Inlet Boundary Layer Research Articles

Effects of Unsteady Wakes on the Secondary Flows in the Linear T106 Turbine Cascade

In modern low pressure turbines the efforts to increase aerodynamic blade loading by increasing blade pitch and optimizing midspan performance in order to reduce weight and complexity can produce increased losses in the endwall region. Airfoils of high flow turning and high pressure gradients between the blades generate strong secondary flows which impair the global aerodynamic performance of the blades. In addition, the unsteady incoming wakes take influence on transition phenomena on the blade surfaces and the inlet boundary layer, and consequently affect the development and the evolution of the secondary flows. In this paper, the T106 cascade is used to identify the effect of unsteady wakes on the development of secondary flows in a turbine cascade. Numerical and experimental results are compared at different flux coefficients and Strouhal numbers, the relative differences and similarities are analyzed.

Numerical investigation of inlet boundary layer skew in axial-flow compressor cascade and the corresponding non-axisymmetric end wall profiling

This paper investigates the influence of boundary layer skew on flow structure, total pressure loss, and flow control technique numerically on a high-loaded axial-flow compressor cascade. We have developed two new models respectively about loss evaluation and end wall flow control mechanism for more specific analysis. The result shows that boundary layer skew weakens the secondary flow and delays the generation of passage vortex when incidence approaches 0°. This results in a reduction of total pressure loss mainly (89.4%) due to relieved corner separation. However, as incidence exceeds a certain value (+7°), severe corner separation or even earlier corner stall can be induced by inlet boundary layer skew. Optimization procedure for profiled end wall at inflow condition of +7° incidence is further carried out to investigate the impact of boundary layer skew on flow control technique. The result shows that boundary layer skew should be counted in the optimization design of profiled end wall because of its significant influence on the development of end wall flow. The optimum profiled end walls for cases with and without boundary layer skew show great difference in the manner of end wall flow control. According to the improvement of cascades’ performance, end wall profiling seems more efficient in reducing loss when influenced by the boundary layer skew.

End Wall Loss Reduction of High Lift Low Pressure Turbine Airfoils Using Profile Contouring—Part II: Validation

The hypothesis, posed in Part I, that excessive end wall loss of high lift low pressure turbine (LPT) airfoils is due to the influence of high stagger angles on the end wall pressure distribution and not front loading is evaluated in a linear cascade at Re = 100,000 using both experimental and computational studies. A nominally high lift and high stagger angle front-loaded profile (L2F) with aspect ratio 3.5 is contoured at the end wall to reduce the stagger angle while maintaining the front loading. The contouring process effectively generates a fillet at the end wall, so the resulting airfoil is referred to as L2F-EF (end wall fillet). Although referred to as a fillet, this profile contouring process is novel in that it is designed to isolate the effect of stagger angle on end wall loss. Total pressure loss measurements downstream of the blade row indicate that the use of the lower stagger angle at the end wall reduces mixed out mass averaged end wall and passage losses approximately 23% and 10%, respectively. This is in good agreement with computational results used to design the contour which predict 18% and 7% loss reductions. The end wall flow field of the L2F and L2F-EF models is measured using stereoscopic particle image velocimetry (PIV) in the passage. These data are used to quantify changes in the end wall flow field due to the contouring. PIV results show that this loss reduction is characterized by reduced inlet boundary layer separation as well as a change in strength and location of the suction side horseshoe vortex (SHV) and passage vortex (PV). The end wall profile contouring also produces a reduction in all terms of the Reynolds stress tensor consistent with a decrease in deformation work and overall flow unsteadiness. These results confirm that the stagger angle has a significant effect on high-lift front-loaded LPT end wall loss. Low stagger profiling is successful in reducing end wall loss by limiting the development and migration of the low momentum fluid associated with the SHV and PV interaction.

Effects of Inlet Turbulence and End-Wall Boundary Layer on Aerothermal Performance of a Transonic Turbine Blade Tip

Most of the previous researches of inlet turbulence effects on blade tip have been carried out for low speed situations. Recent work has indicated that for a transonic turbine tip, turbulent diffusion tends to have a distinctively different impact on tip heat transfer than for its subsonic counterpart. It is hence of interest to examine how inlet turbulence flow conditioning would affect heat transfer characteristics for a transonic tip. The present work is aimed to identify and understand the effects of both inlet freestream turbulence and end wall boundary layer on a transonic turbine blade tip aerothermal performance. Spatially-resolved heat transfer data are obtained at aerodynamic conditions representative of a high-pressure turbine, using the transient infrared thermography technique with the Oxford High-Speed Linear Cascade research facility. With and without turbulence grids, the turbulence levels achieved are 7%–9% and 1%, respectively. On the blade tip surface, no apparent change in heat transfer was observed with high and low inlet turbulence intensity levels investigated. On the blade suction surface, however, substantially different local heat transfer distributions for the suction side near tip surface have been observed, indicating a strong local dependence of the local vortical flow structure on the freestream turbulence. These experimentally observed trends have also been confirmed by CFD examinations using the Rolls-Royce HYDRA. A further CFD analysis suggests that the level of inflow turbulence alters the balance between the passage vortex associated secondary flow and the over tip leakage (OTL) flow. Consequently, an enhanced inertia of near wall fluid at a higher inflow turbulence weakens the cross-passage flow. As such, the weaker passage vortex leads the tip leakage vortex to move further into the mid passage, with the less spanwise coverage on the suction surface, as consistently indicated by the heat transfer signature. Different inlet end wall boundary layer profiles are employed in the computational study with HYDRA. All CFD results indicate the inlet boundary layer thickness has little impact on the heat transfer over the tip surface as well as the pressure side near-tip surface. However, noticeable changes in heat transfer are observed for the suction side near-tip surface. Similar to the inlet turbulence effect, such changes can be attributed to the interaction between the passage vortex and the OTL flow.

Numerical Investigations on Effects of Inlet Boundary Layer Characteristics on Aspirated Compressor Cascade

Numerical Investigations on Effects of Inlet Boundary Layer Characteristics on Aspirated Compressor Cascade

Vortex control and aerodynamic performance improvement of a highly loaded compressor cascade via inlet boundary layer suction

Effects of inlet boundary layer suction on the vortex structure and cascade loss in a highly loaded compressor cascade were investigated experimentally. Ink-track visualization was undertaken on cascade endwall and the blade surface. Ten traverse planes from upstream to downstream of the cascade in a rectangular wind tunnel were measured by an L-shaped five-hole probe. These tested planes revealed the process of emergence, development and decline of several principal vortices as well as the corresponding additional losses. Details of ink-track visualization displaying the secondary flow behavior of boundary layer upon endwall and blade surface assist to make judgment on vortex evolution. Inspection of the vortex structure revealed that highly loaded compressor was characterized by large-scale vortices in the endwall region. After suction, these vortices are all well organized and under control. Among all of them, passage vortex is most sensitive to the variation of the inlet boundary layer, and its main function is to spread low-energy fluid rather than to produce loss. On the other hand, a wall vortex and a concentrated shedding vortex take place inside and after the cascade, respectively, and engender considerable accompanying loss as they dissipate. The effects of inlet boundary layer suction on them are correspondingly weaker. About one forth of the total loss in the baseline cascade was eliminated when boundary layer suction flow rate reaches 2.5 % of the inlet mass flow. The feasibility of simplifying the suction system is also verified through this work.

1023 Condition for Generating an Isolated Turbulent Patch at an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Injection

1023 Condition for Generating an Isolated Turbulent Patch at an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Injection

Laminar-Turbulent Transition of an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Injection (Turbulence within Isolated Turbulent Patches and Its Growth Mechanism)

Laminar-Turbulent Transition of an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Injection (Turbulence within Isolated Turbulent Patches and Its Growth Mechanism)

Laminar-Turbulent Transition of an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Injection (Shape of Isolated Turbulent Patches and Their Growth)

The laminar-turbulent transition of a boundary layer induced by an injection of jet in the inlet region of a circular pipe was experimentally investigated. The jet was periodically injected radially from a small hole in the inlet region into the pipe flow. Axial velocity was measured by a hot-wire anemometer. The turbulence induced by the jet within the boundary layer developed into turbulent patches which then grew in the axial, circumferential and radial directions downstream. The shape of the patches shown by the intermittency factor in the diametrical plane was similar to the turbulent spot in the flat plate boundary layer. Their property became similar to the turbulent slug in the pipe flow developed downstream. The turbulent patches protruded from the boundary layer after they grew and reached the circumferential opposite side, although they stayed within the boundary layer as long as the shape was turbulent spot-like in the diametrical plane. The propagation velocity at the leading edge became faster than the bulk velocity, though it turned slower at the trailing edge. Therefore, the growth rate of its streamwise length varied downstream. The growth rate of the patches' circumferential length was smaller than that in the turbulent spot under zero pressure gradient and was almost the same as the spot under favorable pressure gradient.

Verification of FLUENT in the Computation of Supersonic Inlet

This paper analyzes the simulation results obtained by using FLUENT 6.1, simulation being carried out on supersonic inlet. A combination of turbulence models are adoptedsuch as k-RNG k- and S-A. The computation outcomes are checked by experiment results. Detailed analysis demonstrates that the numerical results obtained from the turbulence models can accurately forecast the pressure distribution trend in the inlet. However, for separation bubbles due to the mutual interference of the shock wave within the inlet boundary layer, the calculation predicts fewer bubbles than actually generated in the experiment. The RNG k- and SA turbulence models produces clearer wave structures in the channel.

Flow Measurements in a First Stage Nozzle Cascade Having Endwall Contouring, Leakage, and Assembly Features

This work supports new gas turbine designs for improved performance by evaluating the use of endwall contouring in a cascade that is representative of a first stage stator passage. Contouring accelerates the flow, reducing the thickness of the endwall inlet boundary layer to the turbine stage and reducing the strength of secondary flows within the passage. The reduction in secondary flows leads to less mixing in the endwall region. This allows for an improved cooling of the endwall and airfoil surfaces with injected and leakage flows. The present paper documents the component misalignment and injected and leakage flow effects on the aerodynamic losses within a passage that has one contoured and one straight endwall. Steps and injected flows within the passage can lead to thicker endwall boundary layers, stronger secondary flows, and possibly additional vortex structures in the passage. The paper compares losses with various steps, gaps, and leakage flows to assess their importance in this contoured passage. In particular, features associated with the combustor-to-turbine transition piece and the slashface on the vane platform are addressed. An n-factorial study is used to quantify the importance of such effects on aerodynamic losses.

Reduction of Tip Clearance Losses in an Unshrouded Turbine by Rotor Casing Contouring

Tip leakage flow is responsible for a number of aerodynamic losses in turbines, and the mixing of tip leakage flow with the passage secondaryflowcauses significant losses. This paperpresents a numerical investigation of the effect of three types of casing contouring: 1) step, 2) trench, and 3) arc casings on aerodynamic performance of anunshrouded turbine rotor at design and off-design incidences, in an attempt to seek an optimal casing contour with the objective both of avoiding direct tip leakage in the tip clearance, and of efficiently using the interaction between the tip leakage vortex and passage vortex inside rotor passage, and then to reduce the total losses in turbines. The results show that the reduction in tip leakage losses obtained by the optimized casing contour significantly overrides the losses caused by the step in theflowpath. Introduction of casing step thickens the inlet boundary layer, which increases the strength of the tip passage vortex and hence suppresses the evolution of the tip leakage vortex. Step and trench casings provide the largest efficiency increase when the inlet flow direction is normal to the casing step while arc casing achieves the best performance improvement at the design incidence.

The Impact of Real Geometries on Three-Dimensional Separations in Compressors

This paper considers the effect of small variations in leading edge geometry, leading edge roughness, leading edge fillet, and blade fillet geometry on the three-dimensional separations found in compressor blade rows. The detrimental effects of these separations have historically been predicted by correlations based on global flow parameters, such as blade loading, inlet boundary layer skew, etc., and thus ignoring small deviations such as those highlighted above. In this paper it is shown that this may not be the case and that certain, engine representative geometry deviations can have an effect equivalent to an increase in blade loading of 10%. Experiments were performed at the stator hub of a low-speed, single-stage compressor. The results show that any deviation which causes suction surface transition to move to the leading edge over the first 30% of span will cause a large growth in the size of the hub separation, doubling its impact on loss. The geometry deviations that caused this, and are thus of greatest concern to a designer, are changes in leading edge quality and roughness around the leading edge, which are characteristic of an eroded blade.

Influence of Chord Length and Inlet Boundary Layer on the Secondary Losses of Turbine Blades

In this paper, results concerning the influence of chord length and inlet boundary layer thickness on the endwall loss of a linear turbine cascade are discussed. The investigations were performed in a low speed cascade tunnel using the turbine profile T40. The turning of 90 deg and 70 deg, the velocity ratio in the cascade from 1.0 to 3.5 as well as the chord length of 100 mm, 200 mm, and 300 mm were specified. In a measurement distance of one chord behind the cascade in main flow direction, an approximate proportionality of endwall loss and chord was observed in a wide range of velocity ratios. At small measurement distances (e.g., s2/l=0.4), this proportionality does not exist. If a part of the flow path within the cascade is approximately incorporated, a proportionality to the chord at small measurement distances can be obtained, too. Then, the magnitude of the endwall loss mainly depends on the distance in main flow direction. At velocity ratios near 1.0, the influence of the chord decreases rapidly, while at a velocity ratio of 1.0, the endwall loss is independent of the chord. By varying the inlet boundary layer thickness, no correlation of displacement thickness and endwall loss was achieved. A calculation method according to the modified integral equation by van Driest delivers the wall shear stress. Its influence on the endwall loss was analyzed.

Effects of Inlet Boundary Layer Thickness on Characteristics of Rotating Stall in Subsonic Axial Compressor

A 3D numerical simulation was conducted to study the effect of inlet boundary layer thickness on rotating stall in an axial compressor. The inlet boundary layer thickness had significant effects on the hub-corner-separation at the corner of hub and suction surfaces. The hub-corner-separation grew considerably for a thick inlet boundary layer as the load increased, while it diminished to become indistinguishable from the rotor wake for a thin inlet boundary layer and another corner-separation originated near the casing. This difference in the internal flow near stall also had a large effect on characteristics of the rotating stall, especially the initial asymmetric disturbance and the size of stall cells. While a pre-stall disturbance arises firstly in the hub-corner-separation for the thick inlet boundary layer, an asymmetric disturbance was initially generated in the tip region because of the corner-separation for the thin inlet boundary layer. This disturbance was transferred to the tip leakage flow and grew to become an attached stall cell, which adheres to the blade passage and rotates at the same speed as the rotor. When this attached stall cell reached a critical size, it started moving along the blade row and became a short-length-scale rotating stall. The size of the stall cell for the thick inlet boundary layer was larger than for the thin inlet boundary layer. Due to the bigger size of the stall cell, the performance of the single rotor for the former case dropped more significantly than for the latter case.

Laminar-Turbulent Transition of an Inlet Boundary Layer in a Circular Pipe Induced by Periodic Ejection (Condition for Generating an Isolated Turbulent Patch)

The laminar-turbulent transition of a boundary layer induced by an ejection of jet in the inlet region of a circular pipe was experimentally investigated. The jet was periodically inserted radially from a small hole in the inlet region into the pipe flow. Axial velocity was monitored by a hot-wire anemometer. The difference of properties in laminar-turbulent transition from developed Hagen-Poiseuille flow was examined. Isolated turbulence patches were generated by the jets, and then they propagated downstream. The leading edge of the turbulent patch was definite, whereas its trailing edge was not. This characteristic was similar to that of a turbulent spot in a flat-plate boundary layer. The threshold value of the jet flow rate to generate the turbulent patch was then obtained. The threshold value decreased and saturated finally with the increase in jet duration. The normalized jet duration when the threshold value was saturated increased with the increase in Reynolds number, contrary to the developed region. The normalized threshold flow rate tended to vary with the Reynolds number among three regions. All tendencies were different from those of the developed region. With the increase in jet flow rate beyond the threshold value, the duration of the turbulent patch increased, though the fluctuating velocity within the patch did not increase. The propagation velocities of the leading and trailing edges, the duration and fluctuating velocity within the turbulent patch were almost constant irrespective of the jet ejection frequency.

Direct numerical simulation of a turbulent boundary layer up to Re θ = 2500

Direct numerical simulations (DNSs) of a turbulent boundary layer (TBL) with Re θ = 570–2560 were performed to investigate the spatial development of its turbulence characteristics. The inflow simulation was conducted in the range Re θ = 570–1600 by using Lund’s method. To resolve the numerical periodicity induced by the recycling method, we adopted a sufficiently long streamwise domain of x/ θ in,i = 1000 (=125 δ 0 ,i ), where θ in,i is the inlet momentum thickness and δ 0 ,i is the inlet boundary layer thickness in the inflow simulation. Furthermore, the main simulation with a length greater than 50 δ 0 was carried out independently by using the inflow data, where δ 0 is the inlet boundary layer thickness of the main simulation. The integral quantities and the first-, second- and higher-order turbulence statistics were compared with those of previous data, and good agreement was found. The present study provides a useful database for the turbulence statistics of TBLs. In addition, instantaneous field and two-point correlation of the streamwise velocity fluctuations displayed the existence of the very large-scale motions (VLSMs) with the characteristic widths of 0.1–0.2 δ and that the flow structure for a length of approximately ∼6 δ fully occupies the streamwise domain statistically.

Investigation of turbulent boundary layer over forward-facing step via direct numerical simulation

This paper presents observations and investigations of the detailed turbulent structure of a boundary layer over a forward-facing step. The present DNSs are conducted under conditions with three Reynolds numbers based on step height, or three Reynolds numbers based on momentum thickness so as to investigate the effects of step height and inlet boundary layer thickness. DNS results show the quantitative turbulent statistics and structures of boundary layers over a forward-facing step, where pronounced counter-gradient diffusion phenomena (CDP) are especially observed on the step near the wall. Also, a quadrant analysis is conducted in which the results indicate in detail the turbulence motion around the step.

Numerical Study of Effect of Streamwise End Wall Fences on Secondary Flow Losses in Two Dimensional Turbine Rotor Cascade

:In a turbine cascade flow losses due to secondary flow are nearly 33% of total losses. In the present study effect of streamwise end wall fence on secondary flow loss reduction is studied numerically with experimental validation. A fence was fixed normal to the end wall and at half pitch away from the blades. Curvature of the fence was the same as that of the blade camberline. Optimization of the fence height for minimum secondary losses is done numerically using coefficient of secondary kinetic energy as the optimization parameter. A fence of height 1/3rd of inlet boundary layer thickness (δ) was found to be optimum and produces a total pressure loss reduction of 15% compared to cascade without fence. This fence trapped the pressure side leg of horseshoe vortex and reduced maximum overturning near the end wall compared to baseline case. A counter-rotating vortex is produced by this fence which weakens the passage vortex and trailing edge vortex. Climbing of passage vortex on suction surface reduced drastically. Fence of height 1/3rd of inlet boundary layer thickness reduced the three dimensionality of the flow and so the losses.

1단 축류 가스터빈내 동익의 허브면에 장착된 경계층 펜스의 효과에 대한 수치 해석적 연구

The objective of this study is to investigate the three-dimensional turbulence flow characteristics of a rotor passage of an one-stage axial flow gas turbine and to investigate the effects of a boundary layer fence installed on the hub endwall of the rotor passage. Secondary flows occurring within the rotor passage (e.g. horseshoe vortex, passage vortex, and cross flow) cause secondary loss and reduce turbine efficiency. To control these secondary flows, a boundary layer fence measuring half the height of the thickness of the inlet boundary layer was installed on the hub end wall of the rotor passage. This study was performed numerically. The results show that the wake and secondary flows generated by the stator reduced the rotor load to constrain the development of cross flow and secondary flow reinforced by the rotor passage. In addition, the secondary vortices occurring within the rotor passage were reduced by the rotation of the rotor. Although, the boundary layer fence induced additional vortices, giving lise to an additional loss of turbine, its presence was shown to reduce the total pressure loss when compared to effects of the case without fence regardless of the relative position of blades by enervating secondary vortices occurred within the rotor passage.