Aiming at the efficient and uniform cooling requirements of high-pressure turbine blades, the flow and heat transfer characteristics of non-uniform and discontinuous sinusoidal ribs in typical U-shaped channels within the internal cooling structure are studied to achieve efficient heat transfer and enhance heat transfer uniformity through an irregular scheme. The results indicate that the uniformity of heat transfer along the channel is appropriately improved by non-uniform amplitude and rib height, and the heat transfer performance is enhanced by 21.9%. However, there is still a problem that overall heat transfer effect decreases significantly due to temperature rise of cooling air along the way in second channel. Based on this, a variety of discontinuous schemes of spoiler ribs are carried out to explore effective methods for heat transfer improvement. The results show that heat transfer performance of first channel is significantly enhanced by double-discontinuities scheme, overall heat transfer performance is improved by 49.7%, while the enhancement in heat transfer for the second channel is restricted. Based on the above research, the optimal strengthening of heat transfer performance is achieved by the coupling of non-uniform and double-discontinuous schemes. Under the premise of raising heat transfer performance by 33%, and phenomenon of heat transfer weakening in second channel of U-shaped channel is alleviated, thus achieving an effective balance between efficient cooling and uniform heat transfer.

Accurate prediction of discharge coefficients (Cd) in rotating orifices is essential for the design of aero-engine internal air systems, yet existing correlations usually treat axial and radial orifices separately and do not fully represent intermediate wall inclination angles. In this study, steady-state RANS simulations in a rotating reference frame, supported by validation against published data and by rotating orifice experiments, are used to investigate the combined effects of wall inclination angle alpha and length-to-diameter ratio L/d on Cd. The numerical results show that, under typical conditions (N = 3000 rpm, Pi = 1.03, L/d = 1.5), Cd increases from 0.301 to 0.340 as alpha increases from pi/2 to pi, corresponding to a 12.96% increase. Under low rotational speeds and high pressure ratios, the Coriolis force reduces the relative tangential velocity and the incidence angle, thereby increasing Cd with alpha; however, at high rotational speeds and low pressure ratios, the centrifugal resistance to radial inflow becomes dominant, and at N = 7000 rpm, the Cd for the alpha = pi orifice is 38.96% lower than that for the alpha = pi/2 orifice. Increasing L/d promotes flow redevelopment and amplifies the Coriolis-force effect, leading to a larger Cd increase for orifices with larger alpha. Based on these mechanisms, a generalized incidence-angle formulation incorporating Coriolis and centrifugal effects is developed, and a Cd prediction model applicable to pi/2 <= alpha <= pi and different L/d values is proposed. Experimental validation shows that the maximum prediction error is reduced to 2.37%, demonstrating the accuracy of the proposed model for rotating inclined orifices.

In this paper, an ultra-large groove with a depth of 31.2% blade height from a real turbine blade, is adopted as the baseline configuration. Winglet structures of varying widths are coupled to investigate the effects of the composite configuration on the aero-thermal characteristics of the turbine blade tip under different tip clearance heights and rotational speeds. The results indicate that compared with the baseline configuration, the composite tip structure improves stage efficiency (eta) by a maximum of 0.64% and reduces relative leakage mass flow by up to 14.29%. Moreover, the leakage suppression effect becomes more significant as the winglet width increases. The heat transfer optimization effect of the composite structure exhibits strong working condition dependence. Under small clearance and low-speed conditions, the winglet alleviates the impingement effect of TLF on the shoulder surface, reducing the shoulder surface Nusselt number (Nu) by up to 26.64%. However, at high rotational speeds and large clearances, the winglet enhances local flow separation, resulting in elevated heat transfer intensity. The wide-winglet composite structure is selected as the optimal configuration for low leakage performance. For composite tip structures targeting low thermal load, the specific tip clearance scale and rotational speed conditions must be fully considered.

This study provides a definitive experimental analysis of turbine squealer rim aerothermal performance under engine-relevant casing relative motion. By focusing specifically on the rim surface, this work demonstrates that the heat transfer on the rim surface is more sensitive; the rim is not only hotter than the squealer bottom, but its heat transfer is also significantly more non-uniform. This study systematically quantifies the effects of tip clearance (S), rim height (h), and rim width (w) on this local non-uniform heat transfer (where h, w, and S are all normalized by the blade height S), thereby revealing key design trade-offs. Results indicate that an increase in S is the primary driver of the rim's thermal non-uniformity, dangerously elevating the local maximum heat transfer coefficient (HTCmax). Furthermore, the study reveals a "selective blindness" in conventional metrics: the true efficacy of geometric parameters is only apparent when focusing on the rim's peak HTC, not the overall average HTC. For instance, increasing w from 1.0%S to 4.0%S is the most powerful method for mitigating hot spots (slashing the local HTCmax by 56.3%), but this comes at an aerodynamic cost. Conversely, increasing h provides a dual aerothermal benefit, reducing both thermal load and leakage. By navigating these complex trade-offs, this study recommends two optimized configurations: h = 10.0%S, w = 2.0%S for the best overall aerothermal balance, and h = 3.0%S, w = 4.0%S for achieving minimum thermal load. This provides a critical framework for mitigating localized ablation risk on turbine squealer rims.

In the turbine blade fractal cooling channels, developing shaped fractal units based on mass flow distribution is crucial for enhancing the cooling effect. Detailed analyses of flow and heat transfer in typical shaped fractal units are conducted in this study, and the performance differences of fractal plates with different shaped units are compared. Additionally, a fully automatic topology optimization process is developed to achieve fractal units' mass flow adjustment and performance optimization. It is indicated that the sensitivity of pressure loss in shaped fractal units to the position of typical control points is significantly higher than that of cooling effects. Within the same range of movement, maximum and minimum pressure losses differ by 279.2%, while cooling effects vary by only 7.3%. Outlet flow uniformity is significantly improved by adding branches to multi-entrance units. When optimization parameters target values are known, such as outlet flow distribution, the gradient descent method can be employed in the automatic optimization process. In contrast, for unknown parameters like minimum pressure loss, accuracy is significantly enhanced through the use of particle swarm optimization. The cooling air across different regions is redistributed by the shaped unit arrangement, thereby improving the fractal plate thermal uniformity.

This study presents an innovative design strategy for diesel engine cooling jackets oriented toward engineering applications. By employing a three-dimensional-to-two-dimensional-to-three-dimensional model simplification and reconstruction approach, the feasibility and design efficiency of topology optimization are significantly enhanced. Traditional jacket designs often suffer from non-uniform cooling and high flow resistance, while existing topology optimization research is largely limited to laminar flow conditions. To address this, the three-dimensional annular jacket is unfolded into a two-dimensional plane for turbulent topology optimization, substantially reducing computational cost, and then reconstructed into a directly applicable three-dimensional arc-shaped model. Penalty functions are introduced into the turbulent k‑ε equations to enable automatic generation and optimization of cooling channels. Results show that at Reynolds number Re=6855, the heat transfer coefficients of optimized configurations TTCPⅠ and TTCPⅡ increase by 9.51% and 19.05%, respectively, while TTCPⅠ reduces pressure drop by 8.46%, achieving synergistic improvement in heat transfer and flow performance. Notably, the optimized jackets exhibit excellent thermal stability. When the Reynolds number drops to 3428, the average temperatures of TTCPⅠ and TTCPⅡ are still 5.18% and 8.55% lower than the traditional jacket at Re=6855. Field synergy analysis further reveals that the optimized flow field exhibits better synergy between velocity and temperature gradients and improved synergy between velocity and pressure gradients, explaining the performance enhancement from a physical mechanism perspective. This work provides a practical technical pathway for the lightweight optimization of high-dimensional, high-cost turbulent cooling structures, offering clear engineering reference value for engine thermal management system design.

High-pressure turbine solid/internal cooling blades are selected as the research objects to systematically explore the laws governing blade thermal response characteristics under transient states with respect to cooling air, temperature rise rate, and temperature ratio. Specifically, the cooling air ratio are 0 %, 5.6 %, and 8.9 %; the temperature rise rates are 50K/60s, 50K/30s, and 50K/10s; and the temperature ratios are 1.25, 1.5, 2.0, 2.5, and 3.0. Increasing cooling air accelerates the heat transfer from mainstream to wall by strengthening convective heat transfer of inner cooling channel, which significantly inhibits the excessive temperature response of the blade surface, but is affected by mainstream temperature rise rate. At large temperature rise rate, cooling air cannot effectively take away the heat and has little effect on the thermal response. At small temperature rise rate, cooling air ratio increases from 0 % to 8.9 %, and temperature thermal response rate decreases by 10.6 %. Rising the temperature ratio, blade thermal response rate is faster, and blade heat load is higher. Reasonable internal cooling structure could effectively avoid the phenomenon. When the temperature ratio rises from 1.5 to 2.0, the overall temperature response rate of the uncooled blade increases by 24.86 %, but the internal cooling blade with a cooling air ratio of 8.9 % only increases by 13.82 %. It is evident that real-time matching design for cooling air ratio and temperature rise rate, taking into account the engine actual operating condition variations, is particularly essential.

To meet efficient uniform cooling demand for high-pressure turbine blades, a numerical investigation is conducted on flow and heat transfer characteristics of non-uniform sinusoidal spoiler ribs in cooling channel. Streamwise variations in geometric parameters amplitude, period, rib height and influence on cooling channel performance are elucidated to mitigate poor downstream heat transfer from streamwise cooling air temperature rise. With increase of the termination amplitude, disturbance of cooling air to near wall is significantly enhanced, and heat transfer performance is improved by 37.2 % compared with traditional uniform scheme. And large amplitude shows better heat transfer performance and does not depend on rib height. The best streamwise heat transfer uniformity is exhibited by moderate rib height and moderate termination amplitude, while high overall heat flux but poor uniformity is presented by large rib height and large termination amplitude. Compared with amplitude and rib height, the ability of non-uniform periodic arrangement to enhance heat transfer is relatively weak. When spoiler period varies drastically in streamwise direction, overall heat transfer reduces, but upstream flow separation is effectively suppressed, significantly reducing pressure loss. Among these, the thermal performance of large rib height and moderate periodic fluctuation scheme is the greatest. On this basis, the optimized arrangement of amplitude combination rib height increases streamwise both improves Nu/Nu0 of downstream ribbed wall, and cooling air diversion effect suppresses pressure loss. Specifically, increasing rib height of streamwise at small amplitude relieves downstream heat transfer deficiency, and enhances the end heat transfer at large amplitude.

When the aero-engine is operating in a dusty service environment, it is inevitable that particles such as sand and dust will invade and deposit on the surface of high-temperature components such as turbines and the internal flow channels of the air system, which will lead to a surge in the risk of various faults. However, there are still few studies on particle deposition in the radial pre-swirl system of aero-engine, and the understanding of its internal deposition mechanism is still unclear. Therefore, this paper studies the deposition and transportation of particles in the typical engine radial pre-swirl system by numerical method, and focuses on the influence of particle diameter, rotor rotational angular velocity and inlet pre-swirl nozzle angle on particle deposition characteristics. The results show that the high-speed air system plays a role similar to that of a separator for large-size particles (> 2 mu m), thus preventing them from reaching the blade root. For smaller particles (< 1 mu m), increasing the rotational angular velocity usually improves the followability of the particles with the flow, thereby increasing the capture efficiency of the particles reaching the system outlet. While for the large particles (>= 1 mu m), the opposite trend is observed. Increasing the pre-swirl angle will significantly increase outlet capture efficiency. When the pre-swirl angle is small, the separation effect of the particles is better, which helps to reduce the risk of erosion and clogging on downstream components.

A real turbine stator is selected as the research object, and cooling performance and thermal stress characteristic of blades with varying film cooling hole geometries-namely circular holes, shaped holes, and 7-7-7 holes-are studied through high temperature experiment and numerical simulation. Compared to circular hole blade, the cooling efficiency of shaped hole blade is improved by 12.99 %, the overall thermal stress is reduced by 17.12 %, and the thermal stress in the maximum 10 % interval is reduced by 24.34 %. It is an optimal scheme for the film holes with great cooling efficiency and low thermal stress. The ability of the shaped hole and 7-7-7 hole blades to improve blade cooling efficiency and reduce surface thermal stress is similar, but the former is slightly better at reducing thermal stress in the blade root and suction-side film hole outflow regions. The aerothermal parameters under different engine operating conditions impact the ability of the shaped hole blades to enhance the cooling efficiency and lower the surface thermal stress. Improvement of blade cooling efficiency is more significant at small temperature ratio and small blowing ratio, and the reduction of the blade surface thermal stress is more obvious at large temperature ratio or large blowing ratio. As the blowing ratio varies along with the engine operating conditions, thermal stress on blade pressure side and suction side with shaped holes are obviously reduced, and the drastic changes of thermal stress in the local area of blade leading edge are also effectively mitigated.

Fractal channels, characterized by high surface-to-volume ratio and geometric flexibility, demonstrate significant potential in turbine blade cooling design. This study experimentally and numerically investigates a turbine blade equipped with a tree-like bionic fractal channel under multi-parameter coupling working conditions, focusing on the effects of blowing ratio, temperature ratio, and mainstream Reynolds number. The results indicate that the three parameters independently influence blade cooling effectiveness with minimal coupling interaction. As temperature ratio increases, the sensitivity of cooling effectiveness to blowing ratio amplifies significantly. Cooling effectiveness improves with higher blowing ratio and lower reynolds number, showing the greatest responsiveness to blowing ratio variations. Consequently, optimal design of bionic fractal cooling architectures must account for operational parameters. The spatial distribution density of tree-like bionic fractal channels at the mid-chord position notably affects localized cooling characteristics, with a 6.5 % enhancement observed in the test region upon increasing channel number. A systematic reduction in fractal channel inlet diameter along the chordwise direction regulates cooling air mass flow rate to match spatially varying heat loads, resulting in a 150 % disparity in internal wall heat transfer coefficients between the blade leading and trailing edges. These findings confirm that fractal channel configuration and geometric parameters are primary factors influencing cooling effectiveness distribution through localised flow modulation, establishing them as critical optimisation targets.

To meet the urgent demand for high-efficiency thermal protection of next-generation gas turbine blades, this study experimentally investigates the cooling performance of porous medium turbine blades under single-phase air transpiration cooling and liquid water phase-change transpiration cooling (PCTC) conditions, with a focus on the topological optimization of the blade front cavity flow channel. The results show that particle size exerts a significant effect on both cooling modes: small particle size leads to high flow resistance and poor cooling uniformity, while large particle size effectively improves the heat dissipation efficiency of porous medium blades. PCTC exhibits superior cooling performance to single-phase transpiration cooling, with the cooling effect of the blade suction side increased by approximately 18% by virtue of the latent heat of vaporization of liquid water, yet local high-temperature zones and uneven cooling caused by gravity-induced liquid pooling still exist. To solve this critical problem, a split-chamber flow channel topological structure for the blade front cavity is proposed and optimally designed. This optimized structure effectively suppresses the liquid pooling effect, realizes uniform coverage of the cooling working fluid on the blade surface, and increases the heat dissipation efficiency of the blade pressure side to 0.60, thus reducing the cooling non-uniformity of each blade region and the local high thermal stress. The research findings provide a feasible structural optimization scheme for the engineering application of PCTC in gas turbine blades.

To address the thickness reduction and thermal-insulation degradation of thermal barrier coatings (TBCs) on turbine-blade surfaces caused by ingested particle erosion, this study extends the classical Neilson particleerosion model by explicitly incorporating the dependence of coating Young's modulus on temperature and the dependence of Vickers hardness on both temperature and service time. The critical velocity and erosion parameters are dynamically updated and then used to analyze coating thinning, surface-temperature variation, and deterioration of thermal-insulation efficiency at the blade scale. On this basis, the effects of particle size and erosion duration on coating thickness reduction, blade-surface temperature distribution, and degradation of TBC thermal-insulation efficiency are numerically investigated. An engineering-oriented analytical correlation for insulation-efficiency degradation is further established using the Buckingham Pi theorem and Lasso regression. The results show that the relative error between the present erosion predictions and representative TBC erosion experiments is 4.77%, whereas the mean relative error between the analytical correlation and the simulation results is 7.16%. The study integrates a classical erosion model, material-property evolution relationships, and blade-scale thermal-insulation analysis into an engineering framework, thereby providing methodological support for protective design and performance assessment of TBCs under particle-erosion conditions.

Aiming at the issue that stress is concentrated at the leading edge of the gas turbine guide blades, a combination of numerical simulation and experimental measurement is employed in the paper to investigate in detail the stress–strain characteristic of a certain type of turbine blade with five different leading edge film hole arrangements, along with optimization design. Compared to the prototype blade, increasing the number of leading edge film holes and hole rows results in a maximum average stress reduction of 5.38 % and 7.58 % in the top 10 %, respectively. It is superior to other schemes in reducing surface temperature and thermal stress of the blade, which is the preferred outcome for low stress cooling design in the study. Appropriately increasing the number of holes and rows of film holes is beneficial to reduce leading edge high strain range and overall stress, with more significant effects observed on the leading edge pressure side. Excessive number of holes and arrangements lead to uneven cooling air distribution, an overall increase in blade stress as well as localized high stress. Compared with circular film holes, shaped holes with a larger film coverage area demonstrate better cooling performance, exhibiting lower strain and stress, making them the preferred hole type for pressure side film cooling. The streamwise strain of blade leading edge is the primary strain direction and increases significantly with rising inlet temperature.

The paper investigates the aerodynamic characteristics of turbine blades with three different surface features under five attack angles and three outlet Mach numbers (Ma) through experimental and simulation approaches. It has been discovered that the surfaces of ceramic matrix composite (CMC) blades fabricated via a specific process and those of ordinary metal blades are extremely smooth, with an exceedingly small difference in surface roughness, leading to a high level of consistency in surface static pressure distributions. For the texture blade, however, the pressure difference between the pressure side and suction side is significantly reduced. At attack angles of 0°, 7°, and 15°, cascade outlet's average total pressure loss coefficients for the texture blade are 7.5% and 13.2%, 7.1% and 6.4%, and 5.1% and 9.2% higher than those of the CMC and metal blades, respectively. This is attributed to the continuous interaction between the airflow and texture features, which disrupts the boundary layer flow, induces vortices, and exacerbates flow losses. Additionally, three types of blades show similar distributions of total pressure loss coefficient along the blade height. For the CMC blade, a larger positive attack angle increases the surface pressure difference and total pressure loss coefficient, whereas a larger negative attack angle produces the opposite effect. The research further confirms that the CMC blade processed by the specified technology demonstrates excellent aerodynamic performance, characterized by a low overall total pressure loss coefficient under high Ma conditions.

This paper addresses the lack of analysis on the impact of blade tip clearance and the insufficiency of blade tip clearance calculation tools in the context of ceramic matrix composite (CMC) shroud used in aircraft engines. To address these issues, a coupled fluid-thermal-solid multi-field prediction model accounting for the physical properties of CMC was established for blade tip clearance, and the thermal-deformation of the shroud under a prescribed temperature boundary was verified using a specialized experimental device. The computational analysis found that due to CMC's low thermal diffusivity and high contact thermal resistance with metals, the conduction of mainstream high-temperature heat into the engine casing was hindered, resulting in a stator thermal deformation response speed that is, on average, about 210.88% of that of traditional metal shrouds. At the same time, mainly because of CMC's low coefficient of thermal expansion, the matching degree of the linear expansion coefficient between the rotor and stator decreased by 54.11% when using CMC shrouds compared to traditional metal shrouds. This caused the minimum blade tip clearance during the transition state to be too small and the clearance range in blade tip clearance to be too large. Therefore, this paper further combined the matching degree of linear expansion coefficients and proposed an optimized blade tip clearance matching method, which successfully increased the minimum blade tip clearance and reduced the clearance range in blade tip clearance under CMC shroud application. Related research reveals the evolution mechanism of blade tip clearance and its key influencing factors under CMC shroud application conditions.

Aero-engine requires maintaining high rotational speeds for warm-up operations prior to takeoff; otherwise, thrust degradation may occur during the climb-out phase. To address this issue, this study focuses on a core engine of a small-bypass-ratio turbofan and develops a fluid-thermal-solid coupling computational methodology that integrates a 1D fluid domain model with a 2D axisymmetric finite element model of the solid structure. This approach enables coupled simulation across different scales for solid domain analysis, air system evaluation, multi-component thermal analysis, and deformation prediction. Simulation results reveal that tip clearance exceeding design specifications constitutes the primary cause of thrust reduction during takeoff. Warm-up operations effectively reduce tip clearances, resulting in a 0.85% improvement in high-pressure turbine efficiency and a 0.5% enhancement in compressor efficiency. These improvements drive a maximum 0.51% increase in high-pressure rotor speed, augment core mass flow, reduce engine bypass ratio, and ultimately elevate minimum takeoff thrust by 2.4%. Further analysis demonstrates that either extending warm-up duration or increasing warm-up rotational speed can enhance takeoff thrust. However, when the warm-up speed remains below 0.75, increasing rotational speed fails to effectively reduce high-pressure turbine tip clearance. Finally, leveraging sequential quadratic programming optimization, a warm-up parameter optimization model constrained by thrust requirements is established to achieve optimal combinations of minimum warm-up time and rotational speed across various operational conditions.

Ceramic matrix composite (CMC) is increasingly applied to high-temperature components of aero-engines, due to their excellent high-temperature resistance. However, molten particle deposition (sand/dust) in complex thermogas environments degrades surface heat transfer and cooling performance. Distinct from traditional metal structures, CMC's high service temperature (approximately 1600 K) and heterogeneous rough surface induce unique deposition behavior. This paper takes the flat-plate film cooling structure of 2.5D braided CMC materials as an example to investigate the particle deposition characteristics of CMC film cooling structures under the influence of fluctuating surface temperature fields of braided CMC components. Additionally, low-melting-point particles were used in a laboratory environment to conduct model validation experiments for particle deposition in CMC film cooling structures. The research results indicate that for the flat-plate film cooling structure of CMC materials, the deposition distribution correlates with the internal woven structure of the CMC material and the deposition amount first decreases and then increases with the increase of the blowing ratio. When the surface roughness exceeds Ra 6.3, the deposition thickness increases significantly. Therefore, in engineering applications, the surface smoothness of CMC hot-end components should be improved as much as possible. Based on this, an integrated co-design of active anti-deposition cooling structures and CMC material braided preforms was carried out, leveraging the designability of CMC material braided structures for high temperature airflow environments. The mechanism of the active anti-deposition cooling structure was analyzed and compared with traditional inclined round holes. Based on comprehensive analysis, it is concluded that the fan-shaped hole with a 10 degrees expansion angle exhibits superior cooling effectiveness and deposition resistance. Specifically, its average deposition thickness is reduced by 35.1% compared to that of the inclined circular hole configuration.

In this paper, experimental and numerical methods are used to research the influence of outer ring cooling air-coupled ultra-large grooves on tip aero-thermal characteristics of small aero-engines. The results show that the relative error between the average Nu measured by TSP experiment and the prototype blade tip in numerical calculation is <5%, and the variation trend is basically the same. 25.35% blade height grooves optimize leakage suppression and tip thermal load reduction. Single-row film holes exert the most significant effect on suppressing leakage flow, while three-row and five-row film holes are more effective in diminishing the thermal load on the shoulder surface. For three-row film holes, a 'gas barrier' is formed under large cooling air flow and high blowing ratio, which weakens leakage vortices, reduces shoulder surface thermal load and improves outer ring cooling efficiency significantly, the aero-thermal characteristics are better than single-row arrangement. Considering the requirements of low leakage and low thermal load for aeroengine rotor, the combination of 'ultra-large groove + single-row holes + large cooling air flow' owns the low leakage, while 'ultra-large grooves + three-row holes + large cooling air flow + high blowing ratio' produces the low thermal load.