Numerical and Experimental Investigations of the Influence of Different Swirl Ratios on the Temperature Streak Development in a 4-Stage Turbine

2000 ◽  
Author(s):  
Dieter Bohn ◽  
Harald Funke ◽  
Tom Heuer ◽  
Jürg Bütikofer
Keyword(s):  
Gas Turbines ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Temperature Inhomogeneity ◽  
Radial Temperature ◽  
Blade Cooling ◽  
Hot Gas ◽  
Segregation Effect ◽  
Temperature Equalization ◽  
Hot Streak

In the development of modern gas turbines the increase in the turbine inlet temperature is restricted by the need for cooling the first stages of the turbine. In addition, the flow leaving the combustor is thermally inhomogeneous. Since the blade cooling has to be designed for the actual local hot gas temperatures, it is important to know how these temperature inhomogeneities develop and attenuate inside the multistage flow passage. In this investigation the development of a circumferential and a radial temperature inhomogeneity inside a 4-stage turbine is analyzed at three different swirl ratios. Since the experimental setup allows a circumferential temperature streak, a radial temperature streak has also been applied at different swirl ratios to the same geometrical configuration for a numerical investigation. The first stage has a significant impact on the attenuation and change in form of a circumferential temperature streak depending on the swirl. For the radial streak the hot streak segregation effect can be eliminated by increasing the swirl. Consequently, the temperature equalization process is weakened.

Numerical and Experimental Investigations on the Flow in a 4-Stage Turbine With Special Focus on the Development of a Radial Temperature Streak

1999 ◽  
Author(s):  
Dieter Bonn ◽  
Harald Funke ◽  
Jochen Gier
Keyword(s):  
Gas Turbines ◽  
Experimental Investigations ◽  
Recent Experimental Data ◽  
Finite Volume Scheme ◽  
Inlet Temperature ◽  
Special Focus ◽  
Test Rig ◽  
Reference Case ◽  
Mixing Processes ◽  
Radial Temperature

In the development of modern gas turbines the increase of the turbine inlet temperature is restricted by the need to cool the first stages of the turbine. In addition the flow leaving the combustor is thermally inhomogeneous. Since the blade cooling has to be designed for the actual local hot gas temperatures, it is important to know how these temperature inhomogeneities develop and attenuate inside the multistage flow passage. In this investigation the flow inside a 4-stage turbine, which is set up in a test rig at the Institute of Steam and Gas Turbines, Aachen University of Technology, is calculated with a state-of-the-art fully three-dimensional Navier-Stokes solver based on an accurate finite volume scheme. The stator and rotor rows are coupled via mixing planes. The turbine is a scaled down original turbine with realistic axial gaps. The homogeneous reference case is qualified by comparison to recent experimental data gathered at the test rig. Therefore, the flow is extensively measured at several locations. In a second step a radial temperature streak is set at the inlet for the same point of operation. The results show the development of the temperature streak through the four stages. With this information the underlying mixing processes are described and analysed. It is found that the hot streak segregation effect is present in all four stages.

Evaluation of the Rotor Temperature Distribution of an Automotive Turbocharger Under Hot Gas Conditions Including Indirect Experimental Validation

2021 ◽  
Vol 143 (3) ◽  
Author(s):  
Christopher Zeh ◽  
Ole Willers ◽  
Thomas Hagemann ◽  
Hubert Schwarze ◽  
Jörg Seume
Keyword(s):  
Temperature Distribution ◽  
Heat Flow ◽  
Internal Combustion Engines ◽  
Fluid Film ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Hot Gas ◽  
Experimental Identification ◽  
Fluid Film Bearings ◽  
Validation Parameters

Abstract While turbocharging is a key technology for improving the performance and efficiency of internal combustion engines, the operating behavior of the turbocharger is highly dependent on the rotor temperature distribution as it directly modifies viscosity and clearances of the fluid film bearings. Since a direct experimental identification of the rotor temperature of an automotive turbocharger is not feasible at an acceptable expense, a combination of numerical analysis and experimental identification is applied to investigate its temperature characteristic and level. On the one hand, a numerical conjugate heat transfer (CHT) model of the automotive turbocharger investigated is developed using a commercial CFD-tool and a bidirectional, thermal coupling of the CFD-model with thermohydrodynamic lubrication simulation codes is implemented. On the other hand, experimental investigations of the numerically modeled turbocharger are conducted on a hot gas turbocharger test rig for selected operating points. Here, rotor speeds range from 64.000 to 168.000 rpm. The turbine inlet temperature is set to 600 °C and the lubricant is supplied at a pressure of 300 kPa with 90 °C to ensure practically relevant boundary conditions. Comparisons of measured and numerically predicted local temperatures of the turbocharger components indicate a good agreement between the analyses. The calorimetrically determined frictional power loss of the bearings as well as the floating ring speed are used as additional validation parameters. Evaluation of heat flow of diabatic simulations indicates a high sensitivity of local temperatures to rotor speed and load. A cooling effect of the fluid film bearings is present. Consequently, results confirm the necessity of the diabatic approach to the heat flow analysis of turbocharger rotors.

A Numerical Study on Conjugate Heat Transfer for Supercritical CO2 Turbine Blade With Cooling Channels

2020 ◽  
Author(s):  
Akshay Khadse ◽  
Andres Curbelo ◽  
Ladislav Vesely ◽  
Jayanta S. Kapat
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Gas Turbines ◽  
Supercritical Co2 ◽  
Conjugate Heat Transfer ◽  
Maximum Temperature ◽  
Inlet Temperature ◽  
Cooling Channels ◽  
Blade Cooling ◽  
Turbine Inlet

Abstract The first stage of turbine in a Brayton cycle faces the maximum temperature in the cycle. This maximum temperature may exceed creep temperature limit or even melting temperature of the blade material. Therefore, it becomes an absolute necessity to implement blade cooling to prevent them from structural damage. Turbine inlet temperatures for oxy-combustion supercritical CO2 (sCO2) are promised to reach blade material limit in near future foreseeing need of turbine blade cooling. Properties of sCO2 and the cycle parameters can make Reynolds number external to blade and external heat transfer coefficient to be significantly higher than those typically experience in regular gas turbines. This necessitates evaluation and rethinking of the internal cooling techniques to be adopted. The purpose of this paper is to investigate conjugate heat transfer effects within a first stage vane cascade of a sCO2 turbine. This study can help understand cooling requirements which include mass flow rate of leakage coolant sCO2 and geometry of cooling channels. Estimations can also be made if the cooling channels alone are enough for blade cooling or there is need for more cooling techniques such as film cooling, impingement cooling and trailing edge cooling. The conjugate heat transfer and aerodynamic analysis of a turbine cascade is carried out using STAR CCM+. The turbine inlet temperature of 1350K and 1775 K is considered for the study considering future potential needs. Thermo-physical properties of this mixture are given as input to the code in form of tables using REFPROP database. The blade material considered is Inconel 718.

Evaluation of the Rotor Temperature Distribution of an Automotive Turbocharger Under Hot Gas Conditions Including Indirect Experimental Validation

2020 ◽  
Author(s):  
Christopher Zeh ◽  
Ole Willers ◽  
Thomas Hagemann ◽  
Hubert Schwarze ◽  
Joerg R. Seume
Keyword(s):  
Temperature Distribution ◽  
Heat Flow ◽  
Internal Combustion Engines ◽  
Fluid Film ◽  
Experimental Investigations ◽  
Inlet Temperature ◽  
Hot Gas ◽  
Experimental Identification ◽  
Fluid Film Bearings ◽  
Validation Parameters

Abstract While turbocharging is a key technology for improving the performance and efficiency of internal combustion engines, the operating behavior of the turbocharger is highly dependent on the rotor temperature distribution as it directly modifies viscosity and clearances of the fluid film bearings. Since a direct experimental identification of the rotor temperature of an automotive turbocharger is not feasible at an acceptable expense, a combination of numerical analysis and experimental identification is applied to investigate its temperature characteristic and level. On the one hand, a numerical conjugate heat transfer (CHT) model of the automotive turbocharger investigated is developed using a commercial CFD-tool and a bidirectional, thermal coupling of the CFD-model with thermohydrodynamic lubrication simulation codes is implemented. On the other hand, experimental investigations of the numerically modelled turbocharger are conducted on a hot gas turbocharger test rig for selected operating points. Here, rotor speeds range from 64.000 to 168.000 rpm. The turbine inlet temperature is set to 600°C and the lubricant is supplied at a pressure of 300 kPa with 90°C to ensure practically relevant boundary conditions. Comparisons of measured and numerically predicted local temperatures of the turbocharger components indicate a good agreement between the analyses. The calorimetrically determined frictional power loss of the bearings as well as the floating ring speed are used as additional validation parameters. Evaluation of heat flow of diabatic simulations indicates a high sensitivity of local temperatures to rotor speed and load. A cooling effect of the fluid film bearings is present. Consequently, results confirm the necessity of the diabatic approach to the heat flow analysis of turbocharger rotors.

scholarly journals Optimum Power Boosting of Gas Turbine Cycles With Refrigerated Inlet Air and Compressor Intercooling

1996 ◽  
Author(s):  
Mohand A. Ait-Ali
Keyword(s):  
Gas Turbine ◽  
Gas Turbines ◽  
Pressure Ratio ◽  
Inlet Temperature ◽  
Turbine Blade Cooling ◽  
High Pressure Ratio ◽  
Blade Cooling ◽  
Air Inlet ◽  
Turbine Inlet ◽  
Compressor Work

With or without turbine blade cooling, gas turbine cycles have consistently higher turbine inlet temperatures than steam turbine cycles. But this advantage is more than offset by the excessive compressor work induced by warm inlet temperatures, particularly during operation on hot summer days. Instead of seeking still higher turbine inlet temperatures by means of sophisticated blade cooling technology and high temperature-resistant blade materials, it is proposed to greatly increase the cycle net work and also improve thermal efficiency by decreasing the compressor work. This is obtained by using refrigerated inlet air and compressor intercooling to an extent which optimizes the refrigerated air inlet temperature and consequently the gas turbine compression ratio with respect to maximum specific net power. The cost effectiveness of this conceptual cycle, which also includes regeneration, has not been examined in this paper as it requires unusually high pressure ratio gas turbines and compressors, as well as high volumetric air flow rate and low temperature refrigeration equipment for which reliable cost data is not easily available.

Gas Turbines Blade Heat Transfer and Internal Swirl Cooling Flow Experimental Study Using Liquid Crystals and Three-Dimensional Stereo-Particle Imaging Velocimetry

2021 ◽  
pp. 1-31
Author(s):  
Daisy Galeana ◽  
Asfaw Beyene
Keyword(s):  
Heat Transfer ◽  
Liquid Crystals ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Convective Heat Transfer Coefficient ◽  
Swirl Flow ◽  
Inlet Temperature ◽  
Cooling Flow ◽  
Blade Cooling ◽  
The Impact

Abstract The challenging engineering intricacies related to improving efficiency of a gas turbine engine comes with the need to maximize the internal cooling of the turbine blade to withstand the high turbine inlet temperature. Understanding the fluid mechanics and heat transfer of internal blade cooling is therefore of paramount importance. This paper presents the impact of swirl cooling flow on the heat transfer of a gas turbine chamber to understand the mechanics of internal blade cooling. The focus is the continuous swirl flow that must be maintained via nonstop injection of tangential flow, whereby swirl flow is generated. The impact of swirl cooling flow variation considers the velocity fields measured using stereo particle image velocimetry, the wall temperature and the convective heat transfer coefficient measured by liquid crystals and system of infrared thermography. Flow behavior and heat transfer at three Reynolds numbers ranging from 7,000 to 21,000 and the average profiles of axial and radial, magnitudes of velocity, and Nusselt numbers are given to research the direct effects of the circular chamber shape. Heat transfer results are measured and collected continuously after the system is heat-soaked to the required temperature. As part of the results relatively low heat transfer rates were observed near the upstream end of the circular chamber, resulting from a low momentum swirl flow as well as crossflow effects. The Thermochromic Liquid Crystal heat transfer results exemplify how the Nu measured favorably at the midstream of the chamber and values decline downstream.

A Simple Method for the Prediction of Wall Temperatures in Gas Turbines

1978 ◽  
Author(s):  
D. Kretschmer ◽  
J. Odgers
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Gas Temperature ◽  
Empirical Formulas ◽  
Simple Method ◽  
Hot Gas ◽  
Manufacturing Tolerances ◽  
Factor Governing

The cited method predicts wall temperatures generally within an accuracy of ± 6 percent, The biggest single factor governing the wall temperature is shown to be the hot gas temperature. Other factors discussed are the effects of changes in inlet temperature, fuel types, the geometry of the film cooling devices and manufacturing tolerances. Empirical formulas are given for the prediction of effective temperatures within the various combustor zones. Some comparisons are made between predictions and measurements of wall temperatures over a range of operating conditions.

scholarly journals Experimental Assessment of Fiber Reinforced Ceramics for Combustor Walls

1997 ◽  
Author(s):  
D. Filsinger ◽  
S. Münz ◽  
A. Schulz ◽  
S. Wittig ◽  
G. Andrees
Keyword(s):  
High Temperature ◽  
Gas Turbine ◽  
Gas Turbines ◽  
Gas Flow ◽  
Ceramic Materials ◽  
Inlet Temperature ◽  
Test Duration ◽  
Fiber Reinforced ◽  
Hot Gas ◽  
Sic Composites

Experimental and theoretical work concerning the application of ceramic components in small high temperature gas turbines has been performed for several years. The significance of some non-oxide ceramic materials for gas turbines in particular is based on their excellent high temperature properties. The application of ceramic materials allows an increase of the turbine inlet temperature resulting in higher efficiencies and a reduction of pollution emissions. The inherent brittleness of monolithic ceramic materials can be virtually reduced by reinforcement with ceramic fibers leading to a quasi-ductile behavior. Unfortunately, some problems arise due to oxidation of these composite materials in the presence of hot gas flow containing oxygen. At the Motoren- und Turbinen Union, München GmbH, comprehensive investigations including strength, oxidation, and thermal shock tests of several materials that seemed to be appropriate for combustor liner applications were undertaken. As a result, C/C, SiC/SiC, and two C/SiC-composites coated with SiC, as oxidation protection, were chosen for examination in a gas turbine combustion chamber. To prove the suitability of these materials under real engine conditions, the fiber reinforced flame tubes were installed in a small gas turbine operating under varying conditions. The loading of the flame tubes was characterized by wall temperature measurements. The materials showed different oxidation behavior when exposed to the hot gas flow. Inspection of the C/SiC-composites revealed debonding of the coatings. The C/C- and the SiC/SiC-materials withstood the tests with a maximum cumulated test duration of 90 hours without damage.

scholarly journals Experimental and Numerical Investigations of the Aerodynamical Effects of Coolant Injection Through the Trailing Edge of a Guide Vane

1995 ◽  
Author(s):  
Dieter E. Bohn ◽  
Volker J. Becker ◽  
Klaus D. Behnke ◽  
Bernhard F. Bonhoff
Keyword(s):  
Gas Turbines ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Turbine Blade Cooling ◽  
Blade Cooling ◽  
Hot Gas ◽  
Flow Situation ◽  
Circumferential Distribution ◽  
Volume Method

Effective turbine blade cooling is necessary to enhance the efficiency of gas turbines. Usually the coolant is mainly ejected through the trailing edge of the vanes. In addition to the desired temperature reduction at the trailing edge there is a 3D-aerodynamical interaction between the hot gas and the coolant. The complex mechanisms of the mixture are a main problem in the numerical prediction of the flow situation in this region. This paper presents the experimental and numerical results of investigations of annular guide vanes. The experiments were conducted in a scaled turbine test rig. The mixing flow of coolant and hot gas was analyzed by measurement of the distribution of both velocity and turbulence very close to the trailing edge using a 2D-LDA measurement technique at different radial positions. The experimental results show that the radial and circumferential distribution of the coolant depends on the pressure gradient in both directions. Inside of the mixture region the turbulence was found to be anisotropic resulting in a non-symmetrical distribution of the coolant. For the numerical calculations a Navier-Stokes-Code was used. The numerical scheme works on the basis of an implicit finite volume method combined with a multi block technique. In order to simulate the aerodynamical effects near the injection slot of the vane it was nessessary to include the coolant flow inside the guide vane.

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.

Sign in / Sign up

Export Citation Format

Share Document