Optical Model of Thermal Radiation Loading System for Turbine Vane Leading Edge

With the increase of combustion temperatures, the thermal radiation effect for hot components in the new generation of aero-engines has become a key factor in the combustion process, cooling structure design, and thermal protection. A radiation loading system can be used as an external heat source to simulate the real thermal environment of hot components in aero-engines. Total receiving power, as well as 3-D heat flux distribution, should better coincide with real conditions. With the aid of freeform optics and the feedback optimization method, the current study develops a concentrating-type radiation heating system fit for the leading-edge surface of a C3X turbine vane. A xenon lamp combined with a freeform reflector was optimized for controllable heat flux. A design method in the area of illumination engineering was innovatively extended for the current model. Considering the effect of polar angular radiative flux distribution of a xenon lamp, a Monte Carlo ray tracing (MCRT) method was adopted to evaluate the optical performance. Feedback modifications based on Bayesian theory were adopted to obtain the optimal shape of the FFS for target heat flux. The current study seeks a feasible way to generate 3-D heat flux distribution for complex curved surfaces, such as turbine vane surfaces, and helps to simulate the real thermal environment of hot components in aero-engines.

Influence of Coolant on Cooling Performance Sensitivity of Internally Convective Turbine Vane

Based on the Brayton cycle for gas-turbine engines, the high thermal efficiency and power output of a gas-turbine engine can be obtainable when the gas-turbine engine operates at high turbine inlet temperatures. However, turbine components e.g., inlet guide vane, rotor blade, and stator vane request high cooling performance. Typically, internal cooling and film cooling are two effective techniques that are widely used to protect high thermal loads for the turbine components in a state-of-the-art gas turbine. Consequently, the high thermal efficiency and power output can be obtained, and the turbine lifespan can be prolonged, also. On top of that, a comprehensive understanding of flow and heat transfer phenomena in the turbine components is very important. As a result, both experiments and simulations have been used to improve the cooling performance of the turbine components. In fact, the cooling air used in the internal cooling and film cooling is partially extracted from the compressor. Therefore, variations in the cooling air affect the cooling performance of the turbine components directly. This paper presents a numerical study on the influence of the cooling air on cooling-performance sensitivity of an internally convective turbine vane, MARK II using the computational fluid dynamics (CFD)/conjugate heat transfer (CHT) with the SST k-ω turbulence model. Result comparisons are conducted in terms of pressure, temperature, and cooling effectiveness under the effects of the inlet temperature, mass flow rate, turbulence intensity, and flow direction of the cooling air. The cooling-performance sensitivity to the coolant parameters is shown through variations of local cooling effectiveness, and area and volume-weighted average cooling effectiveness.

Superposition Effect of the Leading Edge Film On the Downstream Film Cooling of a Turbine Vane Under Combustor Swirling Outflow

To investigate the superposition effect of the leading edge film on the downstream film cooling under swirling inflow, numerical simulations with three vane models (vane with films on the leading edge only, vane with films on the pressure side and suction side only, full-film cooling vane), two inlet conditions (axial inlet and swirling inlet) are conducted. The results indicate that the leading edge is the area where the film is most affected by the swirling inflow. For full-film cooling vane, the film on the leading edge does not always improve or even reduce the downstream film cooling. Flow mechanism analysis shows that the velocity direction near the downstream wall is governed by the interaction between the direction of swirling inflow and the direction of film hole incidence on the leading edge. A new type of leading edge film proposed by the author is also investigated, with the dividing line of the counter-inclined film-hole row coinciding with the twisted stagnant line to ensure that all films are incident at angles inverse to the direction of the swirling inflow. The new leading edge film successfully changes the velocity direction near the downstream wall and suppresses the deflecting effect on the downstream film. The new leading edge film can increase the overall area averaged cooling effectiveness of the full-film cooling vane by 10%, 15%, 18% and reduce the inhomogeneity by 13%, 19%, 27% over the traditional design, as the coolant mass flow increases.

Improvement of Multi-Hole Airflow Impingement on Flow and Heat Transfer Characteristics Inside a Turbine Vane Cavity

The cooling effect of turbine vane is of great importance for ensuring thermal protection and economic operation of gas turbines. This study aims to reveal the influence mechanism and performance of impingement cooling and heat transfer within a turbine guide vane cavity. Then, a turbine guide vane cavity with a complex pin fins structure is numerically investigated at a multi-hole impingement by comparison with experiment verification. The results show that the larger the Reynolds number is, the larger the average Nusselt number is on the upper and lower surfaces of the cavity. The average Nusselt number increased on the upper and lower surfaces as the impingement hole diameter increased. Comparing 1 impingement hole with 16 ones, the average Nusselt number of the lower surface of the latter is 553.9% larger than that of the former. Furthermore, the average Nusselt number of the lower surface for pin fin height of 3 mm is only improved by 11.2% for pin fin height of 24 mm. The heat transfer effect near the impingement holes is better than that far away from the impingement holes. In particular, it is recommended to have 14 impingement holes with a hole diameter of 7.2 mm, as well as circular pin fins with a height of 3 mm and spacing of 25.8 mm. In addition, the entropy generation distribution in impingement cooling is analyzed. This study can provide a reference to enhance the turbine vane cooling performance by optimization design.

Detailed Velocity and Heat Transfer Measurements in an Advanced Gas Turbine Vane Insert Using MRV and IR Thermometry

This work reports the results of paired experiments for a complex internal cooling flow within a gas turbine vane using Magnetic Resonance Velocimetry (MRV) and steady-state Infrared (IR) thermometry. A scaled model of the leading edge insert for a gas turbine vane with multi-pass impingement was designed, built using stereolithography (SLA) fabrication methods, and tested using MRV techniques to collect a three-dimensional, three-component velocity field data set for a fully turbulent test case. Stagnation and recirculation zones were identified and assessed in terms of impact on potential cooling performance. A paired experiment employed an IR camera to measure the temperature profile data of a thin, heated stainless steel impingement surface modeling the inside turbine blade wall cooled by the impingement from the vane cooling insert, providing complementary data sets. The temperature data allow for the calculation of wall heat transfer characteristics, including the Nusselt number distribution for cooling performance analysis to inform design and validate computational models. Quantitative and qualitative comparisons of the paired results show that the flow velocity and cooling performance are highly coupled. Module-to-module variation in the surface Nusselt number distributions are evident, attributable to the complex interaction between transverse and impinging flows within the apparatus. Finally, a comparison with internal heat transfer correlations is conducted using the data from Florschuetz [1]. Measurement uncertainty was assessed and estimated to be approximately 7% for velocity and ranging from 3% to 10% for Nusselt number.