A Novel Multi-Stage Impingement Cooling Scheme—Part II: Design Optimization

2020 ◽  
Vol 142 (12) ◽  
Author(s):  
Kexin Liu ◽  
Qiang Zhang
Keyword(s):  
Design Optimization ◽  
Gas Turbine ◽  
Cross Flow ◽  
Target Surface ◽  
Optimization Strategy ◽  
Working Mechanism ◽  
Impingement Cooling ◽  
Multi Stage ◽  
Cooling Air ◽  
Combustor Liner

Abstract Cross flow and coolant maldistribution are the common design challenges for impingement cooling in modern gas turbine. This paper reports a novel multi-stage impingement cooling scheme for combustor liner. The design concept and general working mechanism are introduced in the Part I paper. This Part II paper presents the design flexibilities and optimization strategies. Conjugate heat transfer (CHT) analysis was conducted at a range of Reynolds numbers to assess the thermal performance, loss penalty, and the working mechanism behind. The results show that varying the jet hole diameter in each cooling stage can be an effective design optimization strategy in balancing the cooling requirement and loss penalty. Inter-stage bypass design is also another design flexibility offered by the multi-stage scheme to regulate the cooling air consumption at different stages. With these optimization strategies, the target surface temperature and local gradient can be effectively reduced with reasonable pressure loss with 50% reduction in the cooling air consumption compared to conventional single-stage impingement design. This multi-stage impingement concept can be practically applied to gas turbine combustor liner and turbine blade cooling.

A Novel Multi-Stage Impingement Cooling Scheme—Part I: Concept Study

2020 ◽  
Vol 142 (12) ◽  
Author(s):  
Kexin Liu ◽  
Qiang Zhang
Keyword(s):  
Cross Flow ◽  
Target Surface ◽  
Cooling Capacity ◽  
Operating Conditions ◽  
Single Stage ◽  
Impingement Cooling ◽  
Flow Effect ◽  
Multi Stage ◽  
Cooling Design ◽  
The Cross

Abstract The impingement cooling for a modern gas turbine component, either a combustor liner or a high-pressure turbine blade, is often not as efficient as required due to strong cross-flow effect and coolant maldistribution. This paper reports a novel multi-stage impingement cooling scheme to effectively use the coolant and minimize the cross-flow effect. The design concept and general working mechanism are introduced in this Part I paper. The extra design flexibilities and optimization strategies are reported in Part II. Numerical simulations on conjugate heat transfer (CHT) were carried out to assess the flow structure and thermal performance between a typical single-stage cooling design and a three-stage cooling design at typical operating conditions. It has been observed that the novel multi-stage cooling design can reinitiate impingement jets at each stage, which greatly reduces the cross-flow impact and local thermal gradient. The staging of cooling air for the target surface also offers better utilization of the cooling capacity. Even by using 50% of the coolant designed for the single-stage impingement cooling, the multi-stage case can still sufficiently cool the target surface. The additional pressure loss penalty introduced in multi-stage design needs further efforts on design optimization.

Continuous adjoint aerodynamic optimization design for multi-stage gas turbine with cooling air

2017 ◽  
Vol 89 (3) ◽  
pp. 444-456
Author(s):  
Lei Chen ◽  
Jiang Chen
Keyword(s):  
Gas Turbine ◽  
Optimization Design ◽  
Adjoint Method ◽  
Air Injection ◽  
Aerodynamic Optimization ◽  
Content Type ◽  
Multi Stage ◽  
Surface Code ◽  
Cooling Air ◽  
Continuous Adjoint

Purpose This paper aims to conduct the optimization of the multi-stage gas turbine with the effect of the cooling air injection based on the adjoint method. Design/methodology/approach Continuous adjoint method is combined with the S2 surface code. Findings The optimization of the stagger angles, stacking lines and the passage can improve the attack angles and restrain the development of the boundary, reducing the secondary flow loss caused by the cooling air injection. Practical implications The aerodynamic performance of the gas turbine can be improved via the optimization of blade and passage based on the adjoint method. Originality/value The results of the first study on the adjoint method applied to the S2 surface through flow calculation including the cooling air effect are presented.

Pulsed Impingement Turbine Cooling and its Effect on the Efficiency of Gas Turbines With Pressure Gain Combustion

2020 ◽  
Author(s):  
Nicolai Neumann ◽  
Dieter Peitsch ◽  
Arne Berthold ◽  
Frank Haucke ◽  
Panagiotis Stathopoulos
Keyword(s):  
Gas Turbine ◽  
Thermal Efficiency ◽  
Gas Turbines ◽  
Turbine Blades ◽  
Impingement Cooling ◽  
Turbine Cooling ◽  
Pressure Gain Combustion ◽  
Cooling Air ◽  
Gas Turbine Cycle ◽  
Cycle Efficiency

Abstract Performance improvements of conventional gas turbines are becoming increasingly difficult and costly to achieve. Pressure Gain Combustion (PGC) has emerged as a promising technology in this respect, due to the higher thermal efficiency of the respective ideal gas turbine cycle. Previous cycle analyses considering turbine cooling methods have shown that the application of pressure gain combustion may require more turbine cooling air. This has a direct impact on the cycle efficiency and reduces the possible efficiency gain that can potentially be harvested from the new combustion technology. Novel cooling techniques could unlock an existing potential for a further increase in efficiency. Such a novel turbine cooling approach is the application of pulsed impingement jets inside the turbine blades. In the first part of this paper, results of pulsed impingement cooling experiments on a curved plate are presented. The potential of this novel cooling approach to increase the convective heat transfer in the inner side of turbine blades is quantified. The second part of this paper presents a gas turbine cycle analysis where the improved cooling approach is incorporated in the cooling air calculation. The effect of pulsed impingement cooling on the overall cycle efficiency is shown for both Joule and PGC cycles. In contrast to the authors’ anticipation, the results suggest that for relevant thermodynamic cycles pulsed impingement cooling increases the thermal efficiency of Joule cycles more significantly than it does in the case of PGC cycles. Thermal efficiency improvements of 1.0 p.p. for pure convective cooling and 0.5 p.p. for combined convective and film with TBC are observed for Joule cycles. But just up to 0.5 p.p. for pure convective cooling and 0.3 p.p. for combined convective and film cooling with TBC are recorded for PGC cycles.

Computational analysis of span-wise hole locations on fluid flow and film cooling of internal channels with crescent ribs

2019 ◽  
Vol 29 (8) ◽  
pp. 2728-2753
Author(s):  
Guohua Zhang ◽  
Xueting Liu ◽  
Bengt Ake Sundén ◽  
Gongnan Xie
Keyword(s):  
Film Cooling ◽  
Cross Flow ◽  
Side Wall ◽  
Target Surface ◽  
Internal Cooling ◽  
Cooling Performance ◽  
Content Type ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cooling Air

Purpose This study aims to clarify the mechanism of film hole location at the span-wise direction of an internal cooling channel with crescent ribs on the adiabatic film cooling performance, three configurations are designed to observe the effects of the distance between the center of the ellipse and the side wall(Case 1, l = w/2, Case 2, l = w/3 and for Case 3, l = w/4). Design/methodology/approach Numerical simulations are conducted under two blowing ratios (i.e. 0.5 and 1) and a fixed cross-flow Reynolds number (Rec = 100,000) with a verified turbulence model. Findings It is shown that at low blowing ratio, reducing the distance increases the film cooling effectiveness but keeps the trend of the effectiveness unchanged, while at high blowing ratio, the characteristic is a little bit different in the range of 0 = x/D = 10. Research limitations/implications These features could be explained by the fact that shrinking the distance between the hole and side wall induces a much smaller reserved region and vortex downstream the ribs and a lower resistance for cooling air entering the film hole. Furthermore, the spiral flow inside the hole is impaired. Originality/value As a result, the kidney-shaped vortices originating from the jet flow are weakened, and the target surface can be well covered, resulting in an enhancement of the adiabatic film cooling performance.

Novel Gas Turbine Blade Leading Edge Cooling Configuration Using Advanced Double Swirl Chambers

2015 ◽  
Author(s):  
Karsten Kusterer ◽  
Gang Lin ◽  
Takao Sugimoto ◽  
Dieter Bohn ◽  
Ryozo Tanaka ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Leading Edge ◽  
Gas Turbine Blade ◽  
Stagnation Line ◽  
Impingement Cooling ◽  
Cooling Technology ◽  
Cooling Air ◽  
Blade Leading Edge

The Double Swirl Chambers (DSC) cooling technology, which has been introduced and developed by the authors, has the potential to be a promising cooling technology for further increase of gas turbine inlet temperature and thus improvement of the thermal efficiency. The DSC cooling technology establishes a significant enhancement of the local internal heat transfer due to the generation of two anti-rotating swirls. The reattachment of the swirl flows with the maximum velocity at the center of the chamber leads to a linear impingement effect on the internal surface of the blade leading edge nearby the stagnation line of gas turbine blade. Due to the existence of two swirls both the suction side and the pressure side of the blade near the leading edge can be very well cooled. In this work, several advanced DSC cooling configurations with a row of cooling air inlet holes have been investigated. Compared with the standard DSC cooling configuration the advanced ones have more suitable cross section profiles, which enables better accordance with the real blade leading edge profile. At the same time these configurations are also easier to be manufactured in a real blade. These new cooling configurations have been numerically compared with the state of the art leading edge impingement cooling configuration. With the same configuration of cooling air supply and boundary conditions the advanced DSC cooling presents 22–26% improvement of overall heat transfer and 3–4% lower total pressure drop. Along the stagnation line the new cooling configuration can generate twice the heat flux than the standard impingement cooling channel. The influence of spent flow in the impinging position and impingement heat transfer value is in the new cooling configurations much smaller, which leads to a much more uniform heat transfer distribution along the chamber axial direction.

Experimental and Numerical Study of Jet Impingement Cooling for Improved Gas Turbine Blade Internal Cooling With In-Line and Staggered Nozzle Arrays

2020 ◽  
Vol 143 (1) ◽  
Author(s):  
Abdel Rahman Salem ◽  
Farah Nazifa Nourin ◽  
Mohammed Abousabae ◽  
Ryoichi S. Amano
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Gas Turbine ◽  
Numerical Study ◽  
Turbine Blades ◽  
Jet Impingement ◽  
Target Surface ◽  
Internal Cooling ◽  
Impingement Cooling ◽  
Gas Turbine Blades

Abstract Internal cooling of gas turbine blades is performed with the combination of impingement cooling and serpentine channels. Besides gas turbine blades, the other turbine components such as turbine guide vanes, rotor disks, and combustor wall can be cooled using jet impingement cooling. This study is focused on jet impingement cooling, in order to optimize the coolant flow, and provide the maximum amount of cooling using the minimum amount of coolant. The study compares between different nozzle configurations (in-line and staggered), two different Reynold's numbers (1500 and 2000), and different stand-off distances (Z/D) both experimentally and numerically. The Z/D considered are 3, 5, and 8. In jet impingement cooling, the jet of fluid strikes perpendicular to the target surface to be cooled with high velocity to dissipate the heat. The target surface is heated up by a direct current (DC) power source. The experimental results are obtained by means of thermal image processing of the captured infra-red (IR) thermal images of the target surface. Computational fluid dynamics (CFD) analysis were employed to predict the complex heat transfer and flow phenomena, primarily the line-averaged and area-averaged Nusselt number and the cross-flow effects. In the current investigation, the flow is confined along with the nozzle plate and two parallel surfaces forming a bi-directional channel (bi-directional exit). The results show a comparison between heat transfer enhancement with in-line and staggered nozzle arrays. It is observed that the peaks of the line-averaged Nusselt number (Nu) become less as the stand-off distance (Z/D) increases. It is also observed that the fluctuations in the stagnation heat transfer are caused by the impingement of the primary vortices originating from the jet nozzle exit.

Experimental and Numerical Investigation of Impingement on a Rib-Roughened Leading-Edge Wall

2003 ◽  
Vol 125 (4) ◽  
pp. 682-691 ◽  
Author(s):  
M. E. Taslim ◽  
K. Bakhtari ◽  
H. Liu
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Cross Flow ◽  
Heat Removal ◽  
Leading Edge ◽  
Target Surface ◽  
Transfer Coefficients ◽  
Heat Transfer Coefficients ◽  
Edge Surface ◽  
Rib Roughened

Effective cooling of the airfoil leading edge is imperative in gas turbine designs. Among several methods of cooling the leading edge, impingement cooling has been utilized in many modern designs. In this method, the cooling air enters the leading edge cavity from the adjacent cavity through a series of crossover holes on the partition wall between the two cavities. The crossover jets impinge on a smooth leading-edge wall and exit through the film holes, and, in some cases, form a cross flow in the leading-edge cavity and move toward the end of the cavity. It was the main objective of this investigation to measure the heat transfer coefficient on a smooth as well as rib-roughened leading-edge wall. Experimental data for impingement on a leading-edge surface roughened with different conical bumps and radial ribs have been reported by the same authors previously. This investigation, however, deals with impingement on different horseshoe ribs and makes a comparison between the experimental and numerical results. Three geometries representing the leading-edge cooling cavity of a modern gas turbine airfoil with crossover jets impinging on (1) a smooth wall, (2) a wall roughened with horseshoe ribs, and (3) a wall roughened with notched-horseshoe ribs were investigated. The tests were run for a range of flow arrangements and jet Reynolds numbers. The major conclusions of this study were: (a) Impingement on the smooth target surface produced the highest overall heat transfer coefficients followed by the notched-horseshoe and horseshoe geometries. (b) There is, however, a heat transfer enhancement benefit in roughening the target surface. Among the three target surface geometries, the notched-horseshoe ribs produced the highest heat removal from the target surface, which was attributed entirely to the area increase of the target surface. (c) CFD could be considered as a viable tool for the prediction of impingement heat transfer coefficients on an airfoil leading-edge wall.

Experimental and Numerical Study of Heat Transfer in a Gas Turbine Combustor Liner

2003 ◽  
Vol 125 (4) ◽  
pp. 994-1002 ◽  
Author(s):  
J. C. Bailey ◽  
J. Intile ◽  
T. F. Fric ◽  
A. K. Tolpadi ◽  
N. V. Nirmalan ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Numerical Study ◽  
Cross Flow ◽  
Turbulence Models ◽  
Flow Distribution ◽  
Transfer Coefficients ◽  
Gas Turbine Combustor ◽  
Heat Transfer Coefficients ◽  
Combustor Liner

Experiments and numerical simulations were conducted to understand the heat transfer characteristics of a stationary gas turbine combustor liner cooled by impingement jets and cross flow between the liner and sleeve. Heat transfer was also aided by trip-strip turbulators on the outside of the liner and in the flowsleeve downstream of the jets. The study was aimed at enhancing heat transfer and prolonging the life of the combustor liner components. The combustor liner and flow sleeve were simulated using a flat-plate rig. The geometry has been scaled from actual combustion geometry except for the curvature. The jet Reynolds number and the mass-velocity ratios between the jet and cross flow in the rig were matched with the corresponding combustor conditions. A steady-state liquid crystal technique was used to measure spatially resolved heat transfer coefficients for the geometric and flow conditions mentioned above. The heat transfer was measured both in the impingement region as well as over the turbulators. A numerical model of the combustor test rig was created that included the impingement holes and the turbulators. Using CFD, the flow distribution within the flow sleeve and the heat transfer coefficients on the liner were both predicted. Calculations were made by varying the turbulence models, numerical schemes, and the geometrical mesh. The results obtained were compared to the experimental data and recommendations have been made with regard to the best modeling approach for such liner-flow sleeve configurations.

Effects of Extended Jet Holes to Heat Transfer and Flow Characteristics of the Jet Impingement Cooling

2019 ◽  
Vol 141 (8) ◽  
Author(s):  
Ahmet Ümit Tepe ◽  
Kamil Arslan ◽  
Yaşar Yetişken ◽  
Ünal Uysal
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Average Nusselt Number ◽  
Flat Surface ◽  
Flow Characteristics ◽  
Cross Flow ◽  
Jet Impingement ◽  
Target Surface ◽  
Impingement Cooling ◽  
Compressor Power

In this study, effects of extended jet holes to heat transfer and flow characteristics of jet impingement cooling were numerically investigated. Cross-flow in the impinging jet cooling adversely affects the heat transfer on the target surface. The main purpose of this study is to reduce the negative effect of cross-flow on heat transfer by extending jet holes toward the target surface with nozzles. This study has been conducted under turbulent flow condition (15,000 ≤ Re  ≤  45,000). The surface of the turbine blade, which is the target surface, has been modeled as a flat plate. The effect of the ribs, placed on the target surface, on the heat transfer has been also investigated, and the results were compared with the flat surface. The parameters such as average and local Nusselt numbers on the target surface, flow characteristics, and compressor power have been examined in detail. It was obtained from the numerical results that the average Nusselt number increases with decreasing the gap between the target surface and the nozzle. In addition, the higher average Nusselt number was obtained on the flat surface than the ribbed surface. The lowest compressor power was achieved in the 5Dj nozzle gap for the flat surface and in the 4Dj nozzle gap for the ribbed surface.

Enhancement of Impingement Cooling in a High Cross Flow Channel Using Shaped Impingement Cooling Holes

2009 ◽  
Vol 132 (2) ◽  
Author(s):  
Andrew C. Chambers ◽  
David R. H. Gillespie ◽  
Peter T. Ireland ◽  
Robert Kingston
Keyword(s):  
Heat Transfer ◽  
Nusselt Number ◽  
Cross Flow ◽  
Target Surface ◽  
Impingement Cooling ◽  
Turbine Cooling ◽  
Impingement Jet ◽  
Coefficient Distribution ◽  
Cooling Passage ◽  
Experimental Rig

Impingement systems are common place in many turbine cooling applications. Generally these systems consist of a target plate that is cooled by the impingement of multiple orthogonal jets. While it is possible to achieve high target surface heat transfer with this configuration, the associated pressure drop is generally high and the cooling efficiency low. Furthermore, especially in large impingement arrays, the buildup of cross flow from upstream jets can be significant and results in deflection of downstream impingement jets reducing the resultant heat transfer coefficient distribution. This paper presents a computational and experimental investigation into the use of shaped elliptical or elongated circular impingement holes designed to improve the penetration of the impinging jet across the coolant passage. This is of particular interest where there is significant cross flow. Literature review and computational investigations are used to determine the optimum aspect ratio of the impingement jet. The improved heat transfer performance of the modified design is then tested in an experimental rig with varying degrees of cross flow at engine representative conditions. In all cases, a 16% increase in the Nusselt number on the impingement target surface in the downstream half of the cooling passage was achieved. Under the first four impingement holes, a Nusselt number enhancement of 28–77% was achieved, provided no additional cross flow was present in the passage. When appropriately aligned, a significant reduction in the stress concentration factor caused by the addition of a hole can be achieved using this design.

Sign in / Sign up

Export Citation Format

Share Document