Patent ID: 8047000

Claim:
A gas turbine combustion system, comprising: an aircraft gas turbine combustion chamber having a combustion chamber wall at least partially formed from a plurality of individual hollow finned tubes connected to each other side-by-side and parallel to one another to form a gas-tight and integrated flame tube and combustion chamber pressure hull, wherein each of the individual tubes is connected to a cooling liquid supply, such that the cooling liquid will flow through the hollow portion of each tube to cool the tube to form a liquid cooling system which includes a closed cooling circuit; wherein a fluid is used in the liquid-cooling system, which is at least one of gaseous and liquid under atmospheric standard conditions; wherein the combustion chamber wall includes a heat-insulating material on a flame side, the heat-insulating material being a ceramic coating; a burner positioned at a head of the combustion chamber and having a burner air inlet; an expanding duct having a funnel shaped contour directly connecting an outlet of a compressor to the burner air inlet and feeding an entirety of the compressed air flowing therethrough to the burner air inlet, the expanding duct including a pre-diffusor casing; a recuperative heat exchanger positioned within the pre-diffusor casing for transferring heat energy absorbed by the cooling liquid in the combustion chamber tubes to combustion air from the compressor.