Blade with wingtip pockets

Abstract

To avoid harmful vibrations thereby allowing longer turbine blades.SOLUTION: An airfoil comprises a pressure sidewall and a suction sidewall that extend axially between corresponding leading and trailing edges and radially between a base and an outboard tip end. The rotor blade further comprises at least one mid-span shroud configured to engage a corresponding mid-span shroud on at least one neighboring rotor blade during operation. The airfoil further comprises: an inboard region between the mid-span shroud and the base of the blade, with an inboard direction of the airfoil toward the base; and an outboard region between the mid-span shroud and the outboard tip end of the blade, with an outboard direction of the airfoil toward the outboard tip end. The outboard region comprises two cavities extending from the outboard tip end inboard of the airfoil toward the mid-span shroud; and the inboard region is substantially solid inboard of the mid-span shroud.SELECTED DRAWING: Figure 4

Description

Technical Field

[0001] This application generally relates to devices, methods, and / or systems related to the design and operation of turbine rotor blades. More specifically, without limitation, this application relates to devices and systems related to turbine blades and to the configuration of turbine blades having tip pockets.

Summary of the Invention

[0002] This application describes a rotor blade for use in a turbine of a combustion turbine engine. The rotor blade includes an airfoil. The airfoil includes a concave pressure sidewall and a convex suction sidewall that extend axially between a corresponding leading edge and trailing edge and radially between a base and an outer tip end. The rotor blade further includes at least one mid-span shroud configured to engage a corresponding mid-span shroud on at least one adjacent rotor blade during operation. The airfoil further includes an inner region between the at least one mid-span shroud and the base of the blade with the inner direction of the airfoil facing the base, and an outer region between the at least one mid-span shroud and the outer tip end of the blade with the outer direction of the airfoil facing the outer tip end. The outer region includes at least two cavities extending from the outer tip end inwardly of the airfoil towards the at least one mid-span shroud, and the inner region is substantially solid inside the at least one mid-span shroud.

[0003] Another aspect of this application describes a combustion turbine engine comprising a rotor blade having an airfoil section having a concave positive pressure sidewall and a convex negative pressure sidewall extending axially between corresponding leading and trailing edges and radially between the base and the outer wingtip. The rotor blade further comprises at least one midspan shroud configured to engage with a corresponding midspan shroud on at least one adjacent rotor blade during operation. The airfoil section further comprises an inner region between at least one midspan shroud and the base of the blade, with the inward direction of the airfoil section toward the base, and an outer region between at least one midspan shroud and the outer wingtip of the blade, with the outward direction of the airfoil section toward the outer wingtip. The outer region comprises at least two cavities extending inward from the outer wingtip toward at least one midspan shroud, and the inner region is substantially solid inside at least one midspan shroud. These and other features of the present application will become apparent upon consideration of the following detailed description of preferred embodiments in conjunction with the drawings and the appended claims.

[0004] These and other features of the Disclosure will be more fully understood and recognized by carefully considering the following more detailed description of exemplary embodiments of the Disclosure in conjunction with the accompanying drawings. [Brief explanation of the drawing]

[0005] [Figure 1] This is a schematic diagram of an exemplary combustion turbine engine in which embodiments of this application may be used. [Figure 2] Figure 1 is a cross-sectional view of the compressor of a combustion turbine engine. [Figure 3] Figure 1 is a cross-sectional view of the turbine of a combustion turbine engine. [Figure 4] This is a schematic diagram of an exemplary rotor blade having a midspan shroud and internal configuration according to one embodiment of the present disclosure. [Figure 5] This is a side view of a blade having a cavity and internal structure according to one embodiment of the present disclosure. [Figure 6] A side view of the outer region of a blade having a cavity and internal structure according to one embodiment of the present disclosure. [Figure 7] This is a partial upper cross-sectional view of the outer region of a rotor blade having a cavity according to an embodiment of the present disclosure. [Modes for carrying out the invention]

[0006] Please note that the drawings in this disclosure are not necessarily to scale. The drawings are intended to illustrate only typical embodiments of this disclosure and should not be considered to limit the scope of this disclosure. In the drawings, similar reference numerals between drawings represent similar elements.

[0007] First, it should be understood that in order to discuss the disclosures of this application, it may be necessary to select terminology to describe specific components within a combustion turbine engine. Wherever possible, common industrial terminology will be used and utilized in the same sense as its accepted meaning. However, it is intended that any such term be given a broad meaning and not to be interpreted narrowly in a way that unduly limits the intended meaning and the scope of the appended claims. Those skilled in the art will understand that a particular component may often be referred to using several different terms. In addition, what may be described as a single part herein may include multiple components and may be referred to in another context as consisting of multiple components, or what may be described herein as consisting of multiple components may be referred to elsewhere as a single part. Therefore, in understanding the scope of this disclosure, attention should be paid not only to the terminology and descriptions presented herein, but also to the appended descriptions, contexts, and in particular to the structure, configuration, function, and / or use of the components that may be provided in the appended claims.

[0008] In addition, several descriptive terms may be used in accordance with the rules of this specification, and it will be useful to define these terms at the beginning of this section. Therefore, these terms and their definitions are as follows, unless otherwise specified. As used herein, “downstream” and “upstream” are terms that indicate the direction of fluid flow relative to the working fluid through a turbine engine, or, for example, the flow of air through a combustor, or the coolant through one of the turbine's component systems. Thus, the term “downstream” corresponds to the direction of fluid flow, and the term “upstream” refers to the opposite direction of flow. The terms “forward” and “rear” refer to direction unless otherwise specified, with “forward” referring to the front end of the engine or compressor end, and “rear” referring to the rear end of the engine or turbine end. The term “radial” refers to movement or position perpendicular to the axis. It is often necessary to describe parts that are at different radial positions relative to the central axis. In such cases, if the first component is located closer to the axis than the second component, this specification states that the first component is “radially inward” or “inboard” the second component. On the other hand, if the first component is located further from the axis than the second component, this specification may state that the first component is “radially outward” or “outboard” the second component. The term “axial” refers to movement or position parallel to the axis. Finally, the term “circumferential” refers to movement or position around the axis. It will be understood that such terms may be applied in relation to the central axis of the turbine, or, when referring to components within the combustor, the central axis of the combustor.

[0009] In combustion turbine engines, it is well known that pressurized air from a compressor is used to burn fuel in a combustor, generating a flow of high-temperature combustion gases, which then flow downstream through one or more turbines from which energy can be extracted. In such turbines, rows of turbine rotor blades, spaced circumferentially, extend radially outward from a supporting rotor disk. Each blade typically includes a dovetail in a corresponding dovetail slot in the rotor disk, allowing for assembly and disassembly of the blade, as well as an airfoil extending radially outward from the dovetail and interacting with the flow of working fluid through the engine. The airfoil has substantially concave positive pressure sidewalls and substantially convex negative pressure sidewalls, extending axially between the corresponding leading and trailing edges and radially between the root and tip. The blade tips are spaced close to a radially outward fixed shroud, where it will be understood that leakage of combustion gases flowing downstream between the turbine blades is minimized.

[0010] Shrouds at the tips of airfoil sections, or tip shrouds, are often implemented in the downstream stage or on the rotor blades to provide damping and reduce tip leakage of working fluid. Given the length of the rotor blades in the downstream stage, the damping function of tip shrouds offers a significant performance advantage. However, considering the weight that tip shrouds add to the assembly, as well as other criteria including thousands of hours of operation exposed to high temperatures and extreme mechanical loads, it can be difficult to fully utilize the damping function. Therefore, while larger tip shrouds are desirable because they can provide a critical connection between adjacent rotor blades that can more effectively seal the gas paths and improve damping, larger tip shrouds can increase the load on the rotor blades.

[0011] The power output and efficiency of a gas turbine engine improve as the engine size, more specifically, the amount of air that can pass through the engine, increases. However, engine size is limited by the operational length of the turbine blades, and the longer the turbine rotor blades, the greater the flow path through the engine. However, longer rotor blades increase the mechanical load, placing further demands on the blades and the rotor disc that holds them. As rotor blades lengthen, the natural frequency of the blades during operation also decreases, which increases the vibration response of the rotor blades. The additional vibration load can place further demands on the rotor blades, which can further shorten the lifespan of the components and, in some cases, lead to vibration load. One way to address the vibration load of longer rotor blades is through shrouds that connect adjacent rotor blades to each other.

[0012] Another way to address this vibration load is to place one or more shrouds on the underside of the airfoil section of the rotor blade. That is, instead of adding a shroud to the tip of the rotor blade, the shroud can be placed near the radial center portion of the airfoil section. As used herein, such a shroud is referred to as a “midspan shroud.” At this lower (or inward) radius, the mass of the shroud reduces the stress level on the rotor blade. However, this type of shroud leaves a portion of the airfoil section of the rotor blade unrestrained, or in a cantilevered state, which is the portion of the airfoil section that extends outward from the midspan shroud towards the wingtip. This cantilevered portion of the airfoil section can typically result in lower frequency vibrations and increased vibration loads.

[0013] To address vibration and load considerations, reducing the overall weight of the outer portion of the blade may be beneficial. This reduction should, ideally, also alter the blade's frequency and mode geometry. This alteration will improve the efficiency of the blade and turbine.

[0014] As background, referring here to the drawings, Figures 1 to 3 show exemplary combustion turbine engines in which embodiments of this application may be used. It will be understood by those skilled in the art that embodiments are not limited to this type of use. As stated, embodiments can be used in combustion turbine engines, steam turbine engines, and other types of rotary engines, such as those used in power generation and aircraft.

[0015] Figure 1 is a schematic and non-limiting diagram of a combustion turbine engine 10. Generally, a combustion turbine engine operates by extracting energy from a pressurized, high-temperature gas flow generated by burning fuel in a stream of compressed air. As shown in Figure 1, the combustion turbine engine 10 can be configured to have an axial-flow compressor 11 mechanically coupled to the downstream turbine section, i.e., the turbine 13, by a common shaft or rotor, and a combustor 12 positioned between the compressor 11 and the turbine 13.

[0016] Figure 2 shows an exemplary, non-limiting, multistage axial-flow compressor 11 that may be used in the combustion turbine engine of Figure 1. As shown, the compressor 11 may have multiple stages. Each stage may comprise a row of compressor rotor blades 14 followed by a row of compressor stator nozzles 15. Thus, the first stage may comprise a row of compressor rotor blades 14 that rotate around a central shaft, followed by a row of compressor stator nozzles 15 that remain stationary during operation. The compressor stator nozzles 15 are generally spaced apart circumferentially from one another and fixed around the axis of rotation. The compressor rotor blades 14 are spaced apart circumferentially and mounted on the shaft. As the shaft rotates during operation, the compressor rotor blades 14 rotate with the shaft. The compressor rotor blades 14 are configured to impart kinetic energy to the air or fluid flowing through the compressor 11 as they rotate around the shaft. The compressor 11 may have other stages after the stages shown in Figure 2. The additional stage may comprise a plurality of compressor rotor blades 14 spaced apart in the circumferential direction, followed by a plurality of compressor stator nozzles 15 spaced apart in the circumferential direction.

[0017] Figure 3 shows a non-limiting partial view of an exemplary turbine section or turbine 13 that may be used in the combustion turbine engine of Figure 1. The turbine 13 may also comprise multiple stages. Three exemplary stages are shown, but these are illustrative and non-limiting and are not intended to limit any embodiments. Thus, the turbine 13 may have more or fewer stages. The first stage comprises a plurality of turbine blades or turbine rotor blades 16 (hereinafter referred to as "blades") that rotate around the shaft during operation, and a nozzle or turbine stator blade 17 (hereinafter referred to as "nozzle") that remains stationary during operation. The nozzles 17 are generally spaced apart from each other in the circumferential direction and fixed around the axis of rotation. The turbine rotor blades 16 may be mounted on a turbine wheel or disk (not shown) to rotate around a shaft (not shown). The second stage of the turbine 13 is also shown. The second stage similarly comprises a plurality of nozzles 17 spaced apart in the circumferential direction, followed by a plurality of turbine rotor blades 16 spaced apart in the circumferential direction, the turbine rotor blades 16 also mounted on a turbine wheel for rotation. A third stage is also shown, similarly comprising a plurality of nozzles 17 and rotor blades 16. It will be understood that the nozzles 17 and turbine rotor blades 16 are located within the hot gas path of the turbine 13. The direction of the hot gas flow through the hot gas path is indicated by arrows. The turbine 13 may have other stages after the stages shown in Figure 3. Each additional stage may comprise a row of nozzles 17 followed by a row of turbine rotor blades 16.

[0018] In a non-exclusive description of its use, the rotation of the compressor rotor blades 14 in an axial-flow compressor 11 can compress the airflow. In the combustor 12, when the compressed air is mixed with fuel and ignited, energy can be released. The resulting flow of hot gas from the combustor 12, which may be called the working fluid, is then directed over the turbine rotor blades 16, and the flow of the working fluid induces the rotation of the turbine rotor blades 16 and the shaft. Thus, the energy of the working fluid flow is converted into mechanical energy for the rotating blades, and furthermore, the shaft rotates due to the connection between the rotor blades and the shaft. The mechanical energy of the shaft can then be used to drive the rotation of the compressor rotor blades 14 and supply the necessary compressed air, and can also be used, for example, to drive a generator to produce electricity.

[0019] As used herein and as shown in Figures 4 to 7, the blade 16 comprises an airfoil section 25. The blade 16 comprises a root or base 121 at one end that can be attached to a rotor, and an outer wingtip 41 at the wingtip opposite the base 121 of the airfoil section 25. The airfoil section 25 comprises at least one midspan shroud 51, preferably a plurality of midspan shrouds 51. In relation to the midspan shroud 51, the airfoil section 25 defines the inner portion of the airfoil section 25 between the midspan shroud 51 and the base 121 of the blade 16, with the "inward" orientation or direction toward the base 121 of the airfoil section 25. Furthermore, in relation to the midspan shroud 51, the airfoil section 25 defines the outer portion of the airfoil section 25 between the midspan shroud 51 and the outer wingtip 41 of the blade 16, with the "outward" orientation or direction toward the outer wingtip 41 of the airfoil section 25.

[0020] Furthermore, the airfoil section 25 has an inner region 58 which is the portion of the airfoil section 25 that is radially inward of at least one of the multiple midspan shrouds 51. The airfoil section 25 also has an outer region 59 which is the portion of the airfoil section 25 that is radially outward of the midspan shrouds 51.

[0021] Embodiments will now be described in relation to blades with mid-span shrouding, but the features of this embodiment are not intended to be limited to blades with mid-span shrouding. The features disclosed herein are applicable to blades with mid-span shrouding as well as blades without mid-span shrouding.

[0022] Figures 4 and 5 are side views of rotor blade 16 having mid-span shrouds 51 and 51' and an internal structure or configuration (shown in dashed lines) according to embodiments of the present disclosure. Figures 6 and 7 show embodiments of the present disclosure focused on the outer region 59 of airfoil 25 as described below.

[0023] According to embodiments of the present disclosure, the inner region 58 is solid or substantially solid, and the outer region 59 is hollow or "substantially" hollow. As used herein, the outer region 59 is a hollow region 64 that is a hollow, any region or space within airfoil 25, for example including cavities 61 and 66 formed therein (described below). The inner region includes a solid region 62 that is any region or space within airfoil 25 that is a substantially solid material. According to a non-limiting aspect of the embodiment, the outer region 59 has a hollowness ratio of at least 40%, and the inner region 58 has a solidness ratio of at least 60%.

[0024] According to embodiments of the present disclosure, the outer region 59 can include at least two cavities 61 and 66 formed therein. As shown in the figures, the at least two cavities 61 and 66 are closed at their inner ends and are open at the outer wing tip end 41 of airfoil 25 in an outer orientation. As shown in Figures 4 - 7, cavities 61 and 66 occupy the space between the positive pressure sidewall 26 and the negative pressure sidewall 27 of airfoil 25. According to a non-limiting aspect of the present disclosure, the entirety of the at least two cavities 61 and 66 can be within the outer region 59 above mid-span shrouds 51 and 51'.

[0025] Although two cavities are shown in the drawings, the two cavities 61 and 66 are merely illustrative and do not limit the scope of the embodiments. The embodiments of this specification are not limited by the drawings. Three or more cavities are within the scope of the present disclosure.

[0026] The cavities 61, 66 in the outer region 59 each comprise side walls 65 and 67 and bottom walls 165 and 167 respectively. As embodied in one non-limiting aspect of the present disclosure, the bottom walls 165 and 167 of the cavities 61, 66 may be at an equal distance within the airfoil portion 25. However, the cavities 61, 66 do not have to extend at approximately equal distances and can extend at various different distances within the airfoil portion 25. As shown in another non-limiting aspect of FIGS. 5 and 6, the bottom walls 165 and 167 of the cavities 61, 66 do not have to be at an equal distance towards the inner region 58.

[0027] Furthermore, according to another non-limiting aspect of the embodiment, the bottom walls 165 and 167 can be substantially orthogonal to the side walls 65 and 67 as shown in FIG. 6. In the figure, the bottom walls are shown as being substantially orthogonal to the side walls, but the scope of the embodiment is not limited to the bottom walls being substantially orthogonal. As further shown in FIG. 6, a further non-limiting aspect of the embodiment can provide bottom walls angled with respect to the side walls 65 and 67, as shown by the bottom walls 265 and 267 drawn as dashed lines.

[0028] In addition, in other non-limiting aspects of the embodiment, the angles of the bottom walls 265 and 267 are angles that cross, an angle that spreads from the leading edge 20 to the trailing edge 21 of the airfoil portion 25, an angle that spreads from the trailing edge 21 to the leading edge 20 of the airfoil portion 25, an angle that spreads from the positive pressure side to the negative pressure side, an angle that spreads from the negative pressure side to the positive pressure side, and / or one of the bottom walls 265 and 267 is angled from the leading edge 20 to the trailing edge 21 of the airfoil portion 25, and the other of the bottom walls 265 and 267 is angled from the trailing edge 21 to the leading edge 20 of the airfoil portion 25.

[0029] Furthermore, non-limiting embodiments of the present disclosure also include bottom walls that are at equal distances from the outer wingtip 41 into the airfoil 25, as shown by the dashed lines in Figure 6, bottom walls 365 and 367. Other alternative configurations of the bottom walls may incorporate orthogonal bottom walls with angled portions, curved bottom walls, and any combination of bottom walls, as described in the non-limiting and exemplary description herein.

[0030] The illustrated embodiment shows two cavities 61 and 66. This configuration of two cavities is not limiting to the embodiment and is merely an example of the number of cavities that may be provided herein. The airfoil section 25 may have three or more cavities 61 and 66. In the configuration shown in Figure 7, reinforcing ribs 68 are formed on the airfoil section 25 at the outer wingtip 41 between each cavity 61 and 66. The reinforcing ribs 68 enhance the structural integrity and overall strength of the airfoil section, particularly the portion of the airfoil section 25 whose mass is reduced in the cavities 61 and 66 in the outer region 59. Naturally, if three or more cavities are provided, a larger number of reinforcing ribs 68 can be formed, further improving the integrity of the airfoil section.

[0031] As shown in Figure 6, the cavities 61 and 66 of the outer region 59 extend a distance A from the leading edge 20 toward the trailing edge 21. Given the configuration of the airfoil 25, it is desirable that distance A be less than or equal to about 60% of the airfoil 25, which should result in a more solid mass near the leading edge 21 compared to the cavities in the region of the airfoil 25 adjacent to the trailing edge 20, thus maintaining the integrity and overall strength of the airfoil 25. Thus, the outer region 59 has a more solid material but is still weight-reduced, thereby reducing vibration and load. Aspects of the embodiment may also include a distance A of less than 60% of the airfoil 25.

[0032] According to another embodiment, each cavity 61 and 66 can have a different volume. As stated above, the bottom walls of cavities 61 and 66 as described herein do not need to be equal. Also, considering the configuration of the airfoil section 25, each cavity has a distinct periphery that matches the shape and profile of the airfoil section. For example, in the non-limiting example of Figure 6, cavity 61 defines a first periphery 161 of the side wall 65, and cavity 66 defines a second periphery 166 of the side wall 65. These peripheries 161 and 166 are not necessarily equal, and therefore, even if the bottom walls extend to the same depth within the hollow region 64, the areas of the respective cavities 61 and 66 may not be equal.

[0033] Furthermore, the side walls of cavities 61 and 66 may vary along the depth of cavities 61 and 66. For example, although not intended to limit the embodiments, Figure 7 shows a cavity 61 having two different side walls 465 and 466, which are not limiting embodiments. Side wall 465 is a roughened, uneven side wall. Side wall 466 is tapered so that its width increases from the outer wingtip 41 toward its bottom wall (omitted for ease of illustration). These side wall structural configurations are merely a number of non-limiting side wall structures, whatever form the end walls may take (as described above). In other words, the geometric shapes of cavities 61 and 66 may differ. The volumes of cavities 61 and 66 may not be equal, the depths of cavities 61 and 66 toward the hollow region 64 may vary, and the configuration, orientation, and angles of the side walls and bottom walls may vary. Furthermore, the area of ​​the periphery of each cavity may vary. All of these modifications allow the volumes of cavities 61 and 66 not to be equal. However, as can be assumed, the volumes of cavities 61 and 66 may be the same if the above factors determining the cavity volumes are chosen to be equal in volume determination.

[0034] The blade 16, as embodied by this disclosure, provides an airfoil section 25 with reduced overall weight by having cavities 61 and 66. The reduction in overall weight should suppress vibration and excitation forces on the blade 16 during operation. The blade 16 with cavities 61 and 66 can reduce vibration forces and internal loads on the blade 16. Furthermore, the lighter blade 16 provided by cavities 61 and 66 can reduce the magnitude of vibration forces (or natural frequencies) so that the mode shape or deflection at the natural frequencies of such a blade is lower (compared to a blade without cavities 61 and 66). Thus, degradation of the blade 16 with cavities 61 and 66 can be suppressed.

[0035] The midspan shrouds 51 and 51' according to this disclosure can be broadly defined to include any shrouds located inside the outer wingtip 41 and outside the base 121 of the airfoil section 25. According to a non-limiting aspect of this disclosure, as described above, the entirety of at least two cavities 61 and 66 should be located in the outer region 59 above the midspan shrouds 51 and 51'.

[0036] The midspan shrouds 51 and 51' according to this disclosure may also be positioned within a range of radial positions on the airfoil section 25. According to certain embodiments of this disclosure, the range of positions for the midspan shrouds 51 and 51' is defined between the inner boundary at about 25% of the radial height of the airfoil section 25 and the outer boundary at about 75% of the radial height of the airfoil section 25. According to other embodiments of this disclosure, as may be defined by the appended claims, the range of positions for the midspan shrouds 51 and 51' is defined between the inner boundary at about 33% of the radial height of the airfoil section 25 and the outer boundary at about 66% of the radial height of the airfoil section 25. According to certain embodiments of this disclosure, the midspan shrouds 51 and 51' are positioned near the substantially radial center region of the airfoil section 25.

[0037] According to some non-limiting and exemplary embodiments described above, it will be understood that the vibration response of the turbine blades can be reduced to limit damaging mechanical loads, thereby enabling the extension of the rotor blades and achieving higher engine efficiency. In other words, the disclosure teaches that cavities 61 and 66 are provided in the turbine blades to limit the vibration response of the outer regions 59 extending beyond one or more midspan shrouds 51 and 51', thereby increasing rigidity and reducing the portions of the airfoil section 25 outside the midspan shrouds 51 and 51'. In this way, harmful vibrations can be avoided, thereby enabling longer turbine blades.

[0038] While this disclosure describes what is considered to be the most practical and preferred embodiment at present, it should be understood that this disclosure is not limited to the disclosed embodiment, but rather is intended to encompass preferred modifications and equivalent configurations that fall within the spirit and scope of the attached claims. [Explanation of Symbols]

[0039] 10 Combustion Turbine Engine 11. Multistage axial flow compressor 12 Combustor 13 Turbine 14 Compressor rotor blades 15 Compressor stator nozzle 16 Turbine rotor blades 17 Nozzles, turbine stator blades 20 Leading edge 21 Trailing edge 25 Airfoil section 26 Positive pressure sidewall 27 Negative pressure sidewall 41 Outer wing tip 51 Mid-span shroud 51' Mid-span shroud 58 Inner area 59 Outer area 61 Cavity 62. Solid area 64 Hollow area 65 side wall 66 Cavity 67 Side wall 68 Reinforcement Ribs 121 Base 161 First peripheral area 165 Bottom wall 166 Second Periphery 167 Bottom wall 265 Bottom wall 267 Bottom wall 365 Bottom wall 367 Bottom wall 465 Side wall 466 Side wall A distance

Claims

1. A rotor blade (16) for use in a turbine (13) of a combustion turbine engine (10), wherein the rotor blade (16) is The airfoil section (25) has a concave positive pressure sidewall (26) and a convex negative pressure sidewall (27) that extend axially between the corresponding leading edge (20) and trailing edge (21), and radially between the base (121) and the outer wingtip (41). Equipped with, The rotor blade (16) At least one midspan shroud (51) configured to engage with a corresponding midspan shroud (51) on at least one adjacent rotor blade (16) during operation. Furthermore, The airfoil portion (25) is positioned such that the inward direction of the airfoil portion (25) is toward the base portion (121), and further comprises an inner region (58) between the at least one midspan shroud (51) and the base portion (121) of the blade (16), The airfoil portion (25) is configured such that the outward direction of the airfoil portion (25) is toward the outer wingtip (41), and further comprises an outer region (59) between the at least one midspan shroud (51) and the outer wingtip (41) of the blade (16), The outer region (59) comprises at least two cavities (61, 66) extending inward from the outer wingtip (41) which is its starting point toward the at least one midspan shroud (51) within the airfoil (25), The two cavities (61, 66) have the outer wingtip (41) as their starting point, The at least two cavities (61, 66) include a first cavity (61) that is close to the leading edge (20) and a second cavity (66) that is far from the leading edge (20), The second cavity (66) is deeper in the inward direction than the first cavity (61), The first cavity (61) comprises side walls (65, 67, 465, 466) and a bottom wall (265), The bottom wall (265) of the first cavity (61) is positioned at an acute angle with respect to the side wall (65, 67, 465, 466) of the first cavity (61) that is close to the trailing edge (21). The second cavity (66) comprises side walls (65, 67, 465, 466) and a bottom wall (267), The bottom wall (267) of the second cavity (66) is positioned at an acute angle with respect to the side wall (65, 67, 465, 466) of the second cavity (66) that is close to the trailing edge (21). A rotor blade (16) having a substantially solid inner region (58) inside at least one midspan shroud (51).

2. The rotor blade (16) according to claim 1, wherein each of the at least two cavities (61, 66) is open at the outer wingtip (41) and closed at their inner ends.

3. The rotor blade (16) according to claim 1, wherein the bottom wall (165, 167, 365, 367) is perpendicular to at least one of the side walls (65, 67, 465, 466).

4. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) have different longitudinal cross-sections and cross-sections in the direction from the leading edge to the trailing edge.

5. The rotor blade (16) according to claim 4, wherein the side walls (65, 67, 465, 466) of the cavities (61, 66) have a roughened, uneven surface.

6. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) comprise three or more cavities (61, 66) formed along the chord of the blade (16).

7. The rotor blade (16) according to claim 1, wherein at the outer wingtip (41), the at least two cavities (61, 66) extend up to 60% of the outer wingtip (41) from the leading edge (20) to the trailing edge (21).

8. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) define a reinforcing rib (68) between them.

9. The rotor blade (16) according to claim 1, wherein the outer region (59) has a hollow ratio of at least 40%.

10. A combustion turbine engine (10), wherein the combustion turbine engine (10) A combustion turbine engine (10) comprising a rotor blade (16) according to any one of claims 1 to 9.

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