Multi-channel spiral cooling for the first stage moving blade of an industrial gas turbine

Abstract

The internal cooling system for the moving blade of a gas turbine is such that the technology used in each part is proportional to the thermal load applied by the hot gases to that part. Thus, in areas that are less sensitive, a smaller amount of cooling air is used, and in critical areas that are more sensitive, cooling with a higher flux is applied. By increasing the efficiency and effectiveness of the cooling system, the amount of coolant flow consumed is reduced, which increases the efficiency of the turbine. Also, the maximum blade temperature is reduced and a more uniform distribution of blade temperature is achieved, which increases blade life at high temperatures. In previous cooling methods, blade cooling is very non-uniform, causing hot spots in the blade and reducing its life. Since the leading edge is exposed to the highest thermal load from the hot gases and is highly sensitive, a separate spiral channel containing V-shaped fins is used to cool this area in order to increase the heat transfer coefficient. The air passing through this spiral channel cools the upper part of the blade and exits from the upper part of the trailing edge. In addition to this spiral channel, the blade cooling system also includes two other spiral channels that are responsible for cooling the middle and trailing edge areas of the blade. In these two spiral channels, the cross-sectional area is large for cooling the middle area of ​​the blade, and rib technology with low density and wide surfaces is used to prevent high pressure drops in these areas. The flow of these two spiral channels also ultimately provides cooling of the trailing edge and is discharged into the main flow through the trailing edge end duct. Trailing edge cooling is also usually of high importance because the possibility of thermal damage is high due to the thinness of the blade in this area. Therefore, a combination of fins and a relatively high density pin-fin array is used to cool the trailing edge. The thinness of the blade at the trailing edge end makes the use of fins difficult. But pin-fins significantly increase the heat transfer coefficient, and the thinness of the fins is not an obstacle to their manufacture.

Description

Description of the invention Title of the invention (as stated in the declaration) Multi-channel spiral cooling for the first stage moving blade of an industrial gas turbine Technical background of the relevant invention Section F: Mechanical Engineering, Subsection of Engine Types Technical problem and stating the objectives of the invention In gas turbines, the most important factor in increasing turbine efficiency is raising the temperature of the exhaust gases from the combustion chamber. At the same time, raising the temperature of the combustion products (which enter the turbine after leaving the chamber) causes damage to turbine parts and greatly reduces the life of these parts. Therefore, to solve this problem, in addition to using superalloys with high thermal resistance in the manufacture of hot turbine parts, it is necessary to use a cooling system for these parts. Therefore, internal cooling of turbine blades has been one of the biggest challenges for gas turbine manufacturers in recent years. Despite many efforts, internal cooling of blades still suffers from some fundamental weaknesses and there is a large scope for further optimizing the cooling system in turbine blades. In turbine blade cooling, in addition to reducing the maximum blade temperature, creating a uniform temperature distribution and preventing the creation of a strong temperature gradient are of great importance. One of the problems with current fin cooling methods is the lack of uniform fin cooling.This non-uniformity in cooling causes hot spots to appear on the surface and inside the blade, which ultimately leads to the possibility of blade cooling failure at these points. Conventional cooling systems use a higher amount of coolant flow rate for each blade to solve this problem. However, this cooling method is not optimal, because by increasing the coolant flow rate, although more effective cooling can be achieved, the turbine efficiency decreases. The present invention provides an optimal cooling system for the inside of the blades that, while maintaining high efficiency in cooling different parts, can provide as uniform a temperature distribution as possible throughout the blade and yet use a lower cooling flow rate. In fact, the use of this invention increases the efficiency of the turbine and increases the life of the turbine blades simultaneously. A description of the state of the prior art and the history of developments related to the claimed invention. The efficiency of gas turbines can be significantly increased by increasing the temperature of the combustion gases. In this regard, manufacturers have succeeded in increasing the inlet temperature from the combustion chamber to the turbine as much as possible, with the help of advances in material manufacturing and the use of cooling technologies. However, increasing the temperature of the hot gas leads to damage to turbine components and significantly reduces the life of these components. Therefore, to avoid this problem, turbine blades require a more efficient and more efficient internal cooling system. Optimal blade cooling means that the maximum blade temperature is reduced without increasing cooling air consumption and that the temperature distribution in the blade is more uniform. A common example of cooling technology, as shown in Figure 1, consists of a number of nearly parallel channels inside the blade. These channels transport cooling air from the root of the blade to its tip and vice versa. In such common systems, the blade has a highly non-uniform temperature distribution. The reason for this phenomenon is that the cooling technology is the same throughout the blade despite significant changes in external thermal load, as well as the sensitivity of different areas around the blade airfoil. For example, the leading edge of the blade, due to the presence of a dead point on it, tolerates a much higher thermal load from the gas than other points. Also, the trailing edge of the blade, although it may have a lower thermal load, is usually subject to significant damage due to the thinner blade thickness in this area and is highly sensitive in determining the blade life. In addition, the blade is exposed to hot gas with varying temperatures from the lower radius to the upper radius, which is caused by changes in the temperature of the exhaust gas from the combustion chamber.Typically, the gas temperature is lower in areas near the blade tip and its inner radius, and the maximum hot gas temperature occurs at or near the middle radius. The combination of temperature and stress levels in the blade sections determine the life. The stress that the blade endures at each section depends on the radius of the section. Thus, the average stress is reduced from the hub surface to the blade tip. Considering the above, in order to achieve the design of an optimal cooling system, it is necessary that the technology used in each section of the blade is proportional to the thermal load and the amount of stress applied to that section. Thus, in areas that have lower sensitivity, a smaller amount of cooling air cooling capacity is spent, and also in critical areas that have higher sensitivity, cooling with a higher flux is applied. Thus, the blade has higher cooling efficiency in its sensitive points. By increasing heat transfer in sensitive points, a more uniform distribution of blade temperature can be achieved. Until now, most cooling systems have not made much difference between cooling different parts of the blade, and therefore, to prevent thermal damage, it is necessary to use a significant amount of coolant to cool the blade.This way, the critical points receive sufficient flow for cooling, but at the same time, other parts of the blade are cooled with a flow that is greater than necessary. This increases the flow rate of the cooling fluid and, as a result, reduces the efficiency of the turbine. The present invention provides a design for an internal cooling system for turbine blades, in which the arrangement and type of technologies used in each part of the blade are proportional to the thermal sensitivity of that part. Thus, with more optimized cooling, the amount of coolant flow rate is reduced and the turbine efficiency is increased. On the other hand, with lower maximum temperatures and more uniform temperature distribution for the blade, the life of the blades is increased. Providing a solution to an existing technical problem along with an accurate, sufficient, and integrated description of the invention The present invention provides a new design for the spiral cooling system and the arrangement of technologies within the ducts to achieve higher cooling efficiency, more uniform temperature distribution within the blade and increase its life. The main goal of this innovation is to increase the cooling efficiency for the first row of moving blades of turbines with a temperature range of up to 1250 degrees Celsius. The main idea of ​​the present invention is the compatibility between the cooling technology with the thermal load on the one hand and the centrifugal stress level on the other. In other words, in any area where a high heat flux is applied to the blade from the hot gas, the cooling intensity from the internal cooling air in that area is higher. In order for the cooling intensity at a specific location of the blade to be higher than at other locations, it is not necessary to increase the coolant flow passing through that point, but in many cases, changing the geometry of the flow passage (cooling technology) at that location can meet this goal. By changing the geometry (cooling technology), the heat transfer coefficient can be increased, resulting in an increase in heat transfer for an equal rate of coolant flow. Figure (2) shows a simplified view of the external geometry of a conventional turbine blade. The details of the cooling system provided by the present invention are shown in Figures (3) to (7) to illustrate the internal structure of the cooling system as well as the cooling technologies used therein. The movable blade shown in Figure (2) consists of a root section 1 with a support surface (hub) 2. The root section 1 is fitted to the rotating disk of the turbine (not shown in the figure) for the movable blade to be mounted. The blade has an outer aerodynamic surface, more or less like that seen in conventional gas turbine blades. The airfoil section 3 extends from the root 1 to the tip 4, which includes a pressure surface 5 and a suction surface 6, which are connected at the leading edge 7 and trailing edge 8. The hot turbine gas flow passes through the space between the support surface 2 and the tip 4 and over the blade. There are numerous channels inside the blade for the passage of cooling air, which will be described below. For greater clarity, the cooling air flow path is shown separately in Figures (3) to (5). Figure (3) is a schematic view of the middle cross-section of the blade, showing the cross-sectional area of ​​the ducts and the thickness of the blade metal. In this figure, duct 9 corresponds to the first spiral duct, which is responsible for cooling the leading edge. Ducts 10 and 11 correspond to the second and third spiral ducts, respectively, which cool the middle region of the blade. Finally, the cooling flow is discharged outside the blade through duct 12, which shows the cooling air path at the trailing edge. In Figure (4) and Figure (5), views of the blade plan are shown in a longitudinal section to provide a better understanding of the invention. Figure (4) shows only the cooling channels without the ribs present. The thickness of the blade metal is shown schematically and hatched in this figure. The cooling air flow pressure is supplied from the channel 14, in which the cooling air flows from the inside of the disk to the paths inside the blade. Then, after entering the blade root, the flow is divided from path 16 into two paths 18 and 20. The latter two paths are separated by a wall 22. The specific shape of this area is important in that the flow rate ratio between the paths depends on the design of this area. Routes 16, 18, and 20 have a larger hydraulic diameter for two purposes: first, the large hydraulic diameter reduces heat transfer from the root to the cooling air (there is practically no need to cool this area), and second, the pressure drop at the branch point of route 16 to routes 18 and 20 is reduced.Path 18 is then responsible for delivering the flow to the beginning of the first spiral channel 24 and the beginning of the second spiral channel 26. Path 20 also delivers the flow to the beginning of the third spiral channel 28. The first spiral channel, starting at 24, is responsible for cooling the leading edge 7. Due to the stagnation of the gas flow outside at the leading edge 7, the heat exchange from the hot gases to the blade is very high in this area and the highest thermal load occurs at this point, so this area is susceptible to high damage. In the present invention, due to the importance of this point, the first spiral channel is considered L-shaped and separate, and the cooling air of this path is directed directly to the outlet after cooling the leading edge to achieve a lower pressure drop. Thus, this spiral channel has an acceptable suction flow rate and is capable of creating a high cooling rate. In addition, due to the high thermal load of the leading edge, the cool flow rapidly increases in temperature while passing through this spiral channel and its potential for cooling is reduced. Therefore, the air from the first spiral channel is then used only to cool the trailing edge at the highest radius, because at a higher radius, the mechanical stress is much lower and there is no need for significant cooling potential.The highest gas temperature and consequently the maximum blade temperature occur near the middle radius, therefore, to achieve a higher internal flow velocity and heat transfer coefficient, the hydraulic diameter of the first spiral channel is minimal in the middle sections of the leading edge to achieve a higher cooling rate. In addition, to increase the cooling efficiency at the leading edge 7, the first spiral channel has V-shaped ribs from 24 to the 90-degree bend 30. These ribs are installed at the leading edge of the first spiral channel with an angle of 45 degrees, a depth of 0.25 mm and a pitch of 2.5 mm. A view of these ribs is shown in Figure 6. As shown in the cross section of this duct in Figure 7, these ribs, in addition to disturbing the boundary layer, increase the mixing of the cool core of the flow with the hot boundary layer near the wall due to their special shape by producing two transverse swirling flows. The orientation of the ribs is such that these two induced vortices hit the front of the duct (region 66 in Figure 7), and as a result, the highest cooling heat transfer coefficient occurs in the front region 66.This special structure also results in lower pressure drop than vertical fins. Another advantage is the high flow resistance resulting from the negative effect of blade rotation, i.e., the Coriolis acceleration. This resistance, which is a direct result of these two induced vortices, prevents the flow from separating from the duct surface under the influence of the Coriolis acceleration and consequently prevents the formation of a hot zone. The second spiral channel, starting at 26, has three channels in the middle region of the fin 36, 38 and 40. It should be noted that each channel refers to a part of the spiral channel that is separated by 180 degree turns. The second spiral channel, with the help of angled ribs 56, is responsible for cooling a part of the middle region of the fin. Since this region has a lower thermal load, low density rib technology has been used in these paths to apply a small pressure drop to the flow. The angle of the ribs in this region is 60 degrees, which has the most optimal performance and, with a significant increase in heat transfer, they apply the least pressure drop to the flow. Also, the depth of the ribs of this spiral channel is 0.5. The use of properly designed angled ribs, by inducing a secondary transverse flow, largely neutralizes the effect of Coriolis acceleration and prevents flow separation, and also creates more balanced cooling in the duct by creating mixing. The flow in the second spiral channel passes through two upward ducts 36 and 40 and one downward duct 38. Therefore, the heating of the internal air is extremely high and is used to cool the trailing edge in the middle radius upward, where its stress is lower than in the lower radii. The flow in the second spiral channel enters the trailing edge after a 90-degree turn. The ribs 62 and the large pin-fins 64 are responsible for spreading the flow and increasing the cooling of the flow in the second spiral channel at the trailing edge. The flow of the third spiral channel, starting from 28, initially enters the upward channel 42 and after passing through the turn 44, enters the downward channel 46. The channel 42 of this spiral channel has a similar structure to the second spiral channel and is responsible for cooling a part of the middle region of the blade by having angled ribs 56 with an angle of 60 degrees and a depth of 0.5 mm. The ribs in this region have a low density, also the third spiral channel has only two channels to prevent the flow from overheating at the beginning of the path and to have a better cooling performance in the continuation of the spiral channel, especially at the trailing edge. The second channel of the third spiral channel 46, which cools the middle region of the blade, has the expanded surface technology 54. By placing the expanded surfaces 54 in the second passage of the third spiral channel 46, the flow turbulence in this area is not significantly increased, so there is little pressure drop.The heat transfer coefficient also remains around the values ​​of a flat plate, but the surface in contact with the coolant fluid is increased, so more heat is absorbed overall than on a flat surface. Given that the heat load in this area is lower, the use of this technology is suitable for this area. In the middle of the duct 46, part of the flow enters the exhaust edge through the path 48 and is discharged from there to directly provide cooling for the central area of ​​the exhaust edge because the sensitivity of this area is very high and supplying cooling air from below the exhaust edge alone is not responsible for cooling this area. A large portion of the flow 46 also enters the trailing edge after passing through the turn 50 and cools this area. Because the flow has a high tendency to exit directly, it is of great importance to create a proper flow distribution at the trailing edge so that the flow covers and cools the entire area. To create a proper flow distribution, ribs 58 and variable density pin-fins 60 are used. At the bottom of this area, three rows of pin-fins are used to prevent the flow from exiting directly, while in the higher areas only two rows of pin-fins are used. Pin-fins 60 increase the contact area and the heat transfer coefficient. The reason for installing pin-fins in this area is that although the trailing edge has a lower thermal load compared to the leading edge, since the thickness of the fin in this area is much less than in other places, increasing the cooling efficiency in this area is of great importance. In other words, the limited space in this area makes it impossible to use other technologies to increase the heat transfer coefficient. Therefore, by installing pin-fins in this area, the heat transfer coefficient and the contact area with the cooling fluid increase, and at the same time, the transverse strength of the fin increases. Finally, the cooling flow is discharged through the end of the trailing edge 52 and enters the hot gas path. Explanation of shapes, maps and diagrams Figure 1: Schematic of a typical cooling system in a moving blade of a gas turbine. Figure 2: Schematic of the external view of a typical moving blade in gas turbines. 1-Spear root 2-Spoke support surface (hub) 3-Airfoil 4-Spade tip 5-Pressure level 6-Suction level 7- Attacking Edge 8-Escape Edge Figure 3: Cross-sectional view of the blade in the present invention to show the cross-section of the cooling system ducts 9-First spiral canal 10-Conduits related to the second spiral canal 11- Channels related to the third spiral canal 12- Current output from the trailing edge Figure 4: Longitudinal cutaway view of the blade to show the cooling channels (ribs and pin-fins are not shown in this image) 14- The direction of flow entering the blade from the chamber between the disk and the moving blade. 16- Cooling flow inlet inside the blade 18-Supplier duct of the first and second spiral canals 20- The duct supplying the third spiral canal 22-Flow channel separating wall 24-Beginning of the first spiral canal 26-Beginning of the second spiral canal 28-Beginning of the third spiral canal 30-90 degree rotation of the first spiral channel 32-Flow path from the second spiral to the first spiral channel 34-Outlet area of ​​the first spiral canal 36- First duct, second spiral canal 38- Second spiral canal 40- Third duct of the second spiral canal 42-First channel of the third spiral canal 44-First turn of the third spiral channel 46- Second duct of the third spiral canal 48-Flow path from the middle of the second channel of the third spiral channel to the edge of the escape 50-Second rotation of the third spiral channel 52-Coolant flow exit from the trailing edge Figure 5: Views of the blade map in longitudinal section. 54-Extended Surface Technology 56-Low density angled ribs 58- Trailing edge ribs to guide the flow of the third spiral channel 60- Dense trailing edge pin-fins 62- Trailing edge ribs to guide the flow of the second spiral channel 64-Large trailing edge pin-fins to guide the flow of the second spiral channel Figure 6: View of angled ribs designed for the leading edge Figure 7: Induced vortices in the cross section of the leading edge cooling duct 66-Leading edge cooling duct leading edge A clear and precise statement of the advantages of the claimed invention over prior inventions. 1- In the present invention, the cooling intensity and technology used in each area of ​​the blade are proportional to the external thermal load applied by the hot gases to that area. In this way, areas with high thermal load are cooled appropriately while preventing excessive cooling of other areas. 2- The temperature distribution becomes as uniform as possible throughout the blade, preventing the creation of high temperature gradients in the blade. 3- This invention, with an optimized design for the blade cooling system, reduces the maximum blade temperature as well as the resulting temperature gradient in the blade and increases the blade life. This increase in blade life reduces the cost of inspecting and replacing hot turbine parts. 4- This design reduces the cooling air flow rate by increasing cooling efficiency and effectiveness. 5- By applying this invention, the amount of cooling air required is reduced, which in itself increases the efficiency of the turbine. Description of at least one implementation method for implementing the invention Given the appropriate performance of the technology presented in this innovation in increasing the life and efficiency of the moving blade, this innovation has the potential to be used in the production of moving blades for the national turbine. In particular, given the proposed scope of application of this technology, its application is suitable for the first moving blade of the upgraded version of the national turbine or similar turbine blades. Explicit mention of the industrial application of the invention The design presented in the present invention has the potential to be used in the gas turbine industry, and specifically in the design of an internal cooling system for moving blades in the first stages of the turbine. The technology presented has high cooling efficiency, and therefore its use in moving blades that are in the high gas temperature range of up to 1250 degrees Celsius is recommended. The geometry and path proposed in this invention for cooling moving blades of gas turbines can replace the common cooling geometry and technologies in blades available in this industry and have a more optimal performance in terms of blade heat transfer. Optimal performance in heat transfer increases the life of the blade by reducing the temperature and temperature gradient in the blade, and also increases the efficiency of the industrial gas turbine by reducing the cooling air flow rate. Brief description of the invention In the design presented in the present invention, the internal cooling system for the moving blade of the gas turbine is installed in such a way that the technology used in each part of the blade is proportional to the thermal load applied by the hot gases to that part. Thus, in areas with lower sensitivity, a smaller amount of cooling air cooling capacity is spent, and also in critical areas with higher sensitivity, cooling with a higher flux is applied. As a result, by increasing the efficiency and effectiveness of the cooling system, the amount of coolant flow rate consumed has been reduced, which has also increased the efficiency of the turbine. At the same time, the maximum blade temperature has been reduced and a more uniform distribution of blade temperature has been achieved, which also increases the blade life against high temperatures. While in previous conventional cooling methods, blade cooling is highly non-uniform, causing hot spots in the blade and ultimately reducing its life.Since the leading edge is exposed to the highest thermal load from the hot gases and is therefore highly sensitive, a separate spiral channel with V-shaped ribs is used to cool this area in order to increase the heat transfer coefficient. The air passing through this spiral channel then cools the upper area of ​​the blade and exits from the upper part of the trailing edge. In addition to this spiral channel, the blade cooling system also includes two other spiral channels that are responsible for cooling the middle area and the trailing edge of the blade. These two spiral channels have a large cross-sectional area for cooling the middle area of ​​the blade, and rib technology with low density and wide surfaces has been used to prevent high pressure drops in these areas. The flow of these two spiral channels also ultimately provides cooling of the trailing edge and is discharged from the end duct of the trailing edge to the main flow. Trailing edge cooling is also usually of high importance because the possibility of thermal damage is high due to the thinness of the blade in this area.Therefore, a combination of ribs and a relatively high density pin-fin array is used to cool the trailing edge. The thinness of the fin at the end of the trailing edge makes it difficult to use ribs. However, pin-fins both significantly increase the heat transfer coefficient and the thinness of the fin is not an obstacle to their manufacture.

Claims

Claims What is claimed: Claim 1) A hollow movable blade for use in an industrial gas turbine, comprising: a blade tip (shad), a support surface (hub), an airfoil extending from the support surface (hub) towards the blade tip and comprising a pressure surface and a suction surface which are connected at the leading edge and trailing edge, a root for holding said blade on a rotating disk, an internal cooling system comprising three spiral channels as follows: the first spiral channel comprising one duct in the radial direction, the second spiral channel comprising three ducts in the radial direction, the third spiral channel comprising two ducts in the radial direction. Claim 2) An internal cooling system for the blade of claim 1, comprising three spiral channels as follows: the first spiral channel comprising a duct in the radial direction for cooling the leading edge and a duct in the axial direction for cooling the upper part of the blade airfoil from the leading edge to the trailing edge, the second spiral channel comprising three ducts in the radial direction for cooling the area after the leading edge (after the area cooled by the first spiral channel) to the middle of the blade (in the flow direction) and also comprising a duct for cooling the trailing edge area of ​​the span (airfoil height of the blade) 50% to 80%, the third spiral channel comprising two ducts in the radial direction for cooling the middle area of ​​the blade in the flow direction to the trailing edge and also a duct for cooling the trailing edge area of ​​the span (airfoil height Blade) 0 to 50 percent. Claim 3) An internal cooling system for the blade of claim 1, which includes three spiral channels with cooling technologies as follows: The first channel of the first spiral channel has V-shaped ribs for cooling the leading edge, and after a 90-degree rotation, the second channel of this spiral channel extends axially towards the trailing edge and has pin-fins in the trailing edge section. The first to third channels of the second spiral channel have angled ribs with an angle of 60 degrees, and the fourth channel of the second spiral channel has pin-fins for spreading the trailing edge flow and cooling this area. The first channel of the third spiral channel has angled ribs with an angle of 60 degrees. The second channel of the third spiral channel has wide-area cooling technology. Also, the third duct of the third spiral channel has pin-fins to spread the trailing edge flow and cool this area. Claim 4) A cooling system according to claim 2, wherein the first spiral channel is an L-shaped spiral channel that first conveys the cool air flow at the leading edge, from the blade root to the blade tip. At the leading edge of this spiral channel, V-shaped ribs are responsible for absorbing the heat load. The second duct of this spiral channel extends in an axial direction at the blade tip towards the trailing edge. Claim 5) A cooling system according to claim 2, wherein the first passage of the first spiral channel has V-shaped ribs with an angle of 45 degrees, a depth of 0.25 mm and a pitch of 2.5 mm. Claim 6) A cooling system according to claim 2, wherein the first duct of the first spiral channel is convergent-divergent in that the cross-sectional area at the entrance of this duct is 19 square millimeters and decreases radially up to a span (airfoil height of the blade) of 50%. The minimum cross-sectional area of ​​this duct is located at a span (airfoil height of the blade) of 50%, which reaches 13 square millimeters. After this section, the cross-sectional area of ​​the duct increases radially and reaches 19 square millimeters at the end of this duct. Claim 7) A cooling system according to claim 2, wherein the second spiral channel has four channels, two 180 degree turns and one 90 degree turn, wherein the first to third channels have angled ribs with an angle of 60 degrees and a depth of 0.5 mm, and the fourth channel of this spiral channel is equipped with pin-fins for flow distribution and heat transfer. Claim 8) A cooling system according to claim 2, wherein the third spiral channel comprises three channels and two 180 degree turns. Wherein the first channel has angled ribs with an angle of 60 degrees and a depth of 0.5 mm, the second channel has wide surface cooling technology and the third channel has pin-fin and rib technologies. Claim 9) A cooling system according to claim 2, wherein at least two spiral channels of the internal cooling system are connected at the trailing edge and their combined flow provides trailing edge cooling. Claim 10) A turbine with at least one row of movable blades of claim 1, wherein the cooling system according to claim 2 is used.

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