Oblique detonation engines, with their significant advantages such as rapid heat release, high specific impulse, and compact combustion chamber structure, have become a research hotspot in the field of hypersonic propulsion. However, the extreme thermal and mechanical loads encountered during operation pose severe challenges to the thermal protection system and structural reliability. Film cooling technology using large-molecule hydrocarbon fuels as the coolant is regarded as a key approach to overcoming this bottleneck. This paper focuses on the combustor wall of an oblique detonation engine and conducts a numerical simulation study on the film cooling and combustion drag reduction characteristics of hydrocarbon fuel (C10H20), employing the SST k-ω turbulence model and the finite-rate chemical reaction model. The effects of injection position, mass flux ratio, and incoming flow equivalence ratio on cooling and drag reduction performance are systematically analyzed. The study sets the incoming flow velocity at 2300 m/s and the fuel injection velocity at 77 m/s. Results indicate that the injection position significantly alters the detonation initiation characteristics and wavefront cellular structures, with cooling and drag reduction efficiencies maintained above 40% and 15%, respectively. Under a fixed mass flux ratio, cooling effectiveness increases nonlinearly with increasing slot height, while drag reduction efficiency decreases accordingly. Increasing the mass flux ratio can substantially enhance cooling effectiveness up to 75%, but drag reduction performance remains only around 20%. An increase in the incoming flow equivalence ratio strengthens combustion intensity without significantly improving cooling or drag reduction effects.

The implementation of internal cooling through a hollow shaft is recognized as an effective approach for managing thermal challenges in gas foil bearings (GFBs) operating at ultra-high rotational speeds. This study conducts a comprehensive numerical parametric comparison of several representative internal cooling configurations applicable to radial GFBs, including axial throughflow, radial injection, and radial impingement schemes. Numerical simulations are conducted to systematically evaluate and compare the flow characteristics and cooling mechanisms of each configuration at a rotational speed of 1.5 & times; 10 5 r/min, under a constant static bearing load of 45 N, with the cooling flow mass flow rates ranging from 0 to 15 kg/h. Results indicate that among all internal cooling configurations, the radial impingement design with the transmission pipe provides superior cooling effectiveness by directly targeting high-temperature zones, but it incurs the highest aerodynamic penalty. The single-nozzle configurations generally achieve better cooling performance and surface temperature uniformity compared to multi-nozzle designs, though at the cost of higher supply pressures. Although axial throughflow requires the lowest supply pressure, it consistently shows the weakest cooling performance and temperature uniformity across most operating conditions. Among all configurations, the multi-nozzle radial impingement scheme provides the best overall trade-off between cooling effectiveness, thermal uniformity, and pressure requirements within the examined mass flow rate range, making it a promising solution for practical thermal management in high-speed GFB systems.

Atmospheric particulate suction inevitably causes particle deposition on the surface of turbine blades, degrading the heat transfer of cooling structures and aero-engine performance. This study investigates particle deposition and its influences on film cooling performance in blades with round and round-to-slot film holes. The discrete phase model coupling with the EI-Batsh model simulates particle motion and deposition, while dynamic mesh reconstructs the evolving deposit morphology. The film cooling performance and particle deposition is strongly influenced by hole geometry and blowing ratio: higher cooling efficiency generally reduces deposition, with slotshaped holes showing lower accumulation than round holes, especially at higher blowing ratios. The decline of particle deposition rate employing slot-shaped film holes reaches to nearly 30% when the blowing ratio is 2.0. The deposition layer formed under the round film hole is significantly thicker than that under the slot-shaped film hole, with a maximum thickness of approximately 0.2 mm observed in leading and trailing edge. The deposition layer generally elevates the blade surface temperature and degrades the film cooling performance, but exerts complex effects on cooling efficiency in different regions of blade surface.

Accurate assessment of rotating disk cavity performance is critical to the effective design of secondary air systems. The traditional fluid network method offers the advantage of fast computation but fails to meet the increasingly stringent requirements for high precision. While computational fluid dynamics (CFD) enables fine-grained modeling, its complex modeling and calculation processes hinder its application in rapid design iterations. To address this gap, this study realizes refined and rapid modeling of flow and heat transfer inside rotating cavities with vortex-reducing tubes, utilizing a lightweight radial basis function neural network (RBFNN). To overcome the challenge of lightweight neural networks in handling large-scale datasets, an innovative one-to-one grid prediction strategy based on independent machine learning grids was proposed, which allows RBFNN to achieve higher modeling accuracy than a deep-learning deconvolutional neural network (Deconv NN). Furthermore, the RBFNN features a simpler topological structure and fewer hyperparameters. The test results show that the mean absolute percentage errors of the total pressure, total temperature, and swirl ratio predicted by the RBFNN are 0.87%, 0.52%, and 1.09%, respectively, and the root-mean-square errors are 3730 Pa, 1.85 K, and 0.0077. Additionally, the parametric influences on RBFNN outputs were examined in detail, and compared with the results from CFD simulations. For small datasets, the proposed lightweight neural network modeling method outperforms deep learning models in both efficiency and accuracy. This approach can also be extended to other components within the secondary air system of gas turbines.

The extreme thermal loads induced by rotating detonation waves pose critical challenges to the thermal protection of rotating detonation combustors (RDCs). This study investigates the feasibility of gas transpiration cooling for RDC thermal protection, establishing a localized physical model of axial porous media to simulate detonation wave-coolant interactions. By comparing transpiration cooling characteristics under varying coolant mass flow rates and porosities, the study reveals distinct cooling mechanisms between rotating detonation wave and oblique shock wave regions. Numerical results indicate that transpiration cooling forms a 200-720 K porous wall thermal barrier, isolates the 2280-2800 K detonation core, and reduces wall heat flux by 72 % and pressure peaks by 20-83 %. Optimal cooling performance requires partitioned porosity: a sloping shock zone of 0.5 maximizes efficiency, while a detonation zone of 0.4 avoids increased heat flux due to flow instability. A 0.5 % coolant-to-mainstream mass flow ratio provides continuous thermal protection with a cooling efficiency of 0.72, while 1.0 % causes film breakage and temperature fluctuations. The porous medium acts as a chemical regulator to inhibit combustion and increase water vapor concentration through reverse osmosis to achieve inert gas buffering. The pressure dissipates within 12 mu s, but the thermal decay lags due to solid-phase inertia, which underscores the need for zonal porosity design and precise coolant control to achieve RDC stability.

Based on the aerodynamic and thermodynamic conditions of the integrated afterburner， the influence of main structural parameters of the insulation screen air film cooling on the cooling performance is systematically studied through numerical simulation methods. This research results indicate that under the same inlet and outlet boundary conditions， an increase in the opening of the upstream injector increases the secondary flow rate， separating the heat shield from the main flow of gas. The increase in the aperture of the gas film increases the depth and momentum of the jet， and reduces the uniformity and continuity of the gas film coverage. The increase in the inclination angle of the gas film hole reduces the wall adhesion of the gas film. The increase in the axial spacing of the gas film holes enhances the superposition effect of the gas film， and increases the coverage range and uniformity of the gas film. The optimal structural parameters are as follows： The aperture of the hole d=1.2 mm， the tilt angle of the hole θ=30°， the axial spacing S=10 mm， and the corrugated plate height H=6 mm. This research uses a forked arrangement for the holes in the screen.

To enhance the waste heat recovery potential of water vapor in hydrogen-fueled turbofan engines, this paper proposes two thermodynamic cycle models for water-enhanced turbofan engines: closed and open cycles, a comparative analysis of their power and economic performance under cruise conditions (altitude: 11 km, Mach number: 0.8) is conducted. Based on a dual-spool hydrogen-fueled turbofan engine, this study incorporates an evaporator-condenser water circulation structure after the low-pressure turbine to achieve water vapor recovery and the re-injection of liquid water into the combustion chamber. Throughout the computational process of this study, the total temperature at the combustor outlet is held constant at 1600 K, and the total air mass flow rate is maintained at 1 kg/s. The research findings indicate that the bypass ratio of the open cycle is significantly lower than that of the closed cycle, and its condenser thermal load is reduced, thereby benefiting the design of both the condenser and the overall engine structure. The open-cycle outperforms the closed-cycle in terms of thrust and overall efficiency but falls short in fuel consumption. As the water-to-gas ratio increases, the performance parameters of the two cycle modes become closer. At a water-to-gas ratio of 0.05, the relative changes are 22.4%, 22.2%, and 155.8%, respectively. At a water-to-gas ratio of 0.4, the relative changes are 4.3%, 4%, and 32.8%, respectively. In the water-to-gas ratio range of 0.15–0.25, the open-cycle demonstrates significant advantages in both power and efficiency while maintaining fuel consumption at a reasonable level, making it the preferred range for balancing performance and operational economy. This study provides a feasible pathway and theoretical basis for the structural design of water-enhanced hydrogen-fueled turbofan engines.

A test investigation is performed to the convective heat transfer in a confined crossflow channel with a spe-cific dimensionless height of 3,produced from a 2×2 square-array impinging jets with dimensionless pitches of 4.Par-ticular focus is played on the heat transfer enhancement by using the passively extended jet pipes and the actively center-positioned synthetic jet in the continuous-jet square array,as well as their combination schemes.Square-array jet Reynolds number(Re)ranges from 3 000 to 10 000,and the synthetic jet acoustic actuator is driven at a fixed fre-quency of 250 Hz.Correspondingly,the synthetic jet velocity ratios(defined as the ratio of synthetic jet characteristic velocity to square-array jet ejecting velocity)are varied from 2.0 to 0.6.From the jet pipe extension,the dimension-less normal distance between jet outlet and targeting wall is adjusted in a range of 1-3.Within the scope of this study,the heat transfer enhancement roles are clearly illustrated.Under Re=3 000,the synthetic jet integration demonstrates a significantly stronger heat transfer augment role than the jet pipe extension,taking on dominant heat transfer en-hancement mechanism in the combination scheme.With respect to the baseline situation(no synthetic jet integration and no jet pipe extension),the area-averaged Nusselt number on a specified zone could be increased up to 200%when the crossflow velocity ratio(defined as the ratio of crossflow inlet velocity to square-array jet ejecting velocity)be-yond 0.67.Whereas under Re=10 000,the jet pipe extension plays dominant heat transfer enhancement mechanism on the otherwise.the area-averaged Nusselt number could be increased up to 100%at crossflow velocity ratio of 0.5 in relation to the baseline situation.Meanwhile,in the square array with the extended jet pipes,the role of synthetic jet integration is very faint.The most possibilities wherein the combination of synthetic jet integration and jet pipe exten-sion could exhibit obviously its significance on heat transfer enhancement appear when both schemes display equiva-lent heat transfer augment roles.For instance,under Re=5 000 and crossflow velocity ratio ranging from 0.6 to 0.8,the combination scheme shows an obviously further improvement on heat transfer enhancement,in related to the single scheme either in active or passive.

In the aviation industry, the end leakage prevails in hyper-rotated rotating machinery, which brings a decline in its performance. However, in the Taylor-Couette flow (TC flow) system with micro-scale and hyper-rotate-speed, end leakage due to self-acting pressure difference has not been systematically elucidated. This study explored flow field driven by self-acting pressure difference in TC flow with micro-scale and hyper-rotate-speed using Large Eddy Simulation. The results showed that the flow field is mainly affected by three types of flow: the planar-bounded shearing flow from the end of rotator, the free cylindrical-bounded shearing flow from the lateral side of rotator as well as the end leakage from the inhomogeneous self-acting pressure difference. The planarbounded shearing flow causes axial flow toward the central and radial flow outward from the edge, which results in the formation and evolution of large-scale vortex structures. Besides, the free cylindrical-bounded shearing flow enhances the exhalation at the upstream region of minimum clearance while weakens the inhalation downstream. The self-acting differential pressure induces the mass and energy transfer between the clearance and the ambiance. The intensity of end leakage at the section of the clearance end is positively correlated with the rotational Reynolds number and eccentricity ratio, but negatively correlated with the dimensionless clearance height. An increase in the dimensionless clearance height weakens the axial velocity while enhances the overall mass transfer between the clearance and ambiance. The relationship between exhaled mass/kinetic energy flow rate driven by the self-acting pressure difference and operating parameters has been established. The sensitivity of the mass/kinetic energy flow rate to rotational Reynolds number is larger than other parameters. More specifically, as rotational Reynolds number increases by 3 times, the corresponding exhaled mass and kinetic energy flow rate increases by 2.59 times and 51.48 times, respectively. These results in this study contribute to design and optimization in the region of airborne rotating machinery such as dynamic pressure gas bearings and air sealing.

Numerical simulations and schlieren imaging experiments are conducted in the present study to investigate the effects of an additional transmission chamber (ATC) on pulsed jet impingement heat transfer in a confined crossflow channel. A confined crossflow channel with H/D = 4 is considered, where the pulsed jet is examined at four pulsation frequencies (f = 20 Hz, 50 Hz, 100 Hz and 200 Hz) and three jet-to-crossflow velocity ratios (VR = 2.5, 5 and 10) with the jet Reynolds number fixed at Rej = 10,000. The results show that, within the range of parameters considered in the present study, the ATC mode is beneficial for pulsed jet impingement heat transfer at higher pulsation frequencies and higher jet-to-crossflow velocity ratios. The inherent flow inertia inside the ATC improves the ejection velocity of the pulsed jet during the off-duty stage of the pulsation cycle and subsequently increases the peak ejection velocity, especially at higher pulsation frequencies. Compared with the conventional long-pipe nozzle(LPN) mode, the critical threshold of Strouhal number under VR = 10 is slightly reduced in the ATC mode. Under VR = 2.5, although jet ejection during the off-duty stage still persists, the off-duty jet is strongly deflected immediately downstream of the exit because of the stronger crossflow effect, thereby losing its normal momentum toward the target stagnation zone. Therefore, the ATC mode has little influence on local Nu distribution in the situation of VR = 2.5, even at higher pulsation frequencies.

Laminated structures, as an advanced cooling technology, are extensively utilized for thermal protection in the hot sections of gas turbines. This study employs a lightweight machine learning approach to swiftly and accurately predict temperature distribution in the laminated structures. To minimize modeling costs, a novel data preprocessing scheme encompassing data compression and one-to-one grid prediction was introduced. Utilizing this scheme, the classical radial basis function neural network (RBFNN) demonstrated superior predictive performance compared to the deconvolutional neural network (DCNN), while also boasting a simpler topological structure and requiring fewer hyperparameters. Testing showed that the mean absolute percentage error (MAPE) and the relative root mean square error (RRMSE) of RBFNN wall temperature predictions were 1.17% and 1.47%, respectively. The impact of changes in the input parameters of RBFNN on the output parameter has been thoroughly investigated. Results indicated that within the studied parameter range, RBFNN can sensitively capture the impact of changes in input parameters on the output parameters, and its predictions are consistent with the CFD calculation results. This study offers an effective method for modeling heat transfer in laminated structures.

Given the growing limitations on cooling air flow and the intensifying problem of gas ingestion, leveraging both sealing flow and surrounding secondary flows is considered to hold significant potential for improving sealing performance. A validated numerical method was employed to investigate a novel sealing structure, incorporating a single auxiliary sealing hole positioned at the high-radius location within the sealing cavity. Single-hole auxiliary sealing flow structures with varying swirl angles (B) and sealing flow distribution ratios (Mr) were analyzed to reveal their influence on the rim seal. Results indicate that the auxiliary sealing flow minimally affects the inner cavity flow field, while primarily influencing the flow behavior within the sealing cavity. Specifically, it obstructs the circumferential flow of the mainstream and influences swirl ratio distribution in the rim clearance. Compared to auxiliary sealing flows at B = 0 degrees and B = 45 degrees, the auxiliary sealing flow at B = 70 degrees consistently improves sealing effectiveness across various Mrs. An optimal flow ratio for peak performance was observed at Mr = 1:2, achieving a 54.2 % improvement in sealing efficiency. Furthermore, auxiliary sealing flow at B = 70 degrees reduces high-temperature zones on the turbine disk, leading to a more uniform wall temperature distribution.

The present study focused on the active heat transfer enhancement roles by integrating a center-positioned acoustic actuator into a specific continuous-jet square array in the crossflow. Experimental tests are conducted on the heat transfer behaviors, wherein four continuous-jet Reynolds numbers (ReCJ=3000, 5000, 7000 and 10,000) are taken into consideration, while the acoustic actuator is kept at a fixed exciation frequency of f = 250 Hz, providing the corresponding synthetic jet velocity ratios as R SJ-CJ =2.0, 1.2, 0.86 and 0.6 respectively. Numerical simulations are also performed on the flow dynamics in such an actively modulated jet impingement configuration. In the crossflow situations, the acoustic actuator integration generally demonstrates two aspects in its active role on heat transfer enhancement. Firstly, the synthetic jet has a more vigorous penetration capacity to exhibit its stronger impingement onto the target. Secondly, the synthetic jet is of high-pulsating capacity to pulsate the nearby flow. Therefore, it could achieve the most possibilities for a more effective heat transfer enhancement in the strong crossflow, unless the synthetic-jet characteristic velocity is far less than the continuous-jet ejecting velocity. In the viewing of spatially-averaged Nu s-av on a specificd zone, an increased about 220 % is identified under Re CJ =3000 (or R SJ-CJ =2.0) when R CF-CJ >= 0.66. Under Re CJ =5000 (or R SJ-CJ =1. 2 ), an increased up to 100 % is achieved when R CF-CJ >= 0.6. Even under Re CJ =7000 (or R SJ-CJ =0.86), greater improvements of Nu s-av are still indicated by an increase of about 60 % (under R CF-CJ =0.43) to 80 % (under R CF- CJ =0.71). However, the increase of Nu s-av becomes faint under Re CJ =10,000 (or R SJ-CJ =0.6), limited within 10 %.

To enhance the survivability of armed helicopters in high-threat environments, integrated infrared (IR) suppressors are increasingly adopted to reduce thermal signatures. However, such integration significantly alters the exhaust flow field, which may in turn affect both the infrared and acoustic characteristics of the helicopter. This study investigates the aerodynamic, infrared, and acoustic impacts of an integrated IR suppressor through the comparative analysis of two helicopter configurations: a conventional design and a design equipped with an integrated IR suppressor. Full-scale models are used to analyze flow field and IR radiation characteristics, while scaled models are employed for aeroacoustic simulations. The results show that although the integrated IR suppressor increases flow resistance and reduces entrainment performance within the exhaust mixing duct, it significantly improves the thermal dissipation efficiency of the exhaust plume. The infrared radiation analysis reveals that the integrated suppressor effectively reduces radiation intensity in both the 3~5 μm and 8~14 μm bands, especially under cruise conditions where the exhaust is more efficiently cooled by ambient airflow. Equivalent radiation temperatures calculated along principal axes confirm lower IR signatures for the integrated configuration. Preliminary acoustic analyses suggest that the slit-type nozzle and integrated suppressor layout may also offer potential benefits in jet noise reduction. Overall, the integrated IR suppressor provides a clear advantage in lowering the infrared observability of armed helicopters, with acceptable aerodynamic and acoustic trade-offs. These findings offer valuable guidance for the future development of low-observable helicopter platforms.

Fluidic oscillators, which can generate a sweeping jet without any moving components, have been utilized to enhance film cooling effectiveness in several research studies. A traditional wall-attachment type oscillator is composed of an inlet nozzle, a mixing chamber, two feedback channels, and an exit nozzle. However, its relatively long structure poses a significant drawback as it cannot be easily applied to thin-wall parts in gas turbines. To address this issue, an improved design of the fluidic oscillator was developed by horizontally arranging its main body. A series of numerical simulations were performed to comprehensively investigate the flow and heat transfer characteristics within this improved structure, and pressure-sensitive paint (PSP) measurements for adiabatic cooling effectiveness were conducted to validate the accuracy of the numerical model. The results demonstrate that horizontally positioning the main body of the fluidic oscillator does not lead to a significant change in the sweeping frequency. However, compared with the traditional structure, it enhances cooling effectiveness by improving the wall attachment performance of the coolant, particularly at high blowing ratio. Specifically, the numerical results indicate that at the blowing ratio of 0.5 and 1.5, the area-averaged adiabatic cooling effectiveness of the improved structure increases by 39.2% and 620.4%, respectively, while the discharge coefficient decreases by only 1.2% and 2.7%. The interaction between multi-row coolant jets contributes to enhancing the cooling effectiveness of the fluidic oscillator at downstream locations, yet its impact on the jet sweeping frequency remains negligible. At the same exit area, the increase of exit width enhances cooling effectiveness. To facilitate the application of the improved fluidic oscillator on curved surfaces, a curved fluidic oscillator was further developed. Moreover, the effects of mainstream velocity, turbulence intensity, and the coolant-to-mainstream density ratio on the cooling performance of the curved fluidic oscillator were discussed in detail.

The primary geometric parameters for shaping non-axisymmetric exhaust nozzles were defined. Based on CFD numerical simulation, the influence mechanisms of geometric parameters on the internal flow characteristics of the exhaust nozzle were explored. The simulation results indicated that the flow losses mainly originated from the wake vortices of the struts, the separation vortices after the center tube cylinder, and the backflow on the inner side of the bend. Within the range of W2/H2 from 1.45 to 2.25, increasing W2/H2 can improve flow distribution in the exhaust nozzle while reducing the backflow area on the inner side of the bend and the wake vortices of the struts, indicating an overall improvement in the performance of the exhaust nozzle. Within the range of A2/A0 from 1.0 to 2.2, increasing A2/A0 can make the total pressure loss coefficient gradually increases, and the static pressure recovery coefficient decreases, indicating that A2/A0 should not be designed too large. Within the range of /1 from 55 degrees to 75 degrees, Increasing /1 results in an increase in the outer contour dimension of the upper wall and an enhancement in the flow capacity, the static pressure recovery coefficient, dynamic pressure coefficient, and uniformity index of the velocity profile initially increase and then decrease. In the design of exhaust nozzle within allowable spatial constraints, to achive optimal performance, it is recommended to prioritize a larger W2/H2, smaller A2/A0, and /1 within the value of approximately 60 degrees.

This study conducted the experimental investigation of aerodynamic heating of Microscale Rotational Shearing Flow with Axial Limited-Length (MRSFALL). The temperature rise of the stator is captured by the high response thermocouples. The eccentricity ratio and clearance height are guaranteed by means of instantaneous trajectory and torsion monitoring of the rotator. The result shows that the maximum temperature rise takes place upstream of the minimum clearance height along circumferential direction. The distribution of temperature rise presents asymmetric curve along axial direction, and peak value occurs near the dimensionless axial position of 0.18. The effect of aerodynamic heating becomes notable as the rotational speed is larger than 3 x 10(4) r/min. The effect of end leakage and the viscous dissipation have great impact on temperature rise of MRSFALL. More specially, the peak value of temperature rise at dimensionless clearance height of 0.008 0 is larger than the case at dimensionless clearance height of 0.004 4. Furthermore, when the eccentricity ratio is too large, the viscous dissipation is induced, and the additional temperature rise is achieved. The heat flux identification of shear flow has been realized by Sequential Function Specification Method (SFSM) and its estimation of thermal load has been given. The heat flux induced by the aerodynamic heating in this study varies from 950 W/m(2) to 1 330 W/m(2). (c) 2024 The Authors. Published by Elsevier Ltd on behalf of Chinese Society of Aeronautics and Astronautics. This is an open access article under the CC BY-NC-ND license (http://creativecommons.org/ licenses/by-nc-nd/4.0/).

Axial throughflow cooling is a practical thermal management solution for gas foil bearings (GFBs) with ultrahigh rotational speeds and small bearing clearances. The present study conducted a numerical investigation to assess a hybrid axial-throughflow cooling design (both inner cooling flow passing through the hollow shaft and outer cooling flow passing through the rotor-stator gap) for a specific radial bump-type gas foil bearing using the fluid-solid coupled modelling methodology. First, the individual effects of each cooling flow (i.e., the outer cooling mode only and the inner cooling mode only) on the GFB thermal behaviours are directly compared under a fixed rotational speed of omega = 1 x 105 rpm with a preset eccentricity ratio of epsilon = 0.9. The outer cooling mode exhibited superior cooling efficiency compared with the inner cooling mode with the same cooling air usage, albeit at the cost of a significantly greater flow pressure drop. Second, the conjugate roles of the hybrid cooling flows on the GFB thermal behaviours are related to the total cooling air mass flow rate of mtotal = 10 kg/h. Nine flow distribution relationships were comparatively studied between the two cooling flows under the same preset static bearing load of F = 31 N, wherein the percentage of the outer cooling flow (mouter/mtotal) varied from 10 % to 90 %. The heat removal pathway of the outer cooling flow was found to be dominant. The outer cooling flow exhibited a heat removal proportion close to 60 % when its distribution percentage was 30 %. The results of a comprehensive performance evaluation that considered the peak temperature reduction and cooling air pressure drop suggest a favourable distribution percentage of the outer cooling flow in the range of 30-40 %.

Sweeping Jet Fluidic Oscillator (SJFO) is a functional device to generate self-excited unsteady oscillatory jets, based on the exact nature of flow instabilities within the device. On account of its inherent advantages, such as unique operation with self-excitation and self-sustainability, strong unsteady flow actuation with high sweep frequency and wide sweep fan angle, simple geometry with no moving parts, etc., SJFO is recognized as one of the promising potential candidates as applied to flow and heat transfer control practices, particularly for the aerospace community. During the past twenty years, vast efforts have been devoted to this issue, advancing the technological development and innovation application of the SJFOs. The current review mainly concentrates on the emerging interests of SJFOs in heat transfer applications, including sweeping jet impingement heat transfer, sweeping jet film cooling and sweeping jet composite cooling. First, a comprehensive overview regarding the recent advancement of sweeping jet heat transfer is provided. Then, from this overview, some general roles of sweeping jet impingement heat transfer and film cooling are presented and the research gaps are briefly addressed. Finally, an outlook on the challenges and future development of sweeping jet heat transfer is put forward to motivate further investigation in four aspects, such as, actively adjustable strategies for the sweeping jet frequency and fan angle, multi-parameter correlation mechanism and optimization of sweeping jets, integration innovation by combining the other enhanced schemes into sweeping jets, and a wide variety of composite cooling configurations by the use of sweeping jets.

Serpentine exhaust systems, known for their infrared and radar stealth capabilities, are becoming standard in flying wing aircraft. However, their design is constrained by the fuselage layout, causing potential offsets between the engine and nozzle exit axes. Developing a universal, high-performance serpentine nozzle design that accommodates various vertical and spanwise offsets (ΔZ, ΔY) presents a significant challenge. A series of ‘Preferred Nozzles’ and ‘Modest Nozzles’ were designed and numerically evaluated to assess the impact of these offsets on flow characteristics. Results show that the ‘Modest Nozzle’ exhibits a complex wave system and significant local losses in the constant-area extension section when subjected to ΔZ > 0.10D0 (D0 is the nozzle inlet diameter) or ΔY > 1.0D0, leading to a rapid thrust coefficient decrease. Vertical offsets significantly affect the Preferred Nozzle’s aerodynamic performance. When ΔZ = −0.50D0, a large vertical offset in the first ‘S’ section creates a recirculation zone, causing significant losses and reducing the thrust coefficient to around 0.96. When ΔZ ≥ −0.25D0, gas flow and wall shear stress distributions transition smoothly. When ΔZ ≥ 0.10D0, as the spanwise offset increases, the thrust coefficient experiences only a 0.17% loss and remains above 0.97.