Numerical Investigation of Base-Setting of Stator’s Stagger Angles for a 15-stage Axial-Flow Compressor简

A 15-stage axial-flow compressor utilized in steel industry was studied in this paper. All the stator＇s stagger an- gles of the compressor are variable to ensure the multistage compressor operate effectively within a wide range of flow rate and meanwhile satisfy the demand for sufficient pressure ratio, adiabatic efficiency and stall margin. Three in all different base-settings of stator＇s stagger angles were presented and commercial CFD soft-ware was applied to obtain the overall performance characteristics. The results showed that both of the optimized base-settings improved the performances both in summer and winter conditions, although the adiabatic efficiency was somewhat decreased. Taking incidence angle and stage loading into consideration, differences among the three cases were analyzed in detail. On the basis of numerical computations, the performance could be effectively improved through adjusting the base-setting of stator＇s stagger angles.     

Investigation on centrifugal impeller in an axial-radial combined compressor with inlet distortion

Assembling an axial rotor and a stator at centrifugal compressor upstream to build an axial-radial combined compressor could achieve high pressure ratio and efficiency by appropriate size augment.Then upstream potential flow and wake effect appear at centrifugal impeller inlet.In this paper,the axial-radial compressor is unsteadily simulated by three-dimensional Reynolds averaged Navier-Stokes equations with uniform and circumferential distorted total pressure inlet condition to investigate upstream effect on radial rotor.The results show that spanwise nonuniform total pressure distribution is generated and radial and circumferential combined distortion is formed at centrifugal rotor inlet.The upstream stator wake deflects to rotor rotation direction and decreases with blade span increases.Circumferential distortion causes different separated flow formations at different pitch positions.The tip leakage vortex is suppressed in centrifugal blade passages.Under distorted inlet condition,flow direction of centrifugal impeller leading edge upstream varies evidently near hub and shroud but varies slightly at mid-span.In addition,compressor stage inlet distortion produces remarkable effect on blade loading of centrifugal blade both along chordwise and pitchwise.     

Numerical investigation of tip clearance effects in an axial transonic compressor

Numerical investigations of the Darmstadt transonic single stage compressor (DTC), in the Rotor1-Stator1 configuration, aimed at advancing the understanding of the effect of different rotor tip gaps and transition modelling on the blade surfaces are presented. Steady three dimensional Reynolds Averaged Navier Stokes (RANS) simulations were performed to obtain the flow fields for the different configurations at different operating conditions using the RANS-Solver TRACE. The stage geometry and the multi-block structured grid were generated by G3DMESH and a grid sensitivity analysis was conducted. For the clearance gap region, a fully gridded special H-grid was chosen. Comparisons were made between the flow characteristic at design speed, representative for a transonic flow regime, and at 65% speed, representative for a subsonic flow regime. The computations were used to analyse the flow phenomena through the tip clearance region for the different configurations and their impact on the performance of the compressor stage.     

Impact of Wake on Downstream Adjacent Rotor in Low-speed Axial Compressor

This paper investigates the effect of upstream wake on unsteady separated flow field of downstream adjacent vanes by solving 2D unsteady Reynolds-averaged Navier-Stokes equations discretized by a high-order scheme. The results indicated that the maximum relative reduction of loss coefficient was 27.2% when the relative passing frequency (the ratio of wake passing frequency to the characteristic frequency of trailing-vortex shedding of downstream adjacent rotor) is less than 2.5. The amplitude of wake defect exists a threshold value and the aerodynamic performance is enhanced monotonically with the amplitude of wake defect basically. The effective range of incidence must be greater than 8°, which is near the stall boundary.     

Experimental Investigation of the Three—Dimensional Flow in an Annular Compressor Cascade at Large Incidence

This paper presents a detailed experimental investigation concerning the influence of blade loading (incidence)on the three-dimensional flow in an annular compressor cascade.The data are acquired at four incidence anglesunder low Mach number and low Reynolds number conditions.Experimental techniquss include the oil-filmvisualization on the profile and the endwall surfaces,the laser-sheet visualization of the flow field inside theblade passage,and the measurement by radial-circumferential traveress using a seven-hole probe.The behaviorand nature of the three-dimensional flow with severe separations inside the blade passage and at the exit areobtained.The distributions of the total pressure loss,static pressure,velocity and outflow angle are also given.These results are valuable for establishing the physical model of the three-dimensional complex flow in axialcompressors and for examining the computational procedures.     

Study on the effects of end-bend cantilevered stator in a 2-stage axial compressor

Leading edge recambering is applied to the cantilevered stator vanes in a 2-stage compressor in this paper. Different curving effects are produced when the end-bend stator vanes are stacked in different ways. Stacking on the leading edge induces a positive curving effect near the casing. When it is stacked on the centre of gravity, a negative curving effect takes place. The numerical investigation shows that the flow field is redistributed when the end-bend stators with leading edge stacking are applied. The variations in the stage matching for the mainstream and near the hub have an impact on the performance of the 2-stage compressor. The isentropic efficiency and the total pressure ratio of the compressor are increased near the design condition. The compressor total pressure ratio is decreased near choke and near stall. The maximum flow rate is reduced and the stall margin is decreased.     

Numerical investigation of the unsteady tip leakage flow and rotating stall inception in a transonic compressor

<正>It is well known that tip leakage flow has a strong effect on the compressor performance and stability. This paper reports on a numerical investigation of detailed flow structures in an isolated transonic compressor rotor-NASA Rotor 37 at near stall and stalled conditions aimed at improving understanding of changes in 3D tip leakage flow structures with rotating stall inception.Steady and unsteady 3D Navier-Stokes analyses were conducted to investigate flow structures in the same rotor.For steady analysis,the predicted results agree well with the experimental data for the estimation of compressor rotor global performance.For unsteady flow analysis, the unsteady flow nature caused by the breakdown of the tip leakage vortex in blade tip region in the transonic compressor rotor at near stall condition has been captured with a single blade passage.On the other hand, the time-accurate unsteady computations of multi-blade passage at near stall condition indicate that the unsteady breakdown of the tip leakage vortex triggered the short length-scale-spike type rotating stall inception at blade tip region.It was the forward spillage of the tip leakage flow at blade leading edge resulting in the spike stall inception. As the mass flow ratio is decreased,the rotating stall cell was further developed in the blade passage.     

Numerical study of the effects of lid oscillation on the periodic flow pattern and convection heat transfer in a triangular cavity

This study is concerned with a periodic flow pattern with mixed convection in a triangular cavity caused by the effects of lid oscillation and buoyancy. The dimensionless stream function–vorticity formulation is adopted, and a curvilinear grid method for solving the stream function–vorticity equations in irregular geometries is used. Attention is in particular focused on the flow behaviour under the interaction between the frequency of the oscillation of the lid velocity and the frequency of the natural periodic flow. Meanwhile, numerical predictions of the thermal characteristics represented by local and average Nusselt numbers on the walls as well as the transient flow patterns are also provided. Results show that the frequency of oscillation of lid velocity and the natural periodic flow frequency both appear to be major frequencies in the frequency spectrum of the cavity flow, for cases with a dimensionless lid oscillation frequency less than 0.5. When the dimensionless frequency of the oscillation of the lid velocity is equal to or higher than 0.5, the flow and thermal fields are completely locked-on to the lid oscillation, and the natural periodic flow frequency is no longer visible in the spectrum.     

Experimental and numerical investigation of the unsteady tip leakage flow in axial compressor cascade

For convenience of both measurement and adjusting the clearance size and incidence, the current research is mainly conducted by experiments on an axial compressor linear cascade. The characteristics and the condition under which the unsteadiness of tip leakage flow would occur were investigated by dynamic measuring in different clearances, inlet velocities and incidences. From the experiment it is found that increasing tip clearance size or reducing rotor tip incidence can affect the strength of the tip clearance flow. Then the experimental results also indicate the tip leakage shows instability in certain conditions, and the frequency of unsteadiness is great influenced by inflow angle. The condition of occurrence of tip leakage flow unsteadiness is when the leakage flow is strong enough to reach the pressure side of the adjacent blade. The main cause of tip leakage flow unsteadiness is the tip blade loading.     

Numerical investigation on the effects of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade

This paper presents a detailed numerical investigation of the influence of re-organized shock waves on the flow separation for a highly-loaded transonic compressor cascade. The boundary layer suction (BLS) was used to control the location and strength of shock waves, with the aspirated slot locating at 49% chord, where is just downstream of the impingement point of shock wave at the leading edge. The numerical simulation is based on NUMECA, a commercial software, where the cell-centered control volume approach with third-order spatial accuracy is used to solve the 3-D Reynolds-averaged Navier-Stokes equations under the Cartesian coordinate system. Several conclusions can be made through the observation of the numerical results. (1) Multiple shock waves in cascade passage leaded the velocity deficits of boundary layer on suction surface downstream of shock wave, resulting in seriously separated flow on the suction side of blade, especially when the front shock wave is much stronger than the rest of the shocks. (2) BLS with small mass flow rate can not effectively improve the boundary layer. When the impingement point of oblique shock wave coming from cascade leading edge is bled to downstream of the passage shock wave by BLS, the boundary layer flow is greatly improved. However, if the BLS mass flow rate exceeds a critical value (1.2%), the boundary layer downstream of shock wave would separate from suction surface. (3) At the blade mid-span, the aerodynamic performance of compressor blade is improved as BLS mass flow rate increases. The optimum BLS is about 1.2%. Compared with the baseline case, the BLS with flow rate of 1.2% increases the total pressure recovery coefficient by 12%, and decreases diffusion factor by 18% and deviation angle to 7 ° while keeping the pressure rise constant. (4) The three dimensional flow structure of the compressor cascade ranged from 25% span to 75% span was improved greatly with the 1.2% BLS flow rate. However it could not control the development of the corner boundary layer effectively.     

Experimental study of effects of one kind of tangentially non-uniform tip clearance on the flow field at an exit of a compressor cascade passage

This paper presents an experimental investigation of effects of one kind of tangentially non-uniform tip clearance on the flow field at an exit of a compressor cascade passage.The tests were performed in a low-speed large-scale cascade with the uniform tip clearance and the non-uniform clearance.The three-dimensional flow field was measured at the exit at three incidence angles of 0°,5°,and 8° using a mini five-hole pressure probe.The measurement results show that the non-uniform tip clearance can moderate the leakage flow and blow down more low-energy fluids at the tip corner and decrease the accumulation of low-energy fluids which cause the flow blockage in the blade passage.In the meantime,the non-uniform clearance can weaken the tangential migration of the low-energy fluids in the endwall boundary layer and reduce the secondary loss and the flow blockage in the tip region.