JP7389574B2 - aircraft gas turbine - Google Patents

Description

The present invention relates to an aircraft gas turbine.

Patent Document 1 discloses an aircraft gas turbine that includes rotor blade stages and stator vane stages alternately in the axial direction. The rotor blade stage of this patent document includes a plurality of rotor blades arranged at intervals in the circumferential direction. Each of the plurality of rotor blades is provided with a tip shroud. These circumferentially adjacent tip shrouds are connected to form an annular shape.

Japanese Patent Application Publication No. 2011-174419

In aircraft gas turbines such as those described in Patent Document 1, there is a demand for further weight reduction. Regarding the shape of the rotor blades of an aircraft gas turbine, generally, when the ratio between the blade pitch and the blade chord length is maintained, the larger the number of rotor blades per one rotor blade stage, the narrower the shape of the rotor blade. Therefore, by increasing the number of rotor blades per rotor blade stage, the weight of the rotor blade stage of an aircraft gas turbine can be reduced.
However, the trailing edge loss of the rotor blade increases as the number of rotor blades increases. Furthermore, from the viewpoint of strength against vibration, it is advantageous to reduce the number of rotor blades because the airfoil can be made thicker.
In other words, there is a trade-off relationship between the number of rotor blades (in other words, the weight) and the performance and vibration strength of the aircraft gas turbine, making it difficult to further reduce the weight of the aircraft gas turbine.

The present invention has been made in view of the above circumstances, and provides an aircraft gas turbine that can further reduce the weight while ensuring the strength of the rotor blades against vibration.

In order to solve the above problems, the following configuration is adopted.
According to a first aspect of the invention, an aircraft gas turbine has a takeoff thrust of 15,000 to 40,000 lbs, and includes a rotor shaft that rotates around an axis, a turbine rotor blade stage, and a turbine casing. , a first upstream stator vane stage and a downstream stator vane stage. The turbine rotor blade stage extends from the rotor shaft toward the outside in a radial direction centered on the axis, and has a plurality of turbine rotor blades arranged in line in a circumferential direction centered on the axis. The turbine casing covers the rotor shaft from the outer peripheral side. The first upstream stator vane stage is arranged upstream of the turbine rotor blade stage in the axial direction. The first upstream stator blade stage is a turbine stator blade stage that includes a plurality of turbine stator blade bodies. The plurality of stator vane bodies of the first upstream stator vane stage extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. The turbine rotor blade includes a rotor blade body having an airfoil cross section extending in the radial direction, and a shroud provided at a tip of the rotor blade body. The turbine rotor blade stage includes shrouds of the turbine rotor blades that are adjacent to each other in the circumferential direction, and includes 50 or more and 72 or less rotor blade bodies. The absolute value of the difference between the number of stator blade bodies in the first upstream stator blade stage and the number of stator blade bodies in the downstream stator blade stage is 1 or more and 3 or less.

In gas turbines for small and medium-sized aircraft, which have so-called integral shroud blades in which the shrouds of the rotor blades are connected to each other in the circumferential direction and have a thrust of 15,000 to 40,000 lbs at takeoff, the rotor blade body of the rotor blade stage has 50 or more blades. In addition, by setting the number of blades to 72 or less, it is possible to avoid resonance of the natural vibration frequency of the primary mode among the vibration modes of the rotor blade at a low rotational order in the main operating range. Generally, when the natural vibration frequency of the first mode is set low, resonance occurs at a low rotational order, and vibration stress tends to increase. However, in the first aspect, since the natural vibration frequency of the first mode can be made higher in the main operating range, resonance at low rotational orders of 3H (harmonics) or less is avoided and it acts on the rotor blades in the main operating range. vibration stress can be reduced. Furthermore, by setting the absolute value of the difference between the number of stator vane bodies in the first upstream stator vane stage and the number of stator vane bodies in the downstream stator vane stage to be 1 or more and 3 or less, fluid turbulence can be reduced. It is possible to prevent vibration from being excited by suppressing it. Therefore, by reducing vibration stress and avoiding resonance, it is possible to further reduce the weight while ensuring the strength of the moving blade against vibration.

According to a second aspect of the invention, in the aircraft gas turbine according to the first aspect, when the height of the rotor blade body is "H" and the shaft cord length of the rotor blade body is "Cx", H/ The value of Cx may be 2 or more and 6 or less.
With this configuration, in a rotor blade that is a so-called integral shroud blade, it is possible to suppress an increase in the size and weight of the rotor blade body.

According to a third aspect of the present invention, an aircraft gas turbine has a takeoff thrust of 15,000 to 40,000 lbs, and includes a rotor shaft that rotates around an axis, a turbine rotor blade stage, and a casing. It includes a first upstream stator vane stage and a downstream stator vane stage. The rotor blade stage extends from the rotor shaft toward the outside in a radial direction about the axis, and has a plurality of turbine rotor blades arranged in parallel in a circumferential direction about the axis. The turbine casing covers the rotor shaft from the outer peripheral side. The first upstream stator vane stage is arranged upstream of the turbine rotor blade stage in the axial direction. The first upstream stator vane stage is a turbine stator vane stage having a plurality of stator vane bodies. The plurality of stator vane bodies of the first upstream stator vane stage extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. The downstream stationary blade stage is arranged downstream of the turbine rotor blade stage in the axial direction. The downstream stationary blade stage is a turbine stationary blade stage that extends inward in the radial direction from the turbine casing and has a plurality of stationary blade bodies arranged in a line in a circumferential direction centered on the axis. The turbine rotor blade includes a rotor blade body having an airfoil cross section extending in the radial direction, and a shroud provided at a tip of the rotor blade body. Assuming that the height of the rotor blade body is "H" and the shaft cord length of the rotor blade body is "Cx", the value of H/Cx is 2 or more and 6 or less. The absolute value of the difference between the number of stator blade bodies in the first upstream stator blade stage and the number of stator blade bodies in the downstream stator blade stage is 1 or more and 3 or less.

In gas turbines for small and medium-sized aircraft, which have so-called integral shroud blades in which the shrouds of rotor blades are connected to each other in the circumferential direction and have a thrust of 15,000 to 40,000 lbs at takeoff, the stator blade body of the first upstream stator blade stage. By setting the absolute value of the difference between the number of blades and the number of stator blade bodies of the downstream stator blade stage to be 1 or more and 3 or less, the flow of combustion gas in the gas turbine is prevented from being disturbed and vibrations are excited. can. Moreover, by setting the value of H/Cx to be 2 or more and 6 or less, it is possible to suppress the rotor blade main body 32 from increasing in size and weight. Therefore, since the vibration stress is reduced, it is possible to further reduce the weight while ensuring the strength of the moving blade against vibration.

According to a fourth aspect of the present invention, the aircraft gas turbine according to any one of the first to third aspects includes an upstream rotor blade stage as the turbine rotor blade stage , and an upstream rotor blade stage as the turbine stationary blade stage. and a second upstream stator vane stage. The upstream moving blade stage is arranged upstream of the first upstream stator vane stage in the axial direction. The second upstream stationary blade stage is arranged upstream of the upstream rotor blade stage in the axial direction. The absolute value of the difference between the number of stator blade bodies of the first upstream stator vane stage and the number of stator blade bodies of the second upstream stator vane stage may be 1 or more and 3 or less. good.
With this configuration, vibration is excited due to a large difference in the number of stator blade bodies between the first upstream stator vane stage and the second upstream stator vane stage, which are adjacent stator vane stages in the axial direction. This can be suppressed.

According to the fifth aspect of the invention, the turbine rotor blade according to any one of the first to fourth aspects may be made of a TiAl alloy.
By doing so, it is possible to further reduce the weight of the rotor blade while ensuring vibration strength.

According to the sixth aspect of the invention, the moving blade main body according to any one of the first to fifth aspects may be a hollow blade.
By doing so, the weight of the rotor blade can be further reduced compared to a rotor blade formed solidly.

According to the seventh aspect of the present invention, the rotor blade body according to any one of the first to fifth aspects may be formed so that at least a portion in the blade height direction is hollow.
By doing so, the weight of the rotor blade can be further reduced compared to a rotor blade formed solidly.

According to an eighth aspect of the invention, the aircraft gas turbine is an aircraft gas turbine having a take-off thrust of 15,000 to 40,000 lbs. The aircraft gas turbine includes a rotor shaft, a turbine rotor blade stage, a turbine casing, a first upstream stator blade stage, an upstream rotor blade stage, and a second upstream stator vane stage. The rotor shaft rotates around the axis. The turbine rotor blade stage extends from the rotor shaft toward the outside in a radial direction centered on the axis, and has a plurality of turbine rotor blades arranged in line in a circumferential direction centered on the axis. The turbine casing covers the rotor shaft from the outer peripheral side. The first upstream stationary blade stage is disposed upstream of the turbine rotor blade stage in the axial direction, extends from the turbine casing toward the inside in the radial direction, and is arranged in a circumferential direction around the axis. This is a turbine stator vane stage having a plurality of stator vane bodies arranged. The upstream rotor blade stage is a turbine rotor blade stage disposed upstream in the axial direction of the first upstream stator vane stage. The second upstream stator vane stage is a turbine stator vane stage disposed upstream in the axial direction of the upstream rotor blade stage. The turbine rotor blade includes a rotor blade body having an airfoil cross section extending in the radial direction, and a shroud provided at a tip of the rotor blade body. The turbine rotor blade stage includes shrouds of the turbine rotor blades that are adjacent to each other in the circumferential direction, and includes 50 or more and 72 or less rotor blade bodies. The absolute value of the difference between the number of stator blade bodies in the first upstream stator blade stage and the number of stator blade bodies in the second upstream stator blade stage is 1 or more and 3 or less.

According to a ninth aspect of the invention, the aircraft gas turbine is an aircraft gas turbine having a take-off thrust of 15,000 to 40,000 lbs. The aircraft gas turbine includes a rotor shaft, a turbine rotor blade stage, a turbine casing, a first upstream stator blade stage, an upstream rotor blade stage, and a second upstream stator vane stage. The rotor shaft rotates around the axis. The turbine rotor blade stage extends from the rotor shaft toward the outside in a radial direction centered on the axis, and has a plurality of turbine rotor blades arranged in line in a circumferential direction centered on the axis. The turbine casing covers the rotor shaft from the outer peripheral side. The first upstream stationary blade stage is disposed upstream of the turbine rotor blade stage in the axial direction, extends from the turbine casing toward the inside in the radial direction, and is arranged in a circumferential direction around the axis. It has a plurality of stator vane bodies arranged. The upstream rotor blade stage is a turbine rotor blade stage disposed upstream in the axial direction of the first upstream stator vane stage. The second upstream stator vane stage is a turbine stator vane stage disposed upstream in the axial direction of the upstream rotor blade stage. The turbine rotor blade includes a rotor blade body having an airfoil cross section extending in the radial direction, and a shroud provided at a tip of the rotor blade body. Assuming that the height of the rotor blade body is "H" and the shaft cord length of the rotor blade body is "Cx", the value of H/Cx is 2 or more and 6 or less. The absolute value of the difference between the number of stator blade bodies in the first upstream stator blade stage and the number of stator blade bodies in the second upstream stator blade stage is 1 or more and 3 or less.

According to the above-mentioned aircraft gas turbine, it is possible to further reduce the weight while ensuring the strength of the rotor blade against vibration.

1 is a configuration diagram showing a schematic configuration of an aircraft gas turbine according to an embodiment of the present invention. FIG. 2 is an enlarged view of a turbine rotor blade stage and a turbine stationary blade stage in FIG. 1; 1 is a diagram showing a schematic configuration of a turbine rotor blade in an embodiment of the present invention. FIG. 3 is a diagram for explaining the cord length of the rotor blade body. It is a Campbell diagram in which the vertical axis represents vibration frequency (Frequency), the horizontal axis represents rotational speed (Rotational Speed), and the oblique axis represents rotational order (H). FIG. 4 is a diagram corresponding to FIG. 3 in a modification of the embodiment of the present invention.

(Embodiment)
Next, an aircraft gas turbine according to an embodiment of the present invention will be described based on the drawings.
FIG. 1 is a configuration diagram showing a schematic configuration of an aircraft gas turbine according to an embodiment of the present invention. FIG. 2 is an enlarged view of the turbine rotor blade stage and the turbine stationary blade stage of FIG. 1.
The aircraft gas turbine 100 according to this embodiment is a gas turbine for small and medium-sized aircraft. This aircraft gas turbine 100 is a gas turbine for obtaining takeoff thrust (for example, about 15,000 to 40,000 lbs) for small and medium-sized aircraft. As shown in FIG. 1, this aircraft gas turbine 100 mainly includes a compressor 1, a combustion chamber 2, and a turbine 3.

The compressor 1 generates high-pressure air by compressing air taken in from the intake duct 10. The compressor 1 includes a compressor rotor shaft 11 , a compressor casing 12 , a compressor moving blade stage 13 , and a compressor stator vane stage 15 . The compressor casing 12 covers the compressor rotor shaft 11 from the outer peripheral side, and extends in the direction in which the axis Am extends (hereinafter referred to as the axial direction Da).

A plurality of compressor rotor stages 13 are provided on the compressor rotor shaft 11 . These compressor blade stages 13 are arranged at intervals in the axial direction Da. Each of the plurality of compressor rotor blade stages 13 includes a plurality of compressor rotor blades 14 . The compressor rotor blades 14 of each compressor rotor blade stage 13 are arranged on the outer peripheral surface of the compressor rotor shaft 11 in a direction centered on the axis Am (hereinafter referred to as the circumferential direction Dc).

A plurality of compressor stator vane stages 15 are provided in the compressor casing 12. These compressor stator vane stages 15 are arranged at intervals in the axial direction Da. The compressor stator blade stages 15 are arranged alternately with the compressor rotor blade stages 13 in the axial direction Da. Each of the plurality of compressor stator blade stages 15 includes a plurality of compressor stator blades 16. The compressor stator vanes 16 of each compressor stator vane stage 15 are arranged on the inner peripheral surface of the compressor casing 12 in the circumferential direction Dc.

The combustion chamber 2 generates combustion gas G by mixing fuel F with high-pressure air generated by the compressor 1 and combusting the mixture. Combustion chamber 2 is provided between compressor casing 12 and turbine casing (casing) 22 of turbine 3 . Combustion gas G generated by this combustion chamber 2 is supplied to a turbine 3.

The turbine 3 is driven by high-temperature, high-pressure combustion gas G generated in the combustion chamber 2 . More specifically, the turbine 3 expands the high-temperature, high-pressure combustion gas G and converts the thermal energy of the combustion gas G into rotational energy. This turbine 3 includes a turbine rotor shaft (rotor shaft) 21, a turbine rotor blade stage (rotor blade stage) 23, a turbine casing (casing) 22, and a turbine stationary blade stage (stator blade stage) 25. There is.

The turbine rotor shaft 21 extends in the axial direction Da. This turbine rotor shaft 21 and the above-mentioned compressor rotor shaft 11 are lined up in the axial direction Da and are immovable relative to each other. These turbine rotor shaft 21 and compressor rotor shaft 11 constitute a gas turbine rotor 91. This gas turbine rotor 91 can rotate integrally around an axis Am inside a gas turbine casing 92.

A plurality of turbine rotor blade stages 23 are provided on the outer peripheral surface of the turbine rotor shaft 21 at intervals in the axial direction Da. These plurality of turbine rotor blade stages 23 each have a plurality of turbine rotor blades (rotor blades) 24 (details will be described later). The plurality of turbine rotor blades 24 included in one turbine rotor blade stage 23 are arranged in line at equal pitches in the circumferential direction Dc. The turbine rotor blades 24 constituting the turbine rotor blade stage 23 can be made of, for example, a TiAl (titanium-aluminum) alloy.

The turbine casing 22 covers the turbine rotor shaft 21 from the outer peripheral side. The turbine casing 22 and the compressor casing 12 described above are integrally connected along the axis Am. The compressor casing 12 and the turbine casing 22 constitute a gas turbine casing 92.

A plurality of turbine stationary blade stages 25 are provided on the inner peripheral surface of the turbine casing 22 at intervals in the axial direction Da. These plurality of turbine stationary blade stages 25 are arranged alternately with the turbine rotor blade stages 23 in the axial direction Da. Each of these turbine stator blade stages 25 includes a plurality of turbine stator blades (stator vanes) 26 . The turbine stator blades 26 provided in each turbine stator blade stage 25 are arranged on the inner peripheral surface of the turbine casing 22 at equal pitches in the circumferential direction Dc.

As shown in FIG. 2, this embodiment illustrates a case where the turbine 3 includes three turbine rotor blade stages 23 and three turbine stationary blade stages 25, respectively. In this turbine 3, the main stream of combustion gas G flows inside the turbine casing 22 from left to right in FIG. Therefore, in the following description, the left side of FIG. 2 may be referred to as the upstream side, and the right side may be referred to as the downstream side.

In this embodiment, the turbine rotor blade stages 23 include, in order from the upstream side, a first turbine rotor blade stage 23A, a second turbine rotor blade stage 23B, and a third turbine rotor blade stage 23C. Similarly, in this embodiment, the turbine stator vane stage 25 includes, in order from the upstream side, a first turbine stator vane stage 25A, a second turbine stator vane stage 25B, and a third turbine stator vane stage 25C. The first turbine stator blade stage 25A is arranged upstream of the first turbine rotor blade stage 23A, the second turbine stator blade stage 25B is arranged upstream of the second turbine rotor blade stage 23B, and the third turbine stator blade stage 25B is arranged upstream of the second turbine rotor blade stage 23B. The blade stage 25C is arranged upstream of the third turbine blade stage 23C. In addition, in the following description, the first turbine rotor blade stage 23A, the second turbine rotor blade stage 23B, and the third turbine rotor blade stage 23C may be collectively referred to as the turbine rotor blade stage 23.

In operating the aircraft gas turbine 100 configured as described above, first, the compressor rotor shaft 11 (gas turbine rotor 91) is rotationally driven by an external drive source. As the compressor rotor shaft 11 rotates, external air is sequentially compressed to generate high-pressure air. This high pressure air is supplied into the combustion chamber 2 through the compressor casing 12. In the combustion chamber 2, this high-pressure air is mixed with fuel and then combusted, producing high-temperature, high-pressure combustion gas. Combustion gas G is supplied into the turbine 3 through the turbine casing 22.

In the turbine 3, the combustion gas G sequentially collides with the turbine rotor blade stage 23 and the turbine stationary blade stage 25, thereby providing rotational driving force to the turbine rotor shaft 21 (gas turbine rotor 91). This rotational energy is mainly used to drive the compressor 1. The combustion gas G that has driven the turbine 3 has its flow velocity increased by the exhaust nozzle 4 to become a jet that generates thrust, and is discharged to the outside from the injection port 27. In this embodiment, a single-shaft turbojet engine has been described as an example of an aircraft gas turbine. However, the present invention is not limited to a single-shaft turbojet engine, and may be any type of aircraft gas turbine. In particular, it is suitable for a low pressure turbine of a multi-shaft turbofan engine.

FIG. 3 is a diagram showing a schematic configuration of a turbine rotor blade in an embodiment of the present invention.
As shown in FIG. 3, the turbine rotor blade 24 of the turbine rotor blade stage 23 described above is a so-called integral shroud blade (ISB), and includes a rotor blade body 32 and a shroud 33.
The rotor blade body 32 has an airfoil-shaped cross section having a pressure surface and a suction surface, and extends in the radial direction Dr. The rotor blade body 32 illustrated in this embodiment extends from the platform 34 of the turbine rotor blade 24 toward the outside in the radial direction Dr. The turbine rotor blade 24 is fixed to the turbine rotor shaft 21 via a blade root (not shown) formed closer to the axis Am than the platform 34 .

The shroud 33 is provided at the tip of the rotor blade body 32. The shroud 33 extends in the circumferential direction Dc and the axial direction Da. The shrouds 33 of the turbine rotor blades 24 adjacent in the circumferential direction Dc are connected to form, for example, an annular shape centered on the axis Am.

On the other hand, the turbine stator blade 26 of the turbine stator blade stage 25 has a stator blade body 42 (see FIG. 2). These stationary blade bodies 42 extend inward from the turbine casing 22 in the radial direction Dr. Since the turbine stator blades 26 are arranged side by side in the circumferential direction Dc, the plurality of stator blade bodies 42 are arranged side by side in the circumferential direction Dc.

The rotor blade main body 32 of the turbine rotor blade 24 provided in the turbine rotor blade stage 23 has a height (in other words, blade height) of "H" and a shaft cord length of "Cx", and has an aspect ratio of H/Cx. ) is set to be 2 or more and 6 or less. The value of H/Cx can also be set to 4 or more and 6 or less. In this embodiment, the axial chord length Cx means the average value of the axial chord lengths in the blade height direction of the rotor blade body 32.

FIG. 4 is a diagram for explaining the cord length of the rotor blade body.
Here, the axial cord length C1 shown in FIG. 4 is the axial cord length of the blade tip. This axial cord length C1 is the length from the leading edge 32f to the trailing edge 32e of the rotor blade body 32 in the axial direction Da. The numbers "100" and "90" shown in FIG. 4 indicate the positions of the leading edge 32f and the trailing edge 32e in the blade height direction in percentage (%). That is, the axial chord length C1 is the axial chord length at the 100% height position, which is the blade tip of the leading edge 32f, and at the 100% height position, which is the blade tip of the trailing edge 32e. Similarly, the shaft cord length C2 is the shaft cord length at the 90% height position of the leading edge 32f and the 90% height position of the trailing edge 32e. Note that the shaft cord length Cx may be the shaft cord length between the 50% height position of the leading edge 32f and the 50% height position of the trailing edge 32e.

Further, the height (H) of the rotor blade body 32 means the length in the radial direction Dr of the portion of the rotor blade body 32 that is exposed to the main flow path through which the combustion gas G flows. This height (H) also means the average value of the height of the rotor blade body 32 from the leading edge 32f to the trailing edge 32e, similarly to the shaft cord length (Cx). Note that the height (H) may be the height at the center of the leading edge 32f and the trailing edge 32e in the direction in which the camber line (not shown) of the airfoil of the rotor blade body 32 extends.

As shown in FIG. 3, the rotor blade main body 32 is formed in a hollow shape with a cavity c1 therein. The rotor blade main body 32 in this embodiment is a so-called hollow blade having a cavity c1 in a range from the platform 34 to the blade tip (in other words, the inner surface of the shroud 33 in the radial direction Dr).

Each turbine rotor blade stage 23 includes 50 or more rotor blade bodies 32 and 72 or less rotor blade bodies 32 . Specifically, the first turbine rotor blade stage 23A includes 50 or more rotor blade bodies 32 and 72 or less rotor blade bodies 32. Similarly, the second turbine rotor blade stage 23B also includes 50 or more rotor blade bodies 32 and 72 or less rotor blade bodies 32. Further, the third turbine rotor blade stage 23C also includes 50 or more rotor blade bodies 32 and 72 or less rotor blade bodies 32.

FIG. 5 is a Campbell diagram in which the vertical axis is vibration frequency (Frequency), the horizontal axis is rotational speed (Rotational Speed), and the oblique axis is rotational order (H; harmonics).
In FIG. 5, the natural vibration frequencies of the primary mode of the turbine rotor blade 24 are shown by solid lines when the number of rotor blade bodies 32 provided in one turbine rotor blade stage 23 is 72, 60, and 50, respectively. There is. Further, in FIG. 5, the rotational speed (rpm) in the main operating range of the aircraft gas turbine 100 whose thrust at takeoff is 15,000 to 40,000 lbs is indicated by hatching.

Here, the ratio of the blade pitch to the shaft chord length of each rotor blade body 32 is constant regardless of the number of rotor blade bodies 32. In other words, as the number of rotor blade bodies 32 increases, the rotor blade body 32 becomes thinner.

In FIG. 5, when the rotational speed reaches a point where the natural vibration frequency of the turbine rotor blade 24 and the oblique axis of the rotational order intersect, resonance tends to occur and the amplitude of vibration tends to increase. Furthermore, the amplitude of vibration during resonance tends to increase as the rotational order becomes lower. Therefore, the higher the natural vibration frequency of the turbine rotor blade 24, the smaller the amplitude of vibration can be.

As shown in FIG. 5, when there are 72 rotor blade bodies 32, the natural vibration frequency of the first mode is 3H (harmonics) or more in the main operating range. Furthermore, when the number of rotor blade bodies 32 is reduced to 60, the natural vibration frequency of the primary mode becomes higher overall than when the number of rotor blade bodies 32 is 72. If the number of rotor blade bodies 32 is further reduced to 50, the natural vibration frequency of the primary mode will be higher overall than in the case of 60 blades, and will be 4H or more in the main operating range.

On the other hand, if the number of rotor blade bodies 32 is 49 or less, the rotor blade body 32 becomes large and the above-described value of H/Cx may not be within the range of 2 or more and 6 or less. In other words, this means an increase in the weight of the turbine blade stage 23. On the other hand, when the rotor blade body 32 has 73 or more blades, there is a possibility that the natural vibration frequency of the rotor blade body 32 intersects the 3H (harmonics) straight line in the main operating range and resonates at a low rotational order. .

As described above, the turbine stationary blade stages 25 are arranged upstream and downstream of the turbine rotor blade stages 23. More specifically, a first turbine rotor blade stage (second upstream stator vane stage) 25A is arranged upstream of the first turbine rotor blade stage 23A, and upstream of the second turbine rotor blade stage 23B (and A second turbine rotor blade stage (first upstream stator vane stage) 25B is arranged downstream of the first turbine rotor blade stage 23A), and a second turbine rotor blade stage 25B is disposed upstream of the third turbine rotor blade stage 23C (and downstream of the second turbine rotor blade stage 23A). A third turbine stator vane stage (downstream stator vane stage) 25C is arranged downstream of the stage 23B.

The absolute value of the difference in the number of stator blade bodies 42 between adjacent turbine stator blade stages 25 in the axial direction Da is greater than or equal to 1 and less than or equal to 3. In this embodiment, the number of stator blade bodies 42 of the second turbine stator blade stage 25B disposed upstream of the second turbine rotor blade stage 23B and the number of stator blade bodies 42 disposed downstream of the second turbine rotor blade stage 23B are determined. The absolute value of the difference from the number of stator blade bodies 42 of the third turbine stator blade stage 25C is greater than or equal to 1 and less than or equal to 3.

Furthermore, in this embodiment, the number of stator blade bodies 42 of the first turbine stator blade stage 25A disposed upstream of the first turbine stator blade stage 23A and the stator blades of the second turbine stator blade stage 25B are The absolute value of the difference from the number of sheets in the main body 42 is greater than or equal to 1 and less than or equal to 3. Here, the absolute value of the difference between the number of stator blade bodies 42 of the first turbine stator blade stage 25A and the number of stator blade bodies 42 of the second turbine stator blade stage 25B may be 3 or less, but may be 3 or less. It is not limited to the following.

Therefore, according to the embodiment described above, in the gas turbine 100 for small and medium-sized aircraft having an integral shroud blade and having a thrust of 15,000 to 40,000 lbs at takeoff, the rotor blade body 32 of the turbine rotor blade stage 23 has 50 or more blades. In addition, by setting the number of blades to 72 or less, the primary mode among the vibration modes of the turbine rotor blade 24 can be made to be 3H (harmonics) or higher in the main operating range. When the natural vibration frequency of the first mode is set low, resonance occurs at a low rotational order and vibration stress tends to increase. However, since the natural vibration frequency of the first mode can be made higher than 3H (harmonics) in the main operating range, it is possible to avoid resonance below 3H (harmonics) and reduce the vibration stress acting on the turbine rotor blade 24 in the main operating range. . Furthermore, the absolute value of the difference between the number of stator blade bodies 42 of the second turbine stator blade stages 25B adjacent to each other in the axial direction Da and the number of stator blade bodies 42 of the third turbine stator blade stage 25C is 1 or more, and By setting it to 3 or less, disturbance of the combustion gas G can be suppressed and vibrations can be prevented from being excited. As a result, by reducing vibration stress and avoiding resonance, it is possible to further reduce the weight while ensuring the strength of the turbine rotor blade 24 against vibration.

Furthermore, by setting the value of H/Cx to 2 or more and 6 or less, it is possible to suppress the increase in size and weight of the rotor blade body 32 in the turbine rotor blade 24, which is an integral shroud blade, and also to prevent the primary mode can be made 3H or more. As a result, further weight reduction can be achieved.

In the above embodiment, the first turbine rotor blade stage 23A is provided upstream of the second turbine rotor blade stage 25B, and the first turbine rotor blade stage 25A is provided upstream of the first turbine rotor blade stage 23A. The absolute value of the difference between the number of stator blade bodies 42 of the second turbine stator blade stage 25B and the number of stator blade bodies 42 of the first turbine stator blade stage 25A is 1 or more and 3 or less. . Therefore, the difference in the number of stator blade bodies 42 between the first turbine stator blade stage 25A and the second turbine stator blade stage 25B, which are the turbine stator blade stages 25 adjacent to each other in the axial direction Da, increases, and vibrations are excited. can be suppressed.

Furthermore, since the turbine rotor blade 24 is formed of a TiAl alloy, it is possible to further reduce the weight of the turbine rotor blade 24 while ensuring vibration strength.
Further, since the rotor blade body 32 is a hollow blade, the weight of the turbine rotor blade 24 can be further reduced compared to a case where the rotor blade body is formed solid.

(Modified example of embodiment)
This invention is not limited to the embodiments described above, but includes various modifications to the embodiments described above without departing from the spirit of the invention. That is, the specific shapes, configurations, etc. mentioned in the embodiments are merely examples, and can be changed as appropriate.

In the above embodiment, one turbine rotor blade stage 23 is provided with rotor blade bodies 32 in the range of 50 to 72 blade bodies, and the stator blade bodies 42 of the second turbine stator blade stages 25B adjacent in the axial direction Da are provided. The case where the absolute value of the difference between the number of stator blade bodies 42 and the number of stator blade bodies 42 of the third turbine stator blade stage 25C is 1 or more and 3 or less has been described. However, the configuration is not limited to the above embodiment. For example, in a gas turbine 100 for a so-called small and medium-sized aircraft with a takeoff thrust of 15,000 to 40,000 lbs, the turbine rotor blades 24 are integral shroud blades, and the aspect ratio of the rotor blade body 32, H/Cx, is 2 or more. and 6 or less, the absolute value of the difference between the number of stator blade bodies 42 of the second turbine stator blade stage 25B and the number of stator blade bodies 42 of the third turbine stator blade stage 25C is 1 or more, and 3 The following may be used.

In the above embodiment, all the rotor blade stages 23, the first turbine rotor blade stage 23A, the second turbine rotor blade stage 23B, and the third turbine rotor blade stage 23C, have rotor blade bodies 32 in the range of 50 to 72 rotor blade bodies 32. We have explained how to prepare for this. However, the first turbine rotor blade stage 23A and the third turbine rotor blade stage 23C do not need to include rotor blade bodies 32 in the range of 50 to 72 blade bodies. Furthermore, although a case has been described in which the first turbine rotor blade stage 23A, the second turbine rotor blade stage 23B, and the third turbine rotor blade stage 23C are provided, the first turbine rotor blade stage 23A and the third turbine rotor blade stage 23C are omitted. You may.

Furthermore, in the above embodiment, the absolute value of the difference between the number of stator blade bodies 42 of the second turbine stator blade stage 25B and the number of stator blade bodies 42 of the third turbine stator blade stage 25C is 1 or more, and 3 The following cases have been explained. However, the configuration is not limited to the above embodiment.
For example, as another aspect, in a gas turbine 100 for a so-called small and medium-sized aircraft with a takeoff thrust of 15,000 to 40,000 lbs, the turbine rotor blade 24 is an integral shroud blade, and the third turbine rotor blade stage (rotor blade stage) 23C has 50 or more and 72 or less rotor blade bodies 32, and the difference between the number of stator blade bodies 42 of the first turbine stator blade stage 25A and the number of stator blade bodies 42 of the second turbine stator blade stage 25B. The absolute value of may be greater than or equal to 1 and less than or equal to 3. In this case, the number of rotor blade bodies 32 of the second turbine rotor blade stage 23B is not limited to 50 or more and 72 or less.

Furthermore, in the other aspect described above, the third turbine rotor blade stage 23C includes 50 or more and 72 or less rotor blade bodies 32, but in the other aspect described above, the third turbine rotor blade stage 23C includes Instead of having the rotor blade body 32 of 50 or more and 72 or less, the value of H/Cx, which is the aspect ratio of the rotor blade body 32 provided in the third turbine rotor blade stage 23C, is set to be 2 or more and 6 or less. You may also adopt a configuration in which: Also in this case, the value of H/Cx, which is the aspect ratio of the rotor blade body 32 of the second turbine rotor blade stage 23B, is not limited to 2 or more and 6 or less.

FIG. 6 is a diagram corresponding to FIG. 3 in a modification of the embodiment of the present invention.
In the embodiment described above, the turbine rotor blade 24 is a hollow blade in which the cavity C is formed over the entire blade height direction. However, the turbine rotor blade 24 is not limited to a hollow blade.
For example, as shown in FIG. 6, a cavity c2 may be formed in a portion of the turbine rotor blade 24 in the blade height direction. By doing so, the weight of the turbine rotor blade 24 can be further reduced compared to a case where the rotor blade body is formed solid. Note that if at least a portion of the rotor blade main body 32 in the blade height direction is formed hollow, the position, shape, and size of the cavity c2 are limited to the position, shape, and size illustrated in FIG. I can't do it.

In the above embodiment, a case has been described in which three turbine rotor blade stages 23 and three turbine stationary blade stages 25 are provided. However, the number of turbine blade stages 23 provided is not limited to three. For example, two or only one turbine rotor blade stage 23 may be provided. Further, the number of turbine stator blade stages 25 provided is not limited to three, similarly to the turbine rotor blade stages 23.

In the above embodiment, the turbine rotor blade 24 is made of a TiAl alloy, but the turbine rotor blade 24 may be made of a material other than the TiAl alloy.
In the above embodiment, the absolute value of the difference in the number of stator blade bodies 42 of adjacent turbine stator blade stages 25 in the axial direction Da is 3 or less, but the absolute value of this difference may be larger than 3. You can.

1 Compressor 2 Combustion chamber 3 Turbine 4 Exhaust nozzle 10 Intake duct 11 Compressor rotor shaft 12 Compressor casing 13 Compressor rotor stage 14 Compressor rotor blade 15 Compressor stator vane stage 16 Compressor stator vane 21 Turbine rotor shaft (rotor shaft )
22 Turbine casing (casing)
23 Turbine rotor blade stage 23A First turbine rotor blade stage (upstream rotor blade stage)
23B Second turbine rotor blade stage (rotor blade stage)
23C Third turbine rotor blade stage (rotor blade stage)
24 Turbine moving blade (moving blade)
25 Turbine stator blade stage (Stator blade stage)
25A First turbine stator vane stage (second upstream stator vane stage)
25B Second turbine stator vane stage (first upstream stator vane stage)
25C Third turbine stator vane stage (downstream stator vane stage)
26 Turbine stator blade 27 Injection port 32 Rotor blade body 32e Trailing edge 32f Leading edge 33 Shroud 34 Platform 42 Stator blade body 91 Gas turbine rotor 92 Gas turbine casing 100 Aircraft gas turbine

Claims (9)

An aircraft gas turbine with a takeoff thrust of 15,000 to 40,000 lbs,
A rotor shaft that rotates around the axis,
a turbine rotor blade stage extending from the rotor shaft toward the outside in a radial direction about the axis and having a plurality of turbine rotor blades arranged in parallel in a circumferential direction about the axis;
a turbine casing that covers the rotor shaft from the outer peripheral side;
A plurality of stationary blade bodies are arranged on the upstream side in the axial direction of the turbine rotor blade stage, extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. a first upstream stator vane stage which is a turbine stator vane stage having;
A plurality of stationary blade bodies are arranged downstream of the turbine rotor blade stage in the axial direction, extend from the turbine casing toward the inside in the radial direction, and are arranged side by side in a circumferential direction centered on the axis. a downstream stator vane stage that is a turbine stator vane stage,
The turbine rotor blade is
a rotor blade body having an airfoil-shaped cross section extending in the radial direction; and a shroud provided at the tip of the rotor blade body,
The turbine rotor blade stage is
The shrouds of the turbine rotor blades adjacent in the circumferential direction are connected to each other, and the rotor blade body has 50 or more and 72 or less shrouds,
An aircraft gas turbine, wherein the absolute value of the difference between the number of stator blade bodies of the first upstream stator blade stage and the number of stator blade bodies of the downstream stator blade stage is 1 or more and 3 or less. . The aircraft gas turbine according to claim 1, wherein the value of H/Cx is 2 or more and 6 or less, where the height of the rotor blade body is "H" and the axial cord length of the rotor blade body is "Cx". . An aircraft gas turbine with a takeoff thrust of 15,000 to 40,000 lbs,
A rotor shaft that rotates around the axis,
a turbine rotor blade stage extending from the rotor shaft toward the outside in a radial direction about the axis and having a plurality of turbine rotor blades arranged in parallel in a circumferential direction about the axis;
a turbine casing that covers the rotor shaft from the outer peripheral side;
A plurality of stationary blade bodies are arranged on the upstream side in the axial direction of the turbine rotor blade stage, extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. a first upstream stator vane stage which is a turbine stator vane stage having;
A plurality of stationary blade bodies are arranged downstream of the turbine rotor blade stage in the axial direction, extend from the turbine casing toward the inside in the radial direction, and are arranged side by side in a circumferential direction centered on the axis. a downstream stator vane stage that is a turbine stator vane stage,
The turbine rotor blade is
a rotor blade body having an airfoil-shaped cross section extending in the radial direction; and a shroud provided at the tip of the rotor blade body,
When the height of the rotor blade body is "H" and the shaft cord length of the rotor blade body is "Cx", the value of H/Cx is 2 or more and 6 or less,
An aircraft gas turbine, wherein the absolute value of the difference between the number of stator blade bodies of the first upstream stator blade stage and the number of stator blade bodies of the downstream stator blade stage is 1 or more and 3 or less. . an upstream rotor blade stage as the turbine rotor blade stage disposed on the axially upstream side of the first upstream stator vane stage;
a second upstream stator vane stage as the turbine stator vane stage disposed upstream of the upstream rotor blade stage in the axial direction;
The absolute value of the difference between the number of stator blade bodies of the first upstream stator vane stage and the number of stator blade bodies of the second upstream stator vane stage is 1 or more and 3 or less. The aircraft gas turbine according to any one of claims 1 to 3. The aircraft gas turbine according to any one of claims 1 to 4, wherein the turbine rotor blade is made of a TiAl alloy. The aircraft gas turbine according to any one of claims 1 to 5, wherein the rotor blade body is a hollow blade. The aircraft gas turbine according to any one of claims 1 to 5, wherein at least a portion of the rotor blade body in the blade height direction is formed hollow. An aircraft gas turbine with a takeoff thrust of 15,000 to 40,000 lbs,
A rotor shaft that rotates around the axis,
a turbine rotor blade stage extending from the rotor shaft toward the outside in a radial direction about the axis and having a plurality of turbine rotor blades arranged in parallel in a circumferential direction about the axis;
a turbine casing that covers the rotor shaft from the outer peripheral side;
A plurality of stationary blade bodies are arranged on the upstream side in the axial direction of the turbine rotor blade stage, extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. a first upstream stator vane stage which is a turbine stator vane stage having;
an upstream rotor blade stage that is a turbine rotor blade stage disposed on the axially upstream side of the first upstream stationary blade stage;
a second upstream stator vane stage that is a turbine stator vane stage disposed on the upstream side of the upstream rotor blade stage in the axial direction;
The turbine rotor blade is
a rotor blade body having an airfoil-shaped cross section extending in the radial direction; and a shroud provided at the tip of the rotor blade body,
The turbine rotor blade stage is
The shrouds of the turbine rotor blades adjacent in the circumferential direction are connected to each other, and the rotor blade body has 50 or more and 72 or less shrouds,
The absolute value of the difference between the number of stator vane bodies of the first upstream stator vane stage and the number of stator vane bodies of the second upstream stator vane stage is greater than or equal to 1 and less than or equal to 3. gas turbine. An aircraft gas turbine with a takeoff thrust of 15,000 to 40,000 lbs,
A rotor shaft that rotates around the axis,
a turbine rotor blade stage extending from the rotor shaft toward the outside in a radial direction about the axis and having a plurality of turbine rotor blades arranged in parallel in a circumferential direction about the axis;
a turbine casing that covers the rotor shaft from the outer peripheral side;
A plurality of stationary blade bodies are arranged on the upstream side in the axial direction of the turbine rotor blade stage, extend inward in the radial direction from the turbine casing, and are arranged in line in a circumferential direction centered on the axis. a first upstream stator vane stage which is a turbine stator vane stage having;
an upstream rotor blade stage that is a turbine rotor blade stage disposed on the axially upstream side of the first upstream stationary blade stage;
a second upstream stator vane stage that is a turbine stator vane stage disposed on the upstream side of the upstream rotor blade stage in the axial direction;
The turbine rotor blade is
a rotor blade body having an airfoil-shaped cross section extending in the radial direction; and a shroud provided at the tip of the rotor blade body,
When the height of the rotor blade body is "H" and the shaft cord length of the rotor blade body is "Cx", the value of H/Cx is 2 or more and 6 or less,
The absolute value of the difference between the number of stator vane bodies of the first upstream stator vane stage and the number of stator vane bodies of the second upstream stator vane stage is greater than or equal to 1 and less than or equal to 3. gas turbine.

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