Improving the efficiency of COMPRESSORs has been one of the most important goals of researchers over the years. In this paper, three different methods are presented for parameterization and BLADE optimization of axial flow COMPRESSOR. All methods consist of flow analysis tool, optimization algorithms, and parametric geometry generation tool that are different in each approach. An objective function is defined based on the aerodynamic performance of BLADE in the acceptable incidence angles range. A double circular arc BLADE is used as the initial guess for all methods. The performance of optimized BLADEs and the initial BLADE are compared to evaluate the capability of various methods, and a good agreement is achieved. The results show that the level of performance improvement in each method depends on the number and type of the chosen parameters. All three methods have improved BLADE performance at the design incidence angle. However, only the first method shows significant performance improvement in off-design conditions.

The helium COMPRESSOR has the inherent characteristics of a lower single-stage pressure ratio and a higher number of stages than an air COMPRESSOR. The highly loaded design method effectively addressed the compressibility issue of the helium COMPRESSOR. However, the COMPRESSOR designed with this technique has narrow passages, short BLADEs, and a large bending angle, making the end-wall secondary flow more intense than a conventional COMPRESSOR. In this paper, numerical simulation and experimental validation has been conducted to identify the effectiveness of the axial slot close to the BLADE root in improving end-wall secondary flow in a high-load helium COMPRESSOR cascade, and to provide data and experimental support for the engineering application of high-load helium COMPRESSORs. The analytical results show that slotting can utilize the self-pressure difference to generate gap leakage vortices, and the axial momentum generated by the leakage vortices blows away the vortices formed due to the separation of corner area. The airflow flows close to the suction surface of the BLADE and breaks away at the trailing edge of the BLADE, merges with the main flow and forms a new vortex. As the height of the channel increases, the blowing away of the vortices in the corner region becomes more pronounced and the cascade improvement performance is better. The test results show that the total pressure loss coefficient at the design operating point is reduced by 6.167% when a slot height of 8.53 mm is positioned at 65% Ca (axial chord length). The improvement effect becomes 16.469% better at a 4° attack angle.

To improve the aerodynamic characteristics of COMPRESSOR BLADEs, a novel asymmetric leading edge (ASYLE) has been introduced and shown to offer superior performance. However, the aerodynamic robustness of such specially designed leading edge (LE) remains unclear due to the considerable uncertainty problems it presents. This paper investigates the robustness of ASYLE BLADE under both geometric and operational uncertainties. Profile deviations within ±0. 05mm were introduced to investigate the influence of manufacturing errors. In addition, the perturbated inflow angles between ±0. 375° were considered for uncertain inflow conditions. The statistic aerodynamic performance as well as operating dispersibilities at Ma=0. 7 were obtained by the non-intrusive polynomial chaos (NIPC) method. The results show that considering uncertain profile errors, the operating range of ASYLE BLADE is 2. 3° wider than original leading edge (ORILE) BLADE and the dispersion of total pressure loss can be reduced by 53. 1% at β1=45. 8°. Regarding uncertain inflow angle variations, the total pressure loss dispersion of ASYLE BLADE can be reduced by 93. 8% at β1=50. 8°. The ASYLE shows better overall aerodynamic robustness than ORILE upon considering uncertainty limits. The influence propagations in the flow fields of both uncertainties were further analysed, which revealed that the variations of separation bubble structure near LE are the direct cause to the aerodynamic uncertainties. The ASYLE design effectively controls the size and variation of LE separation bubble and thus demonstrates better aerodynamic robustness.

This paper exemplified a way to increase pressure ratio and improve efficiency with addition of multi splitters in centrifugal impeller with a vaneless diffuser. DDA 404-III back swept impeller of centrifugal COMPRESSOR, studied through experiment is modified with the addition of splitters BLADEs and a sample impeller is designed and analyzed with big splitter close to pressure surface and small splitter close to suction surface. Keeping all flow conditions and impeller definitions, same as experimentally validated impeller, in order to investigate effects of the location of the splitters between two main BLADEs. It was observed that total pressure ratio is increased from 4. 1 to 4. 5 with 2 % increase in efficiency with big splitter close to pressure surface of main BLADE and small splitter close to suction surface of main BLADE. It was observed that relative Mach number reduces at inlet of modified impeller.

The choice of geometrical shape of the BLADEs has a considerable effect on aerodynamic performance and flow characteristics in axial COMPRESSORs. In this paper, the effects of the BLADEs shape on the aerodynamic design characteristics are investigated based on Streamline Curvature Method (SCM). Initially, the Streamline Curvature Method (SCM) is used for designing a two-stage axial COMPRESSOR. Comparing the current results with experimental ones indicates good agreement. The first stage of the axial COMPRESSOR is selected with three different BLADE profiles. The first case (case I) has the polynomial camber with naca thickness distribution series 6. The second case (case II) has the standard naca profile series 6 and the third case (case III) has the modified standard naca profile series 4. Results reveal that using the standard airfoils in the stators leads to improved flow conditions such as loss coefficient and pressure ratio. On the contrary, this profile selection may cause an increase in the stagger angle that is not favorable. Aerodynamic Design with a polynomial camber line in the rotor demonstrates a better aerodynamic behavior in loss coefficient, pressure ratio and diffusion factor. Whilst the use of such a camber line in the stator leads to the formation of less favorable aerodynamics conditions in comparison to the standard airfoil.

One of the most popular forming processes is the shape rolling process in which the desired shape change is achieved by pressing two rollers with a special shape in the opposite rotational direction. In order to improve the product’ s quality and reduce production costs, accurate analysis of the shape rolling process of the COMPRESSOR BLADEs as well as the investigation of the effective parameters have been done. First, the shape rolling process of a typical COMPRESSOR BLADE was simulated based on the experimental data using the finite element method and Design of Experiment (DOE). Then, the effect of various process parameters, including the thickness and width of the preform, the roller diameter, the thickness and width of the flash channel, and the number of the rolling steps on two objectives, namely the rolling force and the amount of the flash were investigated. The obtained data were analyzed by Analysis of Variance (ANOVA), and the contributory factors of the shape rolling process were identified. The results revealed that all of the considered factors affected the rolling load, but only the initial sheet's width and thickness were the factors with impact on the volume of the flash as the second objective. The required process load decreased by increasing the number of the rolling steps, but the rolling load increased by increasing other factors. Furthermore, increasing the thickness and width of the initial sheet increased the flash volume.