scholarly journals Advanced Compressor Loss Correlations, Part II: Experimental Verifications

1997 ◽  
Vol 3 (3) ◽  
pp. 179-187 ◽  
Author(s):  
M. T. Schobeiri
Keyword(s):  
Total Pressure ◽  
Computational Fluid Mechanics ◽  
Momentum Thickness ◽  
Pressure Losses ◽  
Transonic Compressor ◽  
Tremendous Progress ◽  
Loss Coefficients ◽  
Profile Loss ◽  
Stage Efficiency ◽  
Shock Loss

Reliable efficiency calculation of high-subsonic and transonic compressor stages requires a detailed and accurate prediction of the flow field within these stages. Despite the tremendous progress in turbomachinery computational fluid mechanics, the compressor designer still uses different loss correlations to estimate the total pressure losses and thus the efficiency of the compressor stage. The new shock loss model and the modified diffusion factor, developed in Part I, were implemented into a loss calculation procedure. In this part, correlations for total pressure loss, profile loss, and secondary loss coefficients are presented, using the available experimental data. Based on the profile loss coefficients, correlations were also established for boundary layer momentum thickness. These correlations allow the compressor designer to accurately estimate the blade losses and therefore the stage efficiency.

scholarly journals Advanced Compressor Loss Correlations, Part I: Theoretical Aspects

1997 ◽  
Vol 3 (3) ◽  
pp. 163-177 ◽  
Author(s):  
M. T. Schobeiri
Keyword(s):  
Total Pressure ◽  
Computational Fluid Mechanics ◽  
Loss Model ◽  
Pressure Losses ◽  
Transonic Compressor ◽  
Tremendous Progress ◽  
Profile Losses ◽  
Shock Position ◽  
Experimental Findings ◽  
Shock Loss

Reliable efficiency calculation of high-subsonic and transonic compressor stages requires a detailed and accurate prediction of the flow field within these stages. Despite the tremendous progress in turbomachinery computational fluid mechanics, the compressor designer still uses different loss correlations to estimate the total pressure losses and thus the efficiency of the compressor stage. A new loss model is presented in this article. Special attention is paid to the shock and profile losses, since they contribute significantly to the total pressure loss balance, specifically for transonic compressor stages. A new shock loss model is presented that calculates the shock position and the shock total pressure losses. The available experimental data were used to establish new loss correlations that account for experimental findings.

scholarly journals Aerodynamic Losses in Turbines with and without Film Cooling, as Influenced by Mainstream Turbulence, Surface Roughness, Airfoil Shape, and Mach Number

2012 ◽  
Vol 2012 ◽  
pp. 1-28 ◽  
Author(s):  
Phil Ligrani
Keyword(s):  
Surface Roughness ◽  
Mach Number ◽  
Film Cooling ◽  
Total Pressure ◽  
High Speed ◽  
Density Ratio ◽  
Aerodynamic Loss ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Exit Mach Number

The influences of a variety of different physical phenomena are described as they affect the aerodynamic performance of turbine airfoils in compressible, high-speed flows with either subsonic or transonic Mach number distributions. The presented experimental and numerically predicted results are from a series of investigations which have taken place over the past 32 years. Considered are (i) symmetric airfoils with no film cooling, (ii) symmetric airfoils with film cooling, (iii) cambered vanes with no film cooling, and (iv) cambered vanes with film cooling. When no film cooling is employed on the symmetric airfoils and cambered vanes, experimentally measured and numerically predicted variations of freestream turbulence intensity, surface roughness, exit Mach number, and airfoil camber are considered as they influence local and integrated total pressure losses, deficits of local kinetic energy, Mach number deficits, area-averaged loss coefficients, mass-averaged total pressure loss coefficients, omega loss coefficients, second law loss parameters, and distributions of integrated aerodynamic loss. Similar quantities are measured, and similar parameters are considered when film-cooling is employed on airfoil suction surfaces, along with film cooling density ratio, blowing ratio, Mach number ratio, hole orientation, hole shape, and number of rows of holes.

Modelling Pressure Losses at Arterial Junctions With Application to Junctions of the Fetal Ductus Arteriosus

2012 ◽  
Author(s):  
Jonathan P. Mynard ◽  
Malcolm R. Davidson ◽  
Daniel J. Penny ◽  
Joseph J. Smolich
Keyword(s):  
Total Pressure ◽  
Mechanical Energy ◽  
Ductus Arteriosus ◽  
Acute Angle ◽  
Complex Flow ◽  
Descending Aorta ◽  
Pressure Losses ◽  
High Kinetic Energy ◽  
Loss Coefficients ◽  
Dynamics Simulations

In one-dimensional (1D) models of arterial networks, branch junctions are represented by flow and mechanical energy (or total pressure, i.e. p + 1/2ρu2) coupling conditions. The flow condition simply ensures conservation of mass, but the pressure condition is less trivial because pressure losses are known to occur in the vicinity of junctions, caused by regions of complex flow that depend on the vascular geometry and prevailing flow patterns. These losses are commonly ignored in 1D models under the assumption that area ratios and branching angles of arterial junctions are optimally designed. However, one setting where pressure losses are likely to be important is the junction of the ductus arteriosus (DA) with the aorta in the fetus, considering the high kinetic energy of blood in the DA [1], the acute angle between the aortic isthmus (Aols) and DA, and the redirection of DA blood flow towards the descending aorta (DAo, Figure 1). Previously, pressure losses have been approximated in 1D models by enforcing continuity of static (rather than total) pressure [2] or by using empirical loss coefficients obtained from experiments in 90 degree T-tubes [3]. In the current study, we implemented a loss formulation described by Bassett et al [4] for 1D gas dynamics simulations, which unlike previous methods, can be used to model junctions with any number of branches and any orientation of branch angles, and explicitly accounts for the influence of area and flow ratios on pressure losses. Results of the model are validated against high fidelity measurements in fetal lambs.

scholarly journals Effect of Relative Velocity Distribution on Efficiency and Exit Flow of Centrifugal Impellers

1983 ◽  
Author(s):  
H. Mishina ◽  
H. Nishida
Keyword(s):  
Total Pressure ◽  
Flow Behavior ◽  
Design Parameters ◽  
One Dimensional ◽  
Pressure Losses ◽  
Loss Analysis ◽  
High Stage ◽  
The Relationship ◽  
Stage Efficiency ◽  
Wide Operating

The major problem for designing centrifugal compressors is to attain high stage efficiency as well as a wide operating range. High stage efficiency is customarily attained by the optimization of design parameters using a one-dimensional loss analysis including the relationship between the flow behavior and total pressure losses for limited types of compressors.

Growth of Secondary Flow Losses Downstream of a Turbine Blade Cascade

1986 ◽  
Vol 108 (2) ◽  
pp. 270-276 ◽  
Author(s):  
L. D. Chen ◽  
S. L. Dixon
Keyword(s):  
Boundary Layer ◽  
Aspect Ratio ◽  
Secondary Flow ◽  
Total Pressure ◽  
Turbine Cascade ◽  
Pressure Losses ◽  
Flow Losses ◽  
Blade Cascade ◽  
Flow Loss ◽  
Loss Coefficients

Endwall total pressure losses downstream of a low-speed turbine cascade have been measured at several planes in order to determine the changes in secondary flow loss coefficients and the growth of the mixing loss with distance downstream. The results obtained are compared with various published secondary flow loss correlations in an attempt to explain some of the anomalies which presently exist. The paper includes some new correlations including one for the important gross secondary loss coefficient YSG with loading and aspect ratio parameters as well as the upstream boundary layer parameters.

Off-Design Performance of a Linear Cascade of Turbine Blades

1990 ◽  
Author(s):  
B. Tremblay ◽  
S. A. Sjolander ◽  
S. H. Moustapha
Keyword(s):  
Reynolds Number ◽  
Total Pressure ◽  
Critical Reynolds Number ◽  
Turbine Blades ◽  
Surface Flow ◽  
Turning Angle ◽  
Linear Cascade ◽  
Pressure Losses ◽  
Design Performance ◽  
Loss Coefficients

A recent survey of the literature showed a clear need for additional experimental results on the off-design performance of turbines, particularly for airfoils of recent design. This study presents measurements of the low-speed two-dimensional performance of a linear cascade of turbine blades with a turning angle of 87 degrees. The incidence was varied between −25 and +25 degrees in 5 degree steps. The blade surface pressures, total pressure loss coefficients and trailing-edge deviations are presented for all values of incidence. The influence of incidence on the critical Reynolds number is also examined. Surface flow visualization is presented for different values of Reynolds number and incidence to aid in the physical interpretation of the measurements. The measured total pressure losses agree very well with the new off-design correlation introduced by Moustapha et al. (1989).

Growth of Secondary Flow Losses Downstream of a Turbine Blade Cascade

1985 ◽  
Author(s):  
L. D. Chen ◽  
S. L. Dixon
Keyword(s):  
Boundary Layer ◽  
Secondary Flow ◽  
Total Pressure ◽  
Turbine Cascade ◽  
Pressure Losses ◽  
Flow Losses ◽  
Blade Cascade ◽  
Flow Loss ◽  
Loss Coefficients ◽  
End Wall

End wall total pressure losses downstream of a low-speed turbine cascade have been measured at several planes in order to determine the changes in secondary flow loss coefficients and the growth of the mixing loss with distance downstream. The results obtained are compared with various published secondary flow loss correlations in an attempt to explain some of the anomalies which presently exist. The paper includes some new correlations including one for the important gross secondary loss coefficient YSG with loading and aspect ratio parameters as well as the upstream boundary layer parameters.

Loss Reduction in Compressor Cascades by Means of Passive Flow Control

2008 ◽  
Author(s):  
A. Hergt ◽  
R. Meyer ◽  
M. W. Mu¨ller ◽  
K. Engel
Keyword(s):  
Flow Control ◽  
Total Pressure ◽  
Axial Compressor ◽  
Pressure Rise ◽  
Side Wall ◽  
Vortex Generators ◽  
Test Facility ◽  
Compressor Cascade ◽  
Pressure Losses ◽  
Transonic Compressor

Secondary flow effects like the corner stall between the wall and the vane in a compressor stage are responsible for a large part of total pressure losses. An extensive experimental study of flow control in a highly loaded compressor cascade was performed in order to decrease the separation and reduce the losses by means of vortex generators. The vortex generators were attached at the surface of the cascade side walls. These flow control devices produce strong vortices, which enhance the mixing between the main flow and the decelerated boundary layer at the side wall. Thus, the corner flow separation and the total pressure losses could be reduced. The experiments were carried out with a compressor cascade at a high-speed test facility at the DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. The cascade consisted of five vanes. The blade profiles are comparable to the hub section of the stator vanes used in the transonic compressor test rig running at Technische Universita¨t Darmstadt. In the range between −2° and +4° angle of incidence the total pressure losses of the cascade could be reduced up to 4.6% by means of vortex generators, whereas the static pressure rise was not influenced. Based on the results of the cascade measurements, the vortex generators were applied in front of the stator row of the single stage axial compressor at Technische Universita¨t Darmstadt. A numerical simulation of the compressor flow provided an indication for the adjustment of the vortex generators at the hub and casing. In the experiments the pressure rise and the efficiency of the axial compressor was measured and it could be shown that vortex generators partially improve the efficiency.

scholarly journals Turbulence model performance for ventilation components pressure losses

2021 ◽  
Author(s):  
Karsten Tawackolian ◽  
Martin Kriegel
Keyword(s):  
Reynolds Number ◽  
Turbulence Model ◽  
Pressure Loss ◽  
Low Reynolds Number ◽  
Turbulence Models ◽  
Reynolds Numbers ◽  
Test Case ◽  
Pressure Losses ◽  
Loss Coefficients ◽  
Near Wall

AbstractThis study looks to find a suitable turbulence model for calculating pressure losses of ventilation components. In building ventilation, the most relevant Reynolds number range is between 3×104 and 6×105, depending on the duct dimensions and airflow rates. Pressure loss coefficients can increase considerably for some components at Reynolds numbers below 2×105. An initial survey of popular turbulence models was conducted for a selected test case of a bend with such a strong Reynolds number dependence. Most of the turbulence models failed in reproducing this dependence and predicted curve progressions that were too flat and only applicable for higher Reynolds numbers. Viscous effects near walls played an important role in the present simulations. In turbulence modelling, near-wall damping functions are used to account for this influence. A model that implements near-wall modelling is the lag elliptic blending k-ε model. This model gave reasonable predictions for pressure loss coefficients at lower Reynolds numbers. Another example is the low Reynolds number k-ε turbulence model of Wilcox (LRN). The modification uses damping functions and was initially developed for simulating profiles such as aircraft wings. It has not been widely used for internal flows such as air duct flows. Based on selected reference cases, the three closure coefficients of the LRN model were adapted in this work to simulate ventilation components. Improved predictions were obtained with new coefficients (LRNM model). This underlined that low Reynolds number effects are relevant in ventilation ductworks and give first insights for suitable turbulence models for this application. Both the lag elliptic blending model and the modified LRNM model predicted the pressure losses relatively well for the test case where the other tested models failed.

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.

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