Christian Cornelius

Unsteady CFD Methods in a Commercial Solver for Turbomachinery Applications

Modern CFD flow solvers can be readily used to obtain time-averaged results on industrial size turbomachinery flow problem at low computational cost and overall effort. On the other hand, time-accurate computations are still expensive and require substantial resources in CPU and computer memory. However, numerical techniques such as phase shift and time inclining method can be used to reduce overall computational cost and memory requirements. The unsteady effects of moving wakes, tip vortices and upstream propagation of shock waves in the front stages of multi-stage compressors are crucial to determine the stability and efficiency of gas turbines at part-load conditions. Accurate predictions of efficiency and aerodynamic stability of turbomachinery stages with strong blade row interaction based on transient CFD simulations are therefore of increasing importance today. The T106D turbine profile is under investigation as well as the transonic compressor test rig at Purdue. The main objective of this paper is to contribute to the understanding of unsteady flow phenomena that can lead to the next generation design of turbomachinery blading. Transient results obtained from simulations utilizing shape correction (phase shift) and time inclining methods in an implicit pressure-based solver, are compared with those of a full transient model in terms of computational cost and accuracy. NOMENCLATURE Acronyms DP

CFD Analysis of a 15 Stage Axial Compressor: Part I — Methods

The flow field of a 15 stage axial compressor is analyzed using a 3-D Navier-Stokes CFD tool. The compressor under investigation is a prototype first compressor version before optimization of the Siemens V84.3A family. A methodology is described for steady state and transient flow simulations of the entire 15 stages compressor in one computation (not piece by piece). The simulation includes tip gaps, mass bleeds, hub leakage flows, and ranges from single passage to full 360 degrees analysis. The work is divided into two companion papers. This first part, “CFD Analysis of a 15 Stage Axial Compressor Part I: Methods” describes the overall methodology used, based on a middle portion of the compressor R5S9 (Rotor 5 to Stator 9). Various effects are analyzed: mesh style and refinement, boundary conditions, steady or transient, tip clearance, and numerical issues (turbulence model choice, advection model choice, parallel processing performance). A high sensitivity of the predictions to the tip clearance height was found. Excellent design point predictions are obtained with steady-state frame change interface models (Stage average interface), as well as with transient simulations (transient rotor-stator interface).Copyright

Theory and Application of Axisymmetric Endwall Contouring for Compressors

Engine operating range and efficiency are of increasing importance in modern compressor design for heavy duty gas turbines and aircraft engines. These highly challenging objectives can only be met if all components provide high aerodynamic performance and stability. The aerodynamic losses of highly loaded axial compressors are mainly influenced by the leakage flow through clearance gaps. Especially the leakage flow due to the radial clearances of rotor blades affects negatively both, the efficiency and the operating range of the engine. Recent publications showed that the clearance flow and the clearance vortex can be influenced by an additional static pressure gradient at the outer casing, which is created by an axisymmetric wavy casing shape. A notable performance increase of up to 0.4% stage efficiency at design point conditions was reported for high pressure stages with large tip clearance heights [1] as well as for a transonic stage with a relatively small radial clearance gap [2]. An analytic approach to predict the effects of axisymmetric casing contouring has been developed at DLR, Institute of Propulsion Technology, and is outlined in the first part of this work. The characteristic behavior of the clearance vortex in an adverse pressure gradient is discussed by means of an inviscid vortex model [3]. The critical vortex parameters are isolated and related to the static pressure increase due to the casing contour. The second part illustrates the application of an axisymmetric endwall contour. A three dimensional optimization of the outer casing and the corresponding blade tip airfoil section of a typical gas turbine high pressure compressor stage with a high number of free variables is presented. The optimization led to a significant increase in aerodynamic performance of about 0.8% stage efficiency and to a notable reduction of the endwall blockage at ADP conditions. Furthermore, an improved off-design performance was found and a simple design rule is given to transfer both, the casing contour and the blade tip section modification on similar high pressure compressor blades. Based on these design rules the results of the optimized stages were applied to the rear stages of a Siemens gas turbine compressor CFD model. An increase of 0.3% full compressor performance was reached at design point conditions.© 2011 ASME

CFD Analysis of a 15 Stage Axial Compressor: Part II — Results

The flow field of a 15 stage axial compressor is analyzed using a 3-D Navier-Stokes CFD tool. The compressor under investigation is a prototype engine, first compressor version before optimization of the Siemens V84.3A family. The paper describes steady state and transient flow simulations of the entire 15 stages compressor in one computation (not piece by piece). The simulation includes tip gaps, mass bleeds, hub leakage flows, and ranges from single passage to full 360 degrees analysis. The work is divided into two companion papers. The second paper, “CFD Analysis of a 15 Stage Axial Compressor Part II: Results” describes the application of the methods in Part I to the entire 15 stage compressor (Belamri et al, 2005). The flow in the compressor is modeled first with one blade passage per component (periodicity assumed, an interface pitch change model employed). Steady state and transient models are compared. In a second series of computations, all blade passages in 360 degrees are modeled, (no periodicity or pitch change assumptions required), for portions of the compressor. The various simulation approaches are compared to each other, and to experimental data. Good agreement between predictions and experimental results, both in the details of the flow field and the integral prediction of operating range of the compressor, were found.© 2005 ASME

Efficient Time Resolved Multistage CFD Analysis Applied to Axial Compressors

Unsteady computations are necessary if blade row interactions effects are relevant, for example for detailed optimization of a compressor at off-design conditions towards the aerodynamic stability limit, or for structural mechanical tuning of the blades. Modeling time accurate transient multistage flow is expensive both in terms of computer time and memory. Recently the implicit time-resolved Time Transformation method (based on Giles’ time inclining) has been shown to be computationally efficient and a good alternative for modeling transient flow in a single stage (one pitch ratio) turbomachinery configuration. A further advantage of this time resolved method is its ability to capture not only blade passing frequencies but also self-excited frequencies such as in wakes and tip vortex shedding.In this work, an extension of the Time Transformation method (TT) to multistage modeling has been employed to assess the method’s ability in predicting modern multistage compressor performance speedline curve, as well as its ability in capturing dominant machine frequencies. The multistage TT method is verified on a two and a half stage modified Hannover compressor, followed by an industrial validation on a Siemens Energy half scale six stage axial compressor based on the last stages of the Siemens Platform Compressor (PCO). Reference transient solutions on reduced portions of the compressor and/or modified blade count solutions are obtained and compared directly to single passage multistage Time Transformation predictions for the Hannover compressor. The method is then applied directly to the full six stage Siemens compressor employing the true blade counts for all six stages.The first goal of this work is to investigate the ability and accuracy of the multistage TT method to capture all relevant blades passing frequencies, including the impact of different degrees of pitch change between components. The second goal of this work is to explore how best to apply the method for the prediction of a compressor map, up to the surge line. Solutions are compared to experimental test rig data. Physical explanations of the key flow features observed in the experiment, as well as of the differences between the predictions and experimental data, are given.Copyright

Speed Line Computation of a Transonic Compressor Stage With Unsteady CFD Methods

The flow through a transonic compressor stage is dominated by unsteady effects such as shock propagation and wake shedding. An accurate prediction of the performance of a compressor, i.e. operating range and efficiency, may require the modeling of unsteady effects. Steady CFD methods cease to converge too early when the stall limit is approached. Efficient unsteady CFD methods such as the transient time-inclining (TI) method and the perturbation based non-linear harmonic (NLH) method perform better and are becoming increasingly popular in the industry. Both methods consider the actual blade count ratio for each passage while using a single passage model.The main objective of this paper is to explain these methods and benchmark their performance with respect to reliable near stall predictions. Computed compressor characteristics and blade row interaction effects of the Purdue Transonic Research Compressor are compared to measurement data. The stator row is found to be limited at the casing in all of the unsteady simulation results. This effect is also qualitatively predicted by steady results calculated at a lower back pressure level. The NLH method is significantly faster than the other transient methods and the TI method resolves more flow detail on identical meshes.Copyright

Rotor Casing Contouring in High Pressure Stages of Heavy Duty Gas Turbine Compressors With Large Tip Clearance Heights

Clearance leakage losses of axial compressor rotors and stators have a major impact on the overall compressor performance. The clearance heights in the last stages (high pressure stages) of a gas turbine compressor are very large in comparison to the low pressure stages due to mechanical constraints and small blade heights. The reduction of clearance leakage losses in a high pressure stage still holds an important potential for the overall performance improvement at design point conditions. In the following work, a method for tip clearance loss reduction by circumferential casing contouring above a high pressure stage rotor with a constant clearance height is presented. The subsonic compressor blade provides Siemens HPA-Family [1, 2, 3] airfoils. Starting over with a 3DOptimization of the mentioned rotor casing the work additionally refers to the aerodynamic effects and the off design performance of the optimized geometry. It has been found that an optimized casing and blade tip contour lead to a smaller overall clearance mass flow and lower pressure loss coefficient of the clearance flow so that the endwall blockage is reduced and the stage performance is improved by about 0.35% at design point conditions. Furthermore it was found that the performance improvement drops with increasing exit pressure to about 0.1% close to stall conditions. At lower exit pressure values the optimized geometry provides an additional performance improvement in comparison to the baseline configuration.

Shock Induced Vortices in Transonic Compressors: Physical Interpretations

Unsteady aerodynamics is becoming more and more important for future compressor improvements. It has to be taken into account to reduce the amount of uncertainties for increasing the performance. Uncertainties regarding unsteady blade row interaction are difficult to approximate due to the complex flow phenomena and the high computational effort to evaluate these phenomena. The optimal axial gapping for transonic compressors is particularly difficult to define. A wide range of axial gappings for two different state of the art compressor front stages (F-class and DLR Rig250) have been simulated to understand the underlying physics and to derive design rules for the optimal axial gapping. The incidence angle reduction due to shock induced vortices has not jet been fully investigated. Predicting the exact vortex trajectory in the rotor passage would be a critical step in assessing this reduction in incidence angle. A function to predict the deviation compared to the steady state compressor performance is usefull for future design processes.