JP2718131B2 - Gas turbine disk - Google Patents

Description

Description: TECHNICAL FIELD The present invention relates to a gas turbine disk, and more particularly to a gas turbine disk provided in a gas turbine driven by a high-temperature gas.

[Prior Art] Generally, an axial flow type gas turbine for power generation or the like is provided with a gas turbine disk that is driven to rotate by a gas flow.

As shown in FIG. 4, a conventional gas turbine disk of this type has a disk main body 1 formed of a heat-resistant metal or the like, and rotor blades 2 arranged side by side along the circumferential direction. The rotor blade 2 is provided with a root portion 3 whose base end is bulged, and is fitted to a groove portion 4 which is depressed according to the shape so that the rotor blade 2 is fixed to the disk main body 1. Has become. The blade 2 is formed with a platform 6 protruding substantially in the circumferential direction at a position below the blade portion 5 so as to be connected to the platform 6 of another neighboring blade 2 so as to be connected to the disk body 1. By separating the gas passage 7 from the gas passage 7, heat transfer from the gas flow to the disk main body 1 is reduced.

In particular, when a ceramic rotor blade is used, a cushioning material (not shown) made of metal foil, metal felt, or the like is provided between the groove portion 4 and the root portion 3 in order to make the fitting contact with each other sufficiently uniform. Zu) is inserted.

[Problems to be Solved by the Invention] By the way, metal foil, metal felt and the like are made to uniformly contact the root 3 and the groove 4 by plastic deformation. When driving in
Causes creep deformation. For this reason, there is a problem that the clearance between the tip of the blade portion 5 of the moving blade 2 and the shroud (not shown) is reduced, and eventually damage is caused by mutual contact.

Further, the cushioning material after creep deformation does not sufficiently exhibit a cushioning function, and local stress concentration occurs, thereby reducing the durability of the turbine disk.

In view of the above circumstances, the present invention has been devised to provide a gas turbine disk in which the cushioning material does not undergo creep deformation even at high temperature and high speed operation, and the fitting contact between the rotor blade and the main body is kept sufficiently uniform. It is a thing.

[Means for Solving the Problems] The present invention provides a plurality of metal tapes having different ceramic contents between a root portion formed on a base end side of a rotor blade and a groove portion for fitting the root portion. Are provided with a buffer layer formed by stacking the layers.

[Operation] According to the buffer layer having the above configuration, the metal in the metal tape having a low ceramic content is plastically deformed, so that the fitting contact between the root portion and the groove portion of the moving blade is uniform, and the stress concentration is reduced. In addition, since the ceramic in the metal tape having a high ceramic content exhibits high strength, creep deformation is reduced, the peripheral speed of the turbine disk can be increased, and the durability is improved.

Hereinafter, an embodiment of the present invention will be described with reference to the accompanying drawings.

FIG. 1 shows an embodiment of a gas turbine disk according to the present invention, and the same reference numerals are given to the same components as those in the related art, and the description thereof will be omitted.

This gas turbine disk is provided with a cushioning material 11 containing ceramic between a root 3 formed on the base end side of the rotor blade 2 and a groove 4 formed in the disk main body 1 for fitting the root 3. A buffer layer 12 is provided.

In the present embodiment, the cushioning material 11 is formed as a metal tape containing powdered ceramics (alumina), and three metal tapes having different contents are laminated to form the buffer layer 12. ing. That is, as shown in FIG. 2, a first metal tape 13 containing 10% alumina powder with respect to nickel as a base metal is sandwiched between second metal tapes 14 containing 2% alumina powder. It is.

The moving blade 2 is formed of silicon nitride ceramics.

With this configuration, the cushioning material 11 has the plastic deformation capability of metal, and at the same time, the high strength and high heat resistance of alumina are added, eliminating creep deformation, and under high temperature and high speed operation. Also, the contact between the root 3 and the groove 4 of the rotor blade 2 becomes uniform, so that stress concentration is reduced,
The peripheral speed of the turbine disk can be increased, and the durability can be improved.

Further, since there is no possibility of interference between the tip of the blade portion 5 of the moving blade 2 and the shroud, the clearance can be reduced, the leak can be reduced, and the efficiency of the turbine can be increased.

In manufacturing the metal tapes 13 and 14,
A molten metal (nickel) containing alumina powder at a predetermined ratio is quenched on a cooled roll by a quenching ribbon method, or a mixed powder of alumina powder and nickel powder is dispersed in a liquid to form a slurry. The tape may be formed by a doctor blade method. Alternatively, the same mixed powder may be kneaded with an organic binder or the like and formed into a tape by extrusion molding or calendar roll molding.

Incidentally, the present inventors, using the gas turbine disk according to the configuration of the above embodiment, the turbine inlet temperature 1300 ℃,
When the test was performed at a peripheral speed of 600 m / s, this buffer layer
12 shows a sufficient function, while maintaining a state where creep deformation does not occur even if the continuous operation time reaches 200 hours,
The reduction in the thickness of the cushioning material 11 was at a level that did not cause any problem. In contrast, in the conventional configuration (with nickel foil inserted), under the same conditions, after 50 hours, the tip of the moving blade contacted the shroud and was damaged due to excessive creep deformation of the cushioning material. And have obtained the results.

[Effects of the Invention] In summary, according to the present invention, the following excellent effects are exhibited.

A buffer layer made of a stack of multiple metal tapes with different ceramic contents is provided between the root and the groove, so that creep deformation does not occur and the fitting contact with the rotor blades can be uniform, and gas This can contribute to the improvement of turbine durability.

[Brief description of the drawings]

FIG. 1 is a sectional view showing an embodiment of a gas turbine disk according to the present invention, FIG. 2 is an enlarged view of a main part thereof, and FIG. 3 is a sectional view showing a conventional gas turbine disk. In the figure, 2 is a bucket, 3 is a root, 4 is a groove, 11 is a cushioning material, 12
Is a buffer layer.

 ──────────────────────────────────────────────────続 き Continuing the front page (72) Inventor ▲ Sakaki ▼ Masaru Tadashi 3-1-1-15 Toyosu, Koto-ku, Tokyo Ishikawa Shima-Harima Heavy Industries Co., Ltd. (56) References JP-A-49-105015 (JP) , A)

Claims (1)

(57) [Claims] 1. A buffer layer formed by stacking a plurality of metal tapes having different ceramic contents between a root formed on a base end side of a rotor blade and a groove for fitting the root. A gas turbine disk.