Gas Turbine Blades Research Articles

Effect of hydrogen on nickel oxide reduction on the surface of nozzle blade of a gas turbine unit

Currently, there is a growing interest in the use of hydrogen in the composition of fuel mixtures for turbojet engines and gas turbine units (GTU). The effect of hydrogen on heat-resistant nickel alloys of gas turbine blades has been little studied. In this regard, this work is devoted to studying the effect of hydrogen on nickel oxide reduction on the surface of the nozzle blade of a gas turbine engine. Hydrogen is a good reducing agent. Therefore, this article discusses the effects of hydrogen under various conditions with metal oxides, and methods of metal oxides reduction on the surface of the blades of a gas turbine engine. The thermodynamics of the interaction of aluminum, titanium, nickel and tungsten oxides with hydrogen fluoride and reactions of fluoride with hydrogen was investigated in the temperature range 273 – 1373 K. It was established that the interaction of aluminum oxide with hydrogen fluoride occurs in the temperature range from 273 to 1073 K, titanium oxide with hydrogen fluoride – from 273 to 373 K, nickel oxide with hydrogen fluoride – from 273 to 873 K. In this case, of the resulting fluorides, only nickel fluoride interacts with hydrogen at temperatures above 673 K. Hydrogen interacts with nickel oxide throughout the entire temperature range, and with tungsten oxide at temperatures above 1173 K. We studied the effect of hydrogen on heat-resistant nickel alloys of gas turbine blades subjected to preliminary fluorination and not treated with fluorine compounds. Nickel oxide reduction with hydrogen proceeds better after the preliminary fluorination process. In this case, particles 2 – 5 μm in size containing 90.16 % Ni are formed on the surface of the blade sample. Without fluorination, this process at 1223 K and duration of 1 h does not occur.

The effect of dimple/protrusion arrangements on the comprehensive thermal performance of variable cross-section rotating channels for gas turbine blades

The cooling structure design of turbine blades plays a vital role in ensuring the safety and reliability of the entire gas turbine set. Our team's previous research revealed the influence of the rotating and variable cross-section effect on the aerothermal characteristics of the channel. In order to further enhance the cooling performance, four types of dimple/protrusion arrangement schemes are applied to five channels with different height ratios in this research. The detailed flow-heat transfer-friction characteristics and comprehensive thermal performance of each channel are obtained. The results show that the introduction of heat transfer enhancement element produces obvious heat transfer enhancement effect. Compared with the smooth reduced cross-section channel, the Nu/Nu0 is increased by 10.9 % (height ratio = 4:1, both of two passes are arranged by protrusions). For the rough expand cross-section channel, the Nu/Nu0 is increased by 8.9 % (height ratio = 1:2, the first pass is arranged by protrusions and the second pass is arranged by dimples). Within the scope of this paper, the arrangement scheme of both of two passes are arranged by dimples has the best comprehensive thermal performance. The arrangement scheme of both of two passes are arranged by protrusions has the highest Nu/Nu0, but also the highest f/f0.

Laser cladding of NiCrCoFeNbMoX high-entropy alloy to increase resistance to corrosion of gas turbine blades

To improve the corrosion resistance of gas turbine blades that work in a salty and humid environment, NiCrCoFeNbMoX (X = 0,0.25,0.5,0.75,1) high entropy alloy (HEA) was deposited on GTD-111 superalloy by laser cladding. The results showed that with the increase of Mo, the volume fraction of Laves and eutectic phases increases, and the FCC phase decreases. An increase in the Laves phase shows an increase in the corrosion resistance of the above alloy system. In the meantime, NiCrCoFeNbMo0.75 HEA showed the highest corrosion resistance due to showing the highest self-corrosion potential and the lowest self-corrosion current density. The increase of the Laves phase and the reduction of the FCC phase were the most important factors of corrosion resistance. The NiCrCoFeNbMo1 system showed a lower corrosion resistance than the NiCrCoFeNbMo0.75 HEA due to the creation of fine eutectics. The results of hot corrosion also showed that NiCrCoFeNbMo0.75 coating has the highest resistance to hot corrosion in 90 wt% Na2SO4-10 wt% NaCl at 750 °C due to the uniform distribution of MoO3.

An analysis of microstructural morphology, surface topography, surface integrity, recast layer, and machining performance of graphene nanosheets on Inconel 718 superalloy: Investigating the impact on EDM characteristics, surface characterizations, and optimization

Inconel 718 finds extensive applications in the aviation and aerospace industries, particularly in the manufacturing of jet engines and high-speed airframe components like fasteners, bolts, buckets, instrumentation parts, wheels, and spacers. It is also utilised in the production of cryogenic tankage and gas turbine blades. The present study focuses on investigating the machining performance of graphene nanosheets on Inconel 718. Various aspects of Inconel 718's machinability through electrical discharge machining (EDM) have been examined, including material removal rate (MRR), surface roughness, surface morphology, tool rear Rate (TWR), residual stresses on the machined surface, Vickers hardness, and recast layer thickness. The investigation reveals a significant impact of process parameters on these machining characteristics. The effects of graphene nanosheets have been observed using several analytical instruments such as field emission scanning electron microscopy (FE-SEM), transmission electron microscopy (TEM), particle size analyzer, and X-ray diffraction (XRD). Furthermore, optimization of the response data with respect to input parameters has been performed in this study. TEM analysis is used to determine the size of individual debris particles in deionised water and mixed graphene nanosheet dielectric mediums. To verify that the debris particles are the same as the parent material, energy dispersive X-ray spectroscopy (EDX) is used. To determine the compounds and crystal structures present in the base metal and machined surfaces, XRD analysis is used. A high-resolution X-ray diffractometer (HRXRD) is used to measure the residual stresses on the machined surface. EDX composition testing is used to analyze surface modification. Due to the rapid heating and quenching that takes place in the dielectric medium, the machined surface becomes harder. Deposited materials, microholes, and surface textures can all be observed through FESEM microstructure observation. Comparing conventional EDM to nanosheets mixed dielectric, the thickness of the recast layer is reduced. To recapitulate, the study explores how various machining parameters and dielectric mediums affect EDM processes. It examines debris particle size, compound formation, residual stresses, surface modification, hardness, microstructure, and recast layer thickness. The addition of graphene nanosheets to the dielectric medium produces promising results, reducing the thickness of the recast layer and improving surface quality. The results offer suggestions for improving Inconel 718 material surface properties and EDM machining effectiveness.

Optimizing Crack Detection in Gas Turbine Blades Using Implanted Capsules of Ionizing Gas in Nonsteady Operation at Nonuniform Temperature

Optimizing Crack Detection in Gas Turbine Blades Using Implanted Capsules of Ionizing Gas in Nonsteady Operation at Nonuniform Temperature

Designing brazing filler metal for heat-resistant nickel alloys of new generation marine gas turbines

The object of research is the processes of the formation of brazed joints and the stressed state. The subject of research is structure, chemical composition, long-term high-temperature strength at a temperature of 900 °C, speed of high-temperature salt corrosion. Existing brazing filler metals have a high-temperature performance of 40–50 % of the performance of the SM93-VI and SM96-VI alloys. Despite this, brazing is the main technique of joining modern heat-resistant cast alloys. Therefore, the development of new brazing filler metals that ensure the formation of joints with increased long-term high-temperature strength is relevant. Ship gas turbine blades operate at a temperature of 900 °C. The purpose of the development of the new SBM-4 brazing filler metal is to achieve long-term high-temperature strength of brazed joints at a temperature of 900 °C at the level of 85–90 % of the strength of heat-resistant alloys SM93-VI and SM96-VI. A two-stage method was used in the development of SBM-4 brazing filler metal. At the first stage, the chemical composition of the brazing filler metal base was determined, taking into account the peculiarities of operating conditions of the blades of marine gas turbine engines and the achievements of materials science of heat-resistant alloys. At the second stage, the depressant and its necessary content were selected. Computer software was used to determine the distribution between the γ- and γ'-phases, taking into account the participation of each element in both dispersion and solid-solution strengthening. Rational limits of concentrations of alloying elements were determined. The criterion was the minimum susceptibility of brazing filler metal to the formation of brittle phases, taking into account the influence of chromium, rhenium, and tantalum concentrations on resistance to high-temperature salt corrosion and high-temperature performance. The long-term strength of SM93-VI and CM96-VI alloys brazed with SBM-4 brazing filler metal is 89–91 % of the strength of the base metal. Technologies of brazing and correction of casting defects have been introduced into production.

Conjugated heat transfer simulation of flow mechanism and heat transfer characteristic of sweeping jet impinging on leading edge in turbine cascade

Increasing inlet temperatures of gas turbines is the way to go for the development of advanced aero engines. As the inlet gas temperature of modern gas turbine engines continues to increase, the operating temperature has far exceeded the melting temperature of the typical heated region. Excessive temperatures can cause turbine blade ablation and damage, thus affecting the reliability and safety of the engine. The sweeping jet actuator (SJA) is a fluid oscillator that has attracted considerable attention in the field of enhanced heat transfer and flow control due to its non-constant sweeping characteristics. As a new impinging cooling method, this paper adopts the device to carry out the research work on the flow mechanism and heat transfer characteristics of sweeping jet (SJ) impinging on leading edge in gas turbine blade. Firstly, the numerical investigation of conjugated heat transfer simulation combined with unsteady Reynolds-averaged Navier–Stokes (URANS) is conducted to analyze the cooling performance of SJ. Using a C3X high-pressure turbine blade as a prototype, the temperature drop with the SJ scheme was improved by 1.8 K and 8.7 K compared with the circle jet (CJ) scheme for H/D = 1 and H/D = 5, respectively. Finally, the internal flow field of the blade is reconstructed by combining the proper orthogonal decomposition (POD) method. Through the POD method, we can deeply analyze the mechanism of the effect of SJ on the non-constant heat transfer at the leading edge of the blade, which provides theoretical support for the practical engineering application of SJ in the future industrial field.

Blade tip leakage flow and heat transfer characteristics over a gas turbine blade at subsonic and transonic exit conditions

The present study numerically investigates the blade tip leakage flow (TLF) over a gas turbine blade cascade at low speed ( M ex = 0.1) and transonic ( M ex = 0.92) blade exit Mach numbers, with and without shroud motion. Various blade tip flow phenomena like flow separation and reattachment, flow choking, shock-boundary layer interaction (SBLI), and tip leakage vortex (TLV) are studied, and their influence on the heat transfer coefficient (HTC) distributions over the blade tip and near tip blade suction surface is investigated. The study has found some new zones of heat transfer due to the interaction of horseshoe vortices (HSV) with the blade tip. In addition to the primary TLV, a secondary TLV is formed, which increases near tip blade suction surface heat transfer. The flow transitions to supersonic speed in the aft portion of the blade tip for the transonic case resulting in complex heat transfer distribution due to SBLI, compared to near-uniform distribution for the low speed case. This supersonic flow results in a smaller flow separation zone near the aft portion of the pressure side tip edge for the transonic case compared to the low speed case. The qualitative nature of HTC near the leading edge of the blade tip resembles for low speed and transonic cases. However, near the trailing edge of the blade tip, the qualitative nature of HTC for the transonic case shows a substantial difference due to supersonic flow compared to the low speed case.

The recovery of microstructure and mechanical properties of directionally solidified Ni-based superalloy DZ411 after long-term thermal exposure at different temperatures

Directionally solidified (DS) Ni-based superalloys are widely used in gas turbine blades due to excellent mechanical properties for elevated temperature applications. In harsh service environments, the microstructure degrades and mechanical properties deteriorate. In order to extend the service life of gas turbine blades, rejuvenation heat treatment (RHT) is applied. Considering the existence of temperature gradient in a real gas turbine blade during service, the effects of RHT on the different degraded microstructure at different temperatures are necessary to be investigated. However, few investigations concentrated on it. In this study, to simulate the existence of temperature gradient in a real gas turbine blade during service, DS Ni-based superalloy DZ411 was carried out with long-term thermal exposure at different temperatures. The impact of RHT on the different degrees of degraded microstructure and mechanical properties after thermal exposure at different temperatures was explored. Furthermore, the evolution of dislocations after RHT is studied, which has been rarely investigated. It was found that the degenerated microstructure at different aging temperatures can be significantly restored by RHT, including dissolution and reprecipitation of γ′ and M23C6 carbides and annihilation and rearrangement of dislocations. In addition, creep and low cycle fatigue (LCF) properties could also be effectively recovered by RHT. After RHT, the steady-state creep rate is decreased and creep rupture life has been prolonged. The LCF life and the shape of the stabilized hysteresis loop are restored. However, the creep properties are not completely restored, due to the unrecovered microstructure including the undissolved γ′ in interdendritic regions, the residual dislocations, and irreversible decomposition.

Thermal–Fluid–Solid Coupling Analysis on the Effect of Cooling Gas Temperature on the Fatigue Life of Turbine Blades with TBCs

The application of thermal barrier coatings (TBCs) can increase the blade’s operating temperature and fatigue life. Previous studies have neglected the effects of cooling gas temperature variations on the temperature field, stress field, and fatigue life of blades and TBCs. For this reason, this paper considers the inhomogeneity of the high-temperature gas inlet temperature, the internal cooling gas temperature, the periodicity of the external flow field, etc., and establishes a finite element model of the gas turbine blade with TBCs. Then, the life of the blade and the TBCs is predicted based on the Ncode 2020R2. Finally, the effect of the cooling gas temperature on the temperature, stresses, and life of the blade and the TBCs is analyzed. The results show that the fatigue life of the TBCs is lower than that of the blade, and the low-life region of the TBCs is located at the leading edge of the blade, which is consistent with the TBCs shedding region of the real blade and verifies the accuracy of the life prediction method in this paper. The fatigue life of the blade and TBCs firstly increases and then decreases with the increase in the cooling gas temperature, and the trend of the stress changes in the opposite direction. When the cooling gas temperature is increased from 573 K to 873 K, the minimum life of the blade is increased by 358.5%, and that of the TBCs is increased by 138.7%. The conclusions can provide guidance for the design of long-life turbine blades with TBCs.

High precision thickness evaluation of thermal barrier coating with high frequency eddy current testing method

Consisting of a top ceramic coating (TC) and a bond coating (BC) layer, thermal barrier coating (TBC) is a key component of the heavy duty gas turbine blade to protect the substrate material from high working temperature. Thinning of the top coating and the thermally grown oxide (TGO) layer formed between TC and BC cause TBC thickness reduction, and may significantly affect the in-service performance of TBC. High frequency eddy current testing (HF ECT) is considered as an efficient technique for nondestructive evaluation (NDE) of the TBC thickness, but its measurement accuracy is not satisfied up to now. Aiming to realize a high precision TBC thickness evaluation with HF ECT signals, a numerical method for simulating HF ECT of TBC was developed and implemented with consideration of the dielectric properties of TC and TGO material. Through numerical simulations with the developed code, an ECT probe of optimized configuration and robust against lift-off variation was designed and fabricated for the TBC thickness evaluation. A HF ECT measurement system is then setup to investigate the validity of the prototype ECT probe and sizing method experimentally. It was found that the thickness of the TC layer can be evaluated within 10 μm accuracy for the plate TBC specimens, which reveals the HF ECT method and new ECT probe are efficient for NDE of the TBC thickness.

Study on flow and heat transfer characteristics in rectangular channels with lantern-shaped pin fin array: Part II—Global sensitivity analysis and optimization

Increasing gas turbine inlet temperatures are placing more stringent requirements on the efficiency of internal cooling. Pin fin arrays are a common tool for cooling electronic components or trailing edge cooling channels of gas turbine blades. A crucial concern for researchers is how to achieve a higher Nusselt number with constrained pressure loss to guarantee that the efficiency of the gas turbine will not be impacted by the power dissipation generated by the driving flow passing through the internal cooling loops. In Part I of this study, a lantern-shaped pin is proposed. To seek a more efficient design solution for the pin fin shape, the second part combines the surrogate model, Sobol’s method, and NSGA-II optimization method to perform the overall and spatial sensitivity analysis and optimization for the shape variables on the generatrix of different pin rows. The first-order and total-order sensitivity indices of different variables to Nu T and f are calculated using the surrogate model-based Sobol’s method. The variables that significantly affect the channel performance are identified. The overall and spatial sensitivity analysis results indicate that the variable X 3 i -1 contributes the most to Nu T and f, with total-order sensitivity indices of 70.59–64.26 and 58.16–59.09%, respectively. The overall optimization increases the Nu T and f by 6.95 and 9.91%, respectively, while the spatial optimization leads to an increase in the Nu T by 2.29% and a decrease in the f by 4.56%. Correspondingly, the two optimizations resulted in an increase in the TPI by 3.61 and 3.89%, respectively. The Pareto front solution obtained by the spatial optimization has a more concentrated distribution. Although the TPI obtained by the spatial optimization is only 0.21% higher than that of the overall optimization, it provides designers with a strategy considering the Nusselt number and pressure loss.

Remaining Useful Life Prediction Method for High Temperature Blades of Gas Turbines Based on 3D Reconstruction and Machine Learning Techniques

Turbine blades are crucial components exposed to harsh conditions, such as high temperatures, high pressures, and high rotational speeds. It is of great significance to accurately predict the life of blades for reducing maintenance cost and improving the reliability of gas turbine systems. A rapid and accurate blade life assessment method holds significant importance in the maintenance plan of gas turbine engines. In this paper, a novel on-line remaining useful life (RUL) prediction method for high-temperature blades is proposed based on 3D reconstruction technology and data-driven surrogate mode. Firstly, the 3D reconstruction technology was employed to establish the geometric model of real turbine blades, and the fluid–thermal–solid analysis under actual operational conditions was carried out in ANSYS software. Six checkpoints were selected to estimate the RUL according to the stress–strain distribution of the blade surface. The maximum equivalent stress was 1481.51 MPa and the highest temperature was 1393.42 K. Moreover, the fatigue-creep lifetime was calculated according to the parameters of the selected checkpoints. The RUL error between the simulation model and commercial software (Control and Engine Health Management (CEHM)) was less than 0.986%. Secondly, different data-driven surrogate models (BP, DNN, and LSTM algorithms) were developed according to the results from numerical simulation. The maximum relative errors of BP, DNN, and LSTM models were 0.030%, 0.019%, and 0.014%. LSTM demonstrated the best performance in predicting the RUL of turbine blades with time-series characteristics. Finally, the LSTM model was utilized for predicting the RUL within a gas turbine real operational process that involved five start–stop cycles.

Effect of a curved turning vane on the heat transfer and fluid flow of four-pass internal cooling channels of gas turbine blades

This numerical study examined the effects of a curved turning vane on the heat transfer and pressure loss of a four-pass internal cooling channel. Three-dimensional omega-based Reynolds stress (RSM-ω) turbulence model equations were used in the computation process. Three different curved turning vane configurations were studied using a simple, without a turning vane under stationary and rotating conditions for Reynolds numbers between 20,000 and 60,000 and a rotation number of 0.042. The numerical results showed good agreement with previous experimental data. Under both stationary and rotating conditions, the curved turning vane in a hub turn reduced the pressure drop significantly compared to the conventional turning vanes. Although a curved turning vane attenuated overall heat transfer, the local heat transfer increased it in particular regions, such as the hub turn and the fourth turn of cooling passages, particularly in cases with smaller radii. Coolant flow through the hub turn of the serpentine channel can reduce recirculation, separation, and flow impingement, which are unfavorable factors of pressure loss, owing to the semi-circular configuration of the curved turning vane. Regarding the rotating conditions, four-pass channels with a smaller radius of a curved turning vane provide better overall cooling performance.

Role of surface phenomena in deformation and fracture of gas turbine blades and stamping tools under intense mechanical and thermal loading

Role of surface phenomena in deformation and fracture of gas turbine blades and stamping tools under intense mechanical and thermal loading

Cooling performance of droplet-shaped Kagome truss structure combined with jet array impingement composite cooling structure

Gas turbine is one of the most efficiency conversion equipment of heat into power, which is widely used in aviation propulsion and power generation. With the gas inlet temperature of gas turbine skyrocketing recently, gas turbine can achieve higher efficiency but the super high temperature far exceeds the bearable temperature limits of blade material. This causes a great threaten in lifespan of turbine blades and making advanced and efficient cooling structures for blades essential. To enhance the cooling effectiveness of middle chord region of gas turbine blade, a novel droplet-shaped Kagome truss structure (DKTS) is proposed and investigated through experimental and numerical analysis. Subsequently, within jet impingement channel, DKTS is compared against circular pin-fins (CPF), droplet-shaped pin-fins (DPF), and circular Kagome truss structures (CKTS) for fluid flow and heat transfer performance. The experimental method is used to study effects on cooling characteristics between DKTS and CKTS under different operating conditions. The numerical method is used to study effects on cooling characteristics among CPF, DPF, CKTS and DKTS, and to study influences of investigated parameters on cooling performance of DKTS. Experimental results show that under equivalent windward areas, both heat transfer efficiency and flow performance of DKTS are better than those of CKTS. The results of numerical simulations indicate that at the same porosity across jet Reynolds numbers ranging 5000–50000, DKTS increases thermal-hydraulic performance by 11.1%–17.5% relative to CKTS. Moreover, DPF and DKTS reduce flow resistance and augment heat transfer capabilities as compared to CPF and CKTS. This shows that turbulators with droplet-shaped rods have better thermal-hydraulic performance. Finally, the effects of differing jet Reynolds numbers, diameter, tail angles, and inclination angles on fluid flow and heat transfer characteristics of DKTS are conducted. These results may provide a reference for further optimization researches or industrial applications of DKTS.

Numerical analysis of non-Fourier heat conduction dynamics in the composite layer

This paper presents the numerical analysis of non-Fourier heat conduction in thin composite layers under asymmetrical boundary conditions. In the thermal barriers such as steam and gas turbine blades, thin film coating are used to protect the blade from thermal damage. The coating on the blades are very short in length. Heat conduction across thin composite layer with short time is examined using a finite element approach. With this very small duration with the finite speed of the thermal wave, the Fourier mode of heat conduction is disappeared due to the infinite speed of the thermal wave assumption. Therefore, analyzing the non-Fourier heat conduction in thin layers is essential. The developed model is executed in Python using Newmark's scheme and the constant average acceleration method to predict the temperature variation and temperature contours. The present model is validated with an experimental and numerical solution with good agreement. Besides, the temperature distribution across the composite layer with the entire length of the substrate and the coating for different thermal conductivity values, thermal diffusivity, and relaxation time are examined. It is noted that when the dimensionless

Protection Performance Investigation of an Infrared Temperature Measurement Probe for Gas Turbine

For the temperature monitoring requirements of gas turbine blades, a new infrared temperature measurement probe structure is designed in this paper. Cooling air is used to cool the probe and prevent flue gas flowing into the probe. The computational fluid dynamics method is used to analyze the working process of the temperature probe. Two cases that the optical opening of probe is facing to and on the back of flue gas are considered. The influences of the air inlet total pressure on the pollution protective performance and thermal protective performance of the probe are analyzed. The results show that when the optical opening of probe is facing to flue gas, the pollution protective performance is poorer while the thermal protective performance is better than that when facing to gas. According to the temperature distribution of the probe, three optimization schemes of the probe are proposed: spraying thermal barrier coating on the wall of the probe, drilling film holes at the bottom of the probe, and combining film holes with thermal protective performance. The results show that the probe maximum temperature is reduced by 97, 187 and 328 K by using these three methods, respectively.

Optimization of a novel impulse gas turbine nozzle and blades design utilizing Taguchi method for micro-scale power generation

This study is conducted to optimize the nozzle and blade of a compact non-combustion impulse gas turbine driven by the pressurized gas line with Computational Fluid Dynamics (CFD) approach and Taguchi method. For nozzle, throat diameter, nozzle inlet and outlet diameter and convergent-divergent length were investigated. Meanwhile, number of blades, blade radius, blade curvature angle, blade thickness and surface roughness were evaluated for blade. An L25 orthogonal array was chosen for both optimizations. Once optimized, the corresponding operating envelope was identified and compared against the original turbine. When operating at flow rate of 1.2MMSCFD and pressure of 69 bar, the turbine with optimized nozzle produces a maximum power output of 4383.59W at 5500 rpm, while with optimized blades produce 2058.64W at 5000 rpm. Combining optimized nozzle and blades produced 4928.64W at 6000 rpm. These produced powers are significantly higher than the original turbine maximum power of 1743.81W at 4500 rpm. This indicates potential performance enhancement of the turbine by optimizing its blade and nozzle geometry which is useful for its implementation on the offshore platform. To fully utilize this potential enhancement, the electric generator attached to the turbine shall be fine-tuned to have peak power at the optimum rpm band of the optimized turbine.

Combined effects of blowing ratio and mist diameter with mainstream turbulence intensity in mist-assisted film cooling

Advanced cooling techniques are essential in modern gas turbines to improve the efficiency of the Brayton cycle by raising the turbine inlet temperature. Mist-assisted film cooling offers an innovative solution for cooling high-temperature gas turbine blades. When mist particles with various momenta are injected through the cooling holes, the interplay between mainstream turbulence intensity (Tu) and mist momentum plays a crucial role in mist-assisted film cooling under high Tu conditions. Hence, this numerical study investigates the relationship between mainstream Tu and mist characteristics for the first time, considering the blowing ratio and mist diameter as controlling parameters. Film effectiveness and particle distribution statistics are analysed to understand their correlation. The results demonstrate that mist-assisted film effectiveness follows the trend of air-only film cooling with varying blowing ratios and Tu; however, the magnitude is amplified. The influence of the mist diameter on film effectiveness is determined by the surface-area-to-volume ratio and Stokes number. These factors impact the sensitivity of particles to Tu and subsequently affect film effectiveness. This study provides a novel quantitative analysis of the combined effect of mainstream Tu and mist on film cooling effectiveness, revealing the correlation between mist and mainstream parameters.HighlightsCombined effect of mainstream turbulence intensity (Tu) and mist characteristics is investigated.High mainstream Tu significantly increases the cooling enhancement induced by mist.Mist-assisted film cooling exhibits similar trends to varying mainstream Tu and blowing ratio.Mist diameter significantly influences the response to Tu, governed by surface-area-to-volume ratio and Stokes number.Small particles demonstrate the highest cooling effectiveness and variation by Tu.