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Blade slotting technology is an effective measure to improve the flow structure on the suction surface of a blade and enhance the performance of turbomachinery. To investigate the impact of various slot sizes on the flow field of a single-stage transonic compressor rotor, seven kinds of slot schemes were designed and calculated by numerical simulations. The results show that the above slotting schemes significantly enhance the stability margin of the compressor. In particular, the slotting scheme H9W3 increases the surge margin by 60.9% and slightly reduces peak efficiency by 0.3%, with an almost identical maximum pressure ratio. Slotting promotes high-energy fluid to generate jets from the slot located at the exit of the suction side, effectively controlling blade surface flow separation and reducing channel blockage. Square slots are more effective than elongated slots for controlling separation when using differently shaped slots with equal areas. Increasing slot area gradually decreases outlet total pressure at a constant aspect ratio. A slight increase in the overall blade load causes a backward shift in the front portion load. [ABSTRACT FROM AUTHOR]

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A numerical investigation is conducted to explore the potential of combined slot suction in controlling the shock wave and flow separation in a transonic compressor cascade. The three slots in the combined scheme are arranged in different directions on the suction surface and endwall. The locations are determined by the shock wave and separation point in the baseline cascade. Based on these locations, two combined schemes and three single schemes are provided to explore the control mechanism. For each suction scheme, five bleed mass ratios are examined at the same inlet Mach number. The results suggested that the cascade throughflow loss could be decreased by three single schemes. However, the cascade performance is improved slightly or even deteriorated when the losses generated by the suction are considered. The key reasons are the local effects of single scheme and the opposite trend between corner stall and the suction surface separation. Both corner stall and the suction surface separation are eliminated by the combined scheme with two slots, and the maximum reduction in throughflow losses is 75%. The new corner separation evolved from the horseshoe vortex limited the performance of combined scheme. When the bleed mass ratio excided 5%, the combined scheme with three slots is better than the two slots scheme in loss control. The reason is the improved endwall boundary layer and the eliminated new corner separation. By inducing the horseshoe vortex into the slot, the interaction of the two endwall slots eliminates the new corner separation. [ABSTRACT FROM AUTHOR]

Blade slotting technology is an effective measure to improve the flow structure on the suction surface of a blade and enhance the performance of turbomachinery. To investigate the impact of various slot sizes on the flow field of a single-stage transonic compressor rotor, seven kinds of slot schemes were designed and calculated by numerical simulations. The results show that the above slotting schemes significantly enhance the stability margin of the compressor. In particular, the slotting scheme H9W3 increases the surge margin by 60.9% and slightly reduces peak efficiency by 0.3%, with an almost identical maximum pressure ratio. Slotting promotes high-energy fluid to generate jets from the slot located at the exit of the suction side, effectively controlling blade surface flow separation and reducing channel blockage. Square slots are more effective than elongated slots for controlling separation when using differently shaped slots with equal areas. Increasing slot area gradually decreases outlet total pressure at a constant aspect ratio. A slight increase in the overall blade load causes a backward shift in the front portion load.

The aim of the paper is to explore the influence of the reversed blade-angle slot casing treatment (RBSCT) and its axial deflected angle (ADA) on the compressor performance and stability, and to reveal the mechanism that the change in ADA of the RBSCT influences the effect to broaden the compressor stable working range. The NASA Rotor 35 is used as the object of the investigation, and four RBSCTs with ADA of −15°, −30°, −45° and −60° are designed and investigated by unsteady numerical simulation. The results show that as the absolute value of the axial deflected angle increases, the capacity to improve the compressor stability of the RBSCT increases and then decreases. The unsteadiness of the injection and suction flows formed by the reversed blade angle slot plays an important role in the removal of the low-velocity zone. When ADA is −30°, the unsteadiness amplitude of the injection and suction flows is significantly higher than those of the other three. Consequently, the RBSCT with −30° ADA obtains the maximum stall margin improvement of 17.41% and the maximum design point efficiency improvement of 1.06% among the four RBSCTs. [ABSTRACT FROM AUTHOR]

One of the most commonly used techniques in aerospace engineering is the RANS (Reynolds average Navier–Stokes) approach for calculating the transonic compressor flow field, where the accuracy of the computation is significantly affected by the turbulence model used. In this work, we use SA, SST, k-ɛ, and the PAFV turbulence model developed based on the side-biased mean fluctuations velocity and the mean strain rate tensor to numerically simulate the transonic compressor NASA Rotor 67 to evaluate the accuracy of turbulence modeling in numerical calculations of transonic compressors. The simulation results demonstrate that the four turbulence models are generally superior in the numerical computation of NASA Rotor 67, which essentially satisfies the requirements of the accuracy of engineering calculations; by comparing and analyzing the ability of the four turbulence models to predict the aerodynamic performance of transonic compressors and to capture the details of the flow inside the rotor. The errors of the Rotor 67 clogging flow rate calculated by the SA, SST, k-ɛ, and PAFV turbulence models with the experimental data are 0.9%, 0.8%, 0.7%, and 0.6%, respectively. The errors of the calculated peak efficiencies are 2.2%, 1.6%, 0.9%, and 4.9%. The SA and SST turbulence models were developed for the computational characteristics of the aerospace industry. Their computational stability is better and their outputs for Rotor 67 are comparable. The k-ɛ turbulence model calculates the pressure ratio and efficiency that are closest to the experimental data, but the computation of its details of the flow field near the wall surface is not ideal because the k-ɛ turbulence model cannot accurately capture the flow characteristics of the region of high shear stresses. The PAFV turbulence model has a better prediction of complex phenomena such as rotor internal shock wave location, shock–boundary layer interaction, etc., due to the use of a turbulent velocity scale in vector form, but the calculated rotor efficiency is small. [ABSTRACT FROM AUTHOR]

Performance of a NASA 37 transonic compressor stage with a symmetric Airfoil Vortex Generator (AVG) positioned in the upstream of rotor is investigated through a numerical simulation. Steady state flow simulations were performed with k-ω SST turbulence model in ANSYS CFX flow solver. Grid independence study for the baseline compressor was performed besides CFD predicted performance characteristics were validated against available experimental data. The results of parametric study shows that AVG is able to improve compressor stall margin with a penalty on the stage efficiency. Incorporation of an AVG on the casing surface upstream of rotor reduces its specific work capacity, incurs flow losses and decreases stage peak efficiency by 3.34%. However few AVG configurations have negligible reduction, nearly 0.27-0.4%, in stage efficiency compared to baseline case. At near stall operating point, an AVG affects both rotor and stator flow field. An AVG under loads rotor tip region and decreases tip leakage mass flow rate, leading to the reduction of flow blockage. The flow swallowing capacity of the stator passage increases by diffusing the flow considerably to a low velocity. Consequently, AVG is able to increase the compressor stage stall margin by 8.06%. [ABSTRACT FROM AUTHOR]

High-altitude ice crystal icing of aero engine is a serious threat to flight safety. Previous studies on ice crystal icing focused on the influence of air flow on ice crystals, but ice crystals also affect the engine flow field, which is often ignored in the research on the influence of ice crystals on engines. Numerical simulation based on Eulerian method is adopted to realize the two-way coupling between the compressor air flow and the particles in this study. The approach is demonstrated using the NASA compressor stage 35. The changes of compressor and particle parameters with different inlet total water content, relative humidity, and median volume diameter are calculated and analyzed, and the influence of ice crystal on the compressor performance is studied. The results show that the variation of relative humidity has a great influence on the particle temperature, median volume diameter, and wet bulb temperature. Median volume diameter has a great influence on the melt ratio. The variation of total water content has little effect on particle temperature, total water content, median volume diameter, and wet bulb temperature. The particle parameters are affected by the flow field of the compressor. The parameters show that the icing is easy to occur at the leading edge of NASA stage 35 stator. By contrast, the overall compressor characteristics, after ice particle injection, the total pressure ratio, and isentropic efficiency of the compressor are increased without considering ice crystal accretion, and the chocking boundary and stall boundary are not affected. [ABSTRACT FROM AUTHOR]

• Discussed the mechanism responsible for unsteady flows induced by non-axisymmetric clearance from the casing ovalization on aeroelastic stability in numerical methodology for the first time. • Established a one-way fluid-structure coupled (or partially coupled) method for analyzing turbomachinery flutter (in this manuscript), forced response and stall induced virabtion (future work) in the commercial software ANSYS. • Presented a detailed elucidation of the circumferential non-uniform flow details and flutter behaviors under a non-axisymmetric tip clearance configuration. Other crucial factors such as rotor tip clearance size, operating condition and nodal diameters are also considered. • Primarily revealed the isolated and coupled effects of leakage flow, shock waves and separated flow on aerodynamic characteristics and aerodynamic damping. The aerodynamic and flutter effects caused by the circumferential non-uniform tip clearances from the non-axisymmetric casing geometry, which are more representative of real operational conditions, should receive more attention. In this paper, sensitivity analyses on clearance and non-uniformity regarding rotor performance are conducted. The isentropic efficiency, total pressure ratio and stability margin of the rotor are reduced by 1.96%, 1.06% and 6.76% for an increase of tip clearance by 1% tip chord, respectively. The non-axisymmetric configuration reconfigures the interaction between leakage flow and mainstream in each rotor passage and alters the strength and influence range of the low-speed separation region generated by shock wave. The phase lag phenomenon observed in the aerodynamics of the non-axisymmetric layout also manifests in the aeroelastic effects but requires a larger phase shift at the maximum clearance sector, owing to the longer timescales for flow adjustment relative to the local tip gap considering the flow-induced vibration. Clearance sensitivity on flutter stability varies with the direction of the traveling wave: under forward traveling conditions, damping values initially decrease and then increase with increasing rotor tip clearance; while under backward traveling condition, blade damping linearly decreases with increasing rotor tip clearance. Additionally, variations in flow conditions and nodal diameters affect flutter stability by altering the position of shock waves and their work on the suction surface of the rotors. The mechanism responsible for unsteady flows induced by non-axisymmetric clearances from the casing ovalization on flutter stability is discussed for the first time in this paper, which can guide the industry to evaluate its impact. [ABSTRACT FROM AUTHOR]

One of the most commonly used techniques in aerospace engineering is the RANS (Reynolds average Navier–Stokes) approach for calculating the transonic compressor flow field, where the accuracy of the computation is significantly affected by the turbulence model used. In this work, we use SA, SST, k-ɛ, and the PAFV turbulence model developed based on the side-biased mean fluctuations velocity and the mean strain rate tensor to numerically simulate the transonic compressor NASA Rotor 67 to evaluate the accuracy of turbulence modeling in numerical calculations of transonic compressors. The simulation results demonstrate that the four turbulence models are generally superior in the numerical computation of NASA Rotor 67, which essentially satisfies the requirements of the accuracy of engineering calculations; by comparing and analyzing the ability of the four turbulence models to predict the aerodynamic performance of transonic compressors and to capture the details of the flow inside the rotor. The errors of the Rotor 67 clogging flow rate calculated by the SA, SST, k-ɛ, and PAFV turbulence models with the experimental data are 0.9%, 0.8%, 0.7%, and 0.6%, respectively. The errors of the calculated peak efficiencies are 2.2%, 1.6%, 0.9%, and 4.9%. The SA and SST turbulence models were developed for the computational characteristics of the aerospace industry. Their computational stability is better and their outputs for Rotor 67 are comparable. The k-ɛ turbulence model calculates the pressure ratio and efficiency that are closest to the experimental data, but the computation of its details of the flow field near the wall surface is not ideal because the k-ɛ turbulence model cannot accurately capture the flow characteristics of the region of high shear stresses. The PAFV turbulence model has a better prediction of complex phenomena such as rotor internal shock wave location, shock–boundary layer interaction, etc., due to the use of a turbulent velocity scale in vector form, but the calculated rotor efficiency is small.

Blade lean has been extensively used in axial compressor stators to control flow separations, but its influence mechanism on transonic compressor rotors remains to be revealed. The aim of this study is to numerically explore the influence of blade lean on the performance and shock wave/tip leakage flow interaction in a transonic compressor rotor. The effects of leaned pattern (positively lean and negatively lean), leaned angle and leaned height were studied. Results showed that, compared with baseline configuration, the efficiency and total pressure ratio of the entire constant rotating speed line of positively leaned rotor were both decreased. The absolute value of peak efficiency was reduced by as much as 4.34% at 20° lean angle, whereas the maximum reduction of peak total pressure ratio was 0.1 at 20° lean angle. The tip leakage flow streamlines of baseline transonic rotor can be divided into two parts, i.e., the primary vortex and secondary vortex which arises after the shock. Due to shock/tip leakage vortex interaction, the primary vortex enlarged and low-momentum region showed up after the shock; under near stall (NS) condition, tip leakage vortex breakdown occurred after interacting with shock. As positively leaned angle increased, the shock and the shock/tip leakage vortex interaction point moved upstream. In addition, the phenomenon of tip leakage vortex breakdown was enhanced. For negatively leaned rotors, as negatively leaned angle increased, the peak efficiency and total pressure ratio showed a tendency of first increasing and then decreasing. At 5° leaned angle, the peak efficiency was increased by 0.8% at most, and the maximum increment of total pressure ratio was 0.05 at 5° leaned angle. Besides, the loading of blade tip reduced and the loading moved toward trailing edge, resulting in the downstream movements of primary vortex, shock front and shock/tip leakage vortex interaction location. The results may help to improve the near tip flow field of transonic compressor rotor with leaned blade technology.

To achieve multi-objective aerodynamics design optimization for a 1.5-stage transonic compressor, a design platform incorporating blade parameterization methods, a BPNN surrogate model, and the NSGA-II optimization method was developed. The stagger angle distribution of three blade rows was selected as the optimization variable, with isentropic efficiency at the new design condition and stall margin set as the goal functions. Results demonstrated that, without altering the blade profile shape and endwall contour, the flow rate at design condition increased by 7.1%, stall margin increased by 1.8%, isentropic efficiency decreased by 0.0087, and total pressure ratio experienced a slight increase. The flow field at different conditions before and after optimization was compared and analyzed. The analysis indicated that the tangential velocity of rotor outlet becomes the determining factor for the compressor’s work capacity. The relative Mach number at the rotor inlet emerged as the key parameter affecting shock wave intensity and shock wave/boundary layer interaction, which directly influenced the efficiency of the rotor passage. At near stall condition, the stator vane root’s stagger angle is crucial for the compressor’s performance.

In the boundary layer ingesting propulsion system, the compressor suffers from a non-uniform flow field. The compressor operating with distorted inflow continuously results in the loss of aerodynamic performance and stability margin. In this paper, three non-axisymmetric configurations are described for the stator of a transonic compressor to match the non-uniform flow field. The flow fields with distorted inflow at near stall condition are obtained and analyzed, the effects of the prototype stator and the three non-axisymmetric stators on aerodynamic performance are compared in detail. Results show that the non-axisymmetric stator schemes can effectively improve the stability margin of the transonic compressor and the maximum stability margin is relatively increased by 22.3% in all the three non-axisymmetric stators. The non-axisymmetric stator design is effective on decreasing the aerodynamic losses and improving the performance of the compressor operating with distorted inflow. Overall, the results show that in the design of the non-axisymmetric stator, the adoption of a curved-twisted blade and the increase of cascade solidity have the potential to reduce loss sources caused by distorted inflow. [ABSTRACT FROM AUTHOR]

To reduce the negative impacts of stall and surge on compressor performance, a novel combined casing treatment (CCT) structure with axial skewed slots and injection groove is proposed in this paper. The aerodynamic performance, as well as the mechanisms of loss generation, of a transonic axial compressor with NASA Rotor 67 are investigated numerically. The simulation results indicate that, compared with individual casing treatment method, the CCT works effectively with regard to operation performance. The stall margin (SM) is increased by 14.7% with 1.12% decrease in the peak efficiency. The interaction of axial skewed slots and injection groove can be explained by the enhancement of exchange flow in slots and axial motion of fluid. As a result, the leakage flow near the blade tip is eliminated and the flow separation is further suppressed. What is more, an analysis of entropy generation is also conducted. The results reveal that the effect of CCT on loss reduction mainly concentrates in the tip part of the blade, with the loss decrease about 14.46% compared with the original rotor. The best control effect can be expected by appropriate match between geometrical parameters of axial skewed slots and mass flow rate of injection from the parameter analysis. [ABSTRACT FROM AUTHOR]

To deeply understand the hub leakage flow and its influence on the aerodynamic performance and flow behaviors of a small-scale transonic axial compressor, variations of the performance and the flow field of the compressor with different hub clearance sizes and clearance shapes were numerically analyzed. The results indicate that the hub clearance size has remarkable impacts on the overall performance of the compressor. With the increase of the hub clearance, the intensity of the hub leakage flow increases, resulting in more intense flow blockage near the stator hub, which reduces the compressor efficiency. However, the flow field near the blade mid-span is modified due to the more convergent flow as the reduced effective flow area caused by the passage blockage, and the flow separation range is narrowed, thus the flow stability of the compressor is enhanced. On this basis, two kinds of non-uniform clearance cases of expanding clearance and shrinking clearance with the same circumferential leakage area as the design clearance were investigated. The occurrence position of the double leakage flow which is closely connected with the flow loss and blockage is shifted backward by the expanding clearance, the flow capacity near the stator hub is enhanced, and the unsteady fluctuation intensity of the flow field is attenuated but fluctuation frequency remains. Similarly, the modification of the stator blade root flow field may result in the reduction of stall margin. The effect of the shrinking clearance on compressor performance is opposite to that of the expanding clearance, which reduces the peak efficiency and delays the stall inception. [ABSTRACT FROM AUTHOR]

The current study is aimed at understanding the effect of rotating tip clearance asymmetry on the operability and performance of a transonic compressor. Another objective of this investigation is to determine the influence of tip injection on reducing the detrimental effects of clearance asymmetry. Three dimensional unsteady Reynolds-averaged Navier–stokes simulations have been performed from choke to stall for different arrangements of non-uniform blade heights in a transonic fan. Furthermore, numerical computations have been conducted with endwall injection of air. The numerical results have been validated against experimental data. Results show that having the same mean tip clearance, the asymmetric compressor is less stable than the axisymmetric configuration. However, the peak pressure rise is found to be almost linearly correlated to the mean tip clearance for both the axisymmetric and asymmetric compressors. It is found that tip injection can desensitize the compressor to the tip clearance asymmetry. Results further reveal that tip clearance asymmetry does not change the compressor path to instability. However, endwall injection is found to be able to change the compressor stalling mode. Investigations concerning rotating non-uniformity (caused by non-uniform blade heights) are very few in open literature. The obtained results can assist in predicting the effect of rotating tip clearance asymmetry on the stability and performance of high-speed compressor rotors. Furthermore, the results uncover how tip injection can desensitize the compressor stability and affect its path into instability, which is one of the most important questions in the turbomachinery world. [ABSTRACT FROM AUTHOR]

Blade lean has been extensively used in axial compressor stators to control flow separations, but its influence mechanism on transonic compressor rotors remains to be revealed. The aim of this study is to numerically explore the influence of blade lean on the performance and shock wave/tip leakage flow interaction in a transonic compressor rotor. The effects of leaned pattern (positively lean and negatively lean), leaned angle and leaned height were studied. Results showed that, compared with baseline configuration, the efficiency and total pressure ratio of the entire constant rotating speed line of positively leaned rotor were both decreased. The absolute value of peak efficiency was reduced by as much as 4.34% at 20° lean angle, whereas the maximum reduction of peak total pressure ratio was 0.1 at 20° lean angle. The tip leakage flow streamlines of baseline transonic rotor can be divided into two parts, i.e., the primary vortex and secondary vortex which arises after the shock. Due to shock/tip leakage vortex interaction, the primary vortex enlarged and low-momentum region showed up after the shock; under near stall (NS) condition, tip leakage vortex breakdown occurred after interacting with shock. As positively leaned angle increased, the shock and the shock/tip leakage vortex interaction point moved upstream. In addition, the phenomenon of tip leakage vortex breakdown was enhanced. For negatively leaned rotors, as negatively leaned angle increased, the peak efficiency and total pressure ratio showed a tendency of first increasing and then decreasing. At 5° leaned angle, the peak efficiency was increased by 0.8% at most, and the maximum increment of total pressure ratio was 0.05 at 5° leaned angle. Besides, the loading of blade tip reduced and the loading moved toward trailing edge, resulting in the downstream movements of primary vortex, shock front and shock/tip leakage vortex interaction location. The results may help to improve the near tip flow field of transonic compressor rotor with leaned blade technology. [ABSTRACT FROM AUTHOR]

Since the transition from rotating stall to surge in a transonic compressor at high speed is very quick, quite often there is no time to take measures to prevent the surge. Therefore, it is desired to find any rotating stall precursors, of which the occurrence can offer sufficient time for stall or surge prevention. In this study, a series of unsteady flow analyses were performed on a transonic compressor under operating conditions before rotating stall with unsteady results scrutinized to find rotating stall precursors. Particular attention is paid to the spatial modes and time modes of static pressure near the casing and around the blade leading and trailing edges. The results show that the characteristics of the precursor in both spatial and time domains can be used as rotating stall warnings. [ABSTRACT FROM AUTHOR]

The rotor blade height with low hub-tip ratio is relatively longer, and the aerodynamic parameters change drastically from hub to tip. Especially the organization of flow field at hub becomes more difficult. This paper takes a transonic 1.5-stage axial compressor with low hub-tip ratio as the research object. The influence of four types of rotor hub contouring on the performance of transonic rotor and stage is explored through numerical simulation. The three-dimensional numerical simulation results show that different hub contourings have obvious influence on the flow field of transonic compressor rotor and stage, thus affecting the compressor performance. The detailed comparison is conducted at the rotor peak efficiency point for each hub contouring. Compared with the linear hub contouring, the concave hub contouring can improve the flow capacity, improve the rotor working capacity, and increase the flow rate. The flow field near blade root and efficiency of transonic rotor is improved. The convex hub contouring will reduce the mass flow rate, pressure ratio and efficiency of the transonic rotor. Full consideration should be given to the influence of stator flow field by hub contouring. [ABSTRACT FROM AUTHOR]

To investigate the effect of hub clearance of cantilever stator on the aerodynamic performance and the flow field of the transonic axial-flow compressor, the performance of single-stage compressors with the shrouded stator and cantilever stator was studied numerically. It is found that the hub corner separation on the stator blade suction surface (SS) was modified by introducing the hub leakage flow. The separation vortex on the SS of the stator blade root at about 10% axial chord length caused by the interaction of the shock wave and boundary layer was also controlled. Compared with the tip clearance size of the rotor blade, the stator hub clearance size (HCS) has a much less effect on the overall aerodynamic performance of the compressor, and there is no obvious effect on the flow field in the upstream blade row. With the increase of HCS, the leakage loss and the blockage degree in the flow field near the stator hub are increased and further make the adiabatic efficiency and the total pressure ratio of the compressor gradually decrease. Meanwhile, the stall margin of the compressor was changed slightly, but the response of the stall margin to the change of the HCS is nonlinear and insensitive. The stator hub leakage flow (HLF) can not only change the flow field near the hub but also redistribute the flow law within the range of the entire blade span. It will contribute to further understand the mechanism of the HLF and provide supports for the design of the cantilever stator of transonic compressors. [ABSTRACT FROM AUTHOR]

This study proposes a 1D meanline program for the modeling of modern transonic axial multistage compressors. In this method, an improved blockage factor model is proposed. Work-done factor that varies with the compressor performance conditions is added in this program, and at the same time a notional blockage factor is kept. The coefficient of deviation angle model is tuned according to experimental data. In addition, two surge methods that originated from different sources are chosen to add in and compare with the new method called mass flow separation method. The salient issues presented here deal first with the construction of the compressor program. Three well-documented National Aerodynamics and Space Administration (NASA) axial transonic compressors are calculated, and the speedlines and aerodynamic parameters are compared with the experimental data to verify the reliability and robustness of the proposed method. Results show that consistent agreement can be obtained with such a performance prediction program. It was also apparent that the two common methods of surge prediction, which rely upon either stage or overall characteristic gradients, gave less agreement than the method called mass flow separation method. [ABSTRACT FROM AUTHOR]

The paper presents a design of a small-scale centrifugal compressor to match the performance of up to 20% larger diameter compressor in order to meet the demands for a higher mass flow rate and wide operating range for turbocharging applications. This development considers the 44 mm diameter impeller design involving transonic blading and state-of-the-art blade features. In the course of design, the investigation on the inducer blade angle β 1 s and blade inducer-to-outlet radius ratio has shown to have a dominant influence on the magnitude of inlet relative Mach number. Diffuser width and volute A/R ratio are also observed to be parameters affecting the overall performance of the compressor stage. Based on the diffuser parametric study, the most efficient diffuser width equal to 70% of the blade exducer width is observed. The performance of the present compressor stage is analyzed with computational fluid dynamics simulations and experimental tests. The comparison of the predicted and measured results of the 44 mm compressor stage shows a good agreement for overall performance. Besides, the 44 mm compressor stage having the most efficient diffuser width and enlarged volute A/R ratio shows a good overlap of performance with approximately 20% larger diameter (52 mm) compressor stage performance. This development demonstrates that the impeller with state-of-the-art design features are likely to contribute to enhancement of the compressor performance. [ABSTRACT FROM AUTHOR]

Non-synchronous vibration (NSV) problems in axial compressors are challenging and frequently result in high vibration stress or structural failure within the frequency lock-in region. A deeper understanding of the lock-in phenomenon in NSV is necessary to improve blade reliability and safety. In this study, an experiment was conducted on a multistage transonic axial compressor to obtain frequency lock-in characteristics. The enforced motion of the flexible blade simulation method was used to study the frequency lock-in mechanism between blade vibration and tip unstable flow. The results showed that the aerodynamic frequency characteristic changed from a broadband spectrum to a distinct frequency peak as the NSV onset. The nonlinear vibration pattern could maintain the limited cycle oscillation for 26 s. The vibration stress of the frequency locked-in case was approximately 19 times that of the frequency unlocked one. The pressure fluctuations generated by the large-scale radial separation vortex structures were stronger than those of the tip leakage flow, and the third-order harmonic frequency of the circumferential instability flow was close to the first-order bending model natural frequency. The periodic shedding and reattachment process of flow separation provided the initial non-synchronous aerodynamic excitation source. The frequency lock-in phenomenon occurred at a maximum vibration amplitude of 2 % and 3 % tip chord length. The tip unstable flow characteristics were consistent and entirely dominated by blade vibration. [ABSTRACT FROM AUTHOR]

To reduce the negative impacts of stall and surge on compressor performance, a novel combined casing treatment (CCT) structure with axial skewed slots and injection groove is proposed in this paper. The aerodynamic performance, as well as the mechanisms of loss generation, of a transonic axial compressor with NASA Rotor 67 are investigated numerically. The simulation results indicate that, compared with individual casing treatment method, the CCT works effectively with regard to operation performance. The stall margin (SM) is increased by 14.7% with 1.12% decrease in the peak efficiency. The interaction of axial skewed slots and injection groove can be explained by the enhancement of exchange flow in slots and axial motion of fluid. As a result, the leakage flow near the blade tip is eliminated and the flow separation is further suppressed. What is more, an analysis of entropy generation is also conducted. The results reveal that the effect of CCT on loss reduction mainly concentrates in the tip part of the blade, with the loss decrease about 14.46% compared with the original rotor. The best control effect can be expected by appropriate match between geometrical parameters of axial skewed slots and mass flow rate of injection from the parameter analysis.

A three-dimensional viscous inverse design method is improved and extended to multirow blades environment. The inverse method takes load distribution as optimization objective and is implemented into the time-marching finite-volume Reynolds-averaged Navier–Stokes solver. The camber line of rotor blade is updated by virtual displacement, which is calculated by characteristic compatibility relations according to the difference between target and actual load so as to control the location and intensity of shock wave, and realize the optimization of flow structure and reduction flow separation. The inlet and outlet geometry angles of stator blade are adjusted in real time according to the inlet and outlet flow angles. Thus, it is computationally ensured that the blade row interactions are accounted and optimization process is carried out under the design condition. To preserve the robustness of calculation, the maximum virtual displacement is limited by Y+ <10 and the camber line is smoothed via cubic B-spline interpolation. The complete blade profile is then generated by adding the prescribed blade thickness distribution to the camber line. The effectiveness of the method is demonstrated in the optimization of Stage35 compressor stage. Numerical results showed that this inverse method can effectively improve the internal flow structure and optimize the matching between blade rows, and this method is robust, efficient, and flexible. [ABSTRACT FROM AUTHOR]

Inlet total pressure distortion has great adverse effects on the aero-engine performance. The distorted flow passes through the compressor and becomes non-uniform in the downstream blade rows. Different from previous studies based on the assumption of circumferential uniformity, this study aims to improve circumferential non-uniform flow with the non-axisymmetric structure. Non-axisymmetric stator clearance was adopted to resolve the effects of non-uniform flow caused by inlet total pressure distortion in this paper. The 9 stators with tip clearance were installed in the distorted region and the flow field structure and performance under different operating conditions was studied. The study finds that the non-axisymmetric compressor with 9 tip clearance stators can ensure compressor efficiency while improving compressor stability margin. What's more, the separation range and strength in the distorted region can be reduced significantly and the anti-distortion capability of compressor can be enhanced. [ABSTRACT FROM AUTHOR]

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

The present research investigates the effect of the location and the width of single shallow circumferential groove casing treatment on the flow field and the stability improvement of NASA Rotor 37 utilizing the help of computational fluid dynamics. At first, steady state simulation of Rotor37 was presented for smooth casing (without groove). Then, forty five various grooved casing were simulated and compared with the smooth casing. The results indicated that narrow grooves had slight effect on the adiabatic efficiency but as the width of the groove was increased, a decline in efficiency was observed. The investigation on the stall margin revealed that narrow grooves next to the leading edge could improve the stall margin by a reduction in the size of vortex breakdown zone. Medium-width grooves displayed an effective role in delaying the separation- produced by shock wave and boundary layer interaction- on the blade suction side near the casing. This type of grooves could improve the stall margin more than narrow grooves when located on the top of separation zone near the blade suction side. Wide grooves had negative effect on the stall margin and caused a significant drop in the efficiency and the total pressure ratio of the compressor.

The unfavorable effects of non-uniform temperature inlet flow on gas turbine engine operations have always been a hindrance on the performance of turbo-fan engines. The propulsive efficiency is a function of the overall efficiency of turbofan engine which itself is dependent on other ambient parameters. Variation of inlet compressor temperature due to increase or decrease of aircraft altitude, air density, relative humidity, and geographical climate conditions affects the compressor performance. This research focuses on the turbofan transonic compressor performance due to ambient temperature distortion. A novel predictive approach based on neural network model has been implemented to predict the compressor performance and behavior at different ambient temperature conditions. The model produces substantially accurate results when compared to the results of CFD analysis. Computational results from CFD analysis show that engine thrust decreases at higher altitude, lower density and lower pressure regions. [ABSTRACT FROM AUTHOR]

• Cascade performance has been experimentally tested and numerically analyzed. • Relative humidity ranges from 20 to 53% at Mach number 1.11, and incidence of +5°. • Effects of humidity on blade loading, deviation, and loss are compared. • Loss is evaluated from profile, shock, and thermodynamic loss. Humid air degrades gas turbine aerodynamic performance. However, little is known about humidity effects on the compressor flow field. Blade loading distribution, deviation, and loss coefficient have been measured in a transonic compressor cascade for relative humidity (RH) values of 20, 32, 42, 45, 49 and 53% at Mach number of 1.11 and incidence of +5°. In addition, numerical studies have been conducted for the same conditions by incorporating the Classical Nucleation Theory (CNT) model and Hertz-Knudsen droplet growth model into ANSYS FLUENT. As the relative humidity increases, due to condensation, the pressure of upstream of the shock on the suction surface is increased; the pressure on the pressure surface is decreased; and shock is pushed downstream, reducing shock strength. However, due to the negligible shock-induced boundary layer separation, humidity has little effect on deviation. Loss coefficient is increased as the relative humidity increases because, even though the shock loss is reduced, heat addition to the flow from condensation decreases total pressure. [ABSTRACT FROM AUTHOR]

The adverse impacts of non-uniform inlet flow have been the focus for several decades with the increase of the operating range of engines. A deep understanding of the flow mechanism of distortion passing through a compressor is needed urgently and the improvement of the compressor performance becomes more and more important. In this paper, a non-axisymmetric stator is presented with significant non-axisymmetric characteristics in a transonic compressor to investigate compressor performance and flow field effects. A time-dependent three-dimensional Reynolds-averaged Navier-Stokes equation composed in 'Fluent Software Pack' is validated and used to perform the simulations. The flow fields with distorted inlet are obtained and the effects of original stator and non-axisymmetric stator in a transonic compressor are compared. The results are discussed in terms of the effects of non-axisymmetric stator on compressor performance, blockage of flow passage, rotor and stator. The results show that the non-axisymmetric stator influences not only the stator flow field but also the rotor flow field, so the efficiency and total pressure ratio are improved correspondingly. [ABSTRACT FROM AUTHOR]

Shock wave and tip leakage are important flow features at small length scales. These flow phenomena and their interactions play important roles in the performance of modern transonic fans and compressors. In most numerical predictions of these features, mesh convergence studies are conducted using overall performance data as criteria. However, less effort is made in assessing the quality of the predicted small-scale features using a mesh that yields a fairly accurate overall performance. In this work, this problem is addressed using the adaptive mesh refinement (AMR) method, which automatically refines the local mesh and provides very high resolution for the small-scale flow feature, at much less cost compared with globally refining the mesh. An accurate and robust AMR system suitable for turbomachinery applications is developed in this work and the widely studied NASA Rotor-37 case is investigated using the current AMR method. The complex interactions between the shock wave and the boundary layer, as well as those between the shock wave and the tip vortex, are accurately captured by AMR with a very high local grid resolution, and the flow mechanisms are analyzed in detail. The baseline mesh, which is considered to be "acceptable" according to the commonly used mesh convergence study, is unable to capture the detailed interaction between the shock wave and the boundary layer. Moreover, it falsely predicts the tip leakage vortex breakdown, which is a consequence of inadequate resolution in the tip region. Current work highlights the importance of a careful check of the mesh convergence, if small-scale features are the primary concern. The AMR method developed in this work successfully captures the flow details in the transonic compressor in an automatic fashion, and has been verified to be efficient compared with the globally mesh refinement or manually mesh regeneration. [ABSTRACT FROM AUTHOR]

This paper reports on a time-accurate numerical investigation of axial-skewed slot casing treatment in a high-speed axial fan. Twenty-two axial and radial skewed slots are placed over the casing of the rotor and choke to stall unsteady computations are conducted. Results show that endwall fluid is absorbed by the slots from the downstream part of the blade passage and is injected to the upstream part with a swirl contrary to the blade rotation. This gives the injected fluid a great circumferential velocity component in the relative frame of reference. The shock is, hence, pushed toward the trailing-edge plane and the pressure difference between the pressure and suction surfaces is reduced. Consequently, the occurrence of the leading-edge vortex spillage, which is believed to be the key to stall initiation for the current rotor, is postponed to lower mass flow rates. [ABSTRACT FROM AUTHOR]

Transonic compressors are widely used today in propulsion and industrial applications thanks to their higher specific work compared to subsonic. In this work, the aerodynamic optimization of a two-dimensional Computational Fluid Dynamics (CFD) model of the transonic cascade ARL-SL19 is described. The validated computational model is used for a multi-objective optimization of the cascade at three different inlet Mach numbers using a genetic algorithm and an artificial neural network, with the aim of reducing total pressure loss and increasing maximum pressure ratio. Finally, the optimized shapes on the Pareto fronts are investigated, analyzing mechanisms responsible for loss reduction and enhanced compression. Profiles having the lowest losses have flatter camberlines and reduced acceleration of flow on the suction side, while geometries achieving the highest pressure ratio values have a more cambered shape with a concave suction side.

Tip leakage vortex (TLV) in a transonic compressor rotor was investigated numerically using detached-eddy simulation (DES) method at different working conditions. Strong unsteadiness was found at the tip region, causing a considerable fluctuation in total pressure distribution and flow angle distribution above 80% span. The unsteadiness at near choke point and peak efficiency point is not obvious. DES method can resolve more detailed flow patterns than RANS (Reynolds-averaged Navier–Stokes) results, and detailed structures of the tip leakage flow were captured. A spiral-type breakdown structure of the TLV was successfully observed at the near stall point when the TLV passed through the bow shock. The breakdown of TLV contributed to the unsteadiness and the blockage effect at the tip region.

Instability initiated by rotating stall and surge limits the normal operating range of compressors, and seriously deteriorates the efficiency of gas-turbines. Partial surge is a new type of instability inception observed in a transonic axial flow compressor; it occurs in the form of axisymmetric low-frequency disturbances localized in the hub region and finally grows into rotating stall. For energy saving purposes, it is necessary to enhance the operating margin or stall margin of compressors. In this paper, a new design of casing treatment (axisymmetric arc-shaped slot casing treatment) is investigated to enhance the stall margin of a transonic compressor whose instability is initiated by partial surge. Both compressor performance and instability evolution before and after treatment are analyzed. The stall margin is successfully enhanced by 19.88% with the casing treatment, but the total pressure ratio only drops by 0.3% and the adiabatic efficiency does not quite change (increases by 0.15%), compared with the original solid wall case. The results of instability evolution present that partial surge occurs at nearly the same mass flow rate as the stall point of the solid wall case. As the compressor is further throttled, partial surge disappears. Then, rotating stall cells suddenly appear and trigger the formation of partial surge. In the presence of rotating stall cells and partial surge, the compressor suddenly and substantially turns into the final instability (substantial drop of compressor performance). The reasons for why the stall margin can be enhanced with this casing treatment are discussed, and the functions of the casing treatment are also discussed. [ABSTRACT FROM AUTHOR]

This paper reports on a numerical investigation of the radial penetration of injectors in a high-speed axial fan. Four annular injectors with different radial heights were placed over the casing and choke to stall unsteady computations were performed for each case. Results reveal that as long as the blade is “wall–stall” (i.e., stall is initiated from the tip), range extension is mostly due to injecting into the tip of the blade. If the injector is located close enough to the blade, however, increasing the radial penetration of the injector can provide considerable range extension by changing the rotor stalling mode from “wall–stall” to “blade–stall”. The rotor with zero tip clearance has been also studied in the current work. Results show that endwall injection can desensitize the compressor to the detrimental effects of the tip clearance. [ABSTRACT FROM AUTHOR]

Here, we use numerical analysis to study the effects of a shock control bump (SCB) on the performance of a transonic axial compressor blade section and to optimize its shape and location to improve the compressor performance. A section of the NASA rotor 67 blade is used for this study. Two Bézier curves, each consisting of seven control points, are used to model the suction and pressure surfaces of the blade section. The SCB is modeled with the Hicks-Henne function and, using five design parameters, is added to the suction side. The total pressure loss through a cascade of blade sections is selected as the cost function. A continuous adjoint optimization method is used along with a RANS solver to find a new blade section shape. A grid independence study is performed, and all optimization and flow solver algorithms are validated. Two single-point optimizations are performed in the design condition and in an off-design condition. It is shown that both optimized shapes have overall better performance for both on-design and off-design conditions. An analysis is given regarding how the SCB has changed the wave structure between blade sections resulting in a more favorable flow pattern. [ABSTRACT FROM AUTHOR]

The present research investigates the effect of the location and the width of single shallow circumferential groove casing treatment on the flow field and the stability improvement of NASA Rotor 37 utilizing the help of computational fluid dynamics. At first, steady state simulation of Rotor37 was presented for smooth casing (without groove). Then, forty five various grooved casing were simulated and compared with the smooth casing. The results indicated that narrow grooves had slight effect on the adiabatic efficiency but as the width of the groove was increased, a decline in efficiency was observed. The investigation on the stall margin revealed that narrow grooves next to the leading edge could improve the stall margin by a reduction in the size of vortex breakdown zone. Medium-width grooves displayed an effective role in delaying the separation- produced by shock wave and boundary layer interaction- on the blade suction side near the casing. This type of grooves could improve the stall margin more than narrow grooves when located on the top of separation zone near the blade suction side. Wide grooves had negative effect on the stall margin and caused a significant drop in the efficiency and the total pressure ratio of the compressor. [ABSTRACT FROM AUTHOR]