Abstract: A turbine engine includes a turbine driven by hot gas, a compressor rotating with the turbine to generate compressed air, an annular combustor coaxial with the turbine to combust fuel and compressed air to generate the hot gas, and an annular recuperator to recover heat from the turbine exhaust gas and heat the compressed air for combustion. The annular recuperator surrounds the turbine and includes two contiguous parts made from two materials having different thermal properties and joined to one another to form a single annular structure. One recuperator part is formed from a high-temperature material having a high thermal limit for exposure to high-temperature turbine exhaust gas, and the other recuperator part is formed from a material having a lower thermal limit than the high-temperature material for exposure to reduced-temperature turbine exhaust gas.

Abstract: The expense of fabricating an annular combustor (10) for a gas turbine is minimized by providing a combustor housing (12) including a generally axially extending sleeve (14) and an annular liner (18) disposed within the housing (12) and about the sleeve (14). The annular liner (18) has concentric inner and outer axially extending walls (20, 22) spaced from the sleeve (14 ) and the housing (12), respectively, and also has a radially extending wall (24) spaced from the housing (12 ) and interconnecting the inner and outer walls (20, 22) at one end to define a combustion chamber (26). The liner (18) is spaced from the housing (12 ) and the sleeve (14 ) to define a compressed air flow path (28) extending from a radially outer compressed air inlet (30) to a radially inner compressed air outlet (32) in communication with the combustion chamber (26) axially remote from the radially extending wall (24).

Abstract: The invention is directed to a compressor that recirculates bleed air by way of a fluid distributing unit which is adapted to redirect and uniformly distribute the air bled through passages in a compressor shroud section across a cross-sectional area of a compressor's inlet region to provide uniform distribution of the bleed air into the compressor's impellar to thereby increase efficiency of the compressor.

Abstract: A method of starting a gas turbine (10) having a combustor (12) and a fuel supply system (14) supplying fuel to the combustor in accordance with the invention includes turning on a fuel valve (40) to control fuel flow to the combustor for combustion after a first starting condition (100) is reached; and cyclically turning (116) the main fuel valve on and off when a sensed exhaust gas temperature (114) initially reaches a first set temperature and continuing to cyclically turn the main fuel valve on and off until the sensed exhaust gas temperature falls to a second set temperature (118) lower than the first set temperature or until a second starting condition (102) is reached.

Abstract: The invention is directed to a method of restoring compressor rotor blade/compressor shroud clearance brought on by cold air induced shroud contraction by delivering warmed bleed air through a geodesic array of compressor/turbine support struts which counteracts the cold air induced shroud contraction.

Abstract: Combustors of gas turbine engines in which components are brazed to the combustor are not easily repairable by conventional welding because the brazed joint includes a melting point depressant which may not only cause the joint to be brittle but also the melting point of the joint to be near or below the temperatures required for welding. These problems are solved by a relatively simple and inexpensive process for producing a weld repairable combustor which includes using a brazing powder mix comprising a metal powder compatible with the metals of the combustor and the article to be brazed to the combustor, and suitable melting point depressant in powder form, heating the brazing powder mix to a first temperature sufficient to cause the brazing powder mix to become molten, and finally heating the brazing powder mix at a second temperature for a sufficient time to cause the melting point depressant to substantially diffuse away from the brazed area.

Abstract: An improved tip cap configuration for a turbomachinery blade includes tip cap seats formed integrally with and located on the inner sides of a hollow turbomachinery blade. Tip cap retaining members are also formed integrally with the inner wall of the blade and are spaced a short radial distance from the tip cap seats. The blade is initially provided with a flared open end and the tip caps are positioned on the seats and the flared end is crimped so as to capture the tip cap between the seats and the retaining members. The crimping operation is stopped at a point wherein a smooth outer contour is obtained for the blade in the region of the tip cap. A method of making such a blade is also disclosed.