Tip Clearance Vortex Research Articles

Experimental Studies of the Endwall Flow in a Multistage Low-Speed Axial Compressor

Detailed flowfield measurements are made in the end-wall region of a four-stage low-speed large-scare axial compressor test rig. Two ten-hole survey boundary layer probes, one four-hole pneumatic probe, and six dynamic pressure probes each bedded with a high-frequence response Kulite sensor, are designed and manufactured to aid the test. Boundary layer region and mainflow region can be clearly distinguished from the boundary layer measurement results. A few parameters, including boundary layer thickness, displace thickness, momentum loss thickness, energy loss thickness and blockage coefficient are calculated based on the measurements. Results from turbulent boundary layer empirical formula of plate are in good agreement with the mesurements, indicating a way of boundary layer effect estimation. The evolution trajectory of TCV is identified evidently by dynamic pressure measurements. The maximum aerodynamic load point on the blade moves from the trailing to the leading edge as flowrate decreases, which leads to the forward movement of the tip clearance vortex inception point.

Blade Loading Effects on Axial Turbine Tip Leakage Vortex Dynamics and Loss

Abstract Numerical simulations have been carried out to define the loss generation mechanisms associated with tip leakage in unshrouded axial turbines. Tip clearance vortex dynamics are a dominant feature of two mechanisms important in determining this loss: (i) decreased swirl velocity due to vortex line contraction in regions of decreasing axial velocity, i.e., adverse pressure gradient, and (ii) vortex breakdown and reverse flow in the vortex core. The mixing losses behave differently from the conventional view of flow exiting a turbine tip clearance. More specifically, it is shown through control volume arguments and computations that as a swirling leakage flow passes through a pressure rise, such as in the aft portion of the suction side of a turbine blade, the mixed-out loss can either decrease or increase. For turbines, the latter typically occurs if the deceleration is large enough to initiate vortex breakdown, and it is demonstrated that this can occur in modern turbines. The effect of blade pressure distribution on clearance losses is illustrated through computational examination of turbine blades with forward loading at the tip and with aft loading. A 15% difference in leakage loss is found between the two due to lower clearance vortex deceleration (lower core static pressure rise) with forward loading and, hence, lower vortex breakdown loss. Additional computational experiments, carried out to define the effects of blade loading, incidence, and solidity, are found to be consistent with the proposed ideas linking blade pressure distribution, vortex breakdown, and turbine tip leakage loss.

Investigation of the tip leakage flow at turbine rotor blades with squealer cavity

Understanding of the tip leakage flow (TLF) in turbine rotors is one key aspect in the design for improving the efficiency of turbines. This requires measurements and simulations of the TLF, especially when investigating new rotor blade designs with blade tip treatments. However, flow measurements in the tip gap of a rotating machine are highly challenging because of the small gap size of about 1 mm and the high unsteadiness of the flow requiring a high temporal resolution of about 10 μs. For this purpose, an optimized non-intrusive measurement concept based on frequency modulated Doppler global velocimetry is presented, which fulfills the requirements. Three component velocity fields of the TLF were obtained in a turbine test rig at a blade passing frequency of 930 Hz. The rotor blades were equipped with a squealer tip, and the TLF in the squealer cavity region was successfully measured. The measurement agrees well with calculated results showing gradients in the tip gap above the squealer cavity. Furthermore, the development of the tip clearance vortex was resolved at the suction side of the blades.

Study of Near-Stall Flow Behavior in a Modern Transonic Fan With Compound Sweep

Detailed flow behavior in a modern transonic fan with a compound sweep is investigated in this paper. Both unsteady Reynolds-averaged Navier-Stokes (URANS) and large eddy simulation (LES) methods are applied to investigate the flow field over a wide operating range. The calculated flow fields are compared with the data from an array of high-frequency response pressure transducers embedded in the fan casing. The current study shows that a relatively fine computational grid is required to resolve the flow field adequately and to calculate the pressure rise across the fan correctly. The calculated flow field shows detailed flow structure near the fan rotor tip region. Due to the introduction of compound sweep toward the rotor tip, the flow structure at the rotor tip is much more stable compared to that of the conventional blade design. The passage shock stays very close to the leading edge at the rotor tip even at the throttle limit. On the other hand, the passage shock becomes stronger and detaches earlier from the blade passage at the radius where the blade sweep is in the opposite direction. The interaction between the tip clearance vortex and the passage shock becomes intense as the fan operates toward the stall limit, and tip clearance vortex breakdown occurs at near-stall operation. URANS calculates the time-averaged flow field fairly well. Details of measured rms static pressure are not calculated with sufficient accuracy with URANS. On the other hand, LES calculates details of the measured unsteady flow features in the current transonic fan with compound sweep fairly well and reveals the flow mechanism behind the measured unsteady flow field.

Investigation of the Tip Clearance Flow in a Compressor Cascade Using a Novel Laser Measurement Technique With High Temporal Resolution

The understanding of the tip clearance flow in axial compressors is a key issue for developing new compressors with enhanced efficiency and reduced noise for instance. However, necessary flow measurements in the blade tip region and within the tip clearance are challenging due to the small gap width. The application of a novel optical measurement technique named Doppler global velocimetry with laser frequency modulation is presented, which provides velocity field measurements of all three velocity components nonintrusively in the tip clearance flow of a linear cascade at near stall conditions. These array measurements have a high temporal resolution enabling turbulence analysis such as the evaluation of velocity standard deviations and turbulence spectra up to several kilohertz. Conventional pneumatic and hot-wire measurements in planes at the inlet and the outlet as well as on the blade surface were taken to complete the flow pattern and validate the data of the Doppler global velocimetry. Wake measurements identified a strong flow separation in the rear suction side dominating the transient character of the cascade flow. Towards the endwall this high loss region is reduced by the clearance flow and the resulting vortex, which is obviously not affected by the profile separation and the pulsating blockage frequency. Inside the blade passage and the tip clearance the Doppler global velocimetry measurements allowed a spatial assignment of the origin of the tip leakage flow and the downstream developing vortex. In addition, the tip clearance vortex could be resolved and identified successfully as the most dominant turbulence generating effect in the near endwall region at this high loading operating point of the blading.

Active control of tip clearance flow through casing air injection in axial turbines

This paper presents a numerical investigation of an effective method for controlling tip clearance flow in axial turbines. Cooling air is injected into the rotor passages from the casing wall through 10 discrete holes per rotor pitch, in a direction that is opposite to the tip clearance flow. Emphasis is played on the analysis of the complicated, three‐dimensional flow structures within the tiny tip gap, which is caused by the interaction between tip clearance flow and the high pressure injection flow. The results indicate that due to the obstruction by injection, less passage flow is entrained into tip clearance in all the cases considered. With air injection, the associated losses of both the tip clearance vortex and the tip passage vortex are reduced significantly. The heat transfer condition is improved noticeably caused by the cooling air. Besides that, it can be found that the more downstream the injection is distributed, the more the tip clearance vortex is influenced and the less thermal benefits obtained from the cooling jet. In the current study, the optimum isentropic efficiency occurs when 1% of the passage mass flow is injected at 30% of the axial chord from the leading edge, which has been improved by 0·3341%.

An Investigation on the Effect of Endwall Movement on the Tip Clearance Loss Using Annular Turbine Cascade

The aerodynamic losses in gas turbines are mainly caused by profile loss secondary flow, and tip leakage loss. This study focuses on tip leakage flow of high-pressure turbine stages. An annular turbine cascade was constructed with fixed blades on the casing, and the distance between blade tip and the hub was considered as tip clearance gap. The effect of endwall movement on loss mechanism was investigated by using experimental and numerical techniques. The measurements were obtained while the hub was fixed but the numerical calculations were carried out for both stationary and moving cascades. Upstream and downstream flows were measured by using a calibrated five-hole pressure probe. The steady incompressible turbulent flow was obtained by solving Reynolds averaged Navier-Stokes equations and by employing shear stress transport (SST)k-ωturbulence model. The total pressure loss coefficient obtained from the numerical technique was compared with the experimental measurements, and the comparison showed good agreement. Tip clearance vortices were observed in the tip clearance gap. It was found through this study that end-wall movement reduces tip leakage loss through the cascade.

Experiments and Computations on Large Tip Clearance Effects in a Linear Cascade

Large tip clearances typically in the region of 6% exist in the high pressure (HP) stages of compressors of industrial gas turbines. Due to the relatively short annulus height and significant blockage, the tip clearance flow accounts for the largest proportion of loss in the HP. Therefore increasing the understanding of such flows will allow for improvements in design of such compressors, increasing efficiency, stability, and the operating range. Experimental and computational techniques have been used to increase the physical understanding of the tip clearance flows through varying clearances in a linear cascade of controlled-diffusion blades. This paper shows two unexpected results. First the loss does not increase with clearances greater than 4% and second there is an increase in blade loading toward the tip above 2% clearance. It appears that the loss production mechanisms of the pressure driven tip clearance jet do not increase as the clearance is increased to large values. The increase in blade force is attributed to the effect of the strong tip clearance vortex, which does not move across the blade passage to the pressure surface, as is often observed for high stagger blading. These results may be significant for the design of HP compressors for industrial gas turbines.

Numerical investigation of active tip-clearance control through tip cooling injection in an axial turbine cascade

This paper presents a numerical investigation of an active tip-clearance control method based on cooling injection from the blade tip surface. It aims to study the influences of air injection on controlling tip clearance flow, with emphasis on the effects of the injection location on secondary flow and the potential thermal benefits from the cooling jet. The results show that injection location plays an important role in the redistribution of secondary flow within the cascade passage. Injection located much closer to the pressure-side corner performs better in reducing tip clearance massflow and its associated losses. However, it also intensifies tip passage vortex, due to less restraint deriving from the reduced tip clearance vortex. Lower plenum total pressure is required to inject equivalent amount of cooling air, but the heat transfer condition on the blade tip surface is a bit worse than that with injection from the reattachment region. Thus the optimum location of air injection should be at the tip separation vortex region.

Parametric study of tip cooling injection in an axial turbine cascade: Influences of injection circumferential angle

This article presents a parametric investigation of an innovative tip-clearance control method. Cooling air is injected into the tip clearance to obstruct tip clearance flow, from the blade tip surface through a row of 11 equidistant holes. In order to demonstrate the influences of injection circumferential angle, four different test cases are performed and compared with the baseline case without air injection. The results indicate that injection circumferential angle plays an important role in the redistribution of secondary flow within the cascade passage. Injection at a smaller circumferential angle performs better in reducing tip clearance mass flow and its associated losses. However, it also intensifies the tip passage vortex much more, because of the reduced restraint deriving from the weaker tip clearance vortex. It can also be found that flow under-turning caused by the tip clearance vortex can be reduced greatly with a smaller circumferential angle. But the heat transfer condition is worse than that when cooling air is injected perpendicularly to the blade tip surface. Besides that, a modified tip clearance loss model is observed to be able to perform well in predicting the losses induced by tip clearance flow at all the conditions considered in this article.

Unsteady behaviour of the tip clearance vortex in a rotor equivalent compressor cascade

From earlier experimental investigations in a single-stage axial-flow pump and different numerical calculations of the flow in single-stage axial-flow compressors, it is known that vortex breakdown of the tip clearance vortex can take place in turbomachines, although an experimental proof for subsonic compressors is lacking. Vortex breakdown, if existent, is a source of high instability in the sensitive tip region of axial-flow pumps and compressors and will also play an important role in the stall inception process. Therefore, the flow in a linear compressor cascade with a 3 per cent tip clearance to one side has been investigated at different flow angles from the design point up to the stability limit of the cascade. The cascade resembles the tip section of a single-stage, axial-flow, low-speed compressor that is also in use at the Technical University of Braunschweig. The measuring techniques used were (a) a commercial particle image velocimetry (PIV) system and (b) a pressure measuring system with several flush mounted high-response pressure transducers at selected locations where the vortex was expected. As the cascade approaches its stall limit, the analysis of the pressure signals in the frequency domain revealed a bump of increased amplitude at a certain non-dimensional frequency for some of the measuring positions. The measuring positions that exhibited the bump correlated very well with a paraboloid-shaped region of high standard deviation enveloping an area of very low momentum fluid. It is shown that the frequency of the striking bump corresponds to the rotational frequency of the vortex calculated from the PIV measurements.

Recent German Research on Periodical Unsteady Flow in Turbomachinery

The main source of flow unsteadiness in turbomachinery is the aerodynamical interaction of the rotor and stator blade rows. The blades and vanes, moving relatively to each other, interact because of the viscous wakes and the potential effects of the blades. In addition, the wakes and potential effects superimpose with other flow patterns, for instance the tip clearance vortices and other secondary flow phenomena. Furthermore in transonic compressors the interaction of wakes and shocks plays an important role. As a result, the real flow field is highly periodically unsteady and very complex, especially in multistage turbomachinery. Although this fact has received increasing attention within recent years, blade row interactions effects are not yet typically addressed in current design systems of turbomachinery. Actually, there is a requirement of the ability of modern design methods to predict unsteady flow features. With increasing aerodynamic loading of the blades and higher Mach numbers the consideration of rotor-stator-interactions gains in importance. It is therefore one of the challenges of the present and future design of compressors and turbines to include beneficial unsteady effects to improve the engine parameters. This requires a detailed physical understanding of the unsteady flow field and the resulting effects on the performance and flow stability. In 2000 the joint research program “Periodical Unsteady Flow in Turbomachinery” was initiated. Partners of this project are five research groups from four german universities: Technische Universitat Berlin (Prof. Hourmouziadis), Universitat der Bundeswehr Munchen (Prof. Fottner, Prof. Pfitzner), Universitat Karlsruhe: Institut fur Thermische Stromungsmaschinen (Prof. Wittig, Dr. Dullenkopf), Institut fur Hydromechanik (Prof. Rodi), Technische Universitat Dresden (Prof. Vogeler), and a research group from the: German Aerospace Centre (Deutsches Zentrum fur Luft- und Raumfahrt), (Prof. Weyer, Prof. Monig). This 5-year program was funded by the German Research Foundation (Deutsche Forschungsgemeinschaft) and coordinated by Professor Hourmouziadis (Technische Universitat Berlin). The aim of this joint project is to contribute to an improved physical understanding of the periodical unsteady flow phenomena and to provide more reliable prediction methods of these complex flow conditions in turbomachinery. Selected aspects of flow unsteadiness in turbomachines were investigated with complementing experimental and numerical investigations. Different flow conditions of different complexity were investigated in detail. After a 3-year period of the project, first results of the research group are published in a special issue of the Journal of Flow, Turbulence and Combustion (Flow Turbulence Combust 69, 2002). After the end of the joint project, in the present paper selected results of each research group, which addresses different aspects of periodical unsteady flow in turbomachinery, are discussed. However, it is not the intention of the present paper to give a general survey on this field of research. The following topics are selected to provide insight into the work of the joint research group: (1) Experimental Investigation of Rotor-Stator-Interactions in an Axial Compressor, (2) Influence Of Periodically Unsteady Flow On The Boundary Layer Development Of A Highly Loaded Linear Turbine- And Compressor Cascade, (3) Flow Conditions on a Flat Plate under Oscillating Inlet Conditions, (4) Simultaneous Measurements of Flow and Heat Transfer in a Periodically Unsteady Flow, (5) Turbulence- and Transition Modelling for Unsteady RANS simulations, and (6) Direct Numerical Simulations of Transitional Flow in Turbine-Related Geometries.

Aerodynamic Performance of Low Speed Axial Flow Compressor Rotors with Sweep and Tip Clearance

Blade tip gap, though finite, has great influence on the performance of a turbomachine. Tilting the blade sections to the flow direction (blade sweep) would increase the operating range of an axial compressor by modifying the pressure and velocity fields on the suction surface. The present paper investigates the combined effect of these two factors on various flow characteristics in a low speed axial flow compressor. For the present study, six computational domains were modeled: two rotor sweep configurations (0° and 20°) and three different clearance levels for each rotor. Commercial CFD package ANSYS® CFX 11.0 is used for the simulations. Results indicated that the effect of tip clearance is more predominant in swept rotor than unswept rotor in terms of change in total pressure rise and efficiency. Traced streamlines indicated that blade sweep guides the suction surface boundary layer fluid towards the trailing edge, preventing it from getting accumulated near the tip. Studies near the tip region showed the pattern of velocity streamlines through the tip gap. Tip clearance vortices are found and their trajectory and effect on the flow are studied. Total pressure rise contours at the rotor exit showed that radial migration of low energy boundary layer fluid towards the tip sections is well prevented with sweep.

Time Resolved Experimental Investigations of an Axial Compressor With Casing Treatment

A casing treatment with axial and radial skewed slots ending in a plenum chamber has experimentally been investigated at a highly subsonic axial compressor stage. The aim was to investigate the physical phenomenon of this treatment family that is responsible for the stabilization of the blade passage flow and the drop in efficiency mostly observed. The experimentally gained performance results of this configuration showed an extension of the operating range by approximately 50%, while the efficiency for design conditions is reduced by 1.4%. Apart from this, operating points at part load conditions have been observed nearly without any loss in efficiency. The detailed flow analysis is performed by means of results from a 3D pneumatic probe with temperature sensor and a dynamic total pressure probe. The focus of the investigations is on the incidence flow to the compressor rotor, the tip clearance vortex flow in combination with the wall stall separation region and the blade stall due to suction side separation. The casing treatment configuration is investigated with a special interest in detecting those effects which have an impact on the stability and the compressor overall efficiency, including the interaction of the rotor and the stator flow fields.

Prestall Behavior of a Transonic Axial Compressor Stage via Time-Accurate Numerical Simulation

Computational fluid dynamics calculations using high-performance parallel computing were conducted to simulate the prestall flow of a transonic compressor stage, NASA compressor Stage 35. The simulations were run with a full-annulus grid that models the 3D, viscous, unsteady blade row interaction without the need for an artificial inlet distortion to induce stall. The simulation demonstrates the development of the rotating stall from the growth of instabilities. Pressure rise performance and pressure traces are compared with published experimental data before the study of flow evolution prior to the rotating stall. Spatial fast Fourier transform analysis of the flow indicates a rotating long-length disturbance of one rotor circumference, which is followed by a spike-type breakdown. The analysis also links the long-length wave disturbance with the initiation of the spike inception. The spike instabilities occur when the trajectory of the tip clearance flow becomes perpendicular to the axial direction. When approaching stall, the passage shock changes from a single oblique shock to a dual shock, which distorts the perpendicular trajectory of the tip clearance vortex but shows no evidence of flow separation that may contribute to stall.

Periodical Unsteady Flow Within a Rotor Blade Row of an Axial Compressor—Part II: Wake-Tip Clearance Vortex Interaction

In this two-part paper, results of the periodical unsteady flow field within the third rotor blade row of the four-stage Dresden low-speed research compressor are presented. The main part of the experimental investigations was performed using laser Doppler anemometry. Results of the flow field at several spanwise positions between midspan and rotor blade tip will be discussed. In addition, time-resolving pressure sensors at midspan of the rotor blades provide information about the unsteady profile pressure distribution. In Part II of the paper, the flow field in the rotor blade tip region will be discussed. The experimental results reveal a strong periodical interaction of the incoming stator wakes and the rotor blade tip clearance vortices. Consequently, in the rotor frame of reference, the tip clearance vortices are periodical with the stator blade passing frequency. Due to the wakes, the tip clearance vortices are separated into different segments. Along the mean vortex trajectory, these parts can be characterized by alternating patches of higher and lower velocities and flow turning or subsequent counter-rotating vortex pairs. These flow patterns move downstream along the tip clearance vortex path in time. As a result of the wake influence, the orientation and extension of the tip clearance vortices as well as the flow blockage periodically vary in time.

Stall inception in a transonic axial fan

The current paper reports a numerical investigation of stall inception in a transonic compressor rotor, NASA Rotor-67, by using the whole flow passages in the computations. Surface roughness is added to one of the blades in order to trigger rotating stall. During stall inception, the tip clearance vortex moved away from the suction surface of the roughened blade's upper neighbour, leading to vortex breakdown. The stall cell was found to propagate opposite the direction of the blade rotation at about 30 per cent of rotational speed. The effect of air bleeding on stabilizing the compressor is also studied in this work. The mean mass flowrate removed from the bleed valves was ∼1.2 per cent of the mainstream flowrate. This amount of bleeding was found to effectively suppress the stalling disturbances.

Effects of Tip Clearance on Aerodynamic Damping in a Linear Turbine Cascade

This paper documents an investigation into unsteady flow in a three-dimensional oscillating turbine cascade with emphasis on the influence of tip clearance. Systematic experimental measurements were acquired on a low-speed turbine cascade rig. The cascade consists of seven prismatic turbine blades, with the middle blade being driven to oscillate in a three-dimensional bending/flapping mode. Blades were instrumented with pressure tappings at six spanwise sections to facilitate three-dimensional steady and unsteady pressure measurements on the blade surface. The steady pressure measurements are complemented by computational fluid dynamics simulations. Both are in good agreement and indicate a marked local pressure drop at 70-90 % C on the suction surface resulting from the tip-clearance vortex. The measured unsteady pressure shows that at a small tip clearance (1.25-2.5%C), the tip-clearance flow generates a stabilizing contribution around the midchord on the suction surface near the tip. This behavior is in line with a quasi-steady analysis. Whereas at a large tip clearance (5%C), a dominant destabilizing effect is observed around 80 % C on the suction surface, which is associated with a well developed tip-clearance vortex.

Unsteady tip clearance flow pattern in an isolated axial compressor rotor with micro tip injection

A numerical study of the effect of discrete micro tip injection on unsteady tip clearance flow pattern in an isolated axial compressor rotor is presented, intending to better understand the flow mechanism behind stall control measures that act on tip clearance flow. Under the influence of injection the unsteadiness of self-induced tip clearance flow could be weakened. Also the radial migration of tip clearance vortex is confined to a smaller radial extent near the rotor tip and the trajectory of tip clearance flow is pushed more downstream. So the injection is beneficial to improve compressor stability and increase static pressure rise near rotor tip region. The results of injection with different injected mass flow rates show that for the special type of injector adopted in the paper the effect of injection on tip clearance flow may be different according to the relative strength between these two streams of flow. For a fixed injected mass flow rate, reducing the injector area to increase injection velocity can improve the effect of injection on tip clearance flow and thus the compressor stability. A comparison of calculations between single blade passage and multiple blade passages validates the utility of single passage computations to investigate the tip clearance flow for the case without injection and its interaction with injected flow for the case with tip injection.

Behaviour of tip-leakage flow in an axial flow compressor rotor

The current paper reports on investigations with an aim to advance the understanding of the flowfield near the casing of a small-scale high-speed axial flow compressor rotor. Steady three-dimensional viscous flow calculations are applied to obtain flowfields at various operating conditions. To demonstrate the validity of the computation, the numerical results are first compared with available measured data. Then, the numerically obtained flowfields are analysed to identify the behaviour of tip-leakage flow, and the mechanism of blockage generation arising from flow interactions between the tip clearance flow, the blade/casing wall boundary layers, and non-uniform main flow. The current investigation indicates that the ‘breakdown’ of the tip-leakage vortex occurs inside the rotor passage at the near stall condition. The vortex ‘breakdown’ results in the low-energy fluid accumulating on the casing wall spreads out remarkably, which causes a sudden growth of the casing wall boundary layer having a large blockage effect. A low-velocity region develops along the tip clearance vortex at the near stall condition due to the vortex ‘breakdown’. As the mass flow rate is further decreased, this area builds up rapidly and moves upstream. This area prevents incoming flow from passing through the pressure side of the passage and forces the tip-leakage flow to spill into the adjacent blade passage from the pressure side at the leading edge. It is found that the tip-leakage flow exerts a little influence on the development of the blade suction surface boundary layer even at the near stall condition.