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.

Aeroengine High Pressure Turbine Blade Cooling System Concept

2013 ◽  
Author(s):  
S. Riznyk ◽  
A. Artushenko
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Cooling System ◽  
Systems Design ◽  
Experimental Investigations ◽  
Test Rig ◽  
High Pressure Turbine ◽  
Cooling Systems ◽  
Cooling Channels ◽  
Blade Cooling

Aeroengine high-pressure turbine (HPT) is the key engine component. HPT must withstand high inlet temperatures and mechanical loads providing necessary level of efficiency. To achieve these objectives effective and complex blade cooling systems (internal convective and film cooling) are used in HPT design. Methodology of effective HPT blade cooling system design, numerical and experimental investigations are described in the article. Different HPT cooling systems are considered: internal convective “serpentine” schemes with ribs in channels and wall-cooled system. Two types of film cooling channels (round and shaped) are used in HPT blade cooling systems design. The role of thermal barrier coating is described. Experimental test rig was designed and manufactured to define the heat transfer coefficients and hydraulic parameters for HPT blade with wall-cooled system cooling channels. The results obtained on this test rig were used to determine boundary conditions and temperature fields in advanced HPT blade with wall-cooled cooling system. Numerical and experimental results obtained in HPT blade cooling system research are presented in the article.

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.

scholarly journals Optimization of Gas Turbine Blade Cooling System

2019 ◽  
Vol 8 (11) ◽  
pp. 4176-4181
Keyword(s):  
Turbine Blade ◽  
Gas Turbines ◽  
Cooling System ◽  
Turbine Blades ◽  
Inlet Temperature ◽  
3D Design ◽  
Turbine Blade Cooling ◽  
Pin Fin ◽  
Blade Cooling ◽  
Industrial Sectors

At present, Gas turbines play an essential responsibility in different areas such as jet, generating power and various commercial and industrial sectors. Melting point of the turbine blade may causes the hotness levels which go rapidly raise. Likewise, heavy crack may cause because of Turbine Inlet Temperature (TIT) at turbine blades for the period of expansion procedure of turbine sector. Hence, a highly developed blade cooling system is required for safe operation of turbines. The proposed system deals with the serpentine rip - roughened passage with micro pin fin cooling system and it has been analyzed corresponding to serpentine cooling system. It increases the heat transfer enhancement. Therefore, very warm gases in and around the turbine blade may have a stream temperature at 1500K. On the other side, cool air disclosed to the blade core duct and an entry temperature may have been 650K. The proposed systems with 2D and 3D model were developed by using CATIA. The 3D design is then analyzed using CFD. Further, the corresponding results of serpentine rip - roughened passage and micro pin fin arrangement in serpentine rip-roughened passage are compared and the details are presented.

scholarly journals The Influence of Pressure Pulses to an Innovative Turbine Blade Film Cooling System

1998 ◽  
Author(s):  
Siegfried Moser ◽  
Herbert Jericha ◽  
Jakob Woisetschläger ◽  
Arno Gehrer ◽  
Werner Reinalter
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Cooling System ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Pressure Waves ◽  
Turbine Blade Cooling ◽  
Blade Cooling ◽  
Wall Film ◽  
Turbine Vane

The evolution of increasing turbine inlet temperature has led to the necessity of full-coverage film cooling for the first turbine vane and blade. This paper deals with the investigation of the aerodynamic behaviour of the transonic wall film on a leading edge with and without leading edge pressure waves. Here these films are used for turbine blade cooling. The pressure waves are produced with a rotating „Pressure Wave Generator“. The numerical simulations have been realised with a commercial CFD-Program. The experimental data were obtained in a linear cascade.

Comparison of Two Methods for the Sensitivity Analysis of a One-Dimensional Cooling Flow Network of a High-Pressure-Turbine Blade

2020 ◽  
Author(s):  
Barbara Fiedler ◽  
Yannick Muller ◽  
Matthias Voigt ◽  
Ronald Mailach
Keyword(s):  
High Pressure ◽  
Population Size ◽  
Turbine Blade ◽  
Cooling System ◽  
Design Parameters ◽  
Stochastic Methods ◽  
System Behavior ◽  
High Pressure Turbine ◽  
Cooling Systems ◽  
Cooling Flow

Abstract Efficient cooling of the thermally extremely loaded high-pressure-turbine blades and vanes is required to ensure acceptable service life. The development of cooling systems is a high-dimensional problem, since it includes multiple design parameters. The integration of stochastic methods into the design phase contributes to a better understanding of the complex interaction between the design parameters and the system behavior and thus helps designing more efficient cooling systems. Therefore, the present study compares two stochastic methods, the Elementary-Effect method by Morris and the Coefficient-of-Importance, for the quantification of the sensitivity of the cooling flow to geometric variations using a 1D flow network of a high-pressure-turbine blade. As test case a multipass cooling system with rib-roughened walls is investigated. To achieve geometrically meaningful variations a new parametrization, called Harmonic-Spline-Deformation, is developed and first time presented in the paper at hand. The parametrization proves to be universally applicable to the components of a cooling system and facilitates the interpretation of the physical relationships between variables and resulting system behavior. For a sufficiently large population size, both methods show good agreement in quantifying the importance of design parameters regarding their effect on the cooling flow. The Coefficient-of-Importance, however, proves to be more stable against a decreasing population size and more robust against defects in the population.

Aerothermal Effect of Cavity Welding Beads on a Transonic Squealer Tip

2020 ◽  
Author(s):  
Joao Vieira ◽  
John Coull ◽  
Peter Ireland ◽  
Eduardo Romero
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Turbine Blades ◽  
Reynolds Numbers ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Aerodynamic Loss ◽  
Blade Tip ◽  
Mass Flow Rates

Abstract High pressure turbine blade tips are critical for gas turbine performance and are sensitive to small geometric variations. For this reason, it is increasingly important for experiments and simulations to consider real geometry features. One commonly absent detail is the presence of welding beads on the cavity of the blade tip, which are an inherent by-product of the blade manufacturing process. This paper therefore investigates how such welds affect the Nusselt number, film cooling effectiveness and aerodynamic performance. Measurements are performed on a linear cascade of high pressure turbine blades at engine realistic Mach and Reynolds numbers. Two cooled blade tip geometries were tested: a baseline squealer geometry without welding beads, and a case with representative welding beads added to the tip cavity. Combinations of two tip gaps and several coolant mass flow rates were analysed. Pressure sensitive paint was used to measure the adiabatic film cooling effectiveness on the tip, which is supplemented by heat transfer coefficient measurements obtained via infrared thermography. Drawing from all of this data, it is shown that the weld beads have a generally detrimental impact on thermal performance, but with local variations. Aerodynamic loss measured downstream of the cascade is shown to be largely insensitive to the weld beads.

Modeling and Impact of High-pressure Turbine Blade Trailing Edge Film Cooling Hole Variations

2020 ◽  
Author(s):  
Jan Kamenik ◽  
David J. Toal ◽  
Andy Keane ◽  
Lars Högner ◽  
Marcus Meyer ◽  
...  
Keyword(s):  
High Pressure ◽  
Turbine Blade ◽  
Film Cooling ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Hole

Development of the Advanced Closed-Cycle Gas Turbine System

2001 ◽  
Author(s):  
Miki Koyama ◽  
Toshio Mimaki
Keyword(s):  
Film Cooling ◽  
Cooling System ◽  
Early Stage ◽  
Turbine Blades ◽  
Thermal Conditions ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Turbine Inlet Temperature ◽  
Turbine Blade Cooling ◽  
Turbine Inlet

This aims to put the fruits of the R&D; “The Hydrogen Combustion Turbine” in WE-NET Phase I Program(1993-1998) to practical use at an early stage. The topping regenerating cycle was selected as the optimum cycle, with energy efficiency expected to be more than 60%(HHV) under the conditions of the turbine inlet temperature of 1973K(1700°C) and the pressure of 4.8MPa,in it. • As the turbine inlet temperature and pressure increase, issues to be resolved include the amount of NOx emissions and the durability of super alloys for turbine blades under such thermal conditions. In this respect, the development of the highly efficient methane-oxygen combustion technology, the turbine blade cooling technology, and the ultrahigh-temperature materials including thermal barrier coatings is being carried out. • In 1999, the results made it clear that there are little error among the three analytic programs used to verify the system efficiency, it was verified that the burning rate was going to arrive at over 98% from the methane-oxygen combustion test (under the atmospheric pressure). And the type of vane “Film cooling plus recycle type with internal cooling system” was selected as the most suitable vane.

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.

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