Abstract: Improved fuel atomization for turbine engines utilizing high viscosity fuels is accomplished in an engine having an annular combustor (26) by providing an air injector (52) which produces an air stream (50) generally tangentially directed into the combustor (26) to intersect a fuel spray or film (46) also generally tangentially directed into the combustor (26) by a fuel injector (48) to atomize fuel for ignition by an igniter (60) in an annular outer wall (34) of the combustor (26).

Abstract: Difficulties in achieving reliable starts in gas turbine engines operating at high altitude are avoided in a gas turbine engine including a rotary compressor (10), a rotary turbine wheel (12) coupled to the compressor (10) to drive the same, and a nozzle (42) for directing gases of combustion against the turbine wheel (12). An annular combustor (34) has an outlet (40) connected to the nozzle (42) and an opposed dome (38) that is axially spaced from the outlet (40). At least three sets of air injection openings (80, 92, 100) are axially spaced from one another with one set (82) in close proximity to the dome (38). Fuel injectors (86, 94) are associated with two of the sets (80, 92, 100), including the one set (80) and another of the sets (92) that is nearest the one set (80).

Abstract: Difficulties with fuel atomization at low fuel flow rates in gas turbine fuel injection systems may be overcome by use of an impingement injector (52) including a cylindrical orifice (64, 94, 114) that is smooth, straight and uninterrupted by burrs or other disturbances having an exit opening (68, 96, 118) transverse to the axis of the orifice (64, 94, 114) and facing an impingement surface (72, 101, 126) that is planar and spaced from the exit opening (68, 96, 118).

Abstract: In order to efficiently atomize a fuel/air mixture to reduce smoke, a fuel injection system (26) for a combustor (12) of a turbine engine (10) includes an air blast tube (16), a fuel supply passage (42), and an impingement surface (30). The air blast tube (16) is mounted in relation to an opening (18) in a wall (20) of the combustor (12) such that a first end (22) of the tube (16) is in communication with the combustor (12) and a second end (24) of the tube (16) is in communication with a source of compressed air externally of the combustor (12) to facilitate the delivery of compressed air from the source into the combustor (12).

Abstract: In order to enhance starting and flame propagation at a high altitudes, particularly in small turbine engines such as auxiliary power units and the like, the turbine engine (10) includes a rotor (12), turbine blades (16) a turbine nozzle (20), an annular combustor (22), a fuel injector (40), and a compressed air bleed system (44). The annular combustor (22) is disposed about the rotor (12) and has spaced inner and outer walls (28 and 30) interconnected by a generally radially extending wall (32). A housing (34) substantially surrounds the annular combustor (22) to define an air flow path (36) including a compressed air inlet (38) in communication with a source of compressed air at one end thereof wherein the air flow path (36) extends from the compressed air inlet (38) at least part-way about the annular combustor (22).

Abstract: The high cost of fuel injection nozzles for a gas turbine engine may be reduced by utilizing nozzles (60) formed of simple tubes (63) of capillary cross section having an integral flange (62) formed on one end (86) thereof.

Abstract: Difficulties in starting a gas turbine engine at high altitudes may be avoided by the method of starting such an engine which includes the steps of introducing fuel into the combustor (26) of the engine through at least one fuel injector (46) having an outlet (76, 78) within the combustor (26), atomizing the fuel with a burst of oxidant at elevated pressure from a source (92) of pressurized stored oxidant until ignition is obtained, and thereafter discontinuing the burst of oxidant from the source (92).

Abstract: Difficulties with starting reliability and/or combustion stability in hot gas generating systems are avoided in a construction including a storage vessel (36) for storing an oxidant, a fuel supply (22) and combustion chamber (26) having an out outlet (30) for hot gas. A fuel injection nozzle (28) provides fuel from the supply (22) to the combustion chamber (26) while an oxidant duct (44) connects the vessel (36) to the combustion chamber (26). A choked orifice (46) is disposed in the duct (44) just upstream of the combustion chamber (26) and a bypass duct (50) interconnects the combustion chamber outlet (30) and the oxidant duct (44) between the choke orifice (46) and the combustion chamber (26). A selectively operable valve (53) is disposed in the bypass duct (50) and a pressure regulator (44) is disposed in the oxidant duct (44) for selectively controlling the pressure of oxidant applied to the choked orifice (46).

Abstract: In order to establish a larger dome height to thereby increase combustion volume within the same turbine engine envelope, the inner wall region of an annular combustor is formed to define a thrust nozzle which in turn provides an enhanced combustion zone. The turbine engine will include a rotary compressor and a turbine wheel coupled to the rotary compressor for driven movement thereof along with an annular nozzle proximate the turbine wheel for directing gases of combustion thereat. The annular combustor will be disposed about the turbine wheel and defined by a radially extending wall region connecting the inner wall region to an outer wall region such that the annular combustor has a combustion zone upstream of the annular nozzle and a dilution air zone intermediate the annular nozzle and the combustion zone.

Abstract: The diameter of a gas turbine engine may be reduced through the use of a construction including a monorotor (44) journaled for rotation and mounting compressor blades (48) at one end (46) and turbine blades (56) at an opposite end (54). An axial diffuser (96) receives the discharge from the compressor blades (84) and is contained in a plenum (98). A reverse flow, annular combustor (83) is disposed about the rotor (44) and has a dome (100) whose height is a minor fraction of the axial length of the combustor (83) and which discharges to an annular turbine nozzle (77) having axial turbine nozzle blades (76).

Abstract: Difficulties in achieving reliable starts in gas turbine engines operating at high altitude are avoided in a gas turbine engine including a rotary compressor (10), a rotary turbine wheel (12) coupled to the compressor (10) to drive the same, and a nozzle (42) for directing gases of combustion against the turbine wheel (12). An annular combustor (34) has an outlet (40) connected to the nozzle (42) and an opposed dome (38) that is axially spaced from the outlet (40). At least three sets of air injection openings (80, 92, 100) are axially spaced from one another with one set (82) in close proximity to the dome (38). Fuel injectors (86, 94) are associated with two of the sets (80, 92, 100), including the one set (80) and another of the sets (92) that is nearest the one set (80).

Abstract: To control cooling by oxidant surrounding a combustion chamber (20, 20'), a stored energy combustor (10, 10') includes an oxidant flow path (28) which extends about the combustion chamber (20, 20'). The stored energy combustor (10, 10') comprises a vessel (12, 12') having a wall (14, 14') defining spaced inlet and outlet ends (16 and 18) interconnected by the combustion chamber (20, 20'), an oxidant inlet port (22) upstream of the combustion chamber (20, 20') for directing oxidant thereinto, and a fuel injector (24) upstream of the combustion chamber (20, 20') for directing fuel thereinto. A housing (26) substantially surrounds the vessel (12, 12') in spaced relation to the wall (14, 14') from the inlet end (16) to the outlet end (18) whereby the housing (26) and wall (14, 14') define the oxidant flow path (28) therebetween which is in communication with the oxidant inlet port (22) upstream of the combustion chamber (20, 20').

Abstract: In order to facilitate operation in a wide variety of operating conditions, and to enhance ultra high altitude starting capability while eliminating start injectors, a radial turbine engine (10) includes a pair of fuel injection zones (24 and 26). The radial turbine engine (10) also includes a turbine wheel (12) coupled to a rotary compressor (14) for axially driven movement, an annular nozzle (16) for directing gases of combustion radially at the turbine wheel (12), and an annular combustor (17). The annular combustor (17) defines an annular combustion space disposed about the turbine wheel (12) and in fluid communication with both the compressor (14) and the nozzle (16), and it receives fuel from a source and air from the compressor (14) which are combusted in the combustion space to generate the gases of combustion.

Abstract: In order to provide a simple, low-cost fuel injection system, particularly for a dual zoned annular combustor (10), a single fuel manifold (42) is utilized having first and second manifold portions (42a and 42b) each with a respective plurality of openings (44a and 44b) through which fuel may flow into first and second fuel injection zones (34 and 36). The annular combustor (10) typically includes a combustor housing (12) comprising an axially extending sleeve (14) disposed about a longitudinally extending axis (16) and an annular liner (18) defining a combustion chamber (26) and being spaced from the housing (12) and the sleeve (14) to form a compressed air flow path (28) therebetween.

Abstract: Difficulties in cooling the shrouds (32, 34) and the nozzle (38) of an air breathing gas turbine are minimized or eliminated through the use of a generally toroidal, annular combustor (44) having an outlet (48) and combustion air inlets (90A, 90B) and otherwise being imperforate or free of any means for the admission of cooling air to the interior of the combustor (44).

Abstract: In order to establish a larger dome height to thereby increase combustion volume within the same turbine engine envelope, the inner wall region of an annular combustor is formed to define a thrust nozzle which in turn provides an enhanced combustion zone. The turbine engine will include a rotary compressor and a turbine wheel coupled to the rotary compressor for driven movement thereof along with an annular nozzle proximate the turbine wheel for directing gases of combustion thereat. The annular combustor will be disposed about the turbine wheel and defined by a radially extending wall region connecting the inner wall region to an outer wall region such that the annular combustor has a combustion zone upstream of the annular nozzle and a dilution air zone intermediate the annular nozzle and the combustion zone.

Abstract: Complexity in a fuel injection system for a gas turbine engine including an annular combustor (26) may be reduced by utilizing a plurality of circumferentially spaced fuel injectors (46) within the annular combustion (26). First and second conduit branches (108, 110; 120, 122) in fluid parallel with one another are provided with one of the branches (108; 120) being connected to at least one of the injectors (46-1; 46-1, 46-3, 46-5) and the other of the branches (110; 122) being connected to the remainder of the injectors (46-2 through 46-6, inclusive; 46-2, 46-4, 46-6). A pump (102) interconnects a source of fuel (100) and the branches (108, 110; 120, 122). A normally closed, pressure responsive valve (114) isolates the other branch (110, 122) from the source (100) at low pressures typical of the starting of the turbine engine and is openable at somewhat higher pressures to allow fuel to flow to all of the injectors (46).

Abstract: In order to control engine operation by exhaust gas measurement, while avoiding operating conditions that exceed the safe maximum temperature limit, a gas turbine engine (10) includes a thermocouple temperature probe (28). The gas turbine engine (10) has a rotary compressor (12), a turbine wheel (14) coupled to the compressor (12) for driven movement thereof, and an annular nozzle (16) proximate the turbine wheel (14) for directing gases of combustion at the turbine wheel (14). It also includes an annular combustor (18) defining an annular combustion space (20) disposed about the turbine wheel (14) and in fluid communication with both the compressor (12) and the nozzle (16). The combustor (18) receives fuel from a source and air from the compressor (12) and it combusts fuel and air in the combustion space (20) to generate the gases of combustion.

Abstract: In order to facilitate operation in a wide variety of operating conditions, and to enhance altitude starting capability while eliminating start injectors, a radial turbine engine (10) includes dual fuel injection zones (24 and 26). The radial turbine engine (10) also includes a turbine wheel (12) coupled to a rotary compressor (14) for axially driven movement, an annular nozzle (16) for directing gases of combustion radially at the turbine wheel (12), and an annular combustor (15). The annular combustor (15) defines an annular combustion space disposed about the turbine wheel (12) and in fluid communication with both the compressor (14) and the nozzle (16), and it receives fuel from a source and air from the compressor (14) which are combusted in the combustion space to generate the gases of combustion.

Abstract: In order to optimize ignition capabilities, while at the same time enhancing the life span of an igniter, a stored energy combustor (10) has an igniter (34) disposed within a fuel injector (26). The fuel injector (26) is disposed in a tubular extension (16) leading to a combustion chamber (20) and has a discharge end (28) communicating with a point of entry into the combustion chamber (20). An oxidant inlet port (24) is provided upstream of the combustion chamber (20) for directing oxidant into the combustion chamber (20). The oxidant inlet port (24) directs oxidant through the tubular extension (16) leading to the combustion chamber (20) in surrounding relation to the fuel injector (26). With this arrangement, the igniter (34) is disposed within the fuel injector (26) along a longitudinal axis (22) of the stored energy combustor (10).