Abstract:
A rotary wing aircraft includes a main rotor; an electric motor for rotating the main rotor; an electric generator for supplying electric power to the electric motor; an engine for driving the generator; and battery storage for providing battery power. The aircraft further includes a flight control system for controlling the engine to run at idle and causing the electric motor to receive the battery power to rotate the main rotor during takeoff; for controlling the engine to increase speed above idle and operate the generator to recharge the battery storage during flight; and for controlling the engine to return to idle and controlling the electric motor to receive the battery power for landing.

Description:
[0001]    This is a divisional of copending U.S. patent application Ser. No. 13/916,283 filed 12 Jun. 2013. 
     
    
     BACKGROUND 
       [0002]    Conventional rotary wing aircraft or helicopters are powered with either an internal combustion engine or a turbine engine that is coupled to a gear box system which transmits the rotational output of the engine to the main rotor of the aircraft and to the tail rotor of the aircraft. The gear boxes of rotary wing aircraft require regular monitoring and periodic maintenance to ensure their reliable operation. Gear boxes do not have graceful failure modes and if one piece should fail it is likely the entire gear box will fail. Gear boxes are also expensive to manufacture. Additionally, the tail rotor drive transmission of a rotary wing aircraft is complex, especially for helicopters having an air frame with a folding tail for ground storage. These tail rotor drive transmissions also require regular monitoring and periodic maintenance to guard against failures. The main rotor gear box and the tail rotor drive transmission of a rotary wing aircraft also employ hydraulic actuators and gear box driven hydraulic systems to control the main rotor gear box and tail rotor drive transmission. These hydraulic control systems add to the weight of the aircraft and reduce its operation efficiency. 
       SUMMARY 
       [0003]    A rotary wing aircraft comprises a main rotor; an electric motor for rotating the main rotor; an electric generator for supplying electric power to the electric motor; an engine for driving the generator; and battery storage for providing battery power. The aircraft further comprises a flight control system for controlling the engine to run at idle and causing the electric motor to receive the battery power to rotate the main rotor during takeoff; for controlling the engine to increase speed above idle and operate the generator to recharge the battery storage during flight; and for controlling the engine to return to idle and controlling the electric motor to receive the battery power for landing. 
         [0004]    The features, functions, and advantages that have been discussed can be achieved independently in various embodiments or may be combined in other embodiments further details of which can be seen with reference to the following description and drawings. 
     
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         [0005]      FIG. 1  is a schematic representation of the enhanced fuel efficiency rotary wing aircraft. 
           [0006]      FIG. 2  is a schematic representation of the hybrid system block diagram of the enhanced fuel efficiency rotary wing aircraft. 
       
    
    
     DETAILED DESCRIPTION 
       [0007]    Reference is made to  FIGS. 1 and 2 , which illustrate a hybrid power rotary wing aircraft  10 . The aircraft  10  is an advanced version of a conventional helicopter platform. The design inherently allows a single engine helicopter to be just as safe as a twin-engine helicopter at lower cost and higher fuel efficiency. The following technologies have matured to make this design possible: modern light weight turbocharged diesel engines with electronic fuel controls, efficient high power electrical switching devices (IGBTs), digital commutation controls (IWMC), lightweight high power density lithium ion batteries, carbon fiber structures (advanced rotor blades which allow variable rpm), and efficient brushless dc motors (ring motors). The aircraft  10  has a substantially conventional airframe  12  that supports the component parts of the aircraft. 
         [0008]    A main rotor is mounted to the airframe  12  for rotation of the main rotor relative to the airframe  12 . It is a conventional single main rotor system, multiple blade rotor. The main rotor is comprised of a plurality of rotor blades  14  that are mounted to a rotor shaft  16 . The rotor shaft  16  is mounted to the airframe  12  for rotation of the rotor shaft  16  and the rotor blades  14  relative to the airframe  12 . 
         [0009]    A tail rotor  18  is also mounted to the airframe  12  for rotation of the tail rotor relative to the airframe. The tail rotor  18  is comprised of a plurality of tail rotor blades and a tail rotor shaft  20 . 
         [0010]    A main rotor electric motor  22  is mounted to the airframe  12 . The main rotor electric motor  22  is operatively connected to the main rotor shaft  16 . Running the main rotor electric motor  22  rotates the rotor shaft  16  and the rotor blades  14  of the main rotor of the aircraft. In the preferred embodiment of the aircraft  10 , the main rotor electric motor  22  is connected directly to the main rotor shaft  16 , meaning that there are no intervening gear boxes or equivalent devices between the main rotor electric motor  22  and the main rotor shaft  16 . This eliminates the need for a main gear box. The ratio of rotations of the main rotor electric motor output shaft and the blades of the main rotor is 1 to 1. A rigid rotor design allows for variable rotor speed without the inherent vibration problem associated with articulated rotors. 
         [0011]    A tail rotor electric motor  26  is also mounted to the airframe  12  of the aircraft  10 . The tail rotor electric motor  26  is operatively connected to the tail rotor  18 . Running the tail rotor electric motor  26  rotates the tail rotor  18  of the aircraft  10 . In the preferred embodiment of the aircraft  10 , the tail rotor electric motor  26  is directly connected to the tail rotor  18 , meaning that there are no intervening gear boxes or other equivalent devices between the tail rotor electric motor  26  and the tail rotor  18 . The ratio of the rotations of the tail rotor electric motor output shaft and the blades of the tail rotor  18  is 1 to 1. The tail rotor  18  is a conventional tail fan design with single motor rotation direction and collective control of tail rotor blade pitch via redundant electric linear actuators. 
         [0012]    An electric generator  32  is mounted to the airframe  12  of the aircraft  10 . The generator  32  is an interleaved generator with three phases. The generator phases are isolated and inverted by separate inverters to allow for electric power generation redundancy. The electric generator  32  is operatively, electrically connected to the main rotor electric motor  22 . Operating the electric generator  32  or rotating the rotor of the electric generator  32  generates power that is supplied from the generator  32  through an electrical network  34  of the aircraft  10  (represented in  FIG. 2 ) to the main rotor electric motor  22  to run the main rotor electric motor  22 . The electric generator  32  is also operatively, electrically connected to the tail rotor electric motor  26  through the electrical network  34 . Operating the electric generator  32  or rotating the rotor of the generator  32  produces electric power that is supplied through the network  34  to the tail rotor electric motor  26  to run the tail rotor electric motor  26 . 
         [0013]    A combustion engine  38  is mounted to the airframe  12  of the aircraft  10 . The combustion engine  38  is operatively connected to the generator  32 . Running the combustion engine  38  rotates the rotor of the generator  32  and operates the generator  32 . In the preferred embodiment of the aircraft  10 , the combustion engine  38  is a lightweight turbocharged diesel engine. The diesel engine  38  allows fuel compatibility with existing aviation systems. Turbo charging is employed to provide forced induction and resulting performance gains. The engine  38  and generator  32  are optimized to operate in the 2000 rpm range. The hybrid aircraft  10  allows for the lighter diesel engine  38 , and peak loads can be handled from the battery system to be described. 
         [0014]    For the aircraft  10 , a typical flight would take off at idle using battery power, then ramp up power with the diesel engine to recharge the batteries, then reduce to cruise, then back to idle to land under battery power. The turbocharged diesel provides low specific fuel consumption at a wide range of power, altitude, and rpm settings consistent with aircraft operating over a wide range of mission takeoffs, loiter, surge, etc. Existing turboshaft engines have a specific fuel consumption rating of 0.8 while advanced turbocharged diesel engines have a rating of 0.3 to 0.34. A significant portion of a helicopter gross weight is fuel. The fuel savings will benefit the helicopter design allowing lighter structure and reduced fuel tankage to offset the required volume for the battery system to be described. 
         [0015]    Engine cooling is provided by a conventional radiator system with an electric fan forced air system. The electric fan allows for further efficiencies as the fan motor can be shut down during cruise or at high altitude when not needed to maintain engine temperature. 
         [0016]    The combustion engine  38  includes an electronic engine controller  44 . The electronic engine controller  44  adjusts the speed of the rotation of the engine output shaft  40  in response to electrical signals received by the engine controller  44 . In the preferred embodiment of the aircraft  10 , an output shaft  40  of the combustion engine  38  is directly connected to a rotor shaft  42  of the generator  32 . By being directly connected what is meant is that the engine output shaft  40  is directly connected to the generator rotor shaft  42  with there being no intervening gear boxes or other equivalent devices. The ratio of the rotation of the engine output shaft  40  and the generator rotor shaft  42  is 1 to 1. 
         [0017]    A plurality of batteries  48 ,  50 ,  52  are mounted to the airframe  12  of the aircraft  10 . The batteries  48 ,  50 ,  52  are operatively, electrically connected with the electric generator  32  to receive electric power from the operating generator which charges the batteries. The batteries  48 ,  50 ,  52  are also operatively, electrically connected to the main rotor electric motor  22  to supply power to the main rotor electric motor  22 , and to the tail rotor electric motor  26  to supply power to the tail rotor electric motor  26 . In the preferred embodiment of the aircraft  10 , three batteries  48 ,  50 ,  52  are mounted to the airframe  12  of the aircraft  10 . The three batteries  48 ,  50 ,  52  provide redundancy to the hybrid propulsion system of the aircraft  10 . Additionally, in the preferred embodiment of the aircraft  10  the three batteries  48 ,  50 ,  52  are each high energy density lithium ion batteries. The batteries  48 ,  50 ,  52  also provide backup power in the event of a diesel engine failure and are sized to provide sufficient energy for a take off followed by abort to landing. The batteries operate at the 270 V range. Battery charge control and over voltage protection is managed by generator inverters to be described. 
         [0018]    Three separate, isolated power busses operating in the 270 V range are provided for redundancy in distribution of the generated power to the motors  22 ,  26 . Multiple inverters drive the motors  22 ,  26  using a motor phase interleave technique to provide electrical redundancy and inverter power density efficiency. The integrated motor and inverter combination is fault tolerant to electrical failures. The generator  32  output is commutated, conditioned and controlled by the inverters. 
         [0019]    A plurality of first electric current inverters  56 ,  58 ,  60  are mounted to the airframe  12  of the aircraft  10 . Each of the first inverters  56 ,  58 ,  60  is operatively, electrically connected to the electric generator  32  through the electric network  34  and each of the first inverters  56 ,  58 ,  60  is operatively, electrically connected to the respective batteries  48 ,  50 ,  52  through the network  34 . Each of the inverters  56 ,  58 ,  60  includes a micro-controller unit and converts alternating electric current created by the electric generator  32  to direct electric current and supplies the direct electric current to the respective batteries  48 ,  50 ,  52  to charge the batteries. 
         [0020]    A plurality of second electric current inverters  64 ,  66 ,  68  is also provided on the airframe  12  of the aircraft  10 . Each of the second inverters  64 ,  66 ,  68  is operatively, electrically connected to a respective battery  48 ,  50 ,  52  through the electrical network  34  and is also operatively, electrically connected to the main rotor electric motor  22  through the electric network  34 . Each of the second inverters  64 ,  66 ,  68  includes a microcontroller and receives direct electric current supplied by a respective battery  48 ,  50 ,  52  and converts the direct electric current to alternating electric current. The alternating electric current is supplied by each of the second inverters  64 ,  66 ,  68  to the main rotor electric motor  22  as one phase of a three phase system. The three phases of the alternating current supplied by the second inverters  64 ,  66 ,  68  to the main rotor electric motor  22  run the motor  22 . The inverter control can vary the rpm of the main rotor electric motor  22  to allow operation of the main rotor at the most efficient operating point for a given flight regime and aircraft weight over time. 
         [0021]    A plurality of third electric current inverters  72 ,  74 ,  76  is also provided on the airframe  12  of the aircraft  10 . Each of the third electric current inverters  72 ,  74 ,  76  is operatively, electrically connected to a respective battery  48 ,  50 ,  52  through the electric network  34  and is also operatively, electrically connected to the tail rotor electric motor  26 . Each of the third inverters  72 ,  74 ,  76  includes a microcontroller and receives direct electric current supplied by a respective battery  48 ,  50 ,  52  and converts the direct electric current to alternating electric current. The alternating electric current is supplied by each of the third inverters  72 ,  74 ,  76  as one phase of a three phase system. The three phases of alternating current are supplied to the tail rotor electric motor to run the motor. 
         [0022]    A main rotor tilt actuator  80  is also provided on the airframe  12  of the aircraft  10 . The tilt actuator is operatively, mechanically connected to the main rotor electric motor  22  to selectively tilt the main rotor electric motor  22  and the main rotor shaft  16  forward during high speed flight of the aircraft  10  to improve the aerodynamic efficiency of the airframe  12  and the rotor blades  14 . 
         [0023]    A flight control system  84  of the aircraft  10  communicates through the electrical network  34  with the main rotor electric motor  22 , the tail rotor electric motor  26 , the electric generator  32 , the electronic engine controller  44 , the electric current inverters  56 ,  58 ,  60 ,  64 ,  66 ,  68 ,  72 ,  74 ,  76  and the main rotor tilt actuator  80  and provides control signals from an operator of the aircraft  10  through the electric network  34  to these components of the aircraft. The flight control system  84  includes hardware and software that integrate the generator  32  with control laws pertaining to diesel engine power settings based on system demand and the current operating environment measurements (ambient temperature, altitude, electrical demand, diesel engine temperature and thresholds). The design operates the diesel engine  38  at maximum engine efficiency at any aircraft speed. The hybrid controller logic would automatically vary the diesel engine  38  and rotor motor  22  performance parameters based on takeoff weight, flight conditions, performance settings, etc. Additional efficiencies are gained because the main rotor requires faster rpm in high speed flight and can turn slower in hovering or low speed flight to conserve power. 
         [0024]    The electric propulsion is integrated with electro-mechanical actuators mounted on the main rotor electric motor  22  and tail rotor electric motor  26  to provide additional system efficiencies over conventional helicopters which use hydraulic actuators and a gear box driven hydraulic pump system. The all electric tail rotor design facilitates a folding tail system to aid in aircraft storage on the ground. 
         [0025]    The electrical network  34  of the aircraft is also provided with a backup power connection  86 . When the aircraft  10  is idle, the backup power connection  86  can be connected to a separate source of  24  volt electric power to charge the batteries  48 ,  50 ,  52  as the aircraft is idle. 
         [0026]    The aircraft  10  described above reduces the fuel consumption of the aircraft by as much as fifty percent of that of a conventional rotary wing aircraft of substantially the same configuration. The aircraft  10  lowers the gross takeoff weight by eliminating gear boxes and drive transmissions, and by eliminating hydraulic control systems for the gear boxes and transmissions as well as other operative surfaces of a conventional rotary wing aircraft. Reducing fuel storage needs also reduces the weight of the aircraft and lowers the carbon footprint of the aircraft by reducing engine specific fuel consumption. The acoustics of the aircraft  10  on take off and landing are also lowered by use of electric power. 
         [0027]    Although the hybrid diesel/electric power aircraft has been described above by reference to a specific embodiment, it should be understood that modifications and variations could be made to the aircraft described without departing from the intended scope of the claims appended hereto and their equivalents.