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.

scholarly journals Recontouring of Jet Engine Compressor Blades by Flow Simulation

2001 ◽  
Vol 7 (5) ◽  
pp. 365-374
Author(s):  
Herwart Hönen ◽  
Matthias Panten
Keyword(s):  
Flow Behavior ◽  
Engine Performance ◽  
Flow Simulation ◽  
Leading Edge ◽  
Tip Clearance ◽  
Compressor Blades ◽  
Common Procedure ◽  
Stall Margin ◽  
Leading Edges ◽  
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 by application of a new developed method for the profile definition. 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.

Performance Analysis of Deteriorated High Pressure Compressor Blades

2014 ◽  
Author(s):  
Gerald Reitz ◽  
Jens Friedrichs ◽  
Jonas Marx ◽  
Jörn Städing
Keyword(s):  
High Pressure ◽  
Leading Edge ◽  
Tip Clearance ◽  
Major Influence ◽  
Additional Parameter ◽  
Gas Temperature ◽  
Jet Engine ◽  
High Pressure Compressor ◽  
Stagger Angle ◽  
Pressure Compressor

During the operation of a jet engine, deterioration will constantly reduce its performance. This results in an increase in specific fuel consumption (SFC) and exhaust gas temperature (EGT); the main characteristics to describe the efficiency of a jet engine. Thereby, the high pressure compressor (HPC) is particularly affected by deterioration. Multiple effects take place and decrease the efficiency of the HPC. Erosion is one of the main effects and leads to thinner or thicker leading- and trailing edges, thinner airfoils, a reduction of chord length and an increase in tip clearance. In addition, erosion and fouling may also lead to increased surface roughness on airfoils and endwalls. An additional parameter which is also dependent on the on-wing time are changes in the stagger angle of the different blade heights. The objective is to estimate the quantitative effect of the different wear mechanisms on the stage parameters, like throttle line and efficiency. Therefore, a geometry setup process is implemented to create HPC blade models with independent values of erosion. With these blades, CFD calculations based on realistic boundary conditions were carried out with the CFD solver ANSYS CFX. It could be proven that the deterioration of leading edge thickness has the major influence on stage performance, followed by the max. profile thickness and the stagger angle. The operational blade deterioration of leading edge thickness leads to an efficiency range of about 0.173 %. Moreover, the deterioration of stagger angle leads to an offset of the throttle lines towards higher or smaller loadings, depending on the direction of change.

scholarly journals Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions

2014 ◽  
Author(s):  
Alain Batailly ◽  
Mathias Legrand ◽  
Antoine Millecamps ◽  
Sèbastien Cochon ◽  
François Garcin
Keyword(s):  
High Pressure ◽  
Forced Response ◽  
Compressor Blade ◽  
Design Stage ◽  
Blade Design ◽  
Early Integration ◽  
Nonlinear Structural ◽  
High Pressure Compressor ◽  
Structural Interactions ◽  
Pressure Compressor

Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural blade/casing contacts at the design stage of a high-pressure compressor blade. The blade of interest underwent structural divergence after blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the blade stacking law while maintaining aerodynamic performance of the original blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

Multistage Compressor and Turbine Modeling for the Prediction of the Maximum Turbine Speed Resulting From Shaft Breakage

2009 ◽  
Author(s):  
M. Haake ◽  
R. Fiola ◽  
S. Staudacher
Keyword(s):  
High Pressure ◽  
Tip Clearance ◽  
Maximum Speed ◽  
High Pressure Turbine ◽  
High Pressure Compressor ◽  
Failure Event ◽  
Turbine Speed ◽  
Shaft Failure ◽  
Pressure Compressor ◽  
Vane Tip

A mathematical model for the prediction of the maximum speed of a high pressure turbine following a shaft failure event was developed. The model predicts the high pressure compressor and ducting system pre- and post-stall behavior like rotating stall and surge after the shaft breakage. The corresponding time-dependent high pressure turbine inlet conditions are used to calculate the turbine maximum speed, taking into account friction and blade&vane tip clearance variations as a result of the rearward movement of the turbine and destruction of the turbine blading. The compressor and ducting system is modeled by a 1-dimensional, stage-by-stage approach. The approach uses a finite-difference numerical technique to solve the nonlinear system of equations for continuity, momentum and energy including source terms for the compressible flow through inlet ducting, compressor and combustor. The compressor blade forces and shaft work are provided by a set of quasi steady state stage characteristics being valid for pre-stall and post-stall operations. The maximum turbine speed is calculated from a thermodynamic turbine stand-alone model, derived from a performance synthesis program. Friction and blade&vane tip clearance variations are determined iteratively from graphical data depending on the axial rearward movement of the turbine. The compressor and ducting system model was validated in pre-stall and post-stall operation mode with measured high pressure compressor data of a modern 2-shaft engine. The turbine model was validated with measured intermediate pressure shaft failure data of a 3-shaft engine. The shaft failure model was applied on a modern 2-shaft engine. The model was used to carry out a sensitivity study to demonstrate the impact of control system reactions on the resulting maximum high pressure turbine speed following a shaft failure event.

Failure Analysis of High-Pressure Compressor Blade in an Aero Gas Turbine Engine

2018 ◽  
Vol 18 (3) ◽  
pp. 465-470 ◽  
Author(s):  
R. K. Mishra ◽  
Vaishakhi Nandi ◽  
R. Raghavendra Bhat
Keyword(s):  
High Pressure ◽  
Failure Analysis ◽  
Gas Turbine ◽  
Gas Turbine Engine ◽  
Compressor Blade ◽  
Turbine Engine ◽  
High Pressure Compressor ◽  
Pressure Compressor

scholarly journals Redesign of a High-Pressure Compressor Blade Accounting for Nonlinear Structural Interactions

2014 ◽  
Vol 137 (2) ◽  
Author(s):  
Alain Batailly ◽  
Mathias Legrand ◽  
Antoine Millecamps ◽  
Sébastien Cochon ◽  
François Garcin
Keyword(s):  
High Pressure ◽  
Forced Response ◽  
Compressor Blade ◽  
Design Stage ◽  
Blade Design ◽  
Early Integration ◽  
Nonlinear Structural ◽  
High Pressure Compressor ◽  
Structural Interactions ◽  
Pressure Compressor

Recent numerical developments dedicated to the simulation of rotor/stator interaction involving direct structural contacts have been integrated within the Snecma industrial environment. This paper presents the first attempt to benefit from these developments and account for structural blade/casing contacts at the design stage of a high-pressure compressor blade. The blade of interest underwent structural divergence after blade/abradable coating contact occurrences on a rig test. The design improvements were carried out in several steps with significant modifications of the blade stacking law while maintaining aerodynamic performance of the original blade design. After a brief presentation of the proposed design strategy, basic concepts associated with the design variations are recalled. The iterated profiles are then numerically investigated and compared with respect to key structural criteria such as: (1) their mass, (2) the residual stresses stemming from centrifugal stiffening, (3) the vibratory level under aerodynamic forced response, and (4) the vibratory levels when unilateral contact occurs. Significant improvements of the final blade design are found: the need for an early integration of nonlinear structural interactions criteria in the design stage of modern aircraft engines components is highlighted.

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.

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.

scholarly journals Influence of combined compressor and turbine deterioration on the overall performance of a jet engine using RANS simulation and Pseudo Bond Graph approach

2020 ◽  
Vol 4 ◽  
pp. 296-308
Author(s):  
Jan Goeing ◽  
Hendrik Seehausen ◽  
Vladislav Pak ◽  
Sebastian Lueck ◽  
Joerg R. Seume ◽  
...  
Keyword(s):  
High Pressure ◽  
Differential Equations ◽  
Tip Clearance ◽  
System Stability ◽  
Simplified Model ◽  
Jet Engine ◽  
High Pressure Compressor ◽  
Overall Performance ◽  
The Impact ◽  
Pressure Compressor

In this study, numerical models are used to analyse the influence of isolated component deterioration as well as the combination of miscellaneous deteriorated components on the transient performance of a high-bypass jet engine. For this purpose, the aerodynamic impact of major degradation effects in a high-pressure compressor (HPC) and turbine (HPT) is modelled and simulated by using 3D CFD (Computational Fluid Dynamics). The impact on overall jet engine performance is then modelled using an 1D Reduced Order Model (ROM). Initially, the HPC performance is investigated with a typical level of roughness on vanes and blades and the HPT performance with an increasing tip clearance. Subsequently, the overall performance of the jet engines with the isolated and combined deteriorated domains is computed by the in-house 1D performance tool ASTOR (AircraftEngine Simulation for Transient Operation Research). Degradations have a significant influence on the system stability and transient effects. In ASTOR, a system of differential equations including the equations of motion and further ordinary differential equations is solved. Compared to common ROMs, this enables a higher degree of accuracy. The results of temperature downstream of the high-pressure compressor and low-pressure turbine as well as the specific fuel composition and the HP rotational speed are used to estimate the degree and type of engine deterioration. However, the consideration of the system stability is necessary to analyse the characterisation in more detail. Finally, a simplified model which merges two engines with individual deteriorated domains into one combined deteriorated engine, is proposed. The simplified model predicts the performance of an engine which has been simulated with combined deteriorated components.

1995 ASME Gas Turbine Award Paper: Development and Application of a Multistage Navier–Stokes Flow Solver: Part II—Application to a High-Pressure Compressor Design

1998 ◽  
Vol 120 (2) ◽  
pp. 215-223 ◽  
Author(s):  
C. R. LeJambre ◽  
R. M. Zacharias ◽  
B. P. Biederman ◽  
A. J. Gleixner ◽  
C. J. Yetka
Keyword(s):  
High Pressure ◽  
Three Dimensional ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Flow Solver ◽  
Stall Margin ◽  
High Pressure Compressor ◽  
Compressor Design ◽  
Flow Through ◽  
Pressure Compressor

Two versions of a three-dimensional multistage Navier–Stokes code were used to optimize the design of an eleven-stage high-pressure compressor. The first version of the code utilized a “mixing plane” approach to compute the flow through multistage machines. The effects due to tip clearances and flowpath cavities were not modeled. This code was used to minimize the regions of separation on airfoil and endwall surfaces for the compressor. The resulting compressor contained bowed stators and rotor airfoils with contoured endwalls. Experimental data acquired for the HPC showed that it achieved 2 percent higher efficiency than a baseline machine, but it had 14 percent lower stall margin. Increased stall margin of the HPC was achieved by modifying the stator airfoils without compromising the gain in efficiency as demonstrated in subsequent rig and engine tests. The modifications to the stators were defined by using the second version of the multistage Navier–Stokes code, which models the effects of tip clearance and endwall flowpath cavities, as well as the effects of adjacent airfoil rows through the use of “bodyforces” and “deterministic stresses.” The application of the Navier–Stokes code was assessed to yield up to 50 percent reduction in the compressor development time and cost.

Sign in / Sign up

Export Citation Format

Share Document