JP2001193481A - Gas turbine and heat recovery method thereof - Google Patents

Abstract

PROBLEM TO BE SOLVED: To reduce the size and weight of a gas turbine and to reduce the pressure loss of compressed air by omitting a heat exchanger at the exit side of a turbine part. SOLUTION: This gas turbine is provided with a compressor 1 compressing air A, a combustor 2 combusting a fuel, a turbine part 3 taking out an output from an exhaust gas EG of the combustor 2, an air duct 4 introducing the air A from the compressor 1 to the turbine part 3, and an air duct 4' introducing the air A from the turbine part 3 to the combustor 2.

Description

DETAILED DESCRIPTION OF THE INVENTION

[0001]

BACKGROUND OF THE INVENTION 1. Field of the Invention The present invention relates to a gas turbine comprising a compressor which compresses air with a compressor and expands high-temperature and high-pressure exhaust gas in a combustor to extract output power. The present invention relates to a gas turbine capable of recovering heat in a section and a heat recovery method thereof.

[0002]

2. Description of the Related Art In a conventional gas turbine, as shown in a block diagram of FIG. 5, air A is compressed by a compressor 1 and fuel is burned in a combustor 2 using the compressed air. The output is extracted to the outside by the exhaust gas EG being expanded in the turbine section 3. However, this combustor 2
In the gas turbine of the type that discharges exhaust gas EG into the atmosphere from the outlet side of the turbine unit 3 and draws in new air A, the temperature of the compressed air A sent to the combustor 2 decreases, so that the combustion efficiency is reduced. It was low. Therefore, in order to increase the combustion efficiency in the combustor 2, the air A from the compressor 1
There is a gas turbine using exhaust gas which recovers heat in the heat exchanger 4 on the outlet side of the turbine section 3 and guides it to the compressor 1 again for circulating use.

On the other hand, gas turbines generally require an increase in cycle efficiency and specific output. Therefore, it is necessary to increase the pressure ratio and the gas temperature at the turbine inlet.
That is, it is necessary to efficiently cool the moving blades and the stationary blades that constitute the turbine section 3 that is located immediately after the combustor 2 and is heated by recovering the heat of the exhaust gas EG. By cooling the turbine section 3 in this manner, the temperature of the gas at the turbine inlet can be made higher than the temperature of the members such as the moving blades and the stationary blades of the turbine section 3. Therefore, in order to cool the turbine section 3 efficiently, a part A ′ of the air A from the compressor 1 is
Was discharged from the cooling air outflow holes provided in the moving blades or the stationary blades.

[0004]

However, the turbine section 3
The size of the heat exchanger 4 separately installed in the exhaust part at the outlet was larger than that of the gas turbine body including the compressor 1, the combustor 2, and the turbine unit 3. The heat exchanger 4 having such a large scale is suitable for a gas turbine installed on the ground, but has a problem that it is unsuitable for a gas turbine such as an aircraft which requires small size and light weight.

[0005] Further, in the conventional gas turbine provided with the separate heat exchanger 4, the pressure loss of the compressed air A is liable to occur, so that there is a problem that the gas turbine cannot be used in an aircraft jet engine having a high exhaust speed.

The present invention has been made to solve such a problem. That is, an object of the present invention is to omit the heat exchanger on the turbine section outlet side, thereby reducing the size and weight of the gas turbine and reducing the pressure loss of the compressed air, and the heat recovery method therefor. Is to provide.

[0007]

According to the present invention, a compressor (1) for compressing air (A), a combustor (2) for burning fuel, and an exhaust gas (2) for the combustor (2) are provided. EG), an air passage (4) for guiding air (A) from the compressor (1) to the turbine section (3), and an air passage (4) from the turbine section (3). A) an air passage (4 ') for guiding A) to the combustor (2).

[0008] Air (A) compressed by the compressor (1)
Through the air passage (4) and the turbine section (3)
Can be configured to recover heat in the stationary blade (5) or the moving blade (6).

According to the configuration of the present invention, the air (A) compressed by the compressor (1) is guided to the turbine section (3) through the air passage (4), and the compressed air (A) is used for the turbine section (3). 3) is cooled to recover heat. Thus, the turbine section (3) which is immediately after the combustor (2) and has a high temperature
The stationary blade (5) and the moving blade (6) can be cooled. Therefore, by increasing the turbine inlet gas temperature, it is possible to improve fuel efficiency, increase output, and reduce noise as an engine. At the same time, the compressed air (A), which has been heated by the recovery of heat in the turbine section (3) and heated to a high temperature, is returned to the inlet of the combustor (2) through another air passage (4 '). The combustion efficiency of the combustor 2 can be increased.
Therefore, the turbine section (3) is formed by high-temperature and high-pressure compressed air (A).
Therefore, the cycle efficiency and the specific output of the gas turbine can be increased.

[0010] A part of the air (A) compressed by the compressor (1) may be guided to the turbine section (3) through the air passage (4).

As described above, even if the compressed air (A) guided to the turbine section (3) is a part of the air (A) compressed by the compressor (1), the temperature of the turbine section (3) is high. ) Can be cooled, heat is recovered in the turbine section (3), and the compressed air (A), which has been heated to a high temperature, is returned to the combustor (2) to be cooled. Combustion efficiency can be increased.

Further, according to the method of the present invention, there is provided a heat recovery method for a gas turbine, wherein heat is recovered in a turbine section (3) for extracting output from exhaust gas (EG) of a combustor (2). Is done.

[0013]

DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS Preferred embodiments of the present invention will be described below with reference to the drawings. In the drawings, common members are denoted by the same reference numerals, and duplicate description is omitted.
FIG. 1 is a block diagram showing the configuration of a flow path of only compressed air for cooling in a gas turbine. FIG. 2 is a block diagram showing the configuration of the main passage and the passage of the compressed air for cooling in the gas turbine. The gas turbine of the present invention is applied to a turboprop engine or the like described below, and compresses air A with a compressor 1 and burns fuel with the compressed gas A in a combustor 2 to generate a high-temperature high-pressure gas. The output is extracted to the outside by the exhaust gas EG being expanded in the turbine section 3. The present invention recovers the heat of the compressed air A from the compressor 1 in the turbine section 3 in order to increase the combustion efficiency of the compressed air A in the combustor 2 and to increase the turbine inlet gas temperature by cooling the turbine section 3. ,
The heated air A is returned to the combustor 2 for circulating use.

In the gas turbine according to the present invention, a portion of the air A taken from the compressor 1 is not passed through the combustor 2 and the turbine section 3
An air passage 4 is provided for directly leading to the air passage. That is, as shown in FIG. 2, in addition to the main flow path M that guides the air A in the order of the compressor 1, the combustor 2, and the turbine section 3, a flow path (air passage 4) of the compressed air A for cooling is provided. Things. With the air A from the air passage 4, heat is recovered in the turbine section 3 and the air A returned to the inlet of the combustor 2 can be heated.

In the gas turbine of the present invention, the air A compressed by the compressor 1 is guided to the turbine unit 3 through the air passage 4 by providing the turbine unit 3 with a heat exchange function.
The turbine 3 is cooled by the compressed air A to recover heat. In this manner, the stationary blades 5 and the moving blades 6 of the turbine section 3 which are located immediately after the combustor 2 and have a high temperature can be cooled. Therefore, by increasing the turbine inlet gas temperature, it is possible to improve fuel efficiency, increase output, and reduce noise as an engine. At the same time, the compressed air A, which has recovered heat in the turbine section 3 and has been heated to a high temperature, is returned to the inlet of the combustor 2 through another air passage 4 ′, thereby increasing the combustion efficiency of the combustor 2. be able to. Therefore, since the high-temperature and high-pressure exhaust gas EG can be guided to the turbine section 3, the cycle efficiency and the specific output of the gas turbine can be increased.

FIG. 3 is a partially cutaway sectional view showing an aircraft engine using a gas turbine. In the illustrated example, a gas turbine having the turbine unit 3 having the above heat exchange function is applied to a turboprop engine.
The basic components of this turboprop engine are:
The compressor comprises a compressor 1, a combustor 2, a turbine section 3, an output shaft 7, a fuel control device necessary for operation, auxiliary equipment such as a lubrication system, an air intake port 8, and an exhaust duct 9. Propeller 1
The rotation shaft 0 and the output shaft 7 are connected via a speed reducer 11. Note that the configuration of the air passages 4 and 4 ′ differs depending on the type and shape of the engine, and it is needless to say that the number and arrangement of the illustrated configurations are not limited.

FIG. 4 is an enlarged sectional view showing a stationary blade constituting the turbine section. The stationary blades 5 constituting the turbine section 3 immediately after the combustor 2 and having a high temperature are provided with a cooling passage 12 therein for cooling the same. This cooling passage 1
Numeral 2 guides the above-described compressed air A from the cooling air inlet 13 through the air passage 4. As described above, the stationary blades 5 in which the temperature unevenness due to the combustor 2 is likely to be generated are cooled using the compressed air A, so that the turbine inlet gas temperature can be increased. Therefore, a turbine inlet temperature higher than the operating temperature of the material of the stationary blade 5 can be adopted.
Since the compressed air A that has recovered heat through the cooling passage 12 in the stationary blade 5 is heated, the compressed air A is returned to the combustor 2 through the air passage 4 so that the combustion efficiency of the combustor 2 can be increased. . Also in this stationary blade 5, a cooling air outlet hole (not shown) is provided in the stationary blade 5, and compressed air A is guided to the cooling air outlet hole, and is cooled by blowing out from the outlet hole. It is also possible to use means together.

On the other hand, since the moving blade 6 is rotating at a high speed, the total temperature for the moving blade 6 is lower than that of the stationary blade 5, and the temperature unevenness in the circumferential direction is also averaged. There is no need to provide a heat exchange function. However, the turbine section 3
When it is necessary to increase the overall turbine inlet gas temperature, it is desirable to provide a cooling passage 12 inside the rotor blade 6. The cooling passage 12 in the moving blade 6 guides the above-described compressed air A from the cooling air inlet 13 through the air passage 4.

However, it is not necessary to provide a heat exchange function to all of the plurality of stages of the stationary blades 5 or the moving blades 6 constituting the turbine section 3, and it is not necessary to provide a heat exchange function to the highest gas temperature in the gas turbine. Even if the cooling passage 12 is provided only in the first stage stationary blade 5, a cooling effect can be obtained.

In the embodiment of the present invention, in order to recover the heat of the turbine section 3 of the gas turbine of the aircraft engine, the air passage 4 for recovering the heat by the stationary blade 5 and the moving blade 6 is provided. Has been described. However, a part of the air A from the compressor 1 is guided directly to the turbine unit 3 without passing through the combustor 2, and the heat is recovered in the turbine unit 3 by the air A from the air passage 4 to increase the temperature of the gas at the turbine inlet. The device is not limited to the aircraft engine described above as long as it can heat the air A to the inlet of the combustor 2.

[0021]

As described above, the gas turbine of the present invention and the heat recovery method thereof use the compressed air in the turbine section to recover the heat of the moving blades and the stationary blades which are high in temperature immediately after the combustor. The temperature of the inlet gas can be increased, and at the same time, the compressed air that has been heated and recovered to a high temperature by returning heat to the turbine section can be returned to the combustor, thereby increasing the combustion efficiency of the combustor. As described above, by increasing the turbine inlet gas temperature, it is possible to improve the fuel efficiency of the engine, increase the output, and further reduce the noise.

In the gas turbine and the heat recovery method of the present invention, the heat exchange function is imparted to the turbine section, so that it is not necessary to provide a separate heat exchanger at the exit side of the turbine section. -Since the weight can be reduced and the pressure loss of the compressed air can be reduced, there is an excellent effect that the invention can be applied to an aircraft jet engine having a high exhaust speed.

[Brief description of the drawings]

FIG. 1 is a configuration block diagram showing a flow path of only compressed air for cooling in a gas turbine of the present invention.

FIG. 2 is a block diagram showing a configuration of a main flow path together with a flow path of compressed air for cooling in the gas turbine of the present invention.

FIG. 3 is a partially cutaway sectional view showing an aircraft engine using a gas turbine.

FIG. 4 is an enlarged sectional view showing a stationary blade of a turbine section.

FIG. 5 is a configuration block diagram showing a conventional gas turbine.

[Explanation of symbols]

 DESCRIPTION OF SYMBOLS 1 Compressor 2 Combustor 3 Turbine part 4 Air passage 4 'Air passage 5 Stator blade 6 Moving blade A Air (compressed air)

Claims (5)

[Claims] 1. A compressor (1) for compressing air (A),
A combustor (2) for burning fuel, a turbine section (3) for extracting output from exhaust gas (EG) of the combustor (2),
An air passage (4) for guiding air (A) from the compressor (1) to the turbine section (3); and the turbine section (3).
And an air passage (4 ') for guiding the air (A) from the air to the combustor (2). 2. Air (A) compressed by said compressor (1)
Through the air passage (4) and the turbine section (3)
The gas turbine according to claim 1, wherein heat is recovered in the stationary blade (5) constituting the gas turbine. 3. The air (A) compressed by the compressor (1).
Through the air passage (4) and the turbine section (3)
The gas turbine according to claim 1, wherein heat recovery is performed in the rotor blade (6) constituting the gas turbine. 4. Air (A) compressed by said compressor (1)
The gas turbine according to claim 1, 2 or 3, wherein a part of the gas turbine is guided to the turbine section (3) through the air-air passage (4). 5. A method for recovering heat of a gas turbine, comprising recovering heat in a turbine section (3) for extracting output from exhaust gas (EG) of a combustor (2).