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The kinetic energy loss and the secondary loss downstream of a high turning nozzle cascade were investigated. Influence of the incidence angle (i= −25°, 0° and 25°) and Reynolds number (Re2,is =0.8×105, 1.2×105, 2.5×105 and 4.5×105) were studied on two cascade configurations defined by different pitch to chord ratio (t/c=0.6 and 0.9), respectively. Measurements at the cascade inlet were performed by means of Preston tube, while at the cascade outlet by means of miniature pyramid five-hole pressure probe. An effect of studied parameters on the flow field at the cascade outlet was shown by the distribution of the kinetic energy loss, secondary loss, mix-out loss and streamwise vorticity.

As an effective method to influence end wall flow field, non-axisymmetric profiled end wall can improve the aerodynamic performance of compressor cascades. For a highly loaded low pressure compressor cascade, called V103, the study found the optimal non-axisymmetric profiled end wall decreases total pressure loss coefficient by 4.57%, 5.48% and 3.04% under incidences of −3°, 0°, and 3°, respectively, compared with those of the planar end wall. The optimal non-axisymmetric profiled end wall changes the structure of secondary flow in hub region, generating a corner vortex near suction surface, inhibiting the development of the passage vortex towards suction surface and reducing flow separation. When the inlet Mach numbers are 0.62 and 0.72, the total pressure loss coefficient decreases by 3.19% and 4.58% for optimal non-axisymmetric profiled end wall compared with those of the planar end wall. Though optimal non-axisymmetric profiled end wall increases total pressure loss near hub region in blade passage under different inlet Mach numbers, the peak value and region of high loss coefficient above 10% span in blade passage significantly decrease. In addition, different incidences affect the secondary flow streamlines and vortex structure near the cascade hub region, however, different inlet Mach numbers hardly change the secondary flow streamlines and vortex structure. In short, the optimal non-axisymmetric profiled end wall shows better aerodynamic performance than the planar end wall for the highly loaded compressor cascade in multi-conditions.

In modern high-load high-pressure turbine, the secondary flow in the blade channel is very strong and occupies a large spanwise region. Although high-quality experimental data at the stage interfaces have been obtained in previous research, the influence of the clocking position on the secondary flow patterns is not fully understood. This paper investigates the clocking effect in a 1.5-stage high-pressure turbine and focuses on the variations of secondary flow patterns and their effect on the turbine performance. The detailed flow fields of various clocking positions were obtained by carrying out unsteady flow simulations using an in-house code. Among the four clocking positions in this work, the highest entropy generation was observed when the wakes from stator 1 hit the leading edges of stator 2, which is opposite to the well-known conclusion for the turbine with high-aspect-ratio blades. Detailed flow analysis showed that the wakes and the near tip secondary vortices from stator 1 showed different traces when entering the stator 2 channel and the secondary vortices clearly have a more important influence in determining the performance. The different behaviors of the secondary vortices explained the performance variations due to the clocking effect.

In the current study, thermal boundary conditions are considered in a rotating smooth channel with a square cross-section to investigate the secondary flow and compare it to that of the same channel without heating. The measurement is conducted at three streamwise planes (X=445 mm, 525 mm, 605 mm). The flow parameters are the Reynolds number (Re=4750, which was based on the average longitudinal or primary velocity U and the hydraulic diameter D of the channel cross-section), the rotation number (Ro=ΩD/U, where Ω is the rotational velocity, ranging from 0 to 0.26), and the aspect ratio of the channel cross-section (AR=1, which is calculated by dividing the channel height by the channel width). The leading and trailing walls are heated under a constant heat flux qw=380 W/m2, and the top and bottom walls are isothermal at room temperature. This work is in a series with our previous work without thermal boundary conditions. Based on the experimental data, we obtained a four-vortex regime. There is a counter-rotating vortex pair near the leading side and the trailing side. Because the leading and trailing walls are heated, the buoyancy force increases the relative vertical position of the vortex pair near the trailing side from 5% to 12.5% of the hydraulic diameter. When moving upstream along the streamwise direction, the upper vortex near the trailing wall becomes weaker, whereas the lower vortex becomes stronger. As the rotational speed increases, the vortex pair near the trailing side is inhibited by the Coriolis force. Under heated thermal boundary conditions, the vortex pair near the trailing side reappears due to the effect of buoyancy force. These results indicate that the buoyancy force has a substantial effect on the secondary flow regime under thermal boundary conditions.

For the purpose of improving turbine efficiency, a detailed investigation of flow field near rim seal region in the main annulus is absolutely essential. Purge flow egressed from the seal gap penetrates the hub boundary layer and travels along with the secondary flow in the main path. The propagation of the purge flow and its influence on the turbine performance are crucial and cannot be neglected in the aerodynamic and thermal design. In addition, the interaction of vane wake and blade potential field also has a significant impact on the turbine stage performance. In this study, flow field between inlet guide vanes (IGV) and rotor blades was experimentally investigated at specific positions with the help of a two-dimensional particle image velocimetry (2-D PIV). A phase-locking technique was also employed to capture the images of flow field at different rotor blade positions. Data acquired characterizes the impact of emergent purge flow on the flow structures within the main annulus. The impacts of vane wake and blade potential field on the flow fields were analyzed through characterizations provided by the time-averaged field results. Finally, flow fields at different main flow rates and under different seal gap widths were also presented and assessed.

This paper investigates the flow past two variants of root section profile cascades for a last stage rotor considering three-dimensional flow structures in the near-wall region. Analyses were drawn based on RANS numerical simulations of both variants and on the experimental data obtained by the 3D traversing in the exit flow field of one of the variants. Extent of 3D structures at two different regimes and its influence on aerodynamic characteristics of the blade cascades was assessed. The distributions of Mach number along the profiles were compared with 2D optical measurements and its distortion due to the presence of the sidewall was explored. The interaction between main vortical structures was described and its influence on the loading of the blades, mechanical energy losses and exit flow angle was discussed. The results showed that for a front loaded blade the vortical structures appeared earlier and at a larger extent than for an aft loaded variant. However, due to different Mach number distribution, contribution of end wall flow to the energy losses was lower in the case of the aft loaded variant. The influence of the near wall flow on the loading was found to be rather weak while the deviation of the exit flow angle appeared to be comparable for both of the variants.

Ejector refrigeration has the advantage of low capital cost, simple design, reliable operation, long lifespan and almost no maintenance. The only weakness is the low efficiency and its intolerance to deviations from design operation condition. R134a used in ejector refrigeration system gives better performance in comparison with many other environmental friendly refrigerants as the generation temperature is from 75°C to 80°C. The present work experimentally investigated the on-design and off-design performance of the ejector with fixed geometry using R134a as refrigerant, and cycle performance of the ejector refrigeration system. The experimental prototype was constructed and the effects of primary flow inlet pressure, secondary flow inlet pressure and ejector back pressure on ejector performance and cycle performance were investigated respectively. The operation conditions are: primary flow inlet pressure from 2.2 MPa to 3.25 MPa, secondary flow inlet pressure from 0.36 MPa to 0.51 MPa, ejector back pressure from 0.45 MPa to 0.67 MPa. Conclusions were drawn from the experimental results, and the experimental data can be used for validation of theoretical model for both critical and subcritical mode.

In order to further understand the aerothermodynamic performance and flow loss mechanism of SCO2 turbines, RANS equations and an SST Turbulence Model were chosen for a numerical study on the secondary flow and vortex structure of cascades using the commercial software CFX. The dimensionless vorticity analysis method was used to study the flow characteristics of the logarithmic layer and viscous sublayer in high pressure guide vane cascades. The new vortex structure and formation mechanism of the vortices were given and analyzed. Simulation results indicated that during the motion of the boundary layer in the cascades, the logarithmic layer and viscous sublayer obtain the different rotational direction vorticity, respectively. The endwall logarithmic layer and pressure side leg of the horseshoe votex gradually develop into the passage vortex, with the endwall viscous sublayer gradually developing into the corner viscous sublayer vortex II and the endwall viscous sublayer vortex I. The endwall viscous sublayer that rolled by the passage vortex is encountered with the upstream-side and radial boundary layer of the vane at the suction separation line, forming the suction separation line vortex beside the passage vortex. A pair of radial transition vortices are formed between the wake and the main stream.

In nuclear reactor fuel assemblies, spacer grids are installed among the rod bundles to support the fuel rods and affect the flow field between rods. Mixing vanes, as a swirling device, are set on the upper apex of the spacer grid. Vortexes produced by mixing vanes move along the axial direction in subchannels and enhance the forced convection heat transfer between the rods and cooling-fluid medium. In this paper, a numerical simulation method was used to investigate vortex motion produced by typical AFA-3G spacer grids in a 5×5-rod bundle by Star-CCM+ software. The shear-stress transport k-ω model was used to simulate turbulence phenomena. A dimensionless parameter, Se, based on the absolute vorticity flux, was reported to specify the intensity of secondary flow. Its physical meaning is the ratio of inertial force to viscous force induced by secondary flow. The results are helpful to take advantage of spacer grids in a much more effective way in pressurized water reactors.

A three-dimensional blade of a low solidity circular cascade diffuser in centrifugal blowers is designed by means of a multi-point optimization technique. The optimization aims at improving static pressure coefficient at a design point and at a small flow rate condition. Moreover, a clear definition of secondary flow expressed by positive radial velocity at hub side is taken into consideration in constraints. The number of design parameters for three-dimensional blade reaches to 10 in this study, such as a radial gap, a radial chord length and mean camber angle distribution of the LSD blade with five control points, control point between hub and shroud with two design freedom. Optimization results show clear Pareto front and selected optimum design shows good improvement of pressure rise in diffuser at small flow rate conditions. It is found that three-dimensional blade has advantage to stabilize the secondary flow effect with improving pressure recovery of the low solidity circular cascade diffuser.

The self-induced unsteadiness in tip leakage flow (TLF) of a micro-axial fan rotor is numerically studied by solving Reynolds-averaged Navier-Stokes equations. The micro-axial fan, which is widely used in cooling systems of electronic devices, has a tip clearance of 6% of the axial chord length of the blade. At the design rotation speed, four cases near the peak efficiency point (PEP) with self-induced unsteadiness and four steady cases which have much weaker pressure fluctuations are investigated.

Two examples of the use of vortex control to reduce noise and enhance the stable operating range of a centrifugal compressor are presented in this paper. In the case of high-flow operation of a centrifugal compressor with a vaned diffuser, a discrete frequency noise induced by interaction between the impeller-discharge flow and the diffuser vane, which appears most notably in the power spectra of the radiated noise, can be reduced using a tapered diffuser vane (TDV) without affecting the performance of the compressor. Twin longitudinal vortices produced by leakage flow passing through the tapered portion of the diffuser vane induce secondary flow in the direction of the blade surface and prevent flow separation from the leading edge of the diffuser. The use of a TDV can effectively reduce both the discrete frequency noise generated by the interaction between the impeller-discharge flow and the diffuser surface and the broadband turbulent noise component. In the case of low-flow operation, a leading-edge vortex (LEV) that forms on the shroud side of the suction surface near the leading edge of the diffuser increases significantly in size and blocks flow in the diffuser passage. The formation of an LEV may adversely affect the performance of the compressor and may cause the diffuser to stall. Using a one-side tapered diffuser vane to suppress the evolution of an LEV, the stable operating range of the compressor can be increased by more than 12 percent, and the pressure-rise characteristics of the compressor can be improved. The results of a supplementary examination of the structure and unsteady behavior of LEVs, conducted by means of detailed numerical simulations, are also presented.

An experimental study is conducted to investigate the influences of blade tip winglet on the flow field of a compressor cascade. The tests are performed in a low speed linear cascade with stationary endwall, with three blade tip configurations, including the baseline tip, the suction-side winglet tip and the pressure-side winglet tip. The flowfield downstream of the cascade is measured using five-hole probe, from which the three-dimensional velocity field, vorticity field and pressure field are obtained. Static pressure measurements are made on the endwall above the blade row using pressure taps embedded in the plywood endwall. All measurements are made at both design and off-design conditions for tip clearance level of about 2 percent of the blade chord. The results revealed the incidence variation significantly affects the secondary flow and the associated loss field downstream of the cascade, where the tip leakage vortex and passage vortex exist as the major contributors on the field. The winglet geometry arrangements can change the trajectory of the tip leakage vortex. The suction-side winglet tip blade provides a lower overall total pressure loss coefficient when compared to the baseline tip blade and pressure-side winglet tip blade at all incidence angles.

High-pressure ratio and wide operating range are highly required for compressors and blowers. The technical issue of the design is achievement of suppression of flow separation at small flow rate without deteriorating the efficiency at design flow rate. A numerical simulation is very effective in design procedure, however, cost of the numerical simulation is generally high during the practical design process, and it is difficult to confirm the optimal design which is combined with many parameters. A multi-objective optimization technique is the idea that has been proposed for solving the problem in practical design process. In this study, a Low Solidity circular cascade Diffuser (LSD) in a centrifugal blower is successfully designed by means of multi-objective optimization technique. An optimization code with a meta-model assisted evolutionary algorithm is used with a commercial CFD code ANSYS-CFX. The optimization is aiming at improving the static pressure coefficient at design point and at low flow rate condition while constraining the slope of the lift coefficient curve. Moreover, a small tip clearance of the LSD blade was applied in order to activate and to stabilize the secondary flow effect at small flow rate condition. The optimized LSD blade has an extended operating range of 114 % towards smaller flow rate as compared to the baseline design without deteriorating the diffuser pressure recovery at design point. The diffuser pressure rise and operating flow range of the optimized LSD blade are experimentally verified by overall performance test. The detailed flow in the diffuser is also confirmed by means of a Particle Image Velocimeter. Secondary flow is clearly captured by PIV and it spreads to the whole area of LSD blade pitch. It is found that the optimized LSD blade shows good improvement of the blade loading in the whole operating range, while at small flow rate the flow separation on the LSD blade has been successfully suppressed by the secondary flow effect.

An increase of turbine blade loading can reduce the numbers of blade and stage of gas turbines. However, an increase of blade loading makes the secondary flow much stronger because of the steep pitch-wise pressure gradient in the cascade passage, and consequently deteriorates the turbine efficiency. In this study, the computations were performed for the flow in an ultra-highly loaded turbine cascade with high turning angle in order to clarify the effects of the incidence angle on the two dimensional flow and the secondary flow in the cascade passage, which cause the profile loss and the secondary loss, respectively. The computed results showed good agreement with the experimental surface oil flow visualizations and the blade surface static pressure at mid-span of the blade. The profile loss was strongly increased by the increase of incidence angle especially in the positive range. Moreover, the positive incidences not only strengthened the horseshoe vortex and the passage vortex but also induced a new vortex on the end-wall. Moreover, the newly formed vortex influenced the formation of the pressure side leg of horseshoe vortex.

This paper deals with the application of a non-axisymmetric 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 configuration 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 subsonic axial-flow compressor stator end-wall flow field structure, the coupled flow of the compressor stage with the baseline, axisymmetric and the non-axisymmetric stator hub end-wall was simulated with a state-of-theart multi-block flow 3D CFD solver. Based on the CFD simulations, the optimal compressor hub end-wall configuration is expected to increase the peak efficiency by approximately 2.04 points and a slight increase of the total 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-axisymmetric 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.

Generally speaking, main flow path of gas turbine is assumed to be perfect for standard 3D computation. But in real engine, the turbine annulus geometry is not completely smooth for the presence of the shroud and associated cavity near the end wall. Besides, shroud leakage flow is one of the dominant sources of secondary flow in turbomachinery, which not only causes a deterioration of useful work but also a penalty on turbine efficiency. It has been found that neglect shroud leakage flow makes the computed velocity profiles and loss distribution significantly different to those measured. Even so, the influence of shroud leakage flow is seldom taken into consideration during the routine of turbine design due to insufficient understanding of its impact on end wall flows and turbine performance. In order to evaluate the impact of tip shroud geometry on turbine performance, a 3D computational investigation for 1.5-stage turbine with shrouded blades was performed in this paper. The following geometry parameters were varied respectively: The aim of this paper is to isolate the influence of shroud and cavity geometry modifications on turbine aerodynamicnamic performance and to obtain clear trends of efficiency changes caused by different tip shroud geometry. Moreover, interaction between leakage flow and mainstream for different shroud configuration is also highlighted in order to penetrate into the physical mechanisms producing them. Due to the limitations of the model selected in this paper, the aim of research is not to put forward the design rules of turbine shroud. However, the results obtained from this work will be useful to the integrated design and optimization of turbine with shrouded blades.

As a high specific speed pump, the contra-rotating axial flow pump distinguishes itself in a rear rotor rotating in the opposite direction of the front rotor, which remarkably contributes to the energy conversion, the reduction of the pump size, better hydraulic and cavitation performances. However, with two rotors rotating reversely, the significant interaction between blade rows was observed in our prototype contra-rotating rotors, which highly affected the pump performance compared with the conventional axial flow pumps. Consequently, a new type of rear rotor was designed by the rotational speed optimization methodology with some additional considerations, aiming at better cavitation performance, the reduction of blade rows interaction and the secondary flow suppression. The new rear rotor showed a satisfactory performance at the design flow rate but an unfavorable positive slope of the head — flow rate curve in the partial flow rate range less than 40% of the design flow rate, which should be avoided for the reliability of pump-pipe systems. In the present research, to understand the internal flow field of new rear rotor and its relation to the performances at the partial flow rates, the velocity distributions at the inlets and outlets of the rotors are firstly investigated. Then, the boundary layer flows on rotor surfaces, which clearly reflect the secondary flow inside the rotors, are analyzed through the limiting streamline observations using the multi-color oil-film method. Finally, the unsteady numerical simulations are carried out to understand the complicated internal flow structures in the rotors.

In order to examine the fundamental characteristics of corner separation in a decelerating cascade flow,the experimental apparatus was made and separations around a NACA65 blade in a decelerating channel flow were examined.Two-dimensional calculations show that the NACA65 cascade flow that has 45 deg.of turning,1.24 of solidity and 17 deg.of stagger angle is equivalent to the channel flow that has 14 deg.of stagger angle in terms of pitchwise blade force.Experimental investigation by five-hole probe shows that the accumulations of low energy fluid can be seen around the corner part and the overturning flow due to the secondary flow exists.And,as the periodicity of the blade wake in a pitchwise direction is comparably good,the wake of this channel flow is similar to the wake of the cascade flow for two pitch portion.PIV measurement results shows that a vortex pattern can be seen in the momentary streamline on the suction surface of the blade and in the averaged streamline on the perpendicular to both the suction surface and the endwall.

Since the automobile turbochargers are installed in an engine compartment with limited space, the ducts upstream of the turbocharger compressor may be curved in a complex manner. In the present paper, the effect of a curved duct upstream on performance of small centrifugal compressors for automobile turbochargers is discussed. The computational fluid dynamics (CFD) analysis of a turbocharger compressor validated for the compressor model with the straight pipe applied to the compressor with the curved pipe are executed, and the deterioration of the performance for the curved pipe is confirmed. It is also found that the deterioration of compressor performance is caused by the interaction of the secondary flow and the impeller.

The aerodynamic flow field downstream of a Low-Pressure High-Lift (HL) turbine cascade has been experimentally investigated for different Reynolds numbers under both steady and unsteady inflows, in order to analyse the cascade performance under real engine operating conditions. The Reynolds number has been varied in the range 100000

Taking into account stealth, structure and maintenance, ultra-compact S-shaped intake has been widely used in modern vehicles. In this paper a series of steady numerical simulation were carried out to investigate the complex flow mechanism in s-shaped intake, particular attention was given to examining the effect of centerline curvature on the performance and flowfield of an ultra-compact S-shaped intake. In order to validate the multi-block model, the computational results for Royal Aircraft Establishment intake 2129-M2129 which had modest centerline curvature distribution were corrected with available experimental test data. The numerical simulation results agreed fairly well with the experimental data, and the computational method was then used to investigate the effects of different centerline curvature distributions on performance and flow field in compact S-shaped intake. Detailed analyses of the flow visualization had exposed the different flow topologies between the cases with different centerline curvature. It was found that different centerline curvature distributions changed the initial location and the size of separation bubble, as well as the strength of two counter-rotating vortices at the entrance of engine.

According to the previous experimental works on the low solidity circular cascade diffuser (LSD), a pressure recovery of a centrifugal blower was improved by the LSD significantly in a wide range of flow rate, and the pressure recovery was improved further by the LSD with a tandem cascade in comparison with the LSD with a single- row cascade. In the present study, the flow behavior in the LSD with the tandem cascade has been analyzed numerically by using the commercial CFD code of ANSYS-CFX12. It was shown clearly that the higher pressure recovery was achieved by applying the LSD with the tandem cascade, and the high pressure recovery is based on the high pressure rise in the vaneless space upstream of the LSD and the high blade loading of the front blade of the LSD. The high pressure recovery in the LSD could be achieved by controlling the flow separation on the suction surface of the front blade and also on that of the rear blade due to formation of the favorable secondary flow and due to increase in mass flow passing through the slit section between the front and rear blades., Journal of Thermal Science, 21(4), pp.384-390; 2012

High flow rate aeroengines typically employ axial flow compressors, where aerodynamic loss is predominantly due to secondary flow features such as tip leakage and corner vortices. In very high altitude missions, turbomachinery operates at low density ambient atmosphere, and the recent trend toward more compact engine core inevitably leads to the reduction of blade size, which in turn increases the relative height of the blade tip clearance. Low Reynolds number flowfield as a result of these two factors amplifies the relative importance of secondary flow effects. This paper focuses on the behavior of tip leakage flow, investigating by use of both experimental and numerical approaches. In order to understand the complex secondary flow behavior, cascade tests are usually conducted using intrusive probes to determine the loss. However relatively few experimental studies are published on tip leakage flows which take into account the interaction between a rotating blade row and its casing wall. Hence a new linear cascade facility has been designed with a moving belt casing in order to reproduce more realistic flowfield as encountered by a rotating compressor row. Numerical simulations were also performed to aid in the understanding of the complex flow features. The experimental results indicate a significant difference in the flowfield when the moving belt casing is present. The numerical simulations reveal that the leakage vortex is pulled by the shearing motion of the endwall toward the pressure side of the adjacent blade. The results highlight the importance of casing wall relative motion in analyzing leakage flow effects.

The effect of the secondary flow on the starting pressure of a second-throat supersonic ejector has been investigated by adapting the height of the secondary flow inlet. The obtained results show that an optimum value of the secondary inlet height exists, and the starting pressure of the ejector becomes a minimum at that condition. Based on the results of the pressure measurements, a qualitative analysis has been made to clarify the flow behavior and the physical meaning of the performance diagram. It appears that the choking phenomenon of the secondary flow plays an important role in the starting process of the ejector. When the secondary inlet height is relatively small, the choked secondary flow and the supersonic primary flow could be employed to protect the static pressure in the suction chamber from being disturbed by the back pressure effect at a certain primary stagnation pressure, which is lower than the starting pressure for the case of the zero-secondary flow. However, as the secondary inlet height increases and exceeds a critical value, the static pressure in the suction chamber rapidly increases, and the starting pressure of the ejector increases accordingly.