Impact of Uniform Surface Roughness on the Aerodynamic Performance of an Axial Compressor Rotor at Different Rotational Speeds

2021 ◽  
Author(s):  
Ashima Malhotra ◽  
Shraman Goswami ◽  
Pradeep A M
Keyword(s):  
Surface Roughness ◽  
Numerical Study ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Performance Loss ◽  
Compressor Rotor ◽  
Multistage Compressor ◽  
Design Speed ◽  
The Impact ◽  
Relative Flow

Abstract The aerodynamic performance of a compressor rotor is known to deteriorate due to surface roughness. It is important to understand this deterioration as it impacts the overall performance of the engine. This paper, therefore, aims to numerically investigate the impact of roughness on the performance of an axial compressor rotor at different rotational speeds. In this numerical study, the simulations are carried out for NASA Rotor37 at 100%, 80%, and 60% of its design speed. with and without roughness on the blade surface. These speeds are chosen because they represent different flow regimes. The front stages of a multistage compressor usually have a supersonic or transonic regime whereas the middle and aft stages have a subsonic regime. Thus, these performance characteristics can give an estimate of the impact on the performance of a multistage compressor. At 100% speed (design speed), the relative flow is supersonic, at 80% of design speed, the relative flow is transonic and at 60% of design speed, the relative flow is subsonic. Detailed flow field investigations are carried out to understand the underlying flow physics. The results indicate that, for the same amount of roughness, the degradation in the performance is maximum at 100% speed where the rotor is supersonic, while the impact is minimum at 60% speed where the rotor is subsonic. Thus, the rotor shock system plays an important role in determining the performance loss due to roughness. It is also observed that the loss increases with increased span for 100% and 80% speeds, but for 60% speed, the loss is almost constant from the hub to the shroud. This is because, with the increased span, the shock strength increases for 100% and 80% speeds, whereas at 60% speed flow is subsonic.

scholarly journals The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

1995 ◽  
Vol 117 (4) ◽  
pp. 491-505 ◽  
Author(s):  
K. L. Suder ◽  
R. V. Chima ◽  
A. J. Strazisar ◽  
W. B. Roberts
Keyword(s):  
Surface Roughness ◽  
Surface Finish ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Blade Surface ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Performance Deterioration ◽  
Design Speed

The performance deterioration of a high-speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 rms μm (100–125 rms μin.) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10 percent at the hub and 20 percent at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9 percent reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 rms μm (10–20 rms μin.), compared to the bare metal blade surface finish of 0.508 rms pm (20 rms μin.). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60, 80, and 100 percent of design speed. The results indicate that thickness/roughness over the first 2 percent of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70 percent of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-three-dimensional Navier–Stokes flow solver, which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that adding roughness at the blade leading edge causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage, which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

scholarly journals The Effect of Adding Roughness and Thickness to a Transonic Axial Compressor Rotor

1994 ◽  
Author(s):  
Kenneth L. Suder ◽  
Rodrick V. Chima ◽  
Anthony J. Strazisar ◽  
William B. Roberts
Keyword(s):  
Surface Roughness ◽  
Surface Finish ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Performance Degradation ◽  
Blade Surface ◽  
Suction Surface ◽  
Compressor Rotor ◽  
Performance Deterioration ◽  
Design Speed

The performance deterioration of a high speed axial compressor rotor due to surface roughness and airfoil thickness variations is reported. A 0.025 mm (0.001 in.) thick rough coating with a surface finish of 2.54–3.18 RMS μm (100–125 RMS microinches) is applied to the pressure and suction surface of the rotor blades. Coating both surfaces increases the leading edge thickness by 10% at the hub and 20% at the tip. Application of this coating results in a loss in efficiency of 6 points and a 9% reduction in the pressure ratio across the rotor at an operating condition near the design point. To separate the effects of thickness and roughness, a smooth coating of equal thickness is also applied to the blade. The smooth coating surface finish is 0.254–0.508 RMS μm (10–20 RMS microinches), compared to the bare metal blade surface finish of 0.508 RMS μm (20 RMS microinches). The smooth coating results in approximately half of the performance deterioration found from the rough coating. Both coatings are then applied to different portions of the blade surface to determine which portions of the airfoil are most sensitive to thickness/roughness variations. Aerodynamic performance measurements are presented for a number of coating configurations at 60%, 80%, and 100% of design speed. The results indicate that thickness/roughness over the first 10% of blade chord accounts for virtually all of the observed performance degradation for the smooth coating, compared to about 70% of the observed performance degradation for the rough coating. The performance deterioration is investigated in more detail at design speed using laser anemometer measurements as well as predictions generated by a quasi-3D Navier-Stokes flow solver which includes a surface roughness model. Measurements and analysis are performed on the baseline blade and the full-coverage smooth and rough coatings. The results indicate that coating the blade causes a thickening of the blade boundary layers. The interaction between the rotor passage shock and the thickened suction surface boundary layer then results in an increase in blockage which reduces the diffusion level in the rear half of the blade passage, thus reducing the aerodynamic performance of the rotor.

Numerical Study on the Effect of Seal Leakage Flow on Low Pressure Axial Compressor Performance

2013 ◽  
Author(s):  
Balazs Farkas ◽  
Nicolas Van de Wyer ◽  
Jean-Francois Brouckaert
Keyword(s):  
Numerical Study ◽  
Axial Compressor ◽  
Test Facility ◽  
Leakage Flow ◽  
Stability Range ◽  
Compressor Rotor ◽  
Numerical Predictions ◽  
Aero Engines ◽  
The Stability ◽  
The Impact

This paper presents the extended numerical studies of a one and a half stage axial compressor designed for the LP compressor of a contra-rotating fan engine architecture. The essence of this architecture is given by the fact that the LP compressor rotor is mounted on the same shaft as the second fan stage which results in a lower rotational speed and therefore a much higher loading than in conventional high bypass-ratio aero-engines. The compressor itself was designed at VKI and subsequently tested in the closed loop test facility (VKI-R4) which allowed to compare numerical predictions with experimental data. In this study, particular interest was given to investigate the effect of the seal-leakage flow around the stator hub platform on the performance. To study the effect of the seal-leakage flow three different seal cavity configurations with different seal-tooth gaps sizes were simulated in comparison with no-cavity configuration. This set of investigations allowed to assess the different models by comparison with the results obtained experimentally. This comparison was made on the global performance of the stage, including the impact on the stability range, as well as on the flow field itself in particular in the rotor and stator exit planes. The computations were performed by using the Numeca developed code FINE™/Turbo with steady RANS solver.

Experimental Investigation of the Aerodynamic Performance of a Linear Axial Compressor Cascade With Water Droplet Loading

2010 ◽  
Author(s):  
Tjark Eisfeld ◽  
Franz Joos
Keyword(s):  
Water Injection ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Operating Conditions ◽  
Two Dimensional ◽  
Compressor Cascade ◽  
Compressor Rotor ◽  
Two Dimensional Flow ◽  
Wet Compression ◽  
The Impact

Wet compression operation is a commercially attractive way to increase power output and efficiency of a gas turbine cycle. In recent literature the impact of water loading on the aerodynamic performance of the blading has not been entirely clarified yet. The most significant issues of aerodynamics in wet compression are stage rematching and stability. Therefore, these subjects are investigated in a linear compressor rotor cascade. This setup allows an estimation of the aerodynamic performance of the blading from two-dimensional test data at various operating conditions. Moreover, the impact of droplet flow on the two-dimensional flow field of the blade passage is measured in detail in order to understand the deviation of performance parameters. The results indicate that the effect of water injection on compressor aerodynamics is strongly related to the operating condition. It appears that droplet loading has a beneficial effect on the flow at high blade loading.

scholarly journals The Aerodynamic Performance of an Over-the-Rotor Liner With Circumferential Grooves on a High Bypass Ratio Turbofan Rotor

2013 ◽  
Author(s):  
Rick Bozak ◽  
Christopher Hughes ◽  
James Buckley
Keyword(s):  
Pressure Ratio ◽  
Axial Compressor ◽  
Aerodynamic Performance ◽  
Oscillating Flow ◽  
Acoustic Liners ◽  
Performance Loss ◽  
Acoustic Liner ◽  
Blade Tip ◽  
Fan Blade ◽  
The Impact

While liners have been utilized throughout turbofan ducts to attenuate fan noise, additional attenuation is obtainable by placing an acoustic liner over-the-rotor. Previous experiments have shown significant fan performance losses when acoustic liners are installed over-the-rotor. The fan blades induce an oscillating flow in the acoustic liners which results in a performance loss near the blade tip. An over-the-rotor liner was designed with circumferential grooves between the fan blade tips and the acoustic liner to reduce the oscillating flow in the acoustic liner. An experiment was conducted in the W-8 Single-Stage Axial Compressor Facility at NASA Glenn Research Center on a 1.5 pressure ratio fan to evaluate the impact of this over-the-rotor treatment design on fan aerodynamic performance. The addition of a circumferentially grooved over-the-rotor design between the fan blades and the acoustic liner reduced the performance loss, in terms of fan adiabatic efficiency, to less than 1% which is within the repeatability of this experiment.

Numerical Study on Three-Dimensional Optimization of a Low-Speed Axial Compressor Rotor

2014 ◽  
Author(s):  
Chenkai Zhang ◽  
Jun Hu ◽  
Zhiqiang Wang ◽  
Chao Yin ◽  
Wei Yan
Keyword(s):  
Numerical Optimization ◽  
Large Scale ◽  
Numerical Study ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Leading Edge ◽  
Aerodynamic Performance ◽  
Trailing Edge ◽  
Compressor Rotor ◽  
Flow Loss

This paper presents numerical optimization of a compressor rotor, to deepen the knowledge of endwall flow in the large-scale axial subsonic compressor, accordingly reduce its endwall loss and improve its aerodynamic performance. With numerical simulation and numerical optimization tools, three-dimensional stacking principle is optimized to improve the design operation point performance for the rotor. Results show that, hub region of the rotor cannot undertake large blade loading; compared to the prototype rotor, obvious aerodynamic performance improvements locate near the hub area, and a certain degree of positive dihedral in this region effectively helps to reduce its flow loss. The effect of “loaded leading edge and unloaded trailing edge” due to positive dihedral was shown, which suppresses flow separation near the trailing edge, consequently obviously reduces the flow loss and largely improves the rotor aerodynamic performance.

scholarly journals Impact Investigation of Stator Seal Leakage on Aerodynamic Performance of Multistage Compressor

2021 ◽  
Vol 9 ◽  
Author(s):  
Qi Wang ◽  
Zhou Zhang ◽  
Qingsong Hong ◽  
Lanxue Ren
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Axial Compressor ◽  
Flow Distribution ◽  
Aerodynamic Performance ◽  
Flow Losses ◽  
Multistage Compressor ◽  
The Impact ◽  
Multistage Axial Compressor

In this paper, a numerical model based on the mass flow rate of seal leakage is presented, and a 3D numerical method of a multistage axial compressor with good engineering practicability is established. Validation consists of modeling a nine-stage axial compressor in all operating rotation speeds and calculating results of the performance characteristic curves in good agreement with test data. Comparisons are made against different cases of seal leakage mass flow rate for analyzing the impact of increasing seal leakage on the aerodynamic performance of the multistage axial compressor. The results indicate that the performance of the nine-stage axial compressor is degenerated faster and faster with seal leakage increasing in all operating working points, and the degeneration of performance of this compressor can be evaluated by the relationships of main performance parameters with the mass flow rate of seal leakage. Comparisons of flow distribution in the compressor for different cases of seal leakage also show that stators located in front stages of the multistage axial compressor are affected more seriously by the increasing seal leakage, and it can be confirmed that relatively larger flow losses in front stages bring significant impact on the decay of aerodynamic performance of a multistage axial compressor.

The Impact of Manufacturing Variations on Performance of a Transonic Axial Compressor Rotor

2018 ◽  
Author(s):  
Marcus Lejon ◽  
Niklas Andersson ◽  
Lars Ellbrant ◽  
Hans Mårtensson
Keyword(s):  
Flow Field ◽  
Geometric Mean ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Mean Deviation ◽  
Design Intent ◽  
Compressor Rotor ◽  
Measuring Machine ◽  
The Mean ◽  
The Impact

In this paper, the impact of manufacturing variations on performance of an axial compressor rotor are evaluated at design rotational speed. The geometric variations from the design intent were obtained from an optical coordinate measuring machine and used to evaluate the impact of manufacturing variations on performance and the flow field in the rotor. The complete blisk is simulated using 3D CFD calculations, allowing for a detailed analysis of the impact of geometric variations on the flow. It is shown that the mean shift of the geometry from the design intent is responsible for the majority of the change in performance in terms of mass flow and total pressure ratio for this specific blisk. In terms of polytropic efficiency, the measured geometric scatter is shown to have a higher influence than the geometric mean deviation. The geometric scatter around the mean is shown to impact the pressure distribution along the leading edge and the shock position. Furthermore, a blisk is analyzed with one blade deviating substantially from the design intent, denoted as blade 0. It is shown that the impact of blade 0 on the flow is largely limited to the blade passages that it is directly a part of. The results presented in this paper also show that the impact of this blade on the flow field can be represented by a simulation including 3 blade passages. In terms of loss, using 5 blade passages is shown to give a close estimate for the relative change in loss for blade 0 and neighboring blades.

Fouling on a Multistage Axial Compressor in the Presence of a Third Substance at the Particle/Surface Interface

2018 ◽  
Author(s):  
Nicola Aldi ◽  
Nicola Casari ◽  
Devid Dainese ◽  
Mirko Morini ◽  
Michele Pinelli ◽  
...  
Keyword(s):  
Impact Velocity ◽  
Gas Turbines ◽  
Particle Deposition ◽  
Fluid Dynamic ◽  
Particle Surface ◽  
Axial Compressor ◽  
Micro Particles ◽  
Multistage Compressor ◽  
Particle Ingestion ◽  
The Impact

Solid particle ingestion is one of the principal degradation mechanisms in the compressor and turbine sections of gas turbines. In particular, in industrial applications, the micro-particles not captured by the air filtration system can cause deposits on blades and, consequently, can result in a decrease in compressor performance. It is of great interest to the industry to determine which zones of the compressor blades are impacted by these small particles. However, this information often refers to single stage analysis. This paper presents three-dimensional numerical simulations of the micro-particle ingestion (0.15 μm – 1.50 μm) in a multistage (i.e. eight stage) subsonic axial compressor, carried out by means of a commercial CFD code. Particle trajectory simulations use a stochastic Lagrangian tracking method that solves the equations of motion separately from the continuous phase. The effects of humidity, or more generally, the effects of a third substance at the particle/surface interface (which is considered one of the major promoters of fouling) is then studied. The behavior of wet and oiled particles, in addition to the usual dry particles, is taken into consideration. In the dry case, the particle deposition is established only by using the sticking probability. This quantity links the kinematic characteristics of particle impact on the blade with the fouling phenomenon. In the other two cases, the effect of the presence of a third substance at the particle/surface interface is considered by means of an energy-based model. Moreover, the influence of the tangential impact velocity on particle deposition is analyzed. Introducing the effect of a third substance, such as humidity or oil, the phenomenon of fouling concerns the same areas of the multistage compressor. The most significant results are obtained by combining the effect of the third substance with the effect of the tangential component of the impact velocity of the particles. The deposition trends obtained with these conditions are comparable with those reported in literature, highlighting how the deposits are mainly concentrated in the early stages of a multistage compressor. Particular fluid dynamic phenomena, such as corner separations and clearance vortices, strongly influence the location of particle deposits.

scholarly journals Blockage Development in a Transonic, Axial Compressor Rotor

1997 ◽  
Author(s):  
Kenneth L. Suder
Keyword(s):  
Boundary Layer ◽  
Axial Compressor ◽  
Flow Region ◽  
Tip Clearance ◽  
Surface Boundary ◽  
Core Flow ◽  
The Core ◽  
Compressor Rotor ◽  
Transonic Axial Compressor ◽  
The Impact

A detailed experimental investigation to understand and quantify the development of blockage in the flow field of a transonic, axial flow compressor rotor (NASA Rotor 37) has been undertaken. Detailed laser anemometer measurements were acquired upstream, within, and downstream of a transonic, axial compressor rotor operating at 100%, 85%, 80%, and 60% of design speed which provided inlet relative Mach numbers at the blade tip of 1.48, 1.26, 1.18, and 0.89 respectively. The impact of the shock on the blockage development, pertaining to both the shock / boundary layer interactions and the shock / tip clearance flow interactions, is discussed. The results indicate that for this rotor the blockage in the endwall region is 2–3 times that of the core flow region, and the blockage in the core flow region more than doubles when the shock strength is sufficient to separate the suction surface boundary layer.

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