scholarly journals Erratum to “Optimization of a high pressure turbine blade tip cavity with conjugate heat transfer analysis”

2017 ◽  
Vol 31 (1) ◽  
pp. 479-479
Author(s):  
Jinuk Kim ◽  
Young Seok Kang ◽  
Dongwha Kim ◽  
Jihyeong Lee ◽  
Bong Jun Cha ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Conjugate Heat Transfer ◽  
Heat Transfer Analysis ◽  
High Pressure Turbine ◽  
Blade Tip

Optimization of a high pressure turbine blade tip cavity with conjugate heat transfer analysis

2016 ◽  
Vol 30 (12) ◽  
pp. 5529-5538 ◽  
Author(s):  
Jinuk Kim ◽  
Young Seok Kang ◽  
Dongwha Kim ◽  
Jihyeong Lee ◽  
Bong Jun Cha ◽  
...  
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Conjugate Heat Transfer ◽  
Heat Transfer Analysis ◽  
High Pressure Turbine ◽  
Blade Tip

scholarly journals Computational Modeling and Thermal Paint Verification of Film-Cooling Designs for an Unshrouded High-Pressure Turbine Blade

1999 ◽  
Author(s):  
Steven G. Gegg ◽  
Nathan J. Heidegger ◽  
Ronald A. Mikkelson
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
Data Exchange ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
High Pressure Turbine ◽  
Extreme High Temperature ◽  
Blade Tip

High pressure turbine blades are exposed to an extreme high temperature environment due to increasing turbine inlet temperature. High heat fluxes are likely on the blade pressure surface. Other regions, such as the trailing edge and blade tip may be difficult to cool uniformly. Unshrouded blades present an additional challenge due to the pressure driven transport of hot gas across the blade tip. The blade tip region is therefore prone to severe thermal stress, fatigue and oxidation. In order to develop effective cooling methods, designers require detailed flow and heat transfer information. This paper reports on computational aerodynamics and heat transfer studies for an unshrouded high pressure turbine blade. The emphasis is placed on the application of appropriate 3-D models for the prediction of airfoil surface temperatures. Details of the film cooling model, boundary conditions and data exchange with heat transfer models are described. The analysis approach has been refined for design use to provide timely and accurate results. Film cooling designs are to be tailored for the best coverage of the blade tip region. Designs include near-tip pressure side films and blade tip cooling holes. Hole placement and angle are investigated to achieve the best coolant coverage on the blade tip. Analytical results are compared to a thermal paint test on engine hardware. In addition to film cooling strategies, other aerodynamic/heat transfer design approaches are discussed to address the cooling requirements for an unshrouded blade.

Overtip Shock Wave Structure and Its Impact on Turbine Blade Tip Heat Transfer

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
Q. Zhang ◽  
D. O. O’Dowd ◽  
L. He ◽  
A. P. S. Wheeler ◽  
P. M. Ligrani ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Shock Wave ◽  
High Pressure ◽  
Supersonic Flow ◽  
Turbine Blade ◽  
High Speed ◽  
Wave Structure ◽  
Shock Wave Structure ◽  
High Pressure Turbine ◽  
Blade Tip

In this paper, the transonic flow pattern and its influence on heat transfer on a high-pressure turbine blade tip are investigated using experimental and computational methods. Spatially resolved heat transfer data are obtained at conditions representative of a single-stage high-pressure turbine blade (Mexit=1.0, Reexit=1.27×106, gap=1.5% chord) using the transient infrared thermography technique within the Oxford high speed linear cascade research facility. Computational fluid dynamics (CFD) predictions are conducted using the Rolls-Royce HYDRA/PADRAM suite. The CFD solver is able to capture most of the spatial heat flux variations and gives prediction results, which compare well with the experimental data. The results show that the majority of the blade tip experiences a supersonic flow with peak Mach number reaching 1.8. Unlike other low-speed data in the open literature, the turbine blade tip heat transfer is greatly influenced by the shock wave structure inside the tip gap. Oblique shock waves are initiated near the pressure-side edge of the tip, prior to being reflected multiple times between the casing and the tip. Supersonic flow within the tip gap is generally terminated by a normal shock near the exit of the gap. Both measured and calculated heat transfer spatial distributions illustrate very clear stripes as the signature of the multiple shock structure. Overall, the supersonic part of tip experiences noticeably lower heat transfer than that near the leading-edge where the flow inside the tip gap remains subsonic.

Clocking Effects of Inlet Non-Uniformities in a Fully Cooled High-Pressure Vane: A Conjugate Heat Transfer Analysis

2015 ◽  
Author(s):  
Duccio Griffini ◽  
Massimiliano Insinna ◽  
Simone Salvadori ◽  
Francesco Martelli
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Cooling System ◽  
Hot Spot ◽  
Three Dimensional ◽  
Leading Edge ◽  
Suction Side ◽  
Heat Transfer Analysis

A high-pressure vane equipped with a realistic film-cooling configuration has been studied. The vane is characterized by the presence of multiple rows of fan-shaped holes along pressure and suction side while the leading edge is protected by a showerhead system of cylindrical holes. Steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) simulations have been performed. A preliminary grid sensitivity analysis with uniform inlet flow has been used to quantify the effect of spatial discretization. Turbulence model has been assessed in comparison with available experimental data. The effects of the relative alignment between combustion chamber and high-pressure vanes are then investigated considering realistic inflow conditions in terms of hot spot and swirl. The inlet profiles used are derived from the EU-funded project TATEF2. Two different clocking positions are considered: the first one where hot spot and swirl core are aligned with passage and the second one where they are aligned with the leading edge. Comparisons between metal temperature distributions obtained from conjugate heat transfer simulations are performed evidencing the role of swirl in determining both the hot streak trajectory within the passage and the coolant redistribution. The leading edge aligned configuration is resulted to be the most problematic in terms of thermal load, leading to increased average and local vane temperature peaks on both suction side and pressure side with respect to the passage aligned case. A strong sensitivity of both injected coolant mass flow and heat removed by heat sink effect has also been highlighted for the showerhead cooling system.

Conjugate Heat Transfer Analysis to Assess Overall Cooling Effectiveness of High Pressure Turbine Nozzle With Optimized Film Cooling Hole Arrangements

2019 ◽  
Author(s):  
Young Seok Kang ◽  
Dong-Ho Rhee ◽  
Sanga Lee ◽  
Bong Jun Cha
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Heat Transfer Analysis ◽  
Internal Cooling ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Baseline Design ◽  
Pressure Surface

Abstract Conjugate heat transfer analysis method has been highlighted for predicting heat exchange between fluid domain and solid domain inside high-pressure turbines, which are exposed to very harsh operating conditions. Then it is able to assess the overall cooling effectiveness considering both internal cooling and external film cooling at the cooled turbine design step. In this study, high-pressure turbine nozzles, which have three different film cooling holes arrangements, were numerically simulated with conjugate heat transfer analysis method for predicting overall cooling effectiveness. The film cooling holes distributed over the nozzle pressure surface were optimized by minimizing the peak temperature, temperature deviation. Additional internal cooling components such as pedestals and rectangular rib turbulators were modeled inside the cooling passages for more efficient heat transfer. The real engine conditions were given for boundary conditions to fluid and solid domains for conjugate heat transfer analysis. Hot combustion gas properties such as specific heat at constant pressure and other transport properties were given as functions of temperature. Also, the conductivity of Inconel 718 was also given as a function of temperature to solve the heat equation in the nozzle solid domain. Conjugate heat transfer analysis results showed that optimized designs showed better cooling performance, especially on the pressure surface due to proper staggering and spacing hole-rows compared to the baseline design. The overall cooling performances were offset from the adiabatic film cooling effectiveness. Locally concentrated heat transfer and corresponding high cooling effectiveness region appeared where internal cooling effects were overlapped in the optimized designs. Also, conjugate heat transfer analysis results for the optimized designs showed more uniform contours of the overall cooling effectiveness compared to the baseline design. By varying the coolant mass flow rate, it was observed that pressure surface was more sensitive to the coolant mass flow rate than nozzle leading edge stagnation region and suction surface. The CHT results showed that optimized designs to improve the adiabatic film cooling effectiveness also have better overall cooling effectiveness.

scholarly journals Erratum to: Dynamic probabilistic design approach of high-pressure turbine blade-tip radial running clearance

2016 ◽  
Vol 86 (1) ◽  
pp. 225-225
Author(s):  
Cheng-Wei Fei ◽  
Yat-Sze Choy ◽  
Dian-Yin Hu ◽  
Guang-Chen Bai ◽  
Wen-Zhong Tang
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Design Approach ◽  
Probabilistic Design ◽  
High Pressure Turbine ◽  
Blade Tip

Time-Averaged and Time-Accurate Aero-Dynamics for the Recessed Tip Cavity of a High-Pressure Turbine Blade and the Outer Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Turbine Rotor ◽  
Single Stage ◽  
Cfd Modeling ◽  
Blade Surface ◽  
High Pressure Turbine ◽  
Pressure Transducers ◽  
Blade Tip

The unsteady aero-dynamics of a single-stage high-pressure turbine blade operating at design corrected conditions has been the subject of a thorough study involving detailed measurements and computations. The experimental configuration consisted of a single-stage high-pressure turbine and the adjacent, downstream, low-pressure turbine nozzle row. All three blade-rows were instrumented at three spanwise locations with flush-mounted, high frequency response pressure transducers. The rotor was also instrumented with the same transducers on the blade tip and platform and the stationary shroud was instrumented with pressure transducers at specific locations above the rotating blade. Predictions of the time-dependent flow field around the rotor were obtained using MSU-TURBO, a 3D, non-linear, computational fluid dynamics (CFD) code. Using an isolated blade-row unsteady analysis method, the unsteady surface pressure for the high-pressure turbine rotor due to the upstream high-pressure turbine nozzle was calculated. The predicted unsteady pressure on the rotor surface was compared to the measurements at selected spanwise locations on the blade, in the recessed cavity, and on the shroud. The rig and computational models included a flat and recessed blade tip geometry and were used for the comparisons presented in the paper. Comparisons of the measured and predicted static pressure loading on the blade surface show excellent correlation from both a time-average and time-accurate standpoint. This paper concentrates on the tip and shroud comparisons between the experiments and the predictions and these results also show good correlation with the time-resolved data. These data comparisons provide confidence in the CFD modeling and its ability to capture unsteady flow physics on the blade surface, in the flat and recessed tip regions of the blade, and on the stationary shroud.

Three Dimensional Heat Transfer Analysis of High Pressure Turbine Blade

2009 ◽  
Author(s):  
A. Sipatov ◽  
L. Gomzikov ◽  
V. Latyshev ◽  
N. Gladysheva
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
High Pressure ◽  
Rotor Blade ◽  
Computational Analysis ◽  
Three Dimensional ◽  
Heat Transfer Analysis ◽  
Aircraft Engines ◽  
Turbine Stage ◽  
High Pressure Turbine

The present tendency of creating new aircraft engines with a higher level of fuel efficiency leads to the necessity to increase gas temperature at a high pressure turbine (HPT) inlet. To design such type of engines, the improvement of accuracy of the computational analysis is required. According to this the numerical analysis methods are constantly developing worldwide. The leading firms in designing aircraft engines carry out investigations in this field. However, this problem has not been resolved completely yet because there are many different factors affecting HPT blade heat conditions. In addition in some cases the numerical methods and approaches require tuning (for example to predict laminar-turbulent transition region or to describe the interaction of boundary layer and shock wave). In this work our advanced approach of blade heat condition numerical estimation based on the three-dimensional computational analysis is presented. The object of investigation is an advanced aircraft engine HPT first stage blade. The given analysis consists of two interrelated parts. The first part is a stator-rotor interaction modeling of the investigated turbine stage (unsteady approach). Solving this task we devoted much attention to modeling unsteady effects of stator-rotor interaction and to describing an influence of applied inlet boundary conditions on the blade heat conditions. In particular, to determine the total pressure, flow angle and total temperature distributions at the stage inlet we performed a numerical modeling of the combustor chamber of the investigated engine. The second part is a flow modeling in the turbine stage using flow parameters averaging on the stator-rotor interface (steady approach). Here we used sufficiently finer grid discretization to model all perforation holes on the stator vane and rotor blade, endwalls films in detail and to apply conjugate heat transfer approach for the rotor blade. Final results were obtained applying the results of steady and unsteady approaches. Experimental data of the investigated blade heat conditions are presented in the paper. These data were obtained during full size experimental testing the core of the engine and were collected using two different type of experimental equipment: thermocouples and thermo-crystals. The comparison of experimental data and final results meets the requirements of our investigation.

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