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.

Averaged and Time-Dependent Aerodynamics of a High Pressure Turbine Blade Tip Cavity and Stationary Shroud: Comparison of Computational and Experimental Results

2004 ◽  
Vol 127 (4) ◽  
pp. 736-746 ◽  
Author(s):  
Brian R. Green ◽  
John W. Barter ◽  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Computational Models ◽  
Single Stage ◽  
Time Dependent ◽  
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 three-dimensional (3D), nonlinear, 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.

Using CFD to Reduce Resonant Stresses on a Single-Stage, High-Pressure Turbine Blade

2002 ◽  
Author(s):  
J. P. Clark ◽  
A. S. Aggarwala ◽  
M. A. Velonis ◽  
R. E. Gacek ◽  
S. S. Magge ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Guide Vane ◽  
Turbine Blades ◽  
Suction Side ◽  
Lessons Learned ◽  
Single Stage ◽  
Navier Stokes ◽  
High Pressure Turbine ◽  
Time Resolved

The ability to predict levels of unsteady forcing on high-pressure turbine blades is critical to avoid high-cycle fatigue failures. In this study, 3D time-resolved computational fluid dynamics is used within the design cycle to predict accurately the levels of unsteady forcing on a single-stage high-pressure turbine blade. Further, nozzle-guide-vane geometry changes including asymmetric circumferential spacing and suction-side modification are considered and rigorously analyzed to reduce levels of unsteady blade forcing. The latter is ultimately implemented in a development engine, and it is shown successfully to reduce resonant stresses on the blade. This investigation builds upon data that was recently obtained in a full-scale, transonic turbine rig to validate a Reynolds-Averaged Navier-Stokes (RANS) flow solver for the prediction of both the magnitude and phase of unsteady forcing in a single-stage HPT and the lessons learned in that study.

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

Numerical investigation of high-pressure turbine blade tip flows: analysis of aerodynamics

2011 ◽  
Vol 225 (7) ◽  
pp. 940-953 ◽  
Author(s):  
P Vass ◽  
T Arts
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Three Dimensional ◽  
Suction Side ◽  
Point Of View ◽  
Internal Cooling ◽  
High Pressure Turbine ◽  
Blade Tip ◽  
Design Solutions ◽  
Exit Mach Number

The current contribution reports on the validation and analysis of three-dimensional computational results of the flow around four distinct high-pressure turbine blade tip geometries (TG1, 2, 3, and 4 hereinafter), taking into account the effect of the entire internal cooling setup inside the blade, at design exit Mach number: M = 0.8, and high exit Reynolds number: Re C = 900 000. Three of the four geometries represent different tip design solutions – TG1: full squealer rim; TG2: single squealer on the suction side; TG3: partial suction and pressure side squealer, and one (TG4) models TG1 in worn condition. This article provides a comparison between the different geometries from the aerodynamic point of view, analyses the losses, and evaluates the distinct design solutions. An assessment of the effect of the uneven rubbing of the blade tip was performed as well. TG1 was found to be the top performer followed by TG3 and TG2. According to the investigations, the effect of rubbing increased the kinetic loss coefficient by 10–15 per cent.

Multi-Disciplinary Analyses for the Design of a High Pressure Turbine Blade Tip

2016 ◽  
Author(s):  
Stefano Caloni ◽  
Shahrokh Shahpar
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Turbine Blades ◽  
Suction Side ◽  
High Pressure Turbine ◽  
Cooling Flow ◽  
Barrier Coating ◽  
Flow Features ◽  
Internally Cooled ◽  
Blade Tip

The design of a high pressure turbine blade is a challenging task requiring multiple disciplines to be solved simultaneously. Most recently, conjugate analyses are being developed to tackle such a problem; they are able to resolve both the fluid dynamics in a turbine passage and the thermal distribution in the solid part of the component. In this paper, the in-house Hydra CFD solver is used to analyse a high pressure shroudless turbine blade for a modern jet engine. The turbine is internally cooled and a Thermal Barrier Coating (TBC) is applied on the aerofoil surface. The coupling technique used at the interface in the presence of the TBC is described. The flow features at the tip of the turbine blade are the main focus of this study. Four different tip configurations are analysed. A flat tip and a squealer tip are chosen as reference designs; however the effects of opening the Trailing Edge (TE) on the Suction Side (SS) and the Pressure Side (PS) are also investigated. Both a cooled and an uncooled configuration of the turbine blade are analysed and the effect of the cooling flow on the over tip leakage is studied. Finally, conjugate analyses for the cooled turbine blades are used to predict the temperature reached by the different tip designs. The design with an opened TE on the SS shows a significant aerodynamic improvement over the others without increasing the temperature the tip has to withstand in operation.

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.

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

Heat Transfer Measurements and Predictions for the Vane and Blade of a Rotating High-Pressure Turbine Stage

2003 ◽  
Author(s):  
Charles W. Haldeman ◽  
Michael G. Dunn
Keyword(s):  
Heat Transfer ◽  
Boundary Layer ◽  
High Pressure ◽  
Experimental Results ◽  
Current Paper ◽  
Blade Surface ◽  
Turbine Stage ◽  
High Pressure Turbine ◽  
Transfer Data ◽  
Blade Tip

This paper describes heat-transfer measurements and predictions obtained for the vane and blade of a rotating high-pressure turbine stage. The measurements were obtained with the stage operating at design corrected conditions. A previous paper described the aerodynamics and the blade midspan location heat-transfer data and compared these experimental results with predictions. The intent of the current paper is to concentrate on the measurements and predictions for the 20%, 50%, and 80% span locations on the vane, the vane inner and outer endwall, the 20% and 96% span location on the blade, the blade tip (flat tip), and the stationary blade shroud. Heat-transfer data obtained at midspan for three different TBC coated vanes (fine, medium and coarse) are also presented. Boundary-layer heat transfer predictions at the off-midspan locations are compared with the measurements for both the vane and the blade. The results of a STAR-CD 3D prediction are compared with the 20% and 96% span results for the blade surface. Predictions are not available for comparison with the tip and shroud experimental results.

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.

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
Sign in / Sign up

Export Citation Format

Share Document