Blade Temperature Research Articles

Film Cooling and Thermal Performances of a Blade Tip Winglet Operated in an Annular Test Rig and a Test Gas Turbine Power Plant

Abstract This paper investigates the film cooling performances and thermal behavior of a blade tip winglet, which was initially investigated in a 1.5-stage experimental axial turbine facility and subsequently also tested in a gas turbine power plant. The adiabatic film cooling effectiveness was measured in the rotating 1.5-stage axial turbine facility, whereas in the gas turbine power plant, the tip winglet had active internal cooling coupled to the airfoil and winglet film cooling system. For both test facilities, the tip winglet consisted of several fan-shaped diffuser cooling holes, which were distributed in various regions of the tip winglet and airfoil. The pressure-sensitive paint method was employed to measure the film cooling effectiveness in the 1.5-stage test rig, which was conducted for two tip clearances and a range of blowing ratios. The tip winglet in the gas turbine test power plant was operated for a long duration and under various gas turbine steady-state and transient operating conditions. The blade and tip metal temperatures were measured with embedded thermocrystal sensors and thermocouples for a wide range of operating conditions. Measured local film cooling distributions from the 1.5-stage experimental turbine indicated that it was strongly dependent on the coolant blowing ratios but relatively insensitive to the tip clearances and coolant-to-hot gas density ratio. The measurements also indicated that there was reasonably good coverage of film cooling in large areas of the winglet cavity floor and pressure and suction side tip rims. Metal temperature measurements at various locations of the tip winglet from the gas turbine power plant tests, also indicated that the measured values were within the expected ranges of the predicted metal temperatures.

Jet Engine Technology: Improving Efficiency and Performance through Advanced Design and Cooling Systems

The aviation industry is facing increasing pressure to improve jet engine efficiency while reducing environmental impacts, driven by the need for sustainable development and reduced carbon emissions. Jet engines, central to modern aviation, are evolving through innovative designs and advanced cooling systems. This paper provides an overview of the latest research aimed at improving jet engine efficiency, with a focus on the design and development of advanced turbine blade cooling systems, fan designs, materials, and the utilization of computational tools and techniques. The goal is to promote sustainable development and environmental responsibility in the aviation industry. A systematic review of existing literature, including academic studies and industry reports, was conducted to gather relevant insights. The key findings indicate that enhancing jet engine efficiency can be achieved by lowering turbine blade temperatures, optimizing fan designs and materials, and leveraging advanced computational techniques. These strategies contribute to reducing pollution, improving overall efficiency, and decreasing fuel consumption. The effective implementation of these methods can lead to high-efficiency operations and greater environmental sustainability for jet engines.

Optimal Timing for Safe Bivalving of Fiberglass Casts Is Before the Exothermic Peak.

Cast saw injury is a notable source of medicolegal risk. Previous work with plaster casts demonstrated that cast saw injury was minimized by waiting 12 minutes before removal. In this study, we evaluate the safety parameters of fiberglass casting materials. Eight-ply plaster and fiberglass casts were applied to a pediatric forearm model at variable dip-water temperatures, and the mean time to reach their exothermic peak was determined. Fiberglass casts were then maintained at the manufacturer's recommended dip-water temperature and removed at intervals of 2 (before exothermic peak), 6 (approximately fiberglass's exothermic peak), or 12 (after exothermic peak) minutes. All casts were removed by a pediatric orthopaedic surgeon blinded to the cast set time. Cast/blade temperature, saw force, blade-to-skin contact, bivalve time, cast spreading force, and cut completeness were assessed individually and as short (<6-minutes) or long (≥6-minutes) set times. Fiberglass casts exothermically peaked markedly earlier (5.2 [IQR = 5-5.4] minutes) than plaster (14.8 [IQR=13.7-15.3] minutes), P < 0.0001, at maximum temperatures, which did not markedly differ. Downward force applied during fiberglass cast removal was markedly lower in the short versus long set time groups [average forces of 8.3 (IQR=6.4-10.4) versus 12.9 (IQR=11.1-14.5) Newtons, P < 0.0001, as were maximum forces: 23.2 (IQR=18.9-26.6) versus 43.8 (IQR=38.6-48.5) Newtons, P < 0.0001]. Bivalve time and maximum cast spreading force were decreased in short set times with 40.5 (IQR=39.2-44.7) versus 44.4 (IQR=40.6-47.3) seconds (P = 0.06) and 15.5 (IQR=14-18.5) versus 21.5 (IQR=18-26.5) N (P = 0.07), respectively. Maximum saw blade temperature was markedly lower in the short (99.6°C [IQR=98.2-105.6°C]) versus long (130.6°C [IQR=121.9-141°C]) set times (P = 0.04). No notable differences in blade-to-skin touches or touch duration were detected. Unlike plaster, fiberglass casts cut before exothermically peaking were associated with less downward force, faster bivalve times, and decreased spread force without increased blade temperature or skin contacts. This suggests that fiberglass casts can be bivalved markedly earlier without increased risk of injury.

In-Hole Measurements of Flow Inside Fan-Shaped Film Cooling Holes and Downstream Effects

The study of low-speed jets into crossflow is critical to the performance of gas turbines. Film cooling is a method to maintain manageable blade temperatures in turbine sections while increasing turbine inlet temperatures and turbine efficiencies. Initially, cooling holes were cylindrical. Film cooling jets from these discrete round holes were found to be very susceptible to jet liftoff, which reduces surface effectiveness. Shaped holes have become prominent for improved coolant coverage. Fan-shaped holes are the most common design and have shown good improvement over round holes. However, fan-shaped holes introduce additional parameters to the already complex task of modeling cooling effectiveness. Studies of these flows range in hole lengths from those found in actual turbine blades to very long holes with fully developed flow. The flow within the holes themselves is difficult to study as there is limited optical access. However, the flow within the holes has a strong effect on the resulting properties of the jet. This study presents velocity and vorticity fields measured using high-resolution magnetic resonance velocimetry (MRV) to study three different fan-shaped hole geometries at two blowing ratios. Because MRV does not require line of sight, it provides otherwise hard-to-obtain experimental data of the flow within the film cooling hole in addition to the mainflow measurements. By allowing measurement within the cooling hole, MRV shows how a poor choice of diffuser start point and angle can be detrimental to film cooling if overall hole length and cooling flow velocity are not properly accounted for in the design. The downstream effect of these choices on the jet height and counter-rotating vortex pair is also observed.

Cooled MarkII blade surface pressure and temperature distribution by a conjugate heat transfer analysis using Reynolds stress baseline turbulence model

Cooled MarkII blade surface pressure and temperature distribution by a conjugate heat transfer analysis using Reynolds stress baseline turbulence model

Turbine cooling air estimation in thermodynamic simulations

AbstractModern gas turbines utilize a high amount of the core mass flow rate for component cooling. Thus, a coherent thermodynamic gas turbine representation demands a well-modeled secondary air system, which is able to estimate mass flow rates depending on respective design decisions. In this paper, the focus is set on the estimation of turbine blade cooling air. For this purpose, five different methods are presented and analyzed. The described concepts can be split into empirical and semi-empirical approaches. The semi-empirical approaches, the Horlock [1], Jonsson [2] and the Halliwell [3] method, are able to predict the blade temperature based on a given cooling mass flow rate or the needed cooling air based on a given blade temperature. In contrast, the empirical methods, the Grieb [4] and the Walsh [5] method, can only predict the cooling air consumption. Due to the fully empirical approaches the field of application is limited to the considered engine structures. On the other hand, the empirical methods lead to a better convergence behavior in comparison to the semi-empirical approaches due to their relatively simple calculation methods. The selected cooling air estimations are implemented in the performance code DLRp2 [6–8]. Therefore, processes and methods are deployed that allow to estimate turbines with unlimited cooled stages. Additionally, an off-design procedure is proposed to consider the occurring stagnation temperature drop between stator and rotor based on a reference temperature offset. A simplified thermodynamic gas turbine model is used to analyze the different cooling air estimation methods. For this purpose, sensitivity analyses for the main cooling air parameters are carried out. Moreover, all methods that were developed for the most stressed operating point are compared. Finally, the simplified model is calibrated to NASA’s energy efficient engine [9].

A numerical investigation of heat transfer and clogging in smooth and ribbed cooling channels of the GEF9’s first rotor blade

Analysis of heat transfer and the possibility of clogging in gas turbines play a crucial role in their design and material selection. Due to the high operating temperatures of gas turbine blades, internal blade cooling is utilized to reduce the damage of high temperatures and enhance the durability of gas turbines. Among these methods, the blades' internal cooling is one of the most common approaches. The main objective of this research is to investigate the feasibility of using ribs to increase the internal cooling rate and consequently reduce the surface temperature of the second rotor blade of the GEF9 turbine. Additionally, the experimental observations of damaged blades, proved that some of the internal cooling paths are blocked due to particle deposition of dusty cooling air. So, the main innovation of this research lies in simulating the dust trajectory of dust particles within the cooling channels using an Eulerian-Lagrangian approach to predict blockages in the cooling channels, as well as investigating erosion resulting from the impact of dust particles with the cooling passage walls. Based on the obtained results, ribs increase the convective heat transfer up to 40% but raise the pressure drop in other hands. With analysis of different types of geometrical parameters, the best thermal efficiency is about 0.76 (less than 1). The probability of clogging in the ribbed channels is 100 times more than smooth ones and also the rate of erosion in ribbed channels is 10% more than smooth channels which can increase the stress and reduce the durability of the blade. So using rib in this type of blade is not feasible. By comparing the volume fractions of dust particles near the internal walls in the smooth cooling channels, the first and fourth channels have the highest and lowest likelihood of blockages, respectively and it shows that front channels should be designed with higher diameter.

Mathematical theory and CFD solutions for internal convective cooling of gas turbine blades highlighting the effects of rotation

A quasi-one-dimensional mathematical theory is developed for determining the effects of rotation on the fluid flow and heat transfer in a cooling channel within a rotating gas turbine blade. The three-dimensional intricacies in the solutions of the same problem are captured through accurate computational fluid dynamic simulations. The derived analytical closed-form solutions predict the amount by which the relative total temperature of the coolant changes as a result of rotation. It is shown that such changes due to rotation are significant proportion of the change of coolant temperature due to heat transfer. The computed spatial evolutions of the primary velocity field, secondary velocity field and the force fields are presented as the coolant progresses from the inlet to the outlet so that the complexities and subtleties are exposed to a high degree of quantitative visualizability. It is hoped that the extensive physical discussion given in this paper will be a comprehensive resource for future research, design and understanding. Detailed considerations to various components of the centrifugal, Coriolis and pressure forces are given. The flow field is shown to depend strongly on whether the coolant moves in the spanwise outward or inward direction. The three-dimensionality of the flow field is reflected in complex circumferential variations of the blade temperature and the internal heat transfer coefficient. When the effects of the x–component of the Coriolis force interacts non-linearly with the effects of the z–component of the centrifugal force and a drastic x-gradient in fluid density due to a thin thermal boundary layer, then the primary velocity profile may exhibit great complexity, including flow separation. The present paper thus captures the strong, non-linear, two-way coupling between the fluid dynamics and heat transfer in the internal convective cooling occurring in a channel within a rotating gas turbine blade. Based on about 100 three-dimensional computational fluid dynamics simulations, new heat transfer correlations are developed for average Nusselt number for a rotating channel that will be of practical use.

Numerical study on infrared radiation intensity and suppression methods of nozzle considering multi-component effects

Improving the accuracy of infrared intensity calculation in exhaust system and expanding infrared suppression methods are of great significance for enhancing aircraft stealth performance, this paper studies the infrared radiation (IR) intensity (3–5 µm) of nozzle considering the influence of turbine and afterburner. Initially, the flow field characteristics are computed using commercial software, followed by ray tracing simulations using the reverse Monte Carlo method (RMCM) and transmission calculations using the multi-scale multi-group wideband k distribution model (MSMGWB). The wall emissivity is determined experimentally. This paper quantitatively analyzes the difference between infrared intensity obtained through traditional computational methods and that of the actual model, revealing the impact of low-emissivity coating and turbine cooling on the infrared intensity of the integrated model. It emphasizes the suppression effects and physical mechanisms of the heat shield flanging structure on the infrared characteristics of the integrated model. The results indicate significant differences between infrared intensity derived from traditional computational methods and that from actual models. The influence of the afterburner and turbine infrared characteristics must be considered in infrared numerical studies. When the computation model is a standalone nozzle, the infrared intensity can increase by up to 541.78 % and decrease by a maximum of 14.02 % compared to the integrated model composed of the turbine, afterburner, and nozzle. For the model composed of nozzle and afterburner, the infrared intensity can decrease by up to 10.5 % under non-swirl flow condition, without any increase observed. Compared with turbine wall, changing the emissivity of afterburner wall has more significant effect on infrared suppression of integrated calculation model. Reducing the emissivity of all afterburner walls from 0.7 to 0.1 significantly reduces the infrared characteristics in the range of 0° to 12.5°, while increasing it in the range of 12.5° to 45°. Applying a low-emissivity coating only to the high-temperature walls of the afterburner can inhibit the increase in emitted radiation on the wall, thereby eliminating the increase in infrared intensity mentioned above and reducing the infrared intensity by up to 39.14 %. Compared to low-emissivity coating technology for turbine, implementing cooling measures on turbine blades can effectively reduce the infrared intensity of the integrated model, achieving a maximum reduction of 19.7 % when the blade temperature is lowered by 240 K. Adding flanging structure to the heat shield in the convergent section alters the flow field structure near the nozzle walls, lowering wall temperature and thereby suppressing the infrared characteristics of the integrated model between 10° and 80°. This results in a maximum reduction in infrared intensity of 14.75 % (detection angle of 50°) with negligible impact on thrust coefficients. This is the angle range where afterburner low emissivity coatings and turbine cooling cannot achieve infrared suppression. This is of great significance for infrared stealth in large angle range. Combined with the above effective infrared suppression measures, the infrared intensity can be reduced by up to 49.42 %.

Evolution of Rotating Internal Channel for Heat Transfer Enhancement in a Gas Turbine Blade

To achieve higher thermal efficiency in a gas turbine, increasing the turbine inlet temperature is necessary. The rotor blade at the first stage tolerates the highest temperature, and the serpentine internal channel located in the middle chord of the rotor blade is vital in guaranteeing the blade’s service life. Therefore, it is essential to illustrate the evolution of the rotating internal channel in a gas turbine blade. In the paper, the influence of the Coriolis force, including its mechanisms, on the conventional rotating channel are reviewed and analyzed. A way to utilize the positive heat transfer effect of the Coriolis force is proposed. Recent investigations on corresponding novel rotating channels with a channel orientation angle of 90° (called bilaterally enhanced U-channels) are illustrated. Moreover, numerical investigations about the Re effects on bilaterally enhanced smooth U-channels were carried out in the study. The results indicated that bilaterally enhanced U-channels can utilize the Coriolis force positive heat transfer effect on the leading and the trailing walls at the same time. Re and Ro are vital non-dimensional numbers that influence the performance of bilaterally enhanced U-channels. Re and Ro have an independent influence on the heat transfer performance of the bilaterally enhanced U-channel. Ro is good for the heat transfer of the bilaterally enhanced U-channel on both the leading and the trailing walls. Therefore, the bilaterally enhanced U-channel is suitable for application in the middle chord region of a turbine blade, since it can utilize the rotation effect of the rotating blade to improve the heat transfer ability of the blade and thus reduced the blade temperature. At the same Ro, Re positively affects the Nu on the leading and the trailing walls of the Coriolis-utilization rotating smooth U-channel, but plays a negligible role on Nu/Nu0.

Analysing temperature distributions in turbine first-stage rotor blades of a helicopter turboshaft engine

In helicopter turboshaft engines, turbine blades operate under extreme conditions. With increasing engine power, the gas temperature following the combustion chamber can reach approximately 1300 K. The turbine rotors endure significant centrifugal forces due to their high rotational speeds. Additionally, they experience thermal and aerodynamic loads from the flow of combustible gases, which non-uniformly impact the turbine blades at high temperatures. Furthermore, the mechanical properties of turbine blade materials are limited and strongly influenced by operating temperatures. This article presents a numerical investigation focusing on the temperature distribution of first-stage turbine rotor blades that do not feature internal cooling channels. The results indicate the regions of peak temperatures and evaluate rotor blade strength. Comparative analysis between theoretical and numerical calculations of blade temperature distribution reveals minor disparities: approximately 30 degrees at the blade shroud, 8 degrees at the mid-span section, and 15 degrees at the hub. These variations amount to less than 3% at the shroud, 1% at the mid-span section, and 1.5% at the hub.

Advances in the application of non-contact temperature measurement technology for aero-engine blade.

The advancement of the aviation sector has made the temperature measurement technology for aero-engine turbine blades essential for maintaining the engine's safe and steady performance. The non-contact temperature measurement technology is a trending research focus in turbine blade temperature measurement due to its benefits of not requiring direct touch with the object being measured and its suitability for high-temperature and high-speed conditions. This paper provides a concise overview of various key non-contact temperature measurement methods for aero-engines, such as fluorescence temperature measurement, fiber-optic temperature measurement, and radiation temperature measurement. It discusses the temperature measurement principle, technical characteristics, and the current research status both domestically and internationally. Based on this, this Review further discusses the main challenges faced by the non-contact temperature measurement technology and the development trend of the future.

Study of design methods and heat transfer characteristics of combined cooling structure for turbine blades with turbine inter-vane burner

The turbine inter-vane burner can greatly increase the engine thrust but brings about a harsh thermal environment to the turbine blades. Fuel-air combined cooling structure is an effective way to solve the thermal protection problem of the blades. There is a lack of research on the thermal protection of turbine blades with turbine inter-vane burner, especially systematic research on the fuel–air combined cooling structure. In this paper, the design methods of the fuel–air combined cooling structure of turbine blades with turbine inter-vane burner are proposed, and one-dimensional empirical formulas are used to design specific parameters. The heat transfer characteristics of the designed combined cooling structure are numerically investigated and analyzed without turbine inter-vane combustion, where the mainstream temperature varies from 818 K to 1414 K, the cold air temperature ranges from 500 K to 700 K, and the mass flow rate of cold air changes from 0.005 kg/s to 0.02 kg/s. The results show that in the combined cooling structure with embedded fuel channels, the heat absorbed by fuel is mostly influenced by cold air temperature compared to mainstream temperature and mass flow rate of cold air. The proportion of fuel heat absorption to overall heat absorption decreases from 0.25 to 0.13 when mainstream temperature increases from 818 K to 1418 K, as the increase of mainstream temperature has a minimal effect on the heat flux of the fuel channels, but can significantly increase the overall heat absorption. The proportion of fuel heat absorption to the overall heat absorption changes from 0.15 to 0.31 when cold air temperature increases from 500 K to 700 K, as the increase of cold air temperature enhances the heat exchange between the fuel channels and the cold air but decreases the heat exchange between the mainstream and the cold air. Besides, the flow and heat transfer performance of the combined cooling structure of turbine blades with turbine inter-vane combustion is numerically studied. The results show that the temperature of blade and fuel channels are maintained within the safe range with turbine inter-vane combustion, providing research support for subsequent studies on the fuel–air combined cooling structure of turbine blades with turbine inter-vane burner.

Uniformizing the temperature of turbine blades during the heat treatment process from the perspective of radiation energy distribution

Nickel-based single-crystal turbine blades are critical components of jet engines. Due to element segregation and microstructural defects in as-cast blades, vacuum heat treatment is essential. Controlling the blades’ heat treatment temperature is crucial for improving mechanical performance. However, when treating multiple blades simultaneously, large temperature differences often occur, making it challenging to ensure consistent and uniform temperatures. This paper establishes a numerical algorithm that links blade temperature with radiant energy, optimizing blade arrangement to achieve uniform energy distribution and temperature field homogenization. The improved view factor finite element algorithm is more suitable for blade structures. The Particle Swarm Optimization (PSO) and Coordinate Descent (CD) methods were used to solve the optimization function, aiming to minimize the temperature difference among blades. Finite element simulation was utilized to model the temperature field of the blades during radiative heat transfer from heating elements. The PSO and CD optimization methods were employed to find the local optimal solution that minimizes the view factor variance, reducing it from 2.24 before optimization to about 1. Simultaneously, the optimized maximum average temperature difference of the blades was 71.1 °C, approximately 50 % lower than the 138.8 °C before optimization. By comparing the view factor values with the simulated temperature fields, the strong correlation between the view factor and blade temperature was verified. This paper innovatively proposes a numerical method for optimizing blade arrangement to homogenize the temperature field, effectively enhancing temperature consistency during the heat treatment process.

Thermal image evaluation of crop diseases in the incubation period.

Early blight is one of the main diseases that causes serious pepper yield and quality reduction. The identification of the presymptomatic features during the incubation period is of great research significance, in that scientific prevention measures in the incubation period may effectively reduce the loss of peppers. The present study confirms the feasibility of the identification of presymptomatic features in pepper early blight by infrared thermography during the incubation period. The infrared thermal images of pepper blades were captured during the whole incubation period in our experiment. Evaluation parameters such as temperature uniformity (U) and the variation coefficient (C) were established by infrared thermal images for the blade temperature field. Evaluation parameters such as temperature uniformity U and variation coefficient C of pepper blades change during the time from being inoculated to significant morbidity. For cases of stabbing inoculation, there were no relatively obvious symptoms in visible light images until 96 h after inoculation, whereas some presymptomatic features appeared in infrared thermal images at 60 h after inoculation. Based on the uniformity changes (δU) and the variation coefficient changes (δC) in the temperature field distribution, the characteristic polynomial (CP)-GBDT model has been established by optimizing the gradient boosting decision tree (GBDT) model. Compared with the GBDT mode, the CP-GBDT model accuracy increased from 87.5% to 91.7% on the test set. Evaluation parameters such as temperature uniformity U and variation coefficient C can be used to effectively characterize the presymptomatic features of pepper early blight by infrared thermography during the incubation period. The results of the present study can provide a reference for the detection of crop diseases in the incubation period. © 2024 Society of Chemical Industry.

Application of Heating Uniformity Detection Method for Heating Blade Temperature of Tobacco Products

Application of Heating Uniformity Detection Method for Heating Blade Temperature of Tobacco Products

Numerical simulations on film cooling performance of turbine blade before and after particles deposition

In the present study, a C3X turbine blade with film holes is selected to investigate the effect of blowing ratio on particle deposition and heat transfer characteristics on the blade surface. The deposition characteristics of the blade surface are obtained based on the EI⁃Batsh particle deposition model, and the surface morphology of the deposited layer is reproduced using the dynamic mesh model. The results show that the particle deposition mainly occurs at the leading edge and pressure surface of the blade, but the large-size particles will be deposited on the suction surface due to rebound. The effect of blowing ratio on large-size particles is more obvious, and the deposition rate tends to decrease and then increase with the increase of blowing ratio. Within the range of blowing ratios in this paper, particle deposition leads to an improvement in the cooling efficiency of the air film in the pressure surface along the flow direction near the air film hole region, but a significant decrease in the cooling efficiency near the trailing edge region. Overall, particle deposition caused an increase in overall blade temperature.

Two-dimensional temperature field inversion of turbine blade based on physics-informed neural networks

Correct evaluation of the blades' surface temperature field is crucial to the structural design and operational safety of aero-engine turbine blades. Current high-temperature measurement technology can only measure the limited discrete points temperature of the hot-end turbine blades. If the boundary conditions are completely unknown, it is still difficult to solve equations through traditional computational fluid dynamics methods. It is an inverse problem to predict the temperature field of the turbine engine blade with unknown boundary conditions. This paper proposes a two-dimensional temperature field inversion of turbine blades based on physics-informed neural networks (PINN) and finite discrete temperature measurement points. The PINN is used to model the nonlinear mapping of input variables and output variables. Only a small amount of data is used to train the neural network. It can be found that adding the loss term of the physical governing equation during training can make the neural network better predict the temperature field of the turbine blade, which can also avoid anomalies such as large temperature errors in some areas that may occur when training with only a small amount of data. When the boundary conditions are completely unknown, the average relative error of PINN trained with only 62 points of temperature data on the test set is below 2%, and the R2−Score is above 0.95. When the training data reaches 122 points of temperature data, the average relative error of PINN on the test set is less than 1%, and the R2−Score reaches 0.99.

Application of a Ranque–Hilsch Vortex Flow in an Internal Cooling of a Gas Turbine Engine Blade

Abstract The efficiency of a gas turbine engine is directly impacted by the turbine inlet temperature and the corresponding pressure ratio. A major strategy, aside from the use of costly high-temperature blade materials, is increasing the turbine inlet temperature by internally cooling the blades using pressurized air from the engine compressor. Understanding the fluid mechanics and heat transfer of internal blade cooling is, therefore, of paramount importance for increasing the temperature threshold, hence increasing engine efficiency. This article presents modeling and test results of a novel cooling approach, one in which the Ranque–Hilsch vortex flow is adopted for the first-row gas turbine blade cooling. Simulation and test results demonstrate the successful formation of continuous Ranque–Hilsch vortex flow by injecting compressed air into a cylindrical chamber equipped with seven air inlets. At an inlet pressure of 100 kPa, the outlet temperature from the vortex tube dropped 255 °C, which allowed the blade temperature to cool by 47 °C. When a total inlet pressure of 300 kPa was admitted, the drop-in temperature reached 65 °C. The device has the potential to drop the cooling air temperature below the freezing point with increased inlet pressure. The thermal efficiency of the gas turbine blade increased by about 3% when vortex cooling with 10% mass of partially compressed air was extracted at about 910 kPa. For the tested scenario of a 17 MW power output, the partial extraction had a better efficiency increment than extraction at full compression, which was 1200 kPa.

Simulation of flow and heat transfer characteristics of laminated turbine blades with kerosene cooling channels

An air-kerosene thermal mass coupled turbine blade with kerosene micro-channels added to the traditional laminated structure turbine blade is proposed, and numerical simulations are carried out. The enhanced heat transfer mechanism of the air-kerosene thermal mass coupled turbine blade is studied, and the influence of different kerosene temperatures, blowing ratios, and solid thermal conductivity on the heat transfer of the laminated turbine blades is analyzed. The results show that adding kerosene micro-channels can significantly reduce the blade temperature and change the cooling gas heat transfer direction inside the laminate cooling structure. Compared with the traditional laminate cooling structure, adding kerosene micro-channels can significantly improve the heat transfer performance of the blades, and the integrated cooling efficiency increases by 31.7%. Moreover, when the kerosene temperature decreases from 400-300 K, the cooling efficiency increases by 3.9%. Similar conclusions can be obtained by studying the increases in the blowing ratio and the solid thermal conductivity, respectively.