Assessment of External Heat Transfer Modeling of a Laboratory-Scale Combustor Inside a Pressure-Housing Environment

Many laboratory-scale combustors are equipped with viewing windows to allow for characterization of the reactive flow. Additionally, pressure housing is used in this configuration to study confined pressurized flames. Since the flame characteristics are influenced by heat losses, the prediction of wall temperature fields becomes increasingly necessary to account for conjugate heat transfer in simulations of reactive flows. For configurations similar to this one, the pressure housing makes the use of such computations difficult in the whole system. It is therefore more appropriate to model the external heat transfer beyond the first set of quartz windows. The present study deals with the derivation of such a model which accounts for convective heat transfer from quartz windows external face cooling system, free convection on the quartz windows 2, quartz windows radiative properties, radiative transfer inside the pressure housing and heat conduction through the quartz window. The presence of semi-transparent viewing windows demands additional care in describing its effects in combustor heat transfers. Because this presence is not an issue in industrial-scale combustors with opaque enclosures, it remains hitherto unaddressed in laboratory-scale combustors. After validating the model for the selected setup, the sensitivity of several modeling choices is computed. This enables a simpler expression of the external heat transfer model that can be easily implemented in coupled simulations.

Investigation on the Leading Edge Film Cooling of Counter-Inclined Cylindrical and Laid-Back Holes With and Without Impingement: Part II — Heat Transfer Coefficient

Heat transfer of the counter-inclined cylindrical and laid-back holes with and without impingement on the turbine vane leading edge model are investigated in this paper. To obtain the film cooling effectiveness and heat transfer coefficient, transient temperature measurement technique on complete surface based on double thermochromic liquid crystals is used in this research. A semi-cylinder model is used to model the vane leading edge which is arranged with two rows of holes. Four test models are measured under four blowing ratios including cylindrical film holes with and without impingement tube structure, laid-back film holes with and without impingement tube structure. This is the second part of a two-part paper, the first part paper GT2018-76061 focuses on film cooling effectiveness and this study will focus on heat transfer. Contours of surface heat transfer coefficient and laterally averaged result are presented in this paper. The result shows that the heat transfer coefficient on the surface of the leading edge is enhanced with the increase of blowing ratio for same structure. The shape of the high heat transfer coefficient region gradually inclines to span-wise direction as the blowing ratio increases. Heat transfer coefficient in the region where the jet core flows through is relatively lower, while in the jet edge region the heat transfer coefficient is relatively higher. Compared with cylindrical hole, laid-back holes give higher heat transfer coefficient. Meanwhile, the introduction of impingement also makes heat transfer coefficient higher compared with cross flow air intake. It is found that the heat transfer of the combination of laid-back hole and impingement tube can be very high under large blowing ratio which should get attention in the design process.

Aerothermal Optimization of Fully Cooled Turbine Blade Tips

Optimal turbine blade tip designs have the potential to enhance aerodynamic performance while reducing the thermal loads on one of the most vulnerable parts of the gas turbine. This paper describes a novel strategy to perform a multi-objective optimization of the tip geometry of a cooled turbine blade. The parameterization strategy generates arbitrary rim shapes around the coolant holes on the blade tip. The tip geometry performance is assessed using steady Reynolds-Averaged Navier-Stokes simulations with the k-ω SST model for the turbulence closure. The fluid domain is discretized with hexahedral elements, and the entire optimization is performed using identical mesh characteristics in all simulations. This is done to ensure an adequate comparison among all investigated designs. Isothermal walls were imposed at engine-representative levels to compute the convective heat flux for each case. The optimization objectives were a reduction in heat load and an increase in turbine row efficiency. The multi-objective optimization is performed using a differential evolution strategy. Improvements were achieved in both the aerodynamic efficiency and heat load reduction, relative to a conventional squealer tip arrangement. Furthermore, this work demonstrates that the inclusion of over-tip coolant flows impacts the over-tip flow field, and that the rim-coolant interaction can be used to create a synergistic performance enhancement.

Numerical Investigation on a High Endwall Angle Turbine With Swept-Curved Vanes

In a high endwall angle turbine, large meridional expansion can cause the strong secondary flow at the endwall, which results in a larger endwall flow loss than the small meridional expansion turbine. The endwall heat transfer is strongly affected by secondary flow effect. In order to optimize the endwall flow to reduce the flow loss and optimize the distribution of heat load, the swept-curved method was used in this study. The swept-curved method was investigated on a transonic second stator (S2) with large meridional expansion in a Low-Pressure (LP) Turbine. Validation studies were performed to investigate the aerodynamic and the heat transfer prediction ability of shear stress transport (SST) turbulence model. The influence of different shapes of the stacking line, including forward-swept, backward-swept, positive-curved and negative-curved, were investigated through numerical simulation. The parameterized control of swept-curved height and angle were adopted to optimize the performance of the aerodynamic and heat transfer. 3D flow field calculation captured the relatively accurate flow structures in the parts of endwall and near endwall. Heat transfer behaviors were explored by means of isothermal wall temperature and Nusselt number (Nu) distribution. The results show that the maximal heat transfer coefficient at the leading edge, for the formation of horseshoe vortexes that cause the high velocity towards the endwall. The swept vane can improve the static pressure and heat load distribution at the endwall region, which decreases the area-averaged shroud heat flux by 2.6 percent and the loss coefficient 1.3 percent.

Turbine Nozzle Endwall Phantom Cooling With Compound Angled Pressure Side Injection

Cooling of the turbine nozzle endwall is challenging due to its complex flow field involving strong secondary flows. Increasingly-effective cooling schemes are required to meet the higher turbine inlet temperatures required by today’s gas turbine applications. Therefore, in order to cool the endwall surface near the pressure side of the airfoil and the trailing edge extended area, the spent cooling air from the airfoil film cooling and pressure side discharge slots, referred to as “phantom cooling” is utilized. This paper studies the effect of compound angled pressure side injection on nozzle endwall surface. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Two nozzle test models with a similar film cooling design were investigated, one with an axial pressure side film cooling row and trailing edge slots; the other with the same cooling features but with compound angled injection, aiming at the test endwall. Phantom cooling effectiveness on the endwall was measured using a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Two-dimensional phantom cooling effectiveness distributions on the endwall surface are presented for four MFR (Mass Flow Ratio) values in each test case. Then the phantom cooling effectiveness distributions are pitchwise-averaged along the axial direction and comparisons were made to show the effect of the compound angled injection. The results indicated that the endwall phantom cooling effectiveness increases with the MFR significantly. A compound angle of the pressure side slots also enhanced the endwall phantom cooling significantly. For combined injections, the phantom cooling effectiveness is much higher than the pressure side slots injection only in the endwall downstream extended area.

Investigations on the Influence of Rib Orientation Angle on Film Cooling Performance of Compound Holes

To investigate the effects of the inclined ribs on internal flow structure in film hole and the film cooling performance on outer surface, experimental and numerical studies are conducted on the effects of rib orientation angle on film cooling of compound cylindrical holes. Three coolant channel cases, including two ribbed cross-flow channels (135° and 45° angled ribs) and the plenum case, are studied under three blowing ratios (0.5, 1.0 and 2.0). 2D contours of film cooling effectiveness as well as heat transfer coefficient were measured by transient liquid crystal measurement technique (TLC). The steady RANS simulations with realizable k-ε turbulence model and enhanced wall treatment were performed. The results show that the spanwise width of film coverage is greatly influenced by the rib orientation angle. The spanwise width of the 45° rib case is obviously larger than that of the 135° rib case under lower blowing ratios. When the blowing ratio is 1.0, the area-averaged cooling effectiveness of the 135° rib case and the 45° rib case are higher than that of the plenum case by 38% and 107%, respectively. With the increase of blowing ratio, the film coverage difference between different rib orientation cases becomes smaller. The 45° rib case also produces higher heat transfer coefficient, which is higher than the 135° rib case by 3.4–8.7% within the studied blowing ratio range. Furthermore, the discharge coefficient of the 45° rib case is the lowest among the three cases. The helical motion of coolant flow is observed in the hole of 45° rib case. The jet divides into two parts after being blown out of the hole due to this motion, which induces strong velocity separation and loss. For the 135° rib case, the vortex in the upper half region of the secondary-flow channel rotates in the same direction with the hole inclination direction, which leads to the straight streamlines and thus results in lower loss and higher discharge coefficient.

Effects of Effusion Cooling Pattern Near the Dilution Hole for a Double-Walled Combustor Liner: Part 2 — Flowfield Measurements

The complex flowfield inside a gas turbine combustor creates a difficult challenge in cooling the combustor walls. Many modern combustors are designed with a double-wall that contain both impingement cooling on the backside of the wall and effusion cooling on the external side of the wall. Complicating matters is the fact that these double-walls also contain large dilution holes whereby the cooling film from the effusion holes is interrupted by the high-momentum dilution jets. Given the importance of cooling the entire panel, including the metal surrounding the dilution holes, the focus of this paper is understanding the flow in the region near the dilution holes. Near-wall flowfield measurements are presented for three different effusion cooling hole patterns near the dilution hole. The effusion cooling hole patterns were varied in the region near the dilution hole and include: no effusion holes; effusion holes pointed radially outward from the dilution hole; and effusion holes pointed radially inward toward the dilution hole. Particle image velocimetry (PIV) was used to capture the time-averaged flowfield at approaching freestream turbulence intensities of 0.5% and 13%. Results showed evidence of downward motion at the leading edge of the dilution hole for all three effusion hole patterns. In comparing the three geometries, the outward effusion holes showed significantly higher velocities toward the leading edge of the dilution jet relative to the other two geometries. Although the flowfield generated by the dilution jet dominated the flow downstream, each cooling hole pattern interacted with the flowfield uniquely. Approaching freestream turbulence did not have a significant effect on the flowfield.

Numerical Investigation of Gas Turbine Combustor Liner Film Cooling Slots

A numerical study, based on experimental work of Inanli et al. [1] is conducted to understand the heat transfer characteristics of film cooled test plates that represent the gas turbine combustor liner cooling system. Film cooling tests are conducted by six different slot geometries and they are scaled-up model of real combustor liner. Three different blowing ratios are applied to six different geometries and surface cooling effectiveness is determined for each test condition by measuring the surface temperature distribution. Effects of geometrical and flow parameters on cooling effectiveness are investigated. In this study, Conjugate Heat Transfer (CHT) simulations are performed with different turbulence models. Effect of the turbulent Prandtl Number is also investigated in terms of heat transfer distribution along the measurement surface. For this purpose, turbulent Prandtl number is calculated with a correlation as a function of local surface temperature gradient and its effect also compared with the constant turbulent Prandtl numbers. Good agreement is obtained with two-layered k–ϵ with modified Turbulent Prandtl number.

CFD Investigation of the Flow of Trailing Edge Cooling Slots

Slot film cooling is a popular choice for trailing edge cooling in high pressure (HP) turbine blades because it can provide more uniform film coverage compared to discrete film cooling holes. The slot geometry consists of a cut back in the blade pressure side connected through rectangular openings to the internal coolant feed passage. The numerical simulation of this kind of film cooling flows is challenging due to the presence of flow interactions like step flow separation, coolant-mainstream mixing and heat transfer. The geometry under consideration is a cutback surface at the trailing edge of a constant cross-section aerofoil. The cutback surface is divided into three sections separated by narrow lands. The experiments are conducted in a high speed cascade in Oxford Osney Thermo-Fluids Laboratory at Reynolds and Mach number distributions representative of engine conditions. The capability of CFD methods to capture these flow phenomena is investigated in this paper. The isentropic Mach number and film effectiveness are compared between CFD and pressure sensitive paint (PSP) data. Compared to steady k–ω SST method, Scale Adaptive Simulation (SAS) can agree better with the measurement. Furthermore, the profiles of kinetic energy, production and shear stress obtained by the steady and SAS methods are compared to identify the main source of inaccuracy in RANS simulations. The SAS method is better to capture the unsteady coolant-hot gas mixing and vortex shedding at the slot lip. The cross flow is found to affect the film significantly as it triggers flow separation near the lands and reduces the effectiveness. The film is non-symmetric with respect to the half-span plane and different flow features are present in each slot. The effect of mass flow ratio (MFR) on flow pattern and coolant distribution is also studied. The profiles of velocity, kinetic energy and production of turbulent energy are compared among the slots in detail. The MFR not only affects the magnitude but also changes the sign of production.

Effect of Coolant Injection Angle on Nozzle Endwall Film Cooling: Experimental and Numerical Analysis in Linear Cascade

Gas turbine components can withstand gas temperatures exceeding the melting point of the alloys they’re made of due to increasingly effective cooling methods. Increasing the operating temperature of a gas turbine is key to improving its power density and exhaust heat for steam or combined-cycle efficiency. In the turbine, the component that experiences the highest gas temperature is the vane directly downstream of the combustor; the most complex flow field in a vane occurs near the endwall. In this study, an experimental investigation is carried out to determine the effect of coolant injection angle and mass flow ratio on film effectiveness on the endwall using the pressure sensitive paint technique for various configurations of jump cooling hole configurations. Two rows of angled holes are upstream of an uncooled vane in a three-vane linear cascade. Injection angle including compound angle is varied from 20 to 60 and coolant to mainstream massflux ratio is varied from 0.5% to 3%. Contours of endwall surface film effectiveness are presented along with span-averaged film effectiveness. CFD models of the cascade are developed using a commercial solver to predict film effectiveness for some of the test conditions and comparisons are made between the experimental and numerical results. The CFD models provide further insight into the flow field and explain trends observed in the experiment by understanding the interaction of jump coolant flow with the 3D endwall mainstream flows.