Hydrogen energy is essential to establishing a sustainable and reliable energy system. The continuously growing demand for hydrogen is driven by the challenges associated with greenhouse gas emissions and resource depletion. This paper investigates and summarizes some intriguing hydrogen production processes that have evolved from laboratory stages to mature commercial applications. The analysis of techno-economic, environmental effects and investment trends of these processes are included. Currently, hydrogen is dominantly produced by methods with fossil fuels as feed. These technology processes are relatively mature and account for the majority of the world's hydrogen production, around 99%. However, these results in significant carbon emissions. Around 1400 million tons of carbon dioxide are emitted into the atmosphere. To achieve carbon neutral strategy, the hydrogen production from hydrocarbon fuels needs to become clean. Equipping carbon capture, utilization, and storage system is a promising way to reduce carbon emissions. In addition, hydrogen production schemes with zero carbon emissions like electrolytic and photocatalysis are attracting increasing attention. The survey results suggest that the price of hydrogen production associated with the addition of carbon capture equipment ranges from 1.47 to 6.04 USD/kg, which is higher than the value for the price without the additional facility (1.03–2.08 USD/kg). The introduction of carbon tax is expected to narrow the cost gap between the two. Besides, the cost of electrolysis remains expensive (6.25–12.2 USD/kg), depending on the energy source and electrolytic cell equipment. The high-pressure autothermal reforming technique coupled with carbon capture and electrolytic technique powered by renewable energy are favored by global commercial investment. Finally, key challenges and opportunities for clean hydrogen production are discussed in this paper. More attention should be paid to catalyst blockage or deactivation and the cost of carbon capture equipment for fossil fuel hydrogen production. For the new zero-carbon hydrogen production method, designing efficient, economical catalysts and electrolysis materials is essential for its large-scale application.

A numerical investigation is conducted to uncover the parametric influence of configuration geometry on the thrust performance of annular expansion deflection (ED) nozzles. Based on the classic design principle of Taylor’s ED nozzle configuration, the influences of six geometric elements, covering the expansion channel region, the near-pintle region, and the shroud region, including 13 nozzle configurations, are examined in detail. The flow characteristics in each nozzle are demonstrated, according to which the connections between the geometric changes and nozzle thrust performance are elucidated. The present results show that the nozzle flow pattern is closely related to the nozzle configuration geometry. In the open operation mode of the ED nozzle, the wide expansion channel has very little restriction on the axial expansion of the exhaust gas. The high axial velocity brings strong shock strength and total pressure loss, which are unfavorable to the nozzle thrust performance. The large curvature of the shroud expansion section contributes greatly to the exhaust gas deflection, which increases the mass flow rate of the supersonic core flow zone through the nozzle exit plane, and therefore favors the thrust performance. In the closed operation mode, geometric differences in the expansion channel region have little effect on the supersonic gas, which fills almost the entire nozzle. The shroud region still affects the axial deflection of the gas and its large curvature is associated with superior thrust performance. These investigations suggest that the annular ED nozzles with narrow expansion channels and large shroud curvatures are superior in thrust performance.

An iterative dynamic chemical stiffness removal method (IDCSR) based on quasi-steady-state approximation (QSSA) is proposed. The IDCSR method is built on a previously developed non-iterative method which has proved to work well for small timestep sizes. A novel iterative procedure is designed in IDCSR to enable explicit time integration of stiff chemistry at relatively large timestep sizes relevant to practical reacting flow simulations. The effectiveness of the iterative procedure is first demonstrated with a toy problem and homogeneous auto-ignition with fixed integration step sizes, showing that larger timestep sizes can be allowed for explicit time integration using IDCSR compared with the previous non-iterative method. IDCSR is then compared with existing explicit chemistry solvers for simulations of homogeneous auto-ignition and shows similar or lower computational cost but significantly higher accuracy across a wide range of timestep sizes. IDCSR is further combined with an automatic adaptive time-stepping scheme for simulations of 0-D homogeneous auto-ignition and a 2-D laminar lifted n-dodecane jet flame. For the 0-D auto-ignition simulations, IDCSR is shown to reduce both the error (by 43%–90%) and computational cost (by 6–15 times) compared with existing explicit solvers, while achieving speed-up factors of up to 400 compared with VODE for a wide range of timestep sizes and reaction mechanisms. For the 2-D jet flame simulations, speed-up factors of 15 and 31 for chemistry integration, and 5 and 9 for overall simulation, are achieved by IDCSR compared with CVODE with and without analytic Jacobian, respectively.

The performance prediction accuracy of quasi-3D design is highly dependent on the empirical correlation model. The state of art model from experimental data is hard to meet the request of new blade development and higher aerodynamic design demand. In order to solve the problem by ignoring the influence of blade thickness position on the Reynolds number correction model, numerical research based on the middle section of the low-pressure compressor rotor is conducted. The sample database is constructed based on a uniform sampling method considering the incidence, thickness, and Reynolds number at a certain inlet Mach number, and the new model to design and off-design condition is both established by multiple nonlinear regression and adaptive simulated annealing algorithm. The results show that the root-mean-square error of the total pressure loss is reduced by 78.92% and 60.22%, error of the deviation angle is reduced by 78.44% and 78.56%, under design and off-design incidence angles respectively. The new model can provide more reliable predicted results for modern compressor design and optimization.

Accurate and reliable combustion state monitoring is a key requirement for the development of future rocket-based combined cycle (RBCC) engines. The rapid advancements in deep learning technology have rendered data-driven combustion state sensing a possibility, thus contributing to the realization of intelligent and efficient combustion organization. This paper proposes a multi-path convolutional neural network model that is suitable for the reconstruction of two-dimensional temperature fields. The impact of diverse model architectures on the precision of reconstruction outcomes is examined. The results of the reconstruction of the entire test set demonstrate that the MPFC-CNN model exhibits superior accuracy and extrapolation generalization ability compared to the SP-CNN and MPU-CNN models. The overall test dataset demonstrates an average reconstruction error of 2.84%, a linear correlation coefficient of 0.9901, and a structural similarity index of 0.8842. Validation of the reconstruction was conducted for additional combustor temperature fields with varying strut placements. The reconstruction results also basically largely satisfied the requirements. Additionally, the MPFC-CNN model has fewer parameters, which can provide a reliable basis for combustion state recognition and monitoring.

Variable area turbine vanes are critical components for achieving adaptive cycles in aero-engines, essential for balancing high thrust and efficiency. The variable staggered angle turbine vanes pose challenges in achieving efficient cooling, thereby limiting their applications in high-pressure turbines. Suction-side adjustable turbine vanes can integrate efficient cooling with excellent aerodynamic performance design, potentially solving the challenge of adjusting high-pressure turbines. This paper was based on the VKI LS89 vane and constructed suction-side adjustable turbine vanes. Using a turbulence model named SST k–ω, it analyzed the variations in parameters such as Mass flow rate, aerodynamic losses, and flow angle at different vane openings. It also compared and studied the flow regulation capabilities and aerodynamic loss characteristics between suction-side adjustable turbine vanes and variable staggered angle turbine vanes. The results showed that within the adjustment angle range of 0° to 3°, the suction-side adjustable turbine vane had a flow regulation capability of 10.8%, while the variable staggered angle turbine vane had 15.0%. Considering leakage, the total pressure loss of the suction-side adjustable turbine vane with coolant leakage was much smaller than that of the variable staggered angle vane with end-wall gaps, amounting to only 60% of its total pressure loss. When the vane rotated by 3°, the change in the outlet flow angle of the suction-side adjustable turbine vane was only 1.5°, whereas, for the variable staggered angle vane, the change in the outlet flow angle at the blade tip and root regions was 15°. Additionally, coolant leakage near the blade tip and root regions of the suction-side adjustable turbine vane achieved better cooling efficiency.

Electric vertical take-off and landing (eVTOL) aircraft with ducted fans have high efficiency, high reliability, and low noise, and have great potential to become the majority of future urban air mobility. Nonetheless, the ground effect will cause fluctuations in ducted fan thrust during take-off and landing conditions. Conventional physical models fail to reflect its transient and non-linear nature. In this paper, a physics-inspired-data-driven (PIDD) model for predicting the in-ground-effect (IGE) thrust of ducted fans is presented. The total thrust is treated as the composition of average thrust and transient thrust. For the average part, hypotheses of ideal twist and exponential function are adopted based on the structure of blade element momentum theory (BEMT). For the transient part, a classification task based on cross entropy is added to the meta-learning algorithm. Altitude related and altitude unrelated items are constructed to obtain the PIDD model and enhance its generalization ability. The features of the model are rotary speed, voltage, and current, while the label is IGE thrust. Few-shot training indicates that the proposed model can predict transient thrust accurately. With mean error on training and testing sets is 1.1% and 0.7% respectively, the PIDD model outperforms the conventional physical model by 4.6% and 1.7% and the data-driven model by 1.5% and 2.7%. The PIDD model constructed succeeds in predicting transient IGE thrust and provides new ideas for relevant research.

Micro gas turbines have broad application prospects because of their compact size, low emissions, fuel flexibility, and high reliability. This study focused on a typical reverse flow micro gas turbine combustor and designed six overall structural schemes of micro gas turbine combustors with different evaporation tube arrangements and lengths. Numerical simulations and combustion experiments were also conducted. The data show that a structural scheme in which the evaporation tube is at an angle of 45° to the combustion liner axis provides the best performance. This scheme allows for the formation of multiple swirling regions at the head of the combustion liner, decreasing the flow velocity at the head of the combustion liner from 70 to 20 m/s, thus effectively reducing the flow velocity; in addition, the combustor can organize combustion at a higher inlet flow velocity. This scheme shows the best combustion performance, and the combustion efficiency reaches more than 98%. Increasing the length of the evaporation tube strengthens the circumferential connection and performance of the combustor flame. This also results in an increase in the jet velocity exiting the evaporation tube. The higher jet velocity increases the velocity at the head of the combustion liner. This change affects the temperature distribution within the combustion liner. Specifically, it leads to an increase in the temperature gradient inside the liner.

The efficient mixing of fuel and oxidizer in a scramjet combustor is a critical issue for achieving future wide-speed-range flight. A numerical study is conducted using Reynolds-averaged Navier–Stokes equations coupled with shear stress transport k-ω turbulence model to investigate the mixing characteristics of hydrogen and air in a wide-speed-range supersonic combustor. Validation case using the numerical simulations shows remarkable consistency with experimental observations, confirming the reliability and accuracy of the computational approach. The results indicate that during a wide-speed-range flight, the enhancement effect of shock waves on the growth of the mixing region thickness within the combustor persists. This is attributed to the baroclinic torque and volumetric expansion effects caused by the interaction of shock waves and the turbulent shear layer, which enhance vorticity. When the combustor entrance Mach number is relatively low, this enhancement effect is more pronounced. As the combustor entrance Mach number increases, the enhancement effect gradually decreases. The area of the fuel–air mixing region decreases significantly as the combustor entrance Mach number increases, with this reduction being more pronounced at a low equivalence ratio compared to a higher one. The mixing efficiency of the fuel–air decreases with increasing combustor entrance Mach number. At a low equivalence ratio, a higher combustor entrance Mach number delays the location of full mixing. At a higher equivalence ratio, increasing the combustor entrance Mach number may result in the fuel and air remaining incompletely mixed by the time they reach the combustor exit.

With the rapid advancement of computing power and significant progress in machine learning (ML) algorithms, ML has shown immense potential across a wide range of fields, particularly in simulation, experimental analysis, and condition monitoring. In the realm of combustion science and engineering, ML techniques have driven substantial advancements, particularly in gas turbine engine combustion. The integration of intelligent algorithms in combustion modelling and analysis has led to more accurate predictions, enhanced performance, and the potential for more efficient and environmentally friendly combustion processes. This paper provides a comprehensive review of the current state of research on intelligent algorithms applied to combustion chemical reaction kinetics, combustion simulation, performance prediction, combustion state, and instability monitoring. The review highlights the progress made and offers valuable insights for improving the performance of gas turbine engines. Additionally, the paper discusses the challenges and prospects of applying intelligent algorithms in combustion research, including issues related to data quality, model interpretability, and computational complexity, while identifying avenues for future development and innovation.

Fractal dimensions serve as a metric for evaluating the complexity of turbulent structures, thereby inferring the mixing efficiency of the jet and the mainstream. In this paper, the flow field structures of the interaction between supersonic flow and dual jets were obtained by the nanoparticle-based planar laser scattering (NPLS) technique with high spatiotemporal resolution. Based on the NPLS images, the fractal analysis was conducted. The Canny edge detection algorithm was utilized to delineate turbulent interfaces, and five subregions were defined to investigate the mixing efficiency at different streamwise positions. The results indicate that the 2D average fractal dimension of the entire interaction flow field ranges from 1.4 to 1.61. The larger orifice spacing and greater distance of the dual orifices from the leading edge of the plate, the higher the fractal dimensions. Consequently, it could be inferred that positioning the dual orifices further downstream with greater orifice spacing effectively enhances the mixing efficiency in the interactional flow. Furthermore, the analysis of streamwise subregional fractal dimensions reveals that the interaction between the dual jets and the supersonic mainstream exhibits an “increasing and decreasing” tendency in mixing efficiency along the streamwise direction.

Tip leakage loss is a significant source of aerodynamic loss in the turbine rotor and can be controlled by using a cavity tip. In this paper, three configurations with cutback rims are numerically investigated to control the formation and development of vortices in the tip region. The results indicate that the suction side cutback squealer tips effectively suppress the development of the upper passage vortex (UPV) and its associated losses. The suction side cutback rim increases the leakage rate and strengthens the leading-edge tip leakage vortex (LTLV), which in turn intensifies the interaction between the LTLV and the UPV, thereby inhibiting the development of the UPV. The pressure side cutback rim has a lesser impact on the LTLV but significantly enhances both the tip leakage vortex (TLV) and the UPV, which is unfavorable for controlling local losses in the tip region. When the suction side rim cutback is slightly downstream, it can generate a stronger LTLV for suppressing the development of the UPV. This configuration could effectively reduce the total pressure loss coefficient in the range of 50%–80% span by up to 28%, although it also leads to an increase above 80% span. Overall, the suction side cutback squealer tip could still reduce the average total pressure loss coefficient in the region above 50% span by approximately 5%.

This study investigates the mixing field of a 300MW F-class heavy-duty dry low NO_x (DLN) combustor using PLIF experiments and numerical simulations. Innovative optical schemes, shading designs, and tracer gas generation systems were developed to facilitate the successful execution of the experiments. Numerical simulations assessed various turbulence models by analyzing velocity distribution and mixing characteristics. The results indicate that the combustor’s headend design exhibits strong mixing capabilities, with consistent mixing field profiles observed under varying flow conditions and identical equivalence ratios. Among the turbulence models, large eddy simulation (LES) most accurately reproduced experimental results, especially in terms of velocity distribution, while Reynolds-averaged Navier–Stokes models with a default turbulent Schmidt number of 0.7 significantly underestimated the mixing rate. Additionally, reducing the turbulent Schmidt number enhanced the mixing rate, with a value of 0.2 in the Realizable k-ε model providing results closely aligned with experimental and LES findings. The experimental and numerical methodologies presented in this study provide valuable insights for future research on mixing phenomena in similar combustor designs. Future work may focus on exploring the complex flow and mixing mechanisms within the premixing tube of DLN combustors.

The explicit finite element (FE) simulation method of the flexible rotor with fan blade off (FBO) is developed using LS-DYNA. Three important aspects of the model processing have been discussed, including the setting of blade off and rotational speed, simulation method of bearings, and solution time reduction technologies. An inner cross method is developed to simulate the elastic bearings, which can effectively avoid the problems of the existing simulation methods. Based on the established explicit FE model, the dynamic response, stress distribution characteristics, and impact energy propagation of the shaft are studied after the FBO fault occurs. The numerical results show that the impact energy of the missing blade does not propagate as a wave in the rotating shaft, which is different from the non-rotating beam. The gyroscopic effect can inhibit the typical wave propagation characteristics of impact energy. The bending moment of the rotating shaft is determined by both the gyroscopic moment and the unbalanced load, while the unbalanced load is the dominant factor. Finally, it is analyzed that key factors such as rotational speed, unbalance, and the constraints of fusing structure and fan casing have different effects on the dynamic response of the rotor. The energy concentration phenomenon and the amplification effect of reaction force appear in the rotor with blade off.

The heat transfer performance of the thermal management system plays a crucial role in the hydrogen-powered aviation engine cycle. As an exceptional fuel, the thermophysical parameters of hydrogen change drastically with temperature in the trans-critical state. While previous studies on heat transfer enhancement mainly focused on changing the geometrical structure, few studies have been conducted on realizing heat transfer enhancement based on the properties of the fluid itself. Utilizing the drastic changes in thermophysical parameters of hydrogen in the trans-critical state to achieve heat transfer enhancement could greatly contribute to the thermal management system of the hydrogen-powered cycle. In this study, a trans-critical process control method for heat transfer enhancement based on multidirectional impact flow distribution is proposed. The distributions and variation patterns of temperature, density, specific heat capacity, and equivalent thermal conductivity along the flow directions were investigated, the flow and heat transfer performance of the channel optimized by the proposed method was numerically simulated, and the control of the trans-critical process and the mechanism of heat transfer enhancement were analyzed. The effects of the key design parameters such as flow distribution ratio, number, and spacing of gaps on the flow and heat transfer performance of the heat transfer unit were comparatively analyzed by taking various factors into account, and finally, a relatively optimal combination of key design parameters was obtained.

Regenerative cooling technology is an important approach for thermal protection in scramjet engines. Non-uniform axial and circumferential heat flow distribution on the combustor leads to improper flow distribution, resulting in reduced cooling efficiency and localized overheating, potentially causing thermal protection failure. To effectively utilize the fuel heat sink and enhance cooling efficiency, it is necessary to investigate the mechanism of flow distribution in the channels under different heat flow conditions. This study investigates the flow and heat transfer characteristics in single cooling and parallel channels under various heat flux. The heat transfer characteristics, flow resistance changes, and the influence of heat flux on heat transfer deterioration are analyzed. There are two types of heat transfer deterioration in channels, caused by the inlet effect and the mutation of fuel supercritical properties. Furthermore, the mechanism behind flow deviation induced by non-uniform heat flux in parallel channels is investigated. When a supercritical process occurs, non-uniform heat flux induces temperature deviations, resulting in density distribution variations. This, in turn, influences velocity distribution and leads to disparities in flow resistance distribution. Consequently, flow deviation occurs, and the cracking reaction amplifies flow deviation while reducing temperature deviation.