Thermal Analysis and Durability Design Strategies for Gas Turbine Airfoils

2006 ◽  
Author(s):  
F. J. Cunha ◽  
W. Abdel-Messeh ◽  
M. K. Chyu
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Primary Component ◽  
Coolant Flow ◽  
Design Strategies ◽  
Convective Cooling ◽  
Viscous Forces ◽  
Formidable Challenge ◽  
Durability Design ◽  
Bulk Scale

Aircraft propulsion engines, land-based power generation, and industrial machines have, as a primary component, the turbine as means to produce thrust or generate power. In the turbine section of the engine, airfoil components are subjected to extremely complex and damaging environments. The combination of high gas temperatures and pressures, strong gradients, abrupt geometry changes, viscous forces, rotational forces, and unsteady turbine vane/blade interactions, all combine to offer a formidable challenge in terms of turbine durability. Nevertheless, the ultimate goal is to maintain or even improve the highest level of turbine performance and simultaneously reduce the amount of cooling flow needed to achieve this end. As such, coolant flow is a penalty to the cycle and thermal efficiency. Cooling strategies are developed and presented to determine ways for coolant flow management. The main variables include film cooling configurations, and convective efficiency schemes to balance turbine airfoil thermal loads for target overall cooling effectiveness. The desired targets are determined by the turbine airfoil durability requirements of oxidation and fatigue on a local scale and for creep on the bulk scale. Emphasis is provided to the general modes of cooling including film cooling, impingement cooling, and convective cooling for different parts of the airfoil such as leading edge, mid-body, trailing edge, tip and endwalls. Convective cooling is presented in terms of fundamental cooling enhancements, such as turbulating trip strips and pedestals. Recent literature dealing with these topics is listed.

Effects of Sand Ingestion on the Blockage of Film-Cooling Holes

2006 ◽  
Author(s):  
W. S. Walsh ◽  
K. A. Thole ◽  
Chris Joe
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Flow Parameter ◽  
Pressure Ratio ◽  
Leading Edge ◽  
Coolant Flow ◽  
Aero Engines ◽  
Film Cooling Holes

Gas turbines are often subjected to conditions where dirt and sand are ingested into the engine during takeoffs and landings. Given most aero engines do not have filtration systems, particulates can be present in both the main gas path and coolant streams. Particulates can block coolant passages and film-cooling holes that lead to increased airfoil temperatures caused by reduced coolant available for a given pressure ratio across the cooling holes. This study investigated the effects of sand blockage on film-cooling holes placed in a leading edge coupon. The coupon was tested to determine the reduction in flow parameter for a range of pressure ratios, coolant temperatures, metal temperatures, number of cooling holes, sand amounts, and sand diameters. Depending upon conditions, blockages characterized by reduced coolant flow can be as high as 10%.

scholarly journals Modeling Film-Coolant Flow Characteristics at the Exit of Shower-Head Holes

2000 ◽  
Author(s):  
Vijay K. Garg
Keyword(s):  
Heat Transfer ◽  
Experimental Data ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Three Dimensional ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Coolant Flow ◽  
The Heat Transfer Coefficient

Abstract The coolant flow characteristics at the hole exits of a film-cooled blade are derived from an earlier analysis where the hole pipes and coolant plenum were also discretized. The blade chosen is the VKI rotor with three staggered rows of shower-head holes. The present analysis applies these flow characteristics at the shower-head hole exits. A multi-block three-dimensional Navier-Stokes code with Wilcox’s k-ω model is used to compute the heat transfer coefficient on the film-cooled turbine blade. A reasonably good comparison with the experimental data as well as with the more complete earlier analysis where the hole pipes and coolant plenum were also gridded is obtained. If the 1/7th power law is assumed for the coolant flow characteristics at the hole exits, considerable differences in the heat transfer coefficient on the blade surface, specially in the leading-edge region, are observed even though the span-averaged values of h match well with the experimental data. This calls for span-resolved experimental data near film-cooling holes on a blade for better validation of the code.

Influence of Film Cooling Unsteadiness on Turbine Blade Leading Edge Heat Flux

2012 ◽  
Vol 134 (7) ◽  
Author(s):  
James L. Rutledge ◽  
Paul I. King ◽  
Richard B. Rivir
Keyword(s):  
Heat Flux ◽  
Turbine Blade ◽  
Film Cooling ◽  
Flow Behavior ◽  
Leading Edge ◽  
Coolant Flow ◽  
Adiabatic Wall ◽  
Stagnation Line ◽  
Turbine Engine ◽  
Cooling Flow

Film cooling in the hot gas path of a gas turbine engine can protect components from the high temperature main flow, but it generally increases the heat transfer coefficient h partially offsetting the benefits in reduced adiabatic wall temperature. We are thus interested in adiabatic effectiveness η and h which are combined in a formulation called net heat flux reduction (NHFR). Unsteadiness in coolant flow may arise due to inherent unsteadiness in the external flow or be intentionally introduced for flow control. In previous work it has been suggested that pulsed cooling flow may, in fact, offer benefits over steady blowing in either improving NHFR or reducing the mass flow requirements for matched NHFR. In this paper we examine this hypothesis for a range of steady and pulsed blowing conditions. We use a new experimental technique to analyze unsteady film cooling on a semicircular cylinder simulating the leading edge of a turbine blade. The average NHFR with pulsed and steady film cooling is measured and compared for a single coolant hole located 21.5° downstream from the leading edge stagnation line, angled 20° to the surface and 90° to the streamwise direction. We show that for moderate blowing ratios at blade passing frequencies, steady film flow yields better NHFR. At higher coolant flow rates beyond the optimum steady blowing ratio, however, pulsed film cooling can be advantageous. We present and demonstrate a prediction technique for unsteady blowing at frequencies similar to the blade passing frequency that only requires the knowledge of steady flow behavior. With this important result, it is possible to predict when pulsing would be beneficial or detrimental.

Influence of Large Wake Disturbances Shed From the Combustor Wall on the Leading Edge Film Cooling

2013 ◽  
Author(s):  
Cosimo Maria Mazzoni ◽  
Christian Klostermeier ◽  
Budimir Rosic
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Thermal Load ◽  
Leading Edge ◽  
Coolant Flow ◽  
Test Case ◽  
Trailing Edge ◽  
Industrial Practice ◽  
Vortical Structures ◽  
Numerical Predictions

The first vane leading edge film cooling is challenging because of the highest thermal load and the complex flow interaction between the hot mainstream gas and the coolant flow. This interaction varies significantly from the stagnation region to the regions of high curvature and acceleration further downstream. Additionally, in industrial gas turbines with multiple combustor chambers around the annulus the first vane leading edges may also be exposed to large wake disturbances shed from the upstream combustor walls. The influence of these vortical structures on the first vane leading edge film cooling is numerically analysed in this paper. In order to assess the capabilities of the flow solver TBLOCK to simulate these complex interactions an experimental test case is modelled numerically. The test case is available in the open literature and consists of a cylindrical leading edge and two rows of film cooling holes representative of industrial practice. A LES turbulence modelling strategy with WALE sub-grid scale (SGS) model is applied and compared against experimental results. Based on this validation it is decided to analyse also the wake–leading edge interaction, dominated by large scale unsteady vortical structures, using the same WALE sub-grid scale LES model. The initial flow domain with the cylindrical leading edge and cooling holes is extended to incorporate the effect of the combustor wall, which is modelled as a flat plate with a square trailing edge. The location and the size of the plate are scaled to be representative of industrial practice: the plate is located upstream from the leading edge at a distance twice the leading edge diameter, and the thickness of the plate is one half of the leading edge diameter. Two different clockwise positions of the vertical combustor wall model were investigated and compared with the datum configuration: the former where the axis of the plate and the leading edge are aligned (central wake location), the latter with the combustor wall circumferentially shifted up by a quarter of the leading edge diameter (circumferentially shifted wake location). Numerical predictions show that the shed vortices from the combustor wall trailing edge have a highly detrimental effect on the leading edge film cooling by periodically removing the coolant flow from the leading edge surface. This results in an increased unsteady thermal load. These negative effects are less significant in the case of circumferentially shifted wake, due to the combined action of both shed vortices.

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.

Influence of Large Wake Disturbances Shed From the Combustor Wall on the Leading Edge Film Cooling

2014 ◽  
Vol 136 (8) ◽  
Author(s):  
Cosimo Maria Mazzoni ◽  
Christian Klostermeier ◽  
Budimir Rosic
Keyword(s):  
Film Cooling ◽  
Gas Turbines ◽  
Thermal Load ◽  
Leading Edge ◽  
Coolant Flow ◽  
Test Case ◽  
Subgrid Scale ◽  
Trailing Edge ◽  
Industrial Practice ◽  
Vortical Structures

The first vane leading edge film cooling is challenging because of the highest thermal load and the complex flow interaction between the hot mainstream gas and the coolant flow. This interaction varies significantly from the stagnation region to the regions of high curvature and acceleration further downstream. Additionally, in industrial gas turbines with multiple combustor chambers around the annulus the first vane leading edges may also be exposed to large wake disturbances shed from the upstream combustor walls. The influence of these vortical structures on the first vane leading edge film cooling is numerically analyzed in this paper. In order to assess the capabilities of the flow solver TBLOCK to simulate these complex interactions an experimental test case is modeled numerically. The test case is available in the open literature and consists of a cylindrical leading edge and two rows of film cooling holes representative of industrial practice. A LES turbulence modeling strategy with the WALE subgrid scale (SGS) model is applied and compared against experimental results. Based on this validation it is decided to analyze also the wake–leading edge interaction, dominated by large scale unsteady vortical structures, using the same WALE subgrid scale LES model. The initial flow domain with the cylindrical leading edge and cooling holes is extended to incorporate the effect of the combustor wall, which is modeled as a flat plate with a square trailing edge. The location and the size of the plate are scaled to be representative of industrial practice: the plate is located upstream from the leading edge at a distance twice the leading edge diameter, and the thickness of the plate is one half of the leading edge diameter. Two different clockwise positions of the vertical combustor wall model were investigated and compared with the datum configuration: the former where the axis of the plate and the leading edge are aligned (central wake location), the latter with the combustor wall circumferentially shifted up by a quarter of the leading edge diameter (circumferentially shifted wake location). Numerical predictions show that the shed vortices from the combustor wall trailing edge have a highly detrimental effect on the leading edge film cooling by periodically removing the coolant flow from the leading edge surface. This results in an increased unsteady thermal load. These negative effects are less significant in the case of circumferentially shifted wake, due to the combined action of both shed vortices.

Overall Effectiveness for a Film Cooled Turbine Blade Leading Edge With Varying Hole Pitch

2010 ◽  
Author(s):  
Thomas E. Dyson ◽  
Dave G. Bogard ◽  
Justin D. Piggush ◽  
Atul Kohli
Keyword(s):  
Turbine Blade ◽  
Hot Spots ◽  
Leading Edge ◽  
Stagnation Line ◽  
Convective Cooling ◽  
Impingement Cooling ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Overall Effectiveness ◽  
Blade Leading Edge

Overall effectiveness, φ, for a simulated turbine blade leading edge was experimentally measured using a model constructed with a relatively high conductivity material selected so that the Biot number of the model matched engine conditions. The model incorporated three rows of cylindrical holes with the center row positioned on the stagnation line. Internally the model used an impingement cooling configuration. Overall effectiveness was measured for pitch variation from 7.6d to 9.6d for blowing ratios ranging from 0.5 to 3.0, and angle of attack from −7.7° to +7.7°. Performance was evaluated for operation with a constant overall mass flow rate of coolant. Consequently when increasing the pitch, the blowing ratio was increased proportionally. The increased blowing ratio resulted in increased impingement cooling internally and increased convective cooling through the holes. The increased internal and convective cooling compensated, to a degree, for the decreased coolant coverage with increased pitch. Performance was evaluated in terms of laterally averaged φ, but also in terms of the minimum φ. The minimum φ evaluation revealed localized hot spots which are arguably more critical to turbine blade durability than the laterally averaged results. For small increases in pitch there was negligible decrease in performance.

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.

Numerical simulation of swirling flow effect on the first stage vane film cooling distribution

2016 ◽  
Vol 07 (03) ◽  
pp. 1650031
Author(s):  
Hong Yin
Keyword(s):  
Flow Field ◽  
Film Cooling ◽  
Swirling Flow ◽  
Cooling System ◽  
Leading Edge ◽  
Suction Side ◽  
Local Area ◽  
Swirl Flow ◽  
Flow Effect ◽  
Recirculation Flow

In advanced gas turbine technology, lean premixed combustion is an effective strategy to reduce peak temperature and thus, NO[Formula: see text] emissions. The swirler is adopted to establish recirculation flow zone, enhancing mixing and stabilizing the flame. Therefore, the swirling flow is dominant in the combustor flow field and has impact on the vane. This paper mainly investigates the swirling flow effect on the turbine first stage vane cooling system by conducting a group of numerical simulations. Firstly, the numerical methods of turbulence modeling using RANS and LES are compared. The computational model of one single swirl flow field is considered. Both the RANS and LES results give reasonable recirculation zone shape. When comparing the velocity distribution, the RANS results generally match the experimental data but fail to at some local area. The LES modeling gives better results and more detailed unsteady flow field. In the second step, the RANS modeling is incorporated to investigate the vane film cooling performance under the swirling inflow boundary condition. According to the numerical results, the leading edge film cooling is largely altered by the swirling flow, especially for the swirl core-leading edge aligned case. Compared to the pressure side, the suction side film cooling is more sensitive to the swirling flow. Locally, the film cooling jet is lifted and turned by the strong swirling flow.

Sign in / Sign up

Export Citation Format

Share Document