This study presents a comprehensive numerical investigation of highly loaded tandem blades operating in high-subsonic flow regimes (Ma = 0.75), with emphasis on optimizing gap geometry and blade positioning for integrated intake systems. A two-dimensional Reynolds-Averaged Navier-Stokes (RANS) solver is employed to systematically examine the effects of gap area ratio (AR), contraction angle (CA), axial overlap (AO), and percent pitch (PP) on aerodynamic performance. The results reveal that a convergent gap nozzle with AR = 1.5-2 and CA < 10 degrees minimizes total pressure loss while enhancing boundary layer control on the rear blade suction side. Optimal positioning (AO = 0.08-0.14, PP = 0.6-0.7) maximizes induced circulation of the forward blade and suppresses separation, yielding a balanced load split (LS approximate to 0.5) and a high diffusion factor (DF = 0.64). A design strategy is formulated, recommending an aft-loaded forward blade, a robust rear blade with two-section camber distribution, and a properly tuned gap geometry to exploit circulation and boundary-layer control mechanisms. The proposed tandem blade configuration achieves a turning angle of 45.6 degrees with a wide incidence range (-4 degrees to 3 degrees), demonstrating both high efficiency and robust off-design performance. These findings provide new physical insights into tandem-blade aerodynamics and practical guidelines for designing compact intake guide vanes in next-generation propulsion systems.

This study introduces a novel gas-driven thrust fan (GDTF) propulsion concept for fixed-wing vertical takeoff and landing (VTOL) aircraft, aiming to combine high hovering efficiency with superior cruise performance. The design philosophy integrates a VTOL nacelle concept, encompassing a gas-driven fan along with intake and exhaust systems, into a conventional fixed-wing airframe. A key innovation is the introduction of an aerodynamic cowl, which passively modifies the effective capture area and mitigates spillage drag. Numerical simulations and low-speed wind tunnel experiments validated the proposed configuration, demonstrating that the GDTF-powered VTOL nacelle can achieve a thrust coefficient of 0.8, representing a significant improvement 66 % over the baseline design. The integrated aircraft attains a maximum lift-to-drag ratio of approximately 12.7 at the angle of attack of 4 degrees, and the zero-lift drag coefficient (similar to 0.0228) falls within the typical range for subsonic aircraft. The results confirm that the proposed GDTF integration strategy yields both high hovering efficiency and favorable cruise characteristics. Additionally, these findings provide the foundation for developing an electric VTOL (eVTOL) variant capable of achieving flight speeds of 120 m/s, offering strong potential advantages for future urban air mobility (UAM) applications.

This paper utilizes the Computational Fluid Dynamics (CFD) method to systematically investigate the influence of cold-side (Recold) and hot-side (Rehot) inlet Reynolds numbers on the heat transfer, flow performance, and thermodynamic irreversibility (entropy generation) of an annular plate heat exchanger integrated into a micro aero-engine. The research results indicate that as the cold-side Recold increases, the cold-side convective heat transfer coefficient (hcold) is significantly enhanced; however, the high-velocity fluid experiences a pressure-drop-induced expansion cooling effect in the latter half of the channel, which suppresses the gas temperature rise—this unique physical mechanism is explicitly confirmed through a comparative simulation with an incompressible fluid model. Simultaneously, increasing the hot-side Rehot significantly boosts the system's average heat transfer rate and the overall heat transfer coefficient (U), with an exceptional growth rate of 91.13% for the average heat transfer rate when Rehot increases from 2000 to 5000. Furthermore, the entropy generation analysis shows that the cold-side Reynolds number is more sensitive to the system's total entropy production, suggesting its optimization potential is superior to that of the hot side. This study provides crucial theoretical basis and optimization guidelines for the design of high-efficiency, low-resistance heat exchangers operating under the high-Mach number conditions typical of micro aero-engines.

This study investigates the direct impact of heat transfer on the thermodynamic performance of Micro Swing Rotor Engines (MSRE) through numerical analysis. To comprehensively address the influence of heat transfer, we employ a refined thermodynamic simulation model, incorporating a regressive correlation formula, and introduce a fluid-thermal weak coupling method to yield practical solutions. The numerical analysis reveals that heat transfer has profound effects on the performance of MSRE. Specifically, the temperature cycling curve experiences significant alterations, resulting in an increase in cycle-residual mass by 72.6% and a decrease in intake mass by 10.55% at a working frequency of 100 Hz. The pressure cycling curve is primarily affected during the compression and expansion processes, leading to a substantial rise in pressure during compression (reaching 1.055 MPa) while the contribution of combustion becomes less noticeable. Consequently, these changes increase engine power consumption during compression by 46.41% and reduce overall engine thermal efficiency by 30.23%. Additionally, an increase of the inner wall temperature by 100 K leads to a linear reduction in engine power by 0.1 kW and thermal efficiency by 0.5%. To mitigate these challenges, we propose practical heat management strategies, such as applying heat insulating coatings. The study underscores the critical roles of heat transfer in MSRE operation and provides insights for optimizing its thermodynamic performance, achieving a potential improvement of up to 54.68% in power output and 12.79% in efficiency.

The Front Variable Area Bypass Injector (FVABI) is a key to bypass ratio adjustment for a Variable Cycle Engine (VCE). In order to study the role of the FVABI with the Core Driven Fan Stage (CDFS) duct, firstly, the engine bypass with the CDFS duct model and the equivalent engine bypass without the CDFS duct model are designed using the concept of a jet boundary line. By comparing the difference between airflow driving forces in the two engine bypass models, the quantitative effects of the injection from the CDFS duct on the mass flow rate of the engine bypass airflow are obtained under different combinations of pressure difference and area ratios. Then, the CDFS duct injection characteristic map is obtained through the typical experiment of the FVABI. Based on this map, the performance model of the FVABI is developed. Finally, the turbofan engine model with the Variable Inlet Guide Vane (VIGV), the First Variable Cycle Engine model (VCE1) with the CDFS duct and without the VIGV, and the Second Variable Cycle Engine model (VCE2) with the CDFS duct and VIGV are built. The gain on the engine bypass ratio adjustment range caused by the injection from the CDFS duct is clarified by comparing the three engine models. It is concluded that the bypass ratio adjustment range of the variable cycle engine with the FVABI is about twice that of the traditional turbofan engine.

In order to meet the needs of the development of high-performance micro-engine, the micro-turbine is faced with the technical challenge of high-load and transonic flow. At present, the flow characteristics and loss mechanism of high-load micro-impeller are not clear. In this paper, the flow characteristics of high-load micro-turbine rotor are studied. The results show that the loss in rotor trailing edge is the main factor that affect the operation of the micro transonic turbine, and the key to reduce the loss is to reduce the wake diffusion range and the Mach number in front of the extended shock wave. By adjusting the local curvature of the front in front of the incident point of the back-extension wave in the throat of the suction surface, the convergent-divergent ratio is in the range of 1.04 ~ 1.08. Meanwhile, the Mach number of wave front can be reduced by 3.4% and the reflection of trailing edge inward extension can be effectively suppressed, and the total rotor pressure loss coefficient can be reduced by 14.1% at most. The thickness has an important effect on the wave strength and wake loss in the trailing edge. When the relative thickness of the trailing edge decreases by 50%, the maximum Mach number on the suction surface decreases by 7.8%, the turbine stage efficiency increases by 1.92%, and the relative thickness of trailing edge is 0.124~0.185, the studied micro-transonic turbine performs better.

In this study, firstly, for the axisymmetric RVABI, the change-rule of adverse pressure gradient caused by radial velocity during the transition of internal flow mode in variable geometry is summarized, and a Bypass Ratio (BR) iterative algorithm based on the empirical correlation of non-equilibrium pressure is proposed. The algorithm can estimate the nonlinear relationship between area ratio and BR, with an error range falling below 6.5%. Then, we discuss the favorable effect of uniform mixing on the thrust augmentation of mixed exhaust under variable BR conditions. From this point of view, the characteristics of vortices evolution in different shear strength jets are compared, to clarify the effect of variable cycle parameters on jet mixing. As the results suggest, when Λ is as low as 0.22, the K-H disturbance is of high-frequency wavelet property, and it is difficult to induce large-scale spanwise vortices. The macro migrations of fluid elements in spanwise vortices and the diffusion effect caused by edge tearing is weak, which is not conducive to the energy exchange between the two streams. However, the low Λ jet will also correspondingly weaken the viscous dissipation effect of vortices. It is concluded that the dissipation level is proportional to the 2.31 power of the Λ.

As a propulsion system for vertical take-off and landing (VTOL) aircraft, the gas-driven fan propulsion system has received some attention in recent years due to its simple mechanical structure and good performance. During the operation of the propulsion system, the core turbofan exhaust is directed to the tip turbine to drive the ducted fan to obtain thrust, and the louvered vector exhaust device is used to achieve lift/thrust switching. However, due to the linkage characteristics of the guide vanes of the exhaust device, the exhaust area will gradually change with the deflection of the guide vanes. This may cause uncertain effects on the system. Besides, considering the development cost of the propulsion system, it is necessary to clarify whether the core turbofan needs to be redesigned based on the characteristics of the system. Further, the transitional performance (VTOL to flight) of the propulsion system also needs to be studied. To this end, this study established an overall model of the gas-driven fan propulsion system combined with the characteristics of the core turbofan to analyze these problems. The results indicate that the louvered vector exhaust device can well match the gas-driven fan propulsion system. And when the propulsion system is composed, the gas-driven fan will not significantly affect the operation characteristics of the core turbofan. Besides, the preliminary analysis shows that the gas-driven fan propulsion system can meet the power requirements of the transition state of the fixed-wing VTOL aircraft. And the characteristics of the propulsion system may make the take-off thrust-weight ratio requirement of the fixed-wing VTOL aircraft be as low as 1.1.

This paper presents a novel design method of pressure-controllable bump for hypersonic aircraft forebody. The new developed method can effectively resolve the tradeoff among boundary layer diversion, flow uniformity, and external drag reduction. The classical permeable-boundary method is improved by coupling with the radius-based function, through which the prescribed surface pressure distribution can inversely generate the bump. The improved permeable-boundary method is available for 3D unstructured mesh which increases the solution efficiency and robustness. The accuracy of the method is evaluated through the comparison with experimental results. In addition, five principles to arrange the pressure distribution are proposed. Then, a new pressure-controllable bump is designed. Compared with the typical streamline-tracing bump, the new bump is 44.9% lower in height while diverting identical low kinetic energy flow. According to the space occupied with high kinetic energy flow, the uniform region to preset the inlet of the new bump is 58.0% wider than that of the typical bump. Moreover, the usage efficiency of high kinetic energy flow is 1.5% higher. The new bump is less convex and its equivalent lift-to-drag ratio is two times greater than the typical bump. This research confirms that the new pressure-controllable bump shows better integrating capacity with the hypersonic inlets than the typical bump. The current design method is based on an inviscid condition; the method considering the viscous effects will be further studied in the future.

The exhaust device of gas-driven fan propulsion system (for a VTOL aircraft) adopts vector exhaust guide vane (VEGV) to achieve vector thrust in the range of 0-90°, which requires that the deflection angle of the VEGV reach about ±45∘ and the total pressure recovery coefficient above 0.985. Therefore, the exhaust device needs a wide-range and low-loss VEGV to ensure efficient operation of the propulsion system. And the key of this VEGV is to eliminate the separation of the suction side under large deflection angle. So, this paper designed a new type of VEGV that can meet the demands by reducing the inverse pressure gradient of the suction surface (one of the necessary conditions for two-dimensional separation). In order to realize this design process, the spectral method of small disturbance equation (SMSDE) for two-dimensional subsonic flow was developed. Then, two VEGVs were designed by the SMSDE and verified by numerical simulation with Reynolds number 106 and blade solidity 1.18. The results show that the two VEGVs eliminate the separation of suction side at inlet Mach number of 0.25 and blade stagger angle of 45°. And the VEGV with lateral blade guarantees this characteristic up to Mach 0.3. Besides, when the inlet Mach number is below 0.3, the VEGV with lateral blade can ensure that the total pressure recovery coefficient is greater than 0.985, the outlet area ratio is greater than 0.95 and the exhaust angle error is less than 3.5°. Finally, Experimental verification of the vector exhaust device using the VEGV with lateral blade was carried out. The results show that the new VEGV has a higher total pressure recovery coefficient and outlet area ratio than the traditional VIGV at large deflection angles.

This paper presents a new concept of vertical take-off and landing (VTOL) propulsion system based on gas-driven fan. This system has a simple mechanical actuation mechanism and enables high cruising efficiency. However, compared with the traditional ducted fan propulsion system, the concept may have disadvantages, such as large total pressure loss due to the intake and small installation thrust caused by internal and external flow coupling. An overall performance analysis model of a fixed-wing VTOL aircraft that combines the characteristics of propulsion system components is established to clarify the influences of such factors on the propulsion system. The comprehensive performance of the existing VTOL propulsion system and the new concept is compared. A sensitivity analysis of key aerodynamic parameters is performed for the new propulsion system. Results indicate that the new propulsion system can balance the contradiction between VTOL and flight and obtain a large range and an effective load when the speed requirement is low. Nevertheless, this system is sensitive to the aerodynamic parameters. The energy loss coefficient of the propulsion system should be less than 0.2 to ensure that the propulsion system can be at an efficient operating point.

The turbine based combined cycle propulsion system, a hopeful air-breathing propulsion system for high speed aircrafts, needs to work in a wide Mach number range. To meet the requirement, the variable-geometry internal waverider inlet based on improved Internal Conical of Flowfield C+ is proposed in this paper. Firstly, the axisymmetric basic flowfield is improved to an Internal Conical of Flowfield C+ with a higher external compression ratio, which befit to relieving the mass flow choke under low Mach number conditions.

针对气驱涵道风扇垂直起降动力系统的核心部件叶尖涡轮开展了数值模拟,掌握了低稠度小偏转角叶尖涡轮(具有高反力度特征)的气动特性和损失构成.研究表明:针对低稠度涡轮物理喉道消失、气流易分离的特点,提出低稠度小偏转角涡轮叶型.此叶型可构建气动喉道,在一定程度上弥补物理喉道消失带来的影响,其流动较常规低稠度涡轮得到了极大的改善,不存在明显流动分离,其效率为84.38％,能量利用率为70.13％,其损失构成和常规涡轮有所不同,叶尖泄漏损失高达54.15％,二次流损失为34.41％,叶型损失为11.44％.

Waverider-inlet integration design is an important approach to realize hypersonic flight. The inlet lip is the key factor for internal/external flow field coupling. This paper proposes a novel 3D inverse method of characteristics (MOC) to design the inlet lip with generalized 3D shock. The unit process and marching procedure of the inverse MOC in 2D and 3D focus on obtaining the unique coordinates of the solution points. The accuracy of inverse MOC is verified through its comparison with the analytical solution of a conical flow field. The approach is then applied for a hypersonic bump-inlet integration (The freestream Mach number M-infinity = 6.0), where the inlet lip is inversely generated by a prescribed elliptic-conical shock wave. Inviscid results reveal good performance (The mass capturing ratio phi = 0.813, the Mach number of inlet exit M-exit = an, the total pressure recovery coefficient sigma = 0.752). The incident shock is well attached on the inlet lip, which is of high phi. Viscous results show relatively low performance (phi = 0.738, M-exit = 2.94, sigma = 0.471), which indicates that the new method is a promising solution for the hypersonic internal/external coupling flow. Although the viscous effects should be further considered to improve the design, the proposed method can be applied to 3D surface design with generalized shock shapes.

With low efficiency and insufficient power, the conventional micro swing engine has not been available for portable power source. Therefore a novel hybrid thermodynamic cycle with heat recuperation ability, which is suitable for the micro swing engine, is proposed to overcome the above defects, whose thermodynamic characteristics are investigated in details, indicating a great improvement for the engine performance. Firstly, the micro swing engine is modified to be a three-arm engine with a non-isometric multi-chamber structure. Driven by the hybrid cycle, the engine body and a specially designed spiral recuperator work together successfully in theory. Thus, the entire efficiency is increased by heat recovery from hot exhaust and restriction of structural heat leakage. In consideration of flow leakage effect, simulation models of the steady working process of the engine in hybrid cycle are established. Numerical results reveal that, thermal performance of the hybrid cycle has an obvious advantage over the single thermal cycle. Moreover, the hybrid cycle is composed of two single four-stroke cycles and two two-stroke cycles. The two adjacent cylinders, between whom the flow leakage happens, are with different cycles, and hence distinct thermal performances. Finally, with the application of this heat recovery process, the efficiency of the four-stroke cylinders is improved relatively by 45.31% and the total engine efficiency is increased by 14.09%. Effectiveness and practicability of the hybrid cycle with heat recuperation are demonstrated theoretically, which is of great significance for the progress in micro heat engine field.

A numerical investigation into conjugate heat transfer for a double-tube heat exchanger in mini-scale was conducted, with an asynchronous intermittent laminar flow imposed at the inlets. As a basic component of the micro recuperator for a micro swing engine, a mini double-tube exchanger under the same flow conditions was chosen as the physical model with consideration of variable fluid properties. Results indicate that, for the asynchronous intermittent pulsed flow, the axial wall heat conduction is much more serious and the periodic fluctuation of the solid wall temperature further worsens the heat transfer efficiency. Two dimensionless numbers R* and (C) over bar* are proposed to evaluate the two adverse effects. Then, two key parameters involved in the design of the micro recuperator for the micro swing engine are studied, namely, wall thickness ratio delta/D-c and flow frequency. For certain flow frequencies, as the value of delta/D-c increases, the axial heat conduction is enhanced while the wall temperature fluctuation is reduced. For a given delta/D-c, the flow frequency dually affects the heat exchanger efficiency and the optimum frequency exists. The optimal value increases with the wall thickness ratio. These results lay a good foundation of the micro recuperator design for the micro swing engine. (C) 2019 Elsevier Ltd. All rights reserved.

Internal waverider inlets (IWIs) are novel three-dimensional (3D) high-performance inward turning inlets. However, they possess poor self-starting capacity when applied in a fixed-geometry inlet for ramjets. Firstly, this paper presents an analysis of self-starting capacity for IWIs to demonstrate that IWIs face more difficult challenges when used as ramjet inlets than they do as scramjet inlets. Self-starting capacity must be taken into account when designing ramjet inlets. Secondly, the impact of a fluidic control method on a fixed-geometry IWI was studied by numerical simulation of fluid flow. The numerical results show that the fluidic control mechanism improved the self-starting capacity of the IWI at low Mach numbers: the minimum Mach number of self-starting capacity was reduced from M3.8 to M3.2; furthermore, the compression ratio was increased from 29.9 with no fluidic control to 31.9. By analyzing two different groups of fluidic control positions, it was determined that bleeding before the separation bubble has no impact on the IWI self-starting capacity.

An innovative type of pressure distribution for the hypersonic aircraft forebody (bump) is presented, and this design is based on the newly established known as pressure ridge (PR) flow mechanism. Studies on the low-kinetic-energy fluid over the aircraft surface are summarized. Then, the challenges of the inlet–airframe integration at high speeds are discussed. The flow structure around the bump is analyzed in detail to eliminate the side-embedded shock (SES) effect at the inlet entrance. The concept of PR is proposed to improve the overall aerodynamic characteristics of the bump, namely, the boundary layer removal, the reduction of external drag, and the streamline direction at bump end-section. On the basis of the developed inverse method to generate the bump surface by the prescribed pressure distribution, the improved PR-derived bump is designed and numerically compared with the typical pressure-controllable bump (PCB) while the identical leading edge profile is imposed. Results demonstrate that the PR creates outward and inward pressure gradients. The removed amount of boundary layer increases with the increase in outward pressure gradient. Meanwhile, the inward pressure gradient determines the streamline pattern after the bump. The PR-derived bump is 29.2% lower in height than the typical one by using a proper outward pressure gradient. The uniform area of the new bump is 30% wider than that of the typical PCB. The near-wall streamlines of the new bump are adjusted from expanding to the parallel, thereby relieving the side-compression. Changing the location, width, and peak value of the PR can lead to great flexibility in the design and optimization of aircraft forebody subjected to hypersonic flow.