Abstract: A method of controlling a rotary wing aircraft includes accelerating the aircraft in a fore direction independently of cyclic control of the main rotor and pitch of the aircraft.

Abstract: A method for controlling rotor blades of a co-axial rotor assembly of an aircraft including a first rotor co-axial with a second rotor includes identifying a first zone of rotor rotation angles of the co-axial rotor assembly. The first zone defines a range of rotor rotation angles corresponding to an up-flow of air to the coaxial rotor assembly, and the remainder of the rotor rotation angles other than the first zone of rotation angles is defined as a second zone. The method includes receiving a yaw command to adjust a yaw moment of the aircraft and applying a different rotor blade angle change to rotor blades in the first zone than a rotor blade angle change applied to rotor blades in the second zone to adjust the yaw moment of the aircraft according to the yaw command.

Abstract: A propulsor system for a rotary winged aircraft includes a propeller including having a propeller hub rotatable about a propeller axis and a plurality of propeller blades secured to and extending radially outwardly from the propeller hub. The propeller is oriented on the rotary winged aircraft to provide forward thrust for the rotary winged aircraft when the propeller is rotated about the propeller hub. A plurality of guide vanes non rotatable about the propeller axis are positioned upstream of the plurality of propeller blades. The plurality of guide vanes are positioned to turn an airflow flowing into the plurality or propeller blades, thereby reducing an angle of attack of the plurality of propeller blades relative to the airflow.

Abstract: A main rotor system of an aircraft is provided including a first rotor coupled to a transmission and configured to rotate about an axis in a first direction. A second rotor is similarly coupled to the transmission and is configured to rotate about the axis in a second direction. At least the first rotor includes an individual blade control system (IBCS) configured to adjust a pitch of each of a plurality of blade of the first rotor independently. A standpipe is fixedly attached to the aircraft. The standpipe is arranged such that the first rotor and the second rotor rotate relative to the standpipe. At least one slip ring is configured to transmit electrical power and/or a control signal to the at least one IBCS.

Abstract: A method and system for controlling maneuverability of an aircraft includes receiving one or more signals indicative of commanded peak rotary acceleration at a first timeperiod; determining a signal indicative of an actual peak rotary acceleration for the first timeperiod in response to the receiving of the one or more signals for commanded pilot acceleration; and determining signals indicative of actual rotary acceleration for a second timeperiod.

Abstract: A method for controlling rotor blades of a co-axial rotor assembly of an aircraft including a first rotor co-axial with a second rotor includes identifying a first zone of rotor rotation angles of the co-axial rotor assembly. The first zone defines a range of rotor rotation angles corresponding to an up-flow of air to the coaxial rotor assembly, and the remainder of the rotor rotation angles other than the first zone of rotation angles is defined as a second zone. The method includes receiving a yaw command to adjust a yaw moment of the aircraft and applying a different rotor blade angle change to rotor blades in the first zone than a rotor blade angle change applied to rotor blades in the second zone to adjust the yaw moment of the aircraft according to the yaw command.

Abstract: A main rotor system of an aircraft is provided including a first rotor coupled to a transmission and configured to rotate about an axis in a first direction. A second rotor is similarly coupled to the transmission and is configured to rotate about the axis in a second direction. At least the first rotor includes an individual blade control system (IBCS) configured to adjust a pitch of each of a plurality of blade of the first rotor independently. At least one slip ring is configured to transmit electrical power and/or a control signal to the at least one IBCS.

Abstract: An exhaust system for reducing infrared emissions of a rotary wing aircraft includes a duct assembly having an inlet portion and an outlet portion; the inlet portion configured to receive exhaust from an engine of the aircraft; and the outlet portion coupled to the inlet portion, the outlet portion having an outlet duct with an outlet opening, the outlet duct configured expel an emission containing engine exhaust proximate to a tail fairing of the rotary wing aircraft.

Abstract: A tailwheel assembly for an aircraft includes a ventral fin including a fixed airframe portion of the ventral fin and a moveable tailwheel portion of the ventral fin. A tailwheel strut extends through the ventral fin and includes a fixed first strut portion and an extendable second strut portion having the moveable tailwheel portion of the ventral fin secured thereto. A tailwheel is secured to the extendable second strut portion. A method of operating a tailwheel assembly for an aircraft includes extending an extendable portion of a tailwheel strut from an airframe of the aircraft. The tailwheel strut has a tailwheel secured thereto and is located at least partially in a ventral fin of the aircraft. A tailwheel portion of the ventral fin is moved away from a fixed airframe portion of the ventral fin via extension of the extendable portion of the tailwheel strut.

Abstract: A tailwheel assembly for an aircraft includes a ventral fin including a fixed airframe portion of the ventral fin and a moveable tailwheel portion of the ventral fin. A tailwheel strut extends through the ventral fin and includes a fixed first strut portion and an extendable second strut portion having the moveable tailwheel portion of the ventral fin secured thereto. A tailwheel is secured to the extendable second strut portion. A method of operating a tailwheel assembly for an aircraft includes extending an extendable portion of a tailwheel strut from an airframe of the aircraft. The tailwheel strut has a tailwheel secured thereto and is located at least partially in a ventral fin of the aircraft. A tailwheel portion of the ventral fin is moved away from a fixed airframe portion of the ventral fin via extension of the extendable portion of the tailwheel strut.

Abstract: An interactive aircraft load management system automates calculation and provides simulation capability to changes in an aircraft C.G. limit for display on a three-dimensional aircraft symbology. The aircraft load management system also communicates with a fly by wire (FBW) flight control system wherein the aircraft's control system is programmed to automatically compensate for C.G. excursions and to alter control laws. The aircraft load management system also selectively reels-in and reels out sling lines as the aircraft pitches and rolls to maintain a load vector from a slung load along the aircraft centerline. Load vector travel is accomplished by coupling a winch control system into the flight control system.

Abstract: An interactive aircraft load management system automates calculation and provides simulation capability to changes in an aircraft C.G. limit for display on a three-dimensional aircraft symbology. The aircraft load management system also communicates with a fly by wire (FBW) flight control system wherein the aircraft's control system is programmed to automatically compensate for C.G. excursions and to alter control laws. The aircraft load management system also selectively reels-in and reels out sling lines as the aircraft pitches and rolls to maintain a load vector from a slung load along the aircraft centerline. Load vector travel is accomplished by coupling a winch control system into the flight control system.

Abstract: A system for remote payment of utility usage including a customer terminal and a utility provider server wherein the customer terminal includes a utility meter for collecting utility usage data and an input device configured to receive customer account information. At any time during a billing cycle, a customer may prompt the customer terminal to make payment by inserting account information into the customer terminal. The utility provider server may communicate with the customer terminal to transmit and receive customer account information, payment information, usage data, rate data, and other information.

Abstract: A ducted fan antitorque device for a helicopter having an empennage structure including a shroud configured for internal mounting of the ducted fan antitorque device. The ducted fan antitorque device has a configuration optimized to provide the antitorque thrust required for helicopter yaw stability and maneuverability and to minimize the operating noise levels of the ducted fan antitorque device. The ducted fan antitorque device includes an airflow duct and a fan assembly coaxially mounted within the airflow duct. The airflow duct includes an inlet having a curved lip configuration of constant radius, a divergent duct portion downstream of and contiguous with the inlet having a predetermined divergence angle, and an outlet terminating the divergent duct portion having a curved lip configuration of variable radius.

Abstract: A shroud-fin integration shelf for a helicopter empennage structure that includes a shroud embodying a ducted fan antitorque device, a vertical stabilizer contiguous with the shroud, and a horizontal stabilizer intersecting the vertical stabilizer. The shroud-fin integration shelf is structurally configured and dimensioned to eliminate adverse aerodynamic effects experienced by the vertical stabilizer in reverse thrust operating conditions wherein mass airflow exiting the ducted fan antitorque device may flow along the shroud and interfere with longitudinal mass airflow over the suction surface of the vertical stabilizer. The integration shelf, which provides an acute transition between the shroud and the vertical stabilizer, has an overall predetermined width that includes a first shelf width on the ducted fan antitorque device inlet side of the empennage structure and a second shelf width on the opposite side of the empennage structure.

Abstract: An integrated empennage structure is provided for a helicopter embodying a ducted fan antitorque device. The integrated empennage structure includes a shroud that houses the ducted fan antitorque device, a vertical stabilizer, and a horizontal stabilizer. The aerodynamic configurations and/or orientations and the spatial orientations of the shroud, vertical stabilizer, and the horizontal stabilizer are interactively related and optimized to provide yaw stability and directional control and pitch stability and maneuverability required for helicopter flight operations. The shroud is spatially orientated at a first predetermined cant angle with respect to the vertical plane of symmetry of the helicopter fuselage such that the ducted fan antitorque device provides lateral antitorque thrust to counteract main rotor assembly induced torque and a positive vertical force component to enhance the overall lift capability of the helicopter.