Turbine Blade Temperature Research Articles

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.

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.

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.

Effect of thermal barrier coating on the thermal characteristic of turbine blade and its geometric optimization

Thermal barrier coatings (TBC) and cooling structures are critical factors influencing the performance of aviation turbine blades. In order to investigate the effects of TBC and different cooling structures on the operating temperature and thermal stress of turbine blades, this study establishes a three-dimensional fluid-thermal-solid coupling model for aviation turbine blades. Based on this model, the study analyzes the effect of TBC thicknesses and different design architectures on the blade thermal characters, including temperature distribution and thermal stress analysis. Stress analysis is also conducted on blades with the best heat dissipation performance. The results indicate that as the thickness of the TBC increases, the blade temperature gradually decreases. Cooling channels effectively reduce turbine blade temperatures. It is also found that a reasonable increase in film cooling holes can further reduce blade temperatures. The thermal stress on the blade gradually decreases along the blade height, with the high-stress regions primarily concentrated at the junction between the blade body and the margin plate. The maximum stress is observed at the trailing edge. The conclusions of this study can provide valuable guidance for the design of aviation turbine blades with better thermal performance.

New insight into Allam cycle combined with coal gasification in supercritical water

In this work, we proposed and investigated the power generation system where coal gasification in supercritical water is combined with Allam cycle (CAllam cycle). This CAllam cycle features zero pollutant emission, easy CO2 collection and no extra heating. Our simulations demonstrated the differences between the natural gas Allam cycle and our proposed CAllam cycle, and also elucidated the mechanisms underlying these differences. Results show that recycling CO2 not only reduces the temperature of gas turbine blades but also further improves the system efficiency. Results also show that the gas turbine outlet pressure corresponding to the maximum efficiency is smaller and away from the gas-liquid line. Furthermore, we graphically identified the combination of CO2 recycling ratio and gas turbine outlet pressure, that the CAllam cycle allowed. Built on these findings, our proposed CAllam cycle was optimized graphically. Taking 40 wt% Hongliulin coal as an example, the maximum system efficiency is 53.19% and the corresponding CO2 recycling ratio and gas turbine outlet pressure are 1.158 and 2.43 MPa, respectively.

Flow and Heat Transfer Characteristics of the Turbine Blade Variable Cross-Section Internal Cooling Channel with Turning Vane

The gas turbine blades are scoured by high temperature gas sustainedly and long-term in harsh environment. It is of great significance to explore effective cooling methods to lower the turbine blade temperature so as to ensure safe and stable operation of the gas turbine. However, there are few studies on the cooling channel considering the turning vane, variable cross-section characteristics, and rotation effect. In this paper, five kinds of serpentine cooling channel models with variable cross-section properties and different thickness guide vanes are constructed. The effects of different thickness guide vanes on the overall performance of the channel under stationary and rotating conditions are discussed and compared by numerical method. The result shows that when stationary (Re = 10,000–50,000), the turning vane with suitable thickness can increase the Nu/Nu0 by 56.5%. The f/f0 is decreased by 14.2%, and the comprehensive thermal performance is increased by 4.5%. When rotating (Re = 10,000, Ro = 0–0.5), the turning vane with suitable thickness can increase the Nuup/Nu0 and Nuall/Nu0 by 33.0% and 4.0%, respectively. The comprehensive performance of the variable cross-section serpentine channel can be greatly improved by arranging the turning vane structure with appropriate thickness.

Error Analysis and Correction of Radiation Pyrometer in Measuring Surface Temperature of Turbine Blades With the Influence of Reflection

Error Analysis and Correction of Radiation Pyrometer in Measuring Surface Temperature of Turbine Blades With the Influence of Reflection

Error of Measuring Surface Temperature of Turbine Blades With the Influence of Reflection by Monochromatic Method

A new method is proposed for correcting the error of monochromatic method in measuring the surface temperature of turbine blades with the influence of reflection. The measurement error caused by surrounding components were numerically investigated. The effects of background temperature, adjacent rotor blades, front stator vanes and rear stator vanes on the temperature measurement error are described in detail. Furthermore, the corrected errors of high emissivity and low emissivity are compared and analyzed and the influence of the relative displacement between the turbine blade and the hot-end component on the temperature measurement error is also discussed. Benefiting from the proposed radiation-based error correction method (RECM), we can reduce the monochromatic temperature measurement error to less than 0.3 % at high emissivity and less than 0.8% at low emissivity theoretically. Finally, the RECM method is studied experimentally, and the results show that the temperature measurement error can be corrected to better than 0.6%, which is in good agreement with the Monte Carlo simulation results.

Conjugate Heat Transfer Simulation of Overall Cooling Performance for Cratered Film Cooling Holes

Film cooling is widely applied as an effective way to maintain the turbine blade temperature at an acceptable level. The present paper investigates the overall cooling effectiveness and flow structure by performing conjugate heat transfer simulations for the flat-plate baseline cylindrical and three cratered film-cooling holes. The flow and heat transfer in the fluid domain is obtained using the Shear Stress Transport turbulence model, and the solid conductivity is considered by solving the Laplace equation. Four blowing ratios ranging from 0.5 to 2.0 are studied. The numerical results show that the concentric elliptic cratered hole yields a slightly higher overall cooling effectiveness than the baseline cylindrical hole, but the two contoured cratered holes exhibit great improvements due to the generation of the anti-kidney-shaped vortex pair. The area-averaged overall cooling effectiveness has improved by 5.58–65.30% for the contoured cratered hole. The variation of Biot number results in small change in the area-averaged overall cooling effectiveness. However, the area-averaged overall cooling effectiveness uniformity coefficient depends on the Biot number.

Superiority analysis of the cooled cooling air technology for low bypass ratio aero-engine under typical flight mission

The cooled cooling air technology (CCA technology) is beneficial to increase the aero-engine operating pressure ratio and turbine inlet temperature, while its influence mechanism during the full flight envelope is still ambiguous. A comparative analysis has been completed in this study to assess the flight performance of an aircraft engine with or without CCA technology. Based on the aircraft and aero-engine integrated technique, a novel analytical model is developed by combining the propulsion requirement analysis model with the engine design/off-design point analysis model. The precise submodules of the heat exchanger and the turbine blade in engine model improve the -practicability of analysis, which considers the influence of the flight mission on the heat transfer performance. This literature suggests that CCA technology could enhance the fuel weight at take-off phase, the weight-specific surplus power, the load factors, and the supersonic cruising Mach number by approximately 3–4.85% with re-optimizing the cycle parameters. The inlet temperature of turbine Tt4 rises by around 3% without the turbine blade temperature exceeding the limitation. This work further analyzes the most influential structural factor of the total engine performance and the transverse tube pitch is discovered as a noteworthy feature. Additionally, it is well known that the heat transfer and pressure drop of the CCA heat exchanger (CCA-HEX) change dramatically during the whole flight mission. Our findings provide valuable insights into a better understanding of the mechanisms of CCA-HEX characteristic change, in which poor heat transfer occurs at low altitudes and high Mach numbers, whereas the largest flow resistance appears in the high-altitude and low-Mach-number regionsin theopposite manner. In conclusion, CCA technology is a cost-effective method to improve the propulsion performance of the aircraft engine for flight mission.

Error Analysis and Correction of Radiation Pyrometer in Measuring Surface Temperature of Turbine Blades with the Influence of Reflection

Error Analysis and Correction of Radiation Pyrometer in Measuring Surface Temperature of Turbine Blades with the Influence of Reflection

Thermodynamic Study of a Cooled Micro Gas Turbine for a Range Extended Electric Vehicle

Carbon dioxide released by the transportation sector has a significant impact on the environment. Nowadays, the research efforts are thoroughly investing in the capability of series hybrid electric vehicles in reducing CO2 and NOX emissions. Many regulations are encouraging automakers to substitute internal combustion engines with electric vehicles. Although the latter are zero-emissions energy sources, they are expensive and have a limited autonomy range. Therefore, recent studies are interested in gas turbines (GTs) in series hybrid electric vehicles (SHEV). A gas turbine (GT) coupled to an alternator is considered as an auxiliary power unit capable of charging the battery of the electric vehicle once depleted. The microturbine is generally composed of a centrifugal compressor and a radial turbine mounted on the same shaft with a recuperator and air bearings. Based on previous researches and thermodynamic analysis, the most suitable gas turbine cycle with regard to the efficiency and the net specific work is composed of two compression stages with an intercooler, a regenerator, and two expansion stages with a reheater. It can be deduced that by increasing the turbine inlet temperature, the system efficiency increases. However, the inlet temperature is limited by turbine materials constraints and specifications. The present paper focuses on developing a more efficient GT cycle by reaching higher inlet temperatures by cooling the turbine blade. It consists of using compressed gas from the compressor and introduce them into the turbine blades. This technology takes into consideration the influence of the extracted mass of compressed air, the effectiveness of the coolant, and the turbine blade temperature on thermal efficiency. An increase of the cycle efficiency of 4.78 points is obtained for a turbine inlet temperature of 1,450 °C.

Performance analysis of large-scale industrial gas turbine considering stable combustor operation using novel blended fuel

The allowable Wobbe index (WI) range proposed by gas turbine companies spans from 0 to 5%. System tuning is required when out-of-range fuel is used. In this study, thermodynamic analysis is conducted to study the gas turbine performance and operability of each component with the fuel exceeding 5% of WI. The novel diluting fuel supply system is employed to apply novel fuel compositions for the J-class gas turbine system. High-hydrocarbon and hydrogen blending fuels are used to study the operability and main parameters of each component. The largest diluent mass flow is required when 100% hydrogen fuel is used with nitrogen diluent method. The compressor cannot be operated with 100% hydrogen fuel due to the compressor surge. The maximum turbine blade temperature with steam diluent methods increases up to 80 °C. The power output and efficiency are enhanced with the increasing blending ratio of high-hydrocarbon and hydrogen employing two diluent methods. When the novel fuel supply system is considered for the stable operability of combustor, both its power and efficiency are increased; however, the operability of compressor and turbine is not always stable, especially with hydrogen fuel.

Exploration of turbine blade cooling strategies for performance boosting and CO2 emissions reduction of combined cycle with steam injection based gas turbine

The present work investigates the use of cooling schemes for the STeam Injection based Gas turbine (STIG) blades i.e. Steam Internal Convection Cooling (SICC) and Steam Transpiration Cooling (STC) techniques. The work aims to carry out the energy and exergy analyses considering variations in the design parameters like Gas Turbine Inlet Temperature (GTIT), Compressor Pressure Ratio (CPR), Turbine blade temperature (Tb ) and Steam mass Injection Ratio (SIR). The STIG Combined Cycle Power Plant (CCPP of 355 MW) performance and CO2 emissions are evaluated under the impact of these design parameters with these two cooling schemes, and comparison is made between them. The exergy efficiency increased by 2.98% points and CO2 emissions reduced by 0.0167 kg/kWh with STC scheme when compared with SICC scheme at GTIT = 1823 K, CPR = 12, Tb = 1123 K and SIR = 6 kg/kg of fuel. Exergy efficiency increased and CO2 emissions reduced with increase in SIR. Highlights Performance evaluation of STIG based CCPP with internal convection and transpiration cooling strategies using steam as a coolant applied to gas turbine blades. CO2 emissions reduction with high Steam mass injection ratio and with steam transpiration cooling scheme. The integration of STIG along with open loop steam cooling schemes enhances the performance of CCPP The exergy efficiency increases with increase in SIR, GTIT for STC scheme as compared to SICC scheme. STC is best cooling scheme with minimum coolant requirement.

Research on hybrid optimization and deep learning modeling method in the performance seeking control

A novel performance seeking control method based on hybrid optimization algorithm and deep learning modeling method is proposed to get a better engine performance. The deep learning modeling method, deep neural network, which has strong representation capability and can deal with big training data, is adopted to establish an on-board engine model. A hybrid optimization algorithm—genetic algorithm particle swarm optimization–feasible sequential quadratic programming—is proposed and applied to performance seeking control. The genetic algorithm particle swarm optimization–feasible sequential quadratic programming not only has the global search ability of genetic algorithm particle swarm optimization, but also has the high local search accuracy of feasible sequential quadratic programming. The final simulation experiments show that, compared with feasible sequential quadratic programming, genetic algorithm particle swarm optimization, and genetic algorithm, the proposed optimization algorithm can get more installed thrust, decrease fuel consumption between 2% to 3%, and decrease turbine blade temperature larger than 15k, while meeting all of the constraints. Moreover, it also shows that the proposed modeling method has high accuracy and real-time performance.

Performance augmentation of combined cycle power plant under the control of differing open loop cooling techniques to the gas turbine blades

ABSTRACT The present study investigates the influence of different blade cooling techniques on the combined cycle power plant (CCPP) in the perspective of performance augmentation. The influence of Compressor Pressure Ratio (CPR), Gas Turbine Inlet Temperature (GTIT) and Turbine blade temperature (Tb) are extensively studied on the performance of CCPP with six open loop cooling schemes namely Air Internal Convection Cooling (AICC), Air Film Cooling (AFC), Air Transpiration Cooling (ATC), Steam Internal Convection Cooling (SICC), Steam Film Cooling (SFC) and Steam Transpiration Cooling (STC). The performance characteristics energy efficiency, exergy efficiency, specific work output, coolant flow rate requirement, specific fuel consumption (sfc) and carbon dioxide (CO2) emissions of CCPP are evaluated under the influence of above cooling schemes and comparison is made between them. The results show that the maximum performance was obtained for Steam Transpiration Cooling (STC) scheme with a minimum coolant requirement along with reduced CO2 emissions.

Measurements of turbine blade temperature in an operating aero engine using thermographic phosphors

A diagnostic method is introduced for non-contact temperature measurements on the surface of a rotating turbine blade in an operating engine using laser-induced excitation of thermographic phosphor coatings on the blades. Temperature measurements of between 513 °C and 767 °C are demonstrated on a pair of moving turbine blades with a fiber optic probe mounted in the engine at rotational speeds of between 28 000 and 32 750 RPM. The method involves measuring the decay times of laser-induced luminescence from two thermographic phosphors applied to thermal barrier coatings on the turbine blades, one with a short, temperature-sensitive, decay time and the other with a longer, less temperature-sensitive decay time. The phosphors are excited by laser pulses at a wavelength of 355 nm and luminescence is collected at 456 nm from one phosphor and at 485 nm from the other. The effects of motion on the temperature determination are removed by using a ratio of luminescence intensities of the two phosphors. Potential interferences from chemiluminescence, particles, hot surfaces, and reflections within the engine are avoided by limiting the detection to a narrow spectral window around the luminescence wavelengths. The data exhibit high signal-to-noise ratios and the measured temperatures are within the expected range for this engine.

Turbine Blade Mechanics Research Based on Strain Gauge Bonding Position

Aiming at the high-temperature strain gauge on the turbine blade of aero-engine, which needs to bear the high-temperature effect and reliably carry out the strain test, the turbine blades and adhesive strength of rubber layer of adhesive strain gauge were studied by the theoretical analysis, software simulation and experimental verification method. Three-dimensional model of turbine blade was built firstly. And then, the CFD of the blade fluid domain was analyzed. The bonding position of the strain gauge was determined under the actual working condition by the coupling analysis of hydrodynamics, centrifugal force, thermal stress and aerodynamic force. The reliability of strain position simulation of turbine blade was verified by dynamic strain contact measurement experiment. The effect of high temperature and CFD analysis of turbine blades with rubber layer was carried out. The effects of different rotational speeds on the adhesive strength of the adhesive were studied. It was shown that the strain and stress distribution of the blade’s pressure surface was more uniform than the suction surface.

Effect of fuel composition of LNG on the performances and operating characteristics of a gas turbine cycle

The effect of natural gas compositions on the thermodynamic performance characteristics of the single-shaft gas turbine cycle is analyzed in this study. The fuel characteristics are signified by H/C ratio and heating value, then the system are analyzed, and the full off-design analysis is performed by using TIT, TET, and power control. As the heating value of natural gas decreases, power output increases due to the increased mass flow passing through the turbine for all control methods. Additionally, regardless of the control method, the compressor surge margin is reduced and the turbine blade temperature is raised, making the gas turbine operation extremely vulnerable. Most parameters are sensitively affected by the HC ratio of fuel with TIT control. When the power control is used, gas turbine efficiency is reduced due to the increased operation of the compressor. The results are expressed by polynomial equations relative to each design value. In terms of the performance curves of the gas turbine cycle, the trends of most parameters accord with high H/C ratio and low heating value.

Turbine blade temperature error as measured with an optical pyrometer under different wavelengths and blade TBC thickness.

The optical pyrometer is a widely used instrument for measuring the temperature of gas turbine blade targets. However, one of the limitations of this instrument is the measurement error caused by reflected radiation from the high-temperature surrounding environment. This paper presents an analysis of such measurement error with consideration of a wavelength detection range of 2.5 to 22μm with blade thermal barrier coatings (TBCs) ranging from 18 to 279μm. Surface temperature distributions were simulated using the software ANSYS CFX. Radiation heat transfer was calculated based on discretized triangular surface elements. The variation characteristics of relative temperature error under different wavelengths and TBC thicknesses were obtained. The results obtained are well suited to provide theoretical guidance for choosing the wavelength of a pyrometer when measuring blade temperature with different TBC thickness, and can also be used to provide a method for estimating and correcting the temperature error.