This article reports on numerical investigations of passive control techniques used for the performance enhancement of a large diameter axial flow cooling fan through modification of the blade tip geometry. Using open-source software, a novel meshing strategy is developed to carry out both steady and unsteady numerical computations of a periodic section of an axial flow fan. Analyzing the flow near the blade tip with a reduction in tip clearance, two predominant flow phenomena are identified. These two flow phenomena are further investigated with the aim of controlling them through implementation of a tip appendage design. Both introduced end-plate designs indicate effective control of each relevant flow phenomena. The constant thickness (CT) end-plate design is found to increase all fan performance characteristics at lower than design point (DP) flowrates, while increasing the fan’s peak efficiency plateau toward the rotor stall margin. However, none of the CT end-plate designs are able to improve the fan’s performance characteristics at its DP. The introduction of a novel trailing-edge (TE) end-plate design is found to increase all fan performance characteristics across the entire evaluated stable operating range, with an indicated increase of 37.3 percent in total-to-static pressure rise and a 2.9 percentage point increase in total-to-static efficiency at the fan’s DP flowrate. The aerodynamic performance results attest to the associated benefits of the investigated passive control techniques.

In the drive for lower fuel consumption through increased bypass ratio and increased overall pressure ratio (OPR), engine designs for the next generation of single-aisle aircraft will require core sizes below 3 lb/s and OPRs above 50. Traditionally, these core sizes are the domain of centrifugal compressors, but materials limit pressure ratio in these machines to well below 50. An all-axial high-pressure compressor (HPC) at this core size, however, comes with limitations associated with the small blade spans at the back of the HPC, as clearances, fillets, and leading edges do not scale with the core size. The result is a substantial efficiency penalty, driven primarily by the tip leakage flow produced by the larger clearance-to-span ratio, which negates the cycle efficiency benefits of the high OPR. In order to enable small-core, high-OPR, all-axial compressors mitigating technologies need to be developed and implemented to reduce the large clearance-to-span efficiency penalty. However, for this technology development to be successful, it is imperative that predictive design tools accurately model the overall flow physics and trends of the technologies developed. In this paper, we describe an effort to determine whether different modeling standards are required for a large clearance-to-span ratio, and if so, identify criteria for an appropriate solver and/or mesh. Multiple models are run and results compared with data collected in the NASA Glenn Research Center’s low-speed axial compressor. These comparisons show that steady Reynolds-averaged Navier–Stokes (RANS) solvers can predict the pressure-rise characteristic to an acceptable level of accuracy, if careful attention is paid to mesh topology in the tip region. However, unsteady tools are necessary to accurately capture radial profiles of blockage and total pressure. The Delayed-Detached Eddy Simulation model was also used to run this geometry, but did not resolve any additional features not captured by the unsteady RANS simulation near stall.

The non-uniform disturbance in circumferential direction is main cause for occurrence of rotating stall in turbomachinery. In order to study the effect of tip clearance leakage flow on rotating stall, mixed-flow pump models with different tip clearances are simulated and energy performance curves and internal flow structures are obtained and compared. The results show that the computed pump efficiency and the internal flow field of the pump are in good agreement with experimental results. A saddle region appears in energy performance curves of three tip clearances and with decrease in tip clearance, the head and efficiency of mixed-flow pump increase and critical stall point shifts and stable operating range of mixed-flow pump decreases, which indicates that mixed-flow pump stalls easily for smaller tip clearance. Under deep stall condition, influence of leakage flow in end wall area increases gradually with decrease in clearance. For small clearance, the leakage flow moves away from suction surface to some distance to form number of leakage vortex strips with mainstream flow and flows over the leading edge of next blade and then flows downstream into different flow passages generating back flow and secondary flow separation at the blade inlet, which seriously damages the spatial structure of inlet flow. This results in earlier occurrence of stall. With increase in clearance, the leakage vortex develops along radial direction towards middle of flow channel and large flow separation occurs in downstream channel which induces deep stall.

For an axial compressor stator with tip gap the boundary layer in the hub end-wall region has a significant influence on the development and progression of the tip leakage vortex. Herein the so-called boundary layer skew, which develops through relative motion of the hub, is of particular interest. Therefore, experimental and numerical investigations of a single axial compressor stator row with varying tip gap height (tip gap height/chord length = 2.0%|5.4%|6.7%) have been conducted. Comparing cases with rotating or stationary hub end-wall segments upstream of the examined vanes allowed to determine the effect of skewed and un-skewed inflow boundary layer. The steady state flow fields up- and downstream of the stator row were measured using five-hole pressure probes. For validation and to improve the understanding of the existing flow phenomena 3D-RANS CFD simulations using a commercial flow solver were carried out. Furthermore, analog cases with no tip gap were examined and considered in the comparisons to extend the knowledge on this boundary layer characteristic. The results show that the boundary layer skew has a major influence on the trajectory and size of the tip leakage vortex for the cases with tip clearance. The effect of reduction of the produced losses decreases with increasing tip gap height.

The compressors of power-generation gas turbines (GTs) have a high stage count, blades with low aspect ratios, and large clearances. These features promote strong secondary flows. An important outcome deriving from the convection of intense secondary flows is the enhanced span-wise transport of fluid properties mainly involving the rear stages, generally referred to as “radial mixing.” An incorrect prediction of this key phenomenon may result in inaccurate performance evaluation and could mislead designers. In the rear compressor stages, the stream-wise vorticity associated with tip clearance flows is one of the main drivers of the span-wise transport phenomenon. Limiting it by averaging the flow at row interfaces is the reason why a steady analysis underpredicts radial mixing. To properly forecast the span-wise transport, an unsteady analysis should be adopted. However, this approach has a computational cost not yet suitable for industrial purposes. Currently, only the steady simulation can fit in a lean design chain and any model upgrade improving its radial mixing prediction would be highly beneficial. To attain some progresses in Reynolds-averaged Navier–Stokes (RANS) model, its lack of convection of stream-wise vorticity must be addressed. This can be done by acting on another mixing driver that is turbulent diffusion; by enhancing turbulent viscosity, one can promote span-wise diffusion, thus improving the radial mixing prediction. In this paper, this strategy to update the RANS model and its application on an existing compressor is presented, together with the model tuning that has been performed using unsteady results as the target.

The aim of the present paper is to improve the physical understanding of discrete prestall flow disturbances developing in the tip area of the compressor rotor. For this purpose, a complementary instrumentation was used in a single-stage axial compressor. A set of pressure transducers evenly distributed along the circumference surface mounted in the casing near the rotor tip leading edges measures the time-resolved wall pressures simultaneously to an array of transducers recording the chordwise static pressures. The latter allows for plotting quasi-instantaneous casing pressure contours. Any occurring flow disturbances can be properly classified using validated frequency analysis methods applied to the data from the circumferential sensors. While leaving the flow coefficient constant, a continuously changing number of prestall flow disturbances appears to be causing a unique spectral signature, which is known from investigations on rotating instability. Any arising number of disturbances is matching a specific mode order found within this signature. While the flow coefficient is reduced, the propagation speed of prestall disturbances increases linearly, and meanwhile, the speed seems to be independent from the clearance size. Casing contour plots phase-locked to the rotor additionally provide a strong hint on prestall disturbances clearly not to be caused by a leading edge separation. Data taken beyond the stalling limit demonstrate a complex superposition of stall cells and flow disturbances, which the title “prestall disturbance” therefore does not fit to precisely any more. Different convection speeds allow the phenomena to be clearly distinguished from each other. Furthermore, statistical analysis of the pressure fluctuations caused by the prestall disturbances offer the potential to use them as a stall precursor or to quantify the deterioration of the clearance height between the rotor blade tips and the casing wall during the lifetime of an engine.

Advanced predictions of blade flutter have been continually pursued. It is noted however that validation cases of unsteady CFD methods against experimental cases with detailed 3D unsteady pressures are still rather lacking. The main objectives of the present work are two-folds. First, validate and understand the characteristics of blade tip clearance, as well as a bubble-type flow separation for an unsteady CFD solver against a 3D oscillating cascade experiment. And second, examine the applicability of the influence coefficient method (ICM) as widely used in an oscillating linear cascade setup. In the first part, the capability of a widely used commercial solver (CFX) for unsteady flows induced by a 3D oscillating compressor cascade is examined. The present computations have shown consistently a destabilizing effect of increasing blade tip clearance, in agreement with the experiment. More remarkably, the computational analyses reveal a distinctive interplay between the inlet endwall boundary layer and the tip clearance in relation to the aerodynamic damping. Different inlet endwall boundary layer thicknesses are shown to lead to qualitatively different aeroelastic stability characteristics in relation to tip clearance. The aero-damping variation with the tip clearance under the influence of the inlet endwall boundary layer seems to correlate closely to a balancing act between the passage vortex and the tip leakage vortex. The tip clearance aeroelastic behavior seems also in line with a simple quasi-steady analysis. On the other hand, the mid-chord laminar bubble separation on suction surface, though with a clear signature in the local aero-damping, has negligible effects on the overall stability. The second part aims to examine computationally the applicability of the influence coefficient method in a linear cascade setup. The comparison between the cascade-based ICM data and a baseline “tuned cascade” shows that the differences in the sensitivity to the far-field treatment can be significant, depending on inter-blade phase angles. On the other hand, non-linearity effects closely relevant to the basic linear assumption of the ICM are shown to only have a small influence. The present results suggest that extra caution should be exercised when comparing a CFD-based tuned cascade model with a finite cascade-based ICM model, at conditions close to acoustic resonance. The resultant discrepancies may well arise from the inherently different far-field sensitivities between the two models, rather than those typical numerical and physical modeling aspects of interest (e.g., meshing, spatial and temporal discretization errors as well as turbulence modeling).

In compressor design, a convenient way to save time is to scale an existing geometry to required specifications, rather than developing a new design. The approach works well when scaling compressors of similar size at high Reynolds numbers but becomes more complex when applied to small-scale machines. Besides the well-understood increase in surface friction due to increased relative surface roughness, two other main problems specific to small-scale turbomachinery can be specified: (1) the Reynolds number effect, describing the non-linear dependency of surface friction on Reynolds number and (2) increased relative tip clearance resulting from manufacturing limitations. This paper investigates the role of both effects in a geometric scaling process, as used by a designer. The work is based on numerical models derived from an experimentally validated geometry. First, the effects of geometric scaling on compressor performance are assessed analytically. Second, prediction capabilities of reduced-order models from the public domain are assessed. In addition to design point assessment, often found in other publications, the models are tested at off-design. Third, the impact of tip leakage on compressor performance and its Reynolds number dependency is assessed. Here, geometries of different scale and with different tip clearances are investigated numerically. Fourth, a detailed investigation regarding tip leakage driving mechanisms is carried out and design recommendations to improve small-scale compressor performance are provided.

The pressure difference between suction and pressure sides of a turbine blade leads to tip leakage flow, which adversely affects the first-stage high-pressure (HP) turbine blade tip aerodynamics. In modern gas turbines, HP turbine blade tips are exposed to extreme thermal conditions requiring cooling. If the coolant jet directed into the blade tip gap cannot counter the leakage flow, it will simply add up to the pressure losses due to leakage. Therefore, the compromise between the aerodynamic loss and the gain in tip-cooling effectiveness must be optimized. In this paper, the effect of tip-cooling configuration on the turbine blade tip is investigated numerically from both aerodynamics and thermal aspects to determine the optimum configuration. Computations are performed using the tip cross section of GE-E3 HP turbine first-stage blade for squealer and flat tips, where the number, location, and diameter of holes are varied. The study presents a discussion on the overall loss coefficient, total pressure loss across the tip clearance, and variation in heat transfer on the blade tip. Increasing the coolant mass flow rate using more holes or by increasing the hole diameter results in a decrease in the area-averaged Nusselt number on the tip floor. Both aerodynamic and thermal response of squealer tips to the implementation of cooling holes is superior to their flat counterparts. Among the studied configurations, the squealer tip with a larger number of cooling holes located toward the pressure side is highlighted to have the best cooling performance.

Effects of tip clearance size and flowrate on the flow around the tip of an axial turbomachine rotor are studied experimentally. Visualizations and stereo-particle image velocimetry (PIV) measurements in a refractive index-matched facility compare the performance, leakage velocity, and the trajectory, growth rate, and strength of the tip leakage vortex (TLV) for gaps of 0.49% and 2.3% of the blade chord, and two flowrates. Enlarging the tip clearance delays the TLV breakup in the aft part of the rotor passage at high flowrates but causes earlier breakup under pre-stall conditions. It also reduces the entrainment of endwall boundary layer vorticity from the separation point where the leakage and passage flows meet. Reducing the flowrate or tip gap shifts the location of the TLV detachment from the blade suction side (SS) upstream to points where the leakage velocity is 70–80% of the tip speed. Once detached, the growth rates of the total shed circulation are similar for all cases, i.e., varying the gap or flowrate mostly shifts the detachment point. The TLV migration away from the SS decreases with an increasing gap but not with the flowrate. Two mechanisms dominate this migration: initially, the leakage jet pushes the TLV away from the blade at 50% of the leakage velocity. Further downstream, the TLV is driven by its image on the other side of the endwall. Differences in migration rate are caused by the smaller distance between the TLV and its image for the narrow gap, and the increase in initial TLV strength with decreasing flowrate and gap.

The effect of film cooling on a transonic squealer tip has been examined in a high speed linear cascade, which operates at engine-realistic Mach and Reynolds numbers. Tests have been performed on two uncooled tip geometries with differing pressure side rim edge radii, and a cooled tip matching one of the uncooled cases. The pressure sensitive paint technique has been used to measure adiabatic film cooling effectiveness on the blade tip at a range of tip gaps and coolant mass flow rates. Complementary tip heat transfer coefficients have been measured using transient infrared thermography, and the effects of the coolant film on the tip heat transfer and engine heat flux were examined. The uncooled data show that the tip heat transfer coefficient distribution is governed by the nature of flow reattachments and impingements. The squealer tip can be broken down into three regions, each exhibiting a distinct response to a change in the tip gap, depending on the local behavior of the overtip leakage flow. Complementary computational fluid dynamics (CFD) shows that the addition of casing motion causes no change in the flow over the pressure side rim. Injected coolant interacts with the overtip leakage flow, which can locally enhance the tip heat transfer coefficient. The film effectiveness is dependent on both the coolant mass flow rate and tip clearance. At increased coolant mass flow, areas of high film effectiveness on the pressure side rim coincide strongly with a net heat flux reduction and in the subsonic tip region with low heat transfer coefficient.

Blade flutter in the last stage is an important design consideration for the manufacturers of steam turbines. Therefore, the accurate prediction method for blade flutter is critical. Since the majority of aerodynamic work contributing to flutter is done near the blade tip, resolving the tip leakage flow can increase the accuracy of flutter predictions. The previous research has shown that the induced vortices in the tip region can have a significant influence on the flow field near the tip. The structure of induced vortices due to the tip leakage vortex cannot be resolved by unsteady Reynolds-averaged Navier–Stokes (URANS) simulations because of the high dissipation in turbulence models. To the best of author’s knowledge, the influence of induced vortices on flutter characteristics has not been investigated. In this paper, the results of detached-eddy simulation (DES) and URANS flutter simulations of a realistic-scale last-stage steam turbine are presented, and the influence of induced vortices on the flutter stability has been investigated. Significant differences for the predicted aerodynamic work coefficient distribution on the blade surface, especially on the rear half of the blade suction side near the tip, are observed. At the least stable interblade phase angle (IBPA), the induced vortices show a destabilizing effect on the blade aeroelastic stability. The motion of induced vortices during blade oscillation is dependent on the blade amplitude, and hence, the aerodynamic damping is also dependent on the blade vibration amplitude. In conclusion, the induced vortices can influence the predicted flutter characteristics of the steam turbine test case.

The optimum aspect ratio at which maximum efficiency occurs is relatively low, typically between 1 and 1.5. At these aspect ratios, inaccuracies inherently exist in the decomposition of the flow field into freestream and endwall components due to the absence of a discernible freestream. In this paper, a unique approach is taken: a “linear repeating stage” concept is used in conjunction with a novel way of defining the freestream flow. Through this approach, physically accurate decomposition of the flow field for aspect ratios as low as ∼0.5 can be achieved. This ability to accurately decompose the flow leads to several key findings. First, the endwall flow is found to be dependent on static pressure rise coefficient and endwall geometry, but independent of the aspect ratio. Second, the commonly accepted relationship that endwall loss coefficient varies inversely with the aspect ratio is shown to be physically inaccurate. Instead, a new term, which the authors refer to as the “effective aspect ratio,” should replace the term “aspect ratio.” Moreover, not doing so can result in efficiency errors of ∼0.6% at low aspect ratios. Finally, there exists a low aspect ratio limit below which the two endwall flows interact causing a large separation to occur along the span. From these findings, a low-order model is developed to model the effect of varying aspect ratio on compressor performance. The last section of the paper uses this low-order model and a simple analytical model to show that to a first order, the optimum aspect ratio is just a function of the loss generated by the endwalls at zero clearance and the rate of change in profile loss due to blade thickness. This means that once the endwall configuration has been selected, i.e., cantilever or shroud, the blade thickness sets the optimum aspect ratio.

Impinging jet arrays are typically used to cool several gas turbine parts. Some examples of such applications can be found in the internal cooling of high-pressure turbine airfoils or in the turbine blade tip clearances control of aero-engines. The effect of the wall-to-jets temperature ratio (TR) on heat transfer is generally neglected by the correlations available in the open literature. In the present contribution, the impact of the temperature ratio on the heat transfer for a real engine active clearance control system is analyzed by means of validated computational fluid dynamics (CFD) computations. At different jets Reynolds number and considering several impingement array arrangements, a wide range of target wall-to-jets temperature ratio is accounted for. Computational results prove that both local and averaged Nusselt numbers reduce with increasing. An in-depth analysis of the numerical data shows that the last mentioned evidence is motivated by both the heat transfer incurring between the spent coolant flow and the fresh jets and the variation of gas properties with temperature through the boundary layer. A scaling procedure, based on the TR power law, was proposed to estimate the Nusselt number at different wall temperature levels necessary to correct available open-literature correlations, typically developed with small temperature differences, for real engine applications.

This paper describes the role of tip leakage flow in creating the leading edge separation necessary for the onset of spike-type compressor rotating stall. A series of unsteady multipassage simulations, supported by experimental data, are used to define and illustrate the two competing mechanisms that cause the high incidence responsible for this separation: blockage from a casing-suction-surface corner separation and forward spillage of the tip leakage jet. The axial momentum flux in the tip leakage flow determines which mechanism dominates. At zero tip clearance, corner separation blockage dominates. As clearance is increased, the leakage flow reduces blockage, moving the stall flow coefficient to lower flow, i.e., giving a larger unstalled flow range. Increased clearance, however, means increased leakage jet momentum and contribution to leakage jet spillage. There is thus a clearance above which jet spillage dominates in creating incidence, so the stall flow coefficient increases and flow range decreases with clearance. As a consequence, there is a clearance for maximum flow range; for the two rotors in this study, the value was approximately 0.5% chord. The chordwise distribution of the leakage axial momentum is also important in determining stall onset. Shifting the distribution toward the trailing edge increases flow range for a leakage jet dominated geometry and reduces flow range for a corner separation dominated geometry. Guidelines are developed for flow range enhancement through control of tip leakage flow axial momentum magnitude and distribution. An example is given of how this might be achieved.

Optimal turbine blade tip designs have the potential to enhance aerodynamic performance while reducing the thermal loads on one of the most vulnerable parts of the gas turbine. This paper describes a novel strategy to perform a multi-objective optimization of the tip geometry of a cooled turbine blade. The parameterization strategy generates arbitrary rim shapes around the coolant holes on the blade tip. The tip geometry performance is assessed using steady Reynolds-averaged Navier–Stokes simulations with the k– ω shear stress transport (SST) model for the turbulence closure. The fluid domain is discretized with hexahedral elements, and the entire optimization is performed using identical mesh characteristics in all simulations. This is done to ensure an adequate comparison among all investigated designs. Isothermal walls were imposed at engine-representative levels to compute the convective heat flux for each case. The optimization objectives were a reduction in heat load and an increase in turbine row efficiency. The multi-objective optimization is performed using a differential evolution strategy. Improvements were achieved in both the aerodynamic efficiency and heat load reduction, relative to a conventional squealer tip arrangement. Furthermore, this work demonstrates that the inclusion of over-tip coolant flows impacts the over-tip flow field, and that the rim–coolant interaction can be used to create a synergistic performance enhancement.

The current industrial standard for numerical simulations of axial compressors is the steady Reynolds-averaged Navier–Stokes (RANS) approach. Besides the well-known limitations of mixing planes, namely their inherent inability to capture the potential interaction and the wakes from the upstream blades, there is another flow feature which is lost, and which is a major accountable for the radial mixing: the transport of streamwise vorticity. Streamwise vorticity is generated for various reasons, mainly associated with secondary and tip-clearance flows. A strong link exists between the strain field associated with the vortices and the mixing augmentation: the strain field increases both the area available for mixing and the local gradients in fluid properties, which provide the driving potential for the mixing. In the rear compressor stages, due to high clearances and low aspect ratios, only accounting for the development of secondary and clearance flow structures, it is possible to properly predict the spanwise mixing. In this work, the results of steady and unsteady simulations on a heavy-duty axial compressor are compared with experimental data. Adopting an unsteady framework, the enhanced mixing in the rear stages is properly captured, in remarkable agreement with experimental distributions. On the contrary, steady analyses strongly underestimate the radial transport. It is inferred that the streamwise vorticity associated with clearance flows is a major driver of radial mixing, and restraining it by pitch-averaging the flow at mixing planes is the reason why the steady approach cannot predict the radial transport in the rear part of the compressor.

Determination of a scalable Nusselt number (based on “adiabatic heat transfer coefficient”) has been the primary objective of the most existing heat transfer experimental studies. Based on the assumption that the wall thermal boundary conditions do not affect the flow field, the thermal measurements were mostly carried out at near adiabatic condition without matching the engine realistic wall-to-gas temperature ratio (TR). Recent numerical studies raised a question on the validity of this conventional practice in some applications, especially for turbine blade. Due to the relatively low thermal inertia of the over-tip-leakage (OTL) flow within the thin clearance, the fluids' transport properties vary greatly with different wall thermal boundary conditions and the two-way coupling between OTL aerodynamics and heat transfer cannot be neglected. The issue could become more severe when the gas turbine manufacturers are making effort to achieve much tighter clearance. However, there has been no experimental evidence to back up these numerical findings. In this study, transient thermal measurements were conducted in a high-temperature linear cascade rig for a range of tip clearance ratio (G/S) (0.3%, 0.4%, 0.6%, and 1%). Surface temperature history was captured by infrared thermography at a range of wall-to-gas TRs. Heat transfer coefficient (HTC) distributions were obtained based on a conventional data processing technique. The profound influence of tip surface thermal boundary condition on heat transfer and OTL flow was revealed by the first-of-its-kind experimental data obtained in the present experimental study.

The dimensionless model presented in part I of the corresponding paper to describe the flutter onset of two-fin rotating seals is exploited to extract valuable engineering trends with the design parameters. The analytical expression for the nondimensional work-per-cycle depends on three dimensionless parameters of which two of them are new. These parameters are simple but interrelate the effect of the pressure ratio, the height, and length of the interfin geometry, the seal clearance, the nodal diameter (ND), the fluid swirl velocity, the vibration frequency, and the torsion center location in a compact and intricate manner. It is shown that nonrelated physical parameters can actually have an equivalent impact on seal stability. It is concluded that the pressure ratio can be stabilizing or destabilizing depending on the case, whereas the swirl of the flow is always destabilizing. Finally, a simple method to extend the model to multiple interfin cavities, neglecting the unsteady interaction among them, is described.

In this paper, the effect of a novel honeycomb tip on suppressing tip leakage flow in turbine cascade has been experimentally and numerically studied. Compared to the flat tip cascade with 1%H blade height, the relative leakage flow in honeycomb tip cascade reduces from 3.05% to 2.73%, and the loss also decreases by 8.24%. For honeycomb tip, a number of small vortices are rolled up in the regular hexagonal honeycomb cavities to dissipate the kinetic energy of the clearance flow, and the fluid flowing into and out the cavities create aerodynamic interceptions to the upper clearance flow. As a result, the flow resistance in the clearance increased and the velocity of leakage flow reduced. As the gap height increases, the tip leakage flow and loss changes proportionally, but the growth rate in the honeycomb tip cascade is smaller. Considering its wear resistance of the honeycomb seal, a smaller gap height is allowed in the cascade with honeycomb tip, and that means honeycomb tip has better effect on suppressing leakage flow. Part honeycomb tip structure also retains the effect of suppressing leakage flow. It shows that locally convex honeycomb tip has better suppressing leakage flow effect than the whole honeycomb tip, but locally concave honeycomb tip is slightly less effective.