scholarly journals Aerodynamic performance of chevron geometry trailing edge for a transonic axial compressor impeller

2021 ◽  
Vol 33 ◽  
pp. 184-194
Author(s):  
Ioana Octavia Bucur ◽  
Ion Mălăel ◽  
Valeriu Drăgan
Keyword(s):  
Interaction Mechanism ◽  
Axial Compressor ◽  
Experimental Tests ◽  
Rotor Blades ◽  
Data Availability ◽  
Noise Mitigation ◽  
Trailing Edge ◽  
Compressor Impeller ◽  
Rotor Stator Interaction ◽  
Transonic Axial Compressor

Several studies present how serrated rotor blades mix wakes in order to attenuate noise levels. The current paper analyses how this geometry, applied on the trailing edge, affects the global parameters of a transonic axial compressor impeller. Innovative solutions tackling the rotor-stator interaction mechanism in an axial compressor for noise reduction include serrated trailing edges. Inspired by chevron nozzles, serrations can be transferred to the open-rotor concept in order to reduce tonal noise. Throughout the study we will be focusing on aerodynamic loss estimation while being mindful of the mechanisms which lead to rotor-stator interaction noise, without assessing its per-se effectiveness for noise mitigation. Owing to its qualitative experimental data availability, NASA’s Rotor 37 was chosen as a baseline. A set of fully viscous 3D simulations, using the SST k omega turbulence model and RANS, was carried out to this effect. Spatial discretization was made using a fully structured pre-mesh in order to optimize resolution and accelerate convergence. Full-factorial samples were generated for the geometric variations in order to capture the aerodynamic implications of this concept. Overall, the analysed case provides promising perspectives, pending optimization studies and experimental tests thereof.

Effect of Trailing Edge Crenulation on the Performance and Stall Margin of a Transonic Axial Compressor

2013 ◽  
Author(s):  
S. Subbaramu ◽  
Quamber H. Nagpurwala ◽  
A. T. Sriram
Keyword(s):  
Beneficial Effect ◽  
Total Pressure ◽  
Pressure Ratio ◽  
Axial Compressor ◽  
Compressor Cascade ◽  
Trailing Edge ◽  
Stall Margin ◽  
Positive Effects ◽  
Compressor Rotor ◽  
Transonic Axial Compressor

This paper deals with the numerical investigations on the effect of trailing edge crenulation on the performance of a transonic axial compressor rotor. Crenulation is broadly considered as a series of small notches or slots at the edge of a thin object, like a plate. Incorporating such notches at the trailing edge of a compressor cascade has shown beneficial effect in terms of reduction in total pressure loss due to enhanced mixing in the wake region. These notches act as vortex generators to produce counter rotating vortices, which increase intermixing between the free stream flow and the low momentum wake fluid. Considering the positive effects of crenulation in a cascade, it was hypothesized that the same technique would work in a rotating compressor to enhance its performance and stall margin. However, the present CFD simulations on a transonic compressor rotor have given mixed results. Whereas the peak total pressure ratio in the presence of trailing edge crenulation reduced, the stall margin improved by 2.97% compared to the rotor with straight edge blades. The vortex generation at the crenulated trailing edge was not as strong as reported in case of linear compressor cascade, but it was able to influence the flow field in the rotor tip region so as to energize the low momentum end-wall flow in the aft part of the blade passage. This beneficial effect delayed flow separation and allowed the mass flow rate to be reduced to still lower levels resulting in improved stall margin. The reduction in pressure ratio with crenulation was surprising and might be due to increased mixing losses downstream of the blade.

Analysis of the Inter-Row Flow Field Within a Transonic Axial Compressor: Part 2 — Unsteady Flow Analysis

2000 ◽  
Author(s):  
Xavier Ottavy ◽  
Isabelle Trébinjac ◽  
André Vouillarmet
Keyword(s):  
Flow Field ◽  
Unsteady Flow ◽  
Flow Analysis ◽  
Axial Compressor ◽  
Leading Edge ◽  
Rotor Blades ◽  
Time Averaging ◽  
Flow Phenomena ◽  
Order Of Magnitude ◽  
Transonic Axial Compressor

An analysis of the experimental data, obtained by laser two-focus anemometry in the IGV-rotor inter-row region of a transonic axial compressor, is presented with the aim of improving the understanding of the unsteady flow phenomena. A study of the IGV wakes and of the shock waves emanating from the leading edge of the rotor blades is proposed. Their interaction reveals the increase in magnitude of the wake passing through the moving shock. This result is highlighted by the streamwise evolution of the wake vorticity. Moreover, the results are analyzed in terms of a time averaging procedure and the purely time-dependent velocity fluctuations which occur are quantified. It may be concluded that they are of the same order of magnitude as the spatial terms for the inlet rotor flow field. That shows that the temporal fluctuations should be considered for the 3D rotor time-averaged simulations.

scholarly journals Experimental investigation of unsteady flow phenomena in a three-stage axial compressor

2003 ◽  
Vol 217 (4) ◽  
pp. 341-348 ◽  
Author(s):  
R Niehuis ◽  
A Bohne ◽  
A Hoynacki
Keyword(s):  
Measurement Techniques ◽  
Axial Compressor ◽  
Global Scale ◽  
Rotor Blades ◽  
Tip Clearance ◽  
Tip Clearance Flow ◽  
Clearance Flow ◽  
Flow Phenomena ◽  
Rotor Stator Interaction ◽  
Interaction Phenomena

In the past years, a three-stage axial compressor equipped with a modern controlled diffusion airfoil (CDA) blading has been investigated in much detail, applying state-of-the-art steady and unsteady measurement techniques, at RWTH Aachen University. The compressor under investigation exhibits design features of real industrial compressors. By performing high-resolution measurements both in space and time, a thorough insight into various flow phenomena in the compressor has been achieved, leading to a better understanding of various flow phenomena such as rotor—stator interaction, tip clearance flow and viscous flow effects in a multistage compressor environment. After a short summary of some performance characteristics at design and off-design, this paper focuses on the analysis of interaction phenomena present in the three-stage axial compressor. The interaction phenomena are described on a more global scale. In order to quantify the upstream and downstream influence of the three rotor blades, a suitable parameter is presented.

Squealer Tip Treatment Design for Axial Compressors

2020 ◽  
Author(s):  
Giuseppe Bruni ◽  
James Taylor ◽  
Senthil Krishnababu ◽  
Robert Miller ◽  
Roger Wells
Keyword(s):  
Thickness Distribution ◽  
Axial Compressor ◽  
Leading Edge ◽  
Experimental Tests ◽  
Suction Side ◽  
Rotor Blades ◽  
Superior Performance ◽  
Stall Margin ◽  
Wall Flows ◽  
Multi Stage

Abstract End-wall flows are amongst the main sources of losses in the rear stages of a typical multi-stage axial compressor. Reducing the tip leakage losses in the rotor blades and vanes can provide an increased efficiency and stall margin of a given axial compressor stage. One approach is to use squealer tips, which are traditionally designed to minimize the effect of tip rubbing. However, squealers can also provide a significant performance benefit, when designed considering aerodynamics from the beginning, as shown in this paper. A CFD based methodology, in which the blade or vane thickness distribution is varied in a controlled manner was developed. This design methodology was used to create different types of squealer tip geometry for a representative stage in a low speed compressor rig. Three different tip concepts were designed, based on a Suction Side Squealer, on a Pressure Side Squealer and on the combination of the two being merged between the leading edge and trailing edge, this new design is called the SuPr Tip. Subsequent experimental tests carried out agreed with the predicted relative ranking of the different squealer designs and on the superior performance of the SuPr tip design over the others, thus validating the methodology and the design process.

Effects of Blade Damage on the Performance of a Transonic Axial Compressor Rotor

2012 ◽  
Author(s):  
Yanling Li ◽  
Abdulnaser Sayma
Keyword(s):  
Axial Compressor ◽  
Rotating Stall ◽  
Rotor Blades ◽  
Design Stage ◽  
Compressor Blades ◽  
Stall Margin ◽  
Compressor Rotor ◽  
Surge Margin ◽  
Compressor Performance ◽  
Transonic Axial Compressor

Gas turbine axial compressor blades may encounter damage during service for various reasons. Debris from casing or foreign objects may impact blades causing damage near the rotor’s tip. This may result in deterioration of performance and reduction in the surge margin. Ability to assess the effect of damaged blades on the compressor performance and stability is important at both the design stage and in service. The damage to compressor blades breaks the cyclic symmetry of the compressor assembly. Thus computations have to be performed using the whole annulus. Moreover, if rotating stall or surge occurs, the downstream boundary conditions are not known and simulations become difficult. This paper presents an unsteady CFD analysis of compressor performance with tip curl damage. Tip curl damage typically occurs when rotor blades hit a loose casing liner. The computations were performed up to the stall boundary, predicting rotating stall patterns. The aim is to assess the effect of blade damage on stall margin and provide better understanding of the flow behaviour during rotating stall. Computations for the undamaged rotor are also performed for comparison. A transonic axial compressor rotor is used for the time-accurate numerical unsteady flow simulations, with a variable choked nozzle downstream simulating an experimental throttle. One damaged blade was introduced in the rotor assembly and computations were performed at 60% of the design rotational speed. It was found that there is no significant effect on the compressor stall margin due to one damaged blade despite the differences in rotating stall patterns between the undamaged and damaged assemblies.

scholarly journals Effect of the Rotor Blades Part Span Shrouds on the Transonic Axial Compressor Operation Map

2020 ◽  
Vol 1670 ◽  
pp. 012015
Author(s):  
A.A. Sebelev ◽  
A.A. Schengals ◽  
V.A. Aleksenskiy ◽  
A. Yu. Tamm ◽  
O.I. Klyavin
Keyword(s):  
Axial Compressor ◽  
Rotor Blades ◽  
Transonic Axial Compressor

scholarly journals Stability Enhancement of a Single-Stage Transonic Axial Compressor Using Inclined Oblique Slots

Energies ◽  
2021 ◽  
Vol 14 (9) ◽  
pp. 2346
Author(s):  
Tien-Dung Vuong ◽  
Kwang-Yong Kim
Keyword(s):  
Pressure Ratio ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Single Stage ◽  
Rotor Blades ◽  
Navier Stokes ◽  
Optimized Design ◽  
Stall Margin ◽  
Casing Treatment ◽  
Transonic Axial Compressor

A casing treatment using inclined oblique slots (INOS) is proposed to improve the stability of the single-stage transonic axial compressor, NASA Stage 37, during operation. The slots are installed on the casing of the rotor blades. The aerodynamic performance was estimated using three-dimensional steady Reynolds-Averaged Navier-Stokes analysis. The results showed that the slots effectively increased the stall margin of the compressor with slight reductions in the pressure ratio and adiabatic efficiency. Three geometric parameters were tested in a parametric study. A single-objective optimization to maximize the stall margin was carried out using a Genetic Algorithm coupled with a surrogate model created by a radial basis neural network. The optimized design increased the stall margin by 37.1% compared to that of the smooth casing with little impacts on the efficiency and pressure ratio.

Acoustic Minimization of Rotor-Stator Interaction Noise of Axial Fans by Trailing-Edge Blowing

2014 ◽  
Author(s):  
Michael Kohlhaas ◽  
Thomas H. Carolus
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Rotor Blades ◽  
The Novel ◽  
Optimization Strategy ◽  
Trailing Edge ◽  
Rotor Stator Interaction ◽  
Axial Fans ◽  
Novel Strategy

The work deals with tonal noise of axial turbo fan stages by rotor-stator interaction. The objective of this study is the reduction of this noise by injection of a secondary mass flow at the rotor blades’ trailing edge (trailing edge blowing, TEB). A literature survey suggested that at least the tonal rotor/stator interaction sound can be reduced considerably by perfect blade wake filling which usually is achieved by blowing air through slots in the blade trailing edge region. Our own studies proved that this traditional trailing edge blowing strategy is problematic. Hence, in this paper a novel strategy of trailing edge blowing is described. The key idea is the combination of experimental trailing edge blowing with an evolutionary optimization algorithm. Aiming directly at a minimum of far field sound pressure level we identified spanwise trailing edge flowing distributions that reduced the fundamental tone at blade passing frequency (BPF) by 1.4 dB, and at its first harmonic by 21.4 dB (the latter corresponding to a complete elimination). The required blowing mass flow rate is 2% of the mass flow rate through the stage and hence relatively moderate. 3D hot wire measurement revealed that the acoustically relevant upwash velocity fluctuations as seen by the stator are not only caused by the wakes from the rotor blades but also by vortex structures at the rotor’s hub and tip. Traditional trailing edge blowing primarily aims at eliminating the rotor blade wakes. By applying the novel optimization strategy even some (but not all) of those secondary flow structures could be put out of action which eventually led to the reduction of the spectral components observed.

Influence of Local Surface Roughness of Rotor Blade on Performance of an Axial Compressor Stage

2013 ◽  
Author(s):  
Shaowen Chen ◽  
Shijun Sun ◽  
Hao Xu ◽  
Longxin Zhang ◽  
Songtao Wang ◽  
...  
Keyword(s):  
Surface Roughness ◽  
Mass Flow ◽  
Fluid Dynamic ◽  
Axial Compressor ◽  
Leading Edge ◽  
Rotor Blades ◽  
Compressor Stage ◽  
Local Surface ◽  
Trailing Edge ◽  
Pressure Surface

The influence of local surface roughness of rotor blades on the performance of axial compressor stages were investigated through numerical simulation with local surface roughness added on the suction and pressure surfaces of rotor blades of realistic compressor stage NASA Stage35. First of all, the reliability of a commercial computational fluid dynamic code was validated and the computed performance maps showed a good agreement with experimental data from literatures. Numerical results indicated that the increase in surface roughness in most of the local positions may cause the deterioration of compressor stage performance. The amplitude of decrease in compressor performance due to the addition of surface roughness in outer and inner portions of the span and the area near the leading edge of the rotor blades would be much greater than that in the region near the trailing edge. The roughness added to the pressure surface near the leading edge had less impact on the stage characteristics, including the mass flow rate at the choked point. Thus the compressor characteristic got close to that under normal conditions and showed a wider stable operating range. The mass flow in the choked region and the adiabatic efficiency were less affected by the roughness added to the region near the trailing edge of pressure surface from rotor blades. However, this scheme mentioned before would increase the total pressure ratio to some extent, with the adverse effect of adding roughness on the corresponding suction surface.

Sign in / Sign up

Export Citation Format

Share Document