Effect of Cascade Surface Roughness on Boundary Layer Flow Under Variable Conditions

Under the influence of many factors, the surface roughness of the cascade will change during turbomachinery operation, which will affect the boundary layer flow of the cascade. In this article, the effects of cascade surface roughness on boundary layer flow under variable conditions are analyzed by experiments and numerical simulation. The results show that with the increase of roughness, the total pressure loss coefficient of the cascade decreases first and then increases. The larger the Reynolds number is, the greater the total pressure loss coefficient is, and the sensitive area of loss change is changed. In the sensitive area, the roughness has a greater influence on cascade loss. There are separation bubbles at the suction front edge of smooth cascades. With the increase of roughness, the degree of turbulence increases, and the transition process is accelerated. When the roughness is between 74 and 150 μm, the separation bubble disappears and the separation loss decreases. In conclusion, the aerodynamic loss of the cascade increases with the increase of roughness, and the cascade efficiency decreases. However, roughness can restrain the flow separation and reduce the separation loss. The two have gone through a process of one and the other. When the roughness is 74 μm, the displacement thickness, momentum thickness, and shape factor at the back of the cascade are the minimum.

Maximum Thickness Location Selection of High Subsonic Axial Compressor Airfoils and Its Effect on Aerodynamic Performance

Solidity and camber angle are key parameters with a primary effect on airfoil diffusion. Maximum thickness location has a considerable impact on blade loading distribution. This paper investigates correlations of maximum thickness location, solidity, and camber angle with airfoil performance to choose maximum thickness location quickly for compressor airfoils with different diffusion. The effects of maximum thickness location, solidity, and camber angle on incidence characteristics are discussed based on abundant two-dimensional cascade cases computed through numerical methods. Models of minimum loss incidence, total pressure loss coefficient, diffusion factor, and static pressure rise coefficient are established to describe correlations quantitatively. Based on models, dependence maps of total pressure loss coefficient, diffusion factor, and static pressure rise coefficient are drawn and total loss variation brought by maximum thickness location is analyzed. The study shows that the preferred selection of maximum thickness location can be the most forward one with no serious shock loss. Then, the choice maps of optimal maximum thickness location on different design conditions are presented. The optimal maximum thickness locates at 20–35% chord length. Finally, a database of optimal cases which can meet different loading requirements is provided as a tool for designers to choose geometrical parameters.

Impacts of tandem configurations on the aerodynamic performance of an axial supersonic through-flow fan cascade

In the current study, the tandem blade technology is applied to an STFF tandem cascade for the first time, and a 2D STFF tandem cascade is preliminarily designed. Through the modification design of the tandem airfoils and their configuration (axial overlap, AO and percent pitch, PP), the coefficients of total pressure loss and loading are reduced by 4% and 8.58%, respectively. Furtherly, the impact of tandem configurations on the performance is parametrically investigated by numerical simulations. The results indicate that compared with AO, the performance under design incidence is more sensitive to PP except for the cases with PP exceeding a threshold value (1.15). PP dominates the loss and load by controlling the evolution of the FB wake and the shock structure of FB and RB, while AO mainly adjusts the entire shock system structure through the change of virtual shape, resulting in the variation in load distribution between FB and RB. It is worth noting that the overall loading and the total loss remain unchanged with increasing AO except for the tandem configurations (PP=1.05, AO≤−0.01), which make the flow structure in the gap region undergo a fundamental change. With the optimal tandem configuration (PP=1.05, AO=−0.01) and the modified tandem blades (The ratios of chord length and camber for FB over RB is 0.67 and 0.5, respectively), the total pressure loss coefficient is further reduced by 19.7% in comparison with the preliminary tandem design.

Bayesian Optimization Design of Inlet Volute for Supercritical Carbon Dioxide Radial-Flow Turbine

The radial-flow turbine, a key component of the supercritical CO2 (S-CO2) Brayton cycle, has a significant impact on the cycle efficiency. The inlet volute is an important flow component that introduces working fluid into the centripetal turbine. In-depth research on it will help improve the performance of the turbine and the entire cycle. This article aims to improve the volute flow capacity by optimizing the cross-sectional geometry of the volute, thereby improving the volute performance, both at design and non-design points. The Gaussian process surrogate model based parameter sensitivity analysis is first conducted, and then the optimization process is implemented by Bayesian optimization (BO) wherein the acquisition function is used to query optimal design. The results show that the optimized volute has better and more uniform flow characteristics at design and non-design points. It has a smoother off-design conditions performance curve. The total pressure loss coefficient at the design point of the optimized volute is reduced by 33.26%, and the flow deformation is reduced by 54.55%.

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

Film Cooling Effectiveness Enhancement Using Multi-Longitudinal Vortex Generated by Alternating Elliptical Film Holes

This paper proposed an alternating elliptical film hole for gas turbine blade to restrain kidney vortex and enhance film cooling effectiveness, based on the multi-longitudinal vortexes generated in alternating elliptical tube. The detailed flow structures in film hole delivering tube and out of the film hole, adiabatic film cooling effectiveness distributions as well as the total pressure loss coefficient were investigated. The delivering tube of alternating elliptical film hole consists of two straight sections and a transition section. In the straight sections, the cross section of the film hole is elliptical, and in the transition section, along flow direction, the major axis gradually shortened into the minor axis, and the minor axis gradually expanded to the major axis. But, the cross-section area of the film hole kept constant. Numerical simulations were performed by using 3D steady flow solver of Reynolds-averaged Navier-Stokes equations (RANS) with the SST k-ω turbulence model. To reveal the mechanism of kidney vortex suppression and film cooling effectiveness enhancement, the simulation results were compared with the cylindrical film hole set as the baseline at different mass flow ratios (MFR). Besides, the aerodynamic characteristics of these two kinds of film holes were also investigated. The results showed that obvious jet effect could be found in the cylindrical film hole, and the coolant mainly flowed along the upper wind wall, then interacted with the main flow, forming a strong kidney vortex after flowing out, which made the coolant to lift away from the wall surface and reduced the cooling effectiveness. The alternating elliptical film hole had a good inhibition impact on the jet effect in the hole due to the longitudinal vortices, which made the film adhere to the wall surface better after the coolant flowed out. The longitudinal vortices generated by alternating elliptical film hole have the opposite rotation direction to the vorticity of the kidney vortices, thus the kidney vortices were restrained to a certain extent. The height of kidney vortices is lower, and the size of kidney vortices is also smaller. As a result, the film cooling effectiveness of alternating elliptical film hole is distinctly higher than that of the cylindrical film hole, and the enhancement effect is more significant at higher mass flow ratio. In addition, the total pressure loss coefficient of alternating elliptical film hole is only slightly higher than the cylindrical film hole at the mass flow ratio of 1%, 2% and 3%, and is even lower at the mass flow ratio of 4%, thus inducing an excellent comprehensive performance.

Experimental Measurements and Numerical Investigations on the Aerodynamic Performance and Internal Flow Fields of Tangential Admission Volutes for Steam Turbines

Steam turbines are applied in production plants characterized by very large injections of steam. For this reason, the design and optimization of the admission are fundamentals to obtain an adequate level of turbine efficiency and ensure uniform flow at the inlet of the low pressure stages downstream the injection. In conjunction with a flexible operation and partial load conditions, it is important to estimate the losses appearing at those admissions sufficiently. The aerodynamic performance and flow field of the individual tangential admission volute and tangential admission volute coupled with the downstream vanes were experimentally measured and numerically simulated in this paper. The total pressure loss, outlet flow angle and mass flow rate of the individual tangential admission volute at three different outlet Mach numbers and tangential admission volute coupled with the downstream vanes at four different inlet total pressures were measured. The flow field of the experimental tangential admission volute for the steam turbine was numerically investigated using threedimensional Reynolds-Averaged Navier-Stokes (RANS) and SST turbulence model. The numerical aerodynamic parameters of the tangential admission volute were in good agreement with the experimental data. The accuracy of the presented numerical method was validated. The flow field and aerodynamic parameters of tangential admission volute coupled with the downstream vanes were discussed under different inlet total pressure flow conditions. Then, three volute cases with different transverse distance were designed to investigate the influence of different outlet airflow angles on the aerodynamic performance of the downstream vanes. Results show that the outlet airflow angle of the individual tangential admission volute and tangential admission volute coupled with the downstream vanes usually keep constants when the inlet total pressure is increases. The averaged outlet airflow angle of the individual tangential admission volute and tangential admission volute coupled with the downstream vanes equal 152.0° and 166.6°, respectively. Comparing with the individual tangential admission volute, the outlet airflow angle of the tangential admission volute coupled with the downstream vanes is more uniform. The total pressure loss and mass flow rate of the individual tangential admission volute and tangential admission volute coupled with the downstream vanes increase with the inlet total pressure. With the increase of the inlet total pressure, the total pressure loss coefficient of the individual tangential admission volute increases from 0.73% to 1.64%. In the same case, the total pressure loss of the tangential admission volute coupled with the downstream stator vane increases from 0.82% to 2.66%. The average airflow angle of the volute increases with the increase of the transverse distance. With the increase of the transverse distance, the total pressure loss coefficient of the volute increases and the total pressure loss coefficient of the vanes decreases. At the same time, the total pressure loss coefficient of the whole model decreases at first and then increases. The present work provides the reference for the design and performance analysis of the tangential admission volute for the steam turbines.

Redesigning a compressor inlet guide vane for 0 to 60 degree stagger angles

This report's objective is to reduce the total pressure loss coefficient of an inlet guide vane (IGV) at high stagger angles and to therefore reduce the overall fuel consumption of an aircraft engine. IGVs are usually optimized for cruise where the stagger angle is approximately 0 degrees. To reduce losses, four different methodologies were tested: increasing the leading edge radius, increasing the camber, creating a "drooped nose", and creating an "S" curvature distribution. A baseline IGV was chosen and modified using these methodologies to create 10 new IGV designs. CFX was used to perform a CFD analysis on all 11 IGV designs at 5 stagger angles from 0 to 60 degrees. Typical missions were analyzed and it was discovered that the new designs decreased the fuel consumption of the engine. The IGV with the "S" curvature and thicker leading edge was the best and decreased the fuel consumption by 0.24%.

Redesigning a compressor inlet guide vane for 0 to 60 degree stagger angles

This report's objective is to reduce the total pressure loss coefficient of an inlet guide vane (IGV) at high stagger angles and to therefore reduce the overall fuel consumption of an aircraft engine. IGVs are usually optimized for cruise where the stagger angle is approximately 0 degrees. To reduce losses, four different methodologies were tested: increasing the leading edge radius, increasing the camber, creating a "drooped nose", and creating an "S" curvature distribution. A baseline IGV was chosen and modified using these methodologies to create 10 new IGV designs. CFX was used to perform a CFD analysis on all 11 IGV designs at 5 stagger angles from 0 to 60 degrees. Typical missions were analyzed and it was discovered that the new designs decreased the fuel consumption of the engine. The IGV with the "S" curvature and thicker leading edge was the best and decreased the fuel consumption by 0.24%.

Performance characteristics of flow in annular diffuser using CFD

An annular diffuser is a critical component of the turbomachinery, and its prime function is to reduce the flow velocity. The current work is carried to study the effect of four different geometrical designs of an annular diffuser using the ANSYS Fluent. The numerical simulations were carried out to examine the effect of fully developed turbulent swirling and non-swirling flow. The flow behavior of the annular diffuser is analyzed at Reynolds number 2.5 × 105. The simulated results reveal pressure recovery improvement at the casing wall with adequate swirl intensity at the diffuser inlet. Swirl intensity suppresses the flow separation on the casing and moves the flow from the hub wall to the casing wall of the annulus region. The results also show that the Equal Hub and Diverging Casing (EHDC) annular diffuser in comparison to other diffusers has a higher static pressure recovery (C p = 0.76) and a lower total pressure loss coefficient of (C L = 0.12) at a 17° swirl angle.