Accelerated 3D Aerodynamic Optimization of Gas Turbine Blades

2014 ◽  
Author(s):  
Philipp Amtsfeld ◽  
Michael Lockan ◽  
Dieter Bestle ◽  
Marcus Meyer
Keyword(s):  
Turbine Blades ◽  
Three Dimensional ◽  
Optimization Process ◽  
Design Parameters ◽  
Blade Design ◽  
Aerodynamic Optimization ◽  
Multi Stage ◽  
Design Variables ◽  
Real Flow ◽  
Blade Model

State-of-the-art aerodynamic blade design processes mainly consist of two phases: optimal design of 2D blade sections and then stacking them optimally along a three-dimensional stacking line. Such a quasi-3D approach, however, misses the potential of finding optimal blade designs especially in the presence of strong 3D flow effects. Therefore, in this paper a blade optimization process is demonstrated which uses an integral 3D blade model and 3D CFD analysis to account for three-dimensional flow features. Special emphasis is put on shortening design iterations and reducing design costs in order to obtain a rapid automatic optimization process for fully 3D aerodynamic turbine blade design which can be applied in an early design phase already. The three-dimensional parametric blade model is determined by up to 80 design variables. At first, the most important design parameters are chosen based on a non-linear sensitivity analysis. The objective of the subsequent optimization process is to maximize isentropic efficiency while fulfilling a minimal set of constraints. The CFD model contains both important geometric features like tip gaps and fillets, and cooling and leakage flows to sufficiently represent real flow conditions. Two acceleration strategies are used to cut down the turn-around time from weeks to days. Firstly, the aerodynamic multi-stage design evaluation is significantly accelerated with a GPU-based RANS solver running on a multi-GPU workstation. Secondly, a response surface method is used to reduce the number of expensive function evaluations during the optimization process. The feasibility is demonstrated by an application to a blade which is a part of a research rig similar to the high pressure turbine of a small civil jet engine. The proposed approach enables an automatic aerodynamic design of this 3D blade on a single workstation within few days.

Optimization of Supersonic Axial Turbine Blades Based on Surrogate Models

2020 ◽  
Author(s):  
Markus Waesker ◽  
Bjoern Buelten ◽  
Norman Kienzle ◽  
Christian Doetsch
Keyword(s):  
Reynolds Number ◽  
Turbine Blade ◽  
Surrogate Model ◽  
Cfd Simulation ◽  
Turbine Blades ◽  
Energy System ◽  
Design Parameters ◽  
Blade Design ◽  
Axial Turbine ◽  
Velocity Coefficient

Abstract Due to the transition of the energy system to more decentralized sector-coupled technologies, the demand on small, highly efficient and compact turbines is steadily growing. Therefore, supersonic impulse turbines have been subject of academic research for many years because of their compact and low-cost conditions. However, specific loss models for this type of turbine are still missing. In this paper, a CFD-simulation-based surrogate model for the velocity coefficient, unique incidence as well as outflow deviation of the blade, is introduced. This surrogate model forms the basis for an exemplary efficiency optimization of the “Colclough cascade”. In a first step, an automatic and robust blade design methodology for constant-channel blades based on the supersonic turbine blade design of Stratford and Sansome is shown. The blade flow is fully described by seven geometrical and three aerodynamic design parameters. After that, an automated numerical flow simulation (CFD) workflow for supersonic turbine blades is developed. The validation of the CFD setup with a published supersonic axial turbine blade (Colclough design) shows a high consistency in the shock waves, separation zones and boundary layers as well as velocity coefficients. A design of experiments (DOE) with latin hypercube sampling and 1300 sample points is calculated. This CFD data forms the basis for a highly accurate surrogate model of supersonic turbine blade flow suitable for Mach numbers between 1.1 and 1.6. The throat-based Reynolds number is varied between 1*104 and 4*105. Additionally, an optimization is introduced, based on the surrogate model for the Reynolds number and Mach number of Colclough and no degree of reaction (equal inlet and outlet static pressure). The velocity coefficient is improved by up to 3 %.

Influence of Film Cooling Hole Angles and Geometries on Aerodynamic Loss and Net Heat Flux Reduction

2011 ◽  
Author(s):  
Chia Hui Lim ◽  
Graham Pullan ◽  
Peter Ireland
Keyword(s):  
Heat Flux ◽  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Design Parameters ◽  
Aerodynamic Loss ◽  
Design Engineers ◽  
Net Heat Flux ◽  
Cooling Hole ◽  
Shaped Hole

Turbine design engineers have to ensure that film cooling can provide sufficient protection to turbine blades from the hot mainstream gas, while keeping the losses low. Film cooling hole design parameters include inclination angle (α), compound angle (β), hole inlet geometry and hole exit geometry. The influence of these parameters on aerodynamic loss and net heat flux reduction is investigated, with loss being the primary focus. Low-speed flat plate experiments have been conducted at momentum flux ratios of IR = 0.16, 0.64 and 1.44. The film cooling aerodynamic mixing loss, generated by the mixing of mainstream and coolant, can be quantified using a three-dimensional analytical model that has been previously reported by the authors. The model suggests that for the same flow conditions, the aerodynamic mixing loss is the same for holes with different α and β but with the same angle between the mainstream and coolant flow directions (angle κ). This relationship is assessed through experiments by testing two sets of cylindrical holes with different α and β: one set with κ = 35°, another set with κ = 60°. The data confirm the stated relationship between α, β, κ and the aerodynamic mixing loss. The results show that the designer should minimise κ to obtain the lowest loss, but maximise β to achieve the best heat transfer performance. A suggestion on improving the loss model is also given. Five different hole geometries (α = 35.0°, β = 0°) were also tested: cylindrical hole, trenched hole, fan-shaped hole, D-Fan and SD-Fan. The D-Fan and the SD-Fan have similar hole exits to the fan-shaped hole but their hole inlets are laterally expanded. The external mixing loss and the loss generated inside the hole are compared. It was found that the D-Fan and the SD-Fan have the lowest loss. This is attributed to their laterally expanded hole inlets, which lead to significant reduction in the loss generated inside the holes. As a result, the loss of these geometries is ≈ 50% of the loss of the fan-shaped hole at IR = 0.64 and 1.44.

Non-Dimensional Parameters for Comparing Conventional and Counter-Rotating Turbomachines

2019 ◽  
Author(s):  
J. J. Waldren ◽  
C. J. Clark ◽  
S. D. Grimshaw ◽  
G. Pullan
Keyword(s):  
High Efficiency ◽  
Three Dimensional ◽  
Relative Performance ◽  
Design Parameters ◽  
Blade Design ◽  
Finite Radius ◽  
Counter Rotation ◽  
Dimensional Parameters ◽  
Conventional Machine ◽  
Stage Efficiency

Abstract Counter-rotating turbomachines have the potential to be high efficiency, high power density devices. Comparisons between conventional and counter-rotating turbomachines in the literature make multiple and often contradicting conclusions about their relative performance. By adopting appropriate non-dimensional parameters, based on relative blade speed, the design space of conventional machines can be extended to include those with counter-rotation. This allows engineers familiar with conventional turbomachinery to transfer their experience to counter-rotating machines. By matching appropriate non-dimensional parameters the loss mechanisms directly affected by counter-rotation can be determined. A series of computational studies are performed to investigate the relative performance of conventional and counter-rotating turbines with the same non-dimensional design parameters. Each study targets a specific loss source, highlighting which phenomena are directly due to counter-rotation and which are solely due to blade design. The studies range from two-dimensional blade sections to three-dimensional finite radius stages. It is shown that, at hub-to-tip ratios approaching unity, with matched non-dimensional design parameters, the stage efficiency and work output are identical for both types of machine. However, a counter-rotating turbine in the study is shown to have an efficiency advantage over a conventional machine of up to 0.35 percentage points for a hub-to-tip ratio of 0.65. This is due to differences in absolute velocity producing different spanwise blade designs.

Design Study of Dovetail Geometries of Turbine Blades Using Abaqus and Isight

2012 ◽  
Author(s):  
Youngwon Hahn ◽  
John I. Cofer
Keyword(s):  
Surrogate Model ◽  
High Reliability ◽  
Turbine Blades ◽  
Optimization Methods ◽  
Nonlinear Static Analysis ◽  
Design Parameters ◽  
Second Step ◽  
Nsga Ii ◽  
Python Script ◽  
Design Variables

Blades in gas and steam turbines continually face more challenging requirements for high reliability and efficiency. In order to meet these challenges in an increasingly competitive marketplace, blade design engineers are always looking for more efficient ways to design the blades in the shortest possible time and at the lowest possible cost while meeting multiple design objectives. In this paper, several design studies are performed using Abaqus and Isight to optimize the minimum contact pressure and stress around the dovetail of a typical turbine blade in order to achieved desired goals for stress levels. First, nine design parameters describing the dimensions of the dovetail are set up in a Python script which can be executed in Abaqus/CAE. The Python script generates the entire finite element model including boundary and loading conditions in Abaqus/CAE. A nonlinear static analysis considering centrifugal loading is performed in this work. After setting up the workflow using the Python script and Abaqus/CAE, Isight is used to automate the process to achieve the optimized dimensions of the dovetail. The optimization is performed in two steps. First, a surrogate model using the Optimal Latin Hypercube approximation method is created using tools in Isight. In this step, the surrogate model is used to determine the optimum values of the design variables, as well as the sensitivity of the design to the selected design variables. It also can be observed that the design is especially sensitive to five of the design variables. In the second step of the optimization, the five design variables to which the design is most sensitive are selected for further optimization by setting the other design variables to the optimized values obtained in the first step of the optimization. In this second step, several different optimization methods supported in Isight are used, including the NSGA-II non-dominated sorting genetic algorithm, Downhill Simplex, and an evolutionary optimization algorithm. Results from these methods are compared with those obtained using other common optimization methods in Isight.

A Criterion for Axial Compressor Hub-Corner Stall

2006 ◽  
Author(s):  
V. M. Lei ◽  
Z. S. Spakovszky ◽  
E. M. Greitzer
Keyword(s):  
Flow Control ◽  
Substantial Reduction ◽  
Three Dimensional ◽  
Axial Compressor ◽  
Diffusion Parameter ◽  
Design Parameters ◽  
Diffusion Limit ◽  
Specific Flow ◽  
Multi Stage ◽  
Flow Variables

This paper presents a new criterion for estimating the size and strength of three-dimensional hub-corner stall in rotors and shrouded stators of multi-stage axial compressors. A simple, first-of-a-kind description for the formation of hub-corner stall is derived, consisting of (i) a stall indicator, which quantifies the extent of the reversed flow via the local blade loading and thus indicates whether corner stall occurs, and (ii) a diffusion parameter which defines the diffusion limit. The stall indicator can be cast in terms of a Zweifel loading coefficient. The diffusion parameter is based on preliminary design type flow variables and geometry. Computational simulations and single and multi-stage compressor data are used to show the applicability of the criterion over a range of blade design parameters. The criterion also enables determination of specific flow control actions needed to mitigate hub-corner stall. To illustrate the latter a flow control blade, designed using the ideas developed, is seen to achieve a substantial reduction in the flow non-uniformity associated with hub-corner stall.

An Optimum Aero-Design Process of Turbine Blades

2002 ◽  
Author(s):  
C. Xu ◽  
R. S. Amano
Keyword(s):  
Design Process ◽  
High Efficiency ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Optimization Procedure ◽  
Compressor Blade ◽  
Advanced Technology ◽  
Test Case ◽  
Blade Design

With the development of the advanced technology, the combustion temperature is raised for increased efficiencies. At the same time, the turbine and compressor pressure ratio and the mass flow rate rise; thus causing turbine and compressor blades turning and blade lengths increase. Moreover, the high efficiency requirements had made the turbine and compressor blade design difficult. A turbine airfoil has been custom designed for many years, but an optimization for the section design in a three-dimensional consideration is still a challenge. For a compressor blade design, standard section cannot meet the modern compressor requirements. Modern compressor design has not only needs a custom designed section according to flow situation, but also needs three-dimensional optimizations. Therefore, a good blade design process is critical to the turbines and compressors. A blade design of the turbomachines is one of the important steps for a good turbomachine design. A blade design process not only directly influences the overall machine efficiency but also dramatically impact the design time and cost. In this study, a blade design and optimization procedure was proposed for both turbine and compressor blade design. A compressor blade design was used as a test case. It was shown that the current design process had more advantages than conventional design methodology.

Full Three-Dimensional Optimization Platform of Turbine Blades Considering the Film Cooling

2013 ◽  
Author(s):  
Shaopeng Lu ◽  
Zhongran Chi ◽  
Songtao Wang ◽  
Fengbo Wen ◽  
Guotai Feng
Keyword(s):  
Film Cooling ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Optimization Process ◽  
Pareto Optimal ◽  
Cooling Efficiency ◽  
Objective Functions ◽  
Nsga Ii ◽  
Aerodynamic Efficiency ◽  
Film Cooling Efficiency

In this paper, an optimization platform was established with Isight, cfx and the self-programming program which is used to generate the mesh. Film cooling effect can be taken into account. 15 parameters are selected as optimization variables. During the optimization process, the baseline blade and cooling holes are given by parameterized method. There are two objective functions during the optimization process. The first one is aerodynamic efficiency and the second one is film cooling efficiency. As there are two objective functions, NSGA-II is chosen as the multi-objective optimization algorithm. Then the Pareto-optimal front can be got. The results show that aerodynamic efficiency and film cooling efficiency restrict each other. It’s impossible to get the best solutions in one example, so the Pareto optimal set can provide a lot of choices. Different shapes make different effects on the aerodynamic efficiency and film cooling efficiency. From the above, it can be seen that the platform is helpful especially in the case that aerodynamic efficiency and film cooling efficiency restrict each other. This paper also discusses the prospects for platform applications.

scholarly journals Design, Optimization and Analysis of Supersonic Radial Turbines

2019 ◽  
Author(s):  
Lukas Benjamin Inhestern ◽  
James Braun ◽  
Guillermo Paniagua ◽  
José Ramón Serrano Cruz
Keyword(s):  
High Performance ◽  
Ad Hoc ◽  
Three Dimensional ◽  
Critical Factor ◽  
Design Tool ◽  
Navier Stokes ◽  
Design Parameters ◽  
Design Variables ◽  
Compact Design ◽  
Radial Outflow

Abstract New compact engine architectures such as pressure gain combustion require ad-hoc turbomachinery to ensure an adequate range of operation with high performance. A critical factor for supersonic turbines is to ensure the starting of the flow passages, which limits the flow turning and airfoil thickness. Radial outflow turbines inherently increase the cross section along the flow path, which holds great potential for high turning of supersonic flow with a low stage number and guarantees a compact design. First the preliminary design space is described. Afterwards a differential evolution multi-objective optimization with 12 geometrical design parameters is deducted. With the design tool AutoBlade 10.1, 768 geometries were generated and hub, shroud, and blade camber line were designed by means of Bezier curves. Outlet radius, passage height, and axial location of the outlet were design variables as well. Structured meshes with around 3.7 million cells per passage were generated. Steady three dimensional Reynolds averaged Navier Stokes (RANS) simulations, enclosed by the k-omega SST turbulence model were solved by the commercial solver CFD++. The geometry was optimized towards low entropy and high power output. To prove the functionality of the new turbine concept and optimization, a full wheel unsteady RANS simulation of the optimized geometry exposed to a nozzled rotating detonation combustor (RDC) has been performed and the advantageous flow patterns of the optimization were also observed during transient operation.

Design and Testing of a Multi-Stage IP Turbine for Future Geared Turbofans

2021 ◽  
Author(s):  
D. Torre ◽  
G. García-Valdecasas ◽  
A. Puente ◽  
D. Hernández ◽  
S. Luque
Keyword(s):  
Three Dimensional ◽  
Thermodynamic Cycle ◽  
Reynolds Numbers ◽  
Secondary Flows ◽  
Aerodynamic Optimization ◽  
Trade Offs ◽  
Multi Stage ◽  
Profile Losses ◽  
Previous State ◽  
Design And Testing

Abstract The multi-stage intermediate pressure turbine (IPT) is a key enabler of the thermodynamic cycle in geared turbofan engine architectures, where fan and turbine rotational speeds become decoupled by employing a power gearbox between them. This allows for the separate aerodynamic optimization of both components, an increase in engine bypass ratios, higher propulsive efficiency, and lower specific fuel consumption. Due to significant aerodynamic differences with conventional low pressure turbines (LPTs), multi-stage IPT designs present new aerodynamic, mechanical and acoustic trade-offs. This work describes the aerodynamic design and experimental validation of a fully featured three-stage IP turbine, including a final row of outlet guide vanes. Experiments have been conducted in a highly engine-representative transonic rotating wind tunnel at the CTA (Centro de Tecnologías Aeronáuticas, Spain), in which Mach and Reynolds numbers were matched to engine conditions. The design intent is shown to be fully validated. Efficiency levels are discussed in the context of a previous state-of-the-art LPT, tested in the same facility. It is argued that the efficiency gains of IPTs are due to higher pitch-to-chord ratios, which lead to a reduction in overall profile losses, and higher velocity ratios and lower turning angles, which reduce airfoil secondary flows and three-dimensional losses.

Influence of Geometric Design Parameters Onto Vibratory Response and HCF Safety for Turbine Blades With Friction Damper

2018 ◽  
Author(s):  
Matthias Hüls ◽  
Lars Panning-von Scheidt ◽  
Jörg Wallaschek
Keyword(s):  
Aspect Ratio ◽  
High Cycle Fatigue ◽  
Turbine Blades ◽  
Design Parameters ◽  
System Response ◽  
Blade Design ◽  
Cycle Fatigue ◽  
Resonance Amplitude ◽  
Damped System ◽  
Vibratory Response

Among the major concerns for high aspect-ratio turbine blades are forced and self-excited (flutter) vibrations which can cause failure by high-cycle fatigue. The introduction of friction damping in turbine blades, such as by coupling of adjacent blades via under platform dampers, can lead to a significant reduction of resonance amplitudes at critical operational conditions. In this paper, the influence of basic geometric blade design parameters onto the damped system response will be investigated to link design parameters with functional parameters like damper normal load, frequently used in nonlinear dynamic analysis. The shape of a simplified large aspect-ratio turbine blade is parameterized along with the under platform damper configuration. The airfoil is explicitly included into the parameterization in order to account for changes in blade mode shapes. For evaluation of the damped system response under a typical excitation, a reduced order model for non-linear friction damping is included into an automated 3D FEA tool-chain. Based on a design of experiments approach, the design space will be sampled and a surrogate model is trained on the received dataset. Subsequently, the mean and interaction effects of the true geometric blade design parameters onto the resonance amplitude and safety against high-cycle fatigue will be outlined for a critical first bending type vibrational motion. Design parameters were mainly found to influence the resonance amplitude by their effect onto the tip-to-platform deflection ratio. The HCF safety was affected by those design parameters with large sensitivity onto static and resonant vibratory stress levels. Applying an evolutionary optimization algorithm, it is shown that the optimum blade design with respect to minimum vibratory response at a particular node can differ significantly from a blade designed toward maximum HCF safety.

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