Gas Turbine Blades Research Articles

Augmented cooling performance in gas turbine blade tip with slot cooling

The blade tip is one of the most susceptible parts of gas turbine blade due to the thermally induced failure caused by a hot gas tip leakage flow through tip clearance. A highly sophisticated and efficient blade tip cooling scheme plays a crucial role in reliable operation of a gas turbine blade. In the present study, we adopted a special slot cooling scheme in turbine blade tip. The film cooling effectiveness (FCE) and flow characteristics of turbine blade tip were investigated according to the cooling flow through the slot gap in blade tip rim. A linear cascade was fabricated to measure and evaluate the cooling performance of the blade tip with the film cooling through a slotted rim versus the blade tip with cooling holes. The local FCE values of the blade tip floor and rim upper surface were measured using the pressure-sensitive paint (PSP) method. The Reynolds number based on the axial chord length and the inlet velocity in the experiment was set at 192,000. The tip clearance was varied from 1 to 3% of the blade span and the coolant mass flow rate ranged from 0.3 to 0.7% of the main flow. The slot cooling in the blade tip, which can be incorporated through an additive manufacturing, showed highly enhanced cooling performance ranging from 53 to 116% compared to that of the blade tip with film cooling holes for various tip gaps and coolant mass flow rates. The slot cooling in the blade tip is expected to improve the thermal durability and reliability of blade tip with better coolant coverage and additional convective cooling through the internal cooling passage in the narrow squealer rim.

Mixed-dimensional geometric coupling of port-Hamiltonian systems

We propose a new interconnection relation for infinite-dimensional port-Hamiltonian systems that enables the coupling of ports with different spatial dimensions by integrating over the surplus dimensions. To show the practical relevance, we apply this interconnection to a model system of an actively cooled gas turbine blade. We also show that this interconnection relation behaves well with respect to a discretization in finite element space, ensuring its usability for practical applications.

Review of Advanced Effusive Cooling for Gas Turbine Blades

Turbine inlet temperature has continuously increased to improve gas turbine performance during the past few decades. Although internal convection cooling and traditional film cooling have contributed significantly to the current achievement, advanced cooling schemes are needed to minimize the coolant consumption and maximize the cooling efficiency for future gas turbines. This paper conducts a comprehensive review of advanced effusive cooling schemes for gas turbine blades. First, the background and the history of turbine blade cooling are introduced. Then, the metrics of effusive cooling efficiency are defined. Next, effusion cooling, impingement/effusion cooling, and transpiration cooling are reviewed. The flow and heat transfer mechanisms of the cooling schemes are emphasized, and the design trends of the cooling schemes are revealed. Finally, the conclusions and future research perspectives are summarized.

Data-Driven Conjugate Heat Transfer Analysis of a Gas Turbine Vane

Cooling structures of gas turbine blades have become more complex to achieve a better cooling effect. Therefore, heat transfer analysis tools with higher accuracy and efficiency are required to verify the effectiveness of cooling designs and continuously improve the design. In this work, a data-driven method is combined with a decoupled conjugate heat transfer analysis. The analysis object is a typical air-cooled gas turbine first-stage vane with film cooling, impingement cooling, and pin-fin cooling. In addition, a conventional 3-D conjugate heat transfer simulation of the vane was executed for contrast. Results show that this method shortens the time of the heat transfer analysis process significantly and ensures accuracy. It proves that the data-driven method is effective for the evaluation of a modern gas turbine cooling design and is an improvement compared to the traditional three-dimensional heat transfer analysis method.

Thermal hydraulic performance augmentation by petal-shaped ribs in a two-pass cooling channel

The increase of inlet gas temperature will lead to security problem of turbine blades. Internal cooling channels are usually used to ensure the safe operation temperature of the turbine blades. Rib-roughness can effectively enhance the cooling performance of the internal channels. To further promote the channel performance, a novel petal-shaped rib is proposed and numerically simulated with SST k-ω model in this paper. Three different rib types, including square rib, a-quarter-cylinder rib and petal-shaped rib, are compared under the identical rib height and rib cross-section area. The results showed that the petal-shaped and a-quarter-cylinder ribs have larger heat transfer performance than the normal square rib. The heat transfer of the petal-shaped rib is the largest and increases by up to 21.14% compared with square rib. While the friction factor of the channel with the petal-shaped rib is 10.04% smaller than the square rib. The thermal performance factor of the petal-shaped rib reaches up to 2.02, which concludes a significant increment of 22.91% compared with the square rib. The results prove that the petal-shaped rib has great potential to enhance the thermal performance of the cooling channel and can be applied in the design of cooling channels for gas turbine blades.

Evaluation of Tensile Properties and Microstructure of a Scraped Gas Turbine Blade using Miniature Specimens

A gas turbine (GT) blade is a key hot-pass component for advanced GT engines, and should have stable properties under extreme conditions of 1,350°C and 3,600rpm, etc. The GT blade, after operating with nearly 800 equivalent start-stops (ES) or 24,000 equivalent operation hours (EOH), should be replaced due to degradation of properties and microstructure, particularly, the formation of cracks in the airfoil tip and platform region. To date, the assessment of materials, prototype blade, etc, has been extensively studied in a laboratory simulated environment; however, evaluation of the full-scale blades in a service environment has been rarely reported. Here, the properties and microstructures of an F-class GT first stage blade, with a service history of 800ES were investigated. The results showed decreasing tensile properties at the airfoil part due to its higher temperature exposure. The microstructural characterization results revealed that the finer grain size and dendrite interfaces facilitated the formation of Cr-enriched M<sub>23</sub>C<sub>6</sub> along grain boundaries, as well as the spherical γ' in the airfoil part, resulting in a decrease in tensile properties. The results obtained here provide precious background for assessing the serviced blade and developing advanced blades.

A novel damage identification method for flue gas turbine blades based on tip timing

Tip timing signal analysis has been applied to the online condition monitoring of high-speed blades. However, traditional tip timing analysis methods are not suitable for low-speed flue gas turbines. Therefore, this paper proposes a novel blade tip timing signal analysis method based on an investigation of the dynamic response characteristics of low-speed blades. First, the finite element modal theory is introduced to analyze the characteristics of blade damage. Second, an equivalent cantilever beam analysis model of flue gas turbine blades is established under complex environment and working conditions. In order to monitor the variation of local stiffness, a damage identification method based on the variation of the free end deflection of the equivalent cantilever beam is proposed. Finally, a rotating blade tip timing monitoring testing rig is established to verify the feasibility of the proposed method. The results show that the cracks originating at about 80% of the blade height have the greatest influence on blade stiffness, followed by blade root. The calculated blade damage parameters are 4.8464 mm and 3.7588 mm, and the crack influencing factors are 4.7476 and 3.6822, respectively, indicating that the change trend is consistent with the blade damage rules.

Enhancement of blade tip cooling with different position on tip cooling slot

One of the leading causes of thermally induced failure in gas turbine blade tip is a tip leakage flow. To protect blade tip efficiently from strong tip leakage flow, slot cooling is adopted in blade tip considering the application of additive manufacturing. For slot cooling scheme in blade tip, slot location ranging from PS slot to SS slot is a crucial role in film cooling effectiveness of blade tip. Therefore, we investigated the effect of the cooling-slot location on film cooling effectiveness and flow characteristics of blade tip with slot cooling considering the application of additive manufacturing. The local film cooling effectiveness on the blade tip floor and upper surface of the rim are obtained using pressure sensitive paint method. The case of coolant discharge from the PS slot was both stable and evenly spread, resulting in an outstanding coolant coverage of the entire tip regions. However, under severe effect of hot main gas flow behavior, the case of the SS slot resulted in an uneven FCE distributions over the entire tip region from the LE to the TE region. Under the studied tip gaps and coolant mass flow rates, the PS slot configuration provided a superior blade tip cooling performance under most operating conditions, but the SS slot configuration provided an outstanding cooling performance on blade tip only at a large amount of coolant mass flow rate compared to PS slot. Thus, the blade tip with a PS slot is recommended as a robust cooling design. However, the blade tip with a SS slot may have a limited application for design and off-design conditions.

Cooling of Gas Turbine Blade Leading Edge Via an Array of Air Jets With Variable Diameters

Abstract In this study, the cooling of a gas turbine (GT) blade leading edge using a single row of five impinging air jets was numerically and experimentally investigated. The inner surface of the blade leading edge was depicted as a semicircle profile. The numerical study was performed using a shear stress transport (SST) k–ω turbulence model. On the other side, the experimental work was conducted using the infrared (IR) thermal camera. Additionally, the heat transfer characteristics were investigated with changing the Reynolds number (Re) from 5000 to 40,000. The study was carried out for different cases based on the jet diameter (d), named as fixed diameter (FD) and variable jet diameter in the streamwise flow direction, i.e., ascending diameter (AD) and descending diameter (DD). The local and average Nusselt numbers and heat transfer uniformity were studied at different operating and design parameters. The results revealed that for FD, increasing the jet diameter decreased the average Nusselt number while it enhanced the heat transfer uniformity. Furthermore, for all values of Re, the thermal performance of the AD was better than the DD, which has the worst heat transfer uniformity. Finally, predicted developed correlations to estimate the average Nusselt number, surface heat transfer uniformity, and performance evaluation factor (PEF) for all cases of jet diameter were introduced.

Numerical study of film cooling effectiveness of a gas turbine blade under variable blowing ratio and turbulence intensity

AbstractThe present paper investigates a three‐dimensional simulation of film cooling on a C3X turbine blade with a single hole at a suction surface. The Reynolds averaged Navier–Stokes approach with k–ε realizable turbulence model and enhanced wall function are used for the numerical simulation. To simulate the jet flows, the length of the jet input approximately 4.5 times the diameter of the hole is added to the geometry so that the jet outlet flow is closer to the actual condition. The density ratio of the cooling flow to the mainstream flow is assumed about 2. The numerical results in four blowing ratios of 0.5, 0.7, 1.0, and 1.4, and at the low turbulence intensity (0.02%), and high turbulence intensity (12%) are extracted and compared for the turbine blade with a single hole. The results show that the turbulence intensity has a dual effect on the film cooling effectiveness and a higher blowing ratio increases the strength of the jet against the cross‐flow. Moreover, it is illustrated that the distribution of the film cooling effectiveness in higher blowing ratios and high turbulence intensity is more uniform than the low blowing ratios and low turbulence intensity.

The influence of the size of details on the frontal resistance forces during magnetic abrasive finishing

The magnetic abrasive finishing process (MAF) is an effective method of finishing details. It is especially beneficial of processing important parts with complex shape, such as cutting tools, gas turbine blades, medical products, etc. MAF allows to affect the quality of the treated surface, surface hardness, microgeometry of cutting edges (for a cutting tool), removal of residual stresses in the material of the details. The relative newness of the method and, hence, little awareness of the processes and phenomena that occur during MAF limit the wide use of the method in production. Insufficient awareness of the nature of MAF, especially regarding the forces that arise during the interaction of the processed part and the environment of the magnetic abrasive tool (MAT). Available researches provide only basic information that is not sufficient for a comprehensive study of MAF. This is especially true for the effect of size of the detail on features of the implementation of the process. Conducting research on the processing of parts of various sizes and determining the forces and phenomena accompanying the MAF. To perform the aim, a device with a dynamometric sensor that allows to measure the frontal resistance was used. Ferromagnetic and paramagnetic material cylindrical details with different diameters (8, 12 and 16 mm) were used. Polymam-M and Polymam-T (grain size 400/315 μm, 200/100 μm) powders were used. The speed of movement varied in range 1 - 3 m/s, induction of magnetic field: 0.20 - 0.24 T. It was established that for the specified processing conditions, the influence of the magnitude of the magnetic field induction prevails over the influence of the speed of movement of details. The effect of the size of the parts on the specifics of the MAO is shown, the nature of processing by individual structural formations of the MAO and the difference between them is determined. The question of fluctuations in the magnitude of the resistance force at MAO and the influence of movement speed on them is considered. The influence of technological parameters on the frontal resistance force during processing of parts of different diameters and materials with different magnetic properties has been determined, and further directions for research have been determined.

Cover Picture: Materials and Corrosion. 10/2022

Cover:The internal passageways and holes of the turbine blades have been protected by slurry aluminizing diffusion coatings. The microstructural observations of the aluminized coating, as well as its high‐temperature oxidation behavior, have been investigated. The oxidation test at 1000 °C was performed for 50, 100 and 150 h and the Kp constant for aluminide coatings and un‐coated superalloy are successfully calculated to be 0.0034 and 0.0086 mg2cm−4h, respectively.More detailed information can be found in: Ali Jokar, Farzin Ghadami, Najmeh Azimzadeh and Davood Salehi Doolabi, A novel approach to the protection of internal passageways of gas turbine blades by aluminide coating: Evaluation of oxidation behavior, Materials and Corrosion 2022, 73, 1534.

Numerical Investigation of Heat Transfer Characteristics of Pin-Fins with Roughed Endwalls in Gas Turbine Blade Internal Cooling Channels

In modern jet engines, the effective cooling structure design is essential to both increase turbine efficiency and, in the meantime, maintain structural integrality. Pin-fin arrays play a crucial role in the cooling mechanism of the turbine blade internal cooling system. In the context of investigating the efficiency of cooling techniques using pin-fins, while the other studies focus on the pin-fin configurations, the current study is a step towards optimizing cooling cascade endwalls for better maneuver and reservation of vortices which lead to higher heat transfer near the endwalls. This study presents the results from the investigations of the flow-field and heat transfer characteristics of the pin-fin arrays with roughed endwalls, i.e., recessed endwalls and extruded endwalls types. The heat transfer and pressure drop characteristics of the channel are numerically examined to compare with those of the flat endwalls case within a range of inlet Reynolds number, which varies from 7400 to 36000. Both the leading and trailing sides of the channel are divided into five regions in the streamwise direction to understand further the heat transfer capacity of the pin-fins and the endwalls. The results show that, with these new endwall configurations, the high heat transfer regions near the pin-fins are remarkably enlarged compared to the flat endwalls. And in the meantime, both the new endwall arrangements enhance the heat transfer capacity of the channel near the pin-fins. The two new designs outperform the baseline case, with the heat transfer efficiency index (HTEI) improved by 37.8% for the recessed endwalls and 15.9% for the extruded endwalls. By varying the height of the indentation and the protrusion, it is found that, with smaller heights, both configurations produce a much lower friction factor. However, with much higher heat transfer capacity, leading to relatively higher values of heat transfer efficiency index (HTEI), up to 77.7% and 41.5% of the recessed and extruded endwalls are higher than HTEI of the channel with flat endwalls. These results indicate the great potential for improving the heat transfer capability of pin-fins by optimizing endwall configurations.

Structural and modal analysis of gas turbine blade using ansys

Gas turbine technology has advanced significantly during the last twenty years. The advancement in materials technology, new coatings, and novel cooling techniques are driving the expansion. Material selection is critical since turbomachinery design is complicated, and material performance is directly connected to efficiency. The materials should be able to withstand high vibrations, high temperature, and pressure. In this work, the maximum stress and the deformation induced due to pressure were determined to study the vibrational response, mode shapes, and natural frequencies. Four different materials have been considered, namely Inconel 617, Titanium alloy, Rhenium alloy, and Nimonic 80A. A suitable material for a gas turbine blade was suggested after analyzing these materials. The 3D CAD model is modeled in CATIA V5 and analysis is carried out using FEM-based software ANSYS

3D positioning of defects for gas turbine blades based on digital radiographic projective imaging

During the manufacturing and service process of gas turbine blades, various types of defects may be formed and bring huge threat to the safe operation of gas turbine. To confirm whether the gas turbine blades with defects could be in safe service for some period of time, the residual strength and fatigue life of gas turbine blade should be analyzed and predicted based on the 3D spatial position of defect, which significantly determines the degree of effect on the strength of gas turbine blade. Considering the characteristics of shape (free-from surface) and material (high density superalloy), the digital radiographic imaging techniques could be used in the non-destructive testing of gas turbine blades. However, the projective nature of digital radiography (DR) technique results in the unavailability of position in the transmitting direction of X-ray beam and brings difficulty to the 3D characterization of defect. Aiming at this critical issue, based on the exploration of the changing rule for the projection of a point feature relative to the irradiating angle, a new method is presented to obtain the 3D spatial position of defect. Finally, the accuracy of the proposed method is validated using a metrological computed tomography (CT) system. It can be learned that the proposed DR based method could obtain the 3D spatial position of defect with a relative error less than 2%. The 3D spatial position of defect could be used in the following residual strength analysis and fatigue life prediction of gas turbine blade.

Study on Flow and Heat Transfer Performance of a Rectangular Channel Filled with X-Shaped Truss Array under Operating Conditions of Gas Turbine Blades

In this investigation, the heat transfer and flow capabilities of an X-shaped truss array cooling channel under various operating conditions of gas turbine blades were thoroughly studied. The influence laws of the inlet Reynolds number, inlet turbulence intensity, wall heat flux and cooling medium (air, steam) on the heat transfer and flow performance of the X-shaped truss array channel were analyzed and summarized. The empirical correlations of friction coefficients and average Nusselt numbers with maximum deviations less than ± 14% were fitted. The results show that the inlet Reynolds number has the most significant effect on the flow and heat transfer performance of the X-shaped truss array channel. When the inlet Reynolds number increases from 20,000 to 200,000, the average Nusselt number of the X-shaped truss array channel is increased by 3.92 times, the friction coefficient is decreased by 12.88%, and the comprehensive thermal coefficient is decreased by 31.19%. Compared with the medium turbulence intensity of Tu = 5%, the average Nusselt number, friction coefficient and comprehensive thermal coefficient of the X-shaped truss array channel at Tu = 20% are increased by 3.70%, 2.51% and 2.79%, respectively. With the increase in the wall heat flux, the friction coefficient of the X-shaped truss array channel roughly shows a trend of first decreasing and then increasing, while the average Nusselt number and the comprehensive thermal coefficient show a trend of first rapidly increasing and then slightly decreasing or remaining unchanged. Compared with air cooling, the average Nusselt numbers of the X-shaped truss array channel of steam cooling are increased by 6.30% to 9.54%, and the corresponding friction coefficients and comprehensive thermal coefficients are decreased by 0.11% to 0.55% and 2.63% to 5.59%, respectively.

Surface modification of IN713 LC superalloy with Metco 204NS by laser surface alloying

Ceramics are one of the best engineering materials for coating gas turbine blades. In this study, the Metco204NS ceramic coating (Zr2O + 8%Y2O3) was applied by the laser surface alloying (LSA) method on IN713 LC nickel-based superalloy. To influence the heat input on the structure of the ceramic coating and its substrate, and as well as the rejuvenated zone (RZ), different variables were used in LSA. The results showed that with an increase in heat input by the laser, the sensitivity to liquation cracks in the heat affected zone and solidification cracks in the RZ decreases. The most important reason for this was the increase in backfilling by the molten metal due to its high fluidity. However, with increasing heat input, the hardness increased due to the reduction of the distance between the dendrites. Solidification rate (R) and temperature gradient (G) were identified as the most important microstructure controlling factors in the RZ. So, with increasing the heat input and thus decreasing G and R, the tendency of the structure to change from cellular to columnar and then equiaxed increased. The uniform and homogeneous coating of Metco 204NS significantly increased the wear resistance of IN713 LC superalloy. The higher hardness and wear resistance of melted Metco 204NS coating material relative to the RZ and the base metal was due to the presence of very hard Zr2O and Y2O3 particles in Metco 204NS and the reduction of grain size.

Influence of grooved rib tip structure on tip loss and heat transfer in a gas turbine blade

This study focuses on the effects of three groove tip structures (full rib groove tip, partial rib tip on the suction side, and partial rib tip on the pressure side) on tip leakage flow, aerodynamic characteristics of a cascade, and heat transfer in a gas turbine blade. The groove’s width B = 1.6 mm, while the tip clearance is τ = 1.2 mm. Results of the flow parameters, fluid flow, and heat transfer in the recessed channel are discussed. The results show that all ribbed tips obtain more uniform outlet flow angle distribution and higher aerodynamic performance than the plane tips. The total aerodynamic pressure loss of the ribbed tips on the pressure side is the same as that of complete ribbed tips. The evolution mechanisms are different, although both can improve the turbine efficiency. Although the partial rib tip on the pressure side weakens the mixing of the channel vortex and leakage vortex near the trailing edge and has the best control effect on the leakage vortex, the lack of the suction side rib will make it easier for the low-energy fluid to flow into the gap from the front of the suction side, which is not conducive to reducing the leakage flow inside the gap; the full rib tip not only minimizes the tip relative leakage flow and leakage loss but also increases the channel vortex loss. With the complex vortex system in the groove and the rib blocking effect at the leakage outlet, the suction-side rib tip becomes the tip structure with the best leakage flow control effect under the same clearance, but the leakage vortex loss is the highest.

Processing and Advancements in the Development of Thermal Barrier Coatings: A Review

Thermal barrier coating is critical for thermal insulation technology, making the underlying base metal capable of operating at a melting temperature of 1150 °C. By increasing the temperature of incoming gases, engineers can improve the thermal and mechanical performance of gas turbine blades and the piston cylinder arrangement. Recent developments in the field of thermal barrier coatings (TBCs) have made this material suitable for use in a variety of fields, including the aerospace and diesel engine industries. Changes in the turbine blade microstructure brought on by its operating environment determine how long and reliable it will be. In addition, the effectiveness of multi-layer, composite and functionally graded coatings depends heavily on the deposition procedures used to create them. This research aims to clarify the connection between workplace conditions, coating morphology and application methods. This article presents a high-level overview of the many coating processes and design procedures employed for TBCs to enhance the coating’s surface quality. To that end, this review is primarily concerned with the cultivation, processing and characteristics of engineered TBCs that have aided in the creation of specialized coatings for use in industrial settings.

Investigation of the relationship between the 3D flow structure and surface heat transfer within a realistic gas turbine blade trailing edge internal serpentine cooling channel

Proper internal cooling of the thin trailing edge of gas turbine blades is crucial for blade lifespan extension. In this study, numerical and experimental analyses are conducted to elucidate the relationship between the 3D flow structure and surface heat transfer in a realistic three-pass serpentine channel within the trailing edge of a blade. Magnetic resonance velocimetry is conducted to first experimentally validate the mean velocity field from the Reynolds-averaged Navier Stokes simulation. Heat transfer and friction coefficients are then obtained through numerical analysis. The serpentine channel is investigated at each section. The inlet S-duct contributes to heat-transfer promotion upstream of the first ribbed passage due to an intense vortex pair. This vortex pair generates high streamwise velocity along the pressure and suction side, as well as additional friction due to secondary flow. In the second passage, film cooling holes make the separation bubble attach only to the suction side and create a high streamwise velocity region around the holes by bringing the surrounding flow to the periphery of the holes. This effect also occurs near the ejection holes in the third passage. These flow features within the passages are analyzed using the momentum distortion parameter (D) and secondary flow energy parameter (E). We found that the distortion parameter is not sufficient in predicting the heat transfer trend due to in-plane wall shear. Thus, the estimator is revised as (D+E)/2 to additionally account for secondary flow. This composite parameter shows good performance in estimating heat transfer for all the passages, compared to the friction coefficient which is hard to directly measure.