Surface roughness exerts a significant influence on the aerodynamic performance of compressor blade airfoils. To investigate the influence of leading-edge geometry on the sensitivity of airfoil aerodynamic performance to roughness, this study conducted planar cascade experiments with variable surface roughness to comprehensively measure the total pressure loss characteristics and blade surface pressure distributions of controlled diffusion airfoils with circular and spikeless leading edges. The laws of aerodynamic performance degradation under variations in roughness size and location within the sensitive regions of the blade were obtained. Results indicate that increasing roughness in geometrically sensitive regions induces a linear reduction in the available incidence range. In contrast, the minimum total pressure loss rises sharply once the roughness exceeds a certain value. The spikeless leading-edge airfoil exhibited significantly lower performance degradation in both minimum loss and available incidence range, indicating superior aerodynamic robustness against surface roughness. Increased surface roughness causes laminar separation bubbles (LSBs) to migrate rapidly upstream and collapse prematurely at smaller inlet flow angles. The LSBs on the suction surface of the circular leading-edge airfoil are consistently more susceptible to the disturbance of surface roughness, and turbulent separation is more likely to occur near the trailing edge.

Variable stator vanes (VSVs) represent a mature engineering approach to resolve matching issues in multi-stage axial compressors. However, constrained by the mechanical rotation structure, complex flow phenomena such as gap leakage significantly degrade the efficiency and stability of the compressor. This paper performs high-fidelity Computational Fluid Dynamics (CFD) simulations on the actual geometry of VSV in a specific fan. The results indicate that the drastic variation of the front radial gap during the adjustment process is the primary cause for the performance deterioration of the VSV. To prevent mechanical interference between the blade and the endwall at various adjustment angles, a substantial radial gap must be reserved at certain adjustment angles. The leakage flow intensely mixes with the main flow and bypasses the penny, exacerbating low-energy fluid accumulation in the corner region. To address this, this paper innovatively proposes a non-axisymmetric endwall contouring method to enhance performance across multiple operating conditions. This design effectively mitigates front radial gap variations during rotation and spatially isolates the passage vortex from the leakage vortex, significantly reducing corner loss. At the design incidence and adjustment position, this contouring decreases the total pressure loss coefficient by 11.6% and endwall loss by 28.7%. Ultimately, this approach provides an effective engineering strategy to elevate VSV aerodynamic performance.

Next-generation variable-cycle engines impose stringent requirements on the operating range of compressor systems in terms of mass flow rate and pressure ratio. Variable stators provide an effective means to meet these requirements; however, rapidly and accurately determining the operating domain of multistage compressors with variable stators remains a major challenge. To address this issue, this paper proposes an efficient framework for extracting the operating domain of such compressors. First, an improved one-dimensional performance calculation strategy is developed, which preserves design-point accuracy while significantly enhancing prediction accuracy over a wide stator-angle adjustment range, reducing characteristic prediction errors by more than 50 %. Second, the determination of the operating domain is reformulated as the computation of an envelope surface enclosing performance scatter points across the full adjustment range, which simplifies data handling and facilitates the identification of operating-limit boundaries. Third, a rapid solution strategy for the operating domain is introduced, substantially reducing the computational cost. For a four-stage variable-stator compressor, the required computational effort is reduced to 6.8 % of the original cost, with the reduction becoming more pronounced as the number of adjustable stages increases. Throughout the entire operating domain, the efficiency prediction error remains below 10-6. The proposed method enables fast and accurate determination of the operating domain of multistage compressors with variable stators and provides an effective tool for operatingrange analysis in variable-cycle engine applications.

Inter-stage bleed in compressors serves critical functions in aircraft environmental control, engine inlet antiicing, and cooling of engine hot-end components, thereby ensuring the safety and efficient operation of both aircraft and engines. Consequently, minimizing inter-stage bleed losses and enhancing bleed efficiency have become central challenges in the refined design of high-performance compressors. Building on the typical flow characteristics of the bleed slot, this study classifies the internal flow into three sections: the turning section, the diffusion section, and the sudden expansion mixing section. The flow and loss mechanisms of the primary components of the bleed system are elucidated, covering four aspects: flow separation in the turning section, local separation induced by incidence angle, excessive diffusion induced flow separation, and losses arising from sudden expansion mixing. Targeted flow control strategies are proposed for each of these loss mechanisms to improve the internal flow within the bleed system and thereby mitigate losses. Based on the loss mechanisms and corresponding optimization strategies, a flow-mechanism-based design method for the bleed slot is proposed, which enables parametric profiling of the slot's turning and diffusion sections, respectively. Finally, the methodology was applied to optimize four bleed slots with different overall design parameters, and the optimized configurations were experimentally compared with conventional straight-line and current high-performance quadratic polynomial profiles. The results show that, relative to the straight-line baseline, the optimized bleed slot reduces the loss coefficient by more than 45% and increases the static pressure rise coefficient by more than 0.2, while also outperforming the current advanced quadratic-polynomial profile, thereby confirming the effectiveness of the proposed design method.

Variable stator vanes (VSVs) serve as a critical aerodynamic control mechanism to optimize stage matching in multi-stage axial compressors. However, due to mechanical rotation constraints, complex flow structures such as gap leakage significantly reduce the compressor efficiency and stability margin. In this study, three-dimensional Computational Fluid Dynamics (CFD) simulations were carried out on the actual geometry configuration of VSVs in a specific fan to systematically evaluate the underlying flow physics and loss generation mechanisms. The numerical results indicate that the significant variations in the front radial gap during VSV adjustment are the primary cause of the performance deterioration. To prevent mechanical interference between the blade and the endwall, a substantial radial gap margin is required at certain adjustment angles. Consequently, the large-scale leakage flow strongly mixes with the mainstream and bypasses the penny, further intensifying low-energy fluid accumulation in the corner region. To mitigate these secondary flow losses, a non-axisymmetric endwall contouring control strategy is proposed in this paper. This design effectively accommodates variations in the front radial gap during rotation and spatially isolates the passage vortex from the gap leakage vortex, thereby significantly reducing corner losses. At the design incidence and adjustment position, this contouring scheme reduces the total pressure loss coefficient by 11.6% and the endwall loss by 28.7%. This approach provides an effective engineering strategy for improving the aerodynamic performance of variable stator configurations.

To satisfy the increasingly rigorous requirements for thrust-to-weight ratio and specific fuel consumption in modern advanced aero-engines, the integration of the last-stage stator of the low-pressure compressor within the transition duct has emerged as a critical strategy. This configuration forms an aggressive compressor transition duct, which effectively reduces the axial length of the compression system. However, as aerodynamic loading increases, controlling the stator corner separation to optimize the total pressure loss and outlet flow uniformity has become a significant challenge. This study investigates a high-load aggressive compressor transition duct. Initially, the influence and mechanism of forward sweep configuration on the internal flow of an aggressive transition duct stator were investigated through experimental study and complementary numerical simulations. The results demonstrate that a substantial forward sweep at the stator root effectively suppresses corner stall, significantly reduces loss and enhances flow field uniformity at the transition duct outlet. Nevertheless, this geometry considerably increases the axial length of the stator, thereby compromising primary advantage in length reduction of the aggressive duct. To address this issue, this paper proposes a flow control strategy termed "the integrated curvature-customized contoured endwall and positive lean configuration". This approach optimizes the stator hub curvature to regulate the streamwise adverse pressure gradient at the root. Concurrently, positive lean configuration is utilized to facilitate the radial migration of low-energy fluid away from the hub, thereby suppressing corner separation at the transition duct stator root without increasing axial length. Experimental data reveal that this integrated method maintains the loss at the design condition while achieving an 11.33% reduction in the total pressure loss for the entire transition duct system at the near stall condition. Furthermore, outlet flow uniformity significantly improves under both investigated conditions, with the distortion coefficient decreasing by 32.8% at the design point and 43.3% at the near-stall condition.

Modern high-performance aero-engines are developed for high efficiency, a high thrust-to-weight ratio and low fuel consumption, which leads to a remarkable reduction in blade height at the outlet of the final stage of high-pressure compressors and a pronounced size effect. To explore the influence of the design-point flow coefficient on compressor aerodynamic performance, a single-stage low-speed compressor was used to simulate the final stage of a high-pressure compressor, and four additional configurations with distinct design-point flow coefficients were designed based on this baseline model. The influence mechanisms of the design-point flow coefficient on the internal flow of compressors with different design-point flow coefficients were analyzed under design and near-stall conditions. Results show that compressor efficiency is significantly enhanced with an increase in the design-point flow coefficient, with a 1.9% improvement as the flow coefficient rises from 0.41 to 0.57; the flow margin initially increases with the design-point flow coefficient and stabilizes when the coefficient exceeds 0.45. Efficiency variation is jointly governed by the mainstream and tip clearance flows, while margin variation is dominated primarily by the clearance flow. When accounting for relative clearance variation and stator losses under high-speed conditions, the peak design-point efficiency shifts forward, with a 1.47% improvement in the compressor’s maximum efficiency as the design-point flow coefficient rises from 0.41 to 0.57.

This study employs numerical simulations to investigate the impact of pitch relative positions (PP) and axial relative positions (AO) on tandem stator performance and corner stall. The results indicate that two distinct types of corner stall behavior are exhibited. With PP values of 0.50 and 0.60, the rear blade stalls first, while with values of 0.70 to 0.90, it is the front blade that stalls first. Additionally, there is a trend for corner stall to be delayed as PP value increases, which is attributed to PP changing the load distribution between front and rear blades. Corner stall is also delayed as AO increases. The mechanism is that the adverse pressure gradient of the suction surface boundary layer in the root regions decreases with AO increasing, which delays boundary layer separation. Investigations into tandem stators with various combinations of PP and AO show that the type of corner stall primarily correlates with changes of PP, and type of front blade corner stall first associates with wider operating range and lower losses. Considering both loss and stall margin, the optimal PP range for the tandem stator root is between 0.80 and 0.90, while the optimal AO range is between −0.05 and 0.

Tandem blades have been recognized for their potential to enhance the loading capacity of compressors. However, tandem stators currently do not exhibit advantages due to insufficient understanding of the complex end-wall flow mechanisms. To address this, an extensive study was conducted on tandem stators using experimental and numerical methods. The analysis focused on loss development, three-dimensional flow structures, and interaction mechanisms between front and rear blades. The results indicated the following: (1) The rear blade’s influence on the front blade has contrasting effects in mid-span and end-wall regions. The stagnation action of the rear blade increases front blade load, leading to greater corner separation losses. (2) Mid-span flow losses account for over 60% of total losses near stalls, primarily due to increased mixing losses from the migration of front blade corner separations towards the mid-span region in the rear blade channel. (3) At higher mass flow rates, corner separations occur in the rear blade, driven by significant circumferential pressure differences at the front section of the rear blade, causing end-wall fluid migration towards suction surfaces. (4) Corner stalls predominantly occur in the front blade, with associated losses exceeding those in conventional blades.

Modern compressors evolve toward higher-loading and lower-aspect-ratio designs, significantly amplifying the impact of secondary flows on flow deviation angles. However, existing secondary flow deviation models struggle to mechanistically resolve corner flow physics, which is the primary contributor to this deviation. This study develops a physics-based secondary flow deviation model for throughflow analysis that explicitly predicts both the magnitude and spanwise distribution by rigorously accounting for corner flow mechanisms. The new model decouples secondary flow deviation into two physically distinct components: (1) corner vortex-induced and (2) end wall boundary layer (EWBL) induced deviations. Each component is characterized through dedicated momentum-based submodels that respectively describe corner vortex evolution dynamics and EWBL migration processes. Comprehensive validation using numerical and experimental data from subsonic linear cascades demonstrates an over 70% reduction in root mean square error (RMSE) compared to the conventional approach. Moreover, the model successfully captures abrupt deviation increases during corner separation-to-stall transitions and reveals complex interactions between hub and shroud corner flows. This work establishes a fundamentally sound framework for secondary flow deviation prediction, substantially improving throughflow analysis accuracy in challenging corner flow regimes and providing reliable insights for preliminary compressor design optimization.

The next-generation variable cycle engine imposes extensive adjustable requirements on the transonic fan. This paper presents a numerical simulation of two single-stage fans with identical design-point flow and pressure ratios but differing loading coefficients. The wide-range adjustable capabilities are analyzed, with the mechanism of loading coefficient affecting the fan's adjustable range clarified. The findings reveal that the low-loaded fan demonstrates a superior pressure ratio adjustable range around the design-point flow condition, whereas the high-loaded fan exhibits enhanced pressure ratio adjustable range under small-flow conditions. Around the design-point flow condition, both fans achieve the same upper boundary. However, the high-loaded fan is more susceptible to stator blockage at low-pressure ratio points, resulting in a rapid rise in stator losses and a narrower lower boundary. Under small-flow conditions, the rotor of the high-loaded fan outperforms the low-loaded fan in terms of flow capacity, spanwise matching, and inlet conditions, thereby achieving a wider lower boundary. Furthermore, the high-loaded fan achieves a higher pressure ratio near the stall point and possesses a larger flow margin at the operating point, enabling a wider upper boundary.

The meanline method serves as a fundamental tool for axial compressor design, yet its accuracy is fundamentally constrained by empirical model uncertainties and stage-to-meanline performance discrepancies. To address these limitations, enhanced performance correlations between reference streamlines and complete compressor stages have been developed in this study, while empirical models have been improved through comprehensive database analysis. The refined models exhibit exceptional precision, with incidence angle errors within +/- 1 degrees, deviation angle errors below 2 degrees, and loss coefficient errors of +/- 0.02-0.03. A novel equivalent blockage factor (Kb) is introduced to physically couple reference streamline velocity triangles with stage performance characteristics, effectively accounting for three-dimensional flow effects. Validation results show maximum design point errors of 4.29% in pressure ratio and 1.45% in efficiency for stages 35-38. Multistage simulations exhibit greater error magnification compared to single-stage analyses, reflecting the inherent error propagation characteristics of the meanline methodology. These findings contribute to the advancement of compressor design practice through both improved predictive accuracy and fundamental elucidation of meanline-to-stage performance relationships.

The early design stage of modern compressors urgently requires high-accuracy, low-cost two-dimensional (2D) cascade loss prediction models. However, existing traditional loss models, predominantly based on early profile data, struggle to accurately predict the performance of modern Controlled Diffusion Airfoils (CDA). This study develops a comprehensive loss model system specifically for 2D cascades with modern CDA profiles. A numerical simulation database, encompassing the entire operating range of various subsonic, transonic, and supersonic profile designs, was first constructed to provide the data foundation for the model system development. Eight critical sub-models essential for the system were then identified based on an analysis of loss sources (including blade surface boundary layers and shock waves). A methodology combining physical mechanism analysis and data-driven techniques was applied to determine the final modeling scheme for each sub-model. Validation results demonstrate that within the parameter space covered by the database, the new model system achieves over 70% higher prediction accuracy compared to the traditional model system, with approximately 95% of prediction errors falling within ±0.02. It also accurately captures the variation trend of loss with incidence angle. The entire model system, consisting of a series of explicit formulas with clear physical meanings, can be easily integrated into compressor design processes and effectively support the design and analysis of airfoils during the preliminary stages of compressor development.

Fouling on compressor blades is a critical factor contributing to performance degradation in aero-engines and gas turbines, garnering significant attention. To investigate the fouling characteristics on compressor blades over time and operating conditions, accelerated fouling experiments were conducted on a single-stage low-speed axial compressor. The degradation processes of the static pressure rise coefficient and torque efficiency with fouling time were comparatively analyzed under both the design condition (phi = 0.56) and the low mass-flow rate condition (phi = 0.48). This analysis identified the performance degradation characteristics induced by fouling and quantified the blade surface fouling thickness distribution. Finally, flow fields obtained through numerical simulation were used to explore compressor fouling mechanisms. Results showed that compressor performance exhibits a significant nonlinear decay trend with increasing fouling time, with degradation rates being more pronounced during the initial fouling period. At the maximum fouling time (120 min), the attenuation amplitudes of the static pressure rise coefficient and torque efficiency of the compressor reached 9.13% and 6.92%, respectively, at phi = 0.60. Fouling was mainly deposited on the leading edge and blade tip of the rotor, as well as the leading edge and pressure surface of the stator. In addition, compressors operating at low mass-flow rates were more susceptible to environmental contaminants, which not only enhanced the stability of the fouling layer on the leading edge, but also expanded fouling coverage on the pressure surface.

The increasing performance demands of modern aero engines necessitate the integrated design of compressor transition ducts with upstream components to reduce the axial length of the engine. However, this design approach narrows the spacing between the stator and the strut, making traditional research on transition ducts only with struts unsuitable. The numerical results and experimental oil flow visualization results were utilized to reconstruct the three-dimensional flow structures in the stator passages under various operating conditions. Additionally, numerical methods were employed to analyze the mechanisms of the strut’s effect on the upstream stator in an aggressive transition duct. The results show that the strut potential field increases the load on the upstream stator, leading to severe blade surface separation and corner separation/stall, and redistributes the inflow angle of the upstream stators circumferentially, resulting in significant differences in the flow structures within the stator passages on both sides. The separation flows within the stator passages mainly manifest in five types: pressure surface separation vortex, suction surface concentrated shedding vortex, suction surface separation vortex, suction surface-corner stall separation vortex, and suction surface separation vortex pair. Under different operating conditions, the separation flows within the stator passages are always composed of a part of these five types or a transitional state between two of them.

The operation range of the adaptive cycle engine (ACE) compression system is constrained by both the compression components and the bypass ducts, resulting in intricate matching mechanisms. Conventional analysis methods struggle to adequately evaluate the feasible operating range or the coupled constraints between components. This study employs an integrated hybrid-dimensional approach, combining zero-dimensional bypass analysis with one-dimensional/quasi-two-dimensional component analysis, to systematically investigate the matching effects of a triple-bypass compression system. The influence of key matching parameters, including the compression component operating points, high-pressure (HP) and low-pressure (LP) shaft speeds, and the core-driven fan stage (CDFS) variable inlet guide vane (VIGV) angles, is investigated. Results indicate that compression component matching primarily influences adjacent downstream bypass ratios, while HP/LP shaft speeds and the CDFS VIGV angle predominantly regulate the first and second bypass ratios. The feasible operating envelope is determined by the superimposed effects of these control parameters. To maximize the total bypass ratio, optimal operation requires increasing the front fan stall margin, elevating LP shaft speed, reducing HP shaft speed, and implementing partial CDFS VIGV closure to enhance pre-swirl. These findings provide critical guidance for control logic refinement and design optimization in advanced variable-cycle compression systems.

Corner separation and stall are important phenomena in axial compressors, significantly impacting loading capacity, efficiency, and stall margin. However, most existing loss model systems, applied by two-dimensional (2D) throughflow analysis for compressor preliminary design, highly rely on empirical models to predict secondary loss. These models may produce misleading results under strong three-dimensional (3D) corner flow conditions. In this paper, a physically based semi-empirical model based on corner flow mechanism is proposed for throughflow analysis, which accounts for both the magnitude and spanwise distribution of secondary loss. Secondary loss is decomposed into three components by physically decoupling different flow structures. These components are modeled mainly through simulating the formation of the corner vortex and analyzing the influences of critical factors. The new model is validated employing numerical simulation and experimental data of subsonic linear cascades with various geometric and aerodynamic configurations. Results indicate that the proposed model reduces the prediction errors of secondary loss by over 60%, compared to conventional models. It additionally captures the sharp increase in loss during the transition from corner separation to stall and the variation of spanwise distribution form caused by secondary flow. The introduction of physical mechanism provides a more solid theoretical foundation for secondary loss prediction, beneficial to the accuracy of 2D throughflow calculation and giving designers better guidance at the preliminary stages of design.

The accumulation of particulate contaminants on compressor airfoils constitutes a critical operational challenge for gas turbine engines, inducing progressive aerodynamic performance degradation. To explore the susceptibility of different regions on the blade surface of axial flow compressor to micrometer-sized particles, accelerated fouling experiments were conducted on a four-stage highly-loaded axial compressor. Performance parameters, including static pressure rise and torque efficiency, were analyzed pre- and post-fouling. Non-contact blue light scanning was used to determine fouling distribution and thickness. Results revealed substantial performance deterioration post-fouling, manifesting as 6.8% reduction in stage static pressure rise and 8.9% decrease in torque efficiency at large flow coefficient condition (φ = 0.73). The performance degradation was mainly attributed to fouling on the leading-edge and pressure surface of the blade, where the maximum fouling thickness reached 0.28 mm. Additionally, the susceptibility of the fouling is influenced by complex three-dimensional flow structures—particularly enhanced accumulation in regions affected by tip leakage vortices and corner separation zones.

Conventional one-dimensional mean-line methods struggle to accurately predict the performance of low hub-to-tip ratio transonic fans due to spanwise flow gradients, while two-dimensional methods are computationally costly and require excessive geometric parameters, making them unsuitable for preliminary design phases. This study proposes a quasi-two-dimensional multi-streamline approach that reduces both computational cost and geometric complexity relative to full through-flow solutions, without sacrificing radial flow resolution. By incorporating multiple streamlines under simplified radial equilibrium, it accounts for non-uniform radial distributions during preliminary design. Key innovations include a streamlined resettling algorithm, a stream tube choking model and flow redistribution scheme, and a meridional curvature correction scheme. Validation using single-stage and two-stage transonic fans shows that the method significantly improves prediction accuracy. Compared to traditional streamline curvature through-flow methods, it achieves about 100-fold speedup with the same number of streamlines and closely matches full three-dimensional computational fluid dynamics results in radial aerodynamic parameter distributions. This novel framework reconciles accuracy and computational efficiency, effectively addressing radial flow non-uniformities in the preliminary design phase. It streamlines the fan/compressor design process and enhances performance prediction reliability, offering significant advantages for high-performance turbomachinery development.