Method and apparatus for cooling combustor liner and transition piece of a gas turbine

A turbine engine comprising:

• a combustion section;
• an air discharge section downstream of the combustion section;
• a transition region between the combustion and air discharge section;
• a turbulated combustor liner defining a portion of the combustion section and transition region, said turbulated combustor liner including a plurality of circular ring turbulators arranged in an array axially along a length defining a length of said combustor liner and located on an outer surface thereof;
• a first flow sleeve surrounding said combustor liner with a first flow annulus therebetween, said first flow annulus including a plurality of axial channels extending over a portion of an aft end portion of the liner parallel to each other, the cross-sectional area of each channel is one of substantially constant and varying along the length of the channel, said first flow sleeve having a plurality of rows of cooling holes formed about a circumference of said first flow sleeve for directing cooling air from compressor discharge air into said first flow annulus;
• a transition piece connected to at least one of said combustor liner and said first flow sleeve, said transition piece adapted to carry hot combustion gases to a stage of the turbine corresponding to the air discharge section;
• a second flow sleeve surrounding said transition piece, said second flow sleeve having a second plurality of rows of cooling apertures for directing cooling air from compressor discharge air into a second flow annulus between the second flow sleeve and the transition piece, said first flow annulus connecting to said second flow annulus;
• wherein said first plurality of cooling holes and second plurality of cooling apertures are each configured with an effective area to distribute less than 50% of compressor discharge air to said first flow sleeve and mix with cooling air from said second flow annulus serving to cool air flowing through the transition region of the engine between the combustion and air discharge sections thereof.

A method of cooling a combustor liner of a gas turbine combustor, said combustor liner having a substantially circular cross-section, and a first flow sleeve surrounding said liner in substantially concentric relationship therewith creating a first flow annulus therebetween for feeding air to the gas turbine combustor, and wherein a transition piece is connected to said combustor liner, with the transition piece surrounded by a second flow sleeve, thereby creating a second flow annulus in communication with said first flow annulus;

• the method comprising:
• providing a plurality of axially spaced rows of cooling holes in said flow sleeves, each row extending circumferentially around said flow sleeves, a first of said rows in said second sleeve is located proximate an end where said first flow sleeve and said second flow sleeve interface;
• supplying cooling air from compressor discharge to said cooling holes;
• configuring said cooling holes with an effective area to distribute less than a third of compressor discharge air to said first flow sleeve and mix with a remaining compressor discharge air flowing from said second flow annulus, and
• configuring said first flow annulus with a plurality of axial channels extending over a portion of an aft end portion of the liner parallel to each other, the cross-sectional area of each channel is one of substantially constant and varying along a length of the channel.