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Purpose The geometric parameters of the compressor blade have a noteworthy influence on compressor stability, which should be meticulously designed. However, machining inaccuracies cause the blade geometric parameters to deviate from the ideal design, and the geometric deviation exhibits high randomness. Therefore, the purpose of this study is to quantify the uncertainty and analyze the sensitivity of the impact of blade geometric deviation on compressor stability. Design/methodology/approach In this work, the influence of blade geometric deviation is analyzed based on a subsonic compressor rotor stage, and three-dimensional numerical simulations are used to compute samples with different geometric features. A method of combining Halton sequence and non-intrusive polynomial chaos is adopted to carry out uncertainty quantitative analysis. Sobol’ index and Spearman correlation coefficient are used to analysis the sensitivity and correlation between compressor stability and blade geometric deviation, respectively. Findings The results show that the compressor stability is most sensitive to the tip clearance deviation, whereas deviations in the leading edge radius, trailing edge radius and chord length have minimal impact on the compressor stability. And, the effects of various blade geometric deviations on the compressor stability are basically independent and linearly superimposed. Originality/value This work provided a new approach for uncertainty quantification in compressor stability analysis. The conclusions obtained in this work provide some reference value for the manufacturing and maintenance of rotor blades.

Twist angle errors along the blade radial direction are uncertain and affected by cutting force, tool wear, and other factors. In this paper, the measured twist angle errors of 13 sections of 72 rotor blades were innovatively analyzed to obtain the rational statistical distribution. It is surprisingly found that the under-deflection systematic deviation of twist angle errors shows a gradually increasing W-shaped distribution along the radial direction, while the scatter is nearly linear. Logically, the statistical model is established based on the linear correlation of the scatter by regression analysis to reduce variable dimension from 13 to 1. The influence of the radial non-uniform twist angle errors’ uncertainty on the aerodynamic performance of the three-dimensional compressor rotor is efficiently quantified combining the non-intrusive polynomial chaos method. The results show that the mean values of mass flow rate, total pressure ratio, and isentropic efficiency at the typical operating conditions are lower than the nominal values due to the systematic deviation, indicating that the under-deflection twist angle errors lead to the decrease in compressor thrust. The compressor’s stable operating range is more sensitive to the scatter of twist angle errors, which is up to an order of magnitude greater than that of the total pressure ratio and isentropic efficiency, indicating the compressor’s safe and stable operation risk increases. Additionally, the flow field at the tip region is significantly affected by twist angle errors, especially at the shock wave position of the near-stall condition.

To conduct detailed investigations about tip leakage flow in spatial and temporal scale at near stall, a total pressure probe, stereo particle image velocimetry (SPIV), dynamic pressure sensors array and validated numerical simulation are adopted at the near stall condition in a large scale iso-rotor compressor. Higher pressure regions are found propagating in circumference at near stall through different measurements technology. Completely different leakage vortex propagation characteristics are found assisted with SPIV technology at design and near stall condition, respectively. The rotor outlet total pressure dominant frequency shows an obviously non-synchronous characteristic, which is inconsistent with shaft frequency and blade pass frequency. The wall static pressure results and the vortex propagation frequency are also mutually confirmed nearly at 213 Hz. To investigate the propagation of pressure wave, an azimuthal mode analysis is adopted to get the spatial circular characteristic of disturbance. Furthermore, numerical results also show that the leakage vortex at near stall has propagated closer to the nearby blade pressure side, vortex broken position has advanced 10–20% chord length distance in the passage when compared with the design condition result. Rotating instability and the leakage vortex maintained a close association with the circulation propagation disturbance. Vortex line and blade frontal line intersection angle changed from 13° to 19° in a swing period.

To meet the stringent reliability requirements of rotor blades in turbomachines, greater effort should be devoted to improving both aerodynamic and structural performance in blade design. This paper introduces an aero-structural multi-disciplinary design optimization (MDO) method for compressor rotor blades using a discrete adjoint method and an equivalent stress model (ESM). The principles of the ESM are firstly introduced, and its accuracy in calculating equivalent stress is validated through comparison with a commercial program. Both the aerodynamic performance and the maximum equivalent stress (MES) are selected as optimization objectives. To modify the blade profile, the steepest descent optimization method is utilized, in which the necessary sensitivities of the cost function to the design parameters are calculated by solving the adjoint equations. Finally, the aero-structural MDO of a transonic compressor rotor, NASA Rotor 67, is conducted, and the Pareto solutions are obtained. The optimization results demonstrate that the adiabatic efficiency and the MES are competitive in improving multi-disciplinary performance. For most of the Pareto solutions, the MES can be considerably reduced with increased adiabatic efficiency.

The efficiency of a jet engine heavily depends upon the efficiency of its compressor. This study investigates the impact of leading‐edge tubercles on a transonic axial compressor. For this purpose, a CFD analysis has been performed for NASA Rotor 37. The method of investigation is based on the numerical solution of steady‐state, three‐dimensional Navier–Stokes equations using k‐ω‐SST (Shear Stress Transport) turbulence model. The accuracy of simulations is ascertained by comparing the numerical data with the available experimental data. Several configurations are considered by changing different tubercle parameters: amplitude, wavelength, and span‐wise location of the tubercles on the rotor blade. The results indicate an increase in efficiency for all the configurations considered for the modified rotor as compared to the corresponding baseline rotor with a maximum increase of 0.52%. The improvement in efficiency can be attributed to the higher outlet pressure achieved by the modified blade, which is 1.91% greater than that of the baseline blade. Mach number contours show that the location of the shockwave has been moved further downstream in the optimized case. Furthermore, new vortices are observed to be generated near the middle of the chord on the suction side of the tubercle model. Vortices formation has resulted in the redirection of the surrounding flow in an axial direction. Simultaneously, it has also contributed to loss which is associated with temperature increase. However, pressure gain at the outlet accomplished by redirection of flow outweighs the drawbacks of loss associated with temperature increase thus reducing the overall losses. Amplitude, wavelength, and span‐wise location of the leading‐edge tubercles on a transonic NASA Rotor 37 compressor are numerically optimized. An increase in efficiency by 0.52% compared to the corresponding baseline rotor is observed. The location of the shockwave moves downstream of the rotor with 1.91% higher pressure.

The aerodynamic performance of axial compressor rotors is negatively affected by the ingestion of boundary layer fluids upstream. As the boundary layer becomes thicker, the blade tip load increases and the local loss is aggravated, especially under off-design operating conditions. The major objective of this research is to evaluate the potential for novel blade sweep designs that can tolerate the ingested low-momentum boundary layer fluids. An optimization design approach using a surrogate model and genetic algorithm is employed. By altering the blade stacking line, the optimized sweep design is obtained. The flow mechanisms that enable the performance of the compressor rotor to be improved are fully analyzed, and the findings indicate that the aerodynamic advantages primarily stem from two key aspects. First, in the tip region, the blade loads are decreased at various chordwise locations and the interaction of the tip leakage flow with the mainstream is alleviated. As a result, the loss near the tip is reduced. Second, the blade sweep design alters the distribution of shock intensity across the spanwise direction, leading to a decrease in shock wave intensity in the mid-span region. This is beneficial in reducing the shock wave/boundary layer interaction strength at the trailing edge of the blade airfoil. Overall, after the sweep design has been optimized to ingest the upstream boundary layer, the compressor rotor experiences a 0.8% improvement in adiabatic efficiency compared with the baseline rotor, while preserving the total pressure ratio and stall margin. Additionally, the redesigned compressor retains the overall performance level under clean inlet conditions. This research provides a potentially effective blade sweep optimization design strategy that allows transonic compressor rotors to tolerate low-momentum upstream boundary layer incoming flows.

The flow loss in the blade passage of a supersonic compressor rotor mainly comes from the boundary layers on the blade surface and end wall, the shock wave, the shock wave/boundary layer interaction, and tip leakage flow; the instability is mainly caused by the shock wave near the rotor blade tip exiting the blade passage. This paper adopts an internal slot in the blade, with the inlet of the slot located at the leading edge of the blade and the outlet located on the suction surface of the blade, by using the momentum of the incoming flow to form a high-velocity jet to control the flow loss and improve the stall margin of the supersonic rotor. The mechanism of reducing flow loss by a slotting jet was studied, and a genetic algorithm optimization platform was further used for the coupled optimization design of the slot and blade. The numerical calculation results showed that the slotting jet can effectively suppress the development of the boundary layer on the suction surface while reducing the intensity of the shock wave, thereby reducing the loss of the boundary layer and shock wave, significantly improving the peak efficiency of the rotor, and increasing the mass flow rate at the peak efficiency point. The slotting jet can cause the shock wave in the passage to move downstream, thereby improving the stall margin of the rotor. Due to the strong shock wave in the blade passage near the blade tip, the slot outlet should be near and upstream of the shock wave; the shock wave in the middle and root regions of the blade is weaker, and the slot outlet should be located downstream of the shock wave.

As a core component of large marine engines, the compressor delivers robust and efficient power for propulsion. This study focuses on assessing and quantifying the uncertainty in the aerodynamic performance of a transonic rotor under various operating conditions, with the aim of investigating the impact of blade manufacturing variability on performance. Monte Carlo simulation (MCS) and sensitivity analysis were initially employed to identify parameters that significantly influence airfoil performance. Subsequently, a non-intrusive polynomial chaos (NIPC) uncertainty quantification model was developed to compare the effects of tip clearance deviation and surface geometry deviation on rotor performance. The study then analyzes how the geometric deviation at the different spanwise sections affects aerodynamic performance. The results reveal that geometric deviations have a more profound influence on aerodynamic performance than blade tip clearance. The impact of geometric deviations on average pressure ratio and efficiency of the transonic compressor rotor intensifies as the air mass flow rate approaches the near-stall point, while it decreases near the choking point. Interestingly, fluctuations in pressure ratio exhibit the opposite trend. Regarding spatial distribution, deviations in the upper half of the blade span (near the tip) exert a more dramatic influence on mass flow rate and pressure ratio fluctuation. A conceivable reason is that the inlet airflow velocity increases along the radial direction of the blade, and manufacturing variations in the same magnitude produce more notable relative geometric deviations in the upper half of the blade span. Centered on the machining tolerance guidelines for transonic compressor rotors, this work recommends stricter profile tolerance requirements for the upper half of the blade span.

A Lagrangian method is introduced to analyze the tip leakage vortex (TLV) behavior in a low-speed axial compressor rotor. The finite-time Lyapunov exponent (FTLE) fields are calculated based on the delayed detached-eddy simulation (DDES) results and identifying the FTLE ridges as Lagrangian coherent structures (LCSs). The computational method of the FTLE field in three-dimensional unsteady flow fields is discussed and then applied to the instantaneous flow fields at both the design and near-stall conditions. Results show that the accuracy of the particle trajectory and the density of the initial grid of the particle trajectory greatly affect the results of the FTLE field and, thus, the LCSs. Compared to the Eulerian Q method, which is calculated based on the symmetric and anti-symmetric components of the local velocity gradient tensor, the Lagrangian method has great potential in unraveling the mechanism of complex vortex structures. The LCSs show a transport barrier between the TLV and the secondary TLV, indicating two separate vortices. The aLCSs show the bubble-like and bar-like structure in the isosurfaces corresponding to the bubble and spiral breakdown patterns.

Accurate predictions of the blade response in a multi-row compressor is one of the most important tasks within the design process of compressor blades. Some recent studies have shown that the decoupled method considering only the stator disturbances cannot obtain accurate results for cases with strong rotor–stator interactions, especially for the interaction between the rotor and downstream stator, and the coupled method with multi-row configurations is necessary. Factors determine what computational domains to model need to be clarified to find a balance between accuracy requirements and computational costs. To this end, this study conducted full-annulus unsteady calculations with decoupled and coupled configurations to investigate the forced response of an embedded compressor rotor induced by upstream and downstream stator disturbances and rotor–stator interactions, respectively. The results show that the upstream IGV disturbances were dominated by the wake, and the IGV and S1 potential fields had little effect on the R1 response. Meanwhile, the IGV–R1 interactions and S1–R1 interactions were dominated by one cut-on mode, respectively. The comparisons of the blade vibration amplitude and the unsteady pressure field calculated by decoupled and coupled methods revealed the mechanism of the forced response, namely, for the R1 response induced by upstream aerodynamic disturbances, the dominant excitation source was the IGV wake, and the blade vibration amplitude can be predicted by the decoupled method. In terms of the response induced by downstream disturbances, the cut-on S1-R1-interaction mode was dominant and the use of the decoupled method without considering its influence will lead to an inaccurate prediction. This study concluded that the formation process of rotor–stator interactions was the key factor that determines whether the decoupled method or coupled method should be used, and analogized a process independent of the downstream stator disturbance. The results can provide a preliminary configuration for accurate and efficient blade response predictions and explain the reason why including downstream stator vanes is very important.