Total Pressure Loss Research Articles

Flow disturbance by intrusive instrumentation located on the surface of a compressor blade

In this paper, we present an experimental and numerical investigation into the flow disturbance by intrusive instrumentation in the form of spanwise pressure tubes mounted on the pressure and suction side of a compressor blade in a linear cascade wind tunnel at a moderate subsonic Mach number of 0.5 and a Reynolds number of 790,000. The pressure tubes were modelled as simplified bumps extending over the entire span with a width of 24% of the chord length and a height of 67% of the maximum blade thickness. The leading edge of the bumps is located at 30% of the streamwise surface length. We found that the total pressure losses generated by these pressure tubes are considerably larger when they are placed on the suction side of the blade. The suction-sided bump increases the total pressure loss coefficient by a factor of four, while the pressure-sided bump approximately doubles it. CFD simulations reveal that the suction-sided bump causes the flow to form a large, open recirculation zone, while for the pressure-sided bump, the flow reattaches near the trailing edge. In conclusion, we provide insights into the effects of pressure tubes on the losses generated by compressor blades. The mounted tubes significantly impact the flow and should be carefully considered during instrumentation design.

Endwall film-cooling in a cascade with filleted vane, upstream discrete slots, and holes

ABSTRACT In vane cascades, large filleted vanes influence aerodynamics, while upstream endwall film-cooling arrangement impacts aero-thermal performance. This study explores experimental investigations on endwall flows and film-cooling effectiveness in a filleted vane cascade using discrete slots and holes. Two fillet profiles are examined covering the leading-edge area to either the throat or 62% of the axial chord. Film-cooling employs two slots and two rows of holes in the leading-edge region with inlet blowing ratios ranging from 1.0 to 2.8. Results show variations in total pressure loss and Nusselt number between filleted and un-filleted vanes, highlighting optimal film-cooling effectiveness with the smaller fillet profiles.

Aerodynamic performance evaluation of subsonic compressor cascade blade with leading-edge damage

Abstract Material loss at the leading edge of in-service gas-turbine engine blades is detrimental to the aerodynamic performances of compressors. In this study, computational simulations and experiments using pneumatic probes were conducted to measure the total pressure loss of compressor cascade blades with material loss at their leading edges. The correlations between blade leading-edge damage geometries and cascade blade total pressure losses were analyzed. The geometrical features of blade leading-edge damage that most affect the cascade blade total pressure loss were distinguished, and the influences of leading-edge damage region cross-sectional shapes on the cascade blade total pressure loss were examined. Two methods for evaluating the total pressure loss of damaged cascade blades were developed using experimental and numerical data. In the first method, an attempt is made to correlate the cascade blade total pressure loss with a characteristic parameter that is defined using scale and angle parameters of the leading-edge damage region. The second method is based on the probability distribution function as well as the critical loss boundary of the geometrical parameters of the leading-edge damage region. The results of this study provide a better understanding of the influence of damaged leading edges on compressor cascade blade aerodynamic performance and can help establish low-cost methods for the estimation of compressor performance degradation.

Numerical investigation of the labyrinth leakage flow on the performance of a compressor cascade

This paper examines the effect of labyrinth seal structure on the evolution of vortex systems and the performance of the cascade. It is found that the labyrinth seal leakage leads to the circumferential development of the separation vortex, and the axial influence region of the concentrated shedding vortex also expands, which increases the flow loss and reduces the diffuser capacity. Among all the cases, the deteriorating effect of the cavity without a labyrinth seal is the most significant. This results in an 83% increase in total pressure loss compared to the prototype cascade. The straight tooth labyrinth and the straight tooth labyrinth with block structures can effectively reduce the leakage from the sealing cavity and weaken the deterioration caused by labyrinth seal leakage. Among these schemes, the straight tooth with block scheme has the most considerable effect, resulting in a 19% reduction in total pressure loss. As the incidence angle increases, the recirculation zone on the end wall of each scheme gradually shrinks toward the leading edge. Meanwhile, the static pressure rises, and the velocity decreases near the leading edge. Compared with the straight tooth scheme, the straight tooth with block scheme can reduce the high-loss zone by 14% at most. In addition, as the sealing clearance gradually increases, the inhibitory effect of the straight tooth with block scheme on the deterioration of cascade performance caused by the labyrinth sealing cavity is gradually enhanced compared to the straight tooth structure. The improvement effect is optimal under 0.5 mm sealing clearance.

Investigation on the integrated characteristics of n-decane/air rotating detonation combustor and supersonic turbine

In this study, the two-dimensional numerical simulations are conducted to study the flow field characteristics, turbine performance and loss mechanism of integrated system of rotating detonation combustor and supersonic turbine under two different directions of detonation wave propagation. The results indicate that the propagation direction of RDW affects the incident angle between OSW and guide vanes, resulting in different operating modes for aligned and unaligned modes. OSW in the turbine cascade undergoes the leading-edge shock, blade surface reflection and trailing edge diffraction. The backward-propagating rake-type shock envelope is formed due to leading-edge shock of the rotor. In misaligned mode, the stator has higher damping of both pressure and temperature. The stator significantly improves the circumferential uniformity of both velocity and pressure, particularly when operating in aligned mode. The total pressure loss in aligned mode is less, therefore the turbine achieves higher rim work and efficiency. The viscosity is one of the main sources of loss in the flow field. The reflected shock waves at the leading edge of the rotor and in the stator cascade are the primary factors contributing to the leading-edge vortices on the stator vanes.

Aerodynamic investigation of an S-shaped intake employing sweeping jet actuators based on proper orthogonal decomposition

The central curvature change of the S-shaped intake exacerbates the three-dimensional flow separation, resulting in total pressure and swirl distortions at the outlet. In this paper, an efficient active flow control technique, the sweeping jet actuator, was applied to control flow separation. The instantaneous characteristics of the flow field inside and outside the actuator were revealed by using validated numerical simulation methods. In addition, the effect of the sweeping jet on the aerodynamic performance of the S-shaped intake was investigated. Proper orthogonal decomposition (POD), an analysis method used to extract the main feature components of data, was applied to decompose the flow field structure at the S-shaped intake outlet in this study. The results highlight the external flow field characteristics of the actuator, with a jet sweeping range of up to ±40° and a sweeping frequency of 1 500 Hz, corresponding to a Strouhal number of 0.018. When the jet momentum coefficient is only 0.116%, the total pressure loss coefficient and pressure distortion index are reduced by 5.6% and 24.76%, respectively. This study demonstrates the critical role of momentum injection and discretization in inhibiting flow separation in the channel. The POD-based analysis indicates that a sweeping jet with high momentum can significantly enhance the energy in the flow passage and reduce the vorticity intensity of the first three modes. The stable mode structures exist near the outlet, exhibiting excitation frequency and frequency-doubling properties. The findings of this research provide essential input conditions for subsequent distortion studies.

Effect of Local Concave Non-axisymmetric Endwall Profiling on Endwall Secondary Flows of a Highly-Loaded Turbine Cascade

Non-axisymmetric endwall profiling (NAEP) has been widely utilized in reducing secondary flow loss of turbines. However, most of NAEP are designed for the endwall of the entire blade passage, which presents challenges to the design of cooling structures of turbine endwall. This study aims to explore the local concave non-axisymmetric endwall profiling (LCNP) method with the same effect as whole passage NAEP, and reveal the influence mechanism of LCNP on endwall secondary flow structures of a highly-loaded turbine cascade. Under the condition that the maximum depth of LCNP is unchanged, the axial length effect and pitchwise location effect of LCNP are studied, the influence mechanism of LCNP on the secondary flow loss is analyzed, and the genetic algorithm is utilized to optimize LCNP at the optimal position. Results show that as the axial length of the LCNP increases and the pitchwise location gets closer to the suction surface, the intensity and range of the passage vortex are decreased, and the total pressure loss coefficient (loss coefficient) of the turbine cascade is decreased. When LCNP is 100% axial chord in length and at the position of 2/9 pitch, the loss coefficient is reduced by 5.49%. LCNP was optimized at the optimal position, and the optimal LCNP reduced the loss coefficient of the turbine cascade by 6.73%. After the local concave endwall profiling, the loading in the middle of the endwall of the turbine cascade is reduced, and the intensity of the passage vortex is effectively inhibited, which is the mechanism that the loss coefficient of the turbine cascade is reduced. However, after the local concave endwall profiling, the loading in the trailing of the endwall of the turbine cascade is increased, the transverse migration of the new boundary layer in the endwall is accelerated, and the loss coefficient of the near endwall is increased.

Optimizing Flow and Heat Transfer With Guiding Pin Fins in a Wedge-Shaped Cooling Channel for Gas Turbine Blade Trailing Edge

Abstract In this study, an innovative guiding pin-fin array optimization has been developed through three-dimensional numerical simulations to enhance the cooling efficiency of gas turbine blade trailing edge. The guiding pin-fin array is optimized using the Kriging surrogate model and Genetic Algorithm (GA), aiming to significantly improve the heat transfer rates and uniformity within the wedge-shaped channel. The design parameter chosen for optimization is the deflection angle of each guiding pin fin, and the optimization process is conducted in two rounds. The first-round optimization yields a first-optimized guiding pin-fin array, which exhibits superior overall heat transfer performance and reasonable pressure loss compared to conventional circular, oblong, and parallel pin-fin arrays. For the first-optimized guiding pin-fin (1st-OGP) channel at the Reynolds number of Re = 50,000, the total Nusselt number and pressure loss are 44.1% higher and 9.9% lower than those of the baseline circular pin-fin array (CP), respectively. An experimental validation using the transient liquid crystal (TLC) thermography method is carried out and proves the effectiveness of the optimization process. However, it is noted that the first-optimized guiding pin-fin exhibits even lower heat transfer at a low Reynolds number of Re = 10,000, particularly in the channel middle region, which is mainly due to the incapable turning flow control in the root region of the wedged channel. To address this issue and further improve the heat transfer performance at low Reynolds numbers, a second-round optimization is performed by specifically adjusting the deflection angle of the selected guiding pin fins near the root region of the wedged channel. This secondary optimization demonstrates significant heat transfer improvements over the whole studied Reynolds number range with a reasonably reduced consumption of computational resources. The total Nusselt number and pressure loss are 69.3% higher and 11.9% lower than those of the baseline circular pin-fin array, respectively, at Re = 50,000. The optimization process proposed in this paper produces a high-performance cooling structure design with elaborate guiding pin-fin arrangements in the wedge-shaped channel, which indicates high heat transfer enhancement and relatively lower pressure loss in the wedged channel for the turbine blade trailing edge.

LES investigation of a H2-fueled supersonic combustor: A two-strut injection approach

Efficient combustion schemes are increasingly vital as scramjet expansion into higher Mach numbers emerges as a leading trend. Thus, in the current work, an LES investigation on mixing improvement caused by a two-strut injector has been carried out for a Mach 2.5 model H2 fueled scramjet combustor. The research focuses on understanding the flow field, flame lift-off characteristics and combustion stabilization in the two-strut combustor. The simulation results indicate that the two-strut configuration exhibits a broader pattern of temperature fluctuation, implying a more efficient spread of combustion downstream compared to the single-strut configuration. Analysis of Fast Fourier Transform (FFT) pressure data reveals inherent instability in the combustion processes for two-strut under supersonic air inflow conditions. The findings also demonstrate the significant improvement in mixing performance achieved by the two-strut design compared to conventional struts. The primary mechanism is a two-strut's ability to create significant streamwise vorticity, which can improve the mixing of H2 fuel with supersonic air. According to the mixing and total pressure loss plots, a two-strut produces a rapid complete mixing at 0.182 m at the expense of a minute rise in total pressure loss. Finally, the entropy change plot reveals that the two-strut configuration exhibits a reduced entropy change compared to the single-strut configuration. This is mainly due to enhanced air-fuel mixing, efficient shock wave management, and enhanced flame stabilization in the two-strut setup. These factors collectively contribute to reduced irreversibilities and disorder in the flow, leading to lower entropy within the system.

Aerodynamic Feasibility Investigation of the Novel Experimental Rig on Turbine Blade Tip With High Relative Motion

Abstract The relative casing motion can significantly alter the over-tip-leakage (OTL) flow aerodynamic performances. Traditional tip experimental facilities for relative motion researches, including stationary linear cascade with low-speed moving belt, annular cascade with rotating outer wall, and full-scale experimental rig with rotating blade, were either extremely costly or not capable to reproduce engine-representative casing Mach number. Very recently, a novel design concept for high-speed disk rotor rig has been proposed for tip research. Subsequent to this research, the present study evaluates the aerodynamic performances and the periodicity of the disk rotor rig design using Reynolds-averaged Navier–Stokes computational fluid dynamics simulation. Furthermore, first-of-its-kind experimental results of total pressure loss distributions measured by five-hole probe under high relative casing Mach number are reported. The results demonstrate the applicability of the disk rotor experimental rig for tip aerodynamic investigations.

Numerical investigation on influence of different geometrical cavity flame holders and flame stabilization mechanism in hydrogen fueled multi-strut-based scramjet combustor

A detailed computational study was carried out to explore how the presence of a cavity and cavity angles influence the performance and combustion mechanisms within a scramjet combustor. This research utilized ANSYS Fluent 2021 to create hexahedral meshes and simulate the complex fluid dynamics in scramjet engines equipped with dual strut injectors. The simulations addressed various critical aspects of fluid mechanics, including the continuity, momentum (via the Navier-Stokes equations), and energy equations. These equations were used to model non-reacting flows, while the species transport equations were applied to analyse reacting flows. The turbulence within the combustor was modelled using the standard K-ε model, a common choice for such high-speed aerodynamic evaluations. The study's primary goal was to assess the effects of different geometrical configurations of cavity flame holders on the performance characteristics of an H2-based scramjet combustor overall efficiency, focusing on mixing efficiency, combustion performance, and total pressure loss. Both configurations with and without a cavity were examined. Key findings indicate that incorporating a cavity into the combustor design leads to developing a robust recirculation zone within the cavity area. This recirculation zone is pivotal in enhancing fuel-air mixing and combustion efficiency, with cavity-based combustors showing an earlier onset of combustion and achieving peak combustion efficiencies around 90–95%. The extent of the recirculation region is notably influenced by the proximity of the strut injector to the cavity's length. Additionally, the presence of a shear layer within the cavity generates a weak shock at the cavity's trailing edge. This shock interaction can adversely affect scramjet combustor performance, especially at higher cavity angles (α) and specific geometric configurations, such as an L/D ratio of 4 and α=30°. The L/D ratio of 4 and α=30° combustor model results in a 13.8% increase in turbulence intensity, which raises air-fuel mixing and there by improves mixing and combustion efficiency by 14.23% and 13.34%, compared to the other depth of the cavities. This advantage is critical, especially considering the compact length of the combustor, which is a desirable attribute in scramjet design.

Shock waves characteristics and losses estimation of non-equilibrium condensation flow in nozzle and steam turbine cascade

The investigation of transonic non-equilibrium condensation is paramount due to its profound influence on the shock patterns and the distribution of losses. This study numerically investigates the shock wave morphology, evolution, and quantitative loss calculation under the influence of non-equilibrium condensation within the Moses-Stein nozzle and Dykas cascade. The results indicate that the vapor near the nozzle wall undergoes phase transition first and forms “X-shaped” condensation shock, which impervious to back pressure variations. Conversely, the “λ-shaped” aerodynamic shock moves towards the throat with increasing back pressure, triggering boundary layer separation. The entropy generation of “λ-shocks” accounts for over 98% and represents the primary source of shock losses. Strong pressure side shock induces shock wave-boundary layer interactions with the blade suction wall, resulting in reflected shocks. However, the latent heat released by non-equilibrium condensation mitigates this effect, increasing the total pressure loss coefficient by 0.188. The decrease in Mach number under high back pressure causes a notable increase in the trailing edge shock angle, with the wave foot shifting upstream. Notably, the proportion of shock losses attributed to the suction side of trailing edge exceeds 80%, significantly surpassing other regions, which progressively increases with rising back pressure. Therefore, it is crucial for cascade optimization. The findings can provide references for the optimization design of low-pressure steam turbine blade profiles. The results can serve as a guide for optimizing the blade profiles of low-pressure steam turbine.

Understanding of the asymmetric twin-stroll radial turbine under steady and pulsating inlet conditions

Utilize asymmetrical twin-scroll turbines (ATSTs) to achieve high-pressure exhaust gas recirculation (EGR) with a single scroll while optimizing the other scroll’s design for engine scavenging optimization. ATST performance prediction relies heavily on understanding the turbine flow mechanism, which always faces the challenge posed by time-dependent non-uniform intake. This paper studied its performance and flow field under steady and unsteady intake conditions for a production ATST. Firstly, the steady performance of the ATST with different asymmetries (ASYs) under different scroll pressure ratios (SPRs) is investigated via the computational fluid dynamics (CFD) method. Results demonstrate that the distribution relationship between SPR and turbine flow parameter (MFP) has a certain symmetry based on SPR = 1. For efficiency, the imbalance of inlet conditions will lead to the reduction of turbine efficiency. Moreover, the ASY has little effect on the turbine’s efficiency under the condition of SPR = 1. Also, the unsteady CFD simulations are conducted under pulsating inlet conditions with different scroll averaged pressure ratios (SAPRs). The results show that the highest efficiency was achieved at an ASY of 0.5 under a SAPR of 1.2 conditions, and the maximum, minimum, and average efficiencies were 1.9%, 1.5%, and 0.9% higher than those at an ASY of 0.83. The turbine’s flow field at the optimal incidence angle point is analyzed. The secondary flow vortex in the volute of an ASY of 0.5 is more robust, and the volute outlet ‘s total pressure loss is more significant. However, a relatively good incidence angle is obtained at the rotor inlet, and the loss in the rotor flow field is more minor. Therefore, the turbine efficiency with an ASY of 0.5 is higher.

Effects of Perforated Plates on Shock Structure Alteration for NACA0012 Cascade Configurations

To alleviate the shock boundary layer interaction adverse effects, various active or passive flow control strategies have been investigated in the literature. This research sheds light on the behavior of perforated plates as passive flow control techniques applied to NACA0012 airfoils in cascade configurations. Two identical perforated plates with shallow cavities underneath are accommodated on the upper and lower surfaces of each airfoil in the cascade arrangement. Six different cascade arrangements, including a baseline configuration with no control applied, are additively manufactured, with different perforated plate orifice sizes in the range of 0.5–1.2 mm. A high-speed wind tunnel with Schlieren optical diagnosis and wall static pressure taps is used to investigate the changes in the shock waves pattern triggered by the perforated plates. Steady 3D density-based numerical simulations in Ansys FLUENT are conducted for further analysis and validation. In the cascade configuration, the perforated plates alter the shock structure, and the strong normal shock wave is replaced by a weaker X-type shock structure. Eventually, a 1% penalty in overall total pressure loss is induced by the perforated plates because of the negative loss balance between the reduced shock losses and the enhanced viscous losses. Further studies on perforated plate geometrical features are needed to improve this outcome in a cascade arrangement.

Numerical and experimental study of the combustion performance of an integrated inclined combustor

To shorten the length of an aero-engine combustor and reduce its weight, we propose an integrated inclined combustor. A combination of numerical calculations and experimental research methods is employed to conduct relevant research on the cold and combustion flow fields and ignition dynamic processes of the integrated inclined combustor. Results show that the inclination angle of the swirler has minimal effect on the total pressure loss and combustion efficiency of the integrated inclined combustor. However, as the swirler’s inclination angle increases, the uniformity of the outlet temperature distribution and recirculation flow rate of the primary combustion zone decrease. Furthermore, when the integrated inclined combustor is ignited, the flame does not propagate symmetrically to both sides in the circumferential direction but propagates faster along the inclination direction of the swirler.

The Law of the Total Pressure Loss in Supersonic Streamwise Corner Flow

Abstract The total pressure loss is a critical issue in the design of internal flow channels for hypersonic vehicles. In the flow through the streamwise corner region, the adjacent walls induce secondary flows that exacerbate the viscous effects, thereby intensifying the total pressure loss. This study investigates the impact of the corner’s dihedral angle and the compressibility of the flow on the total pressure loss in the streamwise corner region. The results indicated that as the dihedral angle increases, the distribution of the total pressure contour in the corner region became significantly distorted. The proportion of the boundary layer area within the corner region concurrently increased as the dihedral angle deviates from 90 degrees. Furthermore, we deduce that the total pressure loss is the largest in the supersonic flat plate boundary layer. This paper normalized the total pressure loss coefficient for different dihedral angles in the corner region and established an empirical relationship for the distribution of normalized total pressure loss coefficients along the flow path. It is discovered that under supersonic inflow conditions, the distribution law of the total pressure loss in the corner region exhibits higher predictive accuracy.

Effects of bird feather-like convergent–divergent-riblet surfaces on the total pressure loss and wake turbulence of a transonic compressor cascade

The flow loss caused by the fan blades in a turbofan engine with a large bypass ratio is significant, and the wake considerably affects the inlet flow of downstream components. Surfaces with bird feather-like convergent–divergent (C–D) riblets have been proven to modulate the boundary layer flow by inducing counter-rotating rolling modes; however, the effects of these surfaces on the total pressure loss and wake turbulence of transonic compressor cascades remain unexplored. In this study, the effects of C–D-riblet surfaces on the total pressure loss and wake turbulence of a transonic compressor cascade were experimentally investigated using a five-hole probe and hot-wire measurements. The flow-loss-control effects of C–D-riblet surfaces with different characteristic lengths were also analyzed. The most significant reduction in the area-averaged total pressure loss (11.23%) was achieved using a C–D-riblet surface with a characteristic length of 30 μm at a Mach number of 0.94; this total pressure loss reduction corresponded to an increase in the mean velocity and a decrease in the turbulence intensity of the wake profile. Furthermore, the effectiveness of the loss control varied significantly with the spreading position of the C–D riblets. The optimal control effect was observed in the divergence-line region, and the control was slightly less effective as the measurement position neared the convergence line. This paper demonstrates the promising potential of using C–D riblets to achieve flow loss control in transonic compressor cascades.

Internal flow analysis and design optimization of a membrane oxygenator

Abstract The membrane oxygenator is an essential component in the Extracorporeal Membrane Oxygenation (ECMO) system to offer temporary support to the respiratory system. This study aims to optimize the hemodynamic performance and reduce the thrombosis risk of a membrane oxygenator prototype using computational fluid dynamics. Numerical simulations of steady laminar flow in a full-scale oxygenator prototype (model 1) and two optimized structure Models 2 and 3 at flow rates of 5∼7L/min are carried out using the porous media model. Flow-field-based hydraulic performance indicators and the thrombus risk indicator of the three models are compared extensively. Detailed internal flow analysis revealed that adverse flow states such as insufficient flow circulation in the inlet shunt cone zone, large flow separations at the top corners of heat exchangers, and intensive flow impingement at the exit elbow tube are notably improved in the optimized model 2 and 3. The improvement is more significant at high flow rates. Performance parameters quantitatively validate the effect of optimized configurations. Specifically, at a flow rate of 7L/min, the flow uniformity indexes for the original model at the shunt cone exit increase from 0.884 to 0.923 and 0.890 in the two modified models. The total pressure loss is reduced by over 14%, and maximum wall shear stress is notably reduced from 241.46Pa to 135.9Pa. Additionally, the optimized models exhibited lower thrombus risk. The optimized designs emphasize the importance of smooth transitions between cross-sections and minimizing abrupt changes in flow direction. The employment of flow-field-based parameters allows for the establishment of the relationship between geometric and performance parameters to guide effective design optimization and ensure the safe clinical operation of oxygenator prototypes.

Flow field reconstruction of compressor blade cascade based on deep learning methods

In the design process of compressors, calculating the S1 flow surface involves solving the Navier-Stokes equations, which results in slow convergence and makes determining cascade characteristics time-consuming. However, deep learning offers significant advantages in flow field reconstruction by not only automatically extracting complex flow features and reducing prediction time but also providing high accuracy in reconstruction. This paper implements the rapid reconstruction of the compressor S1 flow surface cascade flow field using two deep learning models: U-Net and 1D-CNN. Using a double-circular arc airfoil as an example, we selected four key design parameters that define the geometry and position of the airfoil, ultimately designing 5,292 sets of cascade geometries. By performing batch meshing and CFD simulations, we built a cascade flow field dataset. The U-Net neural network uses design parameters as input and outputs the aerodynamic distribution of the cascade flow field. After training, it can directly predict the flow field based on the design parameters. Since the U-Net model cannot directly obtain the aerodynamic parameter distribution and flow field aerodynamic coefficients on the airfoil surface, a 1D-CNN model is used as a complementary approach. The 1D-CNN model takes the design parameters as input and outputs the aerodynamic parameter distribution on the airfoil surface and the flow field aerodynamic coefficients. The prediction results show that the U-Net model achieves an average relative error of less than 1% in cascade flow field reconstruction, while the 1D-CNN model achieves an average relative error of less than 1% in predicting the pressure recovery coefficient and less than 2% in predicting the total pressure loss coefficient. This study presents a method for the rapid reconstruction of compressor blade cascade flow fields, which helps improve design efficiency and shorten the design cycle.

Study on the transition mechanism of vibrating low-pressure turbine blades based on large Eddy simulation

The low-pressure turbine blades are susceptible to vibration issues due to their thin profiles and large aspect ratios. Blade vibration will significantly affect the evolution of the boundary layer and the flow state. This paper utilizes large eddy simulation to predict the development of the boundary layer on the suction side of low-pressure turbine blades at low Reynolds numbers (Re = 25,000). It introduces different vibration cases to elucidate the mechanisms by which blade vibrations influence boundary layer separation and transition. The study demonstrates that the introduction of vibration cases significantly reduces both the size of the overall spanwise vortices and their roll-up height. A staggered distribution of spanwise vortices, characterized by alternating high and low regions, is observed near the trailing edge of the vibrating blades. The shorter spanwise vortices develop rapidly, nearly traversing the process of hairpin vortices (Λ vortex) generation and development, and directly breaking down into smaller-scale vortices. This accelerates the transition process. Blade vibration primarily promotes turbulence reattachment by facilitating the transition process dominated by the K-H instability mechanism within the separating shear layer. Consequently, it effectively restricts the growth of the separation bubble on the suction side of the blades, significantly reducing aerodynamic losses. Moreover, increasing the vibration frequency within a certain range can amplify these effects, achieving up to a 23% reduction in total pressure loss compared to stationary blades.