The compressor is a key component in the supercritical carbon dioxide (SCO₂) Brayton cycle. In this paper, the authors designed a series of supercritical CO₂ compressors with different parameters. These compressors are designed for 100 MWe, 10 MWe and 1 MWe scale power systems, respectively. For the 100 MWe SCO₂ Brayton cycle, an axial compressor has been designed by the Smith chart to test whether an axial compressor is suitable for the SCO₂ Brayton cycle. Using a specific speed and a specific diameter, the remaining two compressors were designed as centrifugal compressors with different pressure ratios to examine whether models used for air in the past are applicable to SCO₂. All compressors were generated and analyzed with internal MATLAB programs coupled with the NIST REFPROP database. Finally, the design results are all checked by numerical simulations due to the lack of reliable experimental data. Research has found that in order to meet the de Haller stall criterion, axial compressors require a considerable number of stages, which introduces many additional problems. Thus, a centrifugal compressor is more suitable for the SCO₂ Brayton cycle, even for a 100 MWe scale system. For the performance prediction model of a centrifugal compressor, the stall predictions are compared with steady numerical calculation, which indicates that past stall criteria may also be suitable for SCO₂ compressors, but more validations are needed. However, the accuracy of original loss models is found to be inadequate, particularly for lower flow and higher pressure ratio cases. Deviations may be attributed to the underestimation of clearance loss according to the result of steady simulation. A modified model is adopted which can improve the precision to a certain extent, but more general and reasonable loss models are needed to improve design accuracy in the future.

The aero-engine casing is a key component for carrying loads. With the purpose of improving the thrust-weight ratio of the aero-engine, the casing is required to be designed to be as thin as possible. Therefore, the vibration of aero-engine's rotor, support, and casing will be easily coupled causing the whole engine's vibration to be more serious. Considering the structural vibration propagation is essentially the vibration energy transmission, the structural intensity (SI) method is popular and widely used to investigate the transmission phenomena of vibration energy in vibrating structures. This method combines forces with velocities to quantify the vibrational energy flow (VEF) transmitted in the structures by its directions and magnitude. Therefore, the SI fields are quantified by the developed computation system which combines the finite element design language and the in-house code. And a model of dual-rotor–support–casing coupling system subjected to the unbalanced forces of the rotors is established in this paper. The scalar and vector diagrams of instantaneous SI fields are visualized to show the main vibration energy transmission paths among these three parts. Moreover, the relationship between the SI and the mechanical energy is derived from the kinetic equation. According to this relationship, the phenomenon that the vibration energy and the strain energy are always converted to each other in the middle part of the rotor shaft with the first-order bending mode is discussed, which reveals the cause of the first-order bending mode of the rotor from a microscopic point of view.

The paper presents a multi-objective optimization of circumferential casing grooves geometries for the NASA Rotor 37 transonic compressor. The depth normalized by the tip clearance and the width normalized by the tip chord are selected as the design variables. The stall margin and peak efficiency are used as the objective functions. The Latin Hypercube Sampling technique was used to select the sample points in the design space. Based on the numerical results of the sample points, the radial basis function network model of the artificial neural network was constructed. The NSGA-II multi-objective evolutionary algorithm is then employed to search for Pareto-optimal solutions. The leave-one-out cross validation method was also used to evaluate the precision of the radial basis function network model. The results of the optimization show the present method can be effectively used for the design of circumferential casing grooves to take account of the stall margin and efficiency. From the Pareto-optimal solutions, two groove configurations are selected and the internal flow fields are compared with the smooth casing. The effect mechanism of the circumferential casing grooves on the performance of the transonic compressor is discussed by the analysis of the flow in the blade tip region.

The evolution mechanism of the boundary layer and coherent structures in a low-pressure turbine blade is discussed. Five different incidence angles over the T106A blade for a Mach number Ma = 0.404 and Reynolds number Re = 0.6 × 10⁵ (based on the axial chord and outlet velocity) are performed using large eddy simulation method. The calculation results at +7.8 incidence angle are agreed well with the experimental and direct numerical simulation data. The influence of the incidence angle on the flow field is mainly shown at the front of the suction side and pressure side. As the incidence angle changes from positive to negative, the separation bubble near the leading edge disappears and the blade loading decreases gradually. When the incidence angle reduces to −5°, separation bubble appears near the leading edge of the pressure side. At the case of incidence angle equaling −10°, the length of time-averaged separation bubble on the pressure side grows to 39% axial chord and the evolution process of the coherent structures is extremely complex. The spanwise vortexes roll up near the leading edge and gradually evolve into streamwise vortexes. High-energy fluid in the main flow was driven to near-wall zone by the rotating effect of streamwise vortexes, which increases the fluid momentum inside the boundary layer. The streamwise vortexes are stretched by the strong acceleration of the flow until they transport to the trailing edge.

An improved compressible model for estimating tip clearance loss in transonic compressors is presented with the emphasis on the effects of blade tip loading distribution and double leakage flow. Tip clearance flow is treated as three parts along the chord and the progressive relations from upstream to downstream part is revealed to be responsible for the formation of tip clearance flow. Control volume method is applied to simplify the mixing process and calculate the mixed-out loss for the three parts, separately. Computational study shows that mass flow of the incoming flow entering the control volume is consistent with that passing through an equivalent area of about half of tip leakage vortex region. The new model reveals that the second part of tip clearance flow has a larger mixed-out loss capacity than the two other parts. This difference is attributed to two factors: larger injection flow angle and more enrolled incoming flow, and both factors tend to increase the mixed-out loss. The success of the model implies that blade design or flow control strategies turning the tip clearance/main flow interface’s arrival onto blade tip pressure side downstream and the shock’s impingement point onto blade tip suction side upstream may be beneficial in desensitizing compressor performance to tip clearance size, without trading off pressure rise.