Patent ID: 8739550

Claim:
A combustor for an aerospace gas turbine engine comprising: a) a first stage wherein the first stage defines a reformer comprising a first stage inlet and a first stage exit, the first stage inlet adapted to input all of a fuel to the reformer; the reformer further comprising a metal screen substrate and adapted to dry reform a fuel-rich mixture of fuel and air into a reacted fuel stream comprising carbon monoxide and hydrogen; b) a second stage wherein the second stage defines a second stage inlet and a second stage exit, and wherein the second stage inlet is in fluid communication with the first stage exit thereby defining a first flowpath having a first flowpath inlet, and wherein the first flowpath passes substantially through the second stage and defines a first flowpath exit; c) a plurality of flow channel tubes positioned within the second stage wherein each flow channel tube defines a flow channel tube inlet and a flow channel tube exit, and wherein each flow channel tube inlet passes sealingly through a header plate positioned upstream of the second stage inlet thereby defining a second flowpath having a second flowpath inlet defined by the plurality of flow channel tube inlets, and wherein the second flowpath passes substantially through the second stage and defines a second flowpath exit defined by the plurality of flow channel tube exits, further wherein the first flowpath adapted to pass the reacted fuel stream comprising carbon monoxide and hydrogen within the second stage is configured to be in heat exchange with the second flowpath adapted to pass combustion air; and (d) wherein the first flowpath exit and the second flowpath exit, defined at the termination of the plurality of flow channel tubes, are positioned adjacent and proximate to one another.