Corner Separation Research Articles

Research on identification and quantification methods for loss sources of compressors

Abstract A deep understanding of the internal loss mechanisms and specific values of turbomachinery is crucial for its design and analysis. In this paper, based on the theory of entropy production, the loss sources in the NASA Rotor67 transonic fan and Radiver centrifugal compressor were quantified by using numerical simulation methods and physical flow characteristic description methods. The proportion of losses in various regions was presented. The results show that in the transonic fan, boundary layer losses and wake losses account for more than 70%, while direct shock losses account for less than 2%. Additionally, as the mass flow increases, the losses gradually move downstream, with the proportion of wake losses increasing. Within the wake losses, the influence of tip leakage flow and corner separation flow exceeds 70%. In the centrifugal compressor, as the mass flow decreases, the proportion of impeller losses to the total stage losses increases significantly, from 21.7% at the choking condition to 63.5% near the design condition. Most of the losses within the impeller occur near the casing area, accounting for about 47% of the impeller losses. The areas near the blades and the main flow area are the other two major sources of impeller losses, with both contributing to 36% of the losses.

Design and aerodynamic stability improvement study of a non-axisymmetric stator fan in a distorted insertion plate configuration

The design of anti-distortion fans must ensure the reliable operation and flight safety of propulsion systems under inlet distortion conditions. This research uses an insertion plate to obtain the inlet distortion flow field for a specific type of turbofan engine. Based on the characteristics of distortion effects, a novel strategy is proposed to rapidly implement a non-axisymmetric stator (NAS) quasi-three-dimensional design to create anti-distortion fan designs and improve the performance and stability of the fan in distortion environments. Numerical studies show that under three rotational speed operating conditions, the NAS effectively improves the total pressure ratio and efficiency characteristics of the fan under inlet distortion conditions. Compared to the prototype, the NAS can significantly increase the stability margins by 10.31 %, 8.31 %, and 7.39 % under the three different speed operating conditions, which effectively expands the stability boundaries. The NAS design directly affects the distortion region, effectively improves the incidence angle of stator vanes in the distortion region, suppresses the development and migration of vortices inside the stator passage, and significantly reduces the stator vane diffusion factor. Because it effectively eliminates corner separations and makes the incidence angle distribution of stator vanes relatively circumferentially uniform, the internal flow field of the fan's stator vanes and overall aerodynamic performance of the fan improve. Furthermore, the NAS design can improve the internal flow field downstream of the fan, direct the nonuniform flow field at the inner bypass outlet toward circumferential uniformity, and effectively improve flow field uniformity.

Investigation of mechanisms on leading edge vortex generation and suppression in vaned diffuser of a centrifugal compressor

This study numerically investigates the mechanisms for generating and suppressing the leading edge vortex (LEV) in a vaned diffuser of a centrifugal compressor, aiming to enhance the compressor's stable operating range. Detailed analyses focus on the mechanisms of the rotating stall and the initiation process of the LEV. Notably, the end wall contouring method is applied to vaned diffuser to suppress the LEV and improve diffuser stability. The suppression mechanisms of the LEV are examined to deepen the understanding of stability enhancement mechanisms. Results demonstrate that, under stall conditions, the vaned diffuser experiences two-cell rotating stalls moving opposite to the impeller rotation direction, each rotating at approximately 12% of impeller rotation speed. The formation of diffuser stall is attributed to a longitudinal vortex developing from the LEV and causing the passage blockage. The increasing spanwise distortion of leading edge (LE) incidence angles and circumferential distortion of static pressures at the diffuser inlet promote the evolution of the longitudinal vortex from LEV. The end wall contouring of vaned diffuser effectively enhances the compressor's stable operating range by 15.10% and 8.90% toward lower flow rate conditions for machine Mach numbers 1.3 and 1.2, respectively. At near-stall points, the end wall contoured diffuser reduces spanwise distortion of LE incidence angles and circumferential distortion of static pressures at the diffuser inlet, mitigates the LE flow separation and corner separation, suppresses the generation and development of the LEV, delays the onset of rotating stall in vaned diffuser, and thereby improves the stable operating range of centrifugal compressor stage.

Study on the influence of labyrinth seal structure on the performance of shrouded stator cascades with inter-stage cavities

Numerical simulations of a shrouded stator cascade with an inter-stage cavity were conducted using validated numerical methods to investigate the impact of different labyrinth seal structures and positions on the aerodynamic performance of the stator cascade. The study discussed the development trends of secondary flow motion and corner separation within the cascade under different labyrinth seal structures, evaluated the stator performance using total pressure loss coefficient and entropy-based loss coefficient, and revealed the interaction between cavity leakage flow and mainstream flow. The results indicate that individually changing the labyrinth seal structure or position has a negligible impact on stator performance, but simultaneously altering both the structure and position of the labyrinth seal can effectively reduce stator losses. The structure case 1-position case 2 labyrinth seal structure can reduce total pressure loss by 6.25%, while structure case 4-position case 4 reduces entropy increase loss coefficient by 7.67%. Additionally, it was found that changing the labyrinth seal structure and position can increase the circumferential velocity and momentum of cavity leakage flow at the cavity outlet, enhance the resistance of cavity leakage flow to circumferential secondary flow within the stator passage, and reduce stator losses. Altering the structure and position of the labyrinth seal effectively reduces the cavity leakage vortex at the leading edge of the cascade during the mixing process of cavity leakage flow with the mainstream, causing the passage vortex to move rearward axially.

Numerical investigation of aerodynamic characteristics in tandem cascades with various curving strategies

Tandem configurations with curved blades can improve the flow field in compressor cascades. This study conducts a numerical investigation into the impact of various curving strategies on the losses and flow structures within tandem cascades. Curving strategies of three distinct configurations (curving both blades, solely the rear blade, and solely the front blade) along with three curved angles (5°, 10°, and 15°) are compared. Four primary vortex structures are causing losses in the tandem cascades. Loss weight coefficients, derived from each vortex, are adopted to quantify the proportion of losses. All curving strategies mitigate the accumulation of low-energy fluid near the endwalls, weakening the passage vortex. Various curving strategies have different effects on losses of the endwall spanwise vortex and trailing edge shedding vortices. The influence of curved angles on the three configurations indicates that curving both blades at 5° and curving the front blade at 5° or 10° can mitigate losses. An increasing loss is observed when the rear blade curves at any angle alone. Curving both blades or only the front blade mitigates the corner separation while the curved rear blade exacerbates it. The novelty of this study lies in applying topology analysis to establish correlations between losses, vortices, and topological configurations. This approach is a pivotal research methodology for elucidating the underlying mechanisms of loss generation and vortex dynamics within curved tandem cascades.

Numerical investigation of tip bifurcation on aerodynamic characteristics and flow field structure in a compressor cascade

In this study, the tip bifurcation of the compressor cascade is proposed, referring to the biological characteristics of bird wing bifurcation. The influence of tip bifurcation with different structural parameters on the aerodynamic performance and flow field structure of compressor cascade at different incidence angles is studied by numerical simulation. The research shows that the tip bifurcation can make the starting position of the corner separation move backward, inhibit the separation along the pitch and blade height direction, and reduce the flow loss of the cascade. The separation control effect of the corner region first increases and then decreases as the bifurcation height increases, and gradually diminishes as the arrangement position approaches the trailing edge. When the tip bifurcation is set at the starting position of the corner separation and its height is equal to 4 %H, the flow improvement effect in the passage is the most obvious, and the cascade flow loss is reduced by 14.4 %. At the same time, the tip bifurcation cascade has effective positive incidence angle characteristics, but when the incidence angle is less than the minimum loss incidence angle, the tip bifurcation increases the cascade loss.

Research on Effect of Endwall Contouring of Vaned Diffuser on Stable Operating Range of Centrifugal Compressor

Abstract The study investigates the impact of diverse endwall contouring strategies on the operational stability of a high-pressure ratio centrifugal compressor stage. Employing numerical simulation techniques, this study examines the implications of contoured wall positions, specifically convex profiles at the pressure side and/or concave profiles at the suction side, as well as asymmetric endwall contouring, on both stage performance and internal flow field within the centrifugal compressor. And the stability enhancement mechanism of endwall contouring is clarified. The results demonstrate that employing full circular contoured vaned diffusers on the hub-side wall can achieve a maximum improvement in stall margin of 15.10%. Additionally, the diffuser featuring solely convex profiles at the pressure side on the hub-side wall exhibits an augmented stall margin by 8.39%, with a reduction in performance loss at low flow rate conditions. The asymmetric endwall contouring improves stall margin by 8.39%, and mitigates the performance loss across the entire operating range. At the near-stall point, the contoured endwalls redirect fluid towards the suction side, thereby mitigating the incidence at the vane leading edge and attenuating the interaction between the impeller trailing edge vortex and the diffuser leading edge vortex. Moreover, concave profiles serve to accelerate the fluid on the suction side along the flow direction, thereby suppressing the corner separation and weakening passage blockages. Simultaneously, the advantages of endwall contouring in vane diffusers extend to the improvement in the circumferential non-uniformity of flow fields and the enhancement in diffuser stability.

Investigation of vaned diffuser endwall contouring technology for improving the stable operating range of a centrifugal compressor with an asymmetric volute

AbstractIn this study, nonaxisymmetric endwall contouring technology is employed as a passive control strategy to enhance the stable operating range of a centrifugal compressor with an asymmetric volute. The endwall contouring method is applied to the hub‐side wall of the vaned diffuser within a centrifugal compressor stage. Instead of utilizing a flat hub as the baseline diffuser, various diffuser configurations featuring hub‐side contoured endwalls are explored, with adjustments in the peak radial position and peak height of the convex and concave profiles. Numerical simulations are conducted to evaluate the performance of the centrifugal compressor stage with both the baseline and contoured vaned diffusers. The investigation explores the underlying flow control mechanisms and establishes the effective endwall contouring guidelines. The findings highlight the effectiveness of endwall contouring in enhancing the stable operating range of centrifugal compressors. Notably, a substantial enhancement of 15.1% in the stable operating range is achieved by employing the contoured diffuser with the peak radial position near the vane leading edge (LE) and the peak height at 20% of the vane height. The endwall contoured diffusers reduce the positive LE incidence angles to direct the fluid toward the suction side (SS), and increase the radial velocity near the SS to suppress the hub‐suction corner separations in the diffuser passages near the volute tongue. These improvements collectively contribute to the enhancement of diffuser flow characteristics. Finally, a set of effective endwall contouring guidelines is proposed to guide the endwall contouring design for enhancing the stable operating range of centrifugal compressors.

Stability enhancement using a new combined casing treatment strategy in an ultra-highly loaded transonic compressor rotor

Low-reaction compressors are promising for achieving high loads but face severe flow instability challenges. This study investigates a low-reaction transonic compressor rotor using casing treatment technologies to control unstable flow dynamics precisely. First, a design integrating self-recirculating and circumferential groove casing treatments near the leading edge is implemented. This design enhances flow capacity at the tip passage inlet. However, it causes a “stall transposition” phenomenon. The unstable flow structures shift from shock-tip leakage vortex interactions at the front to corner separation at the rear. Consequently, the stall mechanism transitions from an end-wall stall to a blade stall. To address this issue, a new casing treatment layout is proposed. Grooves are added after the mid-chord of the initial front casing configuration. This adaptation suppresses emerging unstable flow structures and extends the stall margin by approximately 12.07 %. Detailed flow field analysis shows that the rear grooves effectively reduced the trailing edge separation vortex. They also limit downstream corner separation and mitigate disturbances from tip leakage flows in the rear of the passage.

Numerical and experimental investigations of flow separation control through a linear compressor cascade

Modern large-scale gas turbines are equipped with high-pressure ratio compressors to increase engine work and its overall efficiency. Flow separation and energy losses are also two interrelated phenomenon associated with changes in compressor loading level and performance. This paper examines therefore the control of flow separation using a passive-control technique. An arced divergent-convergent slot grooved from the blade pressure side to its suction side was adopted to control flow separation, reducing the losses through a linear compressor cascade. The spanwise location of the slot was selected based on CFD simulations where the corner separation was predicted. The slot height in the spanwise direction was selected to be 8% of the blade height at the end-wall side. The present work was performed experimentally and numerically at an inlet Reynolds number,Rec=ρV∞C/μ=2.98×105\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$${R}_{ec}=\\rho {V}_{\\infty }C/\\mu =2.98\ imes {10}^{5}$$\\end{document}, covering a wide range of incidence angles from +6∘to-6∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$+ 6^{ \\circ } {\ ext{ to}} - 6^{ \\circ }$$\\end{document}. The experimental work was carried out using a linear cascade test section consisting of six NACA 65-009 blade profiles integrated into a low-speed wind tunnel. A five-hole pressure probe system was used to obtain main flow parameters. Numerically, four turbulence models, including Spalart–Allmaras (S–A) model, Realizable (R k-ε) model, Shear-Stress Transport (SST k-ω) model, and Reynolds Stress model (RSM) were tested to predict the velocity and pressure fields. Good agreement between the experimental measurements and the numerical results, which were obtained using the RSM turbulence model in terms of velocity profiles and total pressure downstream of blades. It was observed also that the use of the arced-slotted blades for positive incident angles was more effective in reducing the separation than the negative and zero incident angles, approaching a maximum value of 33% for 6° with enhanced blade loading reaching 17.6%. It is to be concluded that, the use of arced slotted blade improves the compressor performance specially for positive incident angles.

The loss mechanisms and vortex dynamics of rotor platform step geometries in a low-speed research compressor stage

The construction and assembly of an axial compressor blade row introduces real geometric features, which leads to the deviation of flow characteristics from the intended design. According to the realistic assembly errors observed in the four-stage low-speed research compressor developed by Shanghai Jiao Tong University, two representative step geometries at the rotor platform were abstracted. Unsteady simulations were carried out in the stage environment to study the influence of hub discontinuities on compressor performance and loss mechanisms. The comparisons of smooth hub, elevated hub (EH), and a wedged hub (WH) demonstrated that all the step geometries resulted in increased loss inside both rotor and stator rows. The loss induced by WH was more sensitive to step heights than EH. Under the influence of forward facing step, rotor incidence was reduced and horseshoe vortex (HV) was exacerbated, leading to the accumulation and separation of low momentum fluids at the pressure side corner. This led to a redistribution of losses across the span. In the lower half span, the relative total pressure loss increased almost linearly with step heights, while in the tip region, it reduced slightly. The backward facing step generated the step corner vortices (SCV) that shed and propagated downstream. The SCV interacted with the HV and cavity leakage vortex in the stator passage. Coupled with the augmented stator inflow angle, the corner separation at the suction side was dramatically intensified and led to a significant loss increase.

The impact of variable fillets on corner separation in a liner compressor cascade

The impact of variable fillets on corner separation in a liner compressor cascade

Synthesis of Multi-Functional Graphene Monolayers via Bipolar Electrochemistry.

Graphene has gained substantial research interest in many fields due to its remarkable properties among many other two-dimensional materials. In this study, we propose a wireless electrochemical approach, bipolar electrochemistry, for the precise modification of single layers of graphene at predefined locations, such as distinct edges or corners, with a variety of metals or polymers, thus enabling the elaboration of multi-functional monolayer graphene sheets. We illustrate the concept e. g. by depositing multiple metals, or platinum and a catalyst-containing porous polymer on the same graphene sheet, but at separate corners. This configuration allows activating chemiluminescence on the polymer spot, and simultaneously generates the driving force for autonomous motion on the Pt side through the catalytic decomposition of hydrogen peroxide into oxygen bubbles. This integration of different chemical features on the same object, exemplified by these proof-of-principle experiments, enhances the functionality of two-dimensional materials, paving the way for the use of these hybrid materials for a variety of applications, ranging from sensing and catalysis to targeted delivery.

Investigating the mechanism of flow of transonic compact crossover tandem diffuser

Dealing with high transonic airflow at the outlet of a mixed-flow impeller is a challenging issue in research on the design of efficient and compact diffusers. A low aspect ratio and an adverse pressure gradient in the passage of the diffuser can lead to flow separation on the hub side of the bend and the suction side of the aft blade. Tandem blading can be used to attain flow control through an appropriate relative circumferential tandem position (λ). However, a comprehensive understanding of flow in a compact crossover tandem diffuser (CTD) requires further investigation. This study provides a detailed analysis of the mechanism of flow in a compact crossover baseline tandem diffuser (CBTD) as well as the influence of different values of λ on the structure of flow of the CTD. The findings show that the CBTD can suppress secondary flow separation on the hub side of the bend by reasonably distributing the radial load. We also found that a value of λ of 75% can delay the point of separation on the suction side of the aft blade over a wide range of spans, where this inhibits the corner separation vortex and improves the structure of flow and performance of the diffuser. We also preliminarily explore the impacts of excitation of full-span jet intensities with different values of λ on the corner separation vortex on the suction side of the aft blade and offer guidance for the design of a compact and efficient CTD.

Experimental and numerical investigation of two blended blade and endwall configurations

The Blended Blade and Endwall (BBEW) technique has proven to be effective in controlling the intersection of boundary layer in corner region of the compressor endwall. In this study, the experiment and analysis of two different BBEW designs are emphasized. First, based on a linear cascade with 0.7 Mach number inflow, two different configurations, BBEW 1 and BBEW 2, are well conducted. Then, according to the experimental result, control effects of these two BBEW configurations are validated and compared subsequently under various working conditions. The results demonstrate a reduction in total pressure loss of 12.8% and 29% at the design point for BBEW 1 and BBEW 2, respectively. Consequently, the BBEW technique proves effective in suppressing the development of the boundary layer and preventing corner separation. Followed by detailed numerical analysis, improvements around the corner region, especially for the boundary layer, are extracted to show the mechanisms and distinctions between the two configurations. The results indicate that BBEW 1 significantly restrains the development of boundary layer before separation occurs, while BBEW 2 directly controls the strength and scale of the corner separation.

Variations of the shock and secondary flow structure in a transonic compressor cascade with outlet back pressure

The effects of back pressure on the transonic cascade operating state are crucial and can determine the structure of internal shock waves and secondary flows. In this paper, numerical methods validated by experiments were employed to investigate the evolution mechanisms of the inlet flow field, shock structure, secondary flow structure, and cascade performance under different back pressures. Analysis revealed that transonic cascade exhibited unique incidence characteristics in the inlet flow field under both subsonic and supersonic regimes, although these two regimes involved different physical mechanisms. The results revealed that the operating state of the transonic compressor cascade under the unique incidence condition was influenced by the outlet back pressure, and there existed a critical static pressure ratio. The critical static pressure ratio shifted from 1.61 for two-dimensional flow to 1.37 for three-dimensional (3D) flow at M1 = 1.1, due to the corner separation and the characteristics of 3D shocks. The 3D shock structure exhibited a non-uniform distribution along the spanwise direction due to the influence of back pressure and the separated boundary layer. The vortex structures analysis revealed that the secondary flow structure on the sidewalls of the transonic compressor cascade was primarily dominated by corner vortices, whose formation mechanism was related to the interaction between the shock wave and the sidewall boundary layer. Additionally, this interaction also led to the formation of detached shock and lip shock vortex structures. Finally, loss analysis indicated that the wake region of the transonic cascade primarily includes six types of loss, and the total loss of the cascade decreased with the rise in back pressure.

Corner Separation Control Using Sweeping Jet Actuator in a Compressor Cascade

Abstract In the pursuit of advancing active flow control (AFC) technology, a more promising nonsteady actuator known as sweeping jet actuator (SJA) has been equipped on the endwall to manage the corner separation in a compressor cascade, while the steady jet actuators with holes were used for comparison. Experimental investigations have been conducted in a low-speed wind tunnel with an inflow Mach number of 0.23. Five-hole probe measurements and the oil flow visualizations were carried out to demonstrate the performance and physics of steady and unsteady blowing on controlling the corner separation. The transient data at the measurement plane were also obtained using the dynamic pressure sensors to get insight into the unsteady characteristics. Variation of jet momentum of both actuators and the excitation frequency of the SJA allows determination of favorable control parameters. Using the SJA with only 0.13% of the cascade mass flow, the total pressure loss which takes the additional energy input into account is reduced by 14.2%, while the steady jet achieves a reduction of 8.4%. The results highlight the superior characteristics of the SJA in controlling the corner separation.

The Impact of Upstream Static Deformation on Flow Past a Cylinder/Flare

Reynolds-averaged Navier–Stokes simulations are performed for supersonic turbulent flow over a cylinder/flare with upstream surface distortion representative of structural deformation induced via fluid–structural and fluid–thermal–structural behavior. Broad parametric analysis is carried out through the generation of Kriging-response surfaces from a database of general simulations. A posteriori simulations are then carried out at parametric combinations that correspond to extrema in the Kriging response surfaces to gain deeper insights into the interaction between the surface distortion and flow responses. Upstream distortions tend to decrease, rather than increase, the peak pressure and heat flux loads on the flare compared to an undeformed cylinder. Furthermore, decreases in these quantities reach up to O(10%) compared to up to O(1%) for increases. Integrated quantities over the flare are relatively insensitive to upstream distortion. The corner separation length is the most sensitive quantity to upstream distortion, with protrusions tending to increase the separation length and recessions reducing the separation length. Modifications in the separation length of up to 40% are observed. Reductions in peak loads tend to correspond to increases in the corner separation length. The movement of the surface distortion relative to the corner indicates a negligible impact beyond 1.5 distortion lengths from the corner, and the largest impact on the corner separation length occurs when distortion is directly adjacent. These results are an important step toward understanding and quantifying the impact of surface deformations on downstream components.

Transonic Compressor Darmstadt Open Test Case: Experimental Investigation of Stator Secondary Flows and Hub Leakage

Abstract The future trend of clean sky aviation has led to a smaller core engine, resulting in highly loaded stages with increased relative tip gaps and leakage flows, which causes a reinforced influence of secondary flow phenomena. This article experimentally investigates the secondary flow effects in the 3D-optimized shrouded stator of the TUDa-GLR-OpenStage, representative of a transonic high-pressure compressor front stage, and its susceptibility to typical hub leakage flows. The impact of stator hub leakage on efficiency and total pressure ratio is investigated for several operating speeds, from subsonic to nominal transonic operating conditions. Steady five-hole probe and time-resolved virtual multi-hole probe measurements at different operating conditions at nominal speed are conducted to investigate the underlying loss mechanisms. The leakage mass flow is determined using the measured pressure difference within the fin seal. Steady results show that the optimized 3D stator effectively suppresses the hub corner separation compared to the predecessor 2D stator, but its performance is further limited by a near-tip separation. With the increased stator hub leakage (at a leakage-to-main flow ratio of about 0.4%), the compressor isentropic efficiency generally drops (by about −0.5%) at the full operating range due to an enhanced hub corner separation effect. However, the increased leakage also helps alleviate the tip-corner separation when operating near stall, leading to a smaller efficiency penalty under these conditions. Unsteady results reveal the sensitivity of transient compressor performance to the passing rotor wake, but limited interaction between this phenomenon and the increased leakage is observed. These findings provide insight into the stator hub leakage-related penalties on the compressor aerodynamics efficiency.

Simulation and analysis of the over-expanded flow field in asymmetric nozzles with lateral expansion

The study presented a numerical simulation of three-dimensional asymmetric nozzles featuring lateral expansion to investigate the flow field characteristics under over-expanded conditions and assess the influence of the lateral expansion angle. The research findings elucidate the distinct shape of the separation zones within the nozzle resulting from varying lateral expansion angles. In the restricted shock separation with separation bubbles forming on the flap pattern, the central separation bubble and corner separation bubbles exhibit independent characteristics. With an increasing lateral expansion angle under the same nozzle pressure ratio, the corner separation shock wave significantly impacts the shape of the central separation bubble.