Search Results (9)

Since the beginning of the 21st century, the supercritical carbon dioxide (sCO₂) Brayton cycle has emerged as a hot topic of research in the energy field. Among its key components, the sCO₂ compressor has received significant attention. In particular, axial-flow sCO₂ compressors are increasingly being investigated as power systems advance toward high power scaling. This paper reviews global research progress in this field. As for performance characteristics, currently, sCO₂ axial-flow compressors are mostly designed with large mass flow rates (>100 kg/s), near-critical inlet conditions, multistage configurations with relatively low stage pressure ratios (1.1–1.2), and high isentropic efficiencies (87–93%). As for internal flow characteristics, although similarity laws remain applicable to sCO₂ turbomachinery, the flow dynamics are strongly influenced by abrupt variations in thermophysical properties (e.g., viscosities, sound speeds, and isentropic exponents). High Reynolds numbers reduce frictional losses and enhance flow stability against separation but increase sensitivity to wall roughness. The locally reduced sound speed may induce shock waves and choke, while drastic variation in the isentropic exponent makes the multistage matching difficult and disperses normalized performance curves. Additionally, the quantitative impact of a near-critical phase change remains insufficiently understood. As for the experimental investigation, so far, it has been publicly shown that only the University of Notre Dame has conducted an axial-flow compressor experimental test, for the first stage of a 10 MW sCO₂ multistage axial-flow compressor. Although the measured efficiency is higher than that of all known sCO₂ centrifugal compressors, the inlet conditions evidently deviate from the critical point, limiting the applicability of the results to sCO₂ power cycles. As for design and optimization, conventional design methodologies for axial-flow compressors require adaptations to incorporate real-gas property correction models, re-evaluations of maximum diffusion (e.g., the DF parameter) for sCO₂ applications, and the intensification of structural constraints due to the high pressure and density of sCO₂. In conclusion, further research should focus on two aspects. The first is to carry out more fundamental cascade experiments and numerical simulations to reveal the complex mechanisms for the near-critical, transonic, and two-phase flow within the sCO₂ axial-flow compressor. The second is to develop loss models and design a space suitable for sCO₂ multistage axial-flow compressors, thus improving the design tools for high-efficiency and wide-margin sCO₂ axial-flow compressors. Full article

In this study, stall flutter onset prediction in a transonic compressor is carried out using the (uncoupled) energy method with Fourier transform. As the study is conducted computationally using computational fluid dynamics (CFD)-based simulations, the energy method was employed due to its higher computational efficiency by implementing the one-way FSI (Fluid Structure Interaction) model. The energy method is relatively uncommon for determining the aerodynamic damping and flutter prediction, specifically in blade stall conditions for the 3D blade passages. The NASA Rotor 67 was chosen for the validation of the study due to the availability of a wide range of experimental data. A flutter prediction analysis was performed computationally using CFD for the two-blade passages of the rotor in the peak efficiency and stall regions. Prior to this, the modal analysis on the prestressed blade was conducted, considering the centrifugal effects. The modal analysis provided accurate blade frequency and amplitude, which were the inputs of the flutter analysis. The first three modes of blade resonance were studied with a range of nodal diameters within near-peak efficiency and stall regions. The energy method implemented in this study for the flutter analysis was successfully able to predict the aerodynamic damping coefficients of the first three modes for a range of nodal diameters from the periodic-unsteady solution of the defined blade oscillation within the regions of interest (peak efficiency and stall point). The results of the study confirm the rotor blade’s stability within the near-peak region and, most importantly, the prediction of the flutter onset in the stall region. The study concluded that the computationally inexpensive and time-efficient energy method is capable of predicting the stall flutter onset. In the future, further validations of the energy method and investigations related to flow mechanism of stall flutter onset are planned. Full article

This paper presents the development of high-specific-speed mixed flow/centrifugal compressor vaned diffusers. Specifically, the design of a test rig that will make the visualization of shock waves on diffuser vanes manageable is addressed in the current study. In this particular case, linearization of an existing state-of-the-art compressor stage was used. For the computational modeling, a series of RANS analyses were conducted to examine the flow characteristics of the two cases explored: a complete transonic cascade and an idealized periodic passage. The distinct behavior exhibited by each vane passage within the entire cascade offers the opportunity to analyze the shockwave structures across a mass flow range of ±9% around the design point. Overall, the pressure coefficient distributions and flow field patterns appear to align with the single-passage conditions, although there are some minor lateral wall influences, particularly in the first passage close to the suction lateral wall. However, because of the nature of the flow, which is characterized by high velocity and density differences near the vanes, the equivalent mass flow per individual passage was difficult to estimate. This may also be attributed to the small endwall axial vortices; nonetheless, for the purposes of this paper, this was of little consequence. Full article

The aim of this research was to explore the mechanisms underlying the effect of self-circulating casing treatment with different circumferential coverage ratios on the aerodynamic performance of a transonic centrifugal compressor. A three-dimensional unsteady numerical simulation was carried out on a Krain impeller. The circumferential coverage ratios of the self-circulating casings were set to 36%, 54%, 72% and 90%, respectively. The numerical results showed that the Stall Margin Improvement (SMI) increased with the increase in circumferential coverage ratios. The self-circulating casing with a 90% circumferential coverage ratio exhibited the highest SMI at 20.22%. Internal flow field analysis showed that the self-circulating casing treatment improved the compressor stability by sucking the low-speed flow in the blade tip passage and restraining the leakage vortexes breaking, which caused flow blockage. The compressor performance was improved at most of the operating points, and the improvement increased with increase in circumferential coverage ratio. The improvement in compressor performance was mainly attributed to reduction in the area of the high relative total pressure loss in the blade tip passage and significant decrease in the flow loss by the self-circulating casings. Full article

This study aimed to investigate the influence of the bleeding position of a self-circulating casing on the aerodynamic performance of a transonic centrifugal compressor. Three types of self-circulating structures with the bleeding positions of 11% Ca (the axial chord length of the blade tip), 14% Ca and 20% Ca from the leading edge of the blade were studied by using the numerical simulation method, with the Krain impeller taken as the research object. It was found that all three types of self-recirculating casing treatments can expand the stable operating range of the impeller, and that at medium and small flow rates, the total pressure ratio and efficiency of the impeller increase gradually with the backward movement of the bleeding position. The self-circulating casing treatment can restrain the development of tip leakage vortex, reduce the blockage area, and improve the stability of the impeller by sucking low-energy fluid. The farther back the bleeding position is, the greater the bleeding mass flow rate of the self-circulating casing for the low-energy fluid in the blade-tip passage becomes. Additionally, a greater inhibition effect on the tip leakage vortex, and a better effect of improving the performance and stability of the impeller, can be obtained. The best air bleeding position is 20% Ca, but it is not directly above the blade-tip blockage center of the solid wall casing passage. Instead, it is downstream of the blockage area. Full article

A highly loaded transonic centrifugal compressor is aerodynamically designed and numerically investigated. The objectives are to improve the compressor efficiency by using tandem impeller configuration and 3D free-formed blade design concepts. This approach has the potential to control both the transonic and distorted flows within impeller passages. The results suggest that employing the tandem impeller configuration can significantly improve the compressor efficiency by 1.4%. The efficiency gain is mainly contributed by the improved uniformity of the impeller discharge flow achieved when the newly generated inducer-shed vortices rearrange the secondary flow pattern. In addition, the location of the impeller passage shock moves downstream due to the potential effect within the tandem impeller and locally changes the back pressure of the inducer. These factors mitigate flow losses in the impeller and diffuser. Furthermore, the 3D design concepts of forward blade sweep and negative lean are employed in the tandem impeller configuration. The forward sweep design of the inducer weakens flow acceleration before the passage shock, and the negative lean design optimizes the secondary flow pattern, which yield an additional compressor efficiency improvement of 0.7%. The study conducted in this paper provides a valuable reference for future advanced transonic centrifugal compressor designs. Full article

The demands to apply transonic centrifugal compressor have increased in the advanced gas turbine engines. Various techniques are used to increase the aerodynamic performance of the centrifugal compressor. The effects of the inclined leading edges in diffuser vanes of a transonic centrifugal compressor on the flow-field unsteadiness and noise generation are investigated by solving the compressible, three-dimensional, transient Navier–Stokes equations. Diffuser vanes with various inclination angles of the leading edge from shroud-to-hub and hub-to-shroud are numerically modeled. The results show that the hub-to-shroud inclined leading edge improves the compressor performance (2.6%), and the proper inclination angle is effective to increase the stall margin (3.88%). In addition, in this study, the transient pressure variations and radiated noise prediction at the design operating point of the compressor are emphasized. The influences of the inclined leading edges on the pressure waves were captured in time/space domain with different convective velocities. The pressure fluctuation spectra are calculated to investigate the tonal blade passing frequency (BPF) noise, and it is shown that the applied inclination angles in the diffuser blades are effective, not only to improve the aerodynamic performance and stall margin, but also to reduce the BPF noise (7.6 dB sound pressure level reduction). Moreover, it is found that the diffuser vanes with inclination angles could suppress the separation regions and eddy structures inside the passages of the diffuser, which results in reduction of the overall sound pressure level and the broadband noise radiated from the compressor. Full article

In this study, the influences of the flow cut and axial lift of the impeller on the aerodynamic performance of a transonic centrifugal compressor were analyzed. The flow cut is a method to reduce the flow rate by decreasing the impeller passage height. The axial lift is a method of increasing the impeller passage height in the axial direction, which increases the impeller exit width (B2) and increases the total pressure. A NASA CC3 transonic centrifugal compressor with a backswept angle was used as a base compressor. After applying the flow cut, the total pressure at the target flow rate was lower than the total pressure at the design point due to the increase in the relative velocity at the impeller exit. After applying the axial lift, the total pressure at the design flow rate was increased, which was caused by the reduction in the relative velocity as the passage area at the impeller exit was increased. By applying the flow cut and axial lift methods, it was shown that the variation in relative velocity at the impeller exit has a significant effect on the variation in total pressure. In addition, it was found that the relative velocity at the impeller exit of the target flow rate is maintained similar to the base impeller when the flow cut and the axial lift are combined. Therefore, by combining the flow cut and the axial lift, three transonic centrifugal impellers with flow fractions of 0.7, 0.8, and 0.9 compared to the design flow rate were newly designed. Full article

Transonic centrifugal compressors with high performance are required in the oil and gas industries, modern gas turbine engines, and turbochargers. The sweep of the blades is one of the crucial features that have a significant influence on their performance. This paper numerically investigates mechanisms by which sweep affects the performance of a transonic impeller with twin splitters. Sweep is defined as scaling up or down the shroud chord, and the variation range of the sweep angle has been chosen from −25 to +25°. In the current case, results show that the variation of choke mass flow rate, pressure ratio, and efficiency value is around 1%. If the centrifugal compressor has a higher pressure ratio or a higher front loading, the sweep effect on compressor performance will be even stronger. The essential aerodynamic effect of sweep is the spanwise redistribution of the front loading, resulting in effects on the shock structure, the tip leakage vortex, and the flow separation. On the shroud section, forward sweep restricts the front loading, the shock strength, and the tip leakage vortex, which reduces the loss near the casing. On the hub section, aft sweep suppresses the front loading and the flow separation, which reduces the loss near the hub. It is the delicate balance between controlling the loss near the hub and the loss near the casing that determines the optimal sweep angle design. Full article