Experiments and numerical computations are performed to investigate the convective heat transfer characteristics of a gas turbine can combustor under cold flow conditions in a Reynolds number range between 50,000 and 500,000 with a characteristic swirl number of 0.7. It is observed that the flow field in the combustor is characterized by an expanding swirling flow, which impinges on the liner wall close to the inlet of the combustor. The impinging shear layer is responsible for the peak location of heat transfer augmentation. It is observed that as Reynolds number increases from 50,000 to 500,000, the peak heat transfer augmentation ratio (compared with fully developed pipe flow) reduces from 10.5 to 2.75. This is attributed to the reduction in normalized turbulent kinetic energy in the impinging shear layer, which is strongly dependent on the swirl number that remains constant at 0.7 with Reynolds number. Additionally, the peak location does not change with Reynolds number since the flow structure in the combustor is also a function of the swirl number. The size of the corner recirculation zone near the combustor liner remains the same for all Reynolds numbers and hence the location of shear layer impingement and peak augmentation does not change.

The paper presents a detailed experimental and numerical study on the effect of endwall contouring in a quasi 2D cascade, operating at transonic conditions. Aerodynamic performance of two contoured endwalls are studied and compared with a baseline (planar) endwall. The first contoured endwall was generated with the goal of reducing secondary losses (Aero-Optimized contoured endwall) and the second endwall was generated with the objective of reduced overall heat transfer to the endwall (HT-optimized contoured endwall). Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface were measured. The cascade exit Mach numbers range from 0.71 to 0.95 and the turning angle of the airfoil is ∼127°. The inlet span of the airfoils was reduced with respect to the outlet span with the intention of obtaining a realistic inlet/exit Mach number that is observed in a real engine. 3D viscous compressible CFD analysis was carried out to study the detailed behavior of the complex flow structures that develop as a result of endwall contouring. A 3% reduction in area averaged losses was achieved at 0.1 Cax downstream of the trailing edge and a 17% reduction in mixed out losses was achieved at 1.0 Cax downstream location with the Aero-Optimized contoured endwall.

We construct the classical mechanics associated with a conformally flat Riemannian metric on a compact, n-dimensional manifold without boundary. The corresponding gradient Ricci flow equation turns out to equal the time-dependent Hamilton-Jacobi equation of the mechanics so defined.

Comparison of heat transfer performance of a nonaxisymmetric contoured endwall to a planar baseline endwall in the presence of leakage flow through stator–rotor rim seal interface and mateface gap is reported in this paper. Heat transfer experiments were performed on a high turning turbine airfoil passage at Virginia Tech's transonic blow down cascade facility under design conditions for two leakage flow configurations—(1) mateface blowing only, (2) simultaneous coolant injection from the upstream slot and mateface gap. Coolant to mainstream mass flow ratios (MFRs) were 0.35% for mateface blowing only, whereas for combination blowing, a 1.0% MFR was chosen from upstream slot and 0.35% MFR from mateface. A common source of coolant supply to the upstream slot and mateface plenum made sure the coolant temperatures were identical at both upstream slot and mateface gap at the injection location. The contoured endwall geometry was generated to minimize secondary aerodynamic losses. Transient infrared thermography technique was used to measure endwall surface temperature and a linear regression method was developed for simultaneous calculation of heat transfer coefficient (HTC) and adiabatic cooling effectiveness, assuming a one-dimensional (1D) semi-infinite transient conduction. Results indicate reduction in local hot spot regions near suction side as well as area averaged HTC using the contoured endwall compared to baseline endwall for all coolant blowing cases. Contoured geometry also shows better coolant coverage further along the passage. Detailed interpretation of the heat transfer results along with near endwall flow physics has also been discussed.

Contouring of turbine endwalls has been widely studied for aerodynamic performance improvement of turbine passages. However, it is equally important to investigate the effect of contouring on endwall heat transfer, because a substantial increase in endwall heat transfer due to contouring will render the design impractical. In this paper, the effect of contouring on endwall heat transfer performance of a high-turning HP-turbine blade passage, operating under transonic exit Mach number conditions, is reported. Three endwall geometries were experimentally investigated at three different passage exit Mach numbers, 0.71, 0.88(design) and 0.95, for their heat transfer performance. One endwall is a non-contoured baseline endwall and the other two are contoured endwall geometries. One of the contoured endwall geometry was generated with the goal of reduction in stagnation pressure losses and the other was generated with the goal of reduced overall heat transfer through the endwall. The experiments were carried out in Virginia Tech’s transient, blow down, transonic linear cascade facility. Endwall surface temperatures were measured using infrared thermography technique. Local heat transfer coefficient values were calculated using the measured temperatures. The heat transfer coefficient values were then related to the endwall geometries using a camera matrix model. The measurement technique and the methodology for the post-processing of the heat transfer coefficient data have been presented in detail. Details of the flow behavior for these endwalls were obtained using CFD simulations and have been used to assist the interpretation of the experimental results. In this study, the heat transfer performance of the contoured endwalls in comparison to the non-contoured baseline case is presented. Both the contoured endwalls demonstrated a significant reduction in the overall average heat transfer coefficient values. The surface heat transfer coefficient distributions also indicated a reduction in the level of hot spots for most of the endwall surface. However, increase in the heat transfer coefficient values was observed especially in the area near the leading edge. The results indicate that, in addition to a probable improvement in aerodynamic performance, endwall contouring may also be used to improve the heat transfer performance of turbine passages. Additionally, aerodynamic behavior of these endwalls is discussed in detail in the companion paper GT2012-68425, “Effect of endwall contouring on a transonic turbine endwall: Part 1 – Aerodynamic performance.”