Numerical and performance analysis of one row transonic rotor with sweep and lean angle

In this study,aerodynamic behaviors of swept and leaned blades were investigated.Axial and tangential blade curvatures impacts on compressor＇s operating parameters were analyzed separately.A commercial CFD program which solves the Reynolds-averaged Navier-Stokes equations was used to find out the mentioned impact and the complicated flow field of transonic compressor-rotors.The CFD method that was used for solving flow field＇s equation was validated by experimental data of NASA Rotor 67.The results showed that the compressor with curved rotors had higher efficiency,rotor pressure ratio and stable operating range compared to the compressor with un-curved rotors.Using curved rotors mostly had higher impact on the overall stable operating range compared to the other operating parameters.Operating range involves choking point and stall point that were changed separately by using of bended blade.For finding the detailed impact of sweep and lean angle on transonic blades,various forms of lean and sweep angles were exerted to basic rotor.It was found that sweep angles increased overall operating range up to 30%,efficiency up to 2%and pressure ratio up to 1%.Leaning the blades increased the safe operating range,the pressure ratio and efficiency by 14%,4%and 2%respectively.     

A Study of 3D Aerodynamic Design for a Transonic Compressor Blading Optimized by the Locations of Aerofoil Maximum Thickness and Maximum Camber

A design procedure for improving the efficiency of a transonic compressor blading was proposed based on a rapid generation method for three-dimensional blade configuration and computational meshes, a three-dimensional Navier-Stokes solver and an optimization approach. The objective of the present paper is to design a transonic compressor blading optimized only by selection of the locations of maximum camber and maximum thickness for the airfoils at different span heights and to study how do these two design parameters affect the blade performance. The blading configuration and the computational meshes can be obtained very rapidly for any given combination of maximum camber and maximum thickness. The computational grid system generated is used for the Navier-Stokes solution to predict adiabatic efficiency, total pressure ratio and flow rate. As a main result of the optimization, adiabatic efficiency was successfully improved.     

Performance Improvement of Single Screw Compressor by Meshing Clearance Adjustment Used in Refrigeration System

The single screw compressor(SSC)is widely used in air compression and refrigeration systems due to its many advantages.The meshing clearance between the screw groove and gate rotor teeth flank has a significant influence on the compressor performance.In this paper,mathematical calculation models describing the internal working process of the SSC are established in order to evaluate the thermal dynamic characteristics of the compressor under varying meshing clearance heights.The refrigerating capacity,volume efficiency and adiabatic efficiency of the SSC are calculated and discussed.Three prototypes,with different meshing clearance heights,were manufactured to study the internal influence mechanisms.The theoretical model was verified using experimental data and the calculation results strongly agreed with the experimental results.Results demonstrate that comparisons of volume efficiency and adiabatic efficiency between the measured and calculated results exhibited deviations of 3.64%-7.98%and 5.92%-9.4%,respectively.Based on the models,analysis under varying meshing clearance heights and working conditions was performed.Taking into account working performance,actual manufacturing conditions and manufacturing cost limitations,a meshing clearance height range from 0.01 mm to 0.08 mm is suggested.This study can provide important theoretical data and experimental support for the design,manufacturing and optimization of single screw compressors.     

Performance improvements of a subsonic axial-flow compressor by means of a non-axisymmetric stator hub end-wall

This paper deals with the application of a non-axisyrmnetric hub end-wall on the stator of a single stage high subsonic axial-flow compressor. In order to obtain a state-of-the-art stator non-axisymmetric hub end-wall con- figuration fulfilling the requirements for higher efficiency and total pressure ratio, an automated multi-objective optimizer was used, in conjunction with 3D-RANS-flow simulations. For the purpose of quantifying the effect of the optimal stator non axis-symmetric hub contouring on the compressor performance and its effects on the sub- sonic axial-flow compressor stator end-wall flow field structure, the coupled flow of the compressor stage with the baseline, axisymmetric and the non-axisynunetric stator hub end-wall was simulated with a state-of-the- art multi-block flow 3D CFD solver. Based on the CFD simulations, the optimal compressor hub end-wall con- figuration is expected to increase the peak efficiency by approximately 2.04 points and a slight increase of the to- tal pressure ratio. Detailed analyses of the numerical flow visualization at the hub have uncovered the different hub flow topologies between the cases with axisymmetric and non-axisymmetric hub end-walls. It was found that that the primary performance enhancement afforded by the non-axisymmelric hub end-wall is a result of the end-wall flow structure modification. Compared to the smooth wall case, the non-axisymmetric hub end-wall can reduce the formation and development of in-passage secondary flow by aerodynamic loading redistribution.     

Design and optimization of a single stage centrifugal compressor for a solar dish-Brayton system

According to the requirements of a solar dish-Brayton system,a centrifugal compressor stage with a minimum total pressure ratio of 5,an adiabatic efficiency above 75% and a surge margin more than 12% needs to be designed.A single stage,which consists of impeller,radial vaned diffuser,90° crossover and two rows of axial stators,was chosen to satisfy this system.To achieve the stage performance,an impeller with a 6:1 total pressure ratio and an adiabatic efficiency of 90% was designed and its preliminary geometry came from an in-house one-dimensional program.Radial vaned diffuser was applied downstream of the impeller.Two rows of axial stators after 90° crossover were added to guide the flow into axial direction.Since jet-wake flow,shockwave and boundary layer separation coexisted in the impeller-diffuser region,optimization on the radius ratio of radial diffuser vane inlet to impeller exit,diffuser vane inlet blade angle and number of diffuser vanes was carried out at design point.Finally,an optimized centrifugal compressor stage fulfilled the high expectations and presented proper performance.Numerical simulation showed that at design point the stage adiabatic efficiency was 79.93% and the total pressure ratio was 5.6.The surge margin was 15%.The performance map including 80%,90% and 100% design speed was also presented.     

Stall inception process and prospects for active hub-flap control in three-stage axial flow compressor

The possibility to apply the active hub-flap control method, which is a proven rotating stall control method for a single-stage compressor, to a 3-stage axial compressor is experimentally discussed, where complex rotating stall inception processes are observed. The research compressor is a 3-stage one and could change the stagger angle settings for rotor blades and stator vanes. Sixteen rotor blade/stator vane configuration patterns were tested by changing stagger angle for the stator vanes. By measurement of surface-pressure fluctuation, stall inception proc- esses are investigated and the measured pressure fluctuation data is used as a predictive signal for rotating stall. The experimental results show that the stall detection system applied to active hub-flap control in a single-stage compressor could be usefully applied to that in a 3-stage compressor with a more complex stall inception process.     

Computational and Experimental Study of Pinch on the Performance of a Vaneless Diffuser in a Centrifugal Compressor

This study focuses on the vaneless diffuser of a centrifugal compressor.The examined stage consists of an un-shrouded impeller,a parallel wall vaneless diffuser and a volute.The walls of the diffuser were movable allowingdifferent pinch configurations to be investigated.The baseline geometry had no pinch i.e.the height of the dif-fuser was equal to the height of the impeller flow channel plus the axial running clearance.The work consists ofboth numerical and experimental parts.Quasi-steady, turbulent,fully 3D numerical simulations were conducted.The inlet cone,rotor and diffuser were modelled.Six different configurations were studied.The height of thepinch was altered and the pinch made to different walls was tested.Two of the numerically studied cases werealso experimentally investigated.The overall performance of the compressor,the circumferential static and totalpressure and the spanwise total pressure distribution before and after the diffuser were measured.The numericaland experimental studies showed that the pinch improved the efficiency of the compressor.     

Impact of Stagger Angle Nonuniformity on Turbine Aerodynamic Performance

The assembling error may lead to variation in stagger angles, which would affect the aerodynamic performance of the turbine. To investigate this underlying effect, two parallel numerical experiments on two turbines with the same profile, but uniform and nonuniform vane stagger angle respectively, were conducted in both steady and un- steady methods. The results indicate that certain changes in the detailed flow field of the turbine occur when the stagger angles are nonuniform, further, the blade loading distribution of the vane and rotor become markedly dif- ferent from that in uniform vane stagger angle situation. Then these consequences caused by nonuniformity men- tioned above enhance the unsteadiness of the flow, finally, the aerodynamic performance changes dramatically. It also shows that, compared with steady simulation, the unsteady numerical simulation is necessary in this investigation     

Numerical modelling of a supersonic axial turbine stator

In order to improve the turbocharging process,a supersonic axial turbine stator was modelled
numerically with a pulsatile inlet mass flow.The main objectives of the study were to find out how
pulsation affects the flow field and the performance of the stator.At the beginning of the study,a
supersonic turbine stator was modelled using three different techniques：quasi-steady,time-accurate
with constant boundary conditions and time-accurate with a pulsatile inlet mass flow.The time-
averaged and quasi-steady flow fields and performance were compared,and the flow field and stator
performance with a pulsatile inlet mass flow was studied in detail at different time-steps.A hysteresis-
like behaviour was captured when the total-to-static pressure ratio and efficiency were plotted as a
function of the inlet mass flow over one pulse period.The total-to-static pressure ratio and efficiency
followed the sinusoidal shape of the inlet flow as a function of time.It was also concluded that the
stator efficiency decreases downstream from the stator trailing edge and the amplitude of the
pulsating mass flow is decreased at the stator throat.     

Numerical investigation of a high-subsonic axial-flow compressor rotor with non-axisymmetric hub endwall

The major source of loss in modem compressors is the secondary loss. Non-axisymmetric endwall profile contouring is now a well established design methodology in axial flow turbines. However, flow development in axial compressors is differ from turbines, the effects of non-axisymmetric endwall to axial compressors requires flow analysis in detail. This paper presents both experimental and numerical data to deal with the application of a non-axisymmetric hub endwall in a high-subsonic axial-flow compressor. The aims of the experiment here were to make sure the numerically obtained flow fields is the physical mechanism responsible for the improvement in efficiency, due to the non-axisymmetric hub endwall. The computational results were first compared with avail- able measured data of axisymmetric hub endwall. The results agreed well with the experimental data for estima- tion of the global performance. The coupled flow of the compressor rotor with non-axisymmetric hub endwall was simulated by a state-of-the-art multi-block flow solver. The non-axisymmetric hub endwall was designed for a subsonic compressor rotor with the help of sine and cosine functions. This type of non-axisymmetric hub end- wall was found to have a significant improvement in efficiency of 0.45% approximately and a slightly increase for the total pressure ratio. The fundamental mechanisms of non-axisymmetric hub endwall and their effects on the subsonic axial-flow compressor endwall flow field were analyzed in detail. It is concluded that the non-axisymmetric endwall profiling, though not optimum, can mitigate the secondary flow in the vicinity of the hub endwall, resulting in the improvement of aerodynamic performance of the compressor rotor.     

Automatic efficiency optimization of an axial compressor with adjustable inlet guide vanes

The inlet attack angle of rotor blade reasonably can be adjusted with the change of the stagger angle of inlet guide vane (IGV); so the efficiency of each condition will be affected. For the purpose to improve the efficiency, the DSP (Digital Signal Processor) controller is designed to adjust the stagger angle of IGV automatically in order to optimize the efficiency at any operating condition. The A/D signal collection includes inlet static pressure, outlet static pressure, outlet total pressure, rotor speed and torque signal, the efficiency can be calculated in the DSP, and the angle signal for the stepping motor which control the IGV will be sent out from the D/A. Experimental investigations are performed in a three-stage, low-speed axial compressor with variable inlet guide vanes. It is demonstrated that the DSP designed can well adjust the stagger angle of IGV online, the efficiency under different conditions can be optimized. This establishment of DSP online adjustment scheme may provide a practical solution for improving performance of multi-stage axial flow compressor when its operating condition is varied.     

Investigation of a centrifugal compressor and study of the area ratio and TIP clearance effects on performance

In this research, the centrifugal compressor of a turbocharger is investigated experimentally and numerically. Performance characteristics of the compressor were obtained experimentally by measurements of rotor speed and flow parameters at the inlet and outlet of the compressor. Three dimensional flow field in the impeller and dif- fuser was analyzed numerically using a full Navier-Stokes program with SST turbulence model. The performance characteristics of the compressor were obtained numerically, which were then compared with the experimental results. The comparison shows good agreement. Furthermore, the effect of area ratio and tip clearance on the performance parameters and flow field was stud- ied numerically. The impeller area ratio was changed by cutting the impeller exit axial width from an initial value of 4.1 mm to a final value of 5.1 mm, resulting in an area ratio from 0.792 to 0.965. For the rotor with exit axial width of 4.6 mm, performance was investigated for tip clearance of 0.0, 0.5 and 1.0 mm. Results of this simula- tion at design point showed that the compressor pressure ratio peaked at an area ratio of 0.792 while the effi- ciency peaked at a higher value of area ratio of 0.878. Also the increment of the tip clearance from 0 to 1 mm resulted in 20 percent efficiency decrease.     

Clocking Effect in a Two-Stage Compressor with Different Inter-Blade-Row Gaps

A multi-stage axial compressor has inherently unsteady flow fields because of the following main reasons: (1) relative positions between rotor and stator airfoil; (2) the buildup of converted wakes lead to complex wake/wake and wake/airfoil interactions. The distributions of the potential flows and wakes in the flow passage are depended on the relative positions of blade rows in axial and circumference direction, so variations in the relative axial positions (axial gap) and circumferential positions (clocking effect) of stators or rotors can change these distributions, leading to different compressor efficiency. The current study presents the experimental/numerical result of a low-speed axial compressor, considering the combined effects of stator clocking and variation of axial gaps. Testing was conducted in Two-Stage Axial Compressor Facility in Harbin Institute of Technology. In the test, time averaged data were collected. The results of experimental and time accurate flow calculation for 2 axial gaps, 8 clocking positions for each gap are compared. It is shown that clocking determines the degree of interaction of a stator with the wake of another upstream stator for different gaps between the blade rows.     

The effect of variable stator on performance of a highly loaded tandem axial flow compressor stage

Increasing the aerodynamic load on compressor blades helps to obtain a higher pressure ratio in lower rotational speeds.Considering the high aerodynamic load effects and structural concerns in the design process,it is possible to obtain higher pressure ratios compared to conventional compressors.However,it must be noted that imposing higher aerodynamic loads results in higher loss coefficients and deteriorates the overall performance.To avoid the loss increase,the boundary layer quality must be studied carefully over the blade suction surface.Employment of advanced shaped airfoils (like CDAs),slotted blades or other boundary layer control methods has helped the designers to use higher aerodynamic loads on compressor blades.Tandem cascade is a passive boundary layer control method,which is based on using the flow momentum to control the boundary layer on the suction surface and also to avoid the probable separation caused by higher aerodynamic loads.In fact,the front pressure side flow momentum helps to compensate the positive pressure gradient over the aft blade's suction side.Also,in comparison to the single blade stators,tandem variable stators have more degrees of freedom,and this issue increases the possibility of finding enhanced conditions in the compressor off-design performance.In the current study,a 3D design procedure for an axial flow tandem compressor stage has been applied to design a highly loaded stage.Following,this design is numerically investigated using a CFD code and the stage characteristic map is reported.Also,the effect of various stator stagger angles on the compressor performance and especially on the compressor surge margin has been discussed.To validate the CFD method,another known compressor stage is presented and its performance is numerically investigated and the results are compared with available experimental results.     

Three Dimensional Numerical Investigation on the Effect of Wedge Angle of a Passage with Pin-fin Arrays

Heat transfer in passage with pin-fin arrays for cooling blade trailing edge was studied numerically. Three-dimensional numerical simulations were carried out for steady laminar flow in passages with different wedge angles between pressure surface and suction surface of cooling blade trailing edge to study the effect of different wedge angles (from 0°to 30°) on heat transfer and pressure losses. Research was carried out for both in-line array and staggered array. From this investigation, wedge angle 10°gives the best heat transfer performance.     

Aerodynamic Characteristic Research on Final Stage Stator of a Highly Loaded Fan

Due to corner separation and other complex three-dimensional flows existing in the highly loaded stator, which influences the fan performance significantly, highly loaded stator blades of a transonic fan with a maximum camber angle of 57° were studied in this paper and sector cascade experiment was adopted. In order to get the stator aerodynamic parameters as realistic as possible and conduct the experiment without the existence of rotor, an adjustable guide vane was designed to simulate the velocity magnitude and direction of the stator inlet flow. Results show that the adjustable guide vane can simulate the rotor outlet velocity direction and magnitude in most span range. The deviation angle is positive and the maximum value is nearly 21° because the severe separation is at 27% span. Corner separation exists on both pressure side and suction side and the location of separation initiation is determined. Finally, the stator blades were redesigned with some suction slots on the suction side. Experiment results show that the suction slots change the flow field structure, increase the capability of flow turning, and decrease the flow loss.     

Numerical Investigation of the Superimposed Effects on Stator Wake Oscillation in an Axial-radial Combined Compressor ZHAO Ben,YANG Ce,HU Liangjun,ZHOU Mi and ZHANG Jizhong简

The superimposed influences of the blade rows in a multistage compressor are important because different matches of upstream and downstream blades can result in significant differences in the stator wake oscillation. Numerical inves- tigation of the axial stator wake oscillation, which is affected upstream by the axial rotor and downstream by the radial rotor, was performed in an axial-radial combined compressor. Many configurations with different blade numbers and locations, which influence axial stator wake oscillation were investigated. When rotors have equal blade numbers, the axial stator wake oscillates periodically versus time within time T （moving blade passing 1/3 revolution）. In contrast, stator wake oscillates irregularly within T when rotors have different blade numbers. A model-split subtraction method is presented in order to separate the influences of the individual blade rows on the wake oscillation of the axial stator. Analysis from the rotor-stator configuration showed that the unsteady flow angle fluctuation response is caused by the upstream rotor. For the rotor-stator-rotor configuration, the unsteady flow angle fluctuations are influenced by up- and downstream blade rows. With the model-split subtraction method, the up- and downstream influences on the flow angle fluctuation could be clearly separated and quantified. Low amplitudes could be observed when the influences from up- and downstream moving rows were superimposed with the ＂positive peak- negative peak＂ type wave. Clocking investi- gations were carried out to change the relative superimposed phase of influences from the surrounding blade rows in or- der to modulate the amplitudes of the axial stator wake oscillation. However, the amplitudes did not reach the maximum when they were superimposed with ＂positive peak-positive peak＂ type wave due to the impact of the interaction between the two moving blade rows.     

IE3 Efficiency Induction Motors with Aluminum and Copper Rotor Cage： Technical and Economic Comparison

The aim of this study was to design three-phase induction motors with aluminum and copper cage, in the range 0.75 ÷22 kW, to fulfill the 1E3 efficiency level according to typical performance and standard constraints. The proposed study has concerned TEFC （ totally Enclosed Fan-Cooled ）, 400 V, 50 Hz, SI duty three phase squirrel-cage induction motors only. The motors＇ designs, with AI and Cu cage, have been optimized in order to reach the minimum efficiency level IE3 at lowest active material costs and satisfy the physical and performance constraints of the designs, which are the motor specifications. A suitable optimization procedure has been used which allowed to find the ＂best design＂ by chancing the geometric dimensions of the stator, rotor shape, the stator winding and the stack length. In order to guarantee the goodness and feasibility of the optimized designs, several constrains have been imposed.     

The influence of wedge diffuser blade number and divergence angle on the performance of a high pressure ratio centrifugal compressor

Wedge diffuser is widely used in centrifugal compressors due to its high performance and compact size. This paper is aimed to research the influence of wedge diffuser blade number and divergence angle on centrifugal compressor performance. The impact of wedge diffuser blade number on compressor stage performance is investigated, and then the wedge diffusers with different divergence angle are studied by varying diffuser wedge angle and blade number simultaneously. It is found that wedge diffuser with 27 blades could have about 0.8% higher adiabatic efficiency and 0.14 higher total pressure ratio than the wedge diffuser with 19 blades and the best compressor performance is achieved when diffuser divergence angle is 8.3°.These results could give some advices on centrifugal compressor design.     

Design of a Centrifugal Compressor with Low Solidity Vaned Diffuser (LSVD) for Large-Scale Compressed Air Energy Storage (CAES)

Compressed Air Energy Storage(CAES) has tremendous promotional value in the intermittent renewable energy supply systems. CAES has special requirements for compressor(e.g. heavy load, high pressure ratio, wide range). With advantages of higher efficiency and wider operation range, IGC(Integrally Geared Compressors) is selected to fulfill the special requirements of the large-scale CAES. To get a better aerodynamic performance, in this paper, based on the analysis of internal flow of centrifugal compressor, a multi-objective one-dimensional optimization design program was put forward combined with modified Two-Zone model and a low solidity vaned diffuser(LSVD) design method. Then, a centrifugal compressor aerodynamic component optimization design system was established with the three-dimensional blade optimization design method based on neural network and genetic optimization algorithm. Then a validation was done by redesigning the Krain-Impeller to get better performance. Finally, the aerodynamic design of the first stage of IGC was completed. The CFD calculation results indicated that the total-to-total pressure ratio of the first stage was 2.51 and the polytropic efficiency was 91.0% at the design point. What’s more, an operation margin and surge margin of the compressor was about 26.5% and 16.4% respectively.