Compressor Blades Research Articles

An efficient sparse surrogate model for aerodynamic characteristics of a supersonic compressor cascade with uncertain geometric deformations

The increasing demand for high-load compressors necessitates research and design of supersonic compressor blades. Geometric deformations of the blades resulting from manufacturing errors introduce uncertainties in their aerodynamic characteristics. For uncertainty analysis, Polynomial Chaos Expansion (PCE) has gained popularity among engineers in various disciplines. However, as the number of random variables used to describe uncertainties increases, the efficiency of the PCE method diminishes. To overcome this limitation, an Efficient Sparse Surrogate Model (ESSM) was first proposed. The ESSM combines PCE to approximate global characteristics and Gaussian process modeling to capture local variability, resulting in a highly accurate model. Additionally, by utilizing an iteratively diffeomorphic modulation under observable response preserving homotopy, important PCE basis functions are adaptively selected, leading to a sparse surrogate model that significantly reduces the required training samples. Then using the ESSM, efficient Uncertainty Quantification (UQ) and global sensitivity analysis were conducted to evaluate the impact of uncertain geometric deformations on the aerodynamic characteristics of a supersonic compressor cascade. Principal Component Analysis (PCA) was employed in conjunction with an Improved Class Shape Transformation (ICST) to characterize geometric deformations caused by manufacturing errors. The proposed ICST method exhibits enhanced accuracy in fitting near the leading and trailing edges compared to the CST method. PCA enables a 50% reduction in the number of random variables needed to describe geometric deformations. The UQ results establish a quantitative correlation between geometric deformations and supersonic aerodynamics, and the sensitivity analysis identifies some specific forms of geometric deformations that significantly impact the aerodynamics. Finally, the study investigated the flow mechanisms associated with typical geometric deformation forms, further providing recommendations for manufacturing and testing of supersonic blades to prevent performance deterioration.

Unsteady Flows and Component Interaction in Turbomachinery

Unsteady component interaction represents a crucial topic in turbomachinery design and analysis. Combustor/turbine interaction is one of the most widely studied topics both using experimental and numerical methods due to the risk of failure of high-pressure turbine blades by unexpected deviation of hot flow trajectory and local heat transfer characteristics. Compressor/combustor interaction is also of interest since it has been demonstrated that, under certain conditions, a non-uniform flow field feeds the primary zone of the combustor where the high-pressure compressor blade passing frequency can be clearly individuated. At the integral scale, the relative motion between vanes and blades in compressor and turbine stages governs the aerothermal performance of the gas turbine, especially in the presence of shocks. At the inertial scale, high turbulence levels generated in the combustion chamber govern wall heat transfer in the high-pressure turbine stage, and wakes generated by low-pressure turbine vanes interact with separation bubbles at low-Reynolds conditions by suppressing them. The necessity to correctly analyze these phenomena obliges the scientific community, the industry, and public funding bodies to cooperate and continuously build new test rigs equipped with highly accurate instrumentation to account for real machine effects. In computational fluid dynamics, researchers developed fast and reliable methods to analyze unsteady blade-row interaction in the case of uneven blade count conditions as well as component interaction by using different closures for turbulence in each domain using high-performance computing. This research effort results in countless publications that contribute to unveiling the actual behavior of turbomachinery flow. However, the great number of publications also results in fragmented information that risks being useless in a practical situation. Therefore, it is useful to collect the most relevant outcomes and derive general conclusions that may help the design of next-gen turbomachines. In fact, the necessity to meet the emission limits defined by the Paris agreement in 2015 obliges the turbomachinery community to consider revolutionary cycles in which component interaction plays a crucial role. In the present paper, the authors try to summarize almost 40 years of experimental and numerical research in the component interaction field, aiming at both providing a comprehensive overview and defining the most relevant conclusions obtained in this demanding research field.

Direct numerical simulation and mode analysis of turbulent transition flow in a compressor blade channel

The separation and turbulent transition of the flow in a compressor blade channel are investigated through direct numerical simulations (DNS) at a Reynolds number of 1.367 × 105. Based on the original DNS data, both time-averaged statistics and instantaneous vortex structures of the flow field are extensively analyzed. The vortices are visualized and studied by the Liutex method, and the streaming dynamic mode decomposition (SDMD), a low-storage variant of conventional DMD, is applied to the large datasets obtained on both pressure and suction sides. The physical quantity analyzed with SDMD is the Liutex magnitude R. The DNS results indicate that flow separation occurs on both sides of the blade. On the pressure surface, the separation is weak and the flow remains in a natural transition dominated by viscous Tollmien–Schlichting instabilities. In contrast, owing to the presence of a large laminar separation bubble, the flow experiences a separation transition governed by inviscid Kelvin–Helmholtz instabilities on the suction surface. The SDMD results suggest that a broad range of vortex frequencies exist in the transition flow, and the scale of the spatial structures is negatively correlated with the frequency of the mode. On the pressure surface, the extracted SDMD modes are primarily related to Kelvin–Helmholtz rolls, whereas on the suction side, influenced by the separated boundary layer, the modal structures exhibit greater diversity.

“Modeling, microstructural characterization, mechanical and corrosion properties of the CrN/CrAlN multilayer coating produced by physical vapor deposition."

In this research, CrN/CrAlN multilayer coating was deposited on X12CrNiMoV12-3 stainless steel samples and compressor blades by the physical vapor deposition (PVD) method. Based on the computational fluid dynamics model, an optimized thickness of less than 10 μm with a roughness of 0.35 μm was deposited. The coating's microstructure and phase composition were characterized by XRD and SEM. Nanoindentation, adhesion, scratch, and roughness tests as well as residual stress were realized in the coated samples. A Salt spray test was also done on the samples and blades to evaluate the corrosion resistance of the coating. The CrN/CrAlN coating's elastic modulus and hardness were 371 GPa and 1745 HV, respectively. The Corrosion resistance of coated samples and compressor blades was 930 and 50 h, respectively. The induced residual tensile stress of 176.23 MPa due to the required machining process of the compressor blade before the coating was the reason for the lower corrosion resistance of coated blades. to eliminate the residual stresses and improve the corrosion performance of the coating, the blades were subjected to the tumbling process before coating. By optimizing the tumbling process for 3 h, the residual tensile stress converted to a compressive stress of −129.11 MPa, and accordingly corrosion resistance of the coated blades increased to 950 h.

Fringe Projection Profilometry for Three-Dimensional Measurement of Aerospace Blades

The aero-engine serves as the “heart” of an aircraft and is a primary factor determining the aircraft’s performance. Among the crucial components in the core of aero-engines, aero-engine compressor blades stand out as extremely important. They are not only numerous but also characterized by a multitude of parameters, making them the most complex parts in an aero-engine. This paper aims to address the trade-off between accuracy and efficiency in the existing measurement methods for asymmetric blades. Non-contact measurements were conducted using a structured light system composed of a stereo camera and a DLC projector. The point cloud data of the blades are processed using methods such as the PCA (Principal Component Analysis) algorithm, binary search, and least squares fitting. This paper established a fringe-projection profilometry light sensor system for the multi-view measurement of the blades. High-precision rotary tables are utilized to rotate and extract complete spatial point cloud data of aviation blades. Finally, measurements and comparative experiments on the blade body are conducted. The obtained blade point cloud data undergo sorting and denoising processes, resulting in improved measurement accuracy. The measurement error of the blade chord length is 0.001%, the measurement error of blade maximum thickness is 0.895%, compared to CMM (Coordinate Measuring Machine), where the measurement error of chord is 0.06%.

Split Air-conditioner Noise Source Identification through Sound Intensity Mapping

As the global warming takes place in years, the electrical appliances such as air-conditioner have been installed in most residential, industrial, or even in the education institution to provide better air quality and ultimately reduces the heat transfer in the ambient. Common use of split air-conditioner system to give maximum capacity in a space is responsible to the accumulated noise generated by the outdoor unit of the system. Therefore, an experimental study has been done to identify the noise rating and location of the highest noise level emitted by the system through different temperature and fan speed. The sound intensity mapping together with the sound power level will be applied to ensure the reliability of the data for the study purposes. Generally, the results indicated that the noise were concentrated at the compressor and fan blades whereas the level were observed to be higher with the lower temperature and higher fan speed operated. The highest noise level also reported as high as 99.08 dB(A) which located at the compressor. Hence it can be concluded that mitigation action is needed since considerable noise are generated by the application of air-conditioner particularly at the compressor and condenser fan.

Effects of Re on blade load temporal-spatial distribution and correlation with flow stability of transonic compressor under inlet distortion

Compressor blade operating conditions and aerodynamic load change periodically owing to inlet distortion. The effects of low Re are often unavoidable when the inlet distortion is encountered in real compressor operation. Blade load distribution and internal flow field are both affected by variations in Re. In this paper, the effects of Re on the blade load distribution patterns under 60° of total pressure distortion is investigated by numerical simulation. The distortion intensity has been set to 0.05, and four cases with Reynolds number index of 1, 0.3, 0.2 and 0.1 are chosen for further analysis. The findings demonstrate that the variation of Re affects the compressor flow field mainly by modifying the aerodynamic profile of the blade, which also results in the change of the temporal and spatial distribution of the aerodynamic load. Furthermore, the aerodynamic blade profile varies significantly by Reynolds number in different spanwise regions. The blade time constant decreases in the mid-span area, while it increases in the tip. This ultimately leads to a delay in the location of the tip maximum load, changing the starting position of the induced stall.

Energy absorbing nano-porous Ti layers assisted erosion-corrosion resistant Ti/TiN multi-layered coatings for gas turbine compressor blades

Solid particle erosion (SPE) resistant coatings extend the life of gas turbine compressor blades and improve the safety of flights. Among different types of these coatings, nanolayered multi-layered and nanocomposite coatings proved to be more efficient. Further increase in the efficiency of these coatings can be achieved by introducing energy-absorbing metal porous layers. Here, we report the development of three different types of SPE and corrosion resistant Ti/TiN multilayer coatings (thickness ∼9 μm) on Ti6Al4V substrates using a sputtering technique. The coatings contain three stacks of Ti/TiN (7.3 nm) bi-layers separated by two energy-absorbing Ti layers of different porosity in each coating. The Ti film deposited at 4 mTorr chamber pressure showed a mix of ω and α phases with high porosity. The SPE studies were conducted according to ASTM-G76 standards at a temperature of 400 °C. The corrosion properties of the coatings were studied by potentiodynamic polarization in a 3.5 wt% NaCl solution. The coatings with energy-absorbing porous Ti layers showed better SPE-resistant performance compared to the coating with dense Ti layers. Whereas, vice versa behaviour was in the case of corrosion resistance performance. SPE mechanism is also analysed using field emission scanning electron microscopy images of erosion scars.

Collaborative improvement of profile accuracy and aerodynamic performance in robotic grinding of transonic compressor blade leading edge

Blade machining accuracy is critical to engine service performance. However, the intricate curvature characteristics and exceedingly thin dimensions of the leading edge present formidable challenges in achieving precise grinding. This study introduces a novel robotic belt grinding framework aimed at collaboratively enhancing the profile accuracy and aerodynamic performance of the leading edge. The irregular contact area and pressure distribution of the narrow rigid-flexible grinding interface are calculated based on the Semi-Hertz model. Additionally, a time-varying material removal contour (MRC) model is developed for precision grinding with the pyramid-structured abrasive belt. The non-linear equations governing dwell time are transformed into a convex quadratic optimization problem, and the solution is obtained using the fastest gradient method with constraints. Subsequently, the dwell time vector is converted into a feed rate value for robot programming. Experimental results demonstrate that the size and shape accuracy at three typical sections of the leading edge are effectively controlled within the desired range. The mean value of the line profile error is reduced to 0.019 mm, and the standard deviation of the surface profile error shows a noteworthy 32% improvement compared to the previous method. Furthermore, aerodynamic performance analyses are conducted for the ground airfoils using NUMECA. Numerical simulation results reveal that shape accuracy significantly influences the flow field state within the channel, given the precondition of qualified dimensional error. The isentropic efficiency is increased most by 4.67% by this dwell time control grinding.

A data-driven polynomial chaos method for uncertainty quantification of a subsonic compressor cascade with stagger angle errors

The probability-based uncertainty quantification (UQ) methods require a large amount of sampled data to construct the probability distribution of uncertain input parameters. However, it is a common situation that only limited and scarce sampled data are available in engineering applications due to expensive tests. In the present paper, the Data-Driven Polynomial Chaos (DDPC) method is adopted, which can propagate input uncertainty in the case of scarce sampled data. The calculation accuracy and convergence of the self-developed DDPC method are validated by a nonlinear test function. Subsequently, the DDPC method is applied to investigate the uncertain impact of stagger angle errors on the aerodynamic performance of a subsonic compressor cascade. A family of manufacturing error data of stagger angles was obtained from the real compressor blades. Based on the limited measurement data, the DDPC method combined with Computational Fluid Dynamics (CFD) simulation is employed to quantify the performance impact of the compressor cascade. The results show that the performance dispersion under off-design conditions is more prominent than that under design conditions. The actual aerodynamic performance deviating from the nominal performance is not a small probability event, and the probability of deviating from the nominal loss coefficient and exit flow angle by more than 1% can reach up to 47.6% and 36.8% under high incidence i = 7°. Detailed analysis shows that stagger angle errors have a significant effect on the flow state near the leading edge, resulting in variations in separation bubble size and boundary layer thickness.

Performance prediction and loss evaluation of the carbon dioxide supersonic nozzle considering the non-equilibrium condensation

Carbon dioxide (CO2) is being considered as a promising working medium in energy conversion and refrigeration cycles due to its unique properties. When carbon dioxide flows with supersonic in turbo machinery (compressor), the non-equilibrium effect is enhanced due to the large change of fluid velocity, resulting in non-equilibrium condensation of the blade, which will seriously affect the performance of the compressor. Considering the similarities in flow characteristics between the nozzle and the compressor blade, the condensing flow of the blade can be predicted by simulating in a nozzle. The real gas model is used. The pressure and the nucleation rate are predicted based on the modified model, and the flow losses and thermal efficiency are analyzed in different states. The results show that the pressure variation in the nozzle aligns well with the experimental data. When the fluid transitions from subcritical to supercritical, the condensation interval decreases and the peak of the nucleation rate increases. The maximum supercooling decreases gradually. The flow losses are relatively large, and the thermal efficiency is low.

Investigation on unsteady characteristics of flow separation in a supercritical carbon dioxide centrifugal compressor

As one of the core components of supercritical carbon dioxide (S-CO2) closed Brayton cycle, the efficiency of S-CO2 centrifugal compressor plays a crucial role. Affected by the special thermophysical properties of S-CO2 fluid, the flow mechanism of S-CO2 centrifugal compressor is different from the centrifugal compressor of conventional fluid, such as air. Previous studies have found that the flow separation within the flow domain of this kind of compressor is easily to occur downstream the pressure surface. Many steady computational fluid dynamics (CFD) studies have been conducted on the S-CO2 centrifugal compressors, but few studies focused on the unsteady evolution of this flow separation. In this paper, the unsteady CFD simulation is carried out in the flow passage of a S-CO2 centrifugal compressor. The solution domain of CFD simulation includes compressor blades, diffuser and volute. The performance and the unsteady flow behavior of S-CO2 centrifugal compressor is obtained. Under low flow rate conditions, the flow separation on the pressure surface of the blade of S-CO2 centrifugal compressor is more likely to occur, causing the decrease of compressor performance. This flow separation has strong unsteady characteristics, which deteriorates several times during a rotation cycle, meanwhile the vortex shedding happens. At the time steps of vortex shedding, the pressure near the trailing edge of the impeller fluctuates greatly, indicating that this unsteady flow separation brings a large flow loss to the S-CO2 centrifugal compressor.

An investigation of the heat transfer characteristics of a droplet impacting a compressor blade

In this study, the effects of the heat transfer characteristics of jet droplets impacting compressor blades during the on-line washing process of an aero engine are investigated. A three-dimensional transient model of a droplet impinging on a blade is established, the heat transfer effectiveness calculation formula is modified, and the heat transfer process of the impact is numerically simulated using the volume-of-fluid method. The results show that the effective heat transfer area and heat transfer effectiveness increase with increasing blade curvature. The higher the droplet temperature, the higher the average heat flux and the higher the heat transfer effectiveness. For every 10 °C rise in the droplet temperature, the peak average wall heat flux increases by about 0.47 MW/m2, and the heat transfer effectiveness increases by about 0.01. The heat transfer effectiveness is inversely proportional to the diameter of the droplet, and an increase in the diameter reduces the heat transfer effectiveness. The increase in velocity causes the droplet to quickly detach from the blade and the effective heat transfer area to decrease, resulting in a decrease in the amount of heat transfer between the droplet and the blade, and thus a decrease in the heat transfer effectiveness.

Investigation on Aerodynamic Robustness of Compressor Blade with Asymmetric Leading Edge

Investigation on Aerodynamic Robustness of Compressor Blade with Asymmetric Leading Edge

Integrated Design of Axial Slot Casing Treatment and Blades in a Low-Speed Single Stage Axial Compressor

Abstract Casing treatment has been used to enhance the stability of axial-flow compressor by extending the stall margin; it also induces strong interaction with compressor blades, as demonstrated in the open literature. Revealing the mechanism of this interaction is a key to fully exploit the potential of the stability-enhancing capability of casing treatment without efficiency penalty. For this purpose, an integrated optimization design of rotor and axial slot casing treatment is implemented for a low-speed axial-flow compressor, and the stall margin is increased by 10.89% without penalty on the peak efficiency. Through detailed analysis, two important driving factors for the improvement of stall margin are illustrated: the suction effect that decreases the axial momentum of tip leakage flow, and the injection effect that increases the axial momentum of main flow. Meanwhile, it is found that the combination of forward-deflection axial slot casing treatment and forward-sweep rotor can maximize the enhancement of the momentum exchange, thereby adequately reducing the blockage region induced by tip leakage flow. More specially, the stator is then redesigned based on the matching mechanism between the stator and rotor with axial slot casing treatment in a stage. This redesign can considerably reduce entropy generation and flow separation in the passage and achieve an improvement of 0.87% in efficiency. The results hence demonstrate the superiority of integrated design of axial slot casing treatment and blades in improving both stall margin and efficiency.

A novel graph structure data-driven crack damage identification for compressor blade based on vibro-acoustic signal

The evolution of advanced sensing techniques and intelligent algorithms has significantly underpinned the growth of structural health monitoring and damage identification. Modern industry equipment like compressors, which are indispensable to the petrochemical and other process industries, usually operate under complex conditions including variable speed. The more vulnerable compressor components, such as the blades, are prone to diverse levels of damage over time. Existing research usually discusses the damage identification problem of blades under the Euclidean space, facing challenges in linking multi-source heterogeneous signals. This study introduces a novel approach, employing a graph-structured data-based method for identifying compressor blade cracks. It specifically focuses on variable rotating speed conditions, subsequently proposing an intelligent identification framework based on vibro-acoustic graph-structured data. Firstly, the affinity graphs made of vibro-acoustic damage signal are constructed to express the latent damage information beyond Euclidean space after Fourier transform and residual learning-based feature extraction for one-dimensional data. Then the developed multi-order graph convolutional network and domain discriminator layers are used to extract the domain-invariant damage features, which will be fed into the linear layer for class prediction. Finally, the method's efficacy is cross-verified through experiments with actual measurements on a compressor platform, specifically focusing on variable rotating speed cases.

Effects of Water-to-Air Mass Ratio on Long-Term Washing Efficiency and Erosion Risk in an Axial Compressor Under Online Washing Conditions

Abstract One of the main reasons for gas turbines’ performance losses is the deposition of dirt on the compressor blades. Dirt deposit has to be periodically removed to keep the engine performance as high as possible. This is the reason motivating the presence of online water washing systems in most compressor gas turbines. Such systems aim at cleaning the compressor blades to recover efficiency; thus, the larger the water flow, the better it is assumed the compressor is cleaned (fixing all the other conditions). In the present work, we simulate the long-term behavior of a real axial compressor, from the inlet to the first-stage rotor, subject to online water washing with different water flowrates. The frozen rotor approach is adopted to solve the flow field in the rotor region. Simulations are performed by using the unsteady k-ɛ realizable model coupled with a Lagrangian tracking of the injected liquid phase. Water droplet erosion is handled by using a semi-empirical model developed by the authors. In each simulation, 504,000 parcels have been tracked, providing statistically reliable predictions. To simulate the long-term evolution of the washing process, a discrete mesh morphing technique coupled with the use of specific scale factors is adopted. Each of the tested configurations is composed of three successive erosive steps up to the blade compressor end-of-life. By varying the water-to-air mass fraction (WAMF*), six different injection configurations are assessed in terms of long-time average washing efficiency and erosion risk. The results predicted show the dependence of the considered washing indices on water mass flowrate and set the stage for the development of a washing optimization tool, which can help the design and management processes. In scenarios where washing indices are given minimal importance and the objective is to reduce the risk of erosion, the optimal injection configuration was shown to correspond to a WAMF* value of 0.250. Conversely, when washing efficiency is prioritized, the optimal injection configuration has been shown to correspond to the case where WAMF* = 0.750.

Compressor Tip Leakage Mechanisms

Abstract The over tip leakage flow in an unshrouded compressor blade row is highly three dimensional, yet in order for aerodynamicists to analyze and improve designs, they must be able to simplify the problem down to a limited number of mechanisms. In this article, the behaviors of the dominant loss mechanisms are investigated using a multi-order methodology that combines rapid experimental tests of different geometries, detailed measurements in a large rotating rig, large numbers of industry-standard 3D Reynolds-averaged Navier–Stokes (RANS) simulations, and a single direct numerical simulation (DNS) of the datum geometry. The three loss mechanisms identified are ultimately caused by mismatches in flow velocity: separation of the flow as it enters the gap, mixing of the leakage jet with the mainstream close to the suction surface, and endwall shear acting on the jet itself at mid passage. This article is presented in three sections: First, the loss mechanisms are visualized and examined in detail using experiment, simulation, and models. Second, the uncertainty in industry-standard predictions is analyzed and improvements to turbulence modeling are presented. Finally, a matrix of blades with different 3D designs is used to investigate the balance of loss mechanisms and a reduction in total loss generation.

Enhancing heat transfer in MHD Falkner's‐Skan flow with thermal radiation, free convection and dusty fluid between parallel plates

AbstractFalkner's‐Skan flows are one‐dimensional flows widely used in fluid dynamics and boundary layer theory. They optimize heat exchangers and cooling systems, enhance turbine and compressor blades in turbomachinery, and provide consistency and quality in semiconductor fabrication. Additionally, they assist in pollution management and environmental impact assessments. The variety of Falkner's‐Skan flows clarifies fluid flow phenomena and their practical applications. This article aims to explore the impact of surface temperature on the flow behavior of dusty fluid in Falkner's‐Skan flow. The study focuses on the flow of dust particles in a channel formed by two infinite parallel plates. The study assumes that the particles are round and uniformly dispersed in the fluid. Furthermore, the article takes into account the effect of radiation on the energy equation. With the findings of this study, we hope to gain a better understanding of the dynamics of Falkner's‐Skan flow and contribute to the development of effective strategies for managing the flow of dusty fluids. The right plate's movement at free stream velocity causes the fluid to flow. Partial differential equations are used to represent the behavior of the flow. The Poincare‐Light Hill Technique yields exact answers. Visual representations of the temperature and velocity curves show the effects of different factors. It is possible to create graphic pictures of the fluid and dust particles using Mathcad‐15. Furthermore, critical fluid characteristics for engineers, such as skin friction and heat transfer rate, are analyzed and tabulated. These evaluations include the heat transfer rate at the wedge surface and the influence of this enhancement on surface viscous drag forces.

Effect of Leading-Edge Erosion on the Performance of Transonic Compressor Blades

In this paper, an experimental and numerical investigation of the effect of leading-edge erosion in transonic blades was performed. The measurements were carried out on a linear blade cascade in the Transonic Cascade Wind Tunnel of DLR in Cologne at two operating points with an inflow Mach number of 1.05 and 1.12. The numerical simulations were performed by ANSYS Germany. The type and specifications of the erosion for the study were derived from real engine blades and applied to the leading edges of the experimental cascade blades using a waterjet process, as well as modeled in detail and meshed within the numerical setup. Numerical simulations and extensive wake measurements were carried out on the cascades to evaluate the aerodynamic performance. The increase in losses was quantified to be 4 percent, and a reduction in deflection and a rise in pressure were detected at both operating points.