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J. Turbomach.. 2015;137(9):091001-091001-10. doi:10.1115/1.4029598.

The choice of the stator hub configuration (i.e., cantilevered versus shrouded) is an important design decision in the preliminary design stage of an axial compressor. Therefore, it is important to understand the effect of the stator hub configuration on the aerodynamic performance. In particular, the stator hub configuration fundamentally affects the leakage flow across the stator. The effect of the stator hub configuration on the leakage flow and its consequent aerodynamic mixing loss with the main flow within the stator row is systematically investigated in this study. In the first part of the paper, a simple model is formulated to estimate the leakage loss across the stator hub as a function of fundamental stage design parameters, such as the flow coefficient, the degree of reaction, and the work coefficient, in combination with some relevant geometric parameters including the clearance/span, the pitch-to-chord ratio, and the number of seals for the shrouded geometry. The model is exercised in order to understand the effect of each of these design parameters on the leakage loss. It is found that, for a given flow coefficient and work coefficient, the leakage loss across the stator is substantially influenced by the degree of reaction. When a cantilevered stator is compared with a shrouded stator with a single seal at the same clearance, it is shown that a shrouded configuration is generally favored as a higher degree of reaction is selected, whereas a cantilevered configuration is desirable for a lower degree of reaction. Further to this, it is demonstrated that, for shrouded stators, an additional aerodynamic benefit can be achieved by using multiple seals. The second part of the paper investigates the effect of the rotating surfaces. Traditionally, only the pressure loss has been considered for stators. However, the current advanced computational fluid dynamics (CFD) generally includes the leakage path with associated rotating surfaces, which impart energy to the flow. It is shown that the conventional loss coefficient, based on considering only the pressure loss, is misleading when hub leakage flows are modeled in detail, because there is energy addition due to the rotation of the hub or the shroud seals for the cantilevered stator and the shrouded stator, respectively. The calculation of the entropy generation across the stator is a better measure of relative performance when comparing two different stator hub configurations with detailed CFD.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091002-091002-12. doi:10.1115/1.4029713.

In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the exit Mach number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at exit Mach numbers of Mexit = 0.85 (Reexit = 9.75 × 105) and Mexit = 1.0 (Reexit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The exit Reynolds number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with exit Mach number due to high turbulence generation in the tip gap and flow reattachment. While increase in exit Mach number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in exit Mach number was found to also induce strong flow relaminarization on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in exit Mach number but significant increase in local heat transfer was noted in this region.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091003-091003-13. doi:10.1115/1.4029616.

This paper presents a comprehensive assessment of real gas effects on the performance and matching of centrifugal compressors operating in supercritical CO2. The analytical framework combines first principles based modeling with targeted numerical simulations to characterize the internal flow behavior of supercritical fluids with implications for radial turbomachinery design and analysis. Trends in gas dynamic behavior, not observed for ideal fluids, are investigated using influence coefficients for compressible channel flow derived for real gas. The variation in the properties of CO2 and the expansion through the vapor-pressure curve due to local flow acceleration are identified as possible mechanisms for performance and operability issues observed near the critical point. The performance of a centrifugal compressor stage is assessed at different thermodynamic conditions relative to the critical point using computational fluid dynamics (CFD) calculations. The results indicate a reduction of 9% in the choke margin of the stage compared to its performance at ideal gas conditions due to variations in real gas properties. Compressor stage matching is also impacted by real gas effects as the excursion in corrected mass flow per unit area from inlet to outlet increases by 5%. Investigation of the flow field near the impeller leading edge at high flow coefficients shows that local flow acceleration causes the thermodynamic conditions to reach the vapor-pressure curve. The significance of two-phase flow effects is determined through a nondimensional parameter that relates the time required for liquid droplet formation to the residence time of the flow under saturation conditions. Applying this criterion to the candidate compressor stage shows that condensation is not a concern at the investigated operating conditions. In the immediate vicinity of the critical point however, this effect is expected to become more prominent. While the focus of this analysis is on supercritical CO2 compressors for carbon capture and sequestration (CCS), the methodology is directly applicable to other nonconventional fluids and applications.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091004-091004-10. doi:10.1115/1.4029879.

Film cooling and sprayed thermal barrier coatings (TBCs) protect gas turbine components from the hot combustion gas temperatures. As gas turbine designers pursue higher turbine inlet temperatures, film cooling and TBCs are critical in protecting the durability of turbomachinery hardware. One obstacle to the synergy of these technologies is that TBC coatings can block cooling holes when applied to the components, causing a decrease in the film cooling flow area thereby reducing coolant flow for a given pressure ratio (PR). In this study, the effect of TBC blockages was simulated on film cooling holes for widely spaced cylindrical and shaped holes. At low blowing ratios for shaped holes, the blockages were found to have very little effect on adiabatic effectiveness. At high blowing ratios, the area-averaged effectiveness of shaped and cylindrical holes decreased as much as 75% from blockage. The decrease in area-averaged effectiveness was found to scale best with the effective momentum flux ratio of the jet exiting the film cooling hole for the shaped holes.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091005-091005-11. doi:10.1115/1.4029966.

In this paper, the transient IR-thermography method is used to investigate the effect of showerhead cooling on the film-cooling performance of the suction side of a turbine guide vane working under engine-representative conditions. The resulting adiabatic film effectiveness, heat transfer coefficient (HTC) augmentation, and net heat flux reduction (NHFR) due to insertion of rows of cooling holes at two different locations in the presence and absence of the showerhead cooling are presented. One row of cooling holes is located in the relatively high convex surface curvature region, while the other is situated closer to the maximum throat velocity. In the latter case, a double staggered row of fan-shaped cooling holes has been considered for cross-comparison with the single row at the same position. Both cylindrical and fan-shaped holes have been examined, where the characteristics of fan-shaped holes are based on design constraints for medium size gas turbines. The blowing rates tested are 0.6, 0.9, and 1.2 for single and double cooling rows, whereas the showerhead blowing is maintained at constant nominal blowing rate. The adiabatic film effectiveness results indicate that most noticable effects from the showerhead can be seen for the cooling row located on the higher convex surface curvature. This observation holds for both cylindrical and fan-shaped holes. These findings suggest that while the showerhead blowing does not have much impact on this cooling row from HTC enhancement perspective, it is influential in determination of the HTC augmentation for the cooling row close to the maximum throat velocity. The double-row fan-shaped cooling seems to be less affected by an upstream showerhead blowing when considering HTC enhancement, but it makes a major contribution in defining adiabatic film effectiveness. The NHFR results highlight the fact that cylindrical holes are not significantly affected by the showerhead cooling regardless of their position, but showerhead blowing can play an important role in determining the overall film-cooling performance of fan-shaped holes (for both the cooling row located on the higher convex surface curvature and the cooling row close to the maximum throat velocity), for both the single and the double row cases.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091006-091006-8. doi:10.1115/1.4030016.

This paper presents the application of the gradient span analysis (GSA) method to the multipoint optimization of the two-dimensional LS89 turbine distributor. The cost function (total pressure loss) and the constraint (mass flow rate) are computed from the resolution of the Reynolds-averaged Navier–Stokes equations. The penalty method is used to replace the constrained optimization problem with an unconstrained problem. The optimization process is steered by a gradient-based quasi-Newton algorithm. The gradient of the cost function with respect to design variables is obtained with the discrete adjoint method, which ensures an efficient computation time independent of the number of design variables. The GSA method gives a minimal set of operating conditions to insert into the weighted sum model to solve the multipoint optimization problem. The weights associated to these conditions are computed with the utopia point method. The single-point optimization at the nominal condition and the multipoint optimization over a wide range of conditions of the LS89 blade are compared. The comparison shows the strong advantages of the multipoint optimization with the GSA method and utopia-point weighting over the traditional single-point optimization.

Commentary by Dr. Valentin Fuster
J. Turbomach.. 2015;137(9):091007-091007-11. doi:10.1115/1.4029967.

Transition of the state of the boundary layer from laminar to turbulent plays an important role in the aerodynamic loss generation on turbine airfoils. An accurate simulation of the transition process and of the state of the boundary layer is therefore crucial for prediction of the aerodynamic efficiency of components in rotating machines. A lot of the research in the past years dealt with the transition over laminar separation bubbles, especially concerning flows in low pressure turbines (LPTs) of air jet engines. Nevertheless, bypass transition is also frequent in turbomachines at higher Reynolds numbers as well as for properly designed profiles. Compared with transition over a laminar separation bubble, a bypass transition is experimentally much more difficult to detect with standard measurement techniques. In such cases it becomes necessary to use more sophisticated techniques, such as hot-film anemometry, hot wires, or Preston probes in order to obtain accurate information on the state of the boundary layer. The study presented is carried out using a linear cascade with a LPT blade profile with strong front loading and gentle flow deceleration at the rear suction side of the blade. Measurements were performed at the high-speed cascade wind tunnel of the Institute of Jet Propulsion at engine relevant Mach and Reynolds numbers. Emphasis is put on the evaluation of the different transition processes at midspan and its influence on profile losses. The data postprocessing was adapted for compressible flows, which allows a more accurate determination of the transition area as well as qualitatively better distributions of the wall shear stress. Finally, comparisons with simulations, using computational fluid dynamics (CFD) tools, are performed and fields for improvement of the turbulence and transition models are identified.

Commentary by Dr. Valentin Fuster

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