High Pressure Gas Turbine Blade Research Articles

Conjugate heat transfer simulations of a radially cooled gas turbine blade leading edge using a vortex-based fluidic oscillator for sweeping jet impingement

The turbine blades of aero engines are subjected to extremely high temperatures, particularly at the leading edge, where temperatures can reach approximately 1800–2000 K. Therefore, effective heat load management is crucial. A vortex-based fluidic oscillator for sweeping jet impingement was proposed as an innovative cooling method to enhance heat transfer at leading edge of high-pressure gas turbine blades. This numerical investigation evaluates the cooling performance of a vortex-based sweeping jet compared to steady and conventional sweeping jets in a radially cooled high-pressure turbine blade. In this study, a conjugate heat transfer model based on three-dimensional unsteady Reynolds-averaged Navier–Stokes (URANS) equations is employed. The shear stress transport (SST k–ω) model is specifically selected to predict the flow field and heat transfer characteristics of a vortex-based fluidic oscillator applied to the leading edge. To verify the accuracy of numerical calculations, two sets of experimental data were used as benchmark. The results demonstrated strong qualitative and quantitative agreement with experimental data. Various parameters, including coolant mass flow rates (0.171, 0.514, and 0.857 g/s), aspect ratios (0.5, 0.65, and 1), jet-to-wall spacings (H/D = 2, 4, and 6), and pressure drop, were examined to assess overall cooling effectiveness and heat transfer performance. Time-averaged and time-resolved flow field measurements revealed that vortex-based fluidic oscillator significantly enhanced cooling effects and covered a larger impinging area compared to a steady jet. Notably, the vortex-based fluidic oscillator achieved a 24.3% higher heat transfer performance than the steady jet at H/D = 2, with an average temperature decrease in approximately 21 K at leading edge.

The High Pressure Gas Turbine Blades of Jet Engines

The modern turbo-engine is one of the finest examples of engineering ingenuity, complexity and usefulness. The turbo-engine, whether used in land-based power generating equipment or in aircraft propulsion, has had a profound impact on our lives. Increasing air-travel, rising fuel prices, and environmental concerns have all combined to increase the need for more fuel-efficient engines. The blades of the high-pressure turbine are the most severely loaded members in the turbo-engine. The various aspects of these blades are described. First, the major components of a turbo-engine as well as the different types of turbo-engines are briefly described. Then the development of the turbo-engine is placed in its historical perspective. Next, the severity of the blade loading conditions, such as high temperatures, high centrifugal stresses, corrosion, oxidation, sulphidation, foreign object damage, etc., and the resulting failure modes are summarized. Super alloys are the materials of choice for these blades. The incremental improvements in the alloy chemistry are briefly traced. Important developments in the blade manufacture, the different types of blade cooling channels and their machining, thermal barrier coatings to decrease the blade temperature while increasing the Turbine Inlet Temperature (TIT), and future trends are discussed. The future belongs inevitably to advanced ceramics and their composites, with coatings for environmental protection.

Numerical investigation of crack initiation in high-pressure gas turbine blade subjected to thermal-fluid-mechanical low-cycle fatigue

Numerical investigation of crack initiation in high-pressure gas turbine blade subjected to thermal-fluid-mechanical low-cycle fatigue

The Effect of Heat Transfer on Turbine Performance

Abstract For over 50 years, high-pressure gas turbine blades have been cooled using air bled from the compressor. This cooling results in very high rates of heat transfer, both within the fluid and within the blade. The heat transfer often occurs across large differences in temperature and thus is highly irreversible. It is therefore surprising that little is understood about the effect of this heat transfer on turbine performance. This paper solves this problem by applying a new method known as mechanical work potential, or euergy, analysis. The key consequence of the analysis is that the value placed on all heat, relative to work, becomes set by the Joule (Brayton) cycle efficiency. This means that when heat is transferred locally within a flow, or when viscous reheat occurs, the value of this heat should be set by the Joule cycle efficiency. This paper demonstrates how the new method can be implemented both in the preliminary design systems and in the analysis of conjugate CFD solutions of complex engine representative components. The new method provides the cooling designer with a new way of raising turbine efficiency, a form of recuperation locally in the flow. This method offers the exciting potential to design cooling systems that, when added to a blade profile, actually reduces profile loss by up to 7.3%.

The Effect of Heat Transfer on Turbine Performance

Abstract For over 50 years, high-pressure gas turbine blades have been cooled using air bled from the compressor. This cooling results in very high rates of heat transfer, both within the fluid and within the blade, shown in figure 1. The heat transfer often occurs across large temperature differences and is thus highly irreversible. It is therefore surprising that little is understood about the effect of this heat transfer on turbine performance.

Effect of in-service burnout on the transonic tip leakage flows over flat tip model

Un-shrouded turbine blades are more common than shrouded ones in gas turbine aero-engines since they reduce the weight and avoid the centrifugal loading caused by the blades’ shrouds. Despite these important advantages, the absence of the shroud leads to leakage flows across the tip gap and exposes the blade tip to high thermal load and thermal damages. In addition, the leakage flows can contribute up to 30% of the aerodynamic loss in a turbine stage. In this study, the effect of in-service burnout is explored using a fundamental flat tip model of a high-pressure gas turbine blade. This investigation is carried out both experimentally in a transonic wind tunnel and computationally using the Reynolds Averaged Navier-stokes approach at high-speed conditions. It is found that exposing the tip to the in-service burnout effect changes the leakage flow behaviour significantly when compared with the tip with sharp edges (i.e. the tip at the start of its operational life). Different flow acceleration, flow structure and shockwave pattern and interactions are captured for the round-edge flat tip (i.e. the tip exposed to in-service burnout). The effective tip gap is found to be much larger for the round-edge flat tip allowing more leakage flow into the tip gap which results into higher tip leakage losses in comparison to the sharp-edge tip. Experimental and computational flow visualisations, surface pressure distributions and discharge coefficient are given and analysed for several pressure ratios over the tip gap.

Heat Transfer Investigation on an Internal Cooling System of a Gas Turbine Leading Edge Model

Abstract A scaled up test model simulating a realistic leading edge cooling system of a high pressure gas turbine blade was designed with the aim of performing heat transfer measurements in static and rotating conditions. The test model is composed by a trapezoidal supply channel which feeds three large racetrack holes, generating coolant impingement on the internal concave leading edge surface. Four big fins allow to confine the impingement jets impact zones. Air is then extracted through 4 rows of 6 holes each, two of showerhead (SH) and two of film cooling (FC). The test model is installed on a rotating test rig, which allows to reach jet Reynolds numbers (Re j ) up to 40000 and rotation numbers (Ro j ) up to 0.05. The effect of cross-flow in the supply channel is also considered. The heat transfer coefficient (HTC) distribution on the internal concave surface was evaluated by means of a steady state technique, using wide band thermochromic liquid crystals (TLCs) to measure the wall temperature and an electrically heated Inconel sheet to provide a constant heat flux to the investigated surface. This paper reports experimental results obtained in static conditions for Re j 10000 and 30000 and for two cross-flow cases representative of blade tip and hub sections. The effects of different mass flow extraction between pressure and suction side is also investigated by varying the mass flow rate through FC and SH holes. The effect of the coolant extraction holes geometry on the Nusselt number distribution is analyzed by comparing the experimental results, reported as 2D Nusselt number (Nu) maps, with a previous investigation on an analogous test model with similar impingement geometry. A CFD campaign was also carried out on the present test rig, exploiting a previously validated computational model. Both numerical and experimental results reveal that the effects of differentiated mass flow extraction and extraction area modification are secondary with respect to the effects of Re j and crossflow variation.

Using morphological analysis to tackle uncertainty at the design phase for a safety critical application

Abstract The gas turbine engine internal air system provides cooling and sealing air to a series of critical subsystems and components such as high pressure gas turbine blades, as well as controlling the thrust load on the turbine and compressor spool assembly. Many potential variations for the internal air system are possible, depending on the requirement, expertise and command of intellectual property. Some subsystems, such as rim seals, pre-swirl systems, and rotating cavities have been the subject of extensive development and analysis leading to robust design solutions. Nevertheless there remains scope for further consideration of the overall system design, and this paper explores the use of a decision analysis tool called morphological analysis applied to the internal air system. Morphological analysis provides an effective means for tackling issues where there is uncertainty, as is the case with many design scenarios, including the internal air system, with some specific parameters and information not available until later in the design phase, after the key geometry has been defined. The problem space comprising seven principal parameters, and a cross consistency matrix which allows identification of compatible and incompatible states are presented.

Nusselt Correlations in a Trailing Edge Cooling System with Long Pedestals and Ribs

Abstract Detailed TLC measurements have been carried out to estimate the heat transfer coefficient in stationary conditions under high Reynolds numbers in a trailing edge cooling system of a high pressure gas turbine blade. The investigated geometry consists of a 30:1 PMMA scaled model reproducing the typical wedge-shaped discharge duct of a trailing edge cooling system with one row of 7 enlarged pedestals. The section of the channel upstream the pedestal region is used to guide the airflow from the radial hub inlet to the tangential trailing edge outlet; in this section three different surfaces have been studied: one is smooth and the other two are ribbed with +60° and -60° angled ribs respect to the radial direction. This work focuses on the pedestal outlet section, by giving a correlation of the variation of the Nusselt number as a function of the Reynolds number, from 10000 to 40000, in the different inter-pedestal regions along the radial direction. These correlations give the regionally averaged heat transfer coefficients, from the hub to the tip of the model for both smooth and ribbed cases. The interest of these results is the use for the design of the trailing edge blade cooling systems, under the investigated Reynolds number range, that it is the typical case of the industrial application.

Influence of water–air ratio on the heat transfer and creep life of a high pressure gas turbine blade

Abstract An analytical model to investigate the influence of Water–Air Ratio (WAR) on turbine blade heat transfer and cooling processes (and thus the blade creep life) of industrial gas turbines is presented. The effects of WAR are emphasised for the modelling of the gas properties and the subsequent heat transfer process. The approach considers convective/film cooling and includes the influence of a thermal barrier coating. In addition, the approach is based on the thermodynamic outputs of a gas turbine performance simulation, heat transfer model, as well as a method that accounts for the changes in the properties of moist air as a function of WAR. For a given off-design point, the variation of WAR (0.0–0.10) was investigated using the heat transfer model. Results showed that with increasing WAR the blade inlet coolant temperature reduced along the blade span. The blade metal temperature at each section was reduced as WAR increased, which in turn increased the blade creep life. The increase in WAR increased the specific heat of the coolant and increased the heat transfer capacity of the coolant air flow. The model can be implemented by using the thermodynamic cycle of the engine, without knowing the turbine cooling details in the conceptual design stage. Also, this generic method assists the end user to understand the effect of operating conditions and design parameter on the creep life of a high pressure turbine blade.

Heat transfer in internal channel of a blade: Effects of rotation in a trailing edge cooling system

The aerothermal performance of a trailing edge (TE) internal cooling system of a high pressure gas turbine blade was evaluated under stationary and rotating conditions. The investigated geometry consists of a 30:1 scaled model reproducing a typical wedge shaped discharge duct with one row of enlarged pedestals. The airflow pattern inside the device simulates a highly loaded rotor blade cooling scheme with a 90 [deg] turning flow from the radial hub inlet to the tangential TE outlet. Two different tip configurations were tested, the first one with a completely closed section, the second one with a 5 holes outlet surfaces discharging at ambient pressure. In order to assess rotation effects, a rotating test rig, composed of a rotating arm holding both the PMMA TE model and the instrumentation, was purposely developed and manufactured. A thin Inconel heating foil and wide band Thermo-chromic Liquid Crystals are used to perform steady state heat transfer measurements on the blade pressure side. A rotary joint ensures the pneumatic connection between the blower and the rotating apparatus; moreover several slip rings are used for both instrumentation power supply and thermocouple connection. A parallel CFD analysis involving steady-state RANS modeling was conducted to allow an insight of the flow field inside the redirecting channel and the interpedestal ducts to better interpret the developing vortical structures. Low-Reynolds grid clustering permits to integrate up to the wall both the momentum and the thermal boundary layer. Calculations were performed by means of an in-house developed pressure based solver exploiting the k-ω SST turbulence model implemented in the framework of the open-source finite volume discretization toolbox OpenFOAM®. Analyzed flow conditions correspond to Reynolds number of 20000 in the hub inlet section and angular speed varies to obtain rotation numbers in the range from 0 to 0.3. The orientation of the rotation axis is orthogonal to the heated surface as to resemble a 90 [deg] blade metal angle. Results are reported in terms of detailed heat transfer coefficient 2D maps on the suction side surface as well as spanwise profiles inside the pedestal ducts.