TOPOLOGY AND VORTEX STRUCTURES OF A CURVING TURBINE CASCADE WITH TIP CLEARANCE (Ⅰ)- EXPERIMENTAL MODEL AND TOPOLOGICAL FLOW PATTERNS ON BOTH ENDWALLS AND BLADE SURFACES

By means of ink trace visualization of the flows in conventional straight, positively curved and negatively curved cascades with tip clearance, and measurement of the aerodynamic parameters in the transverse section, and by appling topology theory, the structures on both endwalls and blade surfaces were analyzed. Compared with conventional straight cascade, blade positive curving eliminates the separation line of the upper passage vortex and leads the secondary vortex to change from close separation to open separation, while blade negative curving effects merely the positions of singular points and the intensities and scales of vortex.

Optimum Profiles of Endwall Contouring for Enhanced Net Heat Flux Reduction and Aerodynamic Performance

Successfully utilized non-axisymmetric endwalls to enhance turbine efficiencies(aerodynamic and turbine inlet temperatures)by controlling the characteristics of the secondary flow in a blade passage.This is accomplished by steady-state numerical hydrodynamics and deep knowledge of the field of flow.Because of the interaction between mainstream and purge flow contributing supplementary losses in the stage,non-axisymmetric endwalls are highly susceptible to the inception of purge flow exit compared to the flat and any advantage rapidly vanishes.The conclusions reveal that the supreme endwall pattern could yield a lowering of the gross pressure loss at the design stage and is related to the size of the top-loss location being productively lowered.This has led to diminished global thermal exchange lowered in the passage of the vane alone.The reverse flow adjacent to the suction side corner of the endwall is migrated farther from the vane surface,as the deviated pressure spread on the endwall accelerates the flow and progresses the reverse flow core still downstream.The depleted association between the tornado-like vortex and the corner vortex adjacent to the suction side corner of the endwall is the dominant mechanism of control in the contoured end wall.In this publication,we show that the non-axisymmetric endwall contouring by selective numerical shape change method at most prominent locations is advantageous in lowering the thermal load in turbines to augment the net heat flux reduction as well as the aerodynamic performance using multi-objective optimization.

Turbine Passage Secondary Flow Dynamics and Endwall Heat Transfer Under Different Inflow Turbulence

This study presents endwall hydrodynamics and heat transfer in a linear turbine cascade at Re 5×105 at low and high intensities of turbulence.Results are numerically predicted using the standard SST model and Reθ-γtransition model as well as using the high-resolution LES separately.The major secondary flow components,comprising the horseshoe,corner,and passage vortices are recognized and the impact on heat or mass transfer is investigated.The complicated behavior of turbine passage secondary flow generation and establishment are impacted by the perspective of boundary layer attributes and inflow turbulence.The passage vortex concerning the latest big leading-edge vane is generated by the enlargement of the circulation developed at the first instance adjacent to the pressure side becomes powerful and mixes with other vortex systems during its migration towards the suction side.The study conclusions reveal that substantial enhancements are attained on the endwall surface,for the entire spanwise blade extension on the pressure surface,and in the highly 3-D region close to the endwall on the suction surface.The forecasted suction surface thermal exchange depicts great conformity with the measurement values and precisely reproduces the enhanced thermal exchange owing to the development and lateral distribution of the secondary flows along the midspan of the blade passage downstream.The impacts of the different secondary flow structures on the endwall thermal exchange are described in depth.

AITEB - An European Research Project on Aero-thermodynamics of Turbine Endwalls and Blades

The paper delivers an overview on the European research project AITEB - Aerothermal Investigations on Turbine Endwalls and Blades, which started in year 2000 in the course of the 5. Framework Programme (GROWTH). The aim is to submit an integrated technology and design tool package for the advanced, aerothermal highly loaded design of turbines, especially: Experimental/numerical investigation on heat transfer and film-cooling in separated flow for highly loaded blades including advanced trailing edge cooling Heat transfer/ improved cooling of turbine endwalls: Experimental/numerical work on cooling of turbine endwalls, shrouds and recessed blade tips. Optimised CFD-process (drawing-grid-modelling-postprocessing-risk assessment) in order to derive the 'best practice' to use CFD as a time effective tool. After most of the project life, an overview on the project is delivered. Experimental results of test series at various test sites are compared to numerical simulations of the industrial and university partners.

Optimal Design of Non-axisymmetric Endwall with Variable-fidelity CFD Model

Non-axisymmetric endwall is numerically investigated to suppress the secondary endwall flow in a HP turbine suction surface.The endwall contour is represented by a non-uniform rational B-spline surface.Two solution spaces are designed by the rule of hypercube sampling through variation of nodal radial extension.Final most effective endwall contour is obtained through screening all available samples with surrogate-based model.Results show that non axisymmetric endwall reduces the turbulence kinetic energy of the passage vortex by 5.5%and the total pressure loss 1.6%.Local heat transfer coefficient is lowered on the afore passage endwall surface with a certain increment on the aft passage endwall.Overall averaged heat transfer coefficient is reduced.

Effect of Incoming Vortex on Secondary Flows in Turbine Cascades with Planar and Non-Axisymmetric Endwall

Non-Axisymmetric Endwall Profiling(NAEP) is commonly utilized in turbines to eliminate secondary flows.Nevertheless,most of the NAEP methods consider a single-blade row environment without incorporating the effect of the stage environment.This paper aims to investigate the influence mechanism of the incoming vortex on the endwall secondary flow structures of NAEP in a highly loaded turbine cascade.To model the incoming vortex in a stage environment,this study considers a half-delta wing as the vortex generator at the upstream of the turbine cascade.The NAEP is then carried out for a highly loaded turbine cascade with an in-house numerical optimization design platform subject to no incoming vortex.Numerical simulation is also carried out under the influence of the incoming vortex for the turbine cascades with both planar and non-axisymmetric endwall.This paper furthers investigated the pitchwise effect of the incoming vortex on the near endwall secondary flow.The results indicate that the NAEP effectively improves the endwall secondary flow of the turbine cascade,where the total pressure loss coefficient and the secondary kinetic energy(SKE) are reduced by 7.3%,and 45.7%,respectively.It is further seen that with the incoming vortex,the NAEP achieves a considerable control effect on the endwall secondary flow of the turbine cascade.With incoming vortex,the NAEP can still achieve considerable control effect on the endwall secondary flow of the turbine cascade;the averaged reductions of loss coefficient and SKE are 7.8% and 14.2%,respectively.Under some pitchwise locations,incoming vortex can suppress the convection of cross-passage flow toward the suction corner greatly and reduce the loss coefficient of the baseline cascade.The incoming vortex at 4/7 pitch impinged right at the blade leading edge,leading to the generation of low-momentum fluid,which increased the size and the strength of the horseshoe vortex.Under all the pitchwise locations,NAEP can suppress the secondary vortices,e.g.,the passage vortex and the counter vortex,considerably.

Experimental Study of Heat Transfer and Film Cooling Performance of Upstream Ejected Coolant on a Turbine Endwall

An upstream coolant injection that is different from the known leakage flow was introduced to protect the turbine endwall.This coolant is ejected tangentially from a row of cylindrical holes situated at the side of a backward-facing step.In this experiment,the effects of mass flow ratio and leakage slot width on the endwall heat transfer characteristics were investigated.The dimensionless heat transfer coefficient(Nu)and adiabatic film cooling effectiveness(η)on an axisymmetric turbine endwall were measured by the stable-state thermochromic liquid crystal(TLC)technique and the pressure sensitive paint(PSP)technique,respectively.Three mass flow ratios(MFR)of 0.64%,0.85%,and 1.07%,as well as two leakage slot widths(W)of 3.93 mm,and 7.86 mm were considered.Results indicate that the injection film suppresses the strength of the passage vortex,which leads to the coolant covering almost the entire endwall.This result is more evident for the higher MFR cases,meanwhile,the corresponding averaged film cooling effectiveness is increased with the enhancement of the MFR.However,the case with a higher MFR produces a higher heat transfer coefficient distribution,especially in the region close to the leakage slot edge.Besides,when the W is lower,the endwall presents a higherηand a lower Nu for all the cases,which can guide the optimal design of the endwall.

Numerical Investigation on Flow and Cooling Characteristics of a Micro-Ribbed Vane Endwall

The secondary flow originated from the inherent pressure gradient inside the vane cascade has a strong impact on the endwall cooling performance as the crossflow sweeps the upstream coolant jet towards the suction side,resulting in intensifying thermal load near the pressure side endwall.Hence a novel ribbed-endwall is introduced to suppress passage crossflow.The effects of the mass flow ratio and the rib layout were examined using numerical simulations by solving the three-dimensional Reynolds-averaged Navier-Stokes(RANS)equations with the shear stress transport(SST)k-ωturbulence model.The results indicate that the ribs effectively prevent the coolant migrating from the pressure side to the suction side,helping the coolant jet to spread along the lateral orientation.Therefore,the endwall adiabatic film cooling effectiveness is substantially improved.The maximum cooling effectiveness is achieved for the case with three-ribs when the height of the rib equals one hole diameter among all cases.The area-averaged adiabatic cooling effectiveness is enhanced by 31.6%relative to the flat endwall when the mass flow ratio of coolant to mainstream equals to 0.52%.More importantly,the ribbed-endwall obtains a relatively lower level of aerodynamic loss owing to the reduced lateral migration inside the vane cascade.

Static pressure redistribution mechanism of non-axisymmetric endwall based on radial equilibrium

Non-axisymmetric endwall contouring has been proved to be an effective flow control technique in turbomachinery.Several different flow control mechanisms and qualitative design strategies have been proposed.The endwall contouring mechanism based on the flow governing equations is significant for exploring the quantitative design strategies of the nonaxisymmetric endwall contouring.In this paper,the static pressure redistribution mechanism of endwall contouring was explained based on the radial equilibrium equation.A quantified expression of the static pressure redistribution mechanism was proposed.Compressor cascades were simulated using an experimentally validated numerical method to validate the static pressure redistribution mechanism.A geometric parameter named meridional curvature(Cme)is defined to quantify the concave and convex features of the endwall.Results indicate that the contoured endwall changes the streamline curvature,inducing a centrifugal acceleration.Consequently,the radial pressure gradient is reformed to maintain the radial equilibrium.The convex endwall represented by positive Cme increases the radial pressure gradient,decreasing the endwall static pressure,while the concave endwall represented by negative Cme increases the endwall static pressure.The Cme helps to establish the quantified relation between the change in the endwall radial pressure gradient and the endwall geometry.Besides,there is a great correlation between the distributions of the Cme and the change in the endwall static pressure.It can be concluded that the parameter Cme can be considered as a significant parameter to parameterize the endwall surface and to explore the quantitative design strategies of the nonaxisymmetric endwall contouring.

Optimization of endwall contouring in axial compressor S-shaped ducts

This paper presents a numerical investigation of the potential aerodynamic benefits of using endwall contouring in a fairly aggressive duct with six struts based on the platform for endwall design optimization.The platform is constructed by integrating adaptive genetic algorithm（AGA）, design of experiments（DOE）, response surface methodology（RSM） based on the artificial neural network（ANN）, and a 3D Navier–Stokes solver.The visual analysis method based on DOE is used to define the design space and analyze the impact of the design parameters on the target function（response）.Optimization of the axisymmetric and the non-axisymmetric endwall contouring in an S-shaped duct is performed and evaluated to minimize the total pressure loss.The optimal ducts are found to reduce the hub corner separation and suppress the migration of the low momentum fluid.The non-axisymmetric endwall contouring is shown to remove the separation completely and reduce the net duct loss by 32.7%.

Endwall aerodynamic losses from turbine components within gas turbine engines

A survey of research on aerodynamic loss investigations for turbine components of gas tuibine engines is presented.Experimental and numerically predicted results are presented from investigations undertaken over the past 65 plus years.Of particular interest are losses from the development of secondary flows from airfoil/endwall interactions.The most important of the airfoilAmdwall secondary flows are passage vortices,counter voitices,and corner vortices.The structure and development of these secondaiy flows are described as they affect aerodynamic perfonnance within and downstream of turbine passage flows in compressible,high speed flows with either subsonic or transonic Mach number distributions,as well as within low-speed,incompressible flows.Also discussed are methods of endwall contouring,and its consequences in regard to airfoil/endwall secondary flows.

Full blended blade and endwall design of a compressor cascade

In the current state-of-the-art,high-loss flow in the endwall significantly influences compressor performance.Therefore,the control of endwall corner separation in compressor blade rows is important to consider.Based on the previous research of the Blended Blade and End Wall(BBEW)technique,which can significantly reduce corner separation,in combination with a nonaxisymmetric endwall,the full-BBEW technique is proposed in this study to further reduce the separation in endwall region.The principle of the unchanged axial passage area is considered to derive the geometric method for this technique.Three models are further classified based on different geometric characteristics of this technique:the BBEW model,Inclining-Only End Wall(IOEW)model,and full-BBEW model.The most effective design of each model is then found by performing several optimizations at the design point and related numerical investigations over the entire operational conditions.Compared with the prototype,the total pressure loss coefficient decreases by 7%–9%in the optimized full-BBEW at the design point.Moreover,the aerodynamic blockage coefficient over the entire operational range decreases more than the other models,which shows its positive effect for diffusion.This approach has a larger decrease at negative incidence angles where the intersection of the boundary layer plays an important role in corner separation.The analysis shows that the blended blade profile enlarges the dihedral angle and creates a span-wise pressure gradient to move low momentum fluid towards the mainstream.Furthermore,the inclining hub geometry accelerates the accumulated flow in the corner downstream by increasing the pressure gradient.Overall,though losses in the mainstream grow,especially for large incidences,the full-BBEW technique effectively reduces the separation in corners.

Numerical Investigation and Non-Axisymmetric Endwall Profiling of a Turbine Stage

This paper presents an optimization of a high pressure turbine by constructing non-axisymmetric endwalls to the stator row and the rotor hub.The optimization was quantified by using optimization algorithms based on the multi-objective function.The objective was to increase total-to-total efficiency with the constraint on the mass flow rate equal to the design point value.In order to ensure that global optimum could be achieved,the function of parameters was first approximated through the artificial neural network,and then optimum was achieved by implementing the genetic algorithm.It was adopted through the design and optimization environment of FineTM/Design3 D.Three individual treatments of the endwalls were presented.Firstly,the hub and the shroud of the stator were optimized together.Secondly,the hub of the rotor was optimized.Thirdly,the rotor hub was optimized in the presence of the optimized stator.The result of the investigation showed that the optimized shape of the endwalls can significantly help to increase the efficiency up to 0.18%with the help of a reduction of the transverse pressure gradient.The coefficient of secondary kinetic energy,entropy coefficient,spanwise mass averaged entropy were reduced.In order to investigate the periodic effects,the design of the optimized turbine under steady simulations was confirmed through unsteady simulations.The last part of the investigation made sure that the performance improvement remained consistent over the full operating line at off-design conditions by the implementation of non-axisymmetric endwalls.

Effects of Trenched Film Hole Configurations on the Endwall Film Cooling and Suction Side Phantom Cooling

To maximize the turbine thermal efficiency, modern gas turbine’s inlet temperature is significantly augmented within the past few decades. To prolong the lifespan of gas turbines, many efficient cooling techniques have been proposed and applied in the endwall cooling schemes. However, conventional discrete film hole does not take effect at the leading edge nearby region. In this research, how the trenched film hole configurations affects the endwall cooling and phantom cooling characteristics were deeply studied by using a verified approach. Steady 3D Reynolds-averaged Navier-Stokes(RANS) governing equations together with the shear stress transport(SST) k-w turbulence model have been solved. Firstly, results indicate that trenched film holes greatly influence the cooling effectiveness at leading edge nearby region compared to normal case. Nevertheless, suction side phantom cooling is hardly influenced by the trenched film holes. Secondly, the case with a smaller trench width obtains higher endwall cooling effectiveness, particularly at upstream region. More importantly, the cases with W=3D achieve large cooling effectiveness at leading edge nearby region with little influence by trench depth. Additionally, majority of trenched film holes coolant flow is driven towards middle passage. Therefore, the suction side phantom cooling is unaffected by the trenched film holes.

Investigation of flow unsteadiness in a highly-loaded compressor cascade using a dynamic mode decomposition method

Unsteady flow in the hub endwall region has long been a hot topic in the turbomachinery community.However important it is to the performance of the whole engine,the coherent unsteady flow phenomena are still not well understood.In this paper,the complex flow field in the hub endwall of a cantilevered compressor cascade has been investigated through numerical approach.The predicted results were validated by experimental data.To highlight the dominant flow structures among irregular and chaotic motions of various vortices,a Dynamic Mode Decomposition(DMD)method was utilized.The results show that there exist three dominant periodic flow structures:the oscillation of the leakage vortex,a circumferential migration of a Breakdown Induced Vortex(BIV)and the fluctuation of the passage vortex.These three coherent structures all together form a self-sustained closed loop which accounts for the flow unsteadiness of the studied cascade.During this process,the BIV plays a key role in inducing the flow unsteadiness.Only if the BIV is strong enough to affect the passage vortex,the flow unsteadiness occurs.This study expands current knowledge base of flow unsteadiness in a compressor environment,and shows the efficacy of the DMD method for revealing the origin of flow unsteadiness.

Improved Turbine Vane Endwall Film Cooling by Using Sand-Dune-Inspired Design

As continuous of the previous sand-dune-inspired design,the Barchan-Dune-Shaped Injection Compound(BDSIC)’s film cooling performance at the endwall region was further investigated both experimentally and numerically.While the public 777-shaped hole was served as a baseline,the BDSIC’s endwall effectiveness was assessed at various blowing ratios.Experiments were performed in a single-passage transonic wind tunnel using pressure-sensitive paint(PSP)technique.Carbon dioxide was used as coolant with density ratio of DR=1.53.The purge slot’s blowing ratio was fixed at M=0.3,but the coolant holes were adjusted within M=0.5–2.0.The measured experimental results indicate that the film distribution at the endwall is strongly affected by the passage flow structures.The BDSIC holes demonstrate much higher film effectiveness than the 777-shaped holes for all blowing ratios,~30%enhancement for regionally averaged effectiveness at M=1.0 and~26%at M=2.0.As shown by the numerical results,the existence of BDSIC reduced the coolant penetration effect at a higher blowing ratio.Coolant was deflected and its momentum increased in the streamwise direction,therefore providing more robust film coverage over the endwall region.The anti-counter-rotating vortex pair induced by the BDSIC further stabilized the coolant film and increased the coolant spreading downstream.

Numerical Investigation of a Turbine Stator with Nonaxisymmetric Endwall Profiling

Nonaxisymmetric endwall is an effective method to reduce secondary loss and improve aerodynamic performance.In this paper,a nonaxisymmetric endwall automated optimization process based on the nonuniform rational B-spline surface(NURBS)technique was proposed.This technique was applied for the aerodynamic optimization of the turbine stator shroud endwall to reduce total pressure loss and secondary kinetic energy.The flow fields of the datum endwall design(Datum)and optimization endwall design(Opt)were investigated and compared.Quantitative loss analysis was performed with a loss breakdown method.The entropy generation was classified as profile loss,secondary loss and trailing edge loss,all of which were reduced.The secondary loss was much smaller than the profile loss.In general,the blade row total entropy loss decreased by 11.7%.The results showed that the Opt design reduced total pressure loss and coefficient of secondary kinetic energy by 11.1%and 11.0%,respectively.The decrease in secondary kinetic energy could be attributed to the reduction in the horseshoe vortex and the reduced transverse pressure gradient.When the outlet Mach numbers and inlet incidence angles vary,the performance of the profiled endwall design was always better than the datum design.In the turbine stage simulation,the efficiency was increased by 0.28%with nonaxisymmetric endwall.

Influence of Endwall Air Injection with Discrete Holes on Corner Separation of a Compressor Cascade

The aim of this study is to reveal the influence mechanism of endwall air injection with distributed holes on the corner separation of a highly loaded compressor cascade,so as to promote the application of injection in aero-engines.Single-hole and double-hole endwall injection schemes featuring different axial locations,pitchwise locations,injection mass rates and injection directions,were designed and investigated.Results showed that the corner separation was eliminated by endwall injection;the optimal single-hole injection scheme achieved an endwall loss coefficient reduction of 29.7%,with injection coefficient as low as 0.48%.The optimal axial location of single-hole endwall injection was at 82%axial chord,being the center of corner separation.However,as injection hole was located at upstream of it,endwall injection resulted in severer corner separation.The mid-span flow field was deteriorated after endwall injection,which was due to 3D flow effects,i.e.,AVDR(axial velocity density ratio)effect and low-momentum fluid spanwise migration effect.The optimal injection was achieved at low injection angle and from close to the suction surface on pitchwise.Double-hole injection exhibited inferior performance compared with single-hole,which was due to the interaction of the two injection streams and mixing of injection streams with the bulk stream.

Effect of Moving Endwall on Hub Leakage Flow of Cantilevered Stator in a Linear Compressor Cascade

The cantilevered stator has the advantages of reducing mass and axial length of highly loaded com-pressor.The details of the hub leakage flow resulting from the clearance between the high-speed moving hub and the cantilevered stator hub are unclear.In this paper,the effect of a moving endwall on the hub leakage flow of a cantilevered stator in a linear compressor cascade was studied.After the simulation method was verified with the experimental results,the time-averaged results of unsteady Reynolds averaged Navier-Stokes(URANS)were selected to study a case with a hub clearance of 2 mm.The results show that the effect of the moving endwall of the cantilevered cascade on the general characteristics with below 30%span increases the leakage mass flow rate and reduces the static pressure coefficient at three conditions of 0°,6°,and-7°incidences,and the change is most significant at-7°incidence.The effect of the moving endwall on the total pressure loss coefficient varies with different operating conditions,which decreases by 15.94%at 0°incidence,and increases by 4.77%and 18.51%at 6°incidence and-7°incidence,respectively.The influence of the moving endwall is below 14%span at-7°incidence,below 23%span at 0°incidence,and below 30%span at 6°incidence.These effects correspond to the static pressure coefficient and the difference of static pressure coefficient representing the blade loading.In designing the cantilevered stator and matching between the stages of a multistage compressor,the quantitative research results of this paper have certain guiding significance.

Effects of mid-passage gap with a variable surface angle on a turbine vane endwall’s aerothermal and film cooling performances

The mid-passage gap is an inevitable structure in a vane passage due to turbine vanes being manufactured individually.The coolant from this gap is able to prevent hot mainstream ingression and provide cooling protection for the endwall.A novel idea of enlarging the endwall’s coverage area and reducing the endwall’s thermal load by applying the mid-passage gap with variable surface angles is carried out in this paper.The endwall’s aerothermal and film cooling performances under four mid-passage gap modes at three typical mass flow ratio conditions are numerically investigated.Results indicate that under the traditional mid-passage mode,the coolant flows into the mainstream with a perpendicular incidence angle and can’t stick to the endwall.Thus,cooling failure occurs,and the endwall’s thermal load is badly increased.The film cooling level at the suction-side endwall is improved when applying the mid-passage gap of a 45
surface angle due to the secondary vortex being suppressed.In addition,when applying the mid-passage gap of a 135
surface angle,the horseshoe vortex is pushed away,and the coverage area at the pressure-side endwall is enlarged significantly.The best film cooling performance is achieved when the upstream surface angle is 135
and the downstream surface angle is 45
due to the adiabatic film cooling effectiveness being increased at both the pressure-and suction-side endwall.When the mass flow ratio is 1.5%,the coverage area is enlarged by 43%,and the area-averaged adiabatic film cooling effectiveness is increased by 37%,when compared with those under the traditional mid-passage mode.