Blades with tip pits

Abstract

The present invention is entitled "Blade with Tip Recess". The present invention discloses a rotor blade (16) for use in the turbine of a gas turbine engine (10), the rotor blade comprising an airfoil (25). The airfoil (25) comprises a concave pressure sidewall (26) and a convex suction sidewall (27) extending axially between corresponding leading edges (20) and trailing edges (21) and extending radially between the base (121) and the outer tip (41). The rotor blade (16) also includes at least one midspan shroud (51) configured to engage a corresponding midspan shroud (51) on at least one adjacent rotor blade (16) during operation. The airfoil (25) further includes an inner region (58) between the at least one midspan shield (51) and the base (121) of the blade (216), wherein the airfoil (25) faces inward toward the base (121); and an outer region (59) between the at least one midspan shield (51) and the outward tip (41) of the blade (121), wherein the airfoil (25) faces outward toward the outward tip (41). The outer region (59) includes at least two cavities (61, 66) extending from the outward tip (41) of the airfoil (25) toward the at least one midspan shield (51); and the inner side of the at least one midspan shield (51), wherein the inner region (58) is substantially solid.

Description

Background Technology

[0001] This application relates in its entirety to apparatus, methods, and / or systems relating to the design and operation of turbine rotor blades. More specifically, but not limited to, this application relates to apparatus and systems relating to turbine blades and configurations of turbine blades with tip recesses. Summary of the Invention

[0002] This application describes a rotor blade used in a turbine of a gas turbine engine. The rotor blade includes an airfoil. The airfoil includes a concave pressure sidewall and a convex suction sidewall extending axially between corresponding leading and trailing edges and radially between a base and an outer blade tip. The rotor blade also includes at least one midspan shroud configured to engage a corresponding midspan shroud on at least one adjacent rotor blade during operation. The airfoil also includes an inner region located between at least one midspan shroud and the base of the blade, wherein the inward direction of the airfoil faces the base; and an outer region located between at least one midspan shroud and the outer blade tip of the blade, wherein the outward direction of the airfoil faces the outer blade tip. The outer region includes at least two cavities extending from the outer blade tip inside the airfoil toward at least one midspan shroud; and an inner side of at least one midspan shroud, the inner region being substantially solid.

[0003] Another aspect of this application discloses a gas turbine engine including rotor blades, each rotor blade comprising an airfoil having a concave pressure sidewall and a convex suction sidewall extending axially between corresponding leading and trailing edges and radially between a base and an outer blade tip. The rotor blade also includes at least one midspan shroud configured to engage corresponding midspan shrouds on at least one adjacent rotor blade during operation. The airfoil further includes an inner region located between at least one midspan shroud and the base of the blade, wherein the inward direction of the airfoil faces the base; and an outer region located between at least one midspan shroud and an outer blade tip, wherein the outward direction of the airfoil faces the outer blade tip. The outer region includes at least two cavities extending from the outer blade tip inside the airfoil toward at least one midspan shroud; and an inner side of at least one midspan shroud, the inner region being substantially solid. These and other features of this application will become apparent when viewed in conjunction with the accompanying drawings and the appended claims in the following detailed description of preferred embodiments. Attached Figure Description

[0004] These and other features of this disclosure will be more fully understood and appreciated through careful study of the following more detailed description of exemplary embodiments of this disclosure, taken in conjunction with the accompanying drawings, in which:

[0005] Figure 1 This is a schematic diagram of an exemplary gas turbine engine, in which embodiments of this application can be used;

[0006] Figure 2 for Figure 1 A cross-sectional view of the compressor in a gas turbine engine;

[0007] Figure 3 for Figure 1 A cross-sectional view of the turbine in a gas turbine engine;

[0008] Figure 4 This is a schematic diagram of an exemplary rotor blade having a mid-span shield and internal configuration according to an embodiment of this disclosure;

[0009] Figure 5 A side view of a blade having a cavity and internal configuration according to an embodiment of the present disclosure;

[0010] Figure 6 This is a side view of the outer region of a blade having a cavity and internal configuration according to one embodiment of the present disclosure; and

[0011] Figure 7 This is a partial cross-sectional top view of the outer region of a rotor blade with a cavity according to an embodiment of the present disclosure.

[0012] It should be noted that the accompanying drawings of this disclosure are not necessarily drawn to scale. The drawings are intended to depict only typical aspects of this disclosure and should therefore not be considered as limiting the scope of this disclosure. In the drawings, similar numbers denote similar elements between figures. Detailed Implementation

[0013] As an initial point of view, it should be understood that in order to discuss the disclosure of this application, it may be necessary to choose terms to refer to and describe specific components within a gas turbine engine. Wherever possible, common industry terms will be used and adopted in a manner consistent with the accepted meaning of the terms. However, this means that any such term has a broad meaning and should not be interpreted narrowly to unreasonably limit the meaning contemplated herein and the scope of the appended claims. Those skilled in the art will understand that several different terms may be used to refer to a particular component. Furthermore, what is described herein as a single component may include and be referred to in another context as comprising multiple components, or what is described herein as comprising multiple components may elsewhere be referred to as a single component. Therefore, in understanding the scope of this disclosure, attention should be paid not only to the terms and descriptions provided herein, but also to the structure, configuration, function, and / or use of the components, particularly as provided in the appended claims.

[0014] In addition, several descriptive terms may be used periodically throughout this document, and it should prove helpful to define these terms at the beginning of this section. Therefore, unless otherwise stated, these terms and their definitions are as follows. As used herein, “downstream” and “upstream” are terms indicating the direction of fluid flow, such as the working fluid through a turbine engine, or, for example, the airflow through a combustor or the coolant through a component system of a turbine. Thus, the term “downstream” corresponds to the direction of fluid flow, and the term “upstream” refers to the direction opposite to the flow. Without any other particularity, the terms “front” and “rear” refer to directions, where “front” refers to the front end of the engine or compressor end, and “rear” refers 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 located at different radial positions relative to the central axis. In such cases, if the first part resides closer to the axis than the second part, this document will state that the first part is “radially inward” or “inner” of the second part. On the other hand, if the first part resides further away from the axis than the second part, this document may state that the first part is “radially outward” or “outer” of the second part. The term "axial" refers to movement or position parallel to an axis. Finally, the term "circumferential" refers to movement or position about an axis. It should be understood that such terms can be applied relative to the central axis of the turbine, or, when referring to components within the combustor, relative to the central axis of the combustor.

[0015] In gas turbine engines, it is well known that air pressurized in a compressor is used to burn fuel in a combustor to generate a hot combustion gas stream, which then flows downstream through one or more turbines from which energy can be extracted. According to such a turbine, generally, rows of circumferentially spaced turbine rotor blades extend radially outward from a supporting rotor disk. Each blade typically includes a dovetail and an airfoil. The dovetail allows the blade to be assembled and disassembled in a corresponding dovetail-shaped slot in the rotor disk, and the airfoil extends radially outward from the dovetail and interacts with the working fluid flowing through the engine. The airfoil has a generally concave pressure side and a generally convex suction side, which extend axially between corresponding leading and trailing edges and radially between the root and the tip. It should be understood that the blade tip is closely spaced from the radially outer stationary shroud to minimize leakage of combustion gas flowing downstream between the turbine blades between the blade tip and the radially outer stationary shroud.

[0016] A shroud or tip shield at the airfoil tip is typically implemented on the rear tower or rotor blades to provide damping and reduce excessive tip leakage of the working fluid. Given the length of the rotor blades at the rear pole, the damping function of the tip shield provides significant performance benefits. However, fully utilizing the damping function can be challenging considering the weight added to the assembly by the tip shield and other criteria, including thousands of hours of operation exposed to high temperatures and extreme mechanical loads. Therefore, while large tip shields are desirable because they more effectively seal the gas path and provide significant connectivity between adjacent rotor blades, improving damping, larger tip shields can increase the load on the rotor blades.

[0017] The output and efficiency of a gas turbine engine improve with increasing engine size and, more specifically, with the increase in the amount of air it can carry. However, engine size is limited by the operable length of the turbine blades, where longer turbine rotor blades allow for a wider flow path through the engine. However, longer rotor blades introduce increased mechanical loads, placing further demands on the blades and the rotor disks that hold them in place. Longer rotor blades can also lower the blades' natural vibration frequency during operation, increasing the rotor blades' vibrational response. This additional vibrational load can place demands on the rotor blades, further shortening component life and, in some cases, causing vibrational loads. One way to address the vibrational loads of longer rotor blades is by using shrouds that connect adjacent rotor blades to each other.

[0018] Another way to address this vibrational load is to position one or more shrouds on the lower portion of the rotor blade airfoil. That is, instead of adding a shroud to the blade tip, the shroud can be positioned near the middle radial portion of the airfoil. As used herein, such a shroud will be referred to as a “midspan shroud.” At this lower (or more inner) radius, the mass of the shroud results in a reduction in the stress level on the rotor blade. However, this type of shroud leaves a portion of the rotor blade airfoil unrestricted or cantilevered, which is the portion of the airfoil extending outside the midspan shroud. This cantilevered portion of the airfoil typically results in lower frequency vibrations and increased vibrational loads.

[0019] To address vibration and load issues, reducing the overall weight of the outer portion of the blade can be beneficial. This reduction should also advantageously alter the blade's frequency and mode shape. This change will enhance blade and turbine efficiency.

[0020] For background information, please now refer to the attached diagram. Figures 1 to 3An exemplary gas turbine engine in which embodiments of this application can be used is shown. Those skilled in the art will understand that this embodiment is not limited to this type of application. As mentioned above, this embodiment can be used in gas turbine engines, such as engines for power generation and aircraft, steam turbine engines, and other types of rotary engines.

[0021] Figure 1 This is an illustrative, non-limiting representation of a gas turbine engine 10. Generally, a gas turbine engine operates by extracting energy from a pressurized, hot gas stream generated by fuel combustion in a compressed air stream. For example... Figure 1 As shown, the gas turbine engine 10 can be configured to have an axial compressor 11 mechanically connected to a downstream turbine section or turbine 12 via a common shaft or rotor, and a combustor 13 disposed between the compressor 11 and the turbine 12.

[0022] Figure 2 It shows that it can be used Figure 1 A view of an exemplary non-limiting multi-trailer axial compressor 11 for a gas turbine engine. As shown, the compressor 11 may include multiple trays. Each tray may include a row of compressor rotor blades 14, followed by a row of compressor stator nozzles 15. Thus, a first tray may include a row of compressor rotor blades 14 that rotates about a central axis, followed by a row of compressor stator nozzles 15 that remain stationary during operation. The compressor stator nozzles 15 are generally circumferentially spaced from each other and stationary about an axis of rotation. The compressor rotor blades 14 are circumferentially spaced and attached to the shaft. As the shaft rotates during operation, the compressor rotor blades 14 rotate with it. The compressor rotor blades 14 are configured such that, when rotated about the shaft, they impart kinetic energy to air or fluid flowing through the compressor 11. The compressor 11 may have a... Figure 2 Other trays besides the tray shown. Additional trays may include a plurality of circumferentially spaced compressor rotor blades 14, followed by a plurality of circumferentially spaced compressor stator nozzles 15.

[0023] Figure 3 It shows what can be used for Figure 1A partial view of an exemplary non-limiting turbine section or turbine 13 of a gas turbine engine. Turbine 13 may also include multiple towers. Three exemplary towers are shown, but this is merely exemplary and non-limiting, and is not intended to limit the implementation in any way. Therefore, more or fewer towers may be present in turbine 13. The first tower includes a plurality of turbine blades or turbine rotor blades 16 (hereinafter “blades”) that rotate about a shaft during operation, and a plurality of nozzles or turbine stator blades 17 that remain stationary during operation. The nozzles 17 are generally circumferentially spaced from each other and stationary about an axis of rotation. The turbine rotor blades 16 may be mounted on a turbine impeller or disk (not shown) to rotate together with the shaft (not shown). A second tower of turbine 13 is also shown. The second tower similarly includes a plurality of circumferentially spaced nozzles 17, followed by a plurality of circumferentially spaced turbine rotor blades 16, which are also mounted on the turbine impeller for rotation. A third tower is also shown, and similarly includes a plurality of nozzles 17 and rotor blades 16. It should be understood that nozzle 17 and turbine rotor blades 16 are located in the hot gas path of turbine 13. The flow direction of the hot gas through the hot gas path is indicated by arrows. Turbine 13 may have... Figure 3 Other trays besides the tray shown. Each additional tray may include a row of nozzles 17, followed by a row of turbine rotor blades 16.

[0024] In the non-limiting description used, the compressor rotor blades 14 rotate within the axial compressor 11 to compress the airflow. In the burner 12, energy is released when compressed air is mixed with fuel and ignited. The resulting hot gas flow from the burner 12 (which may be referred to as the working fluid) is then directed over the turbine rotor blades 16, causing the turbine rotor blades 16 and the shaft to rotate. Thus, the energy of the working fluid flow is converted into the mechanical energy of the rotating blades, and 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, thereby generating the necessary compressed air supply, and also, for example, to drive a generator to produce electricity.

[0025] As used in this article and as Figures 4 to 7As shown, the blade 16 includes an airfoil 25. The blade 16 includes a root or base 121 attachable at one end to the rotor, and an outer tip portion 41 at the tip portion of the airfoil 25 opposite to the base 121. The airfoil 25 includes at least one midspan shroud 51 and preferably includes multiple midspan shrouds. Referring to the midspan shroud 51, the airfoil 25 defines an inner portion of the airfoil 25 between the midspan shroud 51 and the base 121 of the blade 16, wherein the "inner" orientation or direction is toward the base 121 of the airfoil 25. Furthermore, referring to the midspan shroud 51, the airfoil 25 defines an outer portion of the airfoil 25 between the midspan shroud 51 and the outer tip portion 41 of the blade 16, wherein the "outer" orientation or direction is toward the outer tip portion 41 of the airfoil 25.

[0026] Additionally, the airfoil 25 has an inner region 58, which is the portion of the airfoil 25 that is radially inward of at least one of the plurality of mid-span shields 51. The airfoil 25 also has an outer region 59, which is the portion of the airfoil 25 that is radially outward of the mid-span shield 51.

[0027] The embodiments described below will refer to blades including a midspan shield; however, the features of this embodiment are not intended to be limited to blades having a midspan shield. The components disclosed herein are applicable to blades with or without a midspan shield.

[0028] Figure 4 and Figure 5 The image shows a side view of a rotor blade 16 according to an embodiment of the present disclosure, which has mid-span shrouds 51 and 51' and an internal structure or configuration (shown via dashed lines). Figure 6 and Figure 7 An embodiment of the present disclosure is shown, concentrated on the outer region 59 of the airfoil 25, as described below.

[0029] According to embodiments of this 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 includes a hollow region 64, which is any hollow area or space within the airfoil 25, such as cavities 61 and 66 formed therein (discussed below). Additionally, the inner region includes a solid region 62, which is any area or space within the airfoil 25 that is substantially solid material. According to a non-limiting aspect of the embodiments, the outer region 59 has a hollowness percentage of at least 40%, and the inner region 58 has a solidity percentage of at least 60%.

[0030] According to embodiments of this disclosure, the outer region 59 may include at least two cavities 61 and 66 formed therein. As shown, at least two cavities 61 and 66 are closed at their inner ends and open outwardly at the tip 41 of the outer blade of the airfoil 25. Figures 4 to 7 As shown, cavities 61 and 66 occupy the volume between the pressure sidewall 26 and the suction sidewall 27 of the wing 25. According to a non-limiting aspect of this disclosure, at least two cavities 61 and 66 may be located entirely in the outer region 59 above the mid-span shields 51 and 51'.

[0031] Although the accompanying drawings show two cavities, cavities 61 and 66 are merely non-limiting and illustrate the scope of the embodiments. The embodiments described herein are not intended to be limited by the drawings. More than two cavities are permitted within the scope of this disclosure.

[0032] Cavities 61 and 66 in outer region 59 respectively include sidewalls 65 and 67 and bottom walls 165 and 167. As shown in a non-limiting aspect of this disclosure, the bottom walls 165 and 167 of cavities 61 and 66 can enter the airfoil 25 at approximately equal distances. However, cavities 61 and 66 do not need to extend at approximately equal distances, but can extend into the airfoil 25 by varying distances. Figure 5 and Figure 6 As shown in another non-limiting aspect, the bottom walls 165 and 167 of cavities 61 and 66 do not need to be equidistant from the inner region 58.

[0033] Furthermore, according to other non-limiting aspects of the implementation scheme, bottom walls 165 and 167 may be substantially orthogonal to side walls 65 and 67, such as... Figure 6 As shown in the attached figures. Although the bottom wall is shown to be approximately orthogonal to the side walls, the scope of the embodiment does not limit the bottom wall to approximately orthogonal. Similarly, as... Figure 6 As shown, another non-limiting aspect of the implementation may provide a bottom wall angled to the sidewalls 65 and 67, as indicated by the dashed sidewalls 265 and 267.

[0034] In other non-limiting aspects of the implementation, the angles of the bottom walls 265 and 267 may intersect, extending at an angle from the leading edge 20 of the airfoil 25 to the trailing edge 21, at an angle from the trailing edge 21 of the airfoil 25 to the leading edge 20, at an angle from the pressure side to the suction side, at an angle from the suction side to the pressure side, and / or one bottom wall 265 and 267 extends at an angle from the leading edge 20 of the airfoil 25 to the trailing edge 21, while the other bottom wall 265 and 267 extends at an angle from the trailing edge 21 of the airfoil 25 to the leading edge 20.

[0035] Furthermore, non-limiting aspects of the embodiments disclosed herein also include bottom walls that extend equidistantly into the airfoil 25 from the outer blade tip 41 of the side walls 65 and 67, such as... Figure 7 The dotted-line sidewalls 365 and 367 are indicated in the text. Other alternative arrangements of the base wall may be combined with any combination of orthogonal base walls, curved base walls, and base walls, based on the non-limiting and exemplary description herein.

[0036] The illustrated embodiment shows two cavities 61 and 66. This configuration of two cavities is not limiting for the embodiments, and the number of cavities that can be provided herein is merely exemplary. More than two cavities 61 and 66 may be provided in the airfoil 25. Figure 7 In the configuration shown, stiffening ribs 68 are formed in the airfoil 25 at the tip 41 of the outer blade between each cavity 61 and 66. The stiffening ribs 68 provide enhanced structural integrity and overall strength to the airfoil, especially with the reduction in mass in the airfoil 25 at cavities 61 and 66 in the outer region 59. Of course, providing more than two cavities would create more stiffening ribs 68, further enhancing the airfoil's integrity.

[0037] like Figure 6 As shown, cavities 61 and 66 in the outer region 59 extend a distance A from the leading edge 20 toward the trailing edge 21. Considering the configuration of the airfoil 25, it is desirable that the distance A does not exceed approximately 60% of the airfoil 25, which should maintain integrity and overall strength in the airfoil 25 adjacent to the trailing edge 21, having a greater solid mass compared to the cavities in the region of the airfoil 25 adjacent to the leading edge 20. Therefore, the outer region 59 has more solid material but still has reduced weight, thereby reducing vibration and load. Various aspects of the embodiment may also include a distance A less than 60% of the airfoil 25.

[0038] According to another aspect of the implementation scheme, each cavity 61 and 66 may have a varying volume. As mentioned above, neither of the bottom walls of cavities 65 and 67 described herein needs to be equal. Furthermore, note the configuration of the airfoil 25; each cavity will have a different perimeter to define the shape and profile of the airfoil. For example, in… Figure 6 In a non-limiting example, cavity 61 defines a first periphery 161 of sidewall 65, and cavity 66 defines a second periphery 166 of sidewall 65. These peripheries 161 and 166 are not necessarily equal, so the areas of the corresponding cavities 61 and 66 may not be equal, even if the bottom wall extends to the same depth in the hollow region 64.

[0039] Additionally, the sidewalls of cavities 61 and 66 may vary along their depths. This is for example and is not intended to limit the implementation in any way. Figure 7Cavity 61 is shown with two different non-limiting embodiments of sidewalls 465 and 466. Sidewall 465 is a rough, irregular sidewall. Sidewall 466 tapers in width from the outer leaf tip 41 toward its bottom wall (omitted for illustration). These sidewall arrangements are simply multiple non-limiting sidewall structures, regardless of the form of the end walls (as described above). In other words, the geometries of cavities 61 and 66 can be different. The volumes of cavities 61 and 66 can be unequal because the depths of cavities 61 and 66 entering the hollow region 64 can vary, and the configuration, orientation, and angle of the sidewalls and bottom walls can vary. The additional peripheral area of ​​each cavity can vary. All these variations allow the volumes of cavities 61 and 66 to be not necessarily equal. However, it is conceivable that if the aforementioned factors determining the cavity volume are chosen to be equal in volume determination, the volumes of cavities 61 and 66 can be the same.

[0040] As embodied in this disclosure, the blade 16 provides an airfoil 25 with a reduced overall weight due to cavities 61 and 66. This reduced overall weight should suppress vibrations and excitation forces on the blade 16 during operation. The blade 16, including cavities 61 and 66, is able to reduce vibrational forces and any loads on the blade 16. Furthermore, the weight-reduced blade 16 provided by cavities 61 and 66 reduces the amplitude of vibrational forces (or natural frequencies), thus resulting in a lower mode shape or deflection at the natural frequency (compared to a blade without cavities 61 and 66). Therefore, degradation of the blade 16 with cavities 61 and 66 can be reduced.

[0041] According to this disclosure, the midspan shields 51 and 51' can be broadly defined as any shield located inside the tip portion 41 of the outer blade of the airfoil 25 and outside the base 121. According to a non-limiting aspect of this disclosure, as stated above, at least two cavities 61 and 66 should be entirely located in the outer region 59 above the midspan shields 51 and 51'.

[0042] According to this disclosure, the midspan shields 51 and 51' may also be disposed within a radial position range on the airfoil 25. According to certain embodiments of this disclosure, the position range of the midspan shields 51 and 51' is defined between an inner boundary of approximately 25% of the radial height of the airfoil 25 and an outer boundary of approximately 75% of the radial height of the airfoil 25. According to other embodiments of this disclosure, as defined by the appended claims, the position range of the midspan shields 51 and 51' is defined between an inner boundary of approximately 33% of the radial height of the airfoil 25 and an outer boundary of approximately 66% of the radial height of the airfoil 25. According to certain embodiments of this disclosure, the midspan shields 51 and 51' are positioned near the approximate radial center region of the airfoil 25.

[0043] It should be understood that, according to the non-limiting and exemplary embodiments discussed above, the vibration response of turbine blades can be reduced to limit destructive mechanical loads, which allows for the extension of rotor blades, thereby enabling greater engine efficiency. Specifically, this disclosure teaches providing cavities 61 and 66 in the turbine blades to limit the vibration response of the outer regions 59 extending beyond one or more midspan shrouds 51 and 51', thus increasing stiffness and reducing the outer portions of the airfoil 25's midspan shrouds 51 and 51'. This avoids harmful vibrations, thereby allowing for longer turbine blades.

[0044] Although this disclosure has been described in conjunction with embodiments that are currently considered to be the most practical and preferred, it should be understood that this disclosure is not limited to the disclosed embodiments, but rather is intended to cover various modifications and equivalent arrangements included within the substance and scope of the appended claims.

Claims

1. A rotor blade (16) for use in a turbine (13) of a gas turbine engine (10), the rotor blade (16) comprising: An airfoil (25) comprising a concave pressure sidewall (26) and a convex suction sidewall (27), the concave pressure sidewall and the convex suction sidewall extending axially between corresponding leading edges (20) and trailing edges (21) and extending radially between a base (121) and an outer blade tip (41), the rotor blade (16) further comprising: At least one mid-span guard, the at least one mid-span guard being configured to engage a corresponding mid-span guard on at least one adjacent rotor blade (16) during operation; The airfoil (25) also includes an inner region (58) between the at least one mid-span shield and the base (121) of the blade (216), wherein the inner direction of the airfoil (25) faces the base (121). The airfoil (25) further includes an outer region (59) between the at least one mid-span shield and the outer tip of the blade (216), wherein the outer direction of the airfoil (25) is toward the outer tip of the blade. in: The outer region (59) includes at least two cavities (61, 66) extending from the tip (41) of the outer blade inside the airfoil (25) toward the at least one midspan shield, wherein the at least two cavities (61, 66) include sidewalls (65, 465, 466) and a bottom wall (265), each corresponding bottom wall (265) of the at least two cavities (61, 66) forming different angles relative to each other and the sidewalls (65, 465, 466); and Inside the at least one transverse shield, the inner region (58) is substantially solid.

2. The rotor blade (16) according to claim 1, wherein each of the at least two cavities (61, 66) is open at the tip (41) of the outer blade and closed at its inner end, and the bottom wall (265) extends equidistantly into the airfoil (25).

3. The rotor blade (16) according to claim 1, wherein the bottom wall (265) extends into the airfoil (25) by different distances.

4. The rotor blade (16) according to claim 1, wherein the bottom wall (265) is not orthogonal to the side walls (65, 465, 466).

5. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) comprise varying cross-sections.

6. The rotor blade (16) according to claim 5, wherein the varying cross-section is formed by sidewalls (65, 465, 466) varying in depth along the cavity (61, 66).

7. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) are open at the tip (41) of the outer blade and closed at its inner end.

8. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) comprise more than two cavities (61, 66) formed along the chord of the blade (216).

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

10. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) define stiffening ribs (68) therebetween.

11. The rotor blade (16) according to claim 1, wherein the at least two cavities (61, 66) are entirely located in the outer region (59) above the at least one mid-span shroud.

12. The rotor blade (16) according to claim 1, wherein the outer region (59) has a hollow percentage of at least 40%, and wherein the inner region (58) has a solid percentage of at least 60%.

13. A gas turbine engine (10), the gas turbine engine (10) comprising: Rotor blade (16), the rotor blade including airfoil (25), the airfoil (25) including concave pressure sidewall (26) and convex suction sidewall (27), the concave pressure sidewall and the convex suction sidewall extending axially between corresponding leading edge and trailing edge (21); base (121); outer blade tip portion (41), the rotor blade (16) being the rotor blade (16) according to any one of claims 1 to 12.

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