A one-dimensional analytical shock loss prediction method was proposed to tailor the shock system, i.e. the strength of the first and second passage shock, and reduce the shock loss in a supersonic cascade. To develop the one-dimensional analytical model, the shock system in a supersonic cascade was divided into four processes which can be seen in most supersonic compressor cascade, i.e. the flow upstream the extending-external shock, the flow between the extendingexternal shock and the first passage shock, the accelerating flow from the first passage shock to the second passage shock, and the flow downstream of the second passage shock. Based on some flow assumptions and experimental empirical correlations, the complex flows, containing the shock system, in the blade passage of a supersonic compressor cascade could be described with one-dimensional relationships, which can be used to predict shock losses along the flow passage rapidly and determine the shock system improving direction for achieving lower shock loss while keeping the same cascade static pressure ratio. In order to validate the one-dimensional analytical method, the shock system of two supersonic cascades ARL-SL19 and DLR-PAV-1.5 are modified based on the analysis of the model. The modified cascades achieved about 29% and 25% reduction of shock loss at redesign point compared with baseline cascades, respectively.
Recently, a new type airfoil for variable inlet guide vane (VIGV), featuring “dual-peak” surface velocity pattern at high incidence, is proposed and shows wide low-loss operation range. To further improve its performance, this paper researches the influence of leading edge (LE) thickness and shape on the loss level and surface velocity features of the “dual-peak” type airfoil. Firstly, a polynomial-based continuous-curvature leading edge design method was briefly introduced and used in the LE redesign of sample airfoils. Then, steady simulations based on Reynolds-Averaged Navier-Stokes method (RANS), carried out by commercial software CFX after grid independent study, were used to determine the aerodynamic performance, surface velocity distribution and boundary-layer behaviors of all research airfoils. Simulation results indicate that there exists an optimized range of LE relative thickness that can achieve lower airfoil loss level at high incidence condition. For Case 1 ([Formula: see text]) and Case 2 ([Formula: see text]), the optimized LE relative thickness range is [Formula: see text] and [Formula: see text]. The LE shape optimization can further reduce the maximum incidence condition loss coefficient with proportion up to 18% for airfoils with optimal LE thickness. Analysis of flow mechanism indicates that the optimized LE thickness and shape can reduce the suction spike height and subsequent adverse pressure gradient, therefore, decrease the LE separation scale and result in a lower loss coefficient. As an application, a dual peak VIGV with circular LE, presented in previous paper as the optimized VIGV, is redesigned in the LE portion according to the research findings and achieved 0.6 percent improvement in passage-averaged total pressure recovery coefficient [Formula: see text] at extreme high stagger angle point and the low-loss operation range extends with about 5°, which confirms the effectiveness of the research findings in three-dimensional environment.
The blade geometry design method is an important tool to design high performance axial compressors, expected to have large design space while limiting the quantity of design variables to a suitable level for usability. However, the large design space tends to increase the quantity of the design variables. To solve this problem, this paper utilizes the normalization and subsection techniques to develop a geometry design method featuring flexibility and local adjustability with limited design variables for usability. Firstly, the blade geometry parameters are defined by using the normalization technique. Then, the normalized camber angle f1(x) and thickness f2(x) functions are proposed with subsection techniques used to improve the design flexibility. The setting of adjustable coefficients acquires the local adjustability of blade geometry. Considering the usability, most of the design parameters have clear, intuitive meanings to make the method easy to use. To test this developed geometry design method, it is applied in the design of a transonic, two flow-path axial fan component for an aero engine. Numerical simulations indicate that the designed transonic axial fan system achieves good efficiency above 0.90 for the entire main-flow characteristic and above 0.865 for the bypass flow characteristic, while possessing a sufficiently stable operation range. This indicates that the developed design method has a large design space for containing the good performance compressor blade of different inflow Mach numbers, which is a useful platform for axial-flow compressor blade design.
This paper researches the parametric optimization of a two-stage transonic compressor having a large air bypass at partial rotating speed according to flow analysis for a turbine-based combined cycle engine (TBCC). To obtain adequate thrust, the inlet transonic compressor of the turbofan part of the TBCC is required to have a wider frequently used corrected rotating speed range and a larger mass-flow rate at low rotating speed, which is different from a typical transonic compressor. The one-dimensional blade design parameters and flow path of the baseline two-stage transonic compressor are introduced. With the widely used CFD software Numeca, the three-dimensional flow fields of the baseline transonic compressor and effects of the flow path between Stage 1 and Stage 2 on the inlet mass flow rate are analyzed for indicating the further improvement direction. For design speed (NC = 1.0), to improve the efficiency at the design point, parametric research is carried out on Rotor 2 to optimize the shock structure and strength, resulting in enhanced efficiency at the design point due to reduced shock loss of Rotor 2. For partial speed (NC = 0.8 and 0.7), since the flow field analysis indicates that the flow blockage in S1 limits the entire mass flow rate, the parametric redesign of stator S1 aims at obtaining an increased blade throat width to enhance the flow capacity of S1. Simulation confirms the increase in the mass-flow rate and efficiency at partial speed due to the reduction in flow blockage and related viscous losses. Aerodynamic analysis at representative operation points indicates that the modifications of R2 and S1 lead to obvious aerodynamic improvement at all rotating speeds (NC = 1.0 to 0.7), while maintaining sufficient stall margin.
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