This paper presents a detailed flow and heat transfer characteristic analysis on a gas turbine first stage under hot-streak inlet conditions. Simulations were performed for two locations of hot-streak at turbine inlet with respect to the first stage vane, i.e. i) passage center and ii) blade center. The two kinds of inlet conditions have the same mass-averaged total temperature and total pressure. The passage center hot-streak total pressure and total temperature contours are obtained from the rig data published by Butler. Linear interpolation technique is used to move the hot-streak location from passage center to blade center. The ratio of highest temperature in hot-streak to free stream temperature is 2.0. Mixing plane (MP) and Non-linear harmonic (NLH) approaches are used to address the data transport across the rotor-stator station interface. The numerical solution is validated with the test data obtained from the published rig tests. NLH approach predicted the rotor blade surface temperature distributions close to rig data with a percentage deviation of 3%. The change in hot-streak circumferential position from blade center to passage center lead to decreased attenuation of hot-streak due to pronounced cross momentum transport of fluid across the viscous layers. Turbine flow with blade center hot-streak experiences transient periodic fluctuation of heat load on rotor surface. High temperature gradients are observed at turbine exit station with passage center hot-streak.

In modern gas turbines, film cooling is one of the widely used external cooling techniques for turbine vanes and blades. The turbine airfoil leading edge, which is highly loaded thermally, is currently protected from the hot gas by film cooling schemes, so called showerhead cooling. Flow field in film cooling is very complex and detailed knowledge of heat transfer rates and metal temperatures are required while designing these cooling systems. Computational Fluid Dynamics (CFD) is gaining popularity for modeling these complex cooling systems. However, the application of CFD depends on its accuracy and reliability. This requires the CFD code to be validated with laboratory measurements to ensure its predictive capacity. In this regard, a project has been taken to validate the commercially available CFD code for predicting the blade heat transfer characteristics with shower head film cooling. The validation is accomplished with the test results of Ames [5]. C3X vanes were used for their four vane cascade test facility. The showerhead array used consists of 5 rows of 20° spanwise slanted holes. Experiments were carried out with lower (1%) and higher (12%) turbulence intensities. Results of metal temperatures and heat transfer coefficients were reported. The objective of this study is to validate and calibrate a commercially available CFD code, against the available test data [5] and to understand the relationship between complex flow fields and heat transfer behavior. STAR-CCM+ is used for model generation, mesh generation and solution. Polyhedral elements with prism layers around the wall surfaces are generated. Three turbulence models, Durbin’s v2f model, Menter SST and SST transition models are explored in this study. Simulations are performed for two turbulence intensities available. Typical flow parameters such as blade surface heat transfer coefficient (HTC), surface temperatures and the location of flow transition are compared. Results were compared for two typical cascade exit Mach number conditions such as 0.2 and 0.7, which represents subsonic and transonic conditions respectively. Except in suction side transition region, numerically simulated heat transfer coefficient and Stanton number matched well with test data. Vane wall temperature contours were presented to understand the heat transfer behavior. The heat transfer behavior was numerically investigated for realistic exit Mach numbers. Sensitivity study for two inlet free stream turbulence intensities and three inlet free stream turbulence length scales are performed for realistic exit Mach number and reported heat transfer coefficient and Stanton number.

The knowledge of heat loads on the turbine is of great interest to turbine designers. Turbulence intensity and stator-rotor axial gap plays a key role in affecting the heat loads. Flow field and associated heat transfer characteristics in turbines are complex and unsteady. Computational fluid dynamics (CFD) has emerged as a powerful tool for analyzing these complex flow systems. Honeywell has been exploring the use of CFD tools for analysis of flow and heat transfer characteristics of various gas turbine components. The current study has two objectives. The first objective aims at development of CFD methodology by validation. The commercially available CFD code Fine/Turbo is used to validate the predicted results against the benchmark experimental data. Predicted results of pressure coefficient and Stanton number distributions are compared with available experimental data of Dring et al. [1]. The second objective is to investigate the influence of turbulence (0.5% and 10% Tu) and axial gaps (15% and 65% of axial chord) on flow and heat transfer characteristics. Simulations are carried out using both steady state and harmonic models. Turbulence intensity has shown a strong influence on turbine blade heat transfer near the stagnation region, transition and when the turbulent boundary layer is presented. Results show that a mixing plane is not able to capture the flow unsteady features for a small axial gap. Relatively close agreement is obtained with the harmonic model in these situations. Contours of pressure and temperature on the blade surface are presented to understand the behavior of the flow field across the interface.