Y. S. Li

Adjoint Aerodynamic Design Optimization for Blades in Multistage Turbomachines—Part II: Validation and Application

This is the second part of a two-part paper. First, the design-optimization system basedon the adjoint gradient solution approach as described in Part I is introduced. Severaltest cases are studied for further validation and demonstration of the methodology andimplementation. The base-line adjoint method as applied to realistic 3D configurations isdemonstrated in the redesign of the NASA rotor 67 at a near-choke condition, leading toa 1.77% efficiency gain. The proposed adjoint mixing plane is applied to the redesign ofa transonic compressor stage (DLR compressor stage) and an IGV-rotor-stator configu-ration of a Siemens industrial compressor at a single-operating point, both producingmeasurably positive efficiency gains. An examination on the choice of the operating massflow condition as the basis for the performance optimization, however, highlights thelimitation of the single-point approach for practical applications. For the three-rowcompressor configuration, a near peak-efficiency point based redesign leads to a mea-surable reduction in the choke mass flow, while a near-choke point based redesign leadsto a significant performance drop in other flow conditions. Subsequently, a parallelmultipoint approach is implemented. The results show that a two-point design optimiza-tion can produce a consistently better performance over a whole range of mass flowconditions compared with the original design. In the final case, the effectiveness of thepresent method and system is demonstrated by a redesign applied to a seven-row indus-trial compressor at the design point, leading to a remarkable 2.4% efficiency gain.

Blade Forced Response Prediction for Industrial Gas Turbines

A decoupled aeromechanical system based on an advanced frequency-domain computational fluid dynamics finite element to CFD solver with fully nonlinear capability is presented that allows resonant vibration predictions to be routinely performed during the design process. A new energy method is presented that solves the blade response without the knowledge of original modeshape scales. A robust finite element CFD mesh interface has been developed for industrial use that can accurately deal with differences in mesh geometry, low mesh density, and high modeshape gradients. The capability of the baseline CFD solver for blade row interaction flow prediction is further validated against the von Karman Institute transonic turbine stage. A forced response analysis is carried out on the NASA Rotor 67 transonic fan for demonstration purposes. The system is evaluated for a challenging industrial study of the ALSTOM three-stage transonic test compressor, where the forced response predictions of three crossing points on the Campbell diagram compare well with strain gauge test data. An investigation into aerodynamic damping of the first 10 modes shows a high dependency on modeshape.

Blading Aerodynamics Design Optimization With Mechanical and Aeromechanical Constraints

A blading design optimization system has been developed using an aeromechanical approach and harmonic perturbation method. The developed system has the capability to optimize aero-thermal performance with constraints of mechanical and aeromechanical integrity at the same time. ‘Aerodynamic mode shape’ is introduced to describe geometry deformation which can effectively reduce the number of design parameters while preserving surface smoothness. Compared to the existing design optimization practices, the present system is simpler, more accurate and effective. A redesign practice of the NASA rotor-67 at the peak efficiency point shows that the aero thermal efficiency can be improved by 0.4%, whilst the maximum static stress has been increased by 33%. Aeromechanical analysis of the optimized blade shows that the aerodynamic damping of the least stable first flap mode is still well above the critical value though the natural frequencies of the first 5 modes have been reduced by 1~4%. The present finding highlights the need for more concurrent integrations of mechanics, aerodynamics and aeromechanics design

Predicting Bladerow Interactions Using a Multistage Time-Linearized Navier-Stokes Solver

A multistage frequency domain (time-linearized/nonlinear harmonic) Navier-Stokes unsteady flow solver has been developed for predicting unsteady flows induced by bladerow interactions. In this paper, the time-linearized option of the salver has been used to analyze unsteady flows in a subsonic turbine test stage and the DLR transonic counterrotating shrouded propfan. The numerical accuracy and computational efficiency of the time-linearized viscous methods have been demonstrated by comparing predictions with test data and nonlinear time-marching solutions for these two test cases. It is concluded that the development of efficient frequency domain approaches enables unsteady flow predictions to be used in the design cycles to tackle aeromechanics problems.

Nonlinear Time and Frequency Domain Methods for Multirow Aeromechanical Analysis

An unsteady Navier–Stokes solution system for aeromechanical analysis of multiple blade row configurations is presented. A distinctive feature of the solver is that unified numerical methods and boundary condition treatments are consistently used for both a nonlinear time-domain solution mode and a frequency-domain one. This not only enables a wider range of physical aeromechanical problems to be tackled, but also provides a consistent basis for validating different computational models, identifying and understanding their relative merits and adequate working ranges. An emphasis of the present work is on a highly efficient frequency-domain method for multirow aeromechanical analysis. With a new interface treatment, propagations and reflections of pressure waves between adjacent blade rows are modeled within a domain consisting of only a single passage in each blade row. The computational model and methods are firstly described. Then, extensive validations of the frequency-domain method against both experimental data and the nonlinear time-domain solutions are described. Finally, the computational analysis and demonstration of the intrarow reflection effects on the rotor aerodynamic damping are presented.

Adjoint Aerodynamic Design Optimization for Blades in Multi-Stage Turbomachines: Part II—Validation and Application

This is the second part of a two-part paper. It presents four case studies. The first case is the redesign of a transonic rotor (NASA rotor 67) at a post peak efficiency operating point. The second case is a redesign of a transonic compressor stage originally designed by DLR. The redesign is carried out at the stage peak efficiency point. The third and fourth cases look at the redesign of blade rows within a three-stage transonic test compressor that was originally designed by Siemens Industrial Turbomachinery Ltd known as the ATC compressor. Specifically the third case is a redesign of the IGV-rotor-stator configuration. It is carried out at two operating points: one is at the stage peak efficiency point; the other is at a lower stagnation pressure ratio choked flow point. Initially the redesign at the stage peak efficiency point produces considerable efficiency gain, but leads to noticeably reduced choked mass flow rate. The redesign at a near choked mass flow rate point, on the other hand, leads to considerable performance deterioration at operating points with lower mass flow rate, though the choked mass flow rate is even increased. Subsequently, a parallel multi-point approach has been implemented. Results show that a two-point design optimization avoids unacceptable performance deterioration at off design conditions. In the fourth case a redesign is applied across all 7 blade rows of the ATC compressor at the compressor design point. All these case studies are aimed to increase isentropic efficiency whilst meeting the specified constraints.Copyright

FLOvane: a new approach for HP vane design

This paper presents a new unconventional philosophy for high-pressure (HP) vane design. It is proposed that the most natural design starting point for admitting and accelerating flow with minimum loss and secondary flow is a trumpet-shaped flow-path which gradually turns to the desired angle. Multiple trumpet-shaped inlets are seamlessly blended into the (annular or partitioned) combustor walls resulting in a highly lofted flow-path, rather than a traditional flow-path defined by distinct airfoil and endwall surfaces. We call this trumped-shaped inlet the fully lofted oval vane (FLOvane). The purpose of this paper is to describe the FLOvane concept and to present back-to-back CFD analyses of two current industrial gas turbines with conventional and FLOvanemodified designs. The resulting designs diverge significantly from conventional designs in terms of both process and final geometric form. Computational fluid dynamic predictions for the FLOvane-modified designs show improved aerodynamic performance characteristics, reduced heat load, improved cooling performance, improved thermal-mechanical life, and improved stage/engine efficiency. The mechanisms for improved performance include reduction of secondary flows, reduced mixing of coolant flow with the mainstream flow, reduced skin friction, and improved coolant distribution. In the two current industrial gas turbine engines, the absolute (percentage point) improvement in stage isentropic efficiency when the FLOvane design was included was estimated at 0.33% points and 0.40% points without cooling flow reduction, and 1.5% points in one case and much more is expected for the other case when cooling flow reductions were accounted for.

Improving Purge Air Cooling Effectiveness by Engineered End-Wall Surface Structures ? Part I: Duct Flow

Motivated by the recent advances in additive manufacturing, this study investigated a new turbine end-wall aerothermal management method by engineered surface structures. The feasibility of enhancing purge air cooling effectiveness through a series of small-scale ribs added onto the turbine end-wall was explored experimentally and numerically in this two-part paper. Part I presents the fundamental working mechanism and cooling performance in a 90u2009deg turning duct (part I), and part II of this paper validates the concept in a more realistic turbine cascade case. In part I, the turning duct is employed as a simplified model for the turbine passage without introducing the horseshoe vortex. End-wall heat transfer and temperature were measured by the infrared thermography. Computational fluid dynamics (CFD) simulation was also performed using ANSYS fluent to compliment the experimental findings. With the added end-wall rib structures, purge air flow was observed to be more attached to the end-wall and cover a larger wall surface area. Both experimental and numerical results reveal a consistent trend on improved film cooling effectiveness. The practical design optimization strategy is also discussed in this paper.

Improving Purge Air Cooling Effectiveness by Engineered End-Wall Surface Structures—Part II: Turbine Cascade

Motivated by the recent advances in additive manufacturing, a novel turbine end-wall aerothermal management method is presented in this two-part paper. The feasibility of enhancing purge air cooling effectiveness through engineered surface structure was experimentally and numerically investigated. The fundamental working mechanism and improved cooling performance for a 90u2009deg turning duct are presented in Part I. The second part of this paper demonstrates this novel concept in a low-speed linear cascade environment. The performance in three purge air blowing ratios is presented and enhanced cooling effectiveness and net heat flux reduction (NHFR) were observed from experimental data, especially for higher blow ratios. The Computational fluid dynamics (CFD) analysis indicates that the additional surface features are effective in reducing the passage vortex and providing a larger area of coolant coverage without introducing additional aerodynamic loss.

Non-Linear Time and Frequency Domain Methods for Multi-Row Aeromechanical Analysis

An unsteady Navier-Stokes solution system for aeromechanical analysis of multiple blade row configurations is presented. A distinctive feature of the solver is that unified numerical methods and boundary condition treatments are consistently used for both a nonlinear time-domain solution mode and a frequency-domain one. This not only enables a wider range of physical aeromechanical problems to be tackled, but also provides a consistent basis for validating different computational models, identifying and understanding their relative merits and adequate working ranges. An emphasis of the present work is on a highly efficient frequency-domain method for multi-row aeromechanic analysis. With a new interface treatment, propagations and reflections of pressure waves between adjacent blade rows are modeled within a domain consisting of only a single passage in each blade row. The computational model and methods are firstly described. Then, extensive validations of the frequency-domain method against both experimental data and the nonlinear time-domain solutions are described. Finally the computational analysis and demonstration of the intra-row reflection effects on the rotor aerodynamic damping are presented.© 2012 ASME