Computational Analysis of the Blade Loading on Tip Clearance Effects

To see the tip clearance effect on a centrifugal compressor, time-resolved flow measurements are carried out at the discharge of a centrifugal compressor impeller with a diameter of 280 mm. The impeller is 30° back swept and is operated at a blade tip speed of 260 m/s, which is a rather high speed for measuring time-resolved flow. Six levels of tip clearance ratio ranging from 0.06 to 0.21 were tested. A fast response aerodynamic probe equipped with a single pressure sensor is used for the flow measurements. The correlation between the impeller efficiency and the tip clearance is presented. The mean flow variation from hub to shroud is obtained and its dependence on the tip clearance is discussed. The measured time-resolved flow at impeller exit differs from the classical jet-wake flow model. Four kinds of flow patterns (blade wake, jet, channel wake, and tip clearance flow) are observed at the impeller discharge plane. By increasing the tip clearance, the total pressure loss in the vaneless diffuser decreases, while the impeller loss increases. One-dimensional flow calculations of the vaneless diffuser show that the matching constant of a skin friction coefficient is decreased by increasing the tip clearance.

According to the theory presented by the authors, the tip clearance loss of an un-shrouded centrifugal impeller mainly consists of two kinds of loss; one is the drag due to the leakage flow through the blade tip clearance and the other is the pressure loss to support the fluid in the thin annular clearance space between the shroud and the blade tip against the pressure gradient in the meridional plane without blades. The former is proportional to the leakage flow or the contraction coefficient of leakage flow. The authors have conducted performance tests using an impeller with 16 backward-leaning blades in three configurations of the blade tip: round edge, sharp square edge, and edge with an end-plate. The experimental tip clearance effects can be predicted by the theory assuming reasonable contraction coefficients. They are 0.91, 0.73, and 0.53 for the respective tip configurations. The impeller efficiency is improved by about 1.5 point by reducing the contraction coefficient from 0.91 to 0.53, providing that the tip clearance ratio at the exit of impeller is 0.1. More improvement is expected for an impeller with highly loaded blades where the leakage loss shares the major part of the tip clearance loss.

Multistage compressor is the most important constituent of gas turbines used in land, naval and aeronautical applications. Overall performance of such machinery depends mainly on the axial compressor performance. Due to the relative motion between rotor and stator blades, the flow field in this machinery is highly unsteady. Furthermore several technological effects like tip clearances, complexity of the blade shapes, variation of axial distance between stator and rotor, seal leakages and cooling holes among others complicate the machine. Therefore the study of a complicated, strongly three-dimensional flow field inside a compressor is considered to be one of the most difficult tasks to be performed by a CFD expert. The present work is the extensive numerical study of the effect of: (1) tip clearance of rotor blades and (2) the axial gap between rotor and stator on the overall performance of a multistage axial compressor. A commercial software package is used for this study. Reynolds-averaged Navier–Stokes equations are solved using Spalart–Allmaras model. A number of steady-state viscous flow simulations were run for both the tip clearance effect and different axial gaps between stator and rotor. All simulations were performed for the first stage, i.e. Stator–Rotor–Stator. Simulations were carried out with coarse, medium and fine meshes to find an optimum, mesh-independent solution. It has been found that larger tip clearance has a detrimental effect on the stage pressure ratio and efficiency of a multistage axial compressor. Similarly there exists a certain distance ratio between the stator1–rotor and rotor–stator2, where stage performance is optimum. Overall performance characteristics obtained through simulation for both the tip clearance and axial gap variation were also compared with the experimental studies and found to be in good agreement.

Tip clearance is a critical parameter affecting fan performance and aeroelastic stability. This paper investigates the effect of tip clearance on fan aeroelasticity under different acoustic propagation characteristics through numerical simulations. Focusing on the first bending mode, which is most prone to flutter, the aeroelastic stability under various nodal diameters (ND) is studied using the energy method, with a detailed analysis of the 2 ND mode, which exhibits flutter risk. The study reveals that tip clearance does not alter the position of the most unstable nodal diameter. Under upstream and downstream cut-off mode and upstream cut-on mode, increasing tip clearance enhances aerodynamic damping, whereas under downstream cut-on mode, it reduces aerodynamic damping. The aerodynamic damping of the blade is jointly influenced by the flow field and acoustic propagation characteristics. The impact of the flow field is primarily concentrated near the leading edge of the pressure side at the blade tip, while the acoustic propagation characteristics significantly alter the phase of unsteady pressure, leading to notable changes in energy exchange characteristics. The study concludes that there is no optimal aeroelastic clearance, as the effect of clearance on aeroelasticity is closely related to acoustic propagation characteristics.

An experimental study of partial admission losses with various blade tip clearances using a linear cascade

The blade tip clearance is a crucial factor in the compressor performance and the aeroacoustics. There are usually main blades and splitter blades, two different types of blades in centrifugal compressors. Keeping almost the same work input, three different tip clearance configurations were designed in which the tip clearance gaps of main blades and splitter blades were uneven. Then experimental investigations were carried out to study the influence of uneven tip clearances on the aeroacoustic characteristics of centrifugal compressors. The results showed that larger main blade tip clearances or larger splitter blade tip clearances would induce rotating instabilities (RI) near flow instability boundary at low rotating speeds. However, the broadband noises below the blade passing frequency (BPF) in the case with uneven blade tip clearances were suppressed significantly which might be due to the incongruous interaction between the tip clearance leakage flows of main blades and splitter blades. In addition, the blade nonsynchronous vibrations (NSV) were observed in the analysis of the frequency spectrum of far-field noise in the case which had smaller main blade tip clearances and larger splitter blade tip clearances. Based on the jet core feedback theory, a physical explanation was given for NSV occurred in this case.

The turbine blade tip is considered as one of the most critical areas of gas turbine engines. The tip region often lacks durability and is challenging to cool. To achieve successful blade tip cooling designs, ALSTOM engineers are performing state of the art aero thermal analyses of blade tip cooling configurations. This paper describes the approach used for this analysis and draws conclusion for blade tip cooling optimization. Numerical simulations of flow and heat transfer are presented for a modern industrial gas turbine blade with a film cooled tip. The blade tip metal temperature distribution is analyzed for three different blade tip clearances with a detailed CFD analysis around the blade tip performed. The CFD analysis provides flow streamlines through the blade tip as well as a total blade tip leakage flow. Rough streamlines estimates are then used to define a set of control volumes for which dedicated cooling flow mixing is considered. The total mass flowing through all volumes corresponds to the CFD blade tip leakage. For each control volume corresponds a specific Reynolds number that is used to define a corresponding heat transfer coefficient. The latter is obtained from experimental Nusselt number correlations for the different regions of a blade squealer tip (crown, fillet and cavity). Application of the obtained heat transfer coefficient and mixing temperature boundary conditions on a 3D blade tip finite element model, together with an internal cooling flow network associated to the 3D model allows calculating the blade tip metal temperature. Results for two different tip clearances relative to nominal blade tip gap are presented and discussed. Comparison with experimental data such as thermal paint test and metallurgical data are given, showing good agreement with the blade tip cooling modeling introduced in this paper. Cooling performance of the blade tip is discussed based on the modeling approach proposed in this paper. The latter allows drawing conclusions for blade tip cooling optimization.

A numerical analysis has been conducted in order to simulate the characteristics of complex flow through linear cascades of high performance turbine blade with/without tip clearance by using a pressure-correction based, generalized 3D incompressible Navier-Stokes CFD code. The development and generation of horseshoe vortex, passage vortex, leakage vortex, tip vortex within tip clearance, etc. are clearly identified through the present simulation which uses the RNG k-e turbulent model with wall function method and a second-order linear upwind scheme for convective terms. The present simulation results are consistent with the generally known tendency that occurs in the blade passage and tip clearance. A 3D model for secondary and leakage flows through turbine cascades with/without tip clearance is also suggested from the present simulation results, including the effects of tip clearance height.

The tip clearance of the compressor has always been a key part in both the design and operation of gas turbine. Surge margin and aerodynamic performance of the compressor will be affected if the tip clearance is too large, while risk of rubbing between the blades and casing will arise if the tip clearance is too small. Generally speaking, the calculation and analysis of tip clearance under stable design condition is relatively easy, because the metal temperature field of both rotor and stator components that significantly affect the tip clearance is in stable condition at that time. However, the transient tip clearance under start-up and shut-down condition is difficult to be calculated and analyzed, due to the fact that both the aerodynamic parameters and metal temperature field are instable, and complicated heat-exchange occurs between the metal components and gas. In this paper, the transient deformation of both rotor and stator components of compressor of a certain gas turbine is analyzed in detail by using thermo-solid coupling analysis technology, finally obtained real-time tip clearance. The non-contact measurement technology is used to measure the tip clearance during operation of the unit. The experimental result indicates that the real-time tip clearance value obtained by the calculation method in the paper is in a very good agreement with the measured value. In addition, both the analysis and experimental results show that the moment of smallest tip clearance usually happens at shut-down progress and the influence of tip clearance must be considered when designing unit’s running curve.

Effects of turbine blade tip shape on total pressure loss and secondary flow of a linear turbine cascade

This study focuses on local heat transfer characteristics on the tip and near-tip regions of a turbine blade with a flat tip, tested under transonic conditions in a stationary, 2-D linear cascade with high freestream turbulence. The experiments were conducted at the Virginia Tech transonic blow-down wind tunnel facility. The effects of tip clearance and exit Mach number on heat transfer distribution were investigated on the tip surface using a transient infrared thermography technique. In addition, thin film gages were used to study similar effects in heat transfer on the near-tip regions at 94% height based on engine blade span of the pressure and suction sides. Surface oil flow visualizations on the blade tip region were carried-out to shed some light on the leakage flow structure. Experiments were performed at three exit Mach numbers of 0.7, 0.85, and 1.05 for two different tip clearances of 0.9% and 1.8% based on turbine blade span. The exit Mach numbers tested correspond to exit Reynolds numbers of 7.6 × 105, 9.0 × 105, and 1.1 × 106 based on blade true chord. The tests were performed with a high freestream turbulence intensity of 12% at the cascade inlet. Results at 0.85 exit Mach showed that an increase in the tip gap clearance from 0.9% to 1.8% translates into a 3% increase in the average heat transfer coefficients on the blade tip surface. At 0.9% tip clearance, an increase in exit Mach number from 0.85 to 1.05 led to a 39% increase in average heat transfer on the tip. High heat transfer was observed on the blade tip surface near the leading edge, and an increase in the tip clearance gap and exit Mach number augmented this near-leading edge tip heat transfer. At 94% of engine blade height on the suction side near the tip, a peak in heat transfer was observed in all test cases at s/C = 0.66, due to the onset of a downstream leakage vortex, originating from the pressure side. An increase in both the tip gap and exit Mach number resulted in an increase, followed by a decrease in the near-tip suction side heat transfer. On the near-tip pressure side, a slight increase in heat transfer was observed with increased tip gap and exit Mach number. In general, the suction side heat transfer is greater than the pressure side heat transfer, as a result of the suction side leakage vortices.

The aerodynamic performance results indicate that the fan adiabatic efficiency was highest with nominal tip clearance (about 91.8% at design speed) and decreased as fan tip clearance increased. Differences in fan efficiency between tip clearance configurations was small below 77.5% fan design speed, about 0.5% maximum, and got larger as fan speed increased, to around 1% maximum at 100% fan design speed. The decreases in efficiency are due to a lower blade loading and higher temperature rise over the outer 20% of the blade span; the mechanism is the tip leakage flow from pressure to suction surface of the fan blade, which increases as the fan tip gap increases. Farfield acoustic results show that changes in the noise level are primarily aft-radiating and are on the order of 1 to 5 dB in fan broadband noise level as measured for the rotor-alone configuration (which had no stators). There were very small changes in noise level with tip clearance for the 54-vane Baseline Outlet Guide Vane configuration. Small changes in noise were also seen for the other two Outlet Guide Vane configurations tested - the 26-vane radial Low Count Outlet Guide Vanes and the 26-vane aft swept Low Noise Outlet Guide Vanes. The rotor alone acoustic results suggest that tip clearance changes induce broadband noise changes at the fan tip, but that the noise differences are masked by the rotor wake/stator interaction noise generated with the Outlet Guide Vanes installed. The magnitude of the rotor/stator interaction tones showed small increases with increasing fan tip gap at fan speeds below transonic, indicating slightly stronger interactions with the larger tip vortex wakes that formed with larger tip gaps. The broadband levels, for the most part, only showed the effect of Outlet Guide Vane geometry on the magnitude of the noise as the tip gap changed. The Laser Doppler Velocimetry flow field results show that the flow downstream of the tip of the blades changes very little with changes in the tip clearance when operating at the approach condition. At both the cut-back and take-off speeds, significant changes in the tip flow occur with changes in the tip clearance. Since these changes in the tip flow are not *

It is believed that the rotating instability phenomenon originating in the compressor tip region is due to leakage flow, which is closely associated with the blade tip clearance. In this work, we have studied the correlation between the dynamic characteristics of blade tip flow and the size of tip clearance for a single-stage low-speed compressor rotor, so as to unveil the mechanism of rotating instability. The full-passage numerical simulations were carried out to obtain the variations in frequency, circumferential mode, and spatial flow field associated with rotating instability. The results of spatial mode decomposition with open clearance show the number of predominate instability modes identified are 25 and 30, respectively. By diminishing the blade tip clearance, all these unstable modes greatly diminished. The formation and propagation of the tip leakage vortex were described in detail to show the development of rotating instability. Two flow field reduced-order methods, proper orthogonal decomposition and dynamic mode decomposition, were used to analyze the flow field, energy proportion, and stability of related modes under different tip clearances. The results show that the first several modes with strong stability account for a large proportion of energy and make a major contribution to flow unsteadiness. The energy proportion and stability of rotating instability decrease as the tip clearance becomes smaller. The blade-passing frequency and its multiples emerge as the main components of the flow field.

A three-dimensional viscous flow code was used to study the tip clearance flow in linear turbine cascades. Two test cases with detailed measurements were used for validation. The flow properties inside the tip gap, such as static pressure, velocity, and angle, were compared with the data. The blade loading at different span locations and the downstream flowfield properties, such as loss and angle, were also compared with the data. The clearance flow phenomena, such as leakage vortex, changing of loss distribution, and flow underturning, were predicted by the numerical code. This code not only can be used as an analysis tool to study tip clearance effects in real turbine blade rows but also can be used as a design tool to design the turbine components. HE effort expended to develop a modern gas turbine engine is substantial. A significant portion of this effort is devoted to the preliminary design of aerodynamic components and subsequent modification by iteration to achieve the final design that meets the required engine performance goals. With the rapid growth in computer technology, computational fluid dynamics (CFD) codes are used as an analysis tool to help engineers understand the flow features in turbine components. Use of CFD codes that can accurately predict operation char- acteristics of turbomachiner y components will result in con- siderable reduction of development effort and provide in- creased performance. The validation of a CFD code is necessary to use it as a design tool. To evaluate the performance of the CFD code, reliable experimental data are needed. The data should provide detailed information for flow features in turbine components. Unshrouded turbine rotors operate with a tip clearance between the blade tip and the casing. The fluid that leaks between the blade tip and casing causes large efficiency losses. The tip leakage flow also changes the loss distribution, exit flow angle, and loading. Therefore, it is important for the turbomachinery designer to understand and predict the tip clearance flow phenomena. Hah,1 Pouagare and Delaney,2 and Dawes3 predicted tip clearance flow using a three-dimensional Navier-Stokes code; however, all compressor blade rows are calculated with the assumption of zero thickness on the blade tip. They used one grid line and periodic conditions for the flow inside the tip clearance region. This model is satisfactory for very thin air- foils, such as those found at the tip of a transonic compressor rotor. However, thick airfoils, found in turbine blades, re- quire the full flowfield resolution in the tip gap region. Moore and Moore4 predicted the tip clearance flow in a linear turbine cascade with the three-dimensional incompressible Navier- Stokes code. However, the flow in turbomachiner y compo- nents is usually compressible. In this study, a numerical code that solves three-dimensional, compressible Navier-Stokes

The geometric scaling method is often used in the design and development of axial compressors to reduce high costs. However, Reynolds number, surface roughness, and tip clearance are often difficult to satisfy all the similarity criteria, which breaks the similarity between the Prototype and the Model compressor. This study utilized a 1.5-stage axial compressor as the Prototype to investigate the effects of Reynolds number, surface roughness, and tip clearance on the geometric scaling process. First, the 0.5 scale Model compressor was simulated using RANS method with the scalable wall function. A test facility was constructed for the 0.5 scale Model compressor, and experiments of varying surface roughness and tip clearance were carried out to verify the reliability of the numerical method. Then, the Prototype and Models with scaling factors of 0.4 and 0.33 were simulated using the same numerical method. By analyzing the numerical results of the Prototype and the three Models, a novel correction method for the deviation between the performance of the Prototype and the Model was proposed. This method can be used to correct the deviation of compressors' performance curves caused by the change in Reynolds number, surface roughness, and tip clearance during the geometric scaling process. Meanwhile, both numerical and experimental results were used to validate the accuracy and the universal applicability of the method.