AERODYNAMIC INFLUENCE OF A BLEED ON THE LAST STAGE OF A LOW-PRESSURE COMPRESSOR AND S-DUCT

2021 ◽  
pp. 1-39
Author(s):  
Apostolos Spanelis ◽  
A Duncan Walker
Keyword(s):  
Radial Flow ◽  
Guide Vane ◽  
Low Pressure ◽  
Flow Distortion ◽  
Tip Leakage ◽  
High Pressure Compressor ◽  
Unsteady Rans ◽  
Last Stage ◽  
Flow Variables ◽  
Pressure Compressor

Abstract This paper uses Computational Fluid Dynamics to investigate the effect of an engine handling bleed situated on the outer casing downstream of the last rotor stage of a low-pressure compressor and upstream of the outlet guide vane and S-shaped duct. The model, validated against existing experimental data, utilized an unsteady RANS solver incorporating a Reynolds stress closure to examine the unsteady component interactions. The results showed that at bleed rates less than 25% of the mainstream flow the bleed effects were negligible. However, at higher bleed rates performance was significantly degraded. A uniform flow extraction hypothesis was employed to separate the positional bias effects from the bulk flow diffusion. This revealed that the bleed-induced radial flow distortion can significantly affect the OGV loading distribution, which thereby dictates the position and type of stall within the OGV passage. Extraction of the rotor tip leakage via the shroud bleed, combined with the radial flow distortion, contributed to a 28% reduction in duct loss at 10% bleed and up to 50% reduced loss at 25% bleed. The actual amount of flow required to be extracted for an OGV stall to develop, was 30%. That was independent of the bleed location and the type of stall. For bleeds up to 20%, the S-duct displayed a remarkable resilience and consistency of flow variables at duct exit. However, a stalled OGV deteriorated the radial flow uniformity that was presented to the high-pressure compressor.

Low-Pressure Compressor Near-Stall Predictions Using Unsteady CFD Methods

2021 ◽  
Author(s):  
David Vanpouille ◽  
Dimitrios Papadogiannis ◽  
Stéphane Hiernaux
Keyword(s):  
Low Pressure ◽  
Navier Stokes ◽  
Phase Lag ◽  
Sliding Mesh ◽  
Tip Leakage ◽  
Eddy Simulation ◽  
Surge Margin ◽  
Large Eddy ◽  
Unsteady Rans ◽  
Pressure Compressor

Abstract Surge margin is critical for the safety of aeronautical compressors, hence predicting it early in the design process using CFD is mandatory. However, close to surge, steady-state Reynolds Averaged Navier-Stokes (RANS) simulations are proven inadequate. Unsteady techniques such as Unsteady RANS (URANS) and Large Eddy Simulation (LES) can provide more reliable predictions. Nevertheless, the accuracy of such methods are dependent on the method used to handle the rotor/stator interfaces. The most precise method, the sliding mesh, requires simulating the full annulus or a periodic sector, which can be very costly. Other techniques to reduce the domain exist, such as the phase-lagged approach or geometric blade scaling, but introduce restrictive assumptions on the flow at near-stall conditions. The objective of this paper is to investigate the near-stall flow of a low-pressure compressor using unsteady methods of varying fidelity: URANS with the phase lag assumption, URANS on a periodic sector and a high-fidelity LES on a smaller periodic sector achieved using geometric blade scaling. Results are compared to experimental measurements. An overall good agreement is found. Results show that the tip leakage vortex is not the origin of the stall on the studied configuration and a hub corner separation is initiated. LES further validates the (U)RANS flow predictions and brings additional insight on unsteady flow separations.

Experimental and Computational Investigation of Rayleigh-Bénard Flow in the Rotating Cavities of a Core Compressor

2017 ◽  
Author(s):  
M. R. Puttock-Brown ◽  
M. G. Rose ◽  
C. A. Long
Keyword(s):  
Free Convection ◽  
Nusselt Numbers ◽  
Rotating Cavity ◽  
High Pressure Compressor ◽  
Aero Engine ◽  
Peripheral Surface ◽  
Unsteady Rans ◽  
The University ◽  
Rayleigh Benard ◽  
Pressure Compressor

This paper presents new experimental measurements, at conditions representative of an aero engine, of heat transfer from the inner peripheral surface (shroud) of a rotating cavity. The results are taken from the University of Sussex Multiple Cavity Rig, which is designed to be similar to a gas turbine high pressure compressor internal air system. The shroud Nusselt numbers are shown to be dependent on the shroud Grashof number and insensitive to throughflow axial Reynolds number. The magnitude of the shroud Nusselt numbers are consistent with accepted correlations for turbulent free convection from a horizontal plate, yet show a trend (gradient of Nusselt to Grashof numbers) that is similar to laminar free convection. A supporting high-resolution 3D unsteady RANS simulation was conducted to investigate the cavity flow structure with particular attention paid to the near shroud region. This demonstrated flow structures that are consistent with published work but also show the existence of a type of Rayleigh-Bénard flow that manifests as a series of streaks that propagate along the periphery of the cavity. These structures can be found in the literature albeit in different circumstances. Whilst these streaks have been shown in the simulation their existence cannot be ratified without experimental confirmation.

The Impact of Bleeding on Compressor Stator Corner Separation

2011 ◽  
Author(s):  
Bin Zhao ◽  
Shaobin Li ◽  
Qiushi Li ◽  
Sheng Zhou
Keyword(s):  
Radial Flow ◽  
Single Stage ◽  
Corner Separation ◽  
Transonic Compressor ◽  
High Pressure Compressor ◽  
Compressor Performance ◽  
The Impact ◽  
Large Flow ◽  
Compressor Efficiency ◽  
Pressure Compressor

Bleed air from the high pressure compressor has taken up 3–5% in the air system. However, there are not many studies on the compressor performance after bleeding. By analyzing a low-speed single-stage compressor and a transonic single-stage compressor, this paper presents several plans with different bleeding rates on the casing near stator corner, in order to study the influence of bleeding rates on the compressor stator corner separation. The results showed that for the stators of subsonic compressor with large flow separation in the corner, there is an optimum value in the stator casing bleed air amount. The flow field is better at resisting the radial flow caused by bleed air in the transonic compressor stator. The more the bleeding rates of the stator, the more the compressor efficiency improves.

Corrosion Studies of High and Low Pressure Compressor Blades of a Gas Turbine

2005 ◽  
Author(s):  
A. Boschetti ◽  
E. Y. Kawachi ◽  
M. A. S. Oliveira
Keyword(s):  
High Pressure ◽  
Gas Turbine ◽  
Low Pressure ◽  
Compressor Blade ◽  
Compressor Blades ◽  
X Ray ◽  
Corrosion Studies ◽  
High Pressure Compressor ◽  
Pressure Compressor ◽  
Base Superalloy

This work presents preliminary results of corrosion studies for three blades, one of the low pressure compressor and two of two different stages of the high pressure compressor of a gas turbine, which has been operating for 5,000 hours. Scanning Electron Microscopy (SEM), Energy Dispersive X-ray Spectroscopy (EDS), X-ray diffraction (XRD), Electrochemical Impedance Spectroscopy (EIS) in aqueous solution containing chloride, and Atomic Absorption Spectrometry (AAS) were used to characterize the blades surfaces. The SEM and EDS results showed that the homogeneity and amount of contaminants, such as sodium, potassium, calcium, magnesium, chloride and sulphur are bigger in the high pressure compressor blade surfaces than in the low pressure compressor blade surface. The EIS results showed that the degradation process in turbine compressor blades increases with the temperature and pressure increase inside the compressors and depends of the blade composition. The low pressure compressor blade, which was made of a Ti base superalloy exhibited smaller corrosion resistance (smallest charge transfer resistance value (Rct)) than the two high pressure compressor blades, which were made of a Fe base superalloy. However, despite of its lower resistance to corrosion, after 5,000 hours of service, the low pressure compressor blade did not present pitting corrosion while the high pressure compressor blades did.

Asymmetric Solution for Compressor Station Spare Capacity

2006 ◽  
Author(s):  
Rainer Kurz ◽  
Matt Lubomirsky
Keyword(s):  
High Pressure ◽  
Compression Ratio ◽  
Low Pressure ◽  
Compressor Station ◽  
Spare Capacity ◽  
High Pressure Compressor ◽  
Power Turbine ◽  
In Series ◽  
Asymmetric Solution ◽  
Pressure Compressor

Arranging compressor units in series rather than in parallel can offer a number of advantages in many applications. However, especially if the compression ratio is rather high, it is desirable to use different impellers for the low-pressure and the high-pressure compressor. Otherwise, one or both operate off their best efficiency points. This leaves however the problem of the staging of a spare unit. In the paper, a solution to this dilemma is described. To understand the underlying constraints, the relationships between the system resistance, the compressor characteristic, and the power turbine characteristic are analytically derived.

Advanced Exergy Analysis of a Turbofan Engine (TFE): Splitting Exergy Destruction into Unavoidable/Avoidable and Endogenous/Exogenous

2017 ◽  
Vol 0 (0) ◽  
Author(s):  
Ozgur Balli
Keyword(s):  
High Pressure ◽  
Exergy Analysis ◽  
Low Pressure ◽  
Operating Conditions ◽  
Actual Condition ◽  
Exergy Destruction ◽  
Turbofan Engine ◽  
High Pressure Compressor ◽  
Engine Components ◽  
Pressure Compressor

AbstractA conventional and advanced exergy analysis of a turbofan engine is presented in this paper. In this framework, the main exergy parameters of the engine components are introduced while the exergy destruction rates within the engine components are split into endogenous/exogenous and avoidable/unavoidable parts. Also, the mutual interdependencies among the components of the engine and realistic improvement potentials depending on operating conditions are acquired through the analysis. As a result of the study, the exergy efficiency values of the engine are determined to be 25.7 % for actual condition, 27.55 % for unavoidable condition and 30.54 % for theoretical contion, repectively. The system has low improvement potential because the unavoidable exergy destruction rate is 90 %. The relationships between the components are relatively weak since the endogenous exergy destruction is 73 %. Finally, it may be concluded that the low pressure compressor, the high pressure compressor, the fan, the low pressure compressor, the high pressure compressor and the combustion chamber of the engine should be focused on according to the results obtained.

Automated Aerodynamic Optimization of an Aggressive S-Shaped Intermediate Compressor Duct

2018 ◽  
Author(s):  
T. Stürzebecher ◽  
G. Goinis ◽  
C. Voss ◽  
H. Sahota ◽  
P. Groth ◽  
...  
Keyword(s):  
Guide Vane ◽  
Aerodynamic Optimization ◽  
Baseline Design ◽  
Separating Flow ◽  
High Pressure Compressor ◽  
Important Research Topic ◽  
Final Design ◽  
Endwall Contouring ◽  
Aero Engines ◽  
Pressure Compressor

As bypass-ratio in modern aero engines is continuously increasing over the last decades, the radial offset between low pressure compressor (LPC) and high pressure compressor (HPC), which needs to be overcome by the connecting s-shaped intermediate compressor duct (ICD), is getting higher. Due to performance and weight saving aspects the design of shorter and therefore more aggressive ducts has become an important research topic. In this paper an already aggressive design (with respect to current aero engines) of an ICD with integrated outlet guide vane (OGV) is used as a baseline for an aerodynamic optimization. The aim is to shorten the duct even further while maintaining it separation free. The optimization is broken down into two steps. In the first optimization-step the baseline design is shortened to a feasible extent while keeping weak aerodynamic restrictions. The resulting highly aggressive duct (intermediate design), which is shortened by 19 % in axial length with respect to the baseline, shows separation tendencies of low momentum fluid in the strut/hub region. For the second step, the length of the optimized duct design is frozen. By implementing new design features in the process of the optimizer, this optimization-step aims to eliminate separation and to reduce separation tendencies caused by the aggressive shortening. In particular, these features are: a nonaxisymmetric endwall contouring and parametrization of the strut and the OGV to allow for changes in lift and turning in both blade designs. By comparison of the three designs: Baseline, intermediate (separating flow) and final design, it can be shown, that it is possible to decrease length of the already aggressive baseline design even further, when adding a nonaxisymmetric endwall contouring and changes in blade shape of the strut and OGV. Flow separation can be eliminated while losses are kept low. With a more aggressive and therefore shorter duct the engine length and weight can be reduced. This in turn leads to lighter aircrafts, less fuel consumption and lower CO2 and NOx emissions.

Three-Dimensional Blade Stacking Strategies and Understanding of Flow Physics in Low-Pressure Steam Turbines—Part II: Stacking Equivalence and Differentiators

2015 ◽  
Vol 138 (6) ◽  
Author(s):  
Said Havakechian ◽  
John Denton
Keyword(s):  
Guide Vane ◽  
Three Dimensional ◽  
Low Pressure ◽  
Cfd Modeling ◽  
Steam Turbines ◽  
Flow Parameters ◽  
Pressure Steam ◽  
Computational Fluid Dynamics Cfd ◽  
Last Stage ◽  
Two Stages

Optimization of blade stacking in low-pressure (LP) steam turbine development constitutes one of the most delicate and time-consuming parts of the design process. This is the second part of two papers focusing on stacking strategies applied to the last stage guide vane and represents an attempt to discern the aerodynamic targets that can be achieved by each of the well-known and most often used basic stacking schemes. The effects of lean and twist have been investigated through an iterative process, involving comprehensive 3D computational fluid dynamics (CFD) modeling of the last two stages of a standard LP, where the basic lean and twist stacking schemes were applied on the last stage guide vanes while keeping the throat area (TA) unchanged. It has been found that it is possible to achieve the same target value and pattern of stage reaction by applying either tangential lean or an equivalent value of twist. Moreover, the significance of axial sweep on hub reaction has been found to become pronounced when the blade sweep is carried out at constant TA. The importance of hub-profiling has also been demonstrated and assessed. Detailed analysis of the flow fields has provided an overall picture, revealing the differences in the main flow parameters as produced by each of the alternative basic stacking schemes.

Effect of Tip Clearance Dimensions and Control of Unsteady Flows in a Multi-Stage High-Pressure Compressor

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Nicolas Gourdain ◽  
Fabien Wlassow ◽  
Xavier Ottavy
Keyword(s):  
High Pressure ◽  
Unsteady Flows ◽  
Tip Clearance ◽  
Leakage Flow ◽  
Tip Leakage Flow ◽  
Tip Leakage ◽  
High Pressure Compressor ◽  
Stable Operating ◽  
Overall Performance ◽  
Pressure Compressor

This paper describes the investigations performed to better understand unsteady flows that develop in a three-stage high-pressure compressor. More specifically, this study focuses on rotor-stator interactions and tip leakage flow effects on overall performance and aerodynamic stability. The investigation method is based on three-dimensional unsteady RANS simulations, considering the natural spatial periodicity of the compressor. Indeed, all information related to rotor-stator interactions can be computed. A comparison is first done with experimental measurements to outline the capacity of the numerical method to predict overall performance and unsteady flows. The results show that the simulation correctly estimates most flow features in the multistage compressor. Then numerical data obtained for three configurations of the same compressor are analyzed and compared. Configurations 1 and 2 consider two sets of tip clearance dimensions and a casing treatment based on a honeycomb design is applied for configuration 3. Detailed investigations of the flow at the same operating line show that the tip leakage flow is responsible for the loss of stability in the last stage. An increase by 30% of the tip clearance dimension dramatically reduces the stable operating range (by 40% with respect to the standard configuration). A modal analysis shows that the stall process in this case involves the perturbation of the flow in the last rotor by upstream stator wakes, leading to the development of a rotating instability. The control device designed and investigated in this study allows for reducing the sensitivity of the compressor to tip leakage flow by recovering the initial stable operating range.

An Experimental Investigation of a Tandem Stator Flow Characteristic in a Low Speed Axial Research Compressor

2014 ◽  
Author(s):  
Alrik Tesch ◽  
Martin Lange ◽  
Konrad Vogeler ◽  
Jens Ortmanns ◽  
Erik Johann ◽  
...  
Keyword(s):  
Swirling Flow ◽  
Guide Vane ◽  
Axial Compressor ◽  
Low Speed ◽  
High Pressure Compressor ◽  
Oil Flow ◽  
Compressor Performance ◽  
Operating Points ◽  
Insight Into ◽  
Pressure Compressor

A major goal in axial compressor development is to increase the efficiency and to reduce the weight of the module. In order to do so the power density has to be increased by raising the work per stage. Higher capability to do work can be achieved by increasing the circumferential velocity component of the fluid. Tandem stators might offer the ability to turn high swirling flow with lower losses compared to a single blade stator. In terms of higher aerodynamic loading the use of tandem vanes can be a key feature to allow the design of highly efficient and compact compressor modules. This paper presents the design and experimental validation of a single stage low speed axial compressor with a tandem outlet guide vane, representative for a modern jet engine high pressure compressor. Additionally to the overall compressor performance the 3D flow field of the tandem stator has been measured with a five hole probe at different operating points. The results will be discussed in comparison with numerical results. Furthermore, oil flow pictures are used to get a deeper insight into flow conditions inside the vane passage and to validate the numerically predicted secondary flow structures.

Sign in / Sign up

Export Citation Format

Share Document