The unstable flow with rotating-stall-like (RS) effects in a rotor-cascade of an axial Compressor was numerically investigated. The RS was captured with the reduction in mass flow rate and increasing of exit static pressure with respect to design operating condition of the single rotor. The oscillatory velocity traces during the stall propagation showed that the RS vortices repeat periodically, and the mass flow rate was highly affected by the blockage areas made by stall vortices. The results also showed that large scale vortices highly affects on the generation and growth of the new vortices. An unsteady two-dimensional finite-volume solver was employed for the numerical study which was developed based on Van Leer’s flux splitting algorithm in conjunction with TVD limiters and the κ-ε turbulence model was also employed. The good agreement of the computed mass flow rate with the experimental results validates the numerical study.

Large Eddy Simulation (LES) of a two dimensional supersonic Compressor cascade is performed in the current study. It is found that the Shock Wave Boundary Layer Interaction causes a large scale of total pressure losses and presents strong fluctuation features. Thus the pulsed and steady excitation jets are applied to suppress the flow separations and to reduce the total pressure losses. Several impacting parameters, such as jet axial location, jet hole width, jet angle to the local blade surface and jet mass flowrate are chosen based on the primary analysis by the calculations by the Reynolds Averaged Navier-Stokes equations. In addition, based on the results of frequency spectrum and POD analysis, the excitation jet frequency is chosen for the pulsed excitation jet scheme. It is concluded that the pulsed excitation jet scheme achieves a 9. 8% reduction of total pressure loss in comparison to the steady excitation jet scheme under the same time-averaged excitation jet mass flow rate. The excitation jets affect both the flow field near the jet hole on the suction surface and the flow field on the pressure surface via the management of the reflection shock wave. In addition, the excitation frequency dominates not only the time-averaged flow field, but also the second and third modes which stand for the unsteady structures in the flow field under the POD analysis. The first mode contains most energy in the flow field and the energy percentage decreases dramatically with the increase of the mode number. In comparison to the steady excitation jet scheme, the pulsed excitation jet scheme gathers more energy to the low orders of the modes, especially the first four modes. With the mixing effect and high dissipation rate of the high-frequency signals, the high-frequency signals shrink in the wake and the flow field builds up more uniformity.

The boundary layer's separation loss in Compressor cascades constitutes a significant portion of profile loss, critically influencing aerodynamic performance optimization and control. This study employs Large Eddy Simulation (LES) to examine separation losses at varying attack angles, focusing on a rectangular Compressor cascade. Specifically, it explores the long separation bubble at a 45% blade height cross-section under designed incidence. Analysis of the separation bubble's transition process revealed a notable surge in total pressure loss rate prior to transition, which stabilized following reattachment. The study thoroughly investigates the evolution of long bubbles, employing quadrant analysis of Reynolds stress, critical point theory, and an in-depth examination of individual vortex dynamics. The findings indicate that the peak of cross-flow within the separation bubble acts as the primary mechanism initiating the transition. This insight is corroborated by DNS calculations of natural transitions on flat plates. Building upon these findings, the study discusses the effects of varying attack angles on transition processes. Notably, increased incidence prompted the upstream migration of the long separation bubble, transforming it into a short bubble at the leading edge. This shift led to a fivefold increase in separation loss and doubled the frequency of transverse flow fluctuations.

The impact of the column number of ellipsoidal dimples on a highly-loaded Compressor cascade (NACA65-K48) under design conditions was investigated by using a numerical simulation method. Ellipsoidal dimples with a thickness of 0. 2 mm were located at the position of chord length ranging from 10% to 36%. The span-wise interval was 5. 0 mm. The performance and flow field structures of cascades with 1 to 5 ellipsoidal dimpled columns were compared, and the results showed that the turbulent kinetic energy intensity near the wall was enhanced and the fluid separation resistance was consequently improved. The total pressure loss was reduced by all modified ellipsoidal dimples. In addition, the separation bubble of the suction side was broken or weakened, the corner separation was improved, and the influence range of the passage vortex was reduced. Moreover, the improvement effect of cascade performance parameters initially increased with the increase in the number of dimple columns and then reduced as the number of columns was further increased. The reductions in the total pressure loss of the cascade were 0. 59%, 1. 47%, 1. 69%, 1. 91%, and 1. 73% for column numbers 1 to 5, respectively.

The effects of inflow variations due to the working environment and flight attitude changes on turbomachines are considerable in the real world. Nevertheless, uncertainty quantification can be adopted to assess mean performance changes and perform the aerodynamic shape design as well as optimization. Thus, an uncertainty quantification method of adaptive sparse grid collocation (ASGC) was first introduced to address the inflow uncertainties’ effect issue effectively and accurately. Then, ASGC was utilized to evaluate the impacts of inlet incidence perturbations at different perturbation scales and reference inflow Mach numbers on the aerodynamic performance of a controlled diffusion cascade. The results showed that compared with the Monte Carlo simulation and static sparse gird collocation, the statistical accuracy and response accuracy of ASGC were maintained, and meanwhile its model construction efficiency was significantly improved because of the nested adaptive sampling feature. Under the perturbations of inlet incidences with high reference incidences, the mean aerodynamic loss always aggravates. The changes in aerodynamic loss nonlinearly depend on the inlet incidence perturbations, and the nonlinear dependence becomes greater when the perturbation scale. expands. At the same perturbation scale, the nonlinear dependence on the inlet incidence perturbations is further enhanced when the reference inflow Mach number rises. Finally, uncertainty quantification of the flow field revealed that the fluctuation of flow accelerations at the leading edge plays a fundamental role in determining the uncertainty of the aerodynamic loss.

In this paper, due to the importance of incoming flow turbulence intensity into combustion chamber, tripping wire effect on the flow wake has been experimentally investigated within a linear Compressor cascade. To do this, two wires were implemented along each blade and their effects on average velocity, turbulence intensity and vorticity frequencies at Reynolds number 45500 were accurately considered. To measure wake parameters, single channel hot wire anemometer was used. Turbulence creation in response to the turbulence promoters made the separation to take place within the boundary layer in a distance farther from the edge of attack and also a decrease in wake width was observed. It is found that turbulence promoters increased the maximum turbulence intensity in blades wake and also reduced corresponding frequency in maximum amplitude and the Strouhal number, consequently.

Interest in plasma actuator as an active flow control has grown rapidly in the last years. Plasma actuator consists of a pair of electrodes that are separated by a dielectric material. Applying voltage to the electrodes, results in a body force that act on the flow field and is used in order to control it. Plasma actuator by imparting momentum is able to tangentially accelerate the flow field that can be used for flow control purpose such as boundary layer transition control, drag reduction, lift enhancement, and flow separation control. This work involves the documentation and control of leading-edge separation control that occurs on an axial Compressor cascade at high angle of attack. To study the effect of control technique, a 2-D numerical investigation were performed in presence of varying plasma actuator voltage and location in different flow characteristics such as stream line, pressure and lift-to-drag ratio. The results show that plasma actuator reduce energy losses and a lift-to-drag ratio enhanced of up to 18% can be obtain by using plasma actuator at 15% of the blade chord length. The control effect obtain by the plasma actuator in low Reynolds number is more effective and increasing the applied voltage improves the performance of the Compressor cascade by increasing the induced body force.

The Compressor cascade performance is significantly restricted by the secondary flow mainly presented as the trailing edge separation and corner stall. This paper develops a synthetic flow control approach in a high turning cascade using the vortex generator and slot jet approach. Numerical simulations were conducted to assess the flow control benefits and illustrate the flow control mechanisms. Four configurations, the baseline, the two individual approaches and the synthetic approach, were simulated to compare the separation control effects. The simulations show that all the three configurations achieve considerable improvements of the cascade performance and the cascade sensitivity to incidence angle is greatly decreased. The synthetic approach improves the most among them which is almost the superposition of the two individual ones. In the synthetic approach, the trailing vortex induced by the vortex generator suppresses the end wall cross flow and deflects the passage vortex, and then prevents the production of corner stall; at the same time, the slot jet speeds up the trailing edge separation caused by the cascade high camber. Owing to the combination of the two aspects, the synthetic approach restricts the developments of secondary flow and vortices in the cascade, and improves the outflow uniformity. The synthetic approach nicely utilizes the advantages of the two individual approach while avoids the shortages by the complementation, so it can achieve more powerful flow control effects. At the end, vortices models are established to illustrate the secondary flow structure and the flow control mechanisms.

In this work, measurement and analysis of the dimensionless frequency (Strouhal number) of vortex shedding of an axial Compressor cascade in moderate Reynolds numbers are carried out. Assessment of these measurements can help a more precise prediction of the wake-induced transition on the downstream blades. To measure the flow field in the wake, a hot film anemometry is used. The measurements are conducted at three different incidence angles and the Reynolds numbers ranging from 240000 to 530000 based on the blade chord length and flow velocity. Based upon these measurements, there is linear correlation between the vortex shedding and the Reynolds number, and by increasing the Reynolds number, the vortex shedding frequency increases. The results obtained show that the Strouhal number for the Reynolds number equal or below 360000 has a lower scattering compared with the Reynolds number above 360000. Also decreasing the attack angle increases the wake region. Moreover, the results obtained show that the vortex shedding frequency at moderate and low Reynolds numbers displays different behaviors, which could result in different boundary layer formations at the trailing edge of blades.