JPH02169804A - Gas turbine disc - Google Patents

Abstract

PURPOSE:To restrain heat transmission to a disc body by covering the outer surface of the disc facing high temperature gas streams and the inner surfaces of grooves in which embedding parts of rotor blades are fitted, with a ceramic material so as to form a heat-insulating layer. CONSTITUTION:A heat-insulating layer 24 made of a ceramic material is formed on the outer surface 21 of disc body facing high temperature gas streams and the inner surfaces 23 of grooves 22 in the root parts 3 of rotor blades 2 are fitted. With this arrangement, it is possible to restrain the temperature of the disc body it self without increasing the volume of air for cooling the disc body 1 even through the temperature of gas at the turbine inlet is raised. Accordingly, it is possible to aim at enhancing both efficiency and reliability of the turbine.

Description

DETAILED DESCRIPTION OF THE INVENTION [Field of Industrial Application] The present invention relates to a gas turbine disk, and more particularly to a gas turbine disk installed in a gas turbine driven by high-temperature gas.

[Prior Art] Generally, an axial flow gas turbine for power generation or the like includes a gas turbine disk that is rotationally driven by a gas flow.

As shown in FIG. 3, conventional gas turbine disks of this type include a disk body 1 made of heat-resistant metal or the like, and rotor blades 2 arranged in parallel along the circumferential direction of the disk body 1. The rotor blade 2 is provided with an implant portion 3 whose proximal end side is bulged, and is fixed to the disk body 1 by fitting into a groove portion 4 recessed to match the shape of the implant portion 3. It has become. The rotor blade 2 is provided with a plug (form 6) protruding approximately in the circumferential direction at a position below the blade portion 5.
is formed, and the platforms 6 of other adjacent rotor blades 2
By separating the disk body 1 and the gas passage 7 so as to be continuous with each other, heat transfer from the gas flow to the disk body 1 is alleviated.

In addition, there is also known a structure in which a material such as metal foil or metal felt is inserted between the groove part 4 and the implant part 3.

[Problems to be Solved by the Invention] Nowadays, in order to obtain a gas turbine with higher efficiency, there are cases in which gas is heated to a high temperature and the rotor blades 2 are hardly cooled.

Therefore, the temperature at the turbine inlet where the turbine disk is provided becomes very high (for example, about 1200° C. or higher), and the temperature of the blade portion 5 of the rotor blade 2 and the platform 6 also becomes high. Therefore, there is a problem in that heat transfer from the gas flow directly to the disk body 1 or via the rotor blades 2 is large, and the temperature of the disk body 1 increases significantly, leading to loss of integrity of the gas turbine. Ta.

Using a large amount of cooling air to suppress the temperature rise will reduce the efficiency of the gas turbine, and at the same time
Thermal stress in the vicinity of the fitting portion becomes extremely large, which also poses a problem in the reliability of the turbine.

In view of the above circumstances, the present invention was devised to provide a gas turbine disk that suppresses heat transfer to the disk body.

[Means for Solving the Problems] The present invention provides a heat insulating layer coated with a ceramic material on the outer circumferential surface facing the high-temperature gas flow and on the inner surface of the groove in which the implanted portion of the rotor blade is fitted. It is something.

[Function] With the above configuration, the heat insulating layer provided on the outer peripheral surface prevents heat transfer from the gas flow. Furthermore, the heat insulating layer provided on the inner surface of the groove suppresses heat transfer via the rotor blades.

[Example] Embodiments of the present invention will be described below with reference to the accompanying drawings.

FIG. 1 shows an embodiment of a gas turbine disk according to the present invention, and components similar to those of the conventional system are denoted by the same reference numerals, and their explanations will be omitted.

This gas turbine disk has a heat insulating layer 24 provided on the outer circumferential surface 21 facing the high-temperature gas flow and on the inner surface 23 of the groove 22 into which the implanted portion 3 of the rotor blade 2 is fitted. The heat insulating layer 24 is molded from a ceramic material, and these surfaces 21.
It covers 23.

That is, the disk body 1 is formed of a nila gel alloy or the like.
is coated over its entire surface with a ceramic material.

In this embodiment, zirconia is used as the ceramic material because of its heat insulating properties, stability, and small difference in thermal expansion from the metal forming the disk body 1.

Further, in this embodiment, as shown in FIG. 2, an intermediate layer 25 is provided between the heat insulating layer 24 and the disk body 1. In this intermediate layer 25, on the inner surface 23 of the groove 22, heat-resistant metals H, (N
1crAJI Y), ceramic/heat-resistant metal mixture (
NtCrAjY-50%.

zirconia) are laminated. Similarly, on the outer peripheral surface 21, heat-resistant metal (NiCrAjY), ceramic-heat-resistant metal mixture I (NiCrAjY-30%, zirconia), ceramic-heat-resistant metal mixture I [(Ni
CrAJ Y-50%, zirconia), ceramic
A heat-resistant metal mixture 111 (NiCrAjY-70%, zirconia) is laminated.

Note that the rotor blade 2 is molded from silicon nitride ceramics.

Next, the operation of this embodiment will be explained.

The heated gas rotates the turbine disk to perform turbine work, and the heat is transferred to the rotor blades 2 and the disk body 1. At this time, the heat insulating layer 24 is
The characteristics of the ceramic material block heat transfer via the rotor blades 2 and directly, thereby preventing excessive temperature rise of the disk body 1.

The intermediate layer 25 prevents the generation of harmful thermal stress due to the difference in thermal expansion between the heat insulating layer 24 and the disk body 1. In particular, in the groove portion 22, since the intermediate layer 25 has a multilayer structure with different coefficients of thermal expansion, the distance from the implanted portion 3 to the disk body 1 is
This reduces the occurrence of thermal stress over time and maintains the soundness of the fit.

In addition, when manufacturing this gas turbine disk,
First, disc book 1+ with Niragel alloy.

At the same time, a groove portion 22 having a predetermined shape is formed.

Then, the inner surface 23 of the groove 22 is undercoated with a heat-resistant metal by plasma spraying, followed by coating with a ceramic/heat-resistant metal mixture (I[), and finally coating with zirconia. Polish it again to fit it into the recessed part 3. On the other hand, the outer circumferential surface 21 is also undercoated with a heat-resistant metal by plasma spray, and then coated with the ceramic/heat-resistant metal mixture 1, Il, and I in this order, and the fif& is coated with zirconia.

In this way, a heat insulating layer made of ceramic material is provided on the disk body 1 to improve the heat shielding properties of the gas turbine disk, so even if the gas temperature at the turbine inlet is increased, the amount of air for cooling the disk can be increased. It is possible to suppress the rise in disk temperature without causing any problems. That is, it is possible to simultaneously improve the efficiency and reliability of the turbine.

Note that as the intermediate layer 25, conventional metal felt or the like may be used.

Further, when the present inventors conducted a hot spin test at a turbine inlet temperature of 1350'C using the gas turbine having the configuration of the above embodiment, they found that the conventional configuration (third
The results showed that the temperature at the outer periphery of the disk could be suppressed by more than 100°C compared to the case shown in Figure 1.

[Effects of the Invention] In short, the present invention provides the following excellent effects.

Since the outer peripheral surface and the inner surface of the groove are provided with a heat insulating layer coated with a ceramic material, it is possible to prevent an excessive temperature rise of the gas turbine disk and contribute to improving the efficiency and reliability of the gas turbine.

[Brief explanation of the drawing]

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 the main part thereof, and FIG. 3 is a sectional view showing a conventional gas turbine disk. In the figure, 21 is the outer peripheral surface, 22 is the groove, 23 is the inner surface, 2
4 is a heat insulating layer. Patent applicant Ishikawa & Harima Heavy Industries Co., Ltd. Representative Patent Attorney
Nobuo Kinutani 24...Insulation 1 Figure 3 Figure 2

Claims (1)

[Claims] 1. A gas turbine disk characterized in that a heat insulating layer made of a ceramic material is provided on the outer circumferential surface facing the high-temperature gas flow and on the inner surface of the groove portion into which the implanted portion of the rotor blade is fitted.