A major concern in the Computational Fluid Dynamics (CFD)-based throughflow calculation is the treatment of the incidence and deviation. This paper investigates the effects of the incidence and deviation on the CFD-based throughflow analysis by a time-marching throughflow model. The model is realized by solving complete Navier-Stokes equations with a single grid in the g-wise direction. The inviscid blade force is determined by calculating a pressure difference between the pressure and suction surfaces. The losses are introduced by imposing a blade surface skin friction factor converted from the pressure loss coefficient. And the flow discontinuity problem at the leading and trailing edges is resolved by modifying the mean blade surface to accommodate the incidence and deviation. The sensitivity of the throughflow results to the modification strategy of the mean blade surface is studied through response surface method. And the Kriging-assisted genetic algorithm (GA) is applied to determine the optimal distributions of incidence and deviation in the streamwise direction for NASA Rotor 37. Finally, four examples are provided to validate the throughflow model and to demonstrate the effects of incidence and deviation on CFD-based throughflow analysis at off-design conditions.

The flow fields in rear stages of multi-stage axial compressor is difficult to measure in detail owing to the limited height and space. Thus, low speed research compressor (LSRC) facilities which are modelled from rear stages have been widely used to explore the internal flow fields and improve compressor design. A newly-designed vertical LSRC facility is established and put into used in Shanghai Jiao Tong University. The construction and design features of this LSRC facility are introduced in this paper. A cantilevered stage has been tested in this test rig. Compressor performance, inter-stage parameters distributions and contours are measured at design point and near stall point. Steady single passage simulations for four-stage compressor are carried out to validate numerical methods and further interpret the internal flow fields in test stages. This vertical LSRC facility consists of inlet guide vane (IGV) and four repeated stages with an external diameter of 1.5 meter and a rotating speed of 900 RPM. The third stage is the mainly tested one, while the first and second stages provide the inlet conditions and the fourth stage provides the outlet conditions. Complete measuring methods and systems are established for this newly-built LRSC facility. The measurements of overall performance and inter-stage flow fields are carried out for test stage with cantilevered stator rows. The simulation for four-stage compressor are also performed for cantilevered configuration. The results of steady single-passage simulation have a similar trend with experimental ones, in terms of overall performance and parameters distributions.

Modern axial compressors are designed to be highly loaded in terms of aerodynamics, which can lead to challenges of increasing the compressor efficiency. Losses associated with secondary flow effects are well known to be the major limiting factor of improving the compressor performance. In this study, non-axisymmetric endwall contouring in a linear compressor cascade was generated through the optimization process. Combined with numerical simulation, wind tunnel tests on linear cascades with flat and contoured endwall were performed with various measurement techniques at the design and off-design conditions. The simulation results show that optimal endwall design can provide 3.08% reduction of the total pressure loss at the design condition. The reduction of pressure loss obtained is mainly below 24%span with the size of the high loss region being effectively reduced. At off-design condition, the numerical benefit of the endwall contouring is found less pronounced. The discrepancy is spotted between simulation and experiments. The experimental pressure loss reduction is mainly below 18% at ADP. And the pressure loss for the CEW increases greatly at offdesign condition in experiments. Flow patterns revealed by numerical simulations show that the separation on the blade surface is mitigated with focus point disappearing, and reverse flow on the endwall near the suction side corner is moved away from the blade surface. CFD analysis indicates that the altered pressure distribution on the endwall accelerates the flow at the suction side corner and moves the reverse flow core further downstream. The weakened interaction between the corner vortex and tornado-like vortex from the endwall near the suction side corner is the main control mechanism of the CEW. The performance improvement in the linear compressor is mainly gained from it.

Under the condition of large incidence angle in an axial compressor, corner separation will occur in corner region. When the blade loading increases, there may even be corner stall. This three dimensional complex flow structure is high intermittent and unsteady. In order to study the flow mechanisms, a hybrid RANS/LES turbulence model (SBES) was used to simulate the corner separation in the cascade. Firstly, the evolution of vortical structures in the corner separation region were analyzed. The boundary layer and the backflow in the corner separation region encounter, forming vortices with opposite rotation and developing downstream continuously. This cause the unsteady flow in corner separation. In addition, the shedding of detached eddies at the trailing edge is another sources of unsteadiness. Secondly, based on studying the turbulent characteristics, it can be seen that there is active turbulent fluctuation in the corner separation region. The turbulence is high anisotropic because the distribution and the values of Reynolds stresses in different directions are quite different. Thirdly, the mechanism of bimodal phenomenon was explained in this paper. Bimodal means there are two peaks in the velocity probability histogram, it is found in the region between separated flow and non-separated flow. This indicates that there are two non-periodic switching flow modes. Through the analysis of this paper, we find that it is related to the development of large scale coherent vortical structures.

In this paper, upstream unsteady passing wake effects on the rotor blade tip with film cooling have been numerically examined. The geometry and flow conditions of the first stage of GE-E 3 high-pressure turbine have been used to obtain the unsteady three-dimensional blade tip flow and heat transfer characteristics. The first stage of GE-E 3 high-pressure turbine has 46 guide vanes and 76 rotor blades. In the process of calculation, compromise the computational resources and accuracy to simplify the number of the guide vane and rotor blade to 38:76. Namely, each computational domain comprises of one guide vane and two rotor blades. The computational boundary conditions are consistent with the GE-E 3 annular cascade test conditions. In the case of 1% blade span tip clearance, through comparing the steady results, time-averaged and time-resolved unsteady results, the investigation of the unsteady passing wake effects on flat tip aerodynamic and aerothermal performance without film cooling holes (NC) and three cases with film cooling holes are conducted. To be specific, near the leading edge (C123), at the central area (C456) and near the trailing edge (C789). This paper emphasizes the variation of leakage flow and heat transfer coefficient at different unsteady instants. The results show that the time-averaged leakage flow is pretty similar to the steady results, but the increment of the leakage flow can rise to more than 8% at the maximum envelope. Moreover, the heat transfer coefficient discrepancy between steady results and time-averaged results is almost below 4%, but the dramatic growth of the instantaneous heat transfer coefficient along the pressure side is in excess of 20% due to the shift of the pressure spot.

Recent studies have demonstrated that the turbine blade with a squealer tip has less aerodynamic loss than that with a flat tip. However, there exists an area of high heat transfer coefficient (HTC) in the frontal region of the cavity floor. The present study focuses on the effect of squealer rim slot structure on turbine blade aerothermal performance. An experiment of transient thermal measurement is carried out for a non-slot squealer tip to validate the numerical setup and the choice of turbulence model. ANSYS CFX is utilized to conduct the numerical simulations of cases with and without slot structure and they are both validated by the experimental data. Through the experimental and CFD results, it can be found that the case with a slot on the pressure side of the squealer rim has smaller area of high HTC region and the aerodynamic efficiency is raised. Moreover, the effect of different sizes of the slot is investigated numerically and it also has a great impact on the heat transfer and aerodynamics. The quantity of the flow passing through the slot can largely influence the flow field of the tip gap and the cavity. Thus, the heat transfer of these areas and the aerodynamic efficiency of the blade vary from the slot size.

Numerical simulations with the steady 3D RANS were performed on a 4-stage low speed research compressor (LSRC) with two typical configurations (a shrouded and a redesigned cantilevered stator of the third stator). The shrouded stator (SS) with 0.67% labyrinth seal clearance of the blade height is the prototype, and the cantilevered stator (CS) with 1.2% hub clearance of the blade height is the redesigned cantilevered stator. The fourth rotor that follows after the cantilevered stator was redesigned (RE) according to blade load and inlet flow angle changed based on the redesigned cantilevered stator. The overall performance of the 4-stage LSRC and the distribution of aerodynamic parameters along the blade height were compared between the prototype and the redesigned third stator. Flow characteristics of the third stator and fourth rotor were analyzed in detail. The results indicate that the flow characteristics below the 35% blade height are very different between the prototype and the redesigned due to the effect of leakage flow from seal cavity and hub gap, respectively. The stall margin of CS is 57% higher than SS. The efficiency of CS at the design point is 0.82% higher than SS. Through the redesigned process of R4, the stall margin of RE is 45% higher than CS and the efficiency of R4 is 0.6% higher on average over the entire operating range.

The focus of the present study is to find a solution to the manufacturing tolerances of rotor stagger angle in the high pressure compressor. The manufacturing processes of compressor blades may lead the stagger angle to be out of tolerance, which would reduce the performance of the high-pressure compressor. Meanwhile, the assembly may cause the nonuniform distribution of the variable blades. To investigate the influence of stagger angle variability and its nonuniform distribution in multiple passages, amount of 3D numerical simulations of a high pressure compressor rear-stage rotor were conducted in the present paper. The Gaussian Probabilistic Density Function was used to obtain the sampled blades with variable stagger angles, and then the blades were randomly allocated in a single passage to a half period, respectively. Numerical results show that with the increase of passage number, the total performance varies more stably. In order to get an appropriate tolerance range of the stagger angle to control the deterioration of the total performance, different ranges were tried, and [−1.5°, +1.5°] was determined as the final range to do further analyses. The different stagger angle distribution along the circumferential direction of a specific group of several sampled blades led to the different performance variations. When the stagger angle distribution approximately satisfied the sinusoidal function during assembly, the rotor performance was better than other distributions. Hence, the sinusoidal function during assembly is an effective solution to the manufacturing tolerances of rotor stagger angle.

The compressibility of flow field has an important effect on flow stability. However, when the compressibility is considered, the effect of Mach number is often considered while the effect of heat transfer is always neglected in the existing flow stability studies. Linear stability analysis tools based on compressible Orr-Sommerfeld (O-S) equations and linearized Navier-Stokes equations in cylindrical coordinate system are established in this paper. These equations are numerically solved by using Chebyshev spectral collocation method and pseudo-modes are eliminated. Linear stability analysis of pipe flow with heat transfer whose average flow field is obtained by CFD simulation is carried out. The results show that for spatial modes, the heating effect of the wall makes pipe flow more unstable, while cooling effect of the wall makes pipe flow more stable. For global modes of pipe flow, the frequency of global mode decreases when the wall cools the flow and the decrease of mean temperature of pipe flow leads to the improvement of global mode stability.

Humpback whale flippers’ scalloped tubercles on the leading edge are thought to enhance the whale’s underwater maneuverability. Inspired by the flippers, leading edge tubercles are applied in a low speed annular compressor cascade as a type of passive flow control techniques in this paper. A numerical study is performed to investigate the influence of tubercles on the aerodynamic losses and corner separation in the low speed cascades. Different low speed cascades based on a CDA airfoil profile are built with several hub-tip ratios and aspect ratios. Steady RANS simulations are carried out for these cascades with and without leading edge tubercles. The aerodynamic performance and corner separation features are subsequently investigated in these cascades. The influence of tubercles under the variation of hub-tip ratio and aspect ratio is understood and concluded from the comparison of the performance attained by different cascades. Flow visualizations at a post-stall incidence angle show that the interaction between the tubercle-induced streamwise vortices and corner separation vortices plays a crucial role in attenuating the corner separation and reducing losses. By combining the performance analysis and flow visualizations, this paper discusses the mechanism of leading edge tubercles in a low speed annular compressor cascade with different hub-tip ratios and aspect ratios.

Numerical simulations with the steady 3D RANS were performed on the rear stage of a modern high pressure compressor. The labyrinth seal cavity model of the shrouded stator was simplified according to the actual stator structure, which the seal cavity gap is 1% of blade height. Several typical configurations (shrouded stator, idealized stator and cantilevered stators) were designed and carried out, and cantilevered stators contained no gap, small gap (CS1%), design gap (CS2.5%) and large gap (CS4%/CS5%). The results indicate due to the effect of leakage flow from 1% span seal cavity gap, the total pressure loss of SS is larger than IS, while IS instead of SS in the process of the compressor design, the stall margin will be enlarged nearly 6% numerically. At the design point, when the hub gap is 3.5% span clearance CS has the same loss with IS, and when the hub gap is 4.5% span clearance CS has almost the same loss with SS. Among all operation range, the total pressure loss of S1 increases with the increase of the hub clearance. When the hub gap is 0 (CS0), there is no leakage flow and the loss is the least. At the design point, comparing with SS, the total pressure loss coefficient of CS0 decreases 18.34%, CS2.5% decreases 8.46% and IS decreases 6.45%. It means if the cantilevered stator with 2.5% span hub clearance were adopted in the HPC, the performance would be better than the shrouded stator. However, because of the matching condition, the rotor that follows after cantilevered stator should be redesigned according to blade loading and inlet flow angle changed. The performance of cantilevered stator is impacted of various hub clearance, the loss below 25% span increases significantly with hub clearance, the maximum value of outlet flow angle deviation is 2.3 degree. The stator hub peak loading is shifted upstream toward the leading edge when hub clearance size is increased. The total pressure loss coefficient and pressure coefficient at different axial position had the function relation. When the hub clearance increases, the position of double leakage flow start backwards, in the rear part of stator the secondary flow becomes stronger leading to more mixing loss and lower total pressure.

Slot-type casing treatment generally has a great potential of enhancing the operating range for tip-critical compressor rotors, however, with remarkable efficiency drop. Part I of this two-part paper was committed to develop a slot configuration with desired stall margin improvement and minimized efficiency loss. Steady simulation was carried out in a 1.5 transonic axial compressor stage at part design rotating speed. At this rotating speed this compressor stage operated at a subsonic condition and showed a rather narrow operating range, which needed to be improved badly. Flow fields analysis at peak efficiency and near stall point showed that the development of tip leakage vortex and resulting blockage near casing resulted in numerical stall. Three kinds of skewed slots with same rotor exposure and casing porosity were designed according to the tip flow field and some empirical strategies. Among three configurations, arc-curved skewed slot showed minimum peak efficiency drop with considerable stall margin improvement. Then rotor exposure and casing porosity were varied based on the original arc-curved skewed slot, with a special interest in detecting their impact on the compressor stability and overall efficiency. Result showed that smaller rotor exposure and casing porosity leaded to less efficiency drop. But meanwhile, effectiveness of improving compressor stability was weakened. The relation between efficiency drop and stall margin improvement fell on a smooth continuous curve throughout all slots configurations, indicating that the detrimental effect of casing treatment on compressor was inevitable. Flow analysis was carried out for cases of smooth casing and three arc-curved configurations at smooth casing near stall condition. The strength of suction/injection, tip leakage flow behavior and removal of blockage near casing were detailed examined. Larger rotor tip exposure and slots number contributed to stronger injection flow. The loss generated within the mixing process of injection flow with main flow and leakage flow is the largest source of entropy increase. Further loss mechanisms were interpreted at eight axial cuts, which were taken through the blade row and slots to show the increase in entropy near tip region. Entropy distributions manifested that loss generations with smooth casing were primarily ascribed to low-momentum tip leakage flow/vortex and suction surface separation at leading edge. CU0 slot, the arc-curved slots with 50% rotor tip exposure, was capable of suppressing the suction surface separation loss. Meanwhile, accelerated tip leakage flow brought about additional loss near casing and pressure surface. Upstream high entropy flow would be absorbed into the rear portion of slots repeatedly, resulting in further loss.

Slot-type casing treatment generally has a great potential of enhancing the operating range for tip-critical compressor rotors, however, with remarkable efficiency drop. In the first part of this two-part words, several configurations of slot casing treatment were tested in a 1.5 transonic compressor stage by steady simulations. One kind of arc-curve skewed slot contributed to considerable stall margin improvement with minimum efficiency loss. However, interaction between main passage and casing treatment was inherently unsteady. Steady simulation was inadequate to provide accurate compressor performance prediction and precise flow field details. Thus, this part was aimed at clarifying the differences between steady and unsteady simulations. The unsteady interaction process between main passage flow and slots were also detailed interpreted. Unsteady simulation was conducted by applying sliding interface between rotor passage and arc-curved skewed slots. Firstly, differences of compressor performance were examined between steady and unsteady methods. Results showed that steady simulation underestimated stall margin improvement and efficiency drop by casing treatment. Then analysis on aerodynamic parameters and specific flow fields were carried out at smooth casing peak efficiency and casing treatment near stall conditions. Unsteady simulation provided more than 50% larger mass flow rate entering or exiting slots opening surfaces at both operating conditions. It revealed that in unsteady simulation, casing treatment contributed to stronger suction/injection process, which promoted tip flow fields more effectively than steady simulation. Axial velocity deficit at rotor outlet was refilled by slots more effectively in unsteady simulation. In steady result, a large low momentum blockage existed inside rotor passage near tip region and prevented flow from entering the passage at near stall condition. While in unsteady simulation at the same condition, incoming flow was still able to travel across rotor passage in a high velocity. Further, instantaneous flow fields near tip region and inside the slots were particularly examined during a rotor blade passing period to elaborate the unsteady flow interaction. The mid-pitch surface of a representative slot was selected to represent the re-circulation procedure inside slots. Unsteady flow fields and spectrum analysis manifested that tip flow field was dominated by slots passing, while re-circulation process inside slots was dominated by blade passing. Low pressure region inside the blade passage facilitated the injection process. Circulation inside slots lagged behind the pressure variations beneath slots. When the slot was striding over the blade tip, intense injection didn’t emerge immediately beneath slots’ front portion. Until the high pressure region moved away from the slot opening surface, fluids inside the slots started to inject into the main flow in high speed.

Humpback whales possess bumpy tubercles on the leading edge of their flippers. Due to these leading edge tubercles, the whales are able to perform complex underwater maneuvers agilely. Inspired by the flippers, this paper applies sinusoidal-like tubercles to the leading edge of the blade in an annular compressor cascade, and presents a numerical investigation to explore the effects of tubercles with the aim of controlling the corner separation and reducing losses. A preliminary study by steady 3D RANS simulations is performed. The aerodynamic performance and the behavior of the corner separation are investigated in the baseline compressor cascade. Subsequently, cascades with leading edge tubercles are numerically simulated. A crucial geometry parameter of the tubercles, wavelength, is varied to obtain different configurations. The influence of the parameter is concluded from the comparison of the performance attained by these configurations. Also, several configurations, which are typical in loss characteristics, are selected for further DES simulations so as to obtain more flow details, especially at the separation region. Flow visualizations show that leading edge tubercles could induce the formation of counter-rotating streamwise vortices. The interaction between the streamwise vortices and corner separation is emphatically investigated. By analysis of all the results obtained, this paper tries to figure out the mechanism of leading edge tubercles in loss reduction and separation delay in an annular compressor cascade.

In this paper, we present a study on the aerodynamic and heat transfer performance optimization of one stage air-cooled turbine. In order to tackle the high cost associated with both CFD simulations and the large number of design variables involved, 28 parameters are chosen as the optimization variables including the blade profiles and the direction of film cooling holes. Multiple Pareto Fronts are constructed using the optimization algorithm NSGA-II. A new objective function is proposed to reflect the worst heat transfer condition of the blade except the objective function which is used to reflect the aerodynamic performance. Then three solutions are chosen according to design requirements to compare with the original scheme. The aerodynamic efficiency is improved by 1.5%, and the second objective function is improved by 110.48%. The maximum temperature and the area of the high temperature region is decreased significantly. The result shows that the shock loss is decreased in the stator passage, and the increase of the reaction degree increased the rotor load and changed the separation structure. All the reasons mentioned above can improve the aerodynamic efficiency. The heat transfer effect improvement mainly due to the direction change of the film holes. The direction change makes the coolant distribution more reasonable than the original scheme, so the coolant could cover the blade directly. Meanwhile, the double-jet ejection plays an important role in the improvement of film cooling.

This paper presents a numerical investigation to explore the effects of vortex generators on a high speed compressor cascade. Secondary flow effects like the corner separation vortex have an influence on the performance of a compressor cascade such as leading to increased losses. In order to control the corner separation vortex and reduce losses, an extensive study of vortex generators applied to a compressor cascade is conducted. A preliminary study by steady 3D RANS simulations is performed using the Spalart-Allmaras turbulence model. The aerodynamic performance as well as the behavior of the corner separation vortex is investigated in the compressor cascade without vortex generators. Then, a vortex generator is added to the cascade, which is numerically simulated. Various configurations are considered, which are decided by the height and installation angle of the vortex generator. Comparison of the performance attained by these configurations results in an optimum scheme that has minimum losses. Furthermore, unsteady 3D DES simulations are performed with the optimum configuration. This method that predicts the flow field more precisely could help verify the accuracy of the RANS results. Finally, by analyzing all the resulting aerodynamic performance and numerical flow phenomena, the mechanism of vortex-vortex interaction is presented and discussed, which could be a criterion to reduce the corner separation flow and enhance the performance of axial compressors.

In aeroengine, an s-shaped annular duct is used to connect the high pressure compressor spool to the upstream low pressure compressor. The intermediate casing duct is required because the compressors are designed with different mean radii. Shortening the length of intermediate casing duct is of important significance for improving the performance of aeroengines. Meanwhile, the struts guiding the flow are installed in the intermediate duct, the inlet flow of the duct is not even, and outlet flow requires sufficient uniformity to ensure good working conditions for the high pressure compressor. So the design and optimization of flow path of intermediate casing duct including struts have very important practical significance. This study builds up a system of numerical optimization 3D model based on genetic algorithm and Kriging agent model on the Matlab platform, which automatically couples pre-processing, solver and post-processing of Numeca CFD software. This tool optimizes the intermediate casing duct including struts and a best flow path profile is obtained under the given constrained conditions. The optimization method is validated through experimental investigation. Four kinds of intermediate casing ducts are test, including the original model, optimization model, as well as two models based on optimization model, respectively. The chord of the struts and the length of flow path are changed for the last two models. The internal flow characteristics and aerodynamic performance of four cases above are measured, and the effects of the geometry structure of flow paths and struts to the aerodynamic performance of the ducts are evaluated, and the results may bring some modifications to the optimization method to make it more perfect. Through the experimental and numerical study, two key factors affecting the design of intermediate duct are brought about, which is the midway offset ratio and the maximum to the minimum area ratio of flow path along streamwise including struts. For the given intermediate casing duct in this study, the midway offset ratio is about 0.41, and the maximum to the minimum area ratio of flow path along streamwise including struts should be small.