the complete fan stage with a representative inlet distortion. The final design has also been manufactured and tested experimentally. The results show that a 2D sectional approach can be applied non-axisymmetrically to reduce incidence and diffusion factor at each location. This leads to reduced loss, particularly at the casing and midspan, but it does not eliminate the hub separations that are present within highly distorted regions of the annulus. These are relieved by non-axisymmetric lean where the pressure surface is inclined towards the hub. For the final design, the loss in the stator blades operating with BLI was measured to be 10% lower than for the original stator design operating with undistorted inflow. Overall, the results demonstrate that non-axisymmetric design has the potential to eliminate any additional loss in a BLI fan stator caused by the non-uniform ingested flow-field.

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.

scholarly journals The Effect of the Inlet Boundary Layer on the Tip Flow Field in a Transonic Fan Rotor

1998 ◽  
Author(s):  
Junji Takado ◽  
Toyotaka Sonoda ◽  
Satoshi Nakamura
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Ratio ◽  
Three Dimensional ◽  
Interaction Region ◽  
Operating Conditions ◽  
Laser Doppler Velocimeter ◽  
Intensity Level ◽  
Pressure Surface ◽  
Transonic Fan

Experimental and numerical investigations have been carried out to understand the effects of the inlet boundary layer (IBL) on the tip flow field including the aerodynamic performance in a transonic fan rotor. Both the steady and the unsteady phenomena in the tip flow field have been investigated for operating conditions near peak efficiency and near stall with the two types of tip IBL. In order 10 study these phenomena, high response pressure data with Kulite transducers and laser doppler velocimeter (LDV) data have been acquired around the tip region. Furthermore, three-dimensional Navier-Stokes numerical simulations have been compared with the measured results. The results indicate that the tip IBL significantly influences the spanwise distribution of pressure ratio around the tip region and the stall characteristics including the passage shock / tip leakage vortex interaction, the blockage generation, the wake structure, and the unsteadiness of the tip flow field. In particular, at a near stall condition for the thick IBL with high turbulence intensity level, the tip diffusion level is increased due to a larger blockage, which is generated downstream of a much stronger interaction region. These phenomena are a consequence of the low momentum fluid in the tip IBL, and significantly reduce the stall margin. Furthermore, the unsteadiness drastically increases around the interaction region and around the pressure surface where the blockage migrates. These unsteady phenomena are distinctive features near stall. Downstream of the rotor, the larger and more unsteady blockage is discharged from the pressure surface side, and complicates the three-dimensional rotor exit flow field around the tip region.

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

1998 ◽  
Vol 120 (2) ◽  
pp. 233-246 ◽  
Author(s):  
C. Hah ◽  
D. C. Rabe ◽  
T. J. Sullivan ◽  
A. 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 eight 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 percent 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.

The Measurement of Boundary Layers on a Compressor Blade in Cascade: Part 1—A Unique Experimental Facility

1987 ◽  
Vol 109 (4) ◽  
pp. 520-526 ◽  
Author(s):  
S. Deutsch ◽  
W. C. Zierke
Keyword(s):  
Boundary Layer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Static Pressure ◽  
Compressor Blade ◽  
Boundary Layer Transition ◽  
Two Dimensional ◽  
Suction Surface ◽  
Pressure Surface ◽  
Static Pressure Distribution

A unique cascade facility is described which permits the use of laser-Doppler velocimetry (LDV) to measure blade boundary layer profiles. Because of the need for a laser access window, the facility cannot reply on continuous blade pack suction to achieve two-dimensional, periodic flow. Instead, a strong suction upstream of the blade pack is used in combination with tailboards to control the flow field. The distribution of the upstream suction is controlled through a complex baffling system. A periodic, two–dimensional flow field is achieved at a chord Reynolds number of 500,000 and an incidence angle of 5 deg on a highly loaded, double circular arc, compressor blade. Inlet and outlet flow profiles, taken using five-hole probes, and the blade static-pressure distribution are used to document the flow field for use with the LDV measurements (see Parts 2 and 3). Inlet turbulence intensity is measured, using a hot wire, to be 0.18 percent. The static-pressure distribution suggests both separated flow near the trailing edge of the suction surface and an initially laminar boundary layer profile near the leading edge of the pressure surface. Probe measurements are supplemented by sublimation surface visualization studies. The sublimation studies place boundary layer transition at 64.2 ± 3.9 percent chord on the pressure surface, and indicate separation on the suction surface at 65.6 percent ± 3.5 percent chord.

The Design of Highly Loaded Axial Compressors

2010 ◽  
Vol 133 (3) ◽  
Author(s):  
Tony Dickens ◽  
Ivor Day
Keyword(s):  
Pressure Rise ◽  
High Loading ◽  
Axial Compressors ◽  
Pressure Surface ◽  
High Stage ◽  
Computational Fluid Dynamics Cfd ◽  
Cycle Efficiency ◽  
Stage Efficiency ◽  
Highly Loaded ◽  
Additional Loss

Increasing compressor pressure ratios (thereby gaining a benefit in cycle efficiency), or reducing the number of stages (to reduce weight, cost, etc.), will require an increase in pressure rise per stage. One method of increasing the pressure rise per stage is by increasing the stage loading coefficient, and it is this topic, which forms the focus of the present paper. In the past, a great deal of effort has been expended in trying to design highly loaded blade rows. Most of this work has focused on optimizing a particular design, rather than looking at the fundamental problems associated with high loading. This paper looks at the flow physics behind the problem, makes proposals for a new design strategy, and explains sources of additional loss specific to highly loaded designs. Detailed experimental measurements of three highly loaded stages (Δh0/U2≈0.65) have been used to validate a computational fluid dynamics (CFD) code. The calibrated CFD has then been used to show that, as the stage loading is increased, the flow in the stator passages breaks down first. This happens via a large corner separation, which significantly impairs the stage efficiency. The stator can be relieved by increasing stage reaction, thus shifting the burden to the rotor. Fortunately, the CFD calculations show that the rotor is generally more tolerant of high loading than the stator. Thus, when stage loading is increased, it is necessary to increase the reaction to achieve the optimum efficiency. However, the design exercise using the calibrated CFD also shows that the stage efficiency is inevitably reduced as the stage loading is increased (in agreement with the experimental results). In the second part of the paper, the role that the profile loss plays in the reduction in efficiency at high stage loading is considered. A simple generic velocity distribution is developed from first principles to demonstrate the hitherto neglected importance of the pressure surface losses in highly loaded compressors.

scholarly journals Computer Modeling of Aerosol Emissions Spread in the Atmosphere

2019 ◽  
Vol 97 ◽  
pp. 05023 ◽  
Author(s):  
Daler Sharipov ◽  
Sharofiddin Aynakulov ◽  
Otabek Khafizov
Keyword(s):  
Boundary Layer ◽  
Mathematical Model ◽  
Atmospheric Boundary Layer ◽  
Computer Modeling ◽  
Numerical Experiments ◽  
Climatic Factors ◽  
Numerical Algorithms ◽  
Aerosol Emissions ◽  
The Given ◽  
And Diffusion

The paper deals with the development of mathematical model and numerical algorithms for solving the problem of transfer and diffusion of aerosol emissions in the atmospheric boundary layer. The model takes into account several significant parameters such as terrain relief, characteristics of underlying surface and weather-climatic factors. A series of numerical experiments were conducted based on the given model. The obtained results presented here show how these factors affect aerosol emissions spread in the atmosphere.

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