Turbine Split Rings Thermal Design Using Conjugate Numerical Simulation

One of the key factors ensuring gas turbine engines (GTE) competitiveness is improvement of life, reliability and fuel efficiency. However fuel efficiency improvement and the required increase of turbine inlet gas temperature (T*g) can result in gas turbine engine life reduction because of hot path components structural properties deterioration. Considering circumferential nonuniformity, local gas temperature T*g can reach 2500 K. Under these conditions the largest attention at designing is paid to reliable cooling of turbine vanes and blades. At present in design practice and scientific publications comparatively little attention is paid to detailed study of turbine split rings thermal condition. At the same time the experience of modern GTE operation shows high possibility of defects occurrence in turbine 1st stage split ring. This work objective is to perform conjugate numerical simulation (gas dynamics + heat transfer) of thermal condition for the turbine 1st stage split ring in a modern GTE. This research main task is to determine the split ring thermal condition by defining the conjugate gas dynamics and heat transfer result in ANSYS CFX 13.0 package. The research subject is the turbine 1st stage split ring. The split ring was simulated together with the cavity of cooling air supply from vanes through the case. Besides turbine 1st stage vanes and blades have been simulated. Patterns of total temperature (T*Max = 2000 °C) and pressure and turbulence level at vanes inlet (19.2 %) have been defined based on results of calculating the 1st stage vanes together with the combustor. The obtained results of numerical simulation are well coherent with various experimental studies (measurements of static pressure and temperature in supply cavity, metallography). Based on the obtained performance of the split ring cooling system and its thermal condition, the split ring design has been considerably modified (one supply cavity has been split into separate cavities, the number and arrangement of perforation holes have been changed etc.). All these made it possible to reduce considerably (by 40…50 °C) the split ring temperature comparing with the initial design. The design practice has been added with the methods which make it possible to define thermal condition of GTE turbine components by conjugating gas dynamics and heat transfer problems and this fact will allow to improve the designing level substantially and to consider the influence of different factors on aerodynamics and thermal state of turbine components in an integrated programming and computing suite.

Direct Numerical Simulations of a Transonic Tip Flow With Free-Stream Disturbances

In this paper we use direct numerical simulation to investigate the unsteady flow over a model turbine blade-tip at engine scale Reynolds and Mach numbers. The DNS is performed with a new in-house multi-block structured compressible Navier-Stokes solver purposely developed for exploiting high-performance computing systems. The particular case of a transonic tip flow is studied since previous work has suggested compressibility has an important influence on the turbulent nature of the separation bubble at the inlet to the gap and subsequent flow reattachment. The effects of free-stream turbulence, cross-flow and pressure-side boundary-layer on the tip flow aerodynamics and heat transfer are investigated. For ‘clean’ in-flow cases we find that even at engine scale Reynolds numbers the tip flow is intermittent in nature (neither laminar nor fully turbulent). The breakdown to turbulence occurs through the development of spanwise modes with wavelengths around 25% of the gap height. Cross-flows of 25% of the streamwise gap exit velocity are found to increase the stability of the tip flow, and to significantly reduce the turbulence production in the separation bubble. This is predicted through in-house linear stability analysis, and confirmed by the DNS. For the case when the inlet flow has free-stream turbulence, viscous dissipation and the rapid acceleration of the flow at the inlet to the tip-gap causes significant distortion of the vorticity field and reductions of turbulence intensity as the flow enters the tip gap. This means that only very high turbulence levels at the inlet to the computational domain significantly affect the tip heat transfer. The DNS results are compared with RANS predictions using the Spalart-Allmaras and k–ω SST turbulence models. The RANS and DNS predictions give similar qualitative features for the tip flow, but the size and shape of the inlet separation bubble and shock positions differ noticeably. The RANS predictions are particularly insensitive to free-stream turbulence.

Numerical Study of Heat Transfer and Cooling Effectiveness on a Flat-Tip Turbine Blade

A numerical study is performed to simulate the tip leakage flow and heat transfer on the first stage rotor blade tip of GE-E3 turbine, which represents a modern gas turbine blade geometry. Calculations consist of the flat blade tip without and with film cooling. For the flat tip without film cooling case, in order to investigate the effect of tip gap clearance on the leakage flow and heat transfer on the blade tip, three different tip gap clearances of 1.0%, 1.5% and 2.5% of the blade span are considered. And to assess the performance of the turbulence models in correctly predicting the blade tip heat transfer, the simulations have been performed by using four different models (the standard k-ε, the RNG k-ε, the standard k-ω and the SST models), and the comparison shows that the standard k-ω model provides the best results. All the calculations of the flat tip without film cooling have been compared and validated with the experimental data of Azad[1] and the predictions of Yang[2]. For the flat tip with film cooling case, three different blowing ratio (M = 0.5, 1.0, and 1.5) have been studied to the influence on the leakage flow in tip gap and the cooling effectiveness on the blade tip. Tip film cooling can largely reduce the overall heat transfer on the tip. And the blowing ratio M = 1.0, the cooling effect for the blade tip is the best.

Optimizing the Tip Section Profiles of a Steam Turbine Blading

Though there is not much freedom in shape modification of profiles for tip sections of steam turbines of large output, there is still enough space for optimizing these profiles. The constraints are already in the thermodynamic design of the last stage blading. The layout of the tip cascade indicates that the optimizing strategy should primarily concern the leading edge region and the trailing edge. Proper design can cut down both the front and exit shock wave intensities and the shock wave direct and indirect losses, as well as attenuate the interacting shock waves. This “partial optimization” method has been experimentally verified and is presented in this paper.

Aerothermal Performance of Shroudless Turbine Blade Tips With Effects of Relative Casing Motion

Recent work has indicated qualitatively different heat transfer characteristics between a transonic blade tip and a subsonic one. High resolution experimental data can be acquired for blade tip heat transfer research using a high speed linear cascade. While recognising an important role played by the cascade tests in validating computational models at the same conditions, some questions arise in relation to the effects of relative casing motion: 1) Does the relative casing movement change the main flow physics influencing the blade tip aerothermal performance? 2) Can a cascade set up with stationary casing wall rank different designs? 3) How do the effects of the casing motion depend on tip design configurations? A combined experimental and CFD study on several high pressure blade tip configurations is conducted to address these issues. Firstly, extensive experimental tests with aerodynamic loss and heat transfer measurement in a high speed linear cascade have been carried out for a squealer tip configuration at engine representative aerodynamic conditions. A systematic validation of the CFD solver (Rolls-Royce HYDRA) is presented, which serves as a basis for the computational analyses of the effects of the relative casing motion. Two tip configurations (squealer and flat tip) at three tip gaps (0.5%, 1.0%, 1.5% span) are analysed. The main aerodynamic impact of the casing motion is seen to promote the passage vortex, which consequently supresses the pitchwise reach of the tip leakage vortex. Inside the tip gap, the behaviour is dominated by the extra wall friction in relation of the inertia of the bulk fluid through the gap. As such, the moving casing effect is particularly strong for the flat tip at a small tip gap. For the large and medium tip gaps, both stationary and moving casing results are shown to consistently capture the trends in overall aerothermal performances. The present results confirm that even with relative casing motion, there is still a significant portion of transonic flow over a blade tip. For both the stationary and moving casing cases, the gap dependence of the over-tip heat transfer shows opposite trends for the transonic and subsonic regions respectively. The gap dependence of the blade tip heat transfer is shown to be clearly dependent on tip geometry configurations, as the bulk flow in a squealer cavity is subsonic regardless of the tip gap size, whilst the local flow state over a flat tip is much more responsive to the change of gap size.

Vibration Analysis of Shrouded Turbine Blades for a 30 MW Gas Turbine

Turbine blades are subjected to high static and dynamic loads. In order to reduce the vibration amplitude means of friction damping devices have been developed, e.g. damping wires, interblade friction dampers and shrouds. This paper presents both numerical and experimental results for investigating the dynamical behavior of shrouded turbine blades. The studies are focused on the lowest family of the bladed disk. The aspect of experimental studies, the effect of the shroud contact force on the resonance frequency of the blade was examined by using the simplified blade test stand. Based on the result of the simplified blade studies, the shroud contact force of the real blade was determined in order to stabilize the resonance frequencies of the bladed disk system. The resonance frequencies and mode shapes of the real bladed disk assembly were measured in no rotation and room temperature condition. Finally, the dynamic strains were measured in the actual engine operations by using a telemetry system. The aspect of analytical studies, a non-linear vibration analysis code named DATES was applied to predict vibration behavior of a shrouded blade model which includes contact friction surfaces. The DATES code is a forced response analysis code that employs a 3-dimensional friction contact model. The Harmonic Balance Method (HBM) is applied to solve resulting nonlinear equations of motion in frequency domain. The simulated results show a good agreement with the experimental results.

Design Modification of a Passive Tip-Leakage Control Method for Axial Turbines: Linear Cascade Wind Tunnel Results

Tip-leakage losses can contribute up to one third of the overall losses in unshrouded axial turbine blades. A passive tip-leakage flow control method is used to reduce the tip-leakage loss. Taking into account a modified discharge coefficient model, an inclination of the injection against the tip-leakage flow direction is said to have an even better effect on reducing the tip-leakage loss. To prove the effect, linear cascade measurements have been carried out at three different gap widths from 0.85% to 2.50% chord length. The used geometry is an up-scaled turbine blade tip cross section with weak turning. A single blade is fitted with an injection channel which is inclined by 45° against the tip-leakage flow direction. The flow field of the modified blade was measured 0.31 axial chord length downstream of the cascade using a pneumatic five-hole probe. The tip-leakage loss is reduced by passive tip-injection and further by inclined injection. The reduction can be significant at small gap widths. Detailed results are presented for a gap width of 1.40% chord length.

Experimental and Unsteady Numerical Investigation of the Tip Clearance Noise of an Axial Fan

The aerodynamic and aeroacoustic performance of axial fans are strongly affected by the unavoidable tip clearance. Two identical fan impellers but with different tip clearance ratio were investigated. Unsteady wall pressure fluctuations in the tip region of the rotating blades and on the interior wall of the duct type shroud and the overall sound radiated were analysed by an unsteady numerical Scale-Adaptive Simulation (SAS) and unsteady surface pressure measurements in both, the stationary and rotating system. Based on SAS-predicted pressure fluctuations on the blade surfaces the acoustic analogy according to Ffowcs Williams and Hawkings (FWH) was employed to calculate the sound pressure in the far field. In general, experimentally and numerically determined unsteady flow were found to be a tendentially good agreement. The spatial and temporal structure of the tip vortex system and the resulting unsteady pressure distribution on the surfaces in the vicinity of the blade tips was revealed in good detail. The vortices’ strength and trajectories as well as the unsteadiness are controlled by the size of the tip clearance and the operating point: As tip clearance is increased blade/vortex interaction becomes more prevalent and with it the unsteady surface pressure and eventually the sound radiated into the far field. The broadband tip clearance noise was acceptably predicted from the simulation results, while the prediction at discrete frequency should still be improved in the further work.

Blade Tip Carving Effects on the Aero-Thermal Performance of a Transonic Turbine

Tip leakage flows in unshrouded high speed turbines cause large aerodynamic penalties, induce significant thermal loads and give rise to intense thermal stresses onto the blade tip and casing endwalls. In the pursuit of superior engine reliability and efficiency, the turbine blade tip design is of paramount importance and still poses an exceptional challenge to turbine designers. The ever-increasing rotational speeds and pressure loadings tend to accelerate the tip flow velocities beyond the transonic regime. Overtip supersonic flows are characterized by complex flow patterns, which determine the heat transfer signature. Hence, the physics of the overtip flow structures and the influence of the geometrical parameters on the overtip flow require further understanding to develop innovative tip designs. Conventional blade tip shapes are not adequate for such high speed flows and hence, potential for enhanced performances lays in appropriate tip shaping. The present research aims to quantify the prospective gain offered by a fully contoured blade tip shape against conventional geometries such as a flat and squealer tip. A detailed numerical study was conducted on a modern transonic turbine rotor blade (Reynolds number is 5.5 × 105, relative exit Mach number is 0.9) by means of three-dimensional Reynolds-Averaged Navier-Stokes calculations. The novel contoured tip geometry was designed based on a 2D tip shape optimization in which only the upper 2% of the blade span was modified. This study yields a deeper insight into the application of blade tip carving in high speed turbines and provides guidelines for future tip designs with enhanced aerothermal performances.

Effect of Common Blade Tip Squealer Designs in Terms of Tip Clearance Loss Control

The increase of new gas turbine’s efficiency is connected with further rise of turbine inlet temperature and sometimes as well pressure. In these conditions, first cooled turbine stages of a gas turbine engine usually consist of freestanding airfoils, which do not use an integrated shroud, to avoid risk of shroud overheating. In order to better control the radial gap leakage flow between the rotating blade tip and turbine casing, special design features of the airfoil tip need to be considered in the design process to meet the best possible stage performance. In the general engineering practice, a blade tip squealer provides opportunities to control tip clearance loss. In this paper several simplified types of the tip squealer design are investigated to determine the most effective loss control. At this stage of the investigation, blade tip cooling was not taken into account, but aerodynamic effects were analysed in detail. Based on the most common designs of the blade tip in the literature, four geometry types were investigated: (i) a flat tip design as the reference baseline solution, (ii) full tip squealer, (iii) partial squealer along the pressure side (PS) wall with a cut-out at the pressure side near the trailing edge (TE) and (iv) partial squealer along the suction side (SS) wall with a cut-out at the suction side near TE. All these cases have been compared among each other for two relative radial gaps (gap to blade height) of 0.6% and 1.36%. The flow calculations were done with a full 3D Navier-Stokes CFD code. For the flat tip and for full squealer designs, numerical results were validated against well-known experimental data measured on the GE-E3 blade cascade test rig found in the open literature. By using the 3D numerical data, the special attention was considered to confirm reliability and predictive credibility of the blade tip flow obtained from the analytical model. The obtained loss values and flow details were compared for all studied cases, providing insight into turbine stage aerodynamics with respect to minimal and maximal radial clearance.