JPS62174502A - Gas turbine blade - Google Patents

Abstract

PURPOSE:To alleviate thermal stress by forming a flow passage continuous from the leading edge to the trailing edge of a ceramic hollow blade. CONSTITUTION:A hole 3 is formed on the leading edge 2 of a blade shape member 1 and a hole 5 on the trailing edge 4 thereof. A flow passage wherein high temperature gases run is formed via a hollow part 6. The blade shape member 1 is retained at inner and outer diameter sides via cylindrical surfaces 7 and 8. According to the aforesaid constitution, a temperature difference between the internal and external blade surfaces is reduced at a non-steady state, thereby thermal stress on the blade being alleviated.

Description

[Detailed description of the invention] [Technical field of invention] This invention relates to gas turbine blades.

[Technical background of the invention and its problems]

Efforts have been made in many fields to raise the temperature of the gas in gas turbines in order to increase efficiency, but since metals are mainly used for blade materials, air, steam, or water is used for cooling.

For this reason, even if the temperature increases, the blade is cooled from the point of view of the blade material, and this cooling reduces the effect of the high temperature.

[Purpose of the invention]

This invention is an improvement on the conventional gas turbine blade described above, and an object thereof is to provide a gas turbine blade that does not require cooling.

[Summary of the invention]

By using ceramic for the parts that are directly heated by high-temperature combustion gas and the fixed support parts, we have constructed a gas turbine blade that operates stably even without special cooling.

〔Effect of the invention〕

When the gas turbine combustion gas temperature exceeds 800°C using conventionally used heat-resistant metals, the blades of the gas turbine are cooled with a cooling medium such as air to maintain the metal temperature at around 800°C.

For this reason, research has been conducted to improve the cooling performance of gas turbine blades due to their high temperatures, and this has been put into practical use.

However, by performing cooling, the operating temperature of the gas turbine decreases, making it impossible to make full use of the benefits of increasing the temperature.

The blade of the present invention has a hot gas temperature of about 1 without cooling.
It can operate up to about 300°C and provides a highly efficient gas turbine.

In addition, since it does not have a complicated structure like an air-cooled blade, manufacturing costs can be reduced and a highly reliable gas turbine blade can be provided.

[Embodiments of the invention]

FIG. 1 shows an embodiment of the present invention. A hole 3 is provided in the leading edge 2 of the airfoil member indicated by 1 in the figure, and a hole 5 is similarly provided in the trailing edge 4.
The blade has a structure that allows high temperature gas to flow through the hollow part 6.

The flow path through which high-temperature gas can flow is not limited to the structure of this embodiment, and may be of any type that allows high-temperature gas to flow inside the blade.

The inner and outer diameter sides of the effective part of the blade are fixed by concave surfaces provided on the cylindrical surface.The cylindrical surface is divided in the circumferential direction, and as shown in Fig. 2, the cylindrical surface 7.8 It has a structure in which the effective part is sandwiched between the inner and outer diameter sides. A heat insulating member is inserted between the inner and outer cylindrical surfaces 7, 8 and the inner and outer pressure partitions 11, 12 to prevent the temperature of the pressure partitions from rising.

The phase shift of the blade due to the fluid force acting on the effective part of the blade.

The cylindrical surfaces 7, 8 and the heat insulating member 9,
10 is provided with a step 14 to prevent rotation, and the heat insulating members 9 and 10 and the pressure partition wall are prevented from rotating by pins 13.

These detents are not limited to this method,
A method using screws or the like may also be used.

Since the materials used for the wing parts and the inner and outer cylindrical surfaces are directly exposed to high-temperature combustion gas, materials 9 and 10 that can withstand high temperatures, such as ceramics and tungsten, are preferable.The insulating members 9 and 10 are preferably materials that can withstand high temperatures and have low thermal conductivity.

Table 1 shows the thermal conductivity and thermal expansion coefficient of three types of ceramics and heat-resistant metals.
) has low thermal conductivity and good properties. The airfoil members and inner and outer cylindrical surfaces that are directly exposed to high-temperature combustion gas have high-temperature strength and heat resistance of 1f71! Silicon nitride (SII) has a large temperature difference.
N4), silicon carbide (Sin), etc. are preferable.

Table 1 When the combustion gas is heated to high temperatures, gas turbine blades using ceramics not only have high-temperature strength but also may break due to thermal stress generated inside the ceramic member due to sudden changes in gas temperature. .

For this reason, the thermal stress is reduced by changing the blade from a solid blade to a hollow blade, but in order to achieve higher temperatures, the temperature difference between the outside and outside of the blade during unsteady conditions (startup, stoppage) is reduced and the thermal stress is reduced. In order to reduce this, high-temperature gas is also circulated inside the five blades, and when the gas turbine is started or stopped, heating and cooling are performed from the inside and outside surfaces of the blades, which greatly alleviates thermal stress.

[Other embodiments of the invention]

Although this embodiment shows an example in which the present invention is applied to a stationary blade, the present invention can also be applied to the blade effective portion of a rotor blade in the same manner.

[Brief explanation of drawings]

FIG. 1 is a perspective view showing an embodiment of a gas turbine blade of the present invention, and FIG. 2 is a diagram showing an exploded view of the main parts of FIG. 1. 1... Airfoil member 2... Leading edge portion 3.5...
- Hole 4... Trailing edge 6... Hollow part
7, 8... Cylindrical agent Patent attorney Nori Chika
Ken Yudo Kikuo Takehana Figure 1 Figure 2

Claims (1)

[Scope of Claims] 1. A gas turbine blade, characterized in that a flow path communicating from the leading edge to the trailing edge of an airfoil portion formed of ceramic is provided. 2. The blade of a gas turbine according to claim 1, wherein the airfoil portion is provided separately from the annular flow path forming portion.