Effect of Blade Row Interaction on Rotor Film Cooling

The mechanisms of blade row interaction affecting rotor film cooling are identified in order to make recommendations for the design of film cooling in the real, unsteady turbine environment. Present design practice makes the simplifying assumption of steady boundary conditions, despite intrinsic unsteadiness due to blade row interaction; we argue that if film cooling responds non-linearly to unsteadiness, the time-averaged performance will then be in error. Non-linear behaviour is confirmed using experimental measurements of flat-plate cylindrical film cooling holes, main-stream unsteadiness causing a reduction in film effectiveness of up to 31% at constant time-averaged boundary condition. Unsteady computations are used to identify the blade row interaction mechanisms in a high-pressure turbine rotor: a ‘negative jet’ associated with the upstream vane wake, and frozen and propagating vane potential field interactions. A quasi-steady model is used to predict unsteady excursions in momentum flux ratio of rotor cooling holes, with fluctuations of at least ±30% observed for all hole locations. Computations with modified upstream vanes are used to vary the relative strength of wake and potential field interactions. In general, both mechanisms contribute to rotor film cooling unsteadiness. It is recommended that the designer should choose a cooling configuration which behaves linearly over the expected unsteady excursions in momentum flux ratio as predicted by a quasi-steady hole model.

Conjugate Heat Transfer Analysis to Assess Overall Cooling Effectiveness of High Pressure Turbine Nozzle With Optimized Film Cooling Hole Arrangements

Conjugate heat transfer analysis method has been highlighted for predicting heat exchange between fluid domain and solid domain inside high-pressure turbines, which are exposed to very harsh operating conditions. Then it is able to assess the overall cooling effectiveness considering both internal cooling and external film cooling at the cooled turbine design step. In this study, high-pressure turbine nozzles, which have three different film cooling holes arrangements, were numerically simulated with conjugate heat transfer analysis method for predicting overall cooling effectiveness. The film cooling holes distributed over the nozzle pressure surface were optimized by minimizing the peak temperature, temperature deviation. Additional internal cooling components such as pedestals and rectangular rib turbulators were modeled inside the cooling passages for more efficient heat transfer. The real engine conditions were given for boundary conditions to fluid and solid domains for conjugate heat transfer analysis. Hot combustion gas properties such as specific heat at constant pressure and other transport properties were given as functions of temperature. Also, the conductivity of Inconel 718 was also given as a function of temperature to solve the heat equation in the nozzle solid domain. Conjugate heat transfer analysis results showed that optimized designs showed better cooling performance, especially on the pressure surface due to proper staggering and spacing hole-rows compared to the baseline design. The overall cooling performances were offset from the adiabatic film cooling effectiveness. Locally concentrated heat transfer and corresponding high cooling effectiveness region appeared where internal cooling effects were overlapped in the optimized designs. Also, conjugate heat transfer analysis results for the optimized designs showed more uniform contours of the overall cooling effectiveness compared to the baseline design. By varying the coolant mass flow rate, it was observed that pressure surface was more sensitive to the coolant mass flow rate than nozzle leading edge stagnation region and suction surface. The CHT results showed that optimized designs to improve the adiabatic film cooling effectiveness also have better overall cooling effectiveness.

A Numerical Study on the Effect of Hole Inclination Angle With Imperfection on Film Cooling Effectiveness

The present study numerically evaluates the influence of hole inclination angle with a hole imperfection on film cooling performance. Here, the hole imperfection due to laser percussion drilling is modelled as a half torus. Three hole inclination angles were investigated: 35°, 45° and 55°. Furthermore, every case was evaluated at three blowing ratios: 0.45, 0.90 and 1.25. Each case is compared to a baseline case where the hole imperfection is absent. The results indicate that the hole inclination angle has a strong influence on the film effectiveness performance when a hole imperfection is present. Centerline effectiveness plots reveal a maximum effectiveness deterioration of 89% for a blowing ratio of 0.90 in the vicinity of the hole exit. Dimensionless temperature contours show that the jet produced in the presence of an imperfection is much more compact causing the counter rotating vortex pair to be closer to each other. This enhances the jet to lift off from the plate.

Aerothermal Simulations Comparisons of a Shaped-Hole Film Cooling Flow

Film Cooling is a crucial technology for engine manufacturer to develop high-efficiency gas turbine engines by raising turbine entry temperature. A lot of cooling holes geometries have been studied in the past few years in tests, as well as numerical simulations. Shaped holes are nowadays a standard geometry for protecting the blades, given the performance improvement compared to cylindrical holes. Numerical correlation with physical tests is challenging due to the high sensitivity to thermal mixing and adequate boundary condition predictions. This paper is devoted to numerical simulation comparisons of the 777 shaped holes configuration of Pennsylvania State University, for an incompressible flow with a density ratio of 1.5, a blowing ratio of 1.5 and a free stream turbulence intensity of 0.5%. Two different simulations have been chosen: a state-of-the-art RANS simulation with k-e Realizable model computed with ANSYS Fluent and a high fidelity solver Lattice-Boltzmann Method computed with Simulia PowerFLOW. In order to improve the accuracy of numerical simulations against test results, this article deals with an aerothermal model of the complete test bench. This additional modeling allows to strongly improve thermal prediction and to understand initial discrepancies related to test bench environment. Results show that k-ε Realizable simulation provides a good prediction of average effectiveness, but local differences appear due to inherent RANS modeling limitations. On the other hand, LBM simulation provides excellent results for both aerodynamic and thermal quantities: tests results are very well reproduced.

Characterisation of Waves and Ligaments in Films Close to an Aeroengine Ball Bearing

Surface waves are observed in many situations including natural and engineering applications. Experiments conducted at the Gas Turbine and Transmissions Research Centre (G2TRC) used high speed imaging to observe multiscale wave structures close to an aeroengine ball bearing in a test rig. The dynamic behavior and scale of the waves indicate that these are shear-driven although highly influenced by gravity at low shaft speed. To understand the interactions between gas and liquid phases including momentum and mass transfers, characterization of the observed waves and ligaments was undertaken. Waves were studied at surfaces close to the ball bearing and ligaments were assessed near the cage. Characterization was in terms of frequency and wavelength as functions of speed, flow-rate, bearing axial load and gravity. The assessments confirmed the existence of gravity-capillary waves and capillary waves. Gravity-capillary waves were measured to have a longer mean wavelength on the co-current side of the bearing (gravity and shear acting together) compared to the counter-current side (gravity and shear opposing). Using a published definition of critical wavelength (λcrit), measured wavelengths at 3,000 rpm were 2.56λcrit on the co-current side compared to 1.86λcrit at the countercurrent location. As shaft speed increases, wavelength reduces with transition to capillary waves occurring at around 0.83λcrit. At shaft speeds beyond 5000 rpm, capillary waves were fully visible and the wavelength was obtained as 0.435λcrit. Flow-rate and load did not significantly influence wavelength. Wave frequency was found to be proportional to shaft speed. The gravity-capillary waves had frequencies within the range 13–25 Hz while capillary waves exhibited a frequency well beyond 100 Hz. The frequencies are highly fluctuating with no effect of load and flow rate observed. Ligaments were characterized using Weber number and Stability number. The number of ligaments increased with shaft speed. A correlation for ligament number based on operating conditions is proposed.

Multi-Physics Simulation Based Approach for Life Prediction of a Gas Turbine Combustor Liner

Better life assessment of hot-components of an aero-engine can help improve its reliability and service life, while, reducing associated maintenance cost. Accurate prediction of Thermo-Mechanical Fatigue (TMF) is one of the crucial aspects of life prediction. Therefore, fully resolved simulation methodologies have gained attention as an ingredient for solving TMF problems owing to their potential for providing comprehensive insights into a system having hot components undergoing transient loading during operation. The present work focuses on a multi-physics simulation-based approach for the life-prediction of a representative gas-turbine combustor liner with an objective of providing a complete framework for TMF analysis of an actual aero-engine combustor liner. The presented methodology consists of a coupling between Computational Fluid Dynamics (CFD) and Finite Element Method (FEM). Thermal loads on the representative aero-engine combustor are predicted using Conjugate Heat Transfer (CHT) modeling in the CFD analyses for different operating conditions suitable for a flight cycle. A load cycle is then constructed using these thermal loads and is transferred to the structural analysis to evaluate the stresses in the liner. Results are obtained regarding spatially varying thermal expansion resulting in inelastic strains as governed by temperature and rate dependent material behavior. Stress and plastic strain history information from the structural analysis are processed to predict the life of different regions of the combustor liner. Different simulation methods for conjugate heat-transfer, load-cycle, material property extraction, thermal-stresses, and fatigue are evaluated, and an overall methodology involving accuracy and reasonable computational cost is proposed. The proposed methodology is numerically verified, and the verification results are presented in this work.

Numerical Study of Heat Transfer in Novel Wavy Trailing Edge Design for Gas Turbine Airfoils

The strive to achieve increasingly higher efficiencies in gas turbine power generation has led to a continued rise in the turbine inlet temperature. As a result, novel cooling approaches for gas turbine blades are necessary to maintain them within the material’s thermal mechanical performance envelope. Various internal and external cooling technologies are used in different parts of the blade airfoil to provide the desired levels of cooling. Among the different regions of the blade profile, the trailing edge (TE) presents additional cooling challenges due to the thin cross section and high thermal loads. In this study, a new wavy geometry for the TE has been proposed and analyzed using steady state numerical simulations. The wavy TE structure resembled a sinusoidal wave running along the span of the blade. The troughs on both pressure side and suction side contained the coolant exit slots. As a result, consecutive coolant exit slots provided an alternating discharge between the suction side and the pressure side of the blade. Steady state conjugate heat transfer simulations were carried out using CFX solver for four coolant to mainstream mass flow ratios of 0.45%, 1%, 1.5% and 3%. The temperature distribution and film cooling effectiveness in the TE region were compared to two conventional geometries, pressure side cutback and centerline ejection which are widely used in vanes and blades for both land-based and aviation gas turbine engines. Unstructured mesh was generated for both fluid and solid domains and interfaces were defined between the two domains. For turbulence closer, SST-kω model was used. The wall y+ was maintained < 1 by using inflation layers at all the solid fluid interfaces. The numerical results depicted that the alternating discharge from the wavy TE was able to form protective film coverage on both the pressure and suction side of the blade. As a result, significant reduction in the TE metal was observed which was up to 14% lower in volume averaged temperature in the TE region when compared to the two baseline conventional configurations.

Numerical Investigation of the Effect of Inlet Flow Angle on Film Cooling Performance for the Blade Tip With Suction Surface Squealer Rim

Numerical investigations have been performed to study the effect of incidence angle on the aerodynamic and film cooling performance for the suction surface squealer tip with different film-hole arrangements at τ = 1.5% and BR = 1.0. Meanwhile, the full squealer tip as baseline is also investigated. Three incidence angles at design condition (0 deg) and off-design conditions (± 7 deg) are investigated. The suction surface, pressure surface, and the camber line have seven holes each, with an extra hole right at the leading edge. The Mach number at the cascade inlet and outlet are 0.24 and 0.52, respectively. The results show that the incidence angle has a significant effect on the tip leakage flow characteristics and coolant flow direction. The film cooling effectiveness distribution is altered, especially for the film holes near the leading edge. When the incidence angle changes from +7 deg to 0 and −7 deg, the ‘re-attachment line’ moves downstream and the total tip leakage mass flow ratio decreases, but the suction surface tip leakage mass flow ratio near leading edge increases. In general, the total tip leakage mass flow ratio for suction surface squealer tip is 1% greater than that for full squealer tip at the same incidence angle. The total pressure loss coefficient of suction surface squealer tip is larger than that for full squealer tip. The full squealer tip with film holes near suction surface and the suction surface squealer tip with film hole along camber line show high film cooling performance, and the area averaged film cooling effectiveness at positive incidence angle +7 deg is higher than that at 0 and −7 deg. The coolant discharged from film holes near pressure surface only cools narrow region near pressure surface.

Linear Cascade and Wind Tunnel Development for Turbine Blade Tip Heat Transfer Investigations With and Without Film Cooling

Currently, there is a deficit of experimental data for surface heat transfer characteristics and thermal transport processes associated with tip gap flows, and a lack of understanding of performance and behavior of film cooling as applied to blade tip surfaces. As a result, many avenues of opportunity exist for development of creative tip configurations with innovative external cooling arrangements. Overall goals of the present investigations are to reduce cooling air requirements, and reduce thermal loading, with equivalent improvements of thermal protection and structural integrity. Described is the development of experimental facilities, including a Supersonic/Transonic Wind Tunnel and linear cascade, for investigations of surface heat transfer characteristics of transonic turbine blade tips with unique squealer geometries and innovative film cooling arrangements. Note that data from past investigations are used to illustrate some of the experimental procedures and approaches which will be employed within the investigation. Of interest is development of a two-dimensional linear cascade with appropriate cascade airfoil flow periodicity. Included are boundary layer flow bleed devices, downstream tailboards, and augmented cascade inlet turbulence intensity. The present linear cascade approach allows experimental configuration parameters to be readily varied. Tip gap magnitudes are scaled so that ratios of tip gap to inlet boundary layer thickness, ratios of tip gap to blade axial chord length, and ratios of tip gap magnitudes to blade true chord length match engine hardware configurations. Ratios of inlet boundary layer thickness to tip gap range from 3 to 5. Innovative film cooling configurations are utilized for one blade tip configuration, and scaled engine components are modelled and tested with complete external cooling arrangements. Blade tip and geometry characteristics are also considered, including squealer depth and squealer tip wall thickness. With these experimental components, results will be obtained with engine representative transonic Mach numbers, Reynolds numbers, and film cooling parameters, including density ratios, which are achieved using foreign gas injection with carbon dioxide. Transient, infrared thermography approaches will be employed to measure spatially-resolved distributions of surface heat transfer coefficients, adiabatic surface temperature, and adiabatic film cooling effectiveness.

Numerical Investigation on Cooling Air Flow and Resistance Characteristics Inner Laminated Cooling Structures

Today, laminated cooling structures have been widely used in the designs of advanced gas turbines, because the structures with double walls, pins, impingement holes and film holes can provide much higher overall cooling effectiveness than simple film cooling. Of course, this kind of cooling structures also leads to a higher price due to a larger flow resistance to cooling air injection in comparison with the simple film cooling. The previous investigations concerned with the laminated cooling structures mainly focused on heat transfer performances, the flow resistance characteristics within the complex channel of the structures are relatively less. This paper presents a numerical investigation on the characteristics of the cooling air resistance passing through 6 different laminated structures. The influence factors on the fluid flow and resistance performances of cooling air, such as array, density and shape of film hole, as well as impingement-hole area (diameter), are discussed and compared at the same pressure ratios of the inlet to outlet of the 6 laminated structures. The discussions and comparisons reveal the following interesting phenomena: 1) A larger diameter of impingement hole corresponds to a larger mass flow rate of cooling air at the inlet of the laminated structure, but the inlet velocity is mainly dependent on the density of film hole. At the same total area of film holes, a larger density corresponds to a higher inlet velocity. 2) The flow rate through film hole of laminated structures is influenced more and more obvious by the outlet shape and the inflow angle of film hole as the increasing pressure ratio. 3) The resistance coefficients of the entire laminated structures are dependent on the density and shape of film holes. At the same total area of film holes, a higher density corresponds to a lower resistance coefficient. Although fan-shaped film hole can provide a larger cooling air coverage, the price is a higher resistance coefficient. Therefore, the applications of fan-shaped film holes in the laminated structures should be considered only in the regions with low environment pressures.