Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2008 ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient ◽  
Aerodynamic Losses

The performance of suction-side gill region film cooling is investigated using the University of Utah Transonic Wind Tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77 to 1.99 similar to values present in operating gas turbine engines. Presented are local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated Aerodynamic Losses (IAL) increase anywhere from 4 to 45 percent compared to a smooth blade with no film injection. The performance of each hole type depends upon the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL aerodynamic losses are generally present at any particular wake location for the round radial RR or shaped axial SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA shaped axial holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

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

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract 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.

Numerical Investigation of Flow Separation Control of Low-Pressure High-Lift Blade With Different Grooves and Protrusions

2013 ◽  
Author(s):  
Huancheng Qu ◽  
Ping Li ◽  
Jianhui Chen ◽  
Zhongyang Shen ◽  
Yonghui Xie ◽  
...  
Keyword(s):  
Flow Separation ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Flow Separation Control ◽  
Pressure Loss Coefficient ◽  
Flow Loss ◽  
Total Pressure Loss Coefficient

The shear stress transport (SST) turbulence model and γ-Reθ transition model were employed when solving the Reynolds-averaged Navier-Stokes (RANS) equations. The flow separation in the suction side of the typical high-lift low-pressure gas turbine PakB blade was investigated. Different sets of mesh were adopted and the results of grid independence study show that the precision is maintained when the grid system of 126,780 is adopted. And the computational results were compared with the existing experimental and computational results, which indicate that the numerical method can predict the separated transition flow reliably. Different kinds of structures including V grooves and protrusions, curved grooves and protrusions, rectangular grooves and protrusions were used to passive control of the flow separation in the suction side of the PakB blade. The structures have the same locations including 65%Cax, 68%Cax and 71%Cax on the suction side of the blade as well as length and height for better comparison. All of these cases are compared with the flow of PakB cascade without control with Re = 86,000 and FSTI = 1%. The shear layer is uplifted when the flow passes the passive device. And the separation bubble inside the grooves almost occupies the whole groove space which makes the length of the separation shorter than that in the case without control. The separation inside the grooves joins into the downstream separation in the case of grooves located in 71%Cax. The flow starts to separate in the leeside of the protrusion wherever the protrusion locates. And the attachment point moves forward significantly in the comparison with the case of without control. However, it brings in more flow loss because of the protrusion’s resistance to the boundary flow. In the total pressure loss coefficient comparison with the case without control, the grooves produce less flow loss while the protrusions at all the locations bring more flow loss. The nearer the groove is away from the separation point in the case without control, the higher the efficiency could be in the view of total pressure loss coefficient. The rectangular grooves are considered as a more effective structure for the flow separation control of PakB blade. Moreover, the flow separation bubble length and the total pressure loss coefficient decrease as Reynolds number increases in the cases without control. The total pressure loss coefficient in the different Reynolds numbers cases with rectangular groove is lower than that in the cases without control and the flow control performance gets much better when Reynolds number increases.

The Effect of Upstream Ramps With Different Shapes on Film Cooling Efficiency

2017 ◽  
Author(s):  
Daren Zheng ◽  
Xinjun Wang ◽  
Feng Zhang ◽  
Junfei Zhou ◽  
Qi Yuan
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Adiabatic Cooling ◽  
Film Cooling Holes ◽  
Total Pressure Loss Coefficient

This paper presents a numerical investigation on a concept for enhancing the film cooling performance by modifying the shape of upstream ramps. The novel shape ramp, which is placed in front of the film cooling holes, is presented to alter the approaching boundary-layer flow and its interaction with coolant to increase the lateral spreading of the coolant. Five different shape ramps are investigated, including rectangular, wedge-shaped, convex, concave and wave-shaped ramps. The film cooling performance of different shape ramps are evaluated at the density ratio about 1, with blowing ratios ranging from 0.3 to 1.2. The numerical results for the upstream ramp show an agreement with experiment data when solving three dimensional average Navier-Stokes analysis with the k-ε model. Detailed adiabatic cooling effectiveness and total pressure loss coefficient are simulated. Results obtained indicate that film cooling characteristics in the region downstream of the film cooling holes are sensitive to the ramp shapes. The wave-shaped ramp shows the lowest total pressure loss coefficient among these five ramps. For M = 0.3 and 0.6, the highest centerline adiabatic cooling effectiveness occurs in the convex ramp. And this shape ramp also shows the highest spanwise averaged adiabatic cooling effectiveness at the blowing ratio of 0.3. Compared with the other shape ramps, the concave ramp can greatly increase both the centerline and the spanwise averaged adiabatic cooling effectiveness for M = 1.0 and 1.2.

Blowing Ratio Effect on Film Cooling Performance for a Row Holes Installed in Different Trench Configurations

2020 ◽  
Vol 28 ◽  
pp. 65-75
Author(s):  
Khadidja Boualem ◽  
Abbes Azzi
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Ansys Cfx ◽  
Cooling Hole

In the present study, a numerical analysis was performed to evaluate the performance of cooling hole embedded in different trenched designs (triangular trench, semi-cylindrical trench and corrugated trench) in improving the film cooling efficiency over a flat plate. These concepts are compared to the rectangular trenched and the traditional cylindrical hole. The commercial software ANSYS CFX 18 was used to conduct a series of required numerical calculations. The centerline and laterally averaged film cooling effectiveness and total pressure loss coefficient for the five cases are analyzed at three blowing ratios, M=0.5, M=1.0 and M=1.5. Results show a uniform coverage is obtained by hole installed in trench. The main result obtained in this paper that the cooling hole with corrugated trench enhance the film cooling effectiveness with less total pressure loss. The main result of this study reveals that the jet installed in the trenches yield a better film cooling effectiveness especially at higher blowing ratios (M≥1).

The Influence of Different Rim Seal Geometries on Hot-Gas Ingestion and Total Pressure Loss in a Low-Pressure Turbine

2010 ◽  
Author(s):  
P. Schuler ◽  
W. Kurz ◽  
K. Dullenkopf ◽  
H.-J. Bauer
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Gas Flow ◽  
Low Pressure ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Low Pressure Turbine ◽  
Hot Gas ◽  
Flow Through ◽  
Total Pressure Loss Coefficient

In order to prevent hot-gas ingestion into the rotating turbo machine’s inside, rim seals are used in the cavities located between stator- and rotor-disc. The sealing flow ejected through the rim seal interacts with the boundary layer of the main gas flow, thus playing a significant role in the formation of secondary flows which are a major contributor to aerodynamic losses in turbine passages. Investigations performed in the EU project MAGPI concentrate on the interaction between the sealing flow and the main gas flow and in particular on the influence of different rim seal geometries regarding the loss-mechanism in a low-pressure turbine passage. Within the CFD work reported in this paper static simulations of one typical low-pressure turbine passage were conducted containing two different rim seal geometries, respectively. The sealing flow through the rim seal had an azimuthal velocity component and its rate has been varied between 0–1% of the main gas flow. The modular design of the computational domain provided the easy exchange of the rim seal geometry without remeshing the main gas flow. This allowed assessing the appearing effects only to the change of rim seal geometry. The results of this work agree with well-known secondary flow phenomena inside a turbine passage and reveal the impact of the different rim seal geometries on hot-gas ingestion and aerodynamic losses quantified by a total pressure loss coefficient along the turbine blade. While the simple axial gap geometry suffers considerable hot-gas ingestion upstream the blade leading edge, the compound geometry implying an axial overlapping presents a more promising prevention against hot-gas ingestion. Furthermore, the effect of rim seals on the turbine passage flow field has been identified applying adequate flow visualisation techniques. As a result of the favourable conduction of sealing flow through the compound geometry, the boundary layer is less lifted by the ejected sealing flow, thus resulting in a comparatively reduced total pressure loss coefficient over the turbine blade.

Numerical Simulation of Deposition Phenomena on High-Pressure Turbine Vane Using UPACS

2019 ◽  
Author(s):  
Kenta Mizutori ◽  
Koji Fukudome ◽  
Makoto Yamamoto ◽  
Masaya Suzuki
Keyword(s):  
Numerical Simulation ◽  
High Pressure ◽  
Flow Field ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
High Pressure Turbine ◽  
Pressure Loss Coefficient ◽  
Turbine Vane ◽  
Total Pressure Loss Coefficient

Abstract We performed numerical simulation to understand deposition phenomena on high-pressure turbine vane. Several deposition models were compared and the OSU model showed good adaptation to any flow field and material, so it was implemented on UPACS. After the implementation, the simulations of deposition phenomenon in several cases of the flow field were conducted. From the results, particles adhere on the leading edge and the trailing edge side of the pressure surface. Also, the calculation of the total pressure loss coefficient was conducted after computing the flow field after deposition. The total pressure loss coefficient increased after deposition and it was revealed that the deposition deteriorates aerodynamic performance.

scholarly journals Preliminary Experimental Study on Pressure Loss Coefficients of Exhaust Manifold Junction

2014 ◽  
Vol 2014 ◽  
pp. 1-10
Author(s):  
Xiao-lu Lu ◽  
Kun Zhang ◽  
Wen-hui Wang ◽  
Shao-ming Wang ◽  
Kang-yao Deng
Keyword(s):  
Mass Flow ◽  
Pressure Loss ◽  
Total Pressure ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Flow Ratio ◽  
Pressure Loss Coefficient ◽  
Loss Coefficients ◽  
Total Pressure Loss Coefficient ◽  
Mass Flow Ratio

The flow characteristic of exhaust system has an important impact on inlet boundary of the turbine. In this paper, high speed flow in a diesel exhaust manifold junction was tested and simulated. The pressure loss coefficient of the junction flow was analyzed. The steady experimental results indicated that both of static pressure loss coefficientsL13andL23first increased and then decreased with the increase of mass flow ratio of lateral branch and public manifold. The total pressure loss coefficientK13always increased with the increase of mass flow ratio of junctions 1 and 3. The total pressure loss coefficientK23first increased and then decreased with the increase of mass flow ratio of junctions 2 and 3. These pressure loss coefficients of the exhaust pipe junctions can be used in exhaust flow and turbine inlet boundary conditions analysis. In addition, simulating calculation was conducted to analyze the effect of branch angle on total pressure loss coefficient. According to the calculation results, total pressure loss coefficient was almost the same at low mass flow rate of branch manifold 1 but increased with lateral branch angle at high mass flow rate of branch manifold 1.

Optimization of Tandem Blade Based on Improved Particle Swarm Algorithm

2016 ◽  
Author(s):  
Song Zhaoyun ◽  
Bo Liu ◽  
Mao Xiaochen ◽  
Lu Xiaofeng
Keyword(s):  
Particle Swarm Optimization ◽  
Pressure Loss ◽  
Total Pressure ◽  
Particle Swarm ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Swarm Optimization ◽  
Improved Particle Swarm Optimization ◽  
Pressure Loss Coefficient ◽  
Total Pressure Loss Coefficient

To improve the design quality of high-turning tandem blade, a coupling optimization system for the shape and relative position of tandem blades was developed based on an improved particle swarm optimization algorithm and NURBS parameterization. First of all, to increase convergence speed and avoid local optima of particle swarm optimization (PSO), an improved particle swarm optimization (IPSO) is formulated based on adaptive selection of particle roles, adaptive control of parameters and population diversity control. Then experiments are carried out using test functions to illustrate the performance of IPSO and to compare IPSO with some PSOs. The comparison indicates IPSO can obtain excellent convergence speed and simultaneously keep the best reliability. In addition, the coupling optimization system is validated by optimizing a large-turning tandem blade. Optimization results illustrate IPSO can obviously increase the optimization speed and reduce the time and cost of optimization. After optimization, at design condition, the total pressure loss coefficient of the optimized blade is decreased by 40.4%, and the static pressure ratio of optimized blade is higher and the total pressure loss coefficient is smaller at all incidence angles. In addition, properly reducing the gap area of tandem blade can effectively reduce the friction loss of the blade boundary layer and the mixing loss created by mixing the gap fluid and the mainstream fluid.

Active Separation Control in Circular and Transitioning S-Duct Diffusers Using Vortex Generator Jets

2006 ◽  
Author(s):  
A. M. Pradeep ◽  
R. K. Sullerey
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Vortex Generator ◽  
Loss Coefficient ◽  
Total Pressure Loss ◽  
Separation Control ◽  
Pressure Loss Coefficient ◽  
Distortion Coefficient ◽  
Total Pressure Loss Coefficient

Performance enhancement of three-dimensional S-duct diffusers by separation control using vortex generator jets is the objective of the current experimental investigation. Two different diffuser geometries namely, a circular diffuser and a rectangular–to–circular transitioning diffuser were studied in uniform inflow conditions at a Reynolds number of 7.8 × 105 and the performance evaluation of the diffusers was carried out in terms of static pressure improvement and quality (flow uniformity) of the exit flow. Detailed measurements that included total pressure, velocity distribution, surface static pressure, skin friction and boundary layer measurements were taken and these results are presented here in terms of static pressure rise, distortion coefficient and total pressure loss coefficient at the duct exit. The mass flow rate of the air injected through the VGJ was about 0.06 percent of the main flow for separation control. The distortion coefficient reduced by over 25 percent and the total pressure loss coefficient reduced by about 30 percent in both the diffusers. The physical mechanism of the flow control devices used has been explored using smoke visualization images.

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