Combustor Liner Cooling at Transition Duct Interface
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Solution Overview
Problem
Current gas turbine combustors face challenges in achieving efficient and uniform cooling, particularly at the interface between the combustor liner and transition piece, especially in advanced combustors with low NOx emission systems where cooling air is limited, leading to thermal gradients and pressure loss issues.
Innovation Solution
The implementation of a combustor assembly with an annular seal structure and a hula seal arrangement that directs compressor discharge air through specific channels and apertures to efficiently cool the aft end of the combustor liner and transition piece, minimizing leakage and optimizing cooling air distribution.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Temperature
If film-cooling is used with compressor discharge air, then cooling efficiency is improved, but cooling air availability is reduced due to premixing requirements
Solution Approach 1:
The cooling air flow is segmented into multiple paths: a first path provides cooling air to the combustor liner, and a second path provides cooling air to the transition piece. This segmentation allows independent optimization of cooling for each component while maintaining overall air availability for premixing.
Solution Approach 2:
Different cooling strategies are applied to different locations: the combustor liner receives cooling air through impingement and film cooling, while the transition piece receives cooling air through channels in its wall. This local differentiation ensures optimal cooling efficiency without compromising air availability for fuel premixing.
2Temperature
If cooling air is routed to the seal region from higher-pressure location, then seal cooling is improved, but leakage increases at the interface seal
Solution Approach 1:
A seal cavity is introduced as an intermediary space between the combustor liner and transition piece. Cooling air is supplied to this cavity, which then distributes cooling to the seal region without creating direct high-pressure leakage paths. The cavity acts as a mediator that maintains pressure differential while preventing leakage.
3Temperature
If thermal barrier coatings are used with backside cooling, then heat protection is improved, but thermal gradient requirements become more challenging
Solution Approach 1:
Cooling air is supplied to the backside of the combustor liner and transition piece before the components are exposed to full combustion heat. This preliminary cooling action reduces the thermal gradient that develops during operation, making it more compatible with thermal barrier coatings and preventing thermal shock.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This solution enhances cooling efficiency and uniformity, reduces leakage, and increases air availability for premixing, thereby minimizing emissions while maintaining effective heat transfer and thermal management.
Implementation Method 1
The implementation of a combustor assembly with an annular seal structure and a hula seal arrangement that directs compressor discharge air through specific channels and apertures to efficiently cool the aft end of the combustor liner and transition piece
Implementation Method 2
The air from the plenum passes through apertures in the combustor liner and impinges on the exterior liner surface and then passes as a film over the outer or cold-side surface of the liner
Data Source
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AI summary
A combustor assembly for a turbine includes a combustor and a combustor liner (54); a first flow sleeve (62) surrounding the combustor liner forming a first substantially axially-extending flow annulus (64) radially therebetween. The first flow sleeve (62) has a first plurality of apertures (28) formed about a circumference thereof for directing compressor discharge air as cooling air radially into the first flow annulus. A transition (52) is piece connected to the combustor liner (54), the transition piece adapted to carry hot combustion gases to the turbine, and a second flow sleeve (58) surrounds the transition piece forming a second substantially axially-extending flow annulus (60) radially therebetween. The second flow sleeve has a second plurality of apertures for directing compressor discharge air as cooling air radially into the second flow annulus (60), the first substantially axially-extending flow annulus (64) connecting with the second substantially axially-extending flow annulus (60). A resilient annular seal structure (86) is disposed radially between an aft end portion (56) of the combustor liner and a forward end portion (92) of the transition piece, the resilient annular seal structure configured to form a first annular cavity (104) radially between the forward end portion of the transition piece and the aft end portion of the combustor liner. At least one transfer tube (100) extends radially from the second flow sleeve (58) through the second flow annulus (60) to the transition piece (52), and is arranged to supply compressor discharge cooling air radially from an area outside the first and second substantially axially-extending flow annuli (64, 60) directly to the resilient annular seal structure (86) and to the aft end (92) of the combustor liner.