CROSS-REFERENCE TO RELATED APPLICATION

This application claims priority to Korean Patent Application No. 10-2019-0073078, filed on Jun. 19, 2019, the disclosure of which is incorporated herein by reference in its entirety.

BACKGROUND

Technical Field

Apparatuses and methods consistent with exemplary embodiments relate to a combustor and a gas turbine including the same, and more particularly, to a combustor capable of reducing a loss of pressure while enhancing a cooling efficiency of a liner and transition piece by including first and second channels to guide air for cooling the liner or the transition piece, and a gas turbine including the same.

Description of the Related Art

Turbines are machines that convert an energy of a fluid, such as water, gas, or steam, into mechanical work, and are referred to as turbo machines in which a large number of buckets or blades are mounted to a circumference of each rotor and steam or gas is emitted thereto to rotate the rotor at high speed by impingement or reaction force.

Examples of the turbines include a water turbine using the energy of high-positioned water, a steam turbine using the energy of steam, an air turbine using the energy of high-pressure compressed air, a gas turbine using the energy of high-temperature and high-pressure gas, and the like.

The gas turbine is an internal combustion engine that converts thermal energy into a rotational force as mechanical energy in such a way to rotate a turbine by injecting, into the turbine, high-temperature and high-pressure combustion gas that is produced by mixing fuel with air compressed to high pressure in a compressor and by burning a mixture thereof.

The gas turbine is advantageous in that consumption of lubricant is extremely low due to an absence of mutual friction parts such as a piston-cylinder because the gas turbine does not have a reciprocating mechanism such as a piston which is usually provided in a four-stroke engine, and an amplitude of vibration is markedly reduced unlike reciprocating machines having high-amplitude characteristics. Therefore, high-speed motion is possible.

The gas turbine may include a compressor that compresses air, a combustor that burns a mixture of fuel and the compressed air supplied from the compressor to produce combustion gas, and a turbine that generates electric power by rotating turbine blades with the high-temperature and high-pressure gas supplied from the combustor.

The combustor may mix fuel with compressed air supplied from the compressor, combust the mixture to generate high-temperature and high-pressure combustion gas having high energy, and increase through an isobaric combustion process the temperature of the combustion gas to a heat resistant limit temperature at which the combustor and the turbine can endure.

The combustor of the gas turbine may include a liner and a transition piece. The mixture of fuel and compressed air is burned to produce hot combustion gas in the liner, and the hot combustion gas flows from the liner through the transition piece into the turbine. In order to prevent the liner and the transition piece from being damaged due to the high temperature of the combustion gas, outer walls of the liner and transition piece should be cooled by the compressed air supplied from the compressor.

To this end, the liner and the transition piece are surrounded by a flow sleeve so that the compressed air is introduced through cooling holes formed in the flow sleeve to impinge on the outer walls of the liner and transition piece, thereby enabling the liner and the transition piece to be cooled. The compressed air used to cool the liner and the transition piece flows to a front of the liner along a space between the liner and transition piece and the flow sleeve.

In this case, the compressed air introduced into the flow sleeve through the cooling holes is directed radially inward therefrom, and the compressed air flowing to the front of the liner along the space between the liner and transition piece and the flow sleeve is directed in an axial direction of the flow sleeve. Therefore, a loss of pressure may occur due to an interaction caused if the two types of compressed air flowing in different directions meet each other.

SUMMARY

Aspects of one or more exemplary embodiments provide a combustor capable of reducing a loss of pressure while enhancing a cooling efficiency of a liner and transition piece as each insert includes first and second channels to guide air for cooling the liner or the transition piece, and a gas turbine including the same.

Additional aspects will be set forth in part in the description which follows and, in part, will become apparent from the description, or may be learned by practice of the exemplary embodiments.

According to an aspect of an exemplary embodiment, there is provided a combustor including: a liner configured to define a combustion chamber, a transition piece coupled to a rear end of the liner, a flow sleeve configured to surround the liner and the transition piece, a plurality of impingement holes formed in the flow sleeve, and a plurality of inserts inserted into at least some of the impingement holes, wherein each of the inserts may include a first channel configured to guide combustion air, introduced into an associated one of the impingement holes, in a direction parallel to a direction of extension of an annular passage between the flow sleeve and the liner or an annular passage between the flow sleeve and the transition piece, and a second channel configured to guide the combustion air, introduced into the associated one of the impingement holes, in a direction transverse to the annular passage between the flow sleeve and the liner or the annular passage between the flow sleeve and the transition piece.

The first channel may be formed in front of the second channel in a direction of flow of combustion air in the annular passage.

The insert may have a low air resistance coefficient.

The inserts may be arranged in a plurality of rows in an axial direction of the flow sleeve, and the inserts in adjacent rows may be staggered.

At least some of distances between the adjacent rows may be different from each other.

The second channel may be branched from the first channel to extend toward a surface of the liner or transition piece.

The second channel may include two or more second channels.

The two or more second channels may be formed symmetrically with respect to a plane of symmetry of the first channel.

An angle (α) of a direction of discharge of combustion air from each of the first channels with respect to a surface of the liner or transition piece may be in a range of −10 to 30 degrees.

In the inserts in adjacent rows, angles (α) of directions of discharge of individual types of combustion air from the respective first channels with respect to the surface of the liner or transition piece may be different from each other.

An angle (β) of the horizontal direction of discharge of combustion air from each of the second channels with respect to a plane of symmetry of an associated first channel may be in a range of −30 to 30 degrees.

An angle (γ) of the vertical direction of discharge of combustion air from each of the second channels with respect to a plane of symmetry of an associated first channel may be in a range of −30 to 30 degrees.

According to an aspect of another exemplary embodiment, there is provided a gas turbine including: a compressor configured to compress air, a combustor configured to mix fuel with the air compressed by the compressor and combust a mixture of fuel with the compressed air to produce high-temperature and high-pressure combustion gas, and a turbine configured to generate power using the combustion gas produced by the combustor. The combustor may include a liner configured to define a combustion chamber, a transition piece coupled to a rear end of the liner, a flow sleeve configured to surround the liner and the transition piece, a plurality of impingement holes formed in the flow sleeve, and a plurality of inserts inserted into at least some of the impingement holes, wherein each of the inserts may include a first channel configured to guide combustion air, introduced into an associated one of the impingement holes, in a direction parallel to a direction of extension of an annular passage between the flow sleeve and the liner or an annular passage between the flow sleeve and the transition piece, and a second channel configured to guide the combustion air, introduced into the associated one of the impingement holes, in a direction transverse to the annular passage between the flow sleeve and the liner or the annular passage between the flow sleeve and the transition piece.

The first channel may be formed in front of the second channel in a direction of flow of compressed air in the annular passage.

The inserts may be arranged in a plurality of rows in an axial direction of the flow sleeve, and the inserts in adjacent rows may be staggered.

The second channel may be branched from the first channel to extend toward a surface of the liner or transition piece.

The second channel may include two or more second channels, and the two or more second channels may be formed symmetrically with respect to a plane of symmetry of the first channel.

An angle (α) of a direction of discharge of compressed air from each of the first channels with respect to a surface of the liner or transition piece may be in a range of −10 to 30 degrees.

An angle (β) of a horizontal direction of discharge of compressed air from each of the second channels with respect to a plane of symmetry of the associated first channel may be in a range of −30 to 30 degrees.

An angle (γ) of a vertical direction of discharge of compressed air from each of the second channels with respect to a plane of symmetry of the associated first channel may be in a range of −30 to 30 degrees.

