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.

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.

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.

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.

Review on the aerodynamics of intermediate compressor duct

2020 ◽  
Vol 14 (4) ◽  
pp. 7446-7468
Author(s):  
Manish Sharma ◽  
Beena D. Baloni
Keyword(s):  
High Pressure ◽  
Pressure Loss ◽  
Flow Analysis ◽  
Outer Wall ◽  
Optimization Techniques ◽  
Optimum Pressure ◽  
Research Areas ◽  
Static Pressure Distribution ◽  
High Pressure Compressor ◽  
Pressure Compressor

In a turbofan engine, the air is brought from the low to the high-pressure compressor through an intermediate compressor duct. Weight and design space limitations impel to its design as an S-shaped. Despite it, the intermediate duct has to guide the flow carefully to the high-pressure compressor without disturbances and flow separations hence, flow analysis within the duct has been attractive to the researchers ever since its inception. Consequently, a number of researchers and experimentalists from the aerospace industry could not keep themselves away from this research. Further demand for increasing by-pass ratio will change the shape and weight of the duct that uplift encourages them to continue research in this field. Innumerable studies related to S-shaped duct have proven that its performance depends on many factors like curvature, upstream compressor’s vortices, swirl, insertion of struts, geometrical aspects, Mach number and many more. The application of flow control devices, wall shape optimization techniques, and integrated concepts lead a better system performance and shorten the duct length.  This review paper is an endeavor to encapsulate all the above aspects and finally, it can be concluded that the intermediate duct is a key component to keep the overall weight and specific fuel consumption low. The shape and curvature of the duct significantly affect the pressure distortion. The wall static pressure distribution along the inner wall significantly higher than that of the outer wall. Duct pressure loss enhances with the aggressive design of duct, incursion of struts, thick inlet boundary layer and higher swirl at the inlet. Thus, one should focus on research areas for better aerodynamic effects of the above parameters which give duct design with optimum pressure loss and non-uniformity within the duct.

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.

A Machine Learning Based Approach of Performance Estimation for High-Pressure Compressor Airfoils

2018 ◽  
Author(s):  
Jonas Marx ◽  
Stefan Gantner ◽  
Jörn Städing ◽  
Jens Friedrichs
Keyword(s):  
Machine Learning ◽  
High Pressure ◽  
Machine Learning Algorithms ◽  
Performance Estimation ◽  
Predictive Maintenance ◽  
Support Vector ◽  
K Nearest Neighbor ◽  
Operating Characteristics ◽  
High Pressure Compressor ◽  
Pressure Compressor

In recent years, the demands of Maintenance, Repair and Overhaul (MRO) customers to provide resource-efficient after market services have grown increasingly. One way to meet these requirements is by making use of predictive maintenance methods. These are ideas that involve the derivation of workscoping guidance by assessing and processing previously unused or undocumented service data. In this context a novel approach on predictive maintenance is presented in form of a performance-based classification method for high pressure compressor (HPC) airfoils. The procedure features machine learning algorithms that establish a relation between the airfoil geometry and the associated aerodynamic behavior and is hereby able to divide individual operating characteristics into a finite number of distinct aero-classes. By this means the introduced method not only provides a fast and simple way to assess piece part performance through geometrical data, but also facilitates the consideration of stage matching (axial as well as circumferential) in a simplified manner. It thus serves as prerequisite for an improved customary HPC performance workscope as well as for an automated optimization process for compressor buildup with used or repaired material that would be applicable in an MRO environment. The methods of machine learning that are used in the present work enable the formation of distinct groups of similar aero-performance by unsupervised (step 1) and supervised learning (step 2). The application of the overall classification procedure is shown exemplary on an artificially generated dataset based on real characteristics of a front and a rear rotor of a 10-stage axial compressor that contains both geometry as well as aerodynamic information. In step 1 of the investigation only the aerodynamic quantities in terms of multivariate functional data are used in order to benchmark different clustering algorithms and generate a foundation for a geometry-based aero-classification. Corresponding classifiers are created in step 2 by means of both, the k Nearest Neighbor and the linear Support Vector Machine algorithms. The methods’ fidelities are brought to the test with the attempt to recover the aero-based similarity classes solely by using normalized and reduced geometry data. This results in high classification probabilities of up to 96 % which is proven by using stratified k-fold cross-validation.

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