Hub Boundary Layer Research Articles

Experimental study on three-dimensional blade designs in an aggressive compressor transition duct

The increasing performance requirements for aero-engines necessitate the integrated design of compressor transition ducts with upstream components to reduce the axial length of the engine. This paper presents an experimental investigation of three-dimensional (3D) blade designs within an aggressive compressor transition duct integrated with the upstream stator. The aim is to reduce the total pressure loss and decrease the radial distortion of the flow field at the outlet of the transition duct, thus providing better inflow conditions for the high-pressure compressor. Detailed measurements of the internal flow field were conducted under both near stall (NS) and design (DE) conditions for three 3D blades. The results showed that in the downward path, adopting the 3D blades with positive dihedral and forward sweep contributed to mitigating the stator hub corner stall under the NS condition. Under the same operating conditions, enhanced kinetic energy at the stator root improved the spanwise uniformity of the transition duct outlet profile, while also increasing the anti-separation capability of the downstream hub boundary layer. The corner vortex at the stator root could enhance momentum mixing between the downstream hub boundary layer and the mainstream, thereby providing stronger separation resistance for the hub boundary layer under the NS condition. Compared to the baseline stators, the optimized 3D stators reduced system total pressure loss by 24% under the NS condition, while maintaining essentially unchanged total pressure loss under the DE condition.

Investigations on the unsteady flow mechanism of the high-throughflow fan with partial-height booster rotor

This paper describes the use of numerical simulations to evaluate the complicated steady and unsteady flow in a high-throughflow fan stage. A detailed investigation of the aerodynamic performance is carried out at the choking point, near-choking point, peak efficiency point, and near-stall point. The peak efficiency point is selected as a typical operating point to clarify the unsteady flow mechanism inside the fan stage. The unsteady simulation results show that the mass flow is lower under all operating conditions than in steady flow, and the difference increases as the operation point approaches the near-stall point. The isentropic efficiency decreases by 0.41% at the near-stall point, but improves slightly at the other points considered. The total pressure ratio does not change significantly. The total pressure loss coefficient of the stator increases significantly, but the changes in the isentropic efficiency of the fan stage mean that the stator has only a slight effect on the performance of the overall stage. An in-depth analysis of the peak efficiency point is conducted. First, the distribution of entropy along spanwise at the outlet of the rotor and stator domain is examined, and the unsteady effects are found to be mainly concentrated near the hub, mid-span, and higher-span of the stator. Detailed analysis of these three regions shows that the main influence comes from the distortion of the transmission of flow field information in the hub boundary layer at the rotor/stator interface, the interaction between the tip vortex and other vortexes near the mid-span region, and the interaction between the trailing edge separation vortex of the fan rotor and the surfaces of the stator.

Experimental Investigation of Secondary Flows and Length Reduction for a Low-Pressure Compressor Transition Duct

Abstract The need to reduce fuel-burn and emissions is pushing turbofan engines toward geared architectures with higher bypass ratios and small ultrahigh-pressure ratio cores. However, this increases the radial offset between compressor spools leading to a more challenging design for compressor transition ducts. For the duct connecting the fan to the engine core, this is further complicated by poor-quality flow generated at the fan hub, which is characterized by low total pressure and large rotating secondary flow structures. This paper presents an experimental evaluation of a new rotor designed to produce these larger flow structures and examines their effect on the performance of an engine sector stators (ESS) and compressor transition duct. Aerodynamic data were collected via five-hole probes, for time-averaged pressures and velocities and phase-locked hot-wire anemometry to capture the rotating secondary flows. The data showed that larger structures promoted mixing through the ESS increasing momentum exchange between the core and boundary layer flows. Measurements within the duct showed a continued reduction in the hub boundary layer, suggesting the duct had moved further from separation. Consequently, an aggressive duct with 12.5% length reduction was designed and tested and measurements confirmed the duct remained fully attached. Total pressure loss was slightly increased over the ESS, but this was offset by reduced loss in the duct due to improved flow quality. Overall, this length reduction represents a significant cumulative effect in reduced fuel-burn and emissions over the life of an engine.

STUDI NUMERIK SEPARASI ALIRAN 3D AKIBAT PENAMBAHAN FFST PADA BIDANG TUMPU AIRFOIL ASIMETRI

3D flow separation is a form of flow loss that cannot be avoided on turbo engines. In the axial compressor, 3D flow separation is due to the interaction between the blade boundary layer and the casing boundary layer or the hub boundary layer. The result of the secondary flow causes blockage of the flow which causes the pressure on the compressor to decrease. Efforts to reduce secondary flow are carried out by adding a FFST to endwall. This research was conducted in a numerical simulation using FLUENT 6.3.26 software. The parameters used in the free stream flow Re = 1.64 x 105 and Turbulence Intensity Tu = 0.3% to assess the comparison of the flow characteristics on the endwall of the British 9C7 / 22.5C50 asymmetric airfoil due to the addition of a FFST and without FFST with variations angle of attack (α) of 00, 80, 120, 140, 160 .The results show that the addition of FFST can increase the turbulent intensity in the area near the wall which turns into momentum, so that it has an impact on the ability of the flow to overcome the adverse pressure in the trailing edge area and further backward (delayed) separation which results in smaller wake. With the addition of the angel of attack, the saddle point position is more directed to the lower side and the attachment line is not induced by the horseshoe vortex, so that the flow is more able to follow the contours of the body, as a result the curling flow is weaker and the wake is narrower and the blockage (energy loss) can be reduced. The most effective energy reduction due to secondary flow through FFST occurs at α = 8 ° at 7.36%.

Numerical Analysis of High-Altitude Inlet Air on Boundary Layer Flow Loss in an Aero-Engine Compressor

A numerical analysis is performed to explore the high altitude and high Mach flight on the effect of wall boundary layer loss in the compressor. The accuracy for solution results by the application of the similarity criterion and parameter definition of the air inlet is compared with the existing experimental test result. The results indicate that the radial adverse pressure gradient in the rotor domain gradually increases along the span direction and decreases as flight Mach number increases; meanwhile, the circumferential adverse pressure gradient on the pressure side of the rotor blade is correspondingly larger and less than that on the suction side. In particular, the entropy increase along the streamwise shows a decreasing trend and an increasing trend inside the hub and shroud wall boundary layers, respectively. At 2.1 Ma, the entropy increase in the rotor domains enhances by 24.36–27.80% inside the shroud boundary layer, relative to the hub boundary layer; however, it decreases by 0.97–8.54% in the stator domain. With the increase in flight Mach number from 2.1 to 3.4, the average entropy increase reductions in the rotor domain decrease by 18.99–24.97% within the hub boundary layer and 5.71–8.1% within the shroud boundary layer. In the stator domain, it drops by 18.45–9.03% inside the hub boundary layer and 6.88–8.67% inside the shroud boundary layer. It was therefore found that, as Mach number increases from 2.1 to 3.4, the entropy increase reduction is larger inside the hub boundary layer than inside the shroud boundary layer.

Application of Phase-Locked PIV Technique to the Measurements of Flow Field in a Turbine Stage

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.

The Effect of Trailing Edge Clearance in Suppressing Hub-Corner Stall

Hub corner is a high-loss area located within in the blade passages of turbomachinery. It is well known that flow separation and vortex development in this area directly affects, not only energy loss and efficiency, but also the stability of axial compressors. In this paper, a high diffusion factor cascade is used in order to analyze the vortex structure which causes the hub-corner stall in cascade. The trailing edge clearance method is proposed for controlling the hub-boundary layer and this is compared with full uniform clearance. Our research highlights that the trailing edge clearance suppress the radial development of the suction separation vortex (SSV) in the hub corner and restrains the concentrated separation vortex (CSV). The influence of trailing edge clearance parameters and its effect on control of hub-corner stall is discussed in detail. It is revealed that clearance height avoids flow separation and suppresses radial vortex development, while the clearance length resists the overturning that is caused by transverse pressure gradient and controls the development of CSV. As the radial vortex plays a dominant role in hub-corner stall, clearance height has a better effect in controlling hub-corner stall in comparison to length at the constant flow area.

Effects of rising angle on upstream blades and intermediate turbine duct

With the improvement of requirement, design and manufacture technology, aero-engines for the future are characterized by further reduction in fuel consumption, cost, but increment in propulsion efficiency, which leads to ultra-high bypass ratio. The intermediate turbine duct (ITD), which connects the high pressure turbine (HPT) with the low pressure turbine (LPT), has a critical impact on the overall performances of such future engines. Therefore, it becomes more and more urgent to master the design technique of aggressive, even super-aggressive ITDs. Over the last years, a lot of research works about the flow mechanism in the diffuser ducts were carried out. Many achievements were reported, but further investigation should be performed. With the aid of numerical method, this paper focuses on the change of performance and flow field of ITD, as well as nearby turbines, brought by rising angle (RA). Eight ITDs with the same area ratio and length, but different RAs ranges from 8 degrees to 45 degrees, are compared. According to the investigation, flow field, especially outlet Ma of swirl blade is influenced by RA under potential effect, which is advisable for designers to modify HPT rotor blades after changing ITD. In addition to that, low velocity area moves towards upstream until the first bend as RA increases, while pressure loss distribution at S2 stream surface shows that hub boundary layer is more sensitive to RA, and casing layer keeps almost constant. On the other hand, the overall total pressure loss could keep nearly equivalent among different RA cases, which implies the importance of optimization.

STUDI EKSPERIMENTAL ALIRAN SEKUNDER UNTUK AXIAL COMPRESSOR CASCADE STAGGER KUAT DENGAN DAN TANPA TIP-CLEARANCE: DISTRIBUSI TEKANAN STATIS PADA PERMUKAAN BLADE

Secondary Flow Experimental Study for Axial Compressor Cascade Strong Stagger With and Without Tip-Clearance: Static Pressure Distribution on Blade Surface. The performance of blade cascade is influenced by thegrowth and boundary layer's separation along blade surface and endwall (casing and hub). The secondary flow whichhappens near hub and casing compressors is three dimentional flow separation phenomenon comes from interactionblade boundary layer with casing and hub boundary layers in the compressor. The secondary flow causes secondarylosses, blockage effect, and turning angle (deflection) distribution along blade span. The result of the research showsthat the increase of the angle of attack reveals lines of constan pressure to be curved forward (upstream) or pressuregives rise to spanwise caused by a strong curl flow. Separation is also increase and cross between two blades. Hence,three dimentional separation region in corner also increase.

Turbulence measurements in the inlet plane of a centrifugal compressor vaneless diffuser

Detailed flow measurements at the inlet of a centrifugal compressor vaneless diffuser are presented. The mean 3-d velocities and six Reynolds stress components tensor are used to determine the turbulence production terms which lead to total pressure loss. High levels of turbulence kinetic energy were observed in both the blade and passage wakes, but these were only associated with high Reynolds stresses in the blade wakes. For this reason the blade wakes mixed out rapidly, whereas the passage wake maintained its size, but was redistributed across the full length of the shroud wall. Peak levels of Reynolds stress occurred in regions of high velocity shear and streamline curvature which would tend to destabilize the shear gradient. Four regions in the flow are identified as potential sources of loss - the blade wake, the shear layers between passage wake and jet, the thickened hub boundary layer and the interaction region between the secondary flow within the blade wake and the passage vortex. The blade wakes generate most turbulence, with smaller contributions from the hub boundary layer and secondary flows, but no significant contribution is apparent from the passage wake shear layers.

Non-deterministic kinetic energy within the rotor wakes and boundary layers of low-speed axial fans: frequency-based decomposition of unforced unsteadiness and turbulence

A detailed analysis of the non-deterministic scales of the flow in low-speed axial turbomachinery has revealed the presence of significant large-scale unsteadiness, with periodic features different to the blade passing frequency (BPF), which are superimposed to the turbulent structures. Introducing a frequency-based decomposition, this additional component has been segregated from turbulent phenomena, so the total unsteadiness has been found to be contributed by three components: forced unsteadiness (deterministic scales), unforced unsteadiness (large-scale unsteadiness) and turbulence (small-scale randomness). Dual hot-wire anemometry has been employed intensively within the stage of a low-speed axial fan to provide a valuable experimental database, where the phase-locked averaging technique has been applied to retrieve time-resolved fluctuations and isolate the non-deterministic contribution. The present investigation shows that the presence of the unforced component is mainly related to instabilities of the rotor wakes and tip vortex structures, as well as wake–wake interactions. Moreover, typical eddies size of this component distribute their energy within the frequency range that contains 80% of the total unsteady kinetic energy. As a consequence, it is expected that large eddy simulation (LES) schemes with accurate spatial discretizations may address this component within the resolved scales of the filter, while unsteady Reynolds averaged Navier-Stokes (RANS) modelling could require additional modelling of the unforced mechanisms. In addition, maps and radial distributions of every component illustrate that major flow patterns are identifiable in all of them due to the redistribution of all-range scales throughout the energy cascade. Turbulent and forced mechanisms present important variations with the operating conditions, while the unforced component is barely affected by flow rate variations. It is shown that typical values of unforced unsteadiness reach up to 20% of the total unsteady energy, even for nominal conditions at midspan of the rotor passage. Higher levels of all components are found towards tip and hub boundary layers, as the total unsteadiness is reinforced by massive flow separations, tip blockages and major flow disturbances.

Cylindrical three-hole pressure probe calibration for large angular range

This paper presents a brand new method to calibrate a Cylindrical Three-Hole Probe (CTHP) in a “non-nulling” operating mode. This calibration method employs the definition of a new normalization factor used in the calibration coefficients of the probe. By means of this new normalization, singular points are prevented from appearing in the calibration coefficients. In addition, the angular range attainable when using this new calibration is ±70 ∘, while the typical angular range of traditional calibration is approximately ±30 ∘. To validate this study, an uncertainty analysis of the probe using this new method is carried out, resulting in significantly low uncertainties for the whole angular range of the probe. Complementarily, the influence of the angular distance between the holes of the probe on both angular range and maximum uncertainties is also performed in the paper. From this analysis, it is established that the optimal angular separation for the holes should be chosen to be between 50 and 60 degrees. Finally, to illustrate the usefulness of this procedure, a set of experimental measurements were conducted downstream of the rotor of an axial turbomachine, using the new normalization factor in a CTHP with an angular distance of 60 degrees between the holes. The results revealed that the complex, unsteady flow within the blade passages can be perfectly captured, including those regions with high variations in the flow angle, like the shear layers of the wake fluid or both casing and hub boundary layers. In summary, this newly developed calibration method enables the measurement of large variations of the flow angle, in particular, up to 140 degrees, increasing in 80 degrees the typical angular range of the traditional calibration.

Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows

A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.

Three-Dimensional Flow Measurements in Conical and Straight Wall Centrifugal Compressor Vaneless Diffusers

In the current work three-dimensional flow measurements in two types of centrifugal compressor vaneless diffuser were obtained using hot wire anemometry. The first diffuser was conical, designed to give a constant flow area, while the second straight wall diffuser had a constant axial width. Measurements of mean velocity, flow angle and velocity fluctuation level were obtained on eight cross-sectional planes in each diffuser. The jet-wake flow pattern and the impeller blade wakes are clearly visible at the inlet of both diffusers. Mixing out of the blade wake proceeds more rapidly in the straight diffuser. The hub boundary layer also develops more rapidly in this diffuser because of the adverse pressure gradient. Velocity fluctuation level measurements highlight the mixing regions within the diffusers. Recommendations are also made for the design of vaned diffusers. A larger vaneless space would be required with a straight wall diffuser and significant twisting of the vane would be required for both diffuser geometries if significant incidence losses are to be avoided.

Through-Flow Modeling of Axial Turbomachinery

Through-flow analysis, which is at the heart of the aerodynamic design of turbomachinery, requires as aerodynamic input a row-by-row description of the airfoil loss, deviation, and blockage. Loss and deviation have been investigated extensively in both cascades and rotating rigs as well as in numerous two- and three-dimensional analytical studies. Blockage, however, has received far less attention. As defined herein, blockage is a measure of the departure of the flow field from the condition of axisymmetry which is assumed in the through-flow analysis. The full-span blockage distributions calculated from measured single-stage rotor wake data were used to provide the input to the through-flow analysis, along with the measured full-span distributions of loss and deviation. Measured and computed results are compared for the single-stage rotor operating with both thick and thin inlet hub and tip boundary layers. It is demonstrated that both the level and the spanwise and streamwise distributions of blockage have a strong impact on the computed rotor exit flow field.

A Quasi-Three-Dimensional Blade Surface Boundary Layer Analysis for Rotating Blade Rows

A quasi-three-dimensional, finite difference boundary layer analysis for rotating blade rows has been developed which uses pressure distribution and streamline position data from a three-dimensional Euler equation solver. This analysis uses as coordinate lines the blade normal vector, the local inviscid streamline direction and a crossflow coordinate tine perpendicular to both normal and streamline coordinate lines. The equations solved may be determined either by assuming the crossflow velocity to be small or that its variation in the crossflow direction is small. Thus the analysis would not apply to a region where the boundary layer character changes rapidly such as a corner but could be expected to provide good results away from hub or tip casing boundary layers. Modified versions of Keller’s box scheme are used to solve the streamwise and crossflow momentum equations as well as the energy equation. Results are presented for a high-tip speed, low aspect ratio rotor designed by NASA Lewis Research Center which show that the three-dimensional boundary layer separates significantly sooner and has a much larger influence on rotor performance than would be expected from a two-dimensional analysis.

Treatment of the Annulus Wall Boundary Layer Using a Secondary Flow Hypothesis

Hitherto, theories of annulus wall boundary layer development in axial compressors have assumed an axially-symmetric flow in which the blade action has been replaced by a force field. A more rigorous treatment of the momentum equations in the annulus boundary layer by Mellor and Wood demonstrated the presence of certain terms, after the equations had been averaged in the pitch-wise direction, which arise from the truly three-dimensional character of the flow. These terms, which may be described as the gradients of apparent stresses, were not regarded by them (apart from a discussion of tip clearance) as having importance for the problem. In the present work a second equation of the annulus wall boundary layer is obtained by consideration of the work of these apparent stresses. By integration of the system of equations over a single blade row, two equations are obtained relating various integral quantities at inlet to and exit from the row. Each equation contains terms which depend upon apparent stresses connected with the relative velocity field at the exit plane. An experiment is described in which the six turbulent stresses in the stationary frame downstream of a single rotor, determined by means of a multiple hot wire array, are used to evaluate each term of the aforementioned equations. The integral quantities thus determined are shown to be reasonably consistent with the predictions from the two equations, in particular, for the case of the hub boundary layer. Theoretical solutions of the two integral equations require a secondary flow hypothesis so that the departure from collateral flow at blade row exit is determined by the solution.

End-Wall Boundary Layers in Axial Compressors

Various aspects of an end-wall boundary-layer theory in axial compressors, as formulated originally by Mellor and Wood, are analyzed. This analysis contains three parts. In the first part the original formulation is used with a mainstream flow calculation in order to predict overall performances on a single stage compressor. Comparison with experimental data shows the limitation of the assumption of constant shape factor and skin-friction coefficient. In the second part, through an analysis of existing detailed experimental profiles of flow properties behind blades, empirical values for the coefficients of the theory are deduced. It appears that the skin-friction coefficient takes up unusual high values near the tip of the rotor blades and that the hub boundary layer at the outlet is not collateral. In the last part, the basic equations are reformulated in streamline coordinates. This should allow eventual incorporation of the shape factor variation and the Reynolds number effects as well as the introduction of the wall skewing angle. A modification of the original closure equation is proposed allowing for the noncollateral character of the boundary layer at the blade outlet. The original formulation appears as a limit for low skewing angles.