Cooling Structure Optimization for a Rib-Roughed Channel in a Turbine Rotor Blade

2013 ◽  
Author(s):  
Zhongran Chi ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Rotor Blade ◽  
Cooling System ◽  
Rectangular Channel ◽  
Turbine Blades ◽  
Critical Issue ◽  
Turbine Rotor ◽  
Optimization Study ◽  
Cooling Design ◽  
Rib Arrangement ◽  
Optimization Search

Cooling design for the air-cooled turbine blades is a critical issue in modern gas turbine engineering. Advances in CFD technology is providing new prospects for turbine cooling design, as the optimum cooling structures of the blades could be designed through the optimization search coupled with the Conjugate Heat Transfer (CHT) analysis. In this paper, the optimization study for the rib arrangement of a rib-roughed channel in a rotor blade is discussed. The optimization study introduced is realized utilizing a parametric analysis platform, which consists of the parametric design and mesh generation tool and the commercial CHT solver ANSYS CFX. For the optimization study, firstly a group of Design of Experiments (DoE) analysis of a rib-roughed rectangular channel is performed in order to find the optimum rib arrangement and to explore the objective of the optimization search. Then, the optimization search of the optimum rib arrangement is performed for a rib-roughed channel within a rotor blade based on the multi-island Genetic Algorithms (GA) of iSIGHT. During optimization search, a constant pressure drop is assumed within the cooling system, and the CHT simulations are approached for the interior only in order to make the search computationally faster. According to the DoE analysis, minimizing the averaged wall temperature on blade surface is chosen as the optimization objective for the design of rib arrangement. The results of the GA search shows that the optimal rib arrangement with best cooling performance can be decided, and the optimal mass flow rate for the cooling channel is found simultaneously. The optimum schemes of the rib arrangement found by the DoE analysis and GA search are quite identical, which further validates the feasibility of design optimization for the blade cooling structure with the GA and CHT simulations.

A Resonance Tracking Algorithm for the Prediction of Turbine Forced Response With Friction Dampers

2000 ◽  
Author(s):  
C. Bréard ◽  
J. S. Green ◽  
M. Vahdati ◽  
M. Imregun
Keyword(s):  
Rotor Blade ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Forced Response ◽  
Rotor Blades ◽  
Test Case ◽  
Friction Damper ◽  
Blade Vibration ◽  
Frequency Shifts ◽  
Friction Dampers

This paper presents an iterative method for determining the resonant speed shift when non-linear friction dampers are included in turbine blade roots. Such a need arises when conducting response calculations for turbine blades where the unsteady aerodynamic excitation must be computed at the exact resonant speed of interest. The inclusion of friction dampers is known to raise the resonant frequencies by up to 20% from the standard assembly frequencies. The iterative procedure uses a viscous, time-accurate flow representation for determining the aerodynamic forcing, a look-up table for evaluating the aerodynamic boundary conditions at any speed, and a time-domain friction damping module for resonance tracking. The methodology was applied to an HP turbine rotor test case where the resonances of interest were due to the 1T and 2F blade modes under 40 engine-order excitation. The forced response computations were conducted using a multi-stage approach in order to avoid errors associated with “linking” single stage computations since the spacing between the two bladerows was relatively small. Three friction damper elements were used for each rotor blade. To improve the computational efficiency, the number of rotor blades was decreased by 2 to 90 in order to obtain a stator/rotor blade ratio of 4/9. However, the blade geometry was skewed in order to match the capacity (mass flow rate) of the components and the condition being analysed. Frequency shifts of 3.2% and 20.0% were predicted for the 1T/40EO and 2F/40EO resonances in about 3 iterations. The predicted frequency shifts and the dynamic behaviour of the friction dampers were found to be within the expected range. Furthermore, the measured and predicted blade vibration amplitudes showed a good agreement, indicating that the methodology can be applied to industrial problems.

Coupled Aero-Thermodynamics Optimization for the Cooling System of a Turbine Vane

2013 ◽  
Author(s):  
Zhongran Chi ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Temperature Distribution ◽  
System Design ◽  
Optimization Design ◽  
Cooling System ◽  
Critical Issue ◽  
Operating Conditions ◽  
Critical Parameters ◽  
Pin Fin ◽  
Optimization Search ◽  
Generic Algorithms

Cooling system design for the air-cooled turbine is a critical issue in modern gas turbine engineering. Advances in CFD technology and optimization methodology is providing new prospects for turbine cooling system design, that the optimum cooling system of the vanes and blades could be designed automatically by the optimization search coupled with the Conjugate Heat Transfer (CHT) analysis. An optimization platform consists of the Generic Algorithms (GA), a mesh generation tool (Coolmesh), and the CHT solver (ANSYS CFX) is presented in this paper. The optimization study was aimed at finding the optimum cooling structure for the 2nd stage vane of the E3 engine, with acceptable metal temperature distribution and limited coolant amount simultaneously. The vane was installed with impingement and pin-fin cooling structure. The optimization search involved the design of critical parameters of the cooling system, including the size of impingement tube, diameter and distribution of impingement holes, and the size and distribution of pin-fin near trailing edge. The optimization design was carried under two engine operating conditions to explore the affect of different boundary conditions. A constant pressure drop was assumed within the cooling system during each optimization. To make the problem computationally faster, the simulations were approached for the interior only (solid and coolant). A weighed function of temperature distribution and coolant mass flow was used as the objective of the Single Objective Generic Algorithms (SOGA). The result showed that the optimal cooling system configuration with considerable cooling performance could be designed through SOGA optimization without human interference.

Quantitative MRI Measurements of Hot Streak Development in a Turbine Vane Cascade

2015 ◽  
Author(s):  
Sayuri D. Yapa ◽  
Christopher J. Elkins ◽  
John K. Eaton
Keyword(s):  
Film Cooling ◽  
Thermal Stresses ◽  
Turbine Blades ◽  
Three Dimensional ◽  
Turbine Rotor ◽  
Quantitative Mri ◽  
Inlet Temperature ◽  
Cooling Design ◽  
Pass Through ◽  
Hot Streak

Hot streaks from the combustor and cool streaks from nozzle vane film cooling impose strong inlet temperature variations on high pressure turbine blades, which can lead to local hot or cold spots, high thermal stresses, and fatigue failures. Furthermore, the complex three dimensional flows around the vane may act to concentrate cool or hot fluid exiting the vane row. In order to optimize the cooling design of the turbine blades, the designer must be able to predict the temperature distribution entering the turbine rotor. Therefore, it is important to understand and predict how combustor hot streaks are dispersed as they pass through the vane row. The goal of the present work is to provide detailed three dimensional velocity and temperature data for simulated combustor hot streaks developing through a film cooled vane cascade using the Magnetic Resonance Velocity/Concentration experimental technique. The measurements show that the hot streaks are thinned by acceleration through the vane cascade and diffused by turbulence. The turbulent diffusivity is suppressed by acceleration and leaves significant temperature nonuniformity in the vane wake.

scholarly journals Design Studies for Twist-Coupled Wind Turbine Blades

2003 ◽  
Author(s):  
James Locke ◽  
Ulyses Valencia ◽  
Kosuke Ishikawa
Keyword(s):  
Wind Turbine ◽  
Rotor Blade ◽  
Turbine Blades ◽  
Load Condition ◽  
Turbine Rotor ◽  
Unit Load ◽  
Baseline Design ◽  
Turbine Rotor Blade ◽  
Extreme Wind ◽  
Wind Turbine Rotor

This study presents results obtained for three designs of the Northern Power Systems (NPS) 9.2-meter version of the ERS-100 wind turbine rotor blade. The ERS-100 wind turbine rotor blade was designed and developed by TPI composites. The baseline design uses e-glass unidirectional fibers in combination with ±45-degree and random mat layers for the skin and spar cap. This project involves developing structural finite element models of the baseline design and carbon hybrid designs with twist-bend coupling. All designs were evaluated for a unit load condition and two extreme wind conditions. The unit load condition was used to evaluate the static deflection, twist and twist-coupling parameter. Maximum deflections and strains were determined for the extreme wind conditions. Buckling eigenvalues were determined for a tip load condition. The results indicate that carbon fibers can be used to produce twist-coupled designs with comparable deflections, strains and buckling loads.

Novel Turbine Rotor Shroud Film-Cooling Design and Validation: Part 1

2009 ◽  
Author(s):  
K. S. Chana ◽  
B. Haller
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Turbine Rotor ◽  
Test Facility ◽  
Pressure Probe ◽  
Dimensional Approach ◽  
Flow Angle ◽  
Cooling Hole ◽  
Cooling Design

This paper is part one of a two part paper which considers a shroud film-cooling system designed using a two-dimensional approach. Heat transfer to rotor-casings has reached levels that are causing in-service difficulties to be experienced. Future designs are likely to need to employ film-cooling of some form. There is currently very little information available for film-cooling on shroudless turbine rotor-casing liners. Heat transfer literature on uncooled configurations is not extensive and in particular, spatially-detailed, time-accurate data are rare. This paper describes the aero-thermodynamic design and validation of a rotor casing film-cooling system for a transonic, high-pressure shroudless turbine stage. The design was carried out using a boundary layer code with the film-cooling hole geometry representative of an engine configuration and, has been subjected to mechanical constraints similar to those for an engine component. The design consists of two double rows of cooling holes and two ‘cooling-hole’ shape configurations, cylindrical and fan shaped. The design was tested in the QinetiQ short duration turbine test facility (TTF). Measurements taken include casing heat transfer using thin film gauges and stage exit total pressure, Mach number and flow angle using a three-hole pressure probe. Results showed that while the cooling produced a reduction in the heat transfer rate close to the injection point, the film was stripped off the casing and entrained in nozzle guide vane secondary and rotor overtip flow, where it was transported spanwise towards the hub in the rotor passage. Using the results obtained from this deign a second cooling design was carried out, using a three-dimensional approach this gave significantly better cooling performance. The thee-dimensional design and validation is reported in GT2009-60246 as part 2 of this paper.

Effect of Tip Gap and Squealer Geometry on Detailed Heat Transfer Measurements Over a High Pressure Turbine Rotor Blade Tip

2004 ◽  
Vol 126 (2) ◽  
pp. 221-228 ◽  
Author(s):  
Hasan Nasir ◽  
Srinath V. Ekkad ◽  
David M. Kontrovitz ◽  
Ronald S. Bunker ◽  
Chander Prakash
Keyword(s):  
Heat Transfer ◽  
Rotor Blade ◽  
Pressure Ratio ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Transfer Coefficients ◽  
Blow Down ◽  
Turbine Rotor Blade ◽  
Heat Transfer Coefficients ◽  
Gap Height

The present study explores the effects of gap height and tip geometry on heat transfer distribution over the tip surface of a HPT first-stage rotor blade. The pressure ratio (inlet total pressure to exit static pressure for the cascade) used was 1.2, and the experiments were run in a blow-down test rig with a four-blade linear cascade. A transient liquid crystal technique was used to obtain the tip heat transfer distributions. Pressure measurements were made on the blade surface and on the shroud for different tip geometries and tip gaps to characterize the leakage flow and understand the heat transfer distributions. Two different tip gap-to-blade span ratios of 1% and 2.6% are investigated for a plane tip, and a deep squealer with depth-to-blade span ratio of 0.0416. For a shallow squealer with depth-to-blade span ratio of 0.0104, only 1% gap-to-span ratio is considered. The presence of the squealer alters the tip gap flow field significantly and produces lower overall heat transfer coefficients. The effects of different partial squealer arrangements are also investigated for the shallow squealer depth. These simulate partial burning off of the squealer in real turbine blades. Results show that some partial burning of squealers may be beneficial in terms of overall reduction in heat transfer coefficients over the tip surface.

Experimental Investigation of Two Competitive High Pressure Turbine Blade Cooling Systems

2017 ◽  
Author(s):  
Sergiy Risnyk ◽  
Andriy Artushenko ◽  
Igor Kravchenko ◽  
Sergii Borys
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
Cooling System ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
High Pressure Turbine ◽  
Cooling Systems ◽  
Turbine Blade Cooling ◽  
Blade Cooling

Aeroengine high-pressure turbine (HPT) is the key engine component. HPT blade must withstand high inlet temperatures and mechanical loads providing the necessary level of the efficiency. To achieve these objectives effective and complex blade cooling systems (internal convective and film cooling) are used in the HPT design. The objective of this project is to design and investigate the aeroengine HPT blade cooling system that is able to withstand the blade inlet gas temperature level of approx. 1900K but with the minimal cooling airflow amount. HPT blade of the aeroengine with unducted fan (UDF) was taken as a baseline design, namely, the monocrystal blade with a convective multipass system and the film cooling. Advanced HPT blade inter-wall cooling system was designed, investigated and compared with the typical baseline HPT blade. In the advanced HPT blade inter-wall cooling system special types and structure of cooling channels are used. Both types of cooling systems were investigated experimentally in the turbine rotor of the high temperature core engine. Measurements of turbine blades temperatures were performed using crystal temperature sensors (CTS). HPT blades with two competitive cooling systems incorporated with CTS (0,2–0,3 mm size) were installed in the turbine rotor of the core engine and tested on the engine Maximal rate. After tests and the engine disassembly CTSs were extracted and the characteristics of the CTS crystal lattice were transcribed in temperature values. Thermal state of both two competitive cooling systems was validated by experimental data. Numerical and experimental results obtained in the research of HPT blade cooling system are presented in the article. Aeroengine high pressure turbine blade cooling systems designs are described.

Coupled Aerothermodynamics Optimization for the Cooling System of a Turbine Vane

2013 ◽  
Vol 136 (5) ◽  
Author(s):  
Zhongran Chi ◽  
Jing Ren ◽  
Hongde Jiang
Keyword(s):  
Genetic Algorithm ◽  
Temperature Distribution ◽  
System Design ◽  
Cooling System ◽  
Three Dimensional ◽  
Critical Issue ◽  
Operating Conditions ◽  
Critical Parameters ◽  
Pin Fin ◽  
Optimization Search

The cooling system design for air-cooled turbines is a critical issue in modern gas turbine engineering. Advances in the computational fluid dynamics (CFD) technology and optimization methodology are providing new prospects for turbine cooling system design, in the sense that the optimum cooling system of the vanes and blades could be designed automatically by the optimization search coupled with the full three-dimensional conjugate heat transfer (CHT) analysis. An optimization platform for air-cooled turbines, which consists of the genetic algorithm (GA), a mesh generation tool (Coolmesh), and a CHT solver is presented in this paper. The optimization study was aimed at finding the optimum cooling structure for a 2nd stage vane with, simultaneously, an acceptable metal temperature distribution and limited amount of coolant. The vane was installed with an impingement and pin-fin cooling structure. The optimization search involved the design of the critical parameters of the cooling system, including the size of the impingement tube, diameter and distribution of impingement holes, and the size and distribution of the pin-fin near trailing edge. The design optimization was carried out under two engine operating conditions in order to explore the effects of different boundary conditions. A constant pressure drop was assumed within the cooling system during each optimization. To make the problem computationally faster, the simulations were approached for the interior only (solid and coolant). A weighted function of the temperature distribution and coolant mass flow was used as the objective of the single objective genetic algorithm (SOGA). The result showed that the optimal cooling system configuration with considerable cooling performance could be designed through the SOGA optimization without human interference.

Novel Turbine Rotor Shroud Film-Cooling Design and Validation: Part 2

2009 ◽  
Author(s):  
K. S. Chana ◽  
B. Haller
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Cooling System ◽  
Three Dimensional ◽  
Poor Performance ◽  
Turbine Rotor ◽  
Test Facility ◽  
Test Results ◽  
Film Cooling Effectiveness ◽  
Cooling Design

This paper is part two of a two part paper which considers a shroud film-cooling system design. The design was carried out using test results from a previous two-dimensional (2D) design and optimisation using three-dimensional (3D) CFD. The first cooling design was carried out using a streamline boundary layer approach and tested in the QinetiQ turbine test facility (TTF). The test results showed the design did not function as well as had been predicted and gave a poor performance in terms of film cooling effectiveness. Lessons learnt from the 2D design as well as understanding gained from heat transfer and pressure data taken on the rotor casing led to the formulation of a completely new design philosophy. Accepting, cooling films would not survive rotor passing and therefore concentrating on localised cooling as well as the re-establishment of cooling films between rotor passings. The design concept was validated/optimised with the aid of 3D CFD. Heat transfer instrumentation was implemented in a cooling insert fitted over the test rotor to evaluate the performance of the design. Tests carried out with and without cooling showed an improvement in cooling performance, leading to a 40% reduction in heat transfer rate to the rotor casing across the rotor overtip region. A significant improvement was achieved with the new design over the original with reductions in casing heat transfer rates of up to 44%, with a design coolant mass flow of 1.85% of core flow. Heat transfer data were successfully processed to Nusselt number, allowing the results to be translated to a gas turbine engine design.

Fundamental Gas Turbine Heat Transfer

2013 ◽  
Vol 5 (2) ◽  
Author(s):  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Cooling System ◽  
Turbine Blades ◽  
Turbine Rotor ◽  
Industrial Applications ◽  
Future Research ◽  
Gas Turbine Heat Transfer ◽  
Cooling Air

Gas turbines are used for aircraft propulsion and land-based power generation or industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperatures (RIT). Current advanced gas turbine engines operate at turbine RIT (1700 °C) far higher than the melting point of the blade material (1000 °C); therefore, turbine blades are cooled by compressor discharge air (700 °C). To design an efficient cooling system, it is a great need to increase the understanding of gas turbine heat transfer behaviors within complex 3D high-turbulence unsteady engine-flow environments. Moreover, recent research focuses on aircraft gas turbines operating at even higher RIT with limited cooling air and land-based gas turbines burn coal-gasified fuels with a higher heat load. It is important to understand and solve gas turbine heat transfer problems under new harsh working environments. The advanced cooling technology and durable thermal barrier coatings play critical roles for the development of advanced gas turbines with near zero emissions for safe and long-life operation. This paper reviews fundamental gas turbine heat transfer research topics and documents important relevant papers for future research.

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