Film Cooling Performance Improvement With Optimized Hole Arrangements on Pressure Side Surface of Nozzle Guide Vane: Part II — Experimental Validation

2016 ◽  
Author(s):  
Dong-Ho Rhee ◽  
Young Seok Kang ◽  
Bong Jun Cha ◽  
Jeong-Seek Kang ◽  
Sanga Lee ◽  
...  
Keyword(s):  
Mass Flow Rate ◽  
Performance Improvement ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Side Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Percentage Points ◽  
Nozzle Guide

In the present study, the optimized configurations of film cooled turbine guide vanes proposed in Part I were validated experimentally and the effect of coolant mass flow rate on the performance was examined for those optimized configurations. A set of tests were conducted using an annular sector transonic turbine cascade test facility in Korea Aerospace Research Institute. The mainstream and the secondary air for cooling are supplied by 500 hp and 50 hp compressors, respectively, and the mainstream was heated approximately 20°C above the secondary flow by 300kW heater. To measure the film cooling effectiveness on the pressure side surface, the transient measurement method was used using a FLIR infrared camera system. The test section has five nozzle guide vanes with four passages. The three times scaled-up vane model is manufactured by a stereolithography method. The tests were conducted at mainstream exit Reynolds number based on the chord of 2.2×106 and the coolant mass flow rate ranging from 5 to 13% of the mainstream. The flow periodicity in the cascade passage was verified by surface static pressure measurements. The results showed that the optimized cases present better cooling effectiveness values in the overall region. The effect of coolant mass flow rate also presents the same trend. Comparison with the CFD results shows that the CFD results over-predict film cooling effectiveness by 10∼20 percentage points for baseline and 17∼23 percentage points for the optimized cases. This is probably partly due to the discrepancy of operating conditions such as inlet boundary condition and density ratio and partly due to the limitation of numerical method used in the optimization such as coarse grid near the surface. However, a quite good agreement is obtained qualitatively, which means the optimization process can be utilized as a reliable and efficient method for film cooling performance improvement.

A One-Dimensional Analytical Method for Turbine Blade Preliminary Cooling Design

2016 ◽  
Author(s):  
Chen Li ◽  
Jian-jun Liu
Keyword(s):  
Mass Flow Rate ◽  
Turbine Blade ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Three Dimensional ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
One Dimensional ◽  
Cooling Design

The turbine blade cooling design is a complex procedure including one-dimensional preliminary cooling design, detailed two-dimensional design and fluid network analyses, and three-dimensional conjugate heat transfer and FEM predictions. Frequent alteration and modification of the cooling configurations make it unpractical to obtain all of three-dimensional design results quickly. Preliminary cooling design deals mainly with the coolant requirements and can be knitted into fluid network to look up the expected cooling structural style to promote three-dimensional geometry design. Previous methods to estimate the coolant requirements of the whole turbine blade in the preliminary cooling design were usually based on the semi-empirical air-cooled blade data. This paper combines turbine blade internal and external cooling, and presents a one-dimensional theoretical analytical method to investigate blade cooling performance, assuming that the coolant temperature increases along the blade span. Firstly, a function of non-dimensional cooling mass flow rate is derived to describe the new relationship between adiabatic film cooling effectiveness and overall cooling effectiveness. Secondly, a new variable related to film cooling is found to estimate the required adiabatic film cooling effectiveness without using the empirical correlations. Finally, a theoretical calculation about the relationship between non-dimensional cooling mass flow rate and overall cooling effectiveness well corresponds to semi-empirical air-cooled blade data within regular range of cooling efficiency. The currently proposed method is also a useful tool for the blade thermal analysis and the sensitivity analysis of coolant requirements to various design parameters. It not only can provide all the possible options at the given gas and coolant inlet temperatures to meet the design requirement, but also can give the third boundary conditions for calculating the blade temperature field. It’s convenient to use the heat transfer characteristic of internal cooling structures to estimate the coolant mass flow rate and the channel hydraulic diameter for both convection cooling and film cooling.

An Experimental Investigation of Full-Coverage Film Cooling Effectiveness and Heat Transfer Coefficient of a Turbine Guide Vane in a Linear Transonic Cascade

2016 ◽  
Author(s):  
Zhong-yi Fu ◽  
Hui-ren Zhu ◽  
Cun-liang Liu ◽  
Cong Liu ◽  
Zheng Li
Keyword(s):  
Heat Transfer ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Rate Ratio ◽  
Flow Rate Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mass Flow Rate Ratio

This paper experimentally investigates the film cooling performance of an enlarged turbine guide vane with full-coverage cylindrical hole film cooling in short duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and isentropic exit Mach number independently. The effects of mass flow rate ratio (MFR=4.83%∼8.83%), inlet Reynolds number (Rein= 1.7×105∼5.7×105), and isentropic exit Mach number (Mais=0.81∼1.01) are investigated. There are five rows of cylindrical film cooling holes on the pressure side and four such rows on the suction side respectively. Another four rows of cylindrical holes are provided on the leading edge to obtain a showerhead film cooling. The surface heat transfer coefficient and adiabatic film cooling effectiveness are derived from the surface temperatures measured by the thermocouples mounted in the middle span of the vane surface based on transient heat transfer measurement method. Mass flow rate ratio is shown to have a significant effect on film cooling effectiveness. The increase of mass flow rate ratio increases film cooling effectiveness on pressure side, while increasing this factor has opposite effect on film cooling effectiveness on the suction side. At the same mass flow rate ratio, increasing the Reynolds number can enhance the film cooling performance, the expectation is that at low mass flow rate ratio condition increasing the Reynolds number decreases film cooling effectiveness on the pressure side. The heat transfer coefficient increases with the mass flow rate ratio increasing on both pressure and suction side. At middle and high inlet Reynolds number condition, in the region of 0.4<s<0.6 on suction side, the coolant weakens heat transfer adversely.

Numerical Study of Supersonic Film Cooling in Diverging Section of Nuclear Rocket Laval Nozzle

2018 ◽  
Author(s):  
Xiaokai Sun ◽  
Ping Ye ◽  
Peixue Jiang ◽  
Wei Peng ◽  
Jie Wang
Keyword(s):  
Mach Number ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Laval Nozzle ◽  
Numerical Study ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nuclear Rocket

Nuclear rockets with specific impulse have obvious advantages by greatly reducing the mass of the propellant and potentially decreasing the cost of launching material from the earth’s surface. Nuclear thermal rockets use hydrogen propellant with coolant exit temperature of near 3000 K, which is very high, so the cooling of airframe surfaces in the vicinity of the exhaust is needed, of which film cooling is an effective method. Most of previous studies mainly focus on the film cooling effectiveness using two dimensional backward-facing step model, however, the nuclear rocket exhaust using the converging-diverging Laval nozzle, so the film cooling would be different. The present study numerically investigated the influence of coolant Mach number, coolant inlet height on supersonic film cooling in the diverging section of Laval nozzle, while keeping the coolant mass flow rate constant, with the results showing that: increasing the coolant inlet Mach number and the coolant inlet height can increase the film cooling effectiveness; for the same coolant mass flow rate, reducing the coolant inlet height and increasing the inlet Mach number improves film cooling effectiveness.

Influence of Coolant Jet Pulsation on the Convective Film Cooling of an Adiabatic Wall

2016 ◽  
Vol 139 (2) ◽  
Author(s):  
Qaiser Sultan ◽  
Gildas Lalizel ◽  
Matthieu Fénot ◽  
Eva Dorignac
Keyword(s):  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Convective Heat Transfer Coefficient ◽  
Time Interval ◽  
Adiabatic Wall ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
The Impact

This study investigates the effects of sinusoidal pulsations externally imposed to an oblique round jet. The effectiveness of film coverage of an adiabatic wall onset for a thermally uniform bulk flow is presented in the perspective of gas turbine film cooling. For the injectant fluid, both the temperature and the mass flow rate are controlled prior to entrance to the periodic forcing system using a loudspeaker drive. The characteristic film cooling parameters including the blowing ratios and the temperature ratio are maintained at M=ρiUi/ρ∞U∞ = 0.65, 1, and 1.25, and Ti/T∞=2 respectively. The injection fluid is pulsated to a nondimensionalized frequency of St=f⋅d/U = 0, 0.2, 0.3, and 0.5. In the present investigation, the impact of injectant film modulation is figured out by analyzing the velocity fields measured by a system of time-resolved particle image velocimetry (TR-PIV), as well as analyzing the adiabatic wall temperature and the convective heat transfer coefficient measured by a system of infrared thermography. The overall film-cooling effectiveness is revealed by the time-averaged analysis, in which altered time-averaged jet trajectories and wake behavior are focused. It is observed that the pulsations tend to result in lower effectiveness when the flow remained attached to the wall in steady blowing case. In steady blowing cases with jet liftoff, such as for M= 1.25, rendering low-frequency pulsation helps in increasing film-cooling effectiveness due to the discharge of lower mass flow rate coolant during the significant time interval of the respective pulse cycle.

Overall Cooling Effectiveness Measurements on Pressure Side Surface of the Nozzle Guide Vane With Optimized Film Cooling Hole Arrangements

2017 ◽  
Author(s):  
Dong-Ho Rhee ◽  
Young Seok Kang ◽  
Bong Jun Cha ◽  
Sanga Lee
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Side Surface ◽  
Internal Cooling ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Rib Turbulators ◽  
Nozzle Guide

Most of the optimization researches on film cooling have dealt with adiabatic film cooling effectiveness on the surface. However, the information on the overall cooling effectiveness is required to estimate exact performance of the optimization configuration since hot components such as nozzle guide vane have not only film cooling but also internal cooling features such as rib turbulators, jet impingement and pin-fins on the inner surface. Our previous studies [1,2] conducted the hole arrangement optimization to improve adiabatic film cooling effectiveness values and uniformity on the pressure side surface of the nozzle guide vane. In this study, the overall cooling effectiveness values were obtained at various cooling mass flow rates experimentally for the baseline and the optimized hole arrangements proposed by the previous study [1] and compared with the adiabatic film cooling effectiveness results. The tests were conducted at mainstream exit Reynolds number based on the chord of 2.2 × 106 and the coolant mass flow rate from 5 to 10% of the mainstream. For the experimental measurements, a set of tests were conducted using an annular sector transonic turbine cascade test facility in Korea Aerospace Research Institute. To obtain the overall cooling effectiveness values on the pressure side surface, the additive manufactured nozzle guide vane made of polymer material and Inconel 718 were installed and the surface temperature was measured using a FLIR infrared camera system. Since the optimization was based on the adiabatic film cooling effectiveness, the regions with rib turbulators and film cooling holes show locally higher overall cooling effectiveness due to internal convection and conduction, which can cause non-uniform temperature distributions. Therefore, the optimization of film cooling configuration should consider the effect of the internal cooling to avoid undesirable non-uniform cooling.

Background-Grid Based Mapping Approach to Film Cooling Meshing: Part III — Applications in a Full-Cooled Turbine Vane

2020 ◽  
Author(s):  
Ruiqin Wang ◽  
Xin Yan
Keyword(s):  
Gas Turbine ◽  
Turbine Blade ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Design Flow ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Structured Grids

Abstract To cool a high-pressure gas turbine blade, many rows of cooling holes with different arrangements and configurations are manufactured to achieve higher cooling effect and lower aerodynamic loss. To evaluate the heat transfer and film cooling effect in the full-cooled turbine blade, efficient numerical simulations are required in the design and performance optimization processes. From the view of numerical accuracy, the structured grids have to be employed because of higher resolution in flow and heat transfer than the unstructured grids. Because many splitting, attaching and merging manipulations are involved in meshing the cooling features and curved boundaries, it is very complex and time-consuming for a researcher to generate multi-block structured grids for a full-cooled gas turbine blade. As a result, in the industrial applications, almost all researchers preferred to generate unstructured grids instead of structured grids for the full-cooled blade. Unlike the previous research, the aim of this study is to apply the Background-Grid Based Mapping (BGBM) method proposed in Part I to generate multi-block structured grids for a full-cooled gas turbine vane. With the strategy of BGBM method, meshes were conveniently generated in the computational space with simple geometrical features and plain interfaces, and then were mapped back into physical space to obtain the multi-block structured grids which can be used for numerical simulations. With the experimental data, the present numerical methods and BGBM strategy were carefully validated. Then, the flow and film cooling performance in the full-cooled NASA GE-E3 nozzle guided vane were numerically investigated. The effects of coolant mass flow rate and land extensions on film cooling effectiveness were discussed. The results show that film cooling effectiveness near the stagnation point is the lowest and film cooling effectiveness on the pressure side is slightly higher than that on the suction side. When the coolant mass flow rate increases up to the value of 1.5 design flow, the relative outflow mass flow rates of cooling hole arrays and slots are no longer affected by the increase of the coolant flow rate. At half design flow, the outflow mass flow rates of No.5 hole-array to No.10 hole-array are almost zero, and the area-averaged film cooling effectiveness on vane surface is as low as 0.268. Compared with the cases of half design flow and double design flow, better film cooling performance is obtained in the cases of design flow and 1.5 design flow. Compared with the vane without lands, the area-average cooling effectiveness on vane surface is slightly higher for the vane with lands. Land extensions have a considerable influence on film cooling performance in the cutback region.

Multi-Objective CFD Optimisation of Shaped Hole Film Cooling With Mesh Morphing

2015 ◽  
Author(s):  
D. Proietti ◽  
A. Pranzitelli ◽  
G. E. Andrews ◽  
M. E. Biancolini ◽  
D. B. Ingham ◽  
...  
Keyword(s):  
Design Of Experiments ◽  
Mass Flow Rate ◽  
Mass Flow ◽  
Flow Rate ◽  
Film Cooling ◽  
Maximum Deviation ◽  
Geometrical Parameters ◽  
Shape Parameters ◽  
Cooling Effectiveness ◽  
Adiabatic Cooling

A Computational Fluid Dynamics (CFD) optimisation of a single row of film cooling holes was performed. The aim was to achieve the highest adiabatic cooling effectiveness while minimising the coolant mass flow rate. The geometry investigated by Gritsch et al. [1] was the baseline model. It consisted of a row of cylindrical, 30° inclined holes, with a mainstream inlet Mach number of 0.6, a blowing ratio of 1 and a plenum for the upstream cooling air flow. The predictions agreed with the experimental data with a maximum deviation of 6%. The geometry was then optimised by varying three shape parameters: the injection angle, the lateral hole expansion angle and the downstream compound hole angle. A goal driven optimisation approach was based on a design of experiments table. The minimisation of the coolant mass flow together with the maximisation of the minimum and average cooling effectiveness were the optimisation objectives. The shape modifications were performed directly in the ANSYS Fluent CFD solver by using the software RBF Morph in the commercial software platform ANSYS Workbench. There was no need to generate a new geometry and a new computational mesh for each configuration investigated. The dependency of the average effectiveness along the plane centreline on the three geometrical parameters was investigated based on the metamodel generated from the design of experiments results. The goal driven optimisation led to the optimal combination of the three shape parameters to minimise the coolant flow without reducing the cooling effectiveness. The best results were obtained for a geometry with 20° hole angle and 7.5° compound angle injection, leading to a reduction of 15% in the coolant mass flow rate for an enhanced adiabatic cooling effectiveness. The results also showed the preponderance of the centreline angle over the other two parameters.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

Leading Edge Film Cooling of a Turbine Blade Model Through Single and Double Row Injection: Effects of Coolant Density

1994 ◽  
Author(s):  
M. Salcudean ◽  
I. Gartshore ◽  
K. Zhang ◽  
Y. Barnea
Keyword(s):  
Turbine Blade ◽  
Mass Flow ◽  
Film Cooling ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Double Row ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Staggered Arrangement

Experiments have been conducted on a large model of a turbine blade. Attention has been focussed on the leading edge region, which has a semi-circular shape and four rows of film cooling holes positioned symmetrically about the stagnation line. The cooling holes were oriented in a spanwise direction with an inclination of 30° to the surface, and had streamwise locations of ±15° and ±44° from the stagnation line. Film cooling effectiveness was measured using a heat/mass analogy. Single row cooling from the holes at 15° and 44° showed similar patterns: spanwise averaged effectiveness which rises from zero at zero coolant mass flow to a maximum value η* at some value of mass flow ratio M*, then drops to low values of η at higher M. The trends can be quantitatively explained from simple momentum considerations for either air or CO2 as the coolant gas. Close to the holes, air provides higher η values for small M. At higher M, particularly farther downstream, the CO2 may be superior. The use of an appropriately defined momentum ratio G collapses the data from both holes using either CO2 or air as coolant onto a single curve. For η*, the value of G for all data is about 0.1. Double row cooling with air as coolant shows that the relative stagger of the two rows is an important parameter. Holes in line with each other in successive rows can provide improvements in spanwise averaged film cooling effectiveness of as much as 100% over the common staggered arrangement. This improvement is due to the interaction between coolant from rows one and two, which tends to provide complete coverage of the downstream surface when the rows are placed correctly with respect to each other.

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