Thermal Analysis of a Turbine Blade: Effect of Film Cooling and Internal Convective Cooling

2015 ◽  
Author(s):  
S. Sarkar ◽  
P. Gupta
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Suction Surface ◽  
Transfer Condition ◽  
Convective Cooling ◽  
Cooling Passage

Advanced gas turbines are designed to operate at increasingly higher inlet temperature that poses a greater challenge to the designer for more effective blade cooling strategies. In this paper, a generic high-pressure turbine (HPT) blade of a gas turbine, which is cooled by film cooling in conjunction with internal convective cooling, has been analysed by solving Navier-Stokes and energy equations. The intricate internal cooling passages and a series of holes on the suction surface are considered for the simulations. Large numbers of cell in different zones are used to truly replace the blade with cooling holes and the internal cooling passage. The CFD analysis with conjugate heat transfer condition is accomplished by Fluent, version 6.3. A detailed discussion has been made regarding the aerodynamics and heat transfer. In brief, the suction surface is well protected by film cooling, whereas, the pressure surface demands some additional protection for a longer life. The leading edge is under the metallurgical limit because of internal cooling for the present configuration.

Computational Analysis of Side-Wall Heat Transfer of a Turbine Blade Internal Cooling Passage With Truncated Ribs on Opposite Walls

2012 ◽  
Author(s):  
Shian Li ◽  
Gongnan Xie ◽  
Bengt Sundén ◽  
Weihong Zhang
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Thermal Stresses ◽  
Side Wall ◽  
Leading Edge ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Wall Heat Transfer ◽  
Cooling Passage ◽  
Internal Cooling Passage

A problem involved in the increase of the turbine inlet temperature of gas turbine engine is the failure of material because of excessive thermal stresses. This requires cooling methods to withstand the increase of the inlet temperature. Rib turbulators are often used in the mid-section of internal cooling ducts to augment the heat transfer from blade wall to the coolant. This study numerically investigates side-wall heat transfer of a rectangular passage with the leading/trailing walls being roughened by staggered ribs whose length is less than the passage width. Such a passage corresponds to the internal cooling passage near the leading edge of a turbine blade. The inlet Reynolds number is ranging from 12,000 to 60,000. The detailed 3D fluid flow and heat transfer over the side-wall are presented. The overall performances of several ribbed passages are evaluated and compared. It is found that the side-wall heat transfer coefficients of the passage with truncated (continuous) ribs on opposite walls are about 20%–27% (28%–43%) higher than those of a passage without ribs, while the pressure loss could be reduced compared to a passage with continuous ribs. It is suggested that the usage of truncated ribs is a suitable way to augment the side-wall heat transfer and improve the flow structure near the leading edge.

Unsteady simulations of migration and deposition of fly-ash particles in the first-stage turbine of an aero-engine

2021 ◽  
pp. 1-21
Author(s):  
Z. Hao ◽  
X. Yang ◽  
Z. Feng
Keyword(s):  
Fly Ash ◽  
Film Cooling ◽  
Particle Deposition ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Velocity Model ◽  
Operating Conditions ◽  
Inlet Temperature ◽  
Internal Cooling ◽  
Aero Engine

Abstract Particulate deposits in aero-engine turbines change the profile of blades, increase the blade surface roughness and block internal cooling channels and film cooling holes, which generally leads to the degradation of aerodynamic and cooling performance. To reveal particle deposition effects in the turbine, unsteady simulations were performed by investigating the migration patterns and deposition characteristics of the particle contaminant in a one-stage, high-pressure turbine of an aero-engine. Two typical operating conditions of the aero-engine, i.e. high-temperature take-off and economic cruise, were discussed, and the effects of particle size on the migration and deposition of fly-ash particles were demonstrated. A critical velocity model was applied to predict particle deposition. Comparisons between the stator and rotor were made by presenting the concentration and trajectory of the particles and the resulting deposition patterns on the aerofoil surfaces. Results show that the migration and deposition of the particles in the stator passage is dominated by the flow characteristics of fluid and the property of particles. In the subsequential rotor passage, in addition to these factors, particles are also affected by the stator–rotor interaction and the interference between rotors. With higher inlet temperature and larger diameter of the particle, the quantity of deposits increases and the deposition is distributed mainly on the Pressure Side (PS) and the Leading Edge (LE) of the aerofoil.

Experimental Evaluations of the Relative Contributions to Overall Effectiveness in Turbine Blade Leading Edge Cooling

2018 ◽  
Author(s):  
Carol E. Bryant ◽  
Connor J. Wiese ◽  
James L. Rutledge ◽  
Marc D. Polanka
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Leading Edge ◽  
Internal Surface ◽  
Internal Cooling ◽  
Transfer Coefficients ◽  
Impingement Cooling ◽  
External Cooling ◽  
Turbine Component ◽  
Overall Effectiveness

Gas turbine hot gas path components are protected through a combination of internal cooling and external film cooling. The coolant typically travels through internal passageways, which may involve impingement on the internal surface of a turbine component, before being ejected as film cooling. Internal cooling effects have been studied in facilities that allow measurement of heat transfer coefficients within models of the internal cooling paths, with large heat transfer coefficients generally desirable. External film cooling is typically evaluated through measurements of the adiabatic effectiveness and its effect on the external heat transfer coefficient. Efforts aimed at improving cooling are often focused on either only the internal cooling or the film cooling; however, the common coolant flow means the internal and external cooling schemes are linked and the coolant holes themselves provide another convective path for heat transfer to the coolant. Recently, measurements of overall cooling effectiveness using matched Biot number turbine component models allow evaluation of the nondimensional wall temperature achieved for the fully cooled component. However, the relative contributions of internal cooling, external cooling, and convection within the film cooling holes is not well understood. Large scale, matched Biot number experiments, complemented by CFD simulations, were performed on a fully film cooled cylindrical leading edge model to evaluate the effects of various alterations in the cooling design on the overall effectiveness. The relative influence of film cooling and cooling within the holes was evaluated by selectively disabling individual holes and quantifying how the overall effectiveness changed. Several internal impingement cooling schemes in addition to a baseline case without impingement cooling were also tested. In general, impingement cooling is shown to have a negligible influence on the overall effectiveness in the showerhead region. This indicates that the cost and pressure drop penalties for implementing impingement cooling may not be compensated by an increase in thermal performance. Instead, the internal cooling provided by convection within the holes themselves was shown, along with external film cooling, to be a dominant contribution to the overall cooling effectiveness. Indeed, the numerous holes within the showerhead region impede the ability of internal surface cooling schemes to influence the outside surface temperature. The results of this research may allow improved focus of future efforts on the forms of cooling with the greatest potential to improve cooling performance.

A New Test Facility to Investigate Film Cooling on a Non-Axisymmetric Contoured Turbine Endwall: Part I — Introduction and Aerodynamic Measurements

2015 ◽  
Author(s):  
Franz Puetz ◽  
Johannes Kneer ◽  
Achmed Schulz ◽  
Hans-Joerg Bauer
Keyword(s):  
Heat Transfer ◽  
Flow Field ◽  
Pressure Distribution ◽  
Film Cooling ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Test Facility ◽  
Endwall Contouring

An increased demand for lower emission of stationary gas turbines as well as civil aircraft engines has led to new, low emission combustor designs with less liner cooling and a flattened temperature profile at the outlet. As a consequence, the heat load on the endwall of the first nozzle guide vane is increased. The secondary flow field dominates the endwall heat transfer, which also contributes to aerodynamic losses. A promising approach to reduce these losses is non-axisymmetric endwall contouring. The effects of non-axisymmetric endwall contouring on heat transfer and film cooling are yet to be investigated. Therefore, a new cascade test rig has been set up in order to investigate endwall heat transfer and film cooling on both a flat and a non-axisymmetric contoured endwall. Aerodynamic measurements that have been made prior to the upcoming heat transfer investigation are shown. Periodicity and detailed vane Mach number distributions ranging from 0 to 50% span together with the static pressure distribution on the endwall give detailed information about the aerodynamic behavior and influence of the endwall contouring. The aerodynamic study is backed by an oil paint study, which reveals qualitative information on the effect of the contouring on the endwall flow field. Results show that the contouring has a pronounced effect on vane and endwall pressure distribution and on the endwall flow field. The local increase and decrease of velocity and the reduced blade loading towards the endwall is the expected behavior of the 3d contouring. So are the results of the oil paint visualization, which show a strong change of flow field in the leading edge region as well as that the contouring delays the horse shoe vortex hitting the suction side.

High Rotation Number Effect on Heat Transfer in a Triangular Channel With 45°, Inverted 45°, and 90° Ribs

2009 ◽  
Author(s):  
Yao-Hsien Liu ◽  
Michael Huh ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Enhancement ◽  
Rotation Number ◽  
Leading Edge ◽  
Stationary Condition ◽  
Internal Cooling ◽  
Transfer Enhancement ◽  
Triangular Channel ◽  
Buoyancy Parameter ◽  
Cooling Passage

Heat transfer and pressure drop have been experimentally investigated in an equilateral triangular channel (Dh = 1.83cm), which can be used to simulate the internal cooling passage near the leading edge of a gas turbine blade. Three different rib configurations (45°, inverted 45°, and 90°) were tested at four different Reynolds numbers (10000–40000), each with five different rotational speeds (0–400 rpm). The rib pitch-to-height (P/e) ratio is 8 and the height-to-hydraulic diameter (e/Dh) ratio is 0.087 for every rib configuration. The rotation number and buoyancy parameter achieved in this study were 0–0.58 and 0–2.3, respectively. Both the rotation number and buoyancy parameter have been correlated to predict the rotational heat transfer in the ribbed equilateral triangular channel. For the stationary condition, staggered 45° angled ribs show the highest heat transfer enhancement. However, staggered 45° angled ribs and 90° ribs have the higher comparable heat transfer enhancement at rotating condition near the blade leading edge region.

Experimental and Numerical Conjugate Investigation of the Blowing-Ratio Influence on the Showerhead Cooling Efficiency

1998 ◽  
Author(s):  
Dieter E. Bohn ◽  
Volker J. Becker ◽  
Karsten A. Kusterer ◽  
Agnes U. Rungen
Keyword(s):  
Heat Transfer ◽  
Gas Turbines ◽  
Guide Vane ◽  
Leading Edge ◽  
Suction Side ◽  
Computer Code ◽  
Inlet Temperature ◽  
Cooling Efficiency ◽  
Blowing Ratio ◽  
Turbine Guide Vane

This paper presents the experimental investigation of the flow and the numerical analysis of the flow and heat transfer in a turbine guide vane with showerhead cooling for two different blowing ratios. The aerodynamic results are compared with those of the experiments. Starting with a showerhead design of two rows of ejection holes, two additional rows have to be used in an enhanced design due to hot gas contact in the leading edge area. Thus, the cooling gas mass flow is doubled when keeping the blowing ratio constant at m = 1.5. Lowering the amount of cooling gas needed whilst still guaranteeing sufficient cooling is the motivation for the analysis of the influence of a lower blowing ratio on the cooling efficiency. The investigated blowing ratios are m = 1.5 and m = 1.0. The experiments are conducted using a non-intrusive LDA technique. The numerical results are gained with a conjugate heat transfer and flow computer code that has been developed at the Institute of Steam and Gas Turbines. The results show that the blowing ratio has to be chosen carefully as the leading edge flow pattern and the heat transfer are strongly influenced by the blowing ratio. Lower blowing ratios lead to a better attachment of the cooling film and thus they hardly disturb the main flow. With the lower blowing ratio, the material temperature increases up to 1.5% of the total inlet temperature in the leading edge area on the pressure side, whereas it decreases locally for about 0.8% for the lower blowing ratio on the suction side. This is due to the enhanced attachment of the cooling gas film.

Effects of Jetting Orifice Geometry Parameters and Channel Reynolds Number on Bended Channel Cooling for a Novel Internal Cooling Structure

2019 ◽  
Author(s):  
Wei He ◽  
Qinghua Deng ◽  
Juan He ◽  
Tieyu Gao ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Reynolds Number ◽  
Film Cooling ◽  
Leading Edge ◽  
Jet Impingement ◽  
Suction Side ◽  
Outer Wall ◽  
Internal Cooling ◽  
Cooling Effectiveness ◽  
Blade Leading Edge

Abstract A novel internal cooling structure has been raised recently to enhance internal cooling effectiveness and reduce coolant requirement without using film cooling. This study mainly focuses on verifying the actual cooling performance of the structure and investigating the heat transfer mechanism of the leading edge part of the structure, named bended channel cooling. The cooling performances of the first stage of GE-E3 turbine with three different blade leading edge cooling structures (impingement cooling, swirl cooling and bended channel cooling) were simulated using the conjugate heat transfer method. Furthermore, the effects of jetting orifice geometry and channel Reynolds number were studied with simplified models to illustrate the flow and heat transfer characteristics of the bended channel cooling. The results show that the novel internal cooling structure has obvious advantages on the blade leading edge and suction side under operating condition. The vortex core structure in the bended channel depends on orifice width, but not channel Reynolds number. With the ratio of orifice width to outer wall thickness smaller than a critical value of 0.5, the coolant flows along the external surface of the channel in the pattern of “inner film cooling”, which is pushed by centrifugal force and minimizes the mixing with spent cooling air. Namely, the greatly organized coolant flow generates higher cooling effectiveness and lower coolant demand. Both the Nusselt number on the channel surfaces and total pressure loss increase significantly when the orifice width falls or channel Reynolds increases, but the wall jet impingement distance appears to be less influential.

scholarly journals Experimental Heat Transfer and Bulk Air Temperature Measurements for a Multipass Internal Cooling Model With Ribs and Bleed

2000 ◽  
Author(s):  
Douglas Thurman ◽  
Philip Poinsatte
Keyword(s):  
Heat Transfer ◽  
Turbine Blade ◽  
Air Temperature ◽  
Test Section ◽  
Film Cooling ◽  
External Pressure ◽  
Temperature Measurements ◽  
Internal Cooling ◽  
Model Average ◽  
Cooling Passage

An experimental study was made to obtain heat transfer and air temperature data for a simple 3-leg serpentine test section that simulates a turbine blade internal cooling passage with trip strips and bleed holes. The objectives were to investigate the interaction of ribs and various bleed conditions on internal cooling and to gain a better understanding of bulk air temperature in an internal passage. Steady state heat transfer measurements were obtained using a transient technique with thermochromic liquid crystals. Trip strips were attached to one wall of the test section and were located either between or near the bleed holes. The bleed holes, used for film cooling, were metered to simulate the effect of external pressure on the turbine blade. Heat transfer enhancement was found to be greater for ribs near bleed holes compared to ribs between holes, and both configurations were affected slightly by bleed rates upstream. Air temperature measurements were taken at discreet locations along one leg of the model. Average bulk air temperatures were found to remain fairly constant along one leg of the model.

scholarly journals Development and Aerothermal Investigation of Integrated Combustor Vane Concept

2015 ◽  
Vol 138 (1) ◽  
Author(s):  
Simon Jacobi ◽  
Budimir Rosic
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Gas Turbines ◽  
High Speed ◽  
Leading Edge ◽  
Cfd Simulations ◽  
Pressure Loss Coefficient ◽  
Novel Concept ◽  
Nozzle Guide ◽  
Stage Efficiency

This paper presents the development and aerothermal investigation of the integrated combustor vane concept for power generation gas turbines with individual can combustors. In this novel concept, first introduced in 2010, the conventional nozzle guide vanes (NGVs) are removed and flow turning is achieved by vanes that extend the combustor walls. The concept was developed using the in-house computational fluid dyanamics (CFD) code TBLOCK. Aerothermal experiments were conducted using a modular high-speed linear cascade, designed to model the flow at the combustor–vane interface. The facility is comprised of two can combustor transition ducts and either four conventional vanes (CVs) or two integrated vanes (IVs). The experimental study validates the linear CFD simulations of the IV development. Annular full-stage CFD simulations, used to evaluate aerodynamics, heat transfer, and stage efficiency, confirm the trends of the linear numerical and experimental results, and thus demonstrate the concept's potential for real gas turbine applications. Results show a reduction of the total pressure loss coefficient at the exit of the stator vanes by more than 25% due to a reduction in profile and endwall loss. Combined with an improved rotor performance demonstrated by unsteady stage simulations, these aerodynamic benefits result in a gain in stage efficiency of above 1%. A distinct reduction in heat transfer coefficient (HTC) levels on vane surfaces, on the order of 25–50%, and endwalls is observed and attributed to an altered state of boundary layer (BL) thickness. The development of IV's endwall- and leading edge (LE)-cooling geometry shows a superior surface coverage of cooling effectiveness, and the cooling requirements for the first vane are expected to be halved. Moreover, by halving the number of vanes, simplifying the design and eliminating the need for vane LE film cooling, manufacturing and development costs can be significantly reduced.

scholarly journals Cooling Characteristic of a Wall Jet for Suppressing Crossflow Effect under Conjugate Heat Transfer Condition

Aerospace ◽  
2022 ◽  
Vol 9 (1) ◽  
pp. 29
Author(s):  
Qinghua Deng ◽  
Huihui Wang ◽  
Wei He ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Conjugate Heat Transfer ◽  
Flow Direction ◽  
Leading Edge ◽  
Target Surface ◽  
Suction Side ◽  
Wall Jet ◽  
Internal Cooling ◽  
Transfer Condition ◽  
Jet Cooling

The leading edge is the critical portion for a gas turbine blade and is often insufficiently cooled due to the adverse effect of Crossflow in the cooling chamber. A novel internal cooling structure, wall jet cooling, can suppress Crossflow effect by changing the coolant flow direction. In this paper, the conjugate heat transfer and aerodynamic characteristics of blades with three different internal cooling structures, including impingement with a single row of jets, swirl cooling, and wall jet cooling, are investigated through RANS simulations. The results show that wall jet cooling combines the advantages of impingement cooling and swirl cooling, and has a 19–54% higher laterally-averaged overall cooling effectiveness than the conventional methods at different positions on the suction side. In the blade with wall jet cooling, the spent coolant at the leading edge is extracted away through the downstream channels so that the jet could accurately impinge the target surface without unnecessary mixing, and the high turbulence generated by the separation vortex enhances the heat transfer intensity. The Coriolis force induces the coolant air to adhere to the pressure side’s inner wall surface, preventing the jet from leaving the target surface. The parallel cooling channels eliminate the common Crossflow effect and make the flow distribution of the orifices more uniform. The trailing edge outlet reduces the entire cooling structure’s pressure to a low level, which means less penalty on power output and engine efficiency.

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