Impact of Nonuniform Leading Edge Coatings on the Aerodynamic Performance of Compressor Airfoils

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
Michael E. Elmstrom ◽  
Knox T. Millsaps ◽  
Garth V. Hobson ◽  
Jeffrey S. Patterson
Keyword(s):  
Mach Number ◽  
Computational Fluid Dynamic ◽  
Total Pressure ◽  
Fluid Dynamic ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Transition To Turbulence ◽  
Navier Stokes ◽  
Pressure Gradients ◽  
Small Separation

A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and nonuniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge nonuniformity were added. A nonuniform coating is obtained when a liquid coating is applied to a surface with high curvature, such as an airfoil leading edge. The CFD code used, RVCQ3D, is a Reynolds averaged Navier–Stokes solver, with a k-omega turbulence model. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of a coating nonuniformity parameter. Results are presented over a range of negative and positive incidences and inlet Mach numbers from 0.6 to 0.8. A map is provided that shows the allowable degree of coating nonuniformity as a function of incidence and inlet Mach number.

Impact of Non-Uniform Leading Edge Coatings on the Aerodynamic Performance of Compressor Airfoils

2005 ◽  
Author(s):  
Michael E. Elmstrom ◽  
Knox T. Millsaps ◽  
Garth V. Hobson ◽  
Jeffrey S. Patterson
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Fluid Dynamic ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Transition To Turbulence ◽  
Navier Stokes ◽  
Pressure Gradients ◽  
Uniform Coating ◽  
Small Separation

A computational fluid dynamic (CFD) investigation is presented that provides predictions of the aerodynamic impact of uniform and non-uniform coatings applied to the leading edge of a compressor airfoil in a cascade. Using a NACA 65(12)10 airfoil, coating profiles of varying leading edge non-uniformity were added. A non-uniform coating is obtained when a liquid coating is applied to a surface with high curvature, such as an airfoil leading edge. The CFD code used, RVCQ3D, is a Reynolds Averaged, Navier-Stokes (RANS) solver, with a k-omega turbulence model. The code predicted that these changes in leading edge shape can lead to alternating pressure gradients in the first few percent of chord that create small separation bubbles and possibly early transition to turbulence. The change in total pressure loss and trailing edge deviation are presented as a function of a coating non-uniformity parameter. Results are presented over a range of negative and positive incidences and inlet Mach numbers from 0.6 to 0.8. A map is provided that shows the allowable degree of coating non-uniformity as a function of incidence and inlet Mach number.

Modeling of External Pressure Gradient Influence on the Stability of Disturbances in the Boundary Layers of Compressed Gas

2013 ◽  
Vol 8 (4) ◽  
pp. 64-75
Author(s):  
Sergey Gaponov ◽  
Natalya Terekhova
Keyword(s):  
Mach Number ◽  
Pressure Gradient ◽  
Boundary Layers ◽  
Leading Edge ◽  
External Pressure ◽  
Transition To Turbulence ◽  
Compressed Gas ◽  
High Mach ◽  
The Stability ◽  
Very High

This work continues the research on modeling of passive methods of management of flow regimes in the boundary layers of compressed gas. Authors consider the influence of pressure gradient on the evolution of perturbations of different nature. For low Mach number M = 2 increase in pressure contributes to an earlier transition of laminar to turbulent flow, and, on the contrary, drop in the pressure leads to a prolongation of the transition to turbulence. For high Mach number M = 5.35 found that the acoustic disturbances exhibit a very high dependence on the sign and magnitude of the external gradient, with a favorable gradient of the critical Reynolds number becomes smaller than the vortex disturbances, and at worst – boundary layer is destabilized directly on the leading edge

Aerodynamic Analysis of an Airfoil With Leading Edge Pitting Erosion

2017 ◽  
Vol 139 (6) ◽  
Author(s):  
Yan Wang ◽  
Ruifeng Hu ◽  
Xiaojing Zheng
Keyword(s):  
Wind Turbine ◽  
Indirect Effects ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Navier Stokes ◽  
Distribution Area ◽  
Navier Stokes Equation ◽  
Turbine Performance ◽  
Wind Turbine Airfoil ◽  
Cfd Method

Leading edge erosion is a considerable threat to wind turbine performance and blade maintenance, and it is very imperative to accurately predict the influence of various degrees of erosion on wind turbine performance. In the present study, an attempt to investigate the effects of leading edge erosion on the aerodynamics of wind turbine airfoil is undertaken by using computational fluid dynamics (CFD) method. A new pitting erosion model is proposed and semicircle cavities were used to represent the erosion pits in the simulation. Two-dimensional incompressible Reynolds-averaged Navier–Stokes equation and shear stress transport (SST) k–ω turbulence model are adopted to compute the aerodynamics of a S809 airfoil with leading edge pitting erosions, where the influences of pits depth, densities, distribution area, and locations are considered. The results indicate that pitting erosion has remarkably undesirable influences on the aerodynamic performance of the airfoil, and the critical pits depth, density, and distribution area degrade the airfoil aerodynamic performance mostly were obtained. In addition, the dominant parameters are determined by the correlation coefficient path analysis method, results showed that all parameters have non-negligible effects on the aerodynamics of S809 airfoil, and the Reynolds number is of the most important, followed by pits density, pits depth, and pits distribution area. Meanwhile, the direct and indirect effects of these factors are analyzed, and it is found that the indirect effects are very small and the parameters can be considered to be independent with each other.

scholarly journals Numerical Optimization of a Stall Margin Enhancing Recirculation Channel for an Axial Compressor

Fluids ◽  
2019 ◽  
Vol 4 (2) ◽  
pp. 88
Author(s):  
Motoyuki Kawase ◽  
Aldo Rona
Keyword(s):  
High Speed ◽  
High Performance ◽  
Fluid Dynamic ◽  
Axial Compressor ◽  
Leading Edge ◽  
Navier Stokes ◽  
Test Case ◽  
Dynamic Simulations ◽  
Stall Margin ◽  
Casing Treatment

A proof of concept is provided by computational fluid dynamic simulations of a new recirculating type casing treatment. This treatment aims at extending the stable operating range of highly loaded axial compressors, so to improve the safety of sorties of high-speed, high-performance aircraft powered by high specific thrust engines. This casing treatment, featuring an axisymmetric recirculation channel, is evaluated on the NASA rotor 37 test case by steady and unsteady Reynolds Averaged Navier Stokes (RANS) simulations, using the realizable k-ε model. Flow blockage at the recirculation channel outlet was mitigated by chamfering the exit of the recirculation channel inner wall. The channel axial location from the rotor blade tip leading edge was optimized parametrically over the range −4.6% to 47.6% of the rotor tip axial chord c z . Locating the channel at 18.2% c z provided the best stall margin gain of approximately 5.5% compared to the untreated rotor. No rotor adiabatic efficiency was lost by the application of this casing treatment. The investigation into the flow structure with the recirculating channel gave a good insight into how the new casing treatment generates this benefit. The combination of stall margin gain at no rotor adiabatic efficiency loss makes this design attractive for applications to high-speed gas turbine engines.

scholarly journals The onset of compressibility effects for aerofoils in ground effect

2007 ◽  
Vol 111 (1126) ◽  
pp. 797-806 ◽  
Author(s):  
G. Doig ◽  
T. J. Barber ◽  
E. Leonardi ◽  
A. J. Neely
Keyword(s):  
Mach Number ◽  
High Speed ◽  
Subsonic Flow ◽  
Flow Characteristics ◽  
Ground Effect ◽  
Aerodynamic Performance ◽  
Computational Studies ◽  
Pressure Gradients ◽  
Cfd Simulations ◽  
Compressibility Effects

Abstract The influence of flow compressibility on a highly-cambered inverted aerofoil in ground effect is presented, based on two-dimensional computational studies. This type of problem has relevance to open-wheel racing cars, where local regions of high-speed subsonic flow form under favourable pressure gradients, even though the maximum freestream Mach number is typically considerably less than Mach 0·3. An important consideration for CFD users in this field is addressed in this paper: the freestream Mach number at which flow compressibility significantly affects aerodynamic performance. More broadly, for aerodynamicists, the consequences of this are also considered. Comparisons between incompressible and compressible CFD simulations are used to identify important changes to the flow characteristics caused by density changes, highlighting the inappropriateness of incompressible simulations of ground effect flows for freestream Mach numbers as low as 0·15.

Influence of a Novel Three-Dimensional Leading Edge Geometry on the Aerodynamic Performance of Transonic Cascade Vanes

2012 ◽  
Author(s):  
D. Bouchard ◽  
A. Asghar ◽  
J. Hardes ◽  
R. Edwards ◽  
W. D. E. Allan ◽  
...  
Keyword(s):  
Total Pressure ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Surface Flow ◽  
Pressure Losses ◽  
Low Pressure Turbine ◽  
Edge Geometry ◽  
Turbine Vane ◽  
Transonic Cascade

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference vane. An analysis of tubercles found in nature and some engineering applications was used to synthesize new leading edge geometry. Three variations of the reference low pressure turbine vane were obtained by changing the characteristic parameters of the tubercles. Shock structure, surface flow visualization and total pressure measurements were made through experiments in a cascade rig, as well as through computational fluid dynamics. The tests were carried out at design zero incidence and off-design ±10-deg and ±5-deg incidences. The performance of the new 3D leading edge geometries was compared against the reference vane. Some leading edge tubercle configurations were effective at decreasing total pressure losses at positive inlet incidence angles. Numerical results supplemented experimental results.

Influence of a Novel 3D Leading Edge Geometry on the Aerodynamic Performance of Low Pressure Turbine Blade Cascade Vanes

2014 ◽  
Author(s):  
A. Asghar ◽  
W. D. E. Allan ◽  
M. LaViolette ◽  
R. Woodason
Keyword(s):  
Turbine Blade ◽  
Pressure Loss ◽  
Total Pressure ◽  
Separated Flow ◽  
Leading Edge ◽  
Low Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Low Pressure Turbine ◽  
Edge Geometry

This paper addresses the issue of aerodynamic performance of a novel 3D leading edge modification to a reference low pressure turbine blade. An analysis of tubercles found in nature and used in some engineering applications was employed to synthesize new leading edge geometry. A sinusoidal wave-like geometry characterized by wavelength and amplitude was used to modify the leading edge along the span of a 2D profile, rendering a 3D blade shape. The rationale behind using the sinusoidal leading edge was that they induce streamwise vortices at the leading edge which influence the separation behaviour downstream. Surface pressure and total pressure measurements were made in experiments on a cascade rig. These were complemented with computational fluid dynamics studies where flow visualization was also made from numerical results. The tests were carried out at low Reynolds number of 5.5 × 104 on a well-researched profile representative of conventional low pressure turbine profiles. The performance of the new 3D leading edge geometries was compared against the reference blade revealing a downstream shift in separated flow for the LE tubercle blades; however, total pressure loss reduction was not conclusively substantiated for the blade with leading edge tubercles when compared with the performance of the baseline blade. Factors contributing to the total pressure loss are discussed.

A study of fan-distortion interaction within the fan rotor

2020 ◽  
pp. 095765092092004
Author(s):  
Zhihui Li ◽  
Juan Du ◽  
Qianfeng Zhang ◽  
Guofeng Ji ◽  
Hongwu Zhang
Keyword(s):  
High Performance ◽  
Fluid Dynamic ◽  
Three Dimensional ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Stagnation Pressure ◽  
Detached Eddy Simulation ◽  
Axial Fan ◽  
Fan Performance ◽  
Inlet Distortion

Boundary-layer-ingesting fans and compressors in the next-generation turbofan engines require high-performance operations under distorted inflow. The aim of this work is to study the effects of inlet distortions including inlet stagnation pressure and temperature distortion, on the aerodynamic performance of a transonic axial fan. Firstly, the validated full-annulus, unsteady, three-dimensional computational fluid dynamic code in conjunction with detached Eddy simulation approach is used here to simulate the fan flows assembly with individual inlet stagnation pressure/temperature distortion. Then, the propagation process of the inlet distortion waves is analyzed to understand how the aerodynamic performance degradation is triggered. The simulation results show that the fan performance is remarkably degraded when the inlet distortion is introduced. The leading-edge spillage, the trailing edge back flow and the “tornado vortex” occur when parts of fan blades encounter the incoming distorted flows. Finally, the responses of fan to the combined inlet stagnation distortion effects are discussed in this paper. It is found that the combined distortion effects can be predicted based on the sum of the performance responses to the individual constituent distortions. Furthermore, the relative location of the constituent distortions shows a non-ignorable influence on the overall fan performance, especially for the intensified inlet distortion.

Internal Aerodynamic Performance Evaluation of Double Entrance S-Duct Intake at Moderately High Subsonic Mach Number

2021 ◽  
Author(s):  
Satpreet Sidhu ◽  
Asad Asghar ◽  
William D. E. Allan ◽  
R. A. Stowe ◽  
R. Pimentel
Keyword(s):  
Mach Number ◽  
Total Pressure ◽  
Swirling Flow ◽  
Essential Element ◽  
Aerodynamic Performance ◽  
Geometric Constraints ◽  
Pressure Recovery ◽  
Data Set ◽  
Similar Work ◽  
And Performance

Abstract Inlets are an essential element of aircraft propulsion systems. Aircraft with fuselage-embedded engines require intake ducts with bends to direct oncoming air into the engine. Consequently they often experience flow separation, losses, total pressure distortion, and swirling flow near the engine faces, all of which are detrimental to engine stability and performance. In some aircraft, double-entrance ducts are used to meet geometric constraints and maintain the required airflow. The present paper investigated aerodynamic performance of a bifurcated Y-duct with S-bends in both horizontal and vertical planes. Intake performance was evaluated at inlet Ma = 0.63 by measuring the surface static pressure along the four stream-wise rows of pressure taps and total pressure and 3D velocities using 5-hole probe across the exit plane of the intake duct. The data were used to determine the static and total pressure recovery, together with associated radial and circumferential distortion coefficients and swirl intensity. This work provides a rare experimental data-set for a twin-entrance, moderately high-subsonic, double S-duct intake. It compared reasonably with the most similar work published, that of single-entrance ducts at higher Mach number. Pressure recovery was on par while swirl was noted to be reduced when compared with those geometries. Complementary computational fluid dynamics was useful in the qualitative comparisons as well.

An analysis of aerodynamic forces on a delta wing

1996 ◽  
Vol 316 ◽  
pp. 173-196 ◽  
Author(s):  
Chien-Cheng Chang ◽  
Sheng-Yuan Lei
Keyword(s):  
Mach Number ◽  
Stokes Equations ◽  
Delta Wing ◽  
Leading Edge ◽  
Navier Stokes ◽  
Flow Structures ◽  
Aerodynamic Forces ◽  
Navier Stokes Equations ◽  
Lift And Drag ◽  
Secondary Vortices

The present study aims at relating lift and drag to flow structures around a delta wing of elliptic section. Aerodynamic forces are analysed in terms of fluid elements of non-zero vorticity and density gradient. The flow regime considered is Mα = 0.6 ∼ 1.8 and α = 5° ∼ 19°, where Mα denotes the free-stream Mach number and α the angle of attack. Let ρ denote the density, u velocity, and ω vorticity. It is found that there are two major source elements Re(x) and Ve(x) which contribute about 95% or even more to the aerodynamic forces for all the cases under consideration, \[R_e({\bm x})=-\frac{1}{2} {\bm u}^2 \nabla\rho \cdot \nabla\phi\quad {\rm and}\quad V_e ({\bm x}) = -\rho{\bm u}\times {\bm \omega}\cdot \nabla\phi,\] where θ is an acyclic potential, generated by the delta wing moving with unit velocity in the negative direction of the force (lift or drag). All the physical quantities are non-dimensionalized. Detailed force contributions are analysed in terms of the flow structures and the elements Re(x) and Ve(x). The source elements Re(x) and Ve(x) are concentrated in the following regions: the boundary layer in front of (below) the delta wing, the primary and secondary vortices over the delta wing, and a region of expansion around the leading edge. It is shown that Ve(x) due to vorticity prevails as the source of forces at relatively low Mach number, Mα < 0.7. Above about Mα = 0.75, Re(x) due to compressibility generally becomes the dominating contributor to the lift, while the overall contribution from Ve(x) decreases with increasing Mα, and even becomes negative at Mα = 1.2 for the lift, and at a higher Mα for the drag. The analysis is carried out with the aid of detailed numerical results by solving the Reynolds-averaged Navier–Stokes equations, which are in close agreement with experiments in comparisons of the surface pressure distributions.

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