Turbine Blade Tip Cooling Modeling and Optimization

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.

Purpose The purpose of this paper is to numerically investigate the effect of guide vane unsteady passing wake on the rotor blade tip aerothermal performance with different tip clearances. Design/methodology/approach The geometry and flow conditions of the first stage of GE-E3 high-pressure turbine have been used to obtain the blade tip three-dimensional heat transfer characteristics. The first stage of GE-E3 high-pressure turbine has 46 guide vanes and 76 rotor blades, and the ratio of the vane to the blade is simplified to 38:76 to compromise the computational resources and accuracy. Namely, each computational domain comprises of one guide vane passage and two rotor blade passages. The investigations are conducted at three different tip gaps of 1.0, 1.5 and 2.0 per cent of the average blade span. Findings The results show that the overall discrepancy of the heat transfer coefficient between steady results and unsteady time-averaged results is quite small, but the dramatic growth of the instantaneous heat transfer coefficient along the pressure side is in excess of 20 per cent. The change of the aerothermal performance is mainly driven by turbulence-level fluctuations of the unsteady flow field within gap regions. In addition, the gap size expansion has a marginal impact on the variation ratio of tip unsteady aerothermal performances, even though it has a huge influence on the leakage flow state within the tip region. Originality/value This paper emphasizes the change ratio of unsteady instantaneous heat transfer characteristics and detailed the mechanism of blade tip unsteady heat transfer coefficient fluctuations, which provide some guidance for the future blade tip design and optimization.

Gas turbines normally have the rotating and stationary parts. The rotational blades always need clearance to rotation. Blade tip clearance is one of the important parameters affecting performance, safety and stability of a gas turbine engine. However, it is difficult to measure the tip clearance accurately during the tests. The numerical studies are important methods to study the tip clearance effects. Tip clearance effects have been studied for a long time. However, there are still many aspects that are not fully understood and industries still don’t have a good way to perform design with minimum tip clearance losses. In the gas turbine design, the tip clearance considerations are very critical for both performance and reliability. In this paper, we perform the numerical study on the tip clearance impacts during the aerodynamic design, especially the blade loading through blade counts. The study found that the proper blade loading could reduce the tip clearance sensitivity and reduce the tip clearance losses. In the cases of reasonable number of blades, the effect of clearance on the total pressure loss and efficiency had a linear relationship. But in the cases of small number of blades, the effect of clearance on the total pressure loss and efficiency is no longer linear. The study helped turbine designers to consider blade loading and tip clearance together during the turbine design.

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

The tip clearance between turbine blade and stationary casing leads to the exposure of blade tip to high temperature gas and contributes to a high thermal load in this region. Then it is necessary to perform cooling on the blade tip. Moreover, the tip leakage flow in a high speed brings complex aerothermal conditions to the blade tip. The blade tip film cooling on the transonic squealer tip is investigated in this research. The effects of tip and pressure-side coolant injections on the blade tip heat transfer are discussed. And the influences of total pressure ratio of coolant to main flow and endwall moving on the film cooling are considered. Firstly, the reliability of a commercial CFD code with different turbulence models was first validated by contrasting with the experimental results. The heat flux and isentropic Ma number predicted with the BSL k-ω model shows a better agreement with the test data. Then the mesh independency study was performed. The mesh with tip and pressure-side injection hole is generated by in-house code. For the tip without film cooling at outlet Mach number of 0.96, the squealer tip shows a high heat transfer at front of the tip cavity floor caused by the impingement of tip leakage flow, besides the suction-side rim has high heat transfer coefficient. With the tip coolant injection, the tip coolant is pushed towards the cavity pressure-side and provides better coverage in this region. At the front of cavity floor near suction-side, there is almost no coolant coverage then it shows little cooling effects. The coolant discharges from the tip over the suction-side rim thereby the heat transfer near the trailing edge is less affected by the tip coolant injection. With the pressure-side near tip film cooling, the coolant injection shows high film-cooling effectiveness at low coolant stagnation pressure and lift off blade surface at high coolant stagnation pressure. The endwall motion increases the total heat load of the blade tip by increase heat load at cavity floor and suction-side near tip region and the reduces film-cooling effectiveness.

As turbomachinery OEMs focus efforts to further increase reliability, power and efficiencies, the running clearance between blade tips and stator continue to be of the utmost importance. This paper investigates the capability of capacitive tip clearance systems to perform individual blade tip clearance measurements on high speed rotors of up to 90,000rpm. A rotor was designed using finite element analysis; unique blade responses have been predicted. The objective of this investigation was to consider two different approaches to the application of blade tip clearance measurements and the system requirements to accurately measure low levels of radial displacement of a target rotating between 1,000rpm and 90,000rpm. The first uses the standard approach with passive probes and the second, a new technique using active probes that have demonstrated bandwidths of 1.2MHz and increased measuring range with a lower level of measurement uncertainty. Both systems’ approaches are compared, and their capabilities are evaluated for high-speed applications. The higher bandwidth capabilities of the latter system, combined with smaller sensor diameters, produces comparable signal rise times to the optical systems used in blade tip timing measurements. The difference in approach offers the potential of contamination resistant sensors for long term blade tip timing applications and measurement probes that do not require cooling systems to withstand higher temperature applications. The use of different probe configurations, in a number of applications, has demonstrated a two-fold improvement in the measurement range whilst producing lower levels of noise and uncertainty when applied to blade targets made from composites, aluminium and nickel-alloy materials. The measurement data presented includes individual blade’s radial displacement, identified shaft axial displacement, effects of resonance in the test system and the identification of the main drivers of measurement uncertainty along with an achievable value. The capacitive measurement systems’ performance for blade tip clearance is analysed and reported. The capability to perform other measurement techniques such as blade tip timing with a dual use measurement probe is also analysed and reported. This is done by correlating measurement results between the capacitive systems with that of a repeat measurement of the same target using an optical BTT system.

An experimental study of film cooling is conducted on the tip of a turbine blade with a blade rotation speed of 1200 rpm. The coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes, and six shaped PS holes. The stator–rotor–stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0, and 4.0. The heat transfer result for the no coolant injection shows a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The cooling effectiveness signatures indicate that the tip coolant is pushed toward the blade suction side thereby providing better coverage in that region. This shift in coolant flow toward the blade suction side, as opposed to the PS in stationary studies, can primarily be attributed to the effects of the blade relative motion.

Influences of a multi-cavity tip on the blade tip and the over tip casing aerothermal performance in a high pressure turbine cascade

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.

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.

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.

Steady simulations were performed to investigate tip leakage flow and heat transfer characteristics on the rotor blade tip and casing in a single-stage gas turbine engine. A typical high-pressure gas turbine stage was modeled with a pressure ratio of 3.2. The predicted isentropic Mach number and adiabatic wall temperature on the casing showed good agreement with available experimental data under similar operating condition. The present numerical study focuses extensively on the effects of tip clearance heights and rotor rotational speeds on the blade tip and casing heat transfer characteristics. It was observed that the tip leakage flow structure is highly dependent on the height of the tip gap and the speed of the rotor. In all cases, the tip leakage flow was seen to separate and recirculate just around the corner of the pressure side of the blade tip. This region of re-circulating flow enlarges with increasing clearance heights. The separated leakage flow reattaches afterwards on the tip surface. Leakage flow reattachment was shown to enhance surface heat transfer at the tip. The interaction between tip leakage flow and secondary flows that is induced by the relative casing motion is found to significantly influence the blade tip and casing heat transfer distribution. A region of critical heat transfer exists on the casing near the blade tip leading edge and along the pressure-side edge for all the clearance heights that were investigated. At high rotation speed, the region of critical heat transfer tends to move towards the trailing edge due to the change in inflow angle.

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

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.

The present study considers spatially resolved surface heat transfer coefficients and adiabatic wall temperatures on a turbine blade tip in a linear cascade under transonic conditions. Five different measurement and processing techniques using infrared thermography are considered and compared. Three transient methods use the same experimental setup, using a heater mesh to provide a near-instantaneous step-change in mainstream temperature, employing an infrared camera to measure surface temperature. These three methods use the same data but different processing techniques to determine the heat transfer coefficients and adiabatic wall temperatures. Two of these methods use different processing techniques to reconstruct heat flux from the temperature time trace measured. A plot of the heat flux versus temperature is used to determine the heat transfer coefficients and adiabatic wall temperatures. The third uses the classical solution to the 1D nonsteady Fourier equation to determine heat transfer coefficients and adiabatic wall temperatures. The fourth method uses regression analysis to calculate detailed heat transfer coefficients for a quasi-steady-state condition using a thin-foil heater on the tip surface. Finally, the fifth method uses the infrared camera to measure the adiabatic wall temperature surface distribution of a blade tip after a quasi-steady-state condition is present. Overall, the present study shows that the infrared thermography technique with heat flux reconstruction using the impulse method is the most accurate, computationally efficient, and reliable method to obtain detailed, spatially resolved heat transfer coefficients and adiabatic wall temperatures on a transonic turbine blade tip in a linear cascade.