To mitigate the adverse effects of tip leakage flow and vortices on the internal flow field, numerical simulations are conducted using the delayed detached eddy simulation framework. For a high-subsonic (Ma=0.67) compressor cascade, this study examines tip leakage flow control by endwall distributed pulsed suction under various excitation parameters. Dynamic mode decomposition is employed to examine the mechanism whereby distributed pulsed suction affects the flow. The results indicate that distributed pulsed suction provides secondary control of the leakage vortices that redevelop downstream, further reducing the associated losses. At the same axial position, distributed pulsed suction holes positioned along the leakage vortex trajectory are more effective than those located along the blade camber line, particularly the upstream holes along the vortex trajectory. Distributed pulsed suction requires a certain suction flowrate threshold, when the suction flowrate of the second suction hole is ms=0.25 %, it leads to adverse effects. Based on dynamic mode decomposition (DMD) of the 1Y+excitation frequency control scheme, the intensity of the leakage vortex increases during the development stage in the 1Y+ mode, with the control effect mainly manifested during the vortex breakdown phase. In the 0.5Y+and 2Y+ modes, the intensity of the leakage vortex is effectively controlled throughout all stages.

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 investigates the regulation effect of a self-recirculating passive periodic excitation casing treatment on the stable operating range of a miniature transonic axial-flow fan rotor. Numerical simulations were conducted to analyze the performance impacts and internal flow mechanisms under different rotational speeds (70%, 80%, and 90% of the design speed). The results demonstrate that the self-recirculating casing treatment significantly enhances the rotor’s stall margin by modulating tip shockwave intensity, delaying leakage flow breakdown, and suppressing high-entropy regions. Specifically, the stall margin improvement reaches 7.11% at 70% design speed, while the enhancement diminishes progressively at higher speeds. This method effectively addresses the insufficient stall margin caused by the relatively large tip clearance ratio in miniature rotors, with only a minor efficiency penalty. The proposed strategy provides a feasible unsteady flow control solution for high-load design of miniature turbofan engines, offering critical insights into the interaction mechanisms between periodic excitation and tip leakage flow-shockwave dynamics.

Ensuring the high efficiency and stable operation of a compact diffuser in a wide range of inlet Mach numbers (Ma) is of vital importance. In this study, the effect of flow structure and weight of flow losses in compact Crossover tandem diffuser (CTD) under different inlet Mach numbers at design incidence is analyzed in detail. The results indicate that, compared to subsonic inflow conditions, transonic inflow conditions have a greater impact on the flow losses in the compact CTD. One of the key findings of this task was that the inlet Mach number influences the weight distribution of flow loss from the fore and aft blades within the compact CTD. Among these, secondary vortex losses and end wall losses play the primary and secondary roles, respectively, in the changes weight of flow loss in the compact CTD. Secondly, the secondary vortex losses within the fore blade of compact CTD are primarily determined by the passage vortexes and the concentrated shedding vortex. Meanwhile, passage vortexes also are the key contributors to the secondary vortex loss of aft blade. As the inlet Mach number of compact CTD increases, the reduction in the weight of end wall losses, coupled with the increase in the weight of secondary vortex losses, this can be attributed to the intensified pressure gradient driving more boundary-layer low-energy flow of multi-wall into the secondary vortexes.

In compressors operating under near-stall conditions, the rotor blade tip region is characterized by significant flow expansion and quasi-periodic tip vortex dynamics. This study utilizes large-eddy simulations (LES) to investigate the coupled dynamics of the tip leakage vortex (TLV) and suction-side separation vortex (SSV) in a simplified planar compressor cascade model with tip clearance. Key findings indicate that the self-excited instability of the TLV originates from the periodic shedding of filamentous tip clearance vortices (TCVs), which are formed due to the pressure gradient across the blade. These secondary TLV structures entrain the primary TLV without merging, thereby delaying vortex breakdown and mitigating flow blockage. Application of unsteady pulsed jets on the rotor casing, axially aligned with the blade leading edges, demonstrates that excitation at frequencies corresponding to the dominant TCV frequency (F+=10) induces resonance. This results in a 52.29% reduction in the area of tip high-entropy regions and a shift of the TLV trajectory towards the suction side. This dual mechanism mitigates flow features associated with both spike-initiated and modal-wave stall inception, leading to a significant 26.9% reduction in the total pressure loss coefficient (Cpt). These findings suggest a novel control strategy for enhancing compressor stability margins via targeted unsteady excitation.

This paper experimentally and numerically investigates the starting hysteresis characteristics of a Mach 4.0 mixed-compression variable-geometry TBCC intake under varying contraction ratios (CR) to reveal the mechanisms by which the internal flow structures evolve. A single-degree-of-freedom throat-adjustment mechanism was designed to enable continuous and controllable CR variations from 3.2 to 9.1. The designed mechanism provides an experimental basis for controlling starting paths and analyzing hysteresis behavior. Both experimental and numerical results demonstrate a strong path dependence within the CR dual-solution region. As the CR increases, shock wave/boundary-layer interactions intensify, promoting the development of separation bubble and enhancing the upstream propagation capability of pressure. Once this capability exceeds a critical threshold, the flow field abruptly transitions into an unstart state. Compared to fixed-geometry paths, the variable-geometry starting strategy can sustain the shock wave and boundary layer characteristics based on the pre-established starting flow field. This approach delays both pressure propagation instability and the arrival of the separation critical point, consequently forming a stable starting hysteresis loop and significantly expanding the startable CR range. Under the optimal contraction ratio (CRoptimal = 7.7), the total pressure recovery and kinetic energy efficiency reach 0.4874 and 0.9288, representing increases of 10.82% and 1.25% over the critical fixed-geometry starting condition (CR1 = 5.4), respectively. These findings indicate that a well-designed variable-geometry starting path can effectively utilize hysteresis to release the compression potential of high-CR configurations, ultimately providing a new control strategy to expand the efficient operating range of TBCC intakes.

The effects of steady continuous blowing (SCB) and pulsed blowing (PB) through single endwall hole with low excitation momentum coefficient on the control of flow separation in a highly loaded compressor cascade are investigated experimentally using oil flow visualization and both steady and unsteady pressure measurements. First, excitation location and effects of relatively low values of the excitation momentum coefficient are examined to explore the capacity of SCB for damping flow separation and determine optimal values of the excitation parameters. With the aim of further improving aerodynamic performance, SCB with these optimal parameters is then compared with PB using the same parameters. The experimental results show that significant improvements in aerodynamic performance are achieved with SCB located near the separation region. However, the control effect of SCB becomes less effective as the excitation location moves upstream, and there is a deterioration in the flow field when the excitation is located close to the separation point. The effectiveness of SCB at relatively low values of the excitation momentum coefficient is confirmed. Even at a value of this coefficient as low as 0.1%, there is a 7% reduction in the total pressure loss coefficient. With the optimal parameters for SCB, the total pressure loss coefficient is reduced by 8.83%. PB provides more effective flow control than SCB at the same excitation momentum coefficient of 0.15% when the excitation frequency exceeds 40 Hz. With PB, the maximum loss reduction reaches 1.2% when compared with SCB and 9.85% when compared with the baseline case.

Vertical takeoff and landing (VTOL) aircraft must balance the conflicting demands of hover and cruise performance. To address the lack of integrated design methodologies in the existing literature, a unified design-optimization framework is presented, coupling high-fidelity CFD simulations with a genetic algorithm to refine a gas-driven thrust fan (GDTF) VTOL nacelle. Key geometric parameters-fan pressure ratio pressure ratio, fan tilt, nozzle angle, tail inclination, and tip shape-were varied in a comprehensive parametric study to maximize lift-to-drag ratio and maintain constant mass flow. The optimization reveals that a nearly horizontal fan axis maximizes cruise efficiency (LD approximate to 2.98), a nozzle angle of about 22 degrees offers the best lift-vs-drag compromise during transition, and refining the tip geometry yields a 10-20% performance boost. To validate the numerical predictions, a 1:1.05 scale VTOL nacelle model (fan diameter D = 0.42 m) was fabricated and tested in a low-speed wind tunnel at 52 ms (Re approximate to 5 & times; 106, turbulence intensity approximate to 2%). Total-pressure probes at the intake exit plane and static taps along the inner cowl wall provided detailed pressure distributions, from which exit Mach number, velocity and the equivalent flow coefficient phi (approximate to 0.68 under test conditions) were derived. Oil-flow visualization on the external cowl surface confirmed smooth, attached streamlines with no large separation bubbles. This dual validation combining surface-flow visualization and pressure-recovery mapping demonstrates the accuracy and reliability of the proposed simulation methodology. By successfully bridging detailed CFD with genetic-algorithm-driven design and validating against comprehensive wind-tunnel measurements, this integrated approach paves the way for next-generation VTOL configurations with longer range and lower fuel consumption.

Abstract This paper presents an aerodynamic design and evaluation of a vertical takeoff and landing (VTOL) aircraft featuring fuselage-integrated, gas-driven ducted lift fans and variable-geometry guide vanes. These components are embedded within the airframe to enable coordinated control of inlet and exhaust flows, facilitating a smooth transition between vertical and horizontal flight. A three-dimensional conceptual configuration is established, defining the key geometric and dimensional parameters. To assess both hover and high-subsonic cruise performance, a unified computational fluid dynamics (CFD) framework is developed using boundary-equivalent intake/exhaust conditions with momentum-flux corrections for forces. Simulation results demonstrate sufficient thrust margin for stable hovering and aerodynamic efficiency during cruise, with drag characteristics comparable to low-drag subsonic aircraft of similar scale. These findings contribute to the aerodynamic integration strategies of emerging VTOL and short takeoff/landing (STOL) aircraft.

Micro-turbine engines face significant stability challenges due to increased relative tip clearances and susceptibility to tip-flow instabilities. This study proposes and numerically investigates a novel self-circulating periodic excitation casing treatment (SPEC) designed to enhance the stall margin of a micro axial/diagonal compound compressor rotor. Unlike the previously proposed passive periodic excitation casing (PPEC), which relies on external atmospheric pressure, the SPEC establishes an internal recirculation path from a downstream high-pressure region to discrete upstream jet orifices. The results show that the SPEC reorganizes the near-casing unsteady flow, weakens the interaction between the tip-leakage vortex and the splitter-leading-edge region, and delays passage blockage. For the M1S2 configuration, the computed relative margin increases by 12.0%, while the near-stall efficiency increases by 9.3 percentage points with an approximately 1.3% reduction in pressure ratio. Efficiency is evaluated over the external compressor control volume using the inlet and outlet total quantities, with the SPEC flow treated as internal recirculation. Near-surge measurements at 70% design speed on the present compressor/SPEC hardware provide qualitative, trend-level corroboration of the computed near-surge operating-range shift, while the absolute stall boundary remains evaluated within the numerical framework.

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 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.

This paper investigates the stall mechanism of the first-stage fan of a variable cycle engine (VCE) at a small bypass ratio and the corresponding method for radial variable pre-rotation adjustment. Steady and unsteady three-dimensional numerical simulations of the fan are performed. The results indicate that, under conditions of a small bypass ratio, the pronounced throttling effect of the outer bypass will amplify the interference caused by shock waves, tip leakage, and the boundary layer in the rotor's tip region. This phenomenon is likely to result in substantial flow separation and fan stall. The radial variation trend of the inlet relative airflow angle of the VCE fan in the process of stall with a small bypass ratio is analyzed, and it is found that it is necessary to apply inlet variable pre-rotation to the blade height of the rotor above 60%. A variable inlet guide vane, which can apply variable pre-rotation to the local blade height of the VCE fan, is proposed for the first time. With more than 60% of the guide vanes and a local deflection of 8 degrees, the stability margin of the VCE fan can be increased by 4.33%, and the peak efficiency can be improved by 1.86%.

Endwall-pulsed blowing (EPB) is studied for three different excitation waveforms to improve the aerodynamic performance of highly loaded compressors. Some important excitation parameters include the excitation frequency and momentum coefficient, which were analyzed in detail. The results of the EPB are compared with the endwall steady blowing (ESB) case. For EPBs with the three excitation waveforms (Waveforms sine, triangle and trapezoid), excitation frequencies that are equal to an integral multiple of the natural frequency of the vortex shedding are optimal and provide better performances than the ESB with the same time-mean momentum coefficient. Moreover, the EPBs of the three excitation waveforms have significant differences in their aerodynamic performance improvements. The optimal case is achieved by the EPB with Waveform triangle and provides a total pressure loss coefficient with a reduction of 25.64%.

The current paper proposes a new third-order WENO (weighted essentially non-oscillatory) scheme, denoted as WENO-D3, which constructs the numerical flux using the convex combination of a second-degree polynomial of a three-point stencil with two linear polynomials of two substencils. WENO-D3 comes with a compact formulation of the numerical flux, new compact nonlinear weights, a new reference smoothness indicator, and a simple smoothness indicator for the three-point stencil. The forward differences approach is employed to reformulate the expressions of the polynomials and smoothness indicators of the new scheme. The smoothness indicator of the three-point stencil is designed using a linear combination of the two-point substencil smoothness indicators and compact linear weights. The linear weights of WENO-D3 can be freely selected with one condition: their sum equals one. A detailed analysis of the WENO-D3 scheme is provided, and numerous one- and two-dimensional benchmark numerical experiments are studied. Thirty-one combinations of linear weights are studied to verify the sensitivity of WENO-D3 to linear weights selection. Compared to the WENO-MZQ3 scheme, the WENO-D3 scheme significantly reduced the computational complexity and cost while providing flexibility in linear weights selection. The results show that the new scheme reduces up to 91% of the WENO-MZQ3 scheme's computational time and provides stable results for a wide range of linear weights. The results also show that the proposed scheme has recovered the optimal order at critical points and can capture and resolve sharp discontinuities without spurious oscillations.

The current work evaluates the effectiveness of the pressure-corrected osculating axisymmetric flows method in the design of hypersonic wavecatcher intakes with shape transition. The original osculating axisymmetric flows method, which is essentially used in waverider design, has drawbacks in accurately demonstrating the three-dimensional flowfield due to ignoring the cross-flow effects between the osculating planes. This negatively impacts the intake performance due to the presence of lateral pressure gradients. The pressure-corrected method takes into account these effects by incorporating lateral pressure gradient corrections into the original design methodology using cross-flow velocity information. This results in the generation of three-dimensional streamlines rather than the two-dimensional streamlines of the original method. The semirectangular-to-ellipse wavecatcher intake is selected as the subject of investigation, and the design procedure is reviewed. The characteristics of the wavecatcher intakes with a design point of Mach 6.0 are studied. Computational fluid dynamics analysis of pressure-corrected wavecatcher intakes is presented to assess the design technique. It is found that the initial conical shock impinges at the intake's entrance and the streamtube is completely captured. Furthermore, the comparison with the original wavecatcher intake indicates that the pressure-corrected wavecatcher intake demonstrated better performance in terms of total pressure recovery and flow uniformity. The wavecatcher intake with entrance-to-exit shapes transition showed higher performance than those with the same entrance and exit shapes.

This paper examines the impact of scramjet isolator shape transition on hypersonic internal waverider (IWR) intake. The IWR intake is designed using the osculating axisymmetric flows and streamline-tracing methods. The new Internal Conical Flow “M” basic flowfield is utilized to provide the flow information for the design method. The intake is equipped with three isolators: one with a constant cross section and two with variable cross sections with circular and rectangular exits. The entrance shape and area of the three isolators are fixed to the intake throat shape and area. The exit area of the three isolators is maintained as the entrance one. Numerical computations of three-dimensional configurations reveal that the isolators with variable cross section shapes demonstrate a higher uniformity index than those with constant cross section shape. Thus, the isolator shape transition has decreased the flow distortion of the hypersonic IWR intake system. The three isolators exhibit varied wall pressure distribution depending on the isolator cross section shape, and the total pressure recovery ratios at the three isolators' exit planes are similar. The wall pressure distributions and key performance parameters at the intake throat section, including total pressure recovery, compression ratio, and Mach number, remained consistent across the first part of the intakes. Therefore, changing the cross section shape of the isolator while keeping the area constant could enhance the flow uniformity of compressed air without negatively impacting the intake system's performance. This allows a separate shape selection of the IWR intake throat and the scramjet combustor entrance to fulfill their special requirements.

The tip-leakage vortex, as the dominant coherent structure of the blade tip of a compressor, is an important source of losses in centrifugal compressor rotors; it limits the stall margin and can even cause compressor surge. The miniaturization of mechanical components has attracted significant attention; however, the miniaturization of compressors-particularly centrifugal compressors-further increases the adverse effects of the tip-leakage vortex. This study sought to establish a new passive unsteady flow-control method for a micro centrifugal compressor using equal-circumferential-spacing through-holes on its casing. This approach fully exploits the unsteady characteristics of the flow field. A numerical study of the coherent characteristics of the blade tip was carried out using dynamic mode decomposition. The numerical conclusions were then applied to a physical system. The results of experimental tests of this system indicate that the improved flow-control method can increase the maximum efficiency and stall margin of the compressor by 2.5% and 9.0%, respectively, and it can also increase the maximum pressure ratio.

Three-dimensional numerical simulations of a first-stage fan in a variable-cycle engine (VCE) were performed to reveal the characteristics of the fan under different working conditions and the factors influencing them. The results indicated that several combinations of outer and inner bypass throttling could be used to achieve the same bypass ratio. To the best of our knowledge, this study is the first to establish the characteristic curves, which are dependent on the bypass ratio and throttling of the VCE fan. Consequently, the most suitable method of reducing the bypass ratio while ensuring a sufficient stall margin for the fan was to increase the outer bypass throttling and reduce the inner bypass throttling. Moreover, the only difference between the aerodynamic characteristic curves of the VCE fan for various bypass ratios at the same speed was the stall margin. An analysis of the flow field near the stall point revealed that the fan exhibited flow separation over an area with a large radius. Consequently, based on this flow separation, the reasons for the fan having different stall margins at different bypass ratios were explained. Furthermore, the axial position of the splitter was the primary factor affecting the throttling strength, whereas the radial position of the splitter determined the range of influence of throttling.