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.

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 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.

Measurement and Computation of Heat Transfer in High-Pressure Compressor Drum Geometries With Axial Throughflow

1997 ◽  
Vol 119 (1) ◽  
pp. 51-60 ◽  
Author(s):  
C. A. Long ◽  
A. P. Morse ◽  
P. G. Tucker
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Multigrid Algorithm ◽  
Numerical Predictions ◽  
High Pressure Compressor ◽  
Dependent Form ◽  
Pressure Compressor

This paper makes comparisons between CFD computations and experimental measurements of heat transfer for the axial throughflow of cooling air in a high-pressure compressor spool rig and a plane cavity rig. The heat transfer measurements are produced using fluxmeters and by the conduction solution method from surface temperature measurements. Numerical predictions are made by solving the Navier–Stokes equations in a full three-dimensional, time-dependent form using the finite-volume method. Convergence is accelerated using a multigrid algorithm and turbulence modeled using a simple mixing length formulation. Notwithstanding systematic differences between the measurements and the computations, the level of agreement can be regarded as promising in view of the acknowledged uncertainties in the experimental data, the limitations of the turbulence model and, perhaps more importantly, the modest grid densities used for the computations.

Vorticity Dynamics Based Flow Diagnosis for a 1.5-Stage High Pressure Compressor With an Optimized Transonic Rotor

2016 ◽  
Author(s):  
Huanlong Chen ◽  
Mark G. Turner ◽  
Kiran Siddappaji ◽  
Syed Moez Hussain Mahmood
Keyword(s):  
High Pressure ◽  
Boundary Layer Separation ◽  
Navier Stokes ◽  
Stall Margin ◽  
Transonic Compressor ◽  
High Pressure Compressor ◽  
Blade Row ◽  
Vorticity Dynamics ◽  
Improved Performance ◽  
Pressure Compressor

This paper presents an optimized rotor as part of a 3-blade row optimization (IGV-rotor-stator) of a high-pressure compressor. It is based on modifying blade angles and advanced control of curvature of the airfoil camber line. The effects of these advanced blade techniques on the performance of the transonic 1.5-stage compressor were calculated using a 3D Navier-Stokes solver combined with a vortex/vorticity dynamics diagnosis method. The optimized rotor produces a 3-blade row efficiency improvement over the baseline of 1.45% while also improving stall margin. The throttling range of the compressor is expanded largely because the shock in the rotor tip area is further downstream than that in the baseline case at the operating point. Additionally, optimizing the 3-blade row block while only adjusting the rotor geometry ensures good matching of flow angles allowing the compressor to have more range. The flow diagnostics of the rotor blade based on vortex/vorticity dynamics indicate that the boundary-layer separation behind the shock are verified by on-wall signatures of vorticity and skin-friction vector lines. In addition, azimuthal vorticity and boundary vorticity flux (BVF) are shown to be two vital flow parameters of compressor aerodynamic performance that directly relate to the improved performance of the optimized transonic compressor blade.

Measurement and Computation of Heat Transfer in High Pressure Compressor Drum Geometries With Axial Throughflow

1995 ◽  
Author(s):  
Christopher A. Long ◽  
Alan P. Morse ◽  
Paul G. Tucker
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Stokes Equations ◽  
Three Dimensional ◽  
Navier Stokes ◽  
Multigrid Algorithm ◽  
Numerical Predictions ◽  
High Pressure Compressor ◽  
Dependent Form ◽  
Pressure Compressor

This paper makes comparisons between CFD computations and experimental measurements of heat transfer for the axial throughflow of cooling air in a high-pressure compressor spool rig and a plane cavity rig. The heat transfer measurements are produced using fluxmeters and by the conduction solution method from surface temperature measurements. Numerical predictions are made by solving the Navier-Stokes equations in a full three-dimensional, time-dependent form using the finite-volume method. Convergence is accelerated using a multigrid algorithm and turbulence modelled using a simple mixing length formulation. Notwithstanding systematic differences between the measurements and the computations, the level of agreement can be regarded as promising in view of the acknowledged uncertainties in the experimental data, the limitations of the turbulence model and, perhaps more importantly, the modest grid densities used for the computations.

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 Effects of Stator Pressure Field on Upstream Rotor Performance

1999 ◽  
Author(s):  
M. B. Graf ◽  
E. M. Greitzer ◽  
F. E. Marble ◽  
O. P. Sharma
Keyword(s):  
Steady Flow ◽  
Pressure Field ◽  
Three Dimensional ◽  
Pressure Rise ◽  
Tip Clearance ◽  
Navier Stokes ◽  
Flow Models ◽  
High Pressure Compressor ◽  
Rotor Stator Interaction ◽  
Rotor Loss

Effects of stator pressure field on upstream rotor performance in a high pressure compressor stage have been assessed using three-dimensional steady and time-accurate Reynolds-averaged Navier-Stokes computations. Emphasis was placed on: (1) determining the dominant features of the flow arising from interaction of the rotor with the stator pressure field, and (2) quantifying the overall effects on time averaged loss, blockage, and pressure rise. The time averaged results showed a 20 to 40% increase in overall rotor loss and a 10 to 50% decrease in tip clearance loss compared to an isolated rotor. The differences were dependent on the operating point and increased as the stage pressure rise, and amplitude of the unsteady back pressure variations, was increased. Motions of the tip leakage vortex on the order of the blade pitch were observed at the rotor exit in all the unsteady flow simulations; these were associated with enhanced mixing in the region. The period of the motion scaled with rotor flow-through time rather than stator passing. Three steady flow approximations for the rotor-stator interaction were assessed with reference to the unsteady computations: an axisymmetric representation of the stator pressure field, an inter-blade row averaging plane method, and a technique incorporating deterministic stresses and bodyforces associated with stator flow field. Differences between steady and unsteady predictions of overall rotor loss, tip region loss, and endwall blockage ranged from 5 to 50% of the time average, but the steady flow models gave overall rotor pressure rise and flow capacity within 5% of the time averaged values.

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 Research of the possibility of improving the traction and economic characteristics of a supersonic passenger aircraft engine through minimal modifications to the high-pressure compressor

2021 ◽  
Vol 2094 (4) ◽  
pp. 042055
Author(s):  
D Yu Strelets ◽  
S A Serebryansky ◽  
M V Shkurin
Keyword(s):  
High Pressure ◽  
Numerical Model ◽  
Thermodynamic Model ◽  
Gas Dynamics ◽  
High Speed ◽  
Three Dimensional ◽  
Aircraft Engine ◽  
High Pressure Compressor ◽  
Passenger Aircraft ◽  
Pressure Compressor

Abstract In this paper, the possibilities of improving the traction and economic characteristics of a by-pass turbojet engine of a high-speed passenger aircraft due to minimal modifications of the high-pressure compressor. A thermodynamic model of the investigated engine of a new design in a three-dimensional layout was formed using an automated multicriteria optimization process. A computational assessment of the change in the characteristics of compressor modifications is carried out based on a numerical model of gas dynamics.

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.

Sign in / Sign up

Export Citation Format

Share Document