Lockheed Corporation

Lockheed Aircraft Corporation

Lockhead Corporation

Michael J Cronin

Gordon Seid

Victor W Hamra

The multifunction power system is a combination auxiliary power unit, environmental control system, engine start system, and emergency power system for an aircraft. The system is essentially a four-wheel auxiliary power unit comprising two compressors driven by one turbine by means of a common shaft and in addition a free turbine. The pressurized flow from the second compressor can be directed to parallel combustors and thereafter to the two turbines. A starter-generator and motor are coupled to the compressor drive shaft and the free turbine is used to power a generator and for engine starting. Ducts, heat exchangers, valves, etc., are so arranged as to provide the four functions by proper selection and routing of air.

Multifunction power system for an aircraft

The invention is a two-level power system for electrically powered aircraft sub-systems. An aircraft engine (2) drives a three phase generator (30) at a variable speed. The output of generator (30) is primary variable-voltage/variable-frequency (VV-VF) power which is utilized by various aircraft subsystems. A portion of the primary VV-VF power is converted to constant voltage-constant frequency (CV-CF) power by an inverter (31) which includes front end rectification. The CV-CF power is then available to selected aircraft subsystems while portions of the CV-CF power are converted by rectifier (32) and transformer-rectifier (33) into 270 vdc and 28 vdc power, respectively, which power is also then available to aircraft sub-systems.

Aircraft providing variable and constant electric power

An actuator system for controlling the position of an aerodynamic control surface of an aircraft is disclosed. The actuator system (10) comprises a rotary actuator (12) rotatably mounted to the aircraft structure (20) with the actuator having an output member (22) adapted to be coupled to the control surface. A latch assembly (37) is provided which is mounted to the structure which releasably engages the actuator. Preferably, the latch assembly (37) comprises a grooved flange (38) mounted on the actuator (12) and a solenoid (47) mounted to the aircraft structure which is adapted to move a locking pin (44) from a first position in engagement with the grooved flange (38) to a second position out of engagement with the groove. In a second embodiment, a second rotary actuator (114) is mounted to the aircraft structure and is adapted to rotate the actuator (81) upon its failure and after the latch assembly is activated. In a third embodiment a second rotary actuator (164) is coupled to the input shaft (160) of the actuator (150) and after the latch assembly is activated rotation of the actuator (150) is accomplished by the second rotary actuator (164) via the input shaft. In addition, should the actuator assembly fail electrically, and not be jammed, the second actuator (164) drives the input shaft which in turn drives the output member and thus the latch assembly need not be activated.

Electromechanical actuator system with anti-jamming features

The invention is an ECS system designed to utilize an induction motor (12) and a cabin compressor (14) for providing pressurized air that serves as a source of energy for heating, cooling, pressurizing and other air needs of modern aircraft. An aircraft engine driven generator, preferably a permanent-magnet generator (of the samarium-cobalt, SmCo, type), furnishes variable-voltage/variable frequency, VV/VF, power to the motor (12). The electric motor itself in the preferred embodiment is a highly reliable squirrel-cage induction motor (12) powered directly from the VV/VF supply, with the result that its speed is proportional to the generator frequency, which in turn is dependent on the aircraft engine speed. The main cabin compressor (14) and a secondary (auxiliary) supercharging compressor (18) are both mounted on the motor shaft. The motor-compressor configuration is further characterized in that the ac motor has pole-changing winding (typically two), which permits two speed operation, while the cabin compressor is provided with inlet guide vanes (16). The interactive combination of these two features makes it possible to furnish the requisite air mass flow and pressurization levels over the conditions of variable air density and variable atmospheric pressure, as the airplane operates throughout its flight envelope.

Energy efficient ECS powered by a variable voltage/variable frequency

An actuator system for controlling the position of an aerodynamic control surface of an aircraft, such as an aileron, is disclosed. An actuator (12) having an output shaft (14) is mounted to a frame attached to the aircraft structure. A hydraulic coupling (32) is mounted on one side to the output shaft (14) of the actuator (12) and the other side to the control surface (50) and comprises a hydraulic cylinder (34) having a sealing member (36) movably mounted therein dividing the cylinder (34) into two portions (38, 40). The output shaft (14) of the actuator (12) is coupled to the cylinder (34) and an output shaft is attached to the sealing member (36) at one end and at its opposite end to the control surface or vice versa. A connecting tube (54) is provided coupling the two portions of the hydraulic cylinder together. The tube incorporates a valve (58) having a first position sealing off the two portions from each other and a second position providing a passageway therebetween. Thus, if the actuator (12) is working and the valve (56) is closed, an essentially solid connection is achieved between the actuator (12) and the control surface (50), thus allowing the actuator to move the control surface upon command. Should the actuator fail, the valve (56) can be opened, allowing fluid to flow from portion to portion. Thus, a redundant actuator can continue to control the position of the control surface while the disabled actuator is effectively decoupled from the control surface for the piston (36) will just move back and forth within the cylinder (34) as the control surface moves.

Electric actuator system with hydraulic coupling as protection against

The invention is an electrical connector assembly (10) for coupling flat cable conductors (60) to electrical terminals (40). The invention comprises a receptacle (12) having a dielectric surface (31) with a pattern of first contact pads (33) extending therefrom coupled to the electrical terminals (40). A flat conductor cable (60) is provided having a pattern of second contact pads (64) mounted to a lower side (63), which are in a geometric pattern corresponding to the pattern of the first contact pads (33). A pressure means (72, 76) is provided to force the first and second contact pads (33, 64) into intimate contact, thereby electrically connecting the flat cable conductors (67) to the electrical terminals (40). The pressure means preferably comprises a resilient compression member (72) movably attached to a screw cap cover (76) which is adapted to engage the receptacle (12), such that upon the engagement of the cover (76) with the receptacle (12) the resilient compression member ( 72) forces the first and second contact pads (33, 64) together electrically connecting the flat and round cable conductors (67, 44).

Electrical connector assembly for flat cables

The invention is an all electric environmental control system. The primary power for the system is provided in a first embodiment by a compressor (C.sub.1) driven by an electric motor (14). The system includes the primary heat exchanger (16), an expansion cooling system (20), and a water separator (26). The electric motor (14) may be utilized in another embodiment to drive a compressor (23) within the expansion cooling system (20'). A vapor-cycle cooling system including the freon-compressor (34), the condenser (36) and the evaporator (38) may optionally replace the expansion cooling system (20). To compensate for pressure variations during flight, the electric motor (32) is mounted to a planetary gearbox (44) with a 'makeup'/'subtract' differential electric motor (42) such that the compressor (C.sub.1) is run at the desired speed. The compressor (C.sub.1) is provided with inlet guide vanes (56) which are controlled by the flow meter (52) and flow logic circuit (54) to provide the desired flow rate of air to the environmental control system.

All electric environmental control system for advanced transport aircraft

A cabinet assembly or enclosure which provides an electronically and thermally benign environment for printed circuit boards and the electronic components carried thereby. A wire program device operable from the front of the cabinet permits a changeable interface between the circuits within the cabinet and an external wiring harness. Thermal management for plural solid state power controllers or other heat generating components is achieved through the wire program, and from the use of metal printed circuit boards clamped to a cold plate within the cabinet. A coolant from a remote source may be circulated through the cabinet in contact with the cold plate. The electrical, thermal and mechanical connections are of the zero insertion force type. A novel circuit board for solid state power controllers is also disclosed as is the thermal management of such controllers.

Electronic circuit board and method and apparatus for thermal management

The invention is a multi role primary/auxiliary power system with engine start capability for aircraft that provides electric, mechanical and air conditioning power as well as engine starting functions. The compressor (20), turbine (26) and motor/generator (24) are coupled to a drive shaft (14) of the aircraft's engine (12) with the motor/generator (24) mounted between the compressor (20) and turbine (26). A first clutch (28) is mounted on the drive shaft (14) between the motor/generator (24) and the turbine (26). Coupled to the outlet (29) of the compressor (20) is a two-way valve (34) adapted to couple the compressor (20) to either a first duct (36) or second duct (37). Both of these ducts (36), (37) are coupled at their opposite ends to the turbine (26). Mounted in the first duct (36) is a combustor (40) and in the second duct (37) a first heat exchanger (42). A third duct (44) is also coupled between the compressor (20) and the turbine (26) and which incorporates a valve (50). Mounted between the turbine (26) and the engine (12) is a gear reduction assembly (58). Coupled to the output of the turbine (26) is a fourth duct (100) which is coupled to a second two-way valve. The second two-way valve (102) either couples the fourth duct (100) to an overboard vent (104) or a fifth duct (106), having a heat exchanger (108) mounted therein to the cabin of the aircraft. Coupling the first, second and third ducts (36), (37), and (44) together is a bypass duct (110) (sixth duct) which bypasses the turbine (26) and connects to the fourth duct (100). The sixth duct (110) also incorporates a simple open/close valve (112). Finally, a control circuit (115) for the motor/generator is provided which can couple either onboard battery power or ground supply AC power to the motor/generator.

Multi role primary/auxiliary power system with engine start capability

A quick attach/detach instrument panel and method in which the panel may be both mechanically and electrically connected to the support frame by zero insertion force connectors and in which the individual instruments on the panel may be easily removed from the front of the panel and electrically and mechanically connected thereto by zero insertion force connectors.

Quick attach detach instrument panel method

The invention is an 'electric' constant speed/variable speed drive/generator assembly (10), which finds use in aircraft, and in particular, in 'all-electric' aircraft. The assembly (10) includes a planetary differential (20) whose output is controlled by a 'make-up' induction motor (12). The induction motor (12) includes a squirrel-cage rotor (14) which is rotatable about a rotor-shaped stator (16). The assembly also includes a primary power generator (30) which includes a permanent magnet rotor (18) and an output winding (22). The rotor (18) is co-axial with and rotatable about, the outside of the rotor (14). The outer rotor (18) is driven by the planetary differential (20) in such a manner that the planet-carrier (24) is driven by the aircraft's engine, the ring-gear (26) drives the generator (30), and the sun-gear (32) is connected to the induction motor (12). Constant speed operation of the device is effective when the band-brake (34) is 'off' and the induction motor (12) rotates at a speed, and in a direction, to compensate for the in-flight variations of the engine input-speeds to the planetary differential (20). When the band-brake (34) is 'on,' the sun-gear (32) is locked and the differential acts as a fixed ratio transmission. The result is a dual-speed range all-electric drive which enables the generator (30) to provide constant voltage/constant frequency or variable voltage/variable frequency power (proportional to engine speed).

Electric constant speed/variable speed drive/generator assembly

The invention is an energy-efficient all-electric environmental control system (70). The system includes a cabin compressor (16) and a Freon compressor (14), both of which are commonly driven by a two-speed electric motor (12). The motor (12) operates at low speed at low altitudes and on the ground to drive the Freon compressor (14). Freon fluid then cools the electric motor (12) and the cabin (18) and avionics (56) by means of the evaporator/dehumidifier systems (72) and (74), respectively. At high altitudes the electric motor (12) runs at high speed to drive the cabin compressor (16) to provide pressurization for the cabin (18). In this operation mode, the Freon compressor can be disconnected via an electromagnetic clutch (48). The ECS (70) also includes a liquid Freon pump (52) which can be utilized when the Freon compressor is disconnected or bypassed to circulate liquid Freon to cool the electric motor (12) and to circulate through the evaporators (38) and (76). The pump (52) is driven by an ac or dc motor (54).

Energy-efficient all-electric ECS for aircraft

The invention is an aircraft power generation system which includes a tandem generator (10) which includes left and right generators (14, 15) arranged within a single housing (12). The left and right generators (14, 15) are driven at the anti-drive end of the tandem generator (10) by a gear arrangement including gears (22), (24) and (26). A phase indexer (36) including a phase adjust nut (38) is arranged in conjunction with driven gear (26) to provide for 'vernier' phase-alignment of the generator rotors. In an alternative embodiment, phase-alignment of the generator rotors is accomplished by arranging them in-line to form an in-line tandem generator (10'). In a preferred arrangement, an in-line generator (70) is arranged to be isolated from an aircraft engine via an electromechanical disconnect (108). The disconnect (108) includes a pin (130) activated by a solenoid (110), which upon receipt of a disconnect command or signal, pulls the pin (130) from a hole (131) in ring gear (114) thereby disconnecting the tandem generator (70).

Tandem-generator design for aircraft

The invention relates to a starter-generator machine (10) for starting turbine type aircraft engines. The machine combines an induction motor with a synchronous samarium cobalt generator. In the machine, a rotor-shaped stator (22) is fixed and positioned inside a squirrel-cage induction rotor (18) which has an array of samarium-cobalt magnets (24) attached on the outer diameter thereof. The compound dual machine operates as a starter by using the induction rotor to accelerate the permanent magnet rotor, and thus the aircraft engine via a drive pinion (34), up to some low synchronous speed, when ac power is applied to the outside stator (16) to lock in the permanent-magnet rotor (24) synchronously with the rotating field created in the stator (16) of the synchronous generator. As the speed of the rotor (24) is then increased, the engine speed is also increased via the drive pinion (34). In a second embodiment, a cartridge type induction-motor (70) is utilized to initially start an aircraft engine. The motor (70) includes a gear reduction (74) and a disconnect clutch (76) and drives the engine through a splined pinion (78) which in turn drives an engine connected spline (98). When the speed of the machine is such that the synchronous operation of an outside permanent magnet rotor (96) commences, the clutch (76) is utilized to disconnect the induction motor cartridge (70), leaving the rotor (96) to drive the drive spline (98) via internal splines (94).

Growing community of inventors

Sherman Oaks, CA, United States of America

Michael J Cronin