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.

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.

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.

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.

scholarly journals A Navier-Stokes Analysis of the Effect of Tip Clearance on Compressor Stall Margin

1995 ◽  
Author(s):  
Robert P. Dring ◽  
William D. Sprout ◽  
Harris D. Weingold
Keyword(s):  
Three Dimensional ◽  
Axial Compressor ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Stall Margin ◽  
Multi Stage ◽  
High Pressure Compressor ◽  
Practical Tool ◽  
The Impact ◽  
Pressure Compressor

A three-dimensional Navier-Stokes calculation was used to analyze the impact of rotor tip clearance on the stall margin of a multi-stage axial compressor. This paper presents a summary of: (1) a study of the sensitivity of the results to grid refinement, (2) an assessment of the calculation’s ability to predict stall margin when the stalling row was the first rotor in a multi-stage rig environment, (3) an analysis of the impact of including the effects of the downstream stator through body force effects on the upstream rotor, and (4) the ability of the calculation to predict the impact of tip clearance on stall margin through a calculation of the rear seven airfoil rows of an eleven stage high pressure compressor rig. The result of these studies was that a practical tool is available which can predict stall margin, and the impact of tip clearance, with reasonable accuracy.

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.

Full High Pressure Compressor Investigations to Determine Aerodynamic Changes due to Deterioration

2018 ◽  
Author(s):  
Gerald Reitz ◽  
Andreas Kellersmann ◽  
Jens Friedrichs
Keyword(s):  
High Pressure ◽  
Reverse Engineering ◽  
Test Cell ◽  
Engineering Process ◽  
Cfd Model ◽  
Jet Engine ◽  
High Pressure Compressor ◽  
Compressor Map ◽  
Cell Data ◽  
Pressure Compressor

The Institute of Jet Propulsion and Turbomachinery of the TU Braunschweig owns a jet engine of the type V2500-A1 from the International Aero Engines AG. To conduct research on the jet engine and its components, computer models are necessary. In this paper, the reverse engineering process of the high pressure compressor (HPC) regarding its aerodynamics is presented. Thereby, the reverse engineering process starts from digitizing newly manufactured airfoils, followed by FEM-calculations to enforce the operating forces on the geometries. A computational fluid dynamics (CFD) model using these geometries is set up, considering all relevant geometric and aerodynamic features such as bleed ports and the variable stator vane (VSV) system. Using this CFD-model, the compressor map is calculated and afterwards validated by available manufacturing data [18] and by the institute’s jet engine’s test cell data. Because this jet engine is a highly operated and deteriorated one, a map scaling is necessary before comparing the CFD-model with the test cell data. Nevertheless, an adequate agreement of the operating behavior between scaled compressor map and test cell data is shown. To estimate the deterioration level of the jet engine’s compressor and to evaluate the used scaling factors, the tip gaps inside the CFD-model were doubled and the compressor behavior was simulated. The observed effect of reduced compressor capacity and efficiency is in accordance with literature but is not able to explain the amount of the scaling factors completely.

scholarly journals The comparative analysis of the selected construction types of axial compressor stage including the modal analysis

2017 ◽  
Vol 4 (17) ◽  
pp. 91-97
Author(s):  
Adam KOZAKIEWICZ ◽  
Olga GRZEJSZCZAK ◽  
Tomasz LACKI
Keyword(s):  
High Pressure ◽  
Comparative Analysis ◽  
Vibration Frequency ◽  
Vibration Spectrum ◽  
Axial Compressor ◽  
Point Of View ◽  
Compressor Stage ◽  
Jet Engine ◽  
High Pressure Compressor ◽  
Pressure Compressor

The article concerns the issues of the scope of optimization of the gas turbine jet engine. These issues include limiting the weight and number of engine parts. One way to reduce the weight and number of components, including the compressor assembly, is to use the BLISK's replacement construction. The replacement construction should meet the strength requirement and the vibration spectrum as well. The paper presents a comparative analysis of the influence of rotational speed on the characters and the vibration frequency of the single rotor stage of the high pressure compressor. The analysis was carried out for two different design solutions of the blade-disk connection: the classical and integral. The comparative analysis focused on three important from the point of view of operation, the engine operating ranges: work on the ground (idle) and work during take-off and climb the aircraft.

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.

scholarly journals Development and Application of a Multistage Navier-Stokes Flow Solver: Part II — Application to a High Pressure Compressor Design

1995 ◽  
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% higher efficiency than a baseline machine, but it had 14% 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% reduction in the compressor development time and cost.

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