Solid Vane Research Articles

Design Impacts on Ram Air Vane Cooling in an Ultra-Compact Combustor

Abstract The ultra-compact combustor (UCC) aims to decrease the length of gas turbine combustors using a unique design geometry which wraps a combustion chamber around the central axial flow. This distinctive design enables an out of the box type of cooling scheme to be investigated for the turbine inlet vanes, termed the hybrid guide vane (HGV) in the UCC. The leading edge of the HGV experiences only compressor exit air as combustion products do not interact with the vane upstream of the 14% axial chord location. Previous studies were conducted which computationally evaluated the viability of taking in freestream flow through the HGV stagnation region for use as coolant. Based on these studies, a six vane HGV was manufactured which incorporated a solid vane and five hollow vanes. Each of these vanes incorporated different features to vary the size of the internal plug, trailing edge exit, and film cooling holes. In the present study, the cooled HGV was experimentally analyzed using pressure, thermocouple, and infrared (IR) thermography measurements to evaluate internal coolant flowrates and pressure loss along with cooling performance. Furthermore, the vanes were compared to isolate the impact of design differences on vane cooling. It was found that the location of the internal plug and incorporation of film cooling holes had a minor impact on coolant flow and cooling. Additionally, results showed exit area had the largest impact on surface temperature and coolant mass flow where the largest exit area allowed less restricted coolant flow resulting in the lowest average surface temperature. However, completely blocking the exit slot forced coolant to exit only through film cooling holes, stagnating the majority of the internal flow, resulting in surface temperatures higher than the uncooled, solid vane.

Heat Transfer and Pressure Loss in a Two-Pass, Rectangular Channel Featuring a Reduced Cross-Sectional Area After the 180-Deg Tip Turn With Different Turning Vane Configurations

Abstract Serpentine, multi-pass cooling passages are used in cooling advanced gas turbine blades. In the open literature, most internal cooling studies use a fixed cross-sectional area for multi-pass channels. Studies that use varying aspect ratio channels, along with a guide vane to direct the flow with turning, are scarce. Therefore, this study investigates the effect of using different guide vane designs on both detailed heat transfer distribution and pressure loss in a multi-pass channel with an aspect ratio of (4:1) in the entry passage and (2:1) in the second passage downstream of the vane(s). The first vane configuration is one solid vane with a semi-circular cross section connecting the two flow passages. The second configuration has three broken vanes with a quarter-circular cross section; two broken vanes are located downstream in the first passage, and one broken vane is upstream in the second passage. Detailed heat transfer distributions were obtained on all surfaces within the flow passages by using a transient liquid crystal method. Results show that including the semi-circular vane in the turning region enhanced the overall heat transfer by around 29% with a reduction in pressure loss by around 20%. Moreover, results show that the quarter-circular vane design provides higher overall averaged heat transfer enhancement than the semi-circular vane design by around 9% with penalty of higher pressure drop by 6%, which yields higher thermal performance by 7%, over a Reynolds number range from 15,000 to 45,000.

Aero-thermal-elastic Coupled Simulations of An Air-cooled Turbine with an FDM solver

The aero-thermal-elastic coupled simulations of an air-cooled turbine vane were carried out by a developed coupled solver, HIT3D, employing the finite difference method. The pressure on the vane surface and the temperature in the solid vane were obtained by the coupled heat transfer simulation, then the single-way aero-elastic and thermal-elastic analysis on the turbine vane were performed. It shows that the predicted profile temperature employing the transition model agrees well with the measured one, the blade deformation caused by aerodynamic force is negligible, and that the greater thermal deformation and thermal stress locate at the blade leading edge pressure side and at the blade trailing edge, as the result of the strong temperature gradient and the constraint at the vane endwalls.

Analysis of an Additively Manufactured Cooled Ultra Compact Combustor Vane

The ultra compact combustor (UCC) aims to increase the thrust-to-weight ratio of an aircraft gas turbine engine by decreasing the size, and thus weight, of the engine’s combustor. The configuration of the UCC as a primary combustor enables a unique cooling scheme to be employed for the hybrid guide vane (HGV). A previous effort conducted a computational fluid dynamics (CFD) analysis that evaluated whether it would be possible to cool this vane by drawing in freestream flow at the stagnation region of the airfoil. Based on this study, a cooling scheme was designed and modified with internal supports to make additive manufacturing of the vanes possible. This vane was computationally evaluated comparing the results with those of a solid vane and hollow vane without cooling holes as a demonstration of the improvements offered by this design. Furthermore, the effects of the internal support structure were deemed beneficial to surface cooling when evaluated through comparisons of internal pressure distribution and overall effectiveness. Following the computational study, the vane was manufactured and experimentally evaluated with the results compared to those of an uncooled solid vane. The experimental results validated the computational analysis and demonstrated through pressure and temperature measurements that the cooled vane had a reduced surface temperature compared to the uncooled vane and that pressure distributions supported coolant flow through film-cooling holes.

Numerical studies on ejector with swirl generator

This paper presents three-dimensional numerical studies of ejector with swirl generator (SG) using computational fluid dynamics (CFD) software, ANSYS, FLUENT 15.0. Performance analysis is carried out with two different configurations of swirl generators, namely solid vane type and cavity type. As expected, swirl increases the entrainment performance due to better momentum exchange between the motive and secondary streams compared to the ejector without swirl generator. Also, it is confirmed from the studies that, cavity type SG, imparts higher swirl strength to the flow stream than that of the solid vane type, due to the sweep angle in addition to the camber angle. Hence the studies are further extended with cavity type SG, at camber angle of 10° with various sweep angles like 10°, 20° and 30°, for analyzing the swirl effect over the performance of the ejector. The increase in entrainment ratio observed is about 15% with SG, compared to ejector without SG.

Transition modelling implications in the CFD analysis of a turbine nozzle vane cascade tested over a range of Mach and Reynolds numbers

The aerodynamic performance of a gas turbine nozzle vane cascade was investigated over a range of Mach and Reynolds numbers. The work is part of a vast research project aimed at the analysis of fluid dynamics and heat transfer phenomena in cooled blades. In this paper computed results on the “solid vane” (without cooling devices) are presented and discussed in comparison with experimental data. Detailed measurements were provided by the University of Bergamo where the experimental campaign was carried out by means of a subsonic wind tunnel. The impact of boundary layer transition is investigated by using a novel laminar kinetic energy transport model and the widely used Langtry-Menter γ-Reθ,t model. The comparison between calculations and measurements is presented in terms of blade loading distributions, total pressure loss coefficient contours downstream of the cascade, and velocity/turbulence-intensity profiles within the boundary layer at selected blade surface locations at mid-span. It will be shown how transitional calculations compare favorably with experiments.

Influence of Prehistory and Leading Edge Contouring on Aero Performance of a Three-Dimensional Nozzle Guide Vane

Experiments are conducted to investigate the effect of the prehistory in the aerodynamic performance of a three-dimensional nozzle guide vane with a hub leading edge contouring. The performance is determined with two pneumatic probes (five hole and three hole) concentrating mainly on the end wall. The investigated vane is a geometrically similar gas turbine vane for the first stage with a reference exit Mach number of 0.9. Results are compared for the baseline and filleted cases for a wide range of operating exit Mach numbers from 0.5 to 0.9. The presented data includes loading distributions, loss distributions, fields of exit flow angles, velocity vector, and vorticity contour, as well as mass-averaged loss coefficients. The results show an insignificant influence of the leading edge fillet on the performance of the vane. However, the prehistory (inlet condition) affects significantly in the secondary loss. Additionally, an oil visualization technique yields information about the streamlines on the solid vane surface, which allows identifying the locations of secondary flow vortices, stagnation line, and saddle point.

Heat Transfer for the Film-Cooled Vane of a 1-1/2 Stage High-Pressure Transonic Turbine—Part II: Effect of Cooling Variation on the Vane Airfoil and Inner End Wall

The impact of film cooling on heat transfer is investigated for the high-pressure vane of a 1-1/2 stage high-pressure turbine operating at design corrected conditions. Cooling is supplied through three independently controllable circuits to holes in the inner and outer end wall, vane leading edge showerhead, and the pressure and suction surfaces of the airfoil, in addition to vane trailing edge slots. Four different overall cooling flow rates are investigated and one cooling circuit is varied independently. All results reported in this part of the paper are for a radial inlet temperature profile, one of the four profiles reported in part I of this paper. Part I describes the experimental setup, data quality, influence of inlet temperature profile, and influence of cooling when compared to a solid vane. This part of the paper shows that the addition of coolant reduces airfoil Stanton number by up to 60%. The largest reductions due to cooling are observed close to the inner end wall because the coolant to the majority of the vane is supplied by a plenum at the inside diameter. While the introduction of cooling has a significant impact on Stanton number, the impact of changing coolant flow rates is only observed for gauges near 5% span and on the inner end wall. This indicates that very little of the increased coolant mass flow reaches all the way to 90% span and the majority of the additional mass flow is injected into the core flow near the plenum. Turning off the vane outer cooling circuit that supplies coolant to the outer end wall holes, vane trailing edge slots, and three rows of holes on the pressure surface of the airfoil, has a local impact on Stanton number. Changes downstream of the holes on the airfoil pressure surface indicate that internal heat transfer from the coolant flowing inside the vane is important to the external heat transfer, suggesting that a conjugate heat-transfer solution may be required to achieve good external heat-transfer predictions in this area. Measurements on the inner end wall show that temperature reduction in the vane wake due to the trailing edge cooling is important to many points downstream of the vane.

G1000-2-4 蒸気タービン静翼の材料減衰によるフラッタ抑制効果(機械力学・計測制御部門一般講演(2):振動,社会変革を技術で廻す機械工学)

The vane used in a low pressure end of steam turbine is usually fixed to shroud and casing by welding both ends. In such a vane structure, the damping in operation is comprised of the material damping and the aerodynamic damping, because the structural damping is very small. In this paper, the vane is modeled by the uniform beam fixed at both ends, and the effect of the material damping on the vane flutter is studied. The critical material damping for vane flutter and the allowable vibratory stress are calculated and compared for the solid vane and the hollow vane.

The Sebright bantam chicken and the genetic control of extraglandular aromatase.

I. Sexual Dimorphism in Feather Morphology NO ASPECT of sexual dimorphism is more striking or more variable among species than is the feathering pattern of birds. One such sexual difference involves the morphology of feathers. For example, in some birds the feathers of the male are longer and more deeply fringed than those of the female.1 Species in which such sexual differences are extreme include the lyre bird, the peacock, some types of hummingbird, and the domestic chicken (1). Owing to their accessibility, chickens have been the favored subjects for the study of relationships between gonadal function and feather structure, and it is customary to characterize the plumage of adult chickens as either henny or cocky (Fig. 1) (3). In the hen most feathers are short and straight and have a solid vane with many barbules and hence little fringing. In the rooster, in contrast, the neck hackle feathers, the feathers of the saddle (the median tail coverts), and the major tail coverts and great sickle feathers o...