Impact of Manufacturing Variability on Multi–Stage High–Pressure Compressor Performance

2012 ◽  
Author(s):  
Alexander Lange ◽  
Matthias Voigt ◽  
Konrad Vogeler ◽  
Henner Schrapp ◽  
Erik Johann ◽  
...  
Keyword(s):  
High Pressure ◽  
Pressure Rise ◽  
Leading Edge ◽  
Back Pressure ◽  
Operating Conditions ◽  
Geometric Parameters ◽  
Maximum Efficiency ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Pressure Compressor

The present paper introduces a novel approach for considering manufacturing variability in the numerical simulation of a multi–stage high–pressure compressor (HPC). The manufacturing process is investigated by analyzing three of totally ten rotor rows. Therefore, 150 blades of each of the three rows were 3D scanned to obtain surface meshes of real blades. The deviation of a scanned blade to the design intent is quantified by a vector of 14 geometric parameters. Interpolating the statistical properties of these parameters provides the manufacturing scatter for all ten rotor rows expressed by 140 probability density functions. The probabilistic simulation utilizes the parametric scatter information for generating 200 virtual compressors. The CFD analysis provides the performance of these compressors by calculating speed lines. Post–processing methods are applied to statistically analyze the obtained results. It was found that the global performance parameters show a significantly wider scatter range for higher back pressure levels. The correlation coefficient and the coefficient of importance are utilized to identify the sensitivity of the results to the geometric parameters. It turned out that the sensitivities strongly shift for different operating points. While the leading edge geometry of all rotor rows dominantly influences the overall performance at maximum efficiency, the camber line parameters of the front stages become more important for higher back pressure levels. The analysis of the individual stage performance confirms the determining importance of the front stages — especially for highly throttled operating conditions. This leads to conclusions regarding the robustness of the overall HPC, which is principally determined by the efficiency and pressure rise of the front stages.

Impact of Manufacturing Variability on Multistage High-Pressure Compressor Performance

2012 ◽  
Vol 134 (11) ◽  
Author(s):  
Alexander Lange ◽  
Matthias Voigt ◽  
Konrad Vogeler ◽  
Henner Schrapp ◽  
Erik Johann ◽  
...  
Keyword(s):  
High Pressure ◽  
Pressure Rise ◽  
Leading Edge ◽  
Back Pressure ◽  
Operating Conditions ◽  
Geometric Parameters ◽  
Maximum Efficiency ◽  
Novel Approach ◽  
High Pressure Compressor ◽  
Pressure Compressor

The present paper introduces a novel approach for considering manufacturing variability in the numerical simulation of a multistage high-pressure compressor (HPC). The manufacturing process is investigated by analyzing three of a total of ten rotor rows. Therefore, 150 blades of each of the three rows were 3D scanned to obtain surface meshes of real blades. The deviation of a scanned blade to the design intent is quantified by a vector of 14 geometric parameters. Interpolating the statistical properties of these parameters provides the manufacturing scatter for all ten rotor rows expressed by 140 probability density functions. The probabilistic simulation utilizes the parametric scatter information for generating 200 virtual compressors. The CFD analysis provides the performance of these compressors by calculating speed lines. Postprocessing methods are applied to statistically analyze the obtained results. It was found that the global performance parameters show a significantly wider scatter range for higher back pressure levels. The correlation coefficient and the coefficient of importance are utilized to identify the sensitivity of the results to the geometric parameters. It turned out that the sensitivities strongly shift for different operating points. While the leading edge geometry of all rotor rows dominantly influences the overall performance at maximum efficiency, the camber line parameters of the front stages become more important for higher back pressure levels. The analysis of the individual stage performance confirms the determining importance of the front stages—especially for highly throttled operating conditions. This leads to conclusions regarding the robustness of the overall HPC, which is principally determined by the efficiency and pressure rise of the front stages.

High Pressure Compressor Aerodynamic Performance at Uncertain Boundary Conditions

2019 ◽  
Author(s):  
Philip Magin ◽  
Florian Danner ◽  
Matthias Voigt ◽  
Ronald Mailach
Keyword(s):  
High Pressure ◽  
Boundary Conditions ◽  
Performance Metrics ◽  
Aerodynamic Performance ◽  
Operating Conditions ◽  
Cycle Performance ◽  
Probabilistic Assessment ◽  
Ambient Conditions ◽  
High Pressure Compressor ◽  
Pressure Compressor

Abstract The intended operating point of turbomachinery is subject to numerous kinds of uncertainty. These range from varying ambient conditions, across geometric deviations in a component, to system related loading variability resulting in engine-to-engine variation in component matching. In order to guarantee safe operation at all conditions, it is essential to consider the above uncertainties when designing turbomachinery. In the present work, a probabilistic assessment is performed of the influence of possible operational uncertainties on the aerodynamic performance metrics of an aero-engine multistage high pressure compressor (HPC). To propagate uncertainties, Monte Carlo simulations (MCS) with Latin Hypercube Sampling (LHS) were performed, with both correlated and uncorrelated inputs. Each sample consisted of a steady state computational fluid dynamics (CFD) evaluation of the compressor. The statistical input for the boundary conditions was acquired from a MCS of the engine cycle performance at cruise, accounting for flight-to-flight variations in ambient conditions and engine-to-engine variations in component properties. With the chosen approach, it is possible to quantify the variability in aerodynamic performance of an HPC that is subject to uncertain operating conditions and thus shows the importance of input correlations. Results highlight that deterministically determined performance metrics can differ considerably from the statistical mean, revealing the benefits of a probabilistic assessment. In contrast to performing MCS on the cycle only, a CFD based assessment can also be used to draw conclusions on the aerodynamic mechanisms responsible for changes in efficiency or surge margin.

Jet Engine Compressor Blade Refurbishment by Application of the Advanced Re-Contouring Process

2003 ◽  
Author(s):  
Herwart T. Hoenen ◽  
Karsten Ellenberger
Keyword(s):  
High Pressure ◽  
Flow Behavior ◽  
Leading Edge ◽  
Compressor Blade ◽  
Tip Clearance ◽  
Jet Propulsion ◽  
High Pressure Compressor ◽  
Leading Edges ◽  
Pressure Compressor ◽  
Edge Contours

In modern jet propulsion Systems the core engine has an essential influence on the total engine performance. Especially the high pressure compressor plays an important role in this scheme. Substantial factors here are losses due to tip clearance effects and aerodynamic airfoil quality. During flight Operation the airfoils are subject to wear and tear on the leading edge. These effects cause a shortening of the chord length and the leading edge profiles become deformed. This results in a deterioration of the engine efficiency performance level and a reduced stall margin. The paper deals with the re-contouring of the leading edges of compressor airfoils. Lufthansa Technik AG in cooperation with the Institute of Jet Propulsion and Turbomachinery (RWTH Aachen University) developed a new method for the profile definition for the blade refurbishment. The common procedure of smoothing out the leading edges manually on a wheel grinding machine can not provide a defined contour nor a reproducible result of the overhaul process. In order to achieve optimized flow conditions in the compressor blade rows, suitable leading edge contours have to be defined for the worn airfoils. In an iterative process the flow behavior of these redesigned profiles is checked by numerical flow simulations and the shape of the profiles is improved. The following machining of the new defined leading edge contours is achieved on a grinding station handled by an appropriately programmed robot. Within this Advanced Re-contouring Process (ARP) the worn blades are precision-measured and then provided with an aerodynamically optimized leading edge profile numerically newly developed under computer control. The application of this process enhances the performance and lowers the fuel consumption while prolonging the blades’ service life by 25%. The performance achievable with ARP has been confirmed both through a long term analysis and by a back-to-back comparison test on the engine test stand. For this purpose the stages 3 through 14 of a CF6-50 high pressure compressor were on the one hand fitted with conventionally overhauled blades and on the other with ARP-optimized blades of the same basic geometry. By installing the optimized blades the EGT margin could be increased by 3° to 4° C. This results in an prolongation of the on-wing time by more than 1000 hours.

Aerodynamic Retrofit Design for a High Pressure Compressor Using a High Hub/Tip Ratio Mixed-Flow Compressor

2020 ◽  
Author(s):  
Hang Xiang ◽  
Jiang Chen ◽  
Jinxin Cheng ◽  
Han Niu ◽  
Yi Liu ◽  
...  
Keyword(s):  
High Pressure ◽  
Load Capacity ◽  
Axial Compressor ◽  
Design Flow ◽  
Maximum Efficiency ◽  
Mixed Flow ◽  
High Pressure Compressor ◽  
Flow Compressor ◽  
Retrofit Design ◽  
Pressure Compressor

Abstract A high-load mixed-flow compressor with an extremely high inlet hub/tip ratio (0.889) is designed and analyzed for replacing the rear stages of a multistage high pressure axial compressor. The effects of blade number, splitter blades and dimensionless geometric parameters on the impeller performance are investigated by an improved loss model. A full-surface parameterization control method is adopted for blade optimizations of the mixed-flow impeller and the tandem stator. As a retrofit design of the multistage axial compressor, an unconventional type of axial-co-mixed-flow combined compressor scheme is proposed and discussed. Further, in order to minimize the axial dimension and maximize the load, this paper also proposed preliminary designs of the twin-stage mixed-flow compressor and the twin-stage counter-rotating mixed-flow compressor respectively equipped with the high hub/tip ratio mixed-flow compressor. The results indicate that the mixed-flow impeller configuration with 42 principal blades and splitter blades with a fifth of principal blade length has the maximum efficiency at design flow rate. Blade height/pitch ratio is a considerable parameter which demonstrates the interaction among hub/tip ratio, aspect ratio and solidity especially for high hub/tip ratio cascade designs. The mixed-flow compressor can greatly improve the load capacity of the high pressure compressor with slight impact on efficiency and surge margin. At low rotate speed, the mixed-flow impeller can maintain relatively high efficiency level and even carry a higher proportion of the load, while the tandem stator limits the overall efficiency improvement. Besides, structures with no return channel of the three unconventional combined compressors are beneficial to the reduction of dimension and cost, which shows the potential application prospects of high hub/tip ratio mixed-flow compressors.

Numerical Simulation of Particulates in Multi-Stage Axial Compressors

2016 ◽  
Author(s):  
Swati Saxena ◽  
Giridhar Jothiprasad ◽  
Corey Bourassa ◽  
Byron Pritchard
Keyword(s):  
High Pressure ◽  
Axial Compressor ◽  
Airborne Particulate Matter ◽  
Spherical Particles ◽  
Qualitative Comparison ◽  
Multi Stage ◽  
Sand Erosion ◽  
High Pressure Compressor ◽  
Multi Phase ◽  
Pressure Compressor

Aircraft engines ingest airborne particulate matter, such as sand, dirt, and volcanic ash, into their core. The ingested particulate is transported by the secondary flow circuits via compressor bleeds to the high pressure turbine and may deposit resulting in turbine fouling and loss of cooling effectiveness. Prior publications focused on particulate deposition and sand erosion patterns in a single stage of a compressor or turbine. The current work addresses the migration of ingested particulate through the high pressure compressor and bleed systems. This paper describes a 3D CFD methodology for tracking particles along a multi-stage axial compressor and presents particulate ingestion analysis for a high pressure compressor section. The commercial CFD multi-phase solver ANSYS CFX R has been used for flow and particulate simulations. Particle diameters of 20, 40, and 60 microns are analyzed. Particle trajectories and radial particulate profiles are compared for these particle diameters. The analysis demonstrates how the compressor centrifuges the particles radially towards the compressor case as they travel through the compressor; the larger diameter particles being more significantly affected. Non-spherical particles experience more drag as compared to spherical particles and hence a qualitative comparison between spherical and non-spherical particles is shown.

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.

scholarly journals Surface roughness of real operationally used compressor blade and blisk

2019 ◽  
Vol 233 (14) ◽  
pp. 5321-5330 ◽  
Author(s):  
Philipp Gilge ◽  
Andreas Kellersmann ◽  
Jens Friedrichs ◽  
Jörg R Seume
Keyword(s):  
Surface Roughness ◽  
High Pressure ◽  
Laser Scanning ◽  
Flow Direction ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Aircraft Engine ◽  
Rotor Blades ◽  
High Pressure Compressor ◽  
Pressure Compressor

Deterioration of axial compressors is in general a major concern in aircraft engine maintenance. Among other effects, roughness in high-pressure compressor reduces the pressure rise and thus efficiency, thereby increasing the specific fuel consumption of an engine. Therefore, it is important to improve the understanding of roughness on compressor blading and their impact on compressor performance. To investigate the surface roughness of rotor blades of a compressors, different stages of an axial high-pressure compressor and a first-stage blisk (BLade–Integrated–dISK) of a regional aircraft engine is measured by a three-dimensional laser scanning microscope. Fundamental types of roughness structures can be identified: impacts in different sizes, depositions as isotropically distributed single elements with steep flanks and anisotropic roughness structures direct approximately normal to the flow direction. To characterise and quantify the roughness structures in more detail, roughness parameters were determined from the measured surfaces. The quantification showed that the roughness height varies through the compressor depending on the stage, position and the blade side. Overall complex roughness structures of different shape, height and size are detected regardless of the type of the blades.

Investigation on tip clearance control for the high-pressure rotor of an uncooled vaneless counter-rotating turbine

2020 ◽  
pp. 095441002095050
Author(s):  
Wei Zhao ◽  
Xiuming Sui ◽  
Kai Zhang ◽  
Zeming Wei ◽  
Qingjun Zhao
Keyword(s):  
High Pressure ◽  
Leading Edge ◽  
Back Pressure ◽  
Operating Conditions ◽  
Tip Clearance ◽  
Trailing Edge ◽  
Blade Tip ◽  
Cooling Air ◽  
Clearance Control ◽  
Blade Tip Clearance

In order to develop a tip clearance control system for an uncooled vaneless counter-rotating turbine, tip clearance variation of its high pressure rotor blade at off-design conditions is analyzed. Aero-thermal interaction simulation is performed to predict the temperature and deformation of the solid blade. At operating conditions with rotating speeds greater than 60% design value and expansion ratios greater than 85% design value, the blade tip clearance height at leading edge remains unchanged when the expansion ratio decreases, meanwhile that at trailing edge decreased obviously. However, the tip clearance height variations at the leading edge and trailing edge are almost the same in a conventional subsonic turbine at such conditions. The cause is that the flow in the high-pressure rotor is choked at these conditions. The choked flow results in that the fluid and solid blade temperatures upstream of the throat are not affected by the back pressure and only those downstream of the throat increases with the back pressure. Consequently, the blade height at leading edge keeps constant, and that at trailing edge varies because of thermal expansion. To avoid the rubbing of the blade and case, the blade height at trailing edge is diminished by 30%. As a result, the blade tip clearance height at low speed operating conditions increases in axial direction. Such a design leads to a stronger tip leakage flow. More flow losses might be generated. Therefore, a casing cooling method is proposed to control the blade tip clearance height at leading edge and trailing edge respectively. The deformations of the casing with different mass flow rate of cooling air at design and off-design conditions are calculated. It shows that the blade tip clearance heights at leading edge and at trailing edge of the rotor can be well controlled with appropriate amount of cooling air.

scholarly journals DEVELOPMENT OF THE NUMERICAL METHODOLOGY FOR THE ESTIMATION OF RESONANT PROCESSES AT GASTURBINE ENGINE FLOW DUCT

2020 ◽  
pp. 12-19
Author(s):  
Nikolay Shuvaev ◽  
◽  
Aleksandr Siner ◽  
Ruslan Kolegov ◽  
◽  
...  
Keyword(s):  
High Pressure ◽  
Unsteady Flow ◽  
Aircraft Engine ◽  
Operating Conditions ◽  
Design Stage ◽  
High Pressure Compressor ◽  
Wide Range ◽  
Resonant Processes ◽  
Flow Duct ◽  
Pressure Compressor

Ensuring safety of flights is the most important task that is being solved in the process of designing an aircraft engine and aircraft. The most complex are the physical processes occurring inside the aircraft engine, especially in its gas generator: combustion chamber, high-pressure compressor and high-pressure turbine. The unsteady flow of gas in the flow duct of the aircraft engine is very complex, it is difficult to model, because the flow is characterized by a wide range of time and space scales. Unsteady flow in a high-pressure compressor can cause surge and breakdown of the compressor and the entire engine as a whole. Along with the detachment flows causing the surge, in the flow duct there can be resonant phenomena associated with the propagation of powerful sound waves along the flow duct of the engine, which, when a direct and reflected wave is imposed, create a very powerful standing wave that affects the structure. With a certain combination of conditions, the coincidence of the natural frequencies of the oscillations of the air volume and the solid body, such resonant processes in the flow duct of the gas turbine engine can lead to serious breakdowns, such as breakage of rotor blades and guide vanes, destruction of the aeroengine framework and other. The main difficulty is that it is problematic to identify such processes at the design and debugging stage, since there are no suitable mathematical models, and for experimental verification it is required to withstand the specific operating conditions of the node that are not known in advance. This work is devoted to the creation of a calculation technique that will allow in the future to diagnose resonance phenomena at the design stage and thereby significantly reduce the costs for the design, testing and manufacture of an aircraft engine. The proposed technique is based on the nonstationary Navier-Stokes equations for a compressible gas.

Effects of tip injection on the performance of a multi-stage high-pressure compressor

2011 ◽  
Vol 2 (1-4) ◽  
pp. 99-110 ◽  
Author(s):  
M. Kern ◽  
W. Horn ◽  
S.-J. Hiller ◽  
S. Staudacher
Keyword(s):  
High Pressure ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Tip Injection ◽  
Pressure Compressor
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