Patent Publication Number: US-8991748-B1

Title: Solid lubricated blade pitch control system for use within a compressed airstream of a reaction driven rotorcraft

Description:
RELATED APPLICATIONS 
     This application claims the benefit of U.S. Provisional Application Ser. No. 61/517,413 filed Apr. 19, 2011 and entitled SOLID LUBRICATED BLADE PITCH CONTROL SYSTEM FOR USE WITHIN A COMPRESSED AIR STREAM OF A REACTION DRIVEN ROTORCRAFT. This application further incorporates by reference co-pending U.S. patent application Ser. No. 13/199,671, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/381,291, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,684, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,099, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,678, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,097, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,682, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/381,313, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,681, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,111, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,677, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/381,347, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,679, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,136, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,720, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,134, filed on Sep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/460,572, filed on Jan. 3, 2011, co-pending U.S. patent application Ser. No. 13/199,719, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,098, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,721, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,081, filed on Sep. 9, 2010, co-pending U.S. patent application Ser. No. 13/199,705, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,135, filed on Sep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/466,177, filed on Mar. 22, 2011, U.S. Provisional Patent Application Ser. No. 61/409,475, filed on Nov. 2, 2010, co-pending U.S. patent application Ser. No. 13/199,712, filed on Sep. 7, 2011, which claims the benefit of U.S. Provisional Patent Application Ser. No. 61/403,113, filed on Sep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/409,478, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/409,476, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/409,482, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/409,470, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/517,413, filed on Apr. 19, 2011, U.S. Provisional Patent Application Ser. No. 61/468,964, filed on Mar. 29, 2011, U.S. Provisional Patent Application Ser. No. 61/409,487, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/409,494, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/456,219, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/456,221, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/456,220, filed on Nov. 2, 2010, U.S. Provisional Patent Application Ser. No. 61/432,488, filed on Jan. 13, 2011, U.S. Provisional Patent Application Ser. No. 61/506,572, filed on Jul. 11, 2011, U.S. Provisional Patent Application Ser. No. 61/519,075, filed on May 16, 2011, U.S. Provisional Patent Application Ser. No. 61/519,055, filed on May 16, 2011, U.S. Provisional Patent Application Ser. No. 61/460,573, filed on Jan. 4, 2011, U.S. Provisional Patent Application Ser. No. 61/461,223, filed on Jan. 13, 2011, U.S. Provisional Patent Application Ser. No. 61/429,282, filed on Jan. 3, 2011, U.S. Provisional Patent Application Ser. No. 61/429,289, filed on Jan. 3, 2011, U.S. Provisional Patent Application Ser. No. 61/499,996, filed on Jun. 22, 2011, U.S. Provisional Patent Application Ser. No. 61/575,196, filed on Aug. 17, 2011, U.S. Provisional Patent Application Ser. No. 61/575,204, filed on Aug. 17, 2011, U.S. Provisional Patent Application Ser. No. 61/532,233, filed on Sep. 8, 2011, U.S. Provisional Patent Application Ser. No. 61/539,668, filed on Sep. 27, 2011, and U.S. Provisional Patent Application Ser. No. 61/626,783, filed on Oct. 3, 2011, all of which are hereby incorporated by reference. 
     Additionally, this patent application hereby incorporates by reference U.S. Pat. No. 5,301,900 issued Apr. 12, 1994 to Groen et al., U.S. Pat. No. 1,947,901 issued Feb. 20, 1934 to J. De la Cierva, and U.S. Pat. No. 2,352,342 issued Jun. 27, 1944 to H. F. Pitcairn. 
    
    
     RIGHTS OF U.S. GOVERNMENT 
     This invention was made with Government support under Agreement No. HR0011-06-9-0002 awarded by DARPA. The Government has certain rights in the invention. 
    
    
     BACKGROUND 
     1. The Field of the Invention 
     This invention relates to rotating wing aircraft, also known as rotorcraft, and, more particularly to rotating wing aircraft relying on autorotation of a rotor to provide lift. 
     2. The Background Art 
     Rotorcraft rely on a rotating wing to provide lift. In contrast, fixed wing aircraft rely on air flow over a fixed wing to provide lift. Fixed wing aircraft must therefore achieve a minimum ground velocity on takeoff before the lift on the wing is sufficient to overcome the weight of the plane. Fixed wing aircraft therefore generally require a long runway along which to accelerate to achieve this minimum velocity and takeoff. 
     In contrast, rotating wing aircraft can take off and land vertically or along short runways inasmuch as powered rotation of the rotating wing provides the needed lift. This makes rotating wing aircraft particularly useful for landing in urban locations or undeveloped areas without a proper runway. 
     The most common rotorcraft in use today are helicopters. A helicopter typically includes an airframe or fuselage, housing an engine and passenger compartment, and a rotor, driven by the engine, to provide lift. Forced rotation of the rotor causes a reactive torque on the fuselage. Accordingly, conventional helicopters require either two counter rotating rotors or a tail rotor in order to counteract this reactive torque. 
     Another type of rotorcraft is the autogyro. An autogyro aircraft derives lift from an unpowered, freely rotating rotor or plurality of rotary blades. The energy to rotate the rotor results from a windmill-like effect of air passing through the underside of the rotor. The forward movement of the aircraft comes in response to a thrusting engine such as a motor driven propeller mounted fore or aft. 
     During the developing years of aviation aircraft, autogyro aircraft were proposed to avoid the problem of aircraft stalling in flight and to reduce the need for runways. The relative airspeed of the rotating wing is independent of the forward airspeed of the autogyro, allowing slow ground speed for takeoff and landing, and safety in slow-speed flight. Engines may be tractor-mounted on the front of an autogyro or pusher-mounted on the rear of the autogyro. 
     Airflow passing the rotary wing, alternately called rotor blades, which are tilted upward toward the front of the autogyro, act somewhat like a windmill to provide the driving force to rotate the wing, i.e. autorotation of the rotor. The Bernoulli effect of the airflow moving over the rotor surface creates lift. 
     Various autogyro devices in the past have provided some means to begin rotation of the rotor prior to takeoff, thus further minimizing the takeoff distance down a runway. One type of autogyro is the “gyrodyne,” which includes a gyrodyne built by Fairey aviation in 1962 and the XV-1 convertiplane first flight tested in 1954. The gyrodyne includes a thrust source providing thrust in a flight direction and a large rotor for providing autorotating lift at cruising speeds. To provide initial rotation of the rotor, jet engines were secured to the tip of each blade of the rotor and powered during takeoff, landing, and hovering. 
     Although rotorcraft provide the significant advantage of vertical takeoff and landing (VTOL), they are much more limited in their maximum flight speed than are fixed wing aircraft. The primary reason that prior rotorcraft are unable to achieve high flight speed is a phenomenon known as “retreating blade stall.” As the fuselage of the rotorcraft moves in a flight direction, rotation of the rotor causes each blade thereof to be either “advancing” or “retreating.” 
     That is, in a fixed-wing aircraft, all wings move forward in fixed relation, with the fuselage. In a rotary-wing aircraft, the fuselage moves forward with respect to the air. However, rotor blades on both sides move with respect to the fuselage. Thus, the velocity of any point on any blade is the velocity of that point, with respect to the fuselage, plus the velocity of the fuselage. A blade is advancing if it is moving in the same direction as the flight direction. A blade is retreating if it is moving opposite the flight direction. 
     The rotor blades are airfoils that provide lift that depends on the speed of air flow thereover. The advancing blade therefore experiences much greater lift than the retreating blade. One technical solution to this problem is that the blades of the rotors are allowed to “flap.” That is, the advancing blade is allowed to fly or flap upward in response to the increased air speed thereover such that its blade angle of attack is reduced. This reduces the lift exerted on the blade. The retreating blade experiences less air speed and tends to fly or flap downward such that its blade angle of attack is increased, which increases the lift exerted on the blade. Flap enables rotating wing aircraft to travel in a direction perpendicular to the axis of rotation of the rotor. However, lift equalization due to flapping is limited by a phenomenon known as “retreating blade stall.” As noted above, flapping of the rotor blades increases the angle of attack of the retreating blade. However, at certain higher speeds, the increase in the blade angle of attack required to equalize lift on the advancing and retreating blades results in loss of lift (stalling) of the retreating blade. 
     A second limit on the speed of rotorcraft is the drag at the tips of the rotor. The tip of the advancing blade is moving at a speed equal to the speed of the aircraft and relative to the air, plus the speed of the tip of the blade with respect to the aircraft. That is equal to the sum of the flight speed of the rotorcraft plus the product of the length of the blade and the angular velocity of the rotor. In helicopters, the rotor is forced to rotate in order to provide both upward lift and thrust in the direction of flight. Increasing the speed of a helicopter therefore increases the air speed at the rotor or blade tip, both because of the increased flight speed and the increased angular velocity of the rotors required to provide supporting thrust. 
     The air speed over the tip of the advancing blade can therefore exceed the speed of sound even though the flight speed is actually much less. As the air speed over the tip approaches the speed of sound, the drag on the blade becomes greater than the engine can overcome. In autogyro aircraft, the tips of the advancing blades are also subject to this increased drag, even for flight speeds much lower than the speed of sound. The tip speed for an autogyro is typically smaller than that of a helicopter, for a given airspeed, since the rotor is not driven. Nevertheless, the same drag increase occurs eventually. 
     A third limit on the speed of the rotorcraft is reverse air flow over the retreating blade. As noted above, the retreating blade is traveling opposite the flight direction with respect to the fuselage. At certain high speeds, portions of the retreating blade are moving rearward, with respect to the fuselage, slower than the flight speed of the fuselage. Accordingly, the direction of air flow over these portions of the retreating blade is reversed from that typically designed to generate positive lift. Air flow may instead generate a negative lift, or downward force, on the retreating blade. For example, if the blade angle of attack is upward with respect to wind velocity, but wind is moving over the wing in a reverse direction, the blade may experience negative lift. 
     The ratio of the maximum air speed of a rotorcraft to the maximum air speed of the tips of the rotor blades is known as the “advance ratio. The maximum advance ratio of rotorcraft available today is less than 0.5, which generally limits the top flight speed of rotorcraft to less than 200 miles per hour (mph). For most helicopters, that maximum achievable advance ratio is between about 0.3 and 0.4. 
     In view of the foregoing, it would be an advancement in the art to provide a rotating wing aircraft capable of vertical takeoff and landing and flight speeds in excess of 200 mph. 
     BRIEF SUMMARY OF THE INVENTION 
     The invention has been developed in response to the present state of the art and, in particular, in response to the problems and needs in the art that have not yet been fully solved by currently available apparatus and methods. The features and advantages of the invention will become more fully apparent from the following description and appended claims, or may be learned by practice of the invention as set forth hereinafter. 
     In one aspect of the invention, a rotor system includes a mast and a shroud surrounding the mast. The mast and shroud define an air channel. A hub is mounted to the mast and defines a hub cavity in fluid communication with the air channel. Rotor blades, each defining a blade duct, are coupled to the hub, having the blade ducts thereof in fluid communication with the hub cavity. A swashplate encircles the mast, and a plurality of actuators are coupled to the swashplate by means of actuator rods. A plurality of pitch control rods couple the swashplate to the corresponding plurality of blades. 
     A layer of solid lubricant is fused to at least one surface located at least one of at an interface between the swashplate and the mast, an interface between the swashplate and the plurality of actuator rods, an interface between the swashplate and the plurality of pitch control rods, and an interface between the plurality of pitch control rods and the plurality of blades. 
     In another aspect of the invention, the interface between the swashplate and the mast includes a spherical bearing slidably mounted to the mast and defining a spherical surface engaging the swashplate. The layer of solid lubricant is fused to at least one of a surface of the mast engaging the spherical bearing, a surface of the spherical bearing engaging the mast, the spherical surface of spherical bearing, and a surface of the swashplate engaging the spherical surface. 
     In another aspect of the invention, the interfaces between the swashplate and the pitch control rods, between the swashplate and the actuator rods, and between the pitch control rods and the blades is formed by a pin inserted within an aperture. The layer of solid lubricant may be fused to one or both of the aperture and pin of any of these interfaces. 
     In another aspect of the invention, the layer of solid lubricant includes at least one of molybdenum disulphide, polytetrafluoroethylene, and boron nitride. In embodiments where the layer of solid lubricant includes molybdenum disulphide or boron nitride, the layer of solid lubricant may be baked onto the at least one surface. 
     In another aspect of the invention, a compressed air source is in fluid communication with the air channel, and a plurality of tip jets secure to the plurality of rotor blades in fluid communication with the blade ducts. 
     In another aspect of the invention, a method for operating a rotorcraft, including the above references rotor system, may include directing compressed air through the air channel during at least one of taking off, hovering, and landing of the rotor craft and ceasing to do so during sustained longitudinal flight. The compressed air may have a temperature of above about 300° F. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       The foregoing features of the present invention will become more fully apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. Understanding that these drawings depict only typical embodiments of the invention and are, therefore, not to be considered limiting of its scope, the invention will be described with additional specificity and detail through use of the accompanying drawings in which: 
         FIG. 1  is an isometric view of an aircraft in accordance with an embodiment of the present invention; 
         FIG. 2  is a front elevation view of a compressed or otherwise pressurized air supply for a tip jet in accordance with an embodiment of the present invention; 
         FIG. 3A  is a front elevation view of a rotorcraft illustrating operational parameters describing a rotor configuration suitable for use in embodiments of an apparatus and method in accordance with the present invention, and the system of  FIGS. 1 and 2  in particular; 
         FIG. 3B  is a right side elevation view of the rotorcraft of  FIG. 3A ; 
         FIG. 3C  is a partial cut of a right side elevation view of the rotor of  FIG. 3A ; 
         FIG. 4A  is a side elevation view of an embodiment of a rotor in accordance with the present invention; 
         FIG. 4B  is a side elevation cross-sectional view of an embodiment of a rotor in accordance with the invention; 
         FIG. 4C  is a bottom plan view of an embodiment of a rotor in accordance with the present invention; 
         FIG. 5  is a partial isometric view of an embodiment of a rotor in accordance with the present invention; 
         FIG. 6  is a side elevation cross-sectional view of an embodiment of a rotor and mounting system in accordance with the present invention; 
         FIG. 7  is a top cross-sectional view of an embodiment of a rotor having an angled air inlet in accordance with the present invention; 
         FIG. 8  is a side elevation cross-sectional view of an embodiment of a hub-shroud seal for a rotor in accordance with the present invention; 
         FIG. 9  is a side elevation cross-sectional view of an embodiment of a shroud-mast flange seal for a rotor in accordance with the present invention; 
         FIG. 10  is a side elevation cross-sectional view of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 11  is an isometric view of a rotating ring of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 12  is an isometric view of a non-rotating ring of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 13A  is a bottom plan view of an embodiment of a non-rotating ring having oil ports for a swashplate for use in a rotor in accordance with the present invention; 
         FIG. 13B  is a side elevation cross-sectional view of an embodiment of an oil port of the non-rotating ring of a swashplate for use in a rotor in accordance with the present invention; 
         FIG. 14A  is an isometric view of an upper seal seat of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 14B  is an isometric view of a lower seat of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 15  is a partial side elevation cross-sectional view of an embodiment of a swashplate for a rotor in accordance with the present invention; 
         FIG. 16  is a side elevation cross-sectional view of an embodiment of a mast bearing for a rotor in accordance with the present invention; 
         FIG. 17  is a side elevation cross-sectional view of an embodiment of a rotor hub for a rotor in accordance with the present invention; 
         FIG. 18  is another side elevation cross-sectional view of an embodiment of a mast and hub assembly for a rotor in accordance with the present invention; 
         FIG. 19  is a side elevation cross-sectional view of an alternative embodiment of a mast and hub assembly for a rotor in accordance with an embodiment of the present invention; 
         FIG. 20  is a partial isometric cross-sectional view of an embodiment of a seal mounting ring for a rotor in accordance with the present invention; 
         FIG. 21  is a side elevation cross-sectional view of an embodiment of a spindle bearing for a rotor in accordance with the present invention; 
         FIG. 22  is a schematic diagram of an embodiment of an oil distribution system for a rotor in accordance with the present invention; 
         FIG. 23A  is a schematic diagram of an embodiment of a passive air flow heating system for a rotor in accordance with the present invention; 
         FIG. 23B  is a schematic diagram of an embodiment of a bleed air heating system for a rotor in accordance with the present invention; 
         FIG. 24  is a block diagram of an embodiment of a flight control system for an aircraft incorporating a rotor in accordance with the present invention; 
         FIG. 25  is a process flow diagram of an embodiment of a method for thermal management of a rotor in accordance with the present invention; 
         FIG. 26  is a process flow diagram of an embodiment of another method for extracting heat from a rotor in accordance with the present invention; 
         FIG. 27  is a side elevation cross-sectional view of a swashplate illustrating surfaces suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 28  is a side elevation cross-sectional view of an actuator rod mount suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 29  is a partial isometric view of a pitch control rod suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 30  is a partial isometric view of an interface between a pitch control rod and a pitch control rod mount suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 31  is a side elevation view of a rotating stabilizer linkage having interfaces suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 32  is an isometric view of a spherical bearing seat in a linkage mount suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 33  is a retaining ring for a spherical bearing suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 34  is an isometric view of a non-rotating stabilizer linkage having interfaces suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 35  is an isometric view of a linkage mount suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 36  is an isometric view of a leg of a stabilizer linkage suitable for application of a solid lubricant in accordance with an embodiment of the present invention; 
         FIG. 37  is an isometric view of an interface between a pitch control rod and a pitch horn suitable for application of a solid lubricant in accordance with an embodiment of the present invention; and 
         FIG. 38  is a side cross-sectional view of a portion of a component bearing a solid lubricant layer in accordance with an embodiment of the present invention. 
     
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     It will be readily understood that the components of the present invention, as generally described and illustrated in the drawings herein, could be arranged and designed in a wide variety of different configurations. Thus, the following more detailed description of the embodiments of the system and method of the present invention, as represented in the drawings, is not intended to limit the scope of the invention, as claimed, but is merely representative of various embodiments of the invention. The illustrated embodiments of the invention will be best understood by reference to the drawings, wherein like parts are designated by like numerals throughout. 
     Referring to  FIG. 1 , an aircraft  10  includes a fuselage  12  or airframe  12  defining a cabin for carrying an operator, passengers, cargo, or the like. The airframe  12  may include one or more fixed wings  14  shaped as airfoils for providing lift to the aircraft. The wings  14  may be configured such that they provide sufficient lift to overcome the weight of the aircraft  10  only at comparatively high speeds inasmuch as the aircraft  10  is capable of vertical takeoff and landing (VTOL) and does not need lift from the fixed wings  14  at low speeds, e.g. below 50 mph or even 100 mph upon taking off. 
     In this manner, the wings  14  may be made smaller than those of fixed wing aircraft requiring a high velocity takeoff, which results in lower drag at higher velocities. 
     In some embodiments the wings  14  provide sufficient lift to support at least 50 percent, preferably 90 percent, of the weight of the aircraft  10  at air speeds above 200 mph. 
     Control surfaces  16  may secure to one or both of the airframe  12  and wings  14 . For example a tail structure  18  may include one or more vertical stabilizers  20  and one or more rudders  22 . The rudders  22  may be adjustable as known in the art to control the yaw  24  of the aircraft  10  during flight. As known in the art, yaw  24  is defined as rotation about a vertical axis  26  of the aircraft  10 . In the illustrated embodiment, the rudders  22  may comprise hinged portions of the vertical stabilizers  20 . 
     The tail structure  18  may further include a horizontal stabilizer  28  and an elevator  30 . The elevator  30  may be adjustable as known in the art to alter the pitch  32  of the aircraft  10 . As known in the art, pitch  32  is defined as rotation in a plane containing the vertical axis  26  and a longitudinal axis  34  of the airframe of an aircraft  10 . In the illustrated embodiment, the elevator  30  is a hinged portion of the horizontal stabilizer  28 . In some embodiments, twin rudders  22  may be positioned at an angle relative to the vertical axis  26  and serve both to adjust the yaw  24  and pitch  32  of the aircraft  10 . 
     The control surfaces  16  may also include ailerons  36  on the wings  14 . As known in the art, ailerons  36  are used to control roll  38  of the airplane. As known in the art, roll  38  is defined as rotation about the longitudinal axis  34  of the aircraft  10 . 
     Lift during vertical takeoff and landing and for augmenting lift of the wings  14  during flight is provided by a rotor  40  comprising a number of individual blades  42 . The blades are mounted to a rotor hub  44 . The hub  44  is coupled to a mast  46  which couples the rotor hub  44  to the airframe  12 . The rotor  40  may be selectively powered by one or more engines  48  housed in the airframe  12 , or adjacent nacelles, and coupled to the rotor  40 . In some embodiments, jets  50  located at or near the tips of the blades  42  power the rotor  40  during takeoff, landing, hovering, or when the flight speed of the aircraft is insufficient to provide sufficient autorotation to develop needed lift. 
     Referring to  FIG. 2 , while still referring to  FIG. 1 , in the illustrated embodiment, the engines  48  may be embodied as jet engines  48  that provide thrust during flight of the aircraft. The jet engines  48  may additionally supply compressed air to the jets  46  by driving a bypass turbine  62  or auxiliary compressor. Air compressed by the bypass turbine  62  may be transmitted through ducts  54  to a plenum  56  in fluid communication with the ducts  54 . 
     The plenum  56  is in fluid communication with the mast  46  that is hollow or has another passage to provide for air conduction. A mast shroud  58  positioned around the mast  46  may provide one or both of an air channel and a low drag profile for the mast  46 . The mast  46  or mast shroud  58  is in fluid communication with the rotor hub  44 . The rotor hub  44  is in fluid communication with blade ducts  60  extending longitudinally through the blades  42  to feed the tip jets  50 . 
     Referring to  FIGS. 3A-3C , rotation of the rotor  40  about its axis of rotation  72  occurs in a rotor disc  70  that is generally planar but may be contoured due to flexing of the blades  42  during flight. In general, the rotor disc  70  may be defined as a plane in which the tips of the blades  42  travel. Inasmuch as the blades  42  flap cyclically upward and downward due to changes in lift while advancing and retreating, the rotor disc  70  is angled with respect to the axis of rotation  72  when viewed along the longitudinal axis  34 , as shown in  FIG. 3A . 
     Referring to  FIG. 3B , the angle  74  of the rotor disc  70 , relative to a flight direction  76  in the plane containing the longitudinal axis  34  and vertical axis  26 , is defined as the angle of attack  74  or rotor disk angle of attack  74 . For purposes of this application, flight direction  76  and air speed refer to the direction and speed, respectively, of the airframe  12  of the aircraft  10  relative to surrounding air. In autogyro systems, the angle of attack  74  of the rotor disc  70  is generally positive in order to achieve autorotation of the rotor  40 , which in turn generates lift. 
     Referring to  FIG. 3C , the surfaces of the blades  42 , and particularly the chord of each blade  42 , define a pitch angle  78 , or blade angle of attack  78 , relative to the direction of movement  80  of the blades  42 . In general, a higher pitch angle  78  will result in more lift and higher drag on the blade up to the point where stalling occurs, at which point lift has declined below a value necessary to sustain flight. The pitch angle  78  of the blade  42  may be controlled by both cyclic and collective pitch control as known in the art of rotary wing aircraft design. 
     Referring to  FIGS. 4A through 4C , the mast shroud  58  encircles the mast  46  such that the shroud and mast define an annular air channel  92 . In an alternative embodiment, the air channel  92  passes through the center of the mast  46 . The hub  44  defines a cavity  94  in fluid communication with the air channel  92 . The blade ducts  60  are in fluid communication with the cavity  94  enabling air flow from the air channel  92  through the cavity  94  and blade ducts  60  to the tip jets  50 . In the illustrated embodiment, the blades  42  define a hollow blade spar  96  extending through the hub  44  into the cavity  94  and the blade duct  60  is embodied as a hollow channel extending longitudinally through the blade spar  96 . 
     A mast flange  98  may be rigidly secured to, or formed monolithically with the mast  46 . The shroud may create a substantially continuous barrier to air flow between the hub  44  and the mast flange  46 , but for an inlet  100  coupled to the plenum  56 . By substantially continuous barrier to air flow, what is meant is that the shroud ensures that at least 90%, preferably at least 95%, of all air entering the air channel  92  from the plenum  56  passes into the cavity  94  of the hub  44 . The shroud may additionally include one or more sealed hatches  102  that are selectively openable to service internal components of the rotor  40  without requiring removal of the entire mast shroud  58 . 
     Referring specifically to  FIG. 4B , the one or more hatches  102  may be located adjacent a swashplate  104  that encircles the mast  46  to facilitate servicing of the swashplate  104 . The swashplate  104  may engage a spherical bearing  106  or gimbal slidingly mounted to the mast  46 . The swashplate  104  includes a rotating ring  108  and a non-rotating ring  110  rotatably secured to one another. The swashplate  104  may encircle the mast  46  between the hub  44  and the mast flange  98 . The non-rotating ring  110  is coupled to swashplate actuators  112 , such as by means of actuator rods  114 . The rotating ring  108  is coupled to the blade spars  96 , such as by means of pitch horns  118  coupled near proximal ends  120  of the blade spars  96  and pitch control rods  116  connecting the pitch horns  118  to the rotating ring  108 . The swashplate actuators  112  are selectively actuated to raise and lower the swashplate  104  and spherical bearing  106  in order to change the pitch angles  78  of all of the blades  42  by a uniform amount, i.e., the collective pitch of the blades  42 . The swashplate actuators  112  are also selectively actuated to change the angle of the swashplate  104  in order to change the amplitude and phase of cyclic variation of the pitch angles  78 , i.e., cyclic pitch, of the blades  42  as they rotate around the mast  46 . The swashplate actuators  112  may be selectively activated to change the collective and cyclic pitch simultaneously. 
     The swashplate actuators  112  are rigidly mounted to the mast  46 , such as by rigidly mounting the swashplate actuators  112  to the mast flange  98 . In the illustrated embodiment, the swashplate actuators  112  are secured on an opposite side of the flange  98  as the swashplate  104  and the mast flange  98  defines apertures  122  permitting the actuator rods  114  to pass from the swashplate actuator  112  to the swashplate  104 . The swashplate actuators  112  may be embodied as hydraulic pistons  124  and cylinders  126 . 
     Referring specifically to  FIG. 4C , in the illustrated embodiment, three swashplate actuators  112  secure to the flange  98  inasmuch as three independent points are sufficient to define any plane, i.e., any orientation of the swashplate  104 . The swashplates  112  may be distributed at equal or unequal angular intervals around the axis of rotation  72  of the hub  44  and may be located at equal or unequal distances from the axis of rotation of the hub  44 . 
     Referring again to  FIG. 4B , a central channel  128  may extend through both the mast  46  and the mast flange  98 . A plurality of electrical lines  130  and fluid lines  132 , such as oil and fuel lines, may pass through the central channel  128 . The lines  130 ,  132  may couple to the hub  44  or structures fixedly mounted to the hub  44 . Accordingly, the lines  130 ,  132  may rotate with the hub  44 . The electrical lines  130  may couple to a slip ring assembly  134  coupling signals from stationary lines  136  to the rotating electrical lines  130 . In a like manner, a hydraulic rotary union  138  may couple stationary fluid ports  140  to the rotating fluid lines  132 . 
     The air channel  92  enables the flow of air from the inlet  100  to the cavity  94  and into the blade ducts  60 . In some embodiments, structures of the rotor  40  include fairings to reduce drag on moving components and pressure losses incurred on air moving from the inlet  100  to the blade ducts. For example, a mast fairing  142  may secure to the hub  44  and encircle that mast  46 . The mast fairing  142  extends along the mast  46  and has a contour effective to reduce pressure losses in air flowing from along the mast  46  to along the hub  44 . For example, the mast fairing may have an outer surface  144  that decreases smoothly in diameter with distance from the hub  44  along the axis of rotation  72 . By smoothly, what is meant is that the slope of the change in outer diameter with distance from the hub  44  along the axis of rotation  72  does not exceed 1.0, except for possibly discontinuities at the upper and lower edges of the mast fairing  142 . In some embodiments, the mast fairing  142  is fixed to the mast and the hub  44  is free to rotate relative to the mast fairing  142 . 
     Pressure losses in airflow from the cavity  94  to the blade ducts  60  may also be reduced by means of a blade duct fairing  146  covering the proximal end  120  of the blade spar  96 . The blade duct fairing  146  defines a “bell mouth” as known in the art of aerodynamics. The blade duct fairing  146  may define an aperture  148  through which the pitch horns  118  protrude. 
       FIG. 5  illustrates an isometric view of the swashplate  104 . In addition to the pitch control rods  116  and actuator rods  114  coupled to the rotating and non-rotating rings  108 ,  110 , respectively, rotating stabilizer linkages  150  and non-rotating stabilizer linkages  152  may secure to the rotating and non-rotating rings  108 ,  110 , respectively. The linkage  150  hinders rotation of the rotating ring  108  relative to the hub  44  while still permitting vertical movement of the rotating ring  108  relative to the hub  44 . Likewise, the linkage  152  hinders rotation of the non-rotating ring  110  relative to the swashplate actuators  112 . The linkages  150 ,  152  may each comprise a leg  154   a  coupled to the rotating ring  108  and non-rotating ring  110 , respectively, and a leg  154   b  coupled to the hub  44  and mast  46 , respectively. In the illustrated embodiment, the leg  154   b  of the linkage  150  is mounted to the mast fairing  142  secured to the hub  44  and the leg  154   b  of the linkage  152  is mounted to the mast flange  98 . 
     The legs  154   a ,  154   b  are coupled to one another by a hinge  156  defining a pivoting axis  158  that is perpendicular to the axis of rotation  72  of the hub  44  and tangent to a circle centered on the axis of rotation  72  of the hub  44 . The linkages  150 ,  152  reduce torque on the pitch control rods  116  and actuator rods  114  which are oriented vertically and are not well suited to bear such loads without hindering their ability to move freely in response to actuator inputs. 
     The rotating ring  108  may have one or more linkage mounts  160  secured thereto, or formed monolithically therewith. The legs  154   a  of one or more linkages  150  may secure to the linkage mounts  160 , such as by means of a spherical joint. The non-rotating ring may likewise have one or more linkage mounts  162  secured thereto, or formed monolithically therewith. The legs  154   a  of one or more linkages  152  may secure to the linkage mounts  162 , such as by means of a spherical joint. 
     Referring to  FIG. 6 , the mast flange  98  may include a pivot  170  pivotally secured to a mount  172  enabling changing of the angle of attach  74  of the rotor disc  70 . An actuator  174  may secure to a mount  172  and flange  98  for changing the pitch of the hub  44  and  46 . The actuator  174  may be embodied as a hydraulic cylinder  176  and piston  178  having one of the cylinder  176  and piston  178  coupled to the flange  98  and the other of the hydraulic cylinder  176  and piston  178  coupled to the mount  172 . The actuator  174  and pivot  170  may secure to separate mounts  172  or may be attached to the same mount  172 . The one or more mounts  172  are secured to a structural member  180  of the fuselage  12  by means of one or more vibration suppression devices  182  or dampers  182 . 
     The above described arrangement of the pivot  170 , actuator  174 , mount  172 , and vibration suppression  182  in conjunction with the mounting of the swashplate actuators  112  to the mast  46 , such as the mast flange  98 , provide a rotor  40  that is exceptionally rigid with very little play or slop between the blades  42  and the mast  46  and between the pitch horns  118  and the swashplate actuators  112 . Due to the rigidity of the rotor  40 , the frequency response of the rotor  40  in the same range of frequencies as the cyclic loads on the rotor  40  induced by cyclic variation in lift on the blades  42  may be damped, rather than resonant. For high speed flight, the frequency of rotation of the blades  42  is generally reduced to a minimum frequency of rotation in order to avoid the problems mentioned hereinabove that occur at high advance ratios such as retreating blade stall and high tip speed mach numbers. Reduction of rotation frequency may also facilitate equalization of roll moments exerted by the blades  42  as described in U.S. Prov. Pat. App. Ser. No. 61/403,136, filed Sep. 9, 2010, and entitled ROLL MOMENT EQUALIZATION AT HIGH ADVANCE RATIOS. Accordingly, low frequency cyclic loading may occur with large amplitudes during high speed flight. The rigidity of the rotor system described hereinabove, increases the harmonic frequencies of the rotor in order to reduce the risk of destructive resonance for low rotational frequencies during high speed flight. The rigidity of the coupling between the swashplate actuators  112  and the pitch horns  118  likewise enables very precise control of collective and cyclic pitch and raises harmonic frequencies of the linkage between the swashplate actuators  112  and the pitch horns  118  in order to avoid destructive resonances during high speed flight. 
     Referring to  FIG. 7 , in some embodiments, the inlet  100  defines a center axis  190  having a non-zero angle  192  with respect to the vertical plane  196  containing the vertical axis  26  and longitudinal axis  34  of the aircraft  10 . In some embodiments the angle is between 5 and 45 degrees, preferably between 10 and 30 degrees. In some embodiments, the extension of the center axis  190  of the inlet across the air channel  92  is offset a distance  198  from the axis of rotation  72  of the hub  44  at its point of greatest proximity to the axis of rotation  72 . The distance  190  may be between may be between 5 and 40 percent, preferably between 15 and 35 percent of the diameter of the shroud  58 . In some embodiments, one or both of the angle  192  and offset distance  190  may be chosen effective to reduce the pressure losses of air flowing through the air channel  92  and cavity  94  by between five percent and fifty percent, preferably between twenty and fifty percent. 
     As a result of the angle  192  and/or offset distance  198 , air flow  196  within the inlet  100  will be forced to rotate within the annular air channel  92  defined between the mast  46  and the mast shroud  58 . During operation, the rotor  40  includes a number of rotating components including the pitch control rods  116 , rotating ring  108 , stabilizer linkages hub  44 , and pitch horns  118 , that are rotating at high speeds. Due to the rotational velocity of the air within the inlet, the relative air speed between the air flow  196  and the rotating components of the rotor  40  is reduced, provided the rotational velocity of the air flow  196  is in the same direction as the tangential velocity of the rotating components of the rotor  40 . Accordingly, drag on the rotating components and pressure losses of the air flow  196  over the rotating components, which are proportional to the relative velocity squared, will be reduced. Pressure losses of the air flow  196  will also be reduced inasmuch as the air entering the air channel  92  from the inlet  100  is not required to make a 90 degree turn, which is very aerodynamically inefficient. In some embodiments, one or both of the angle  192  and offset distance  190  may be chosen effective to reduce pressure losses of air flowing through the air channel  92  and cavity  94  by between five percent and twenty percent, preferably between ten percent and thirty percent. 
     In some embodiments, the inlet  100  may have curved walls such that a center axis is not readily identified. In such embodiments, and in embodiments having a generally straight inlet  100 , the contour and orientation of the inlet  100  may be such that air from the inlet  100  flowing through the air channel  92  while the tip jets  50  are ignited has an average angular velocity that is in the same direction as the angular velocity of the hub and has a magnitude greater than 50 percent, preferably greater than 80 percent, of the magnitude of the angular velocity of the hub  44 . 
     Referring to  FIG. 8 , in order to reduce pressure losses due to air leakage, a seal  200  may be secured to one of the hub  44  and shroud  58 . The seal  202  may include a ring  202  of flexible material secured to one of the hub  44  and mast shroud  58  and overlapping a portion of the other of the hub  44  and the mast shroud  58 . In the illustrated embodiment, the ring  202  secures to an outer surface  204  of the mast shroud  58  and overlaps an inner surface  206  of the hub  44 . When pressurized air is within the air channel  92 , the ring  202  is urged against the inner surface  206  of the hub  44  and hinders air leakage. 
     In some embodiments, the ring  202  is an inner ring  202  and the seal  200  includes an outer ring  208 . The inner and outer rings  202 ,  208  may include cuts  210  enabling the rings  202 ,  208  to flair outwardly in response to air pressure within the shroud  58 . The cuts  210  of the inner ring  202  may be offset from the cuts  210  of the outer ring  208  such that air flow through aligned cuts  210  is prevented. The rings  202 ,  208  may be made of a flexible polymer with high wear resistance or coated with a wear resistant material. For example, the rings  202 ,  208  may be made of wear resistant polymer. 
     Referring to  FIG. 9 , air leakage between the mast shroud  58  and the mast  46  may be hindered by a flange  220  secured to, or formed monolithically with, the mast shroud  58  and encircling the mast shroud  58 . The flange  220  is secured to a seat  222  formed on the mast flange  98 . A seal or sealant material may be interposed between the flange  220  and seat  222 . Alternatively, mating surfaces of the flange  220  and seat  222  may be polished sufficiently to create an adequate seal. Fasteners  224 , such as screws or bolts, distributed circumferentially around the flange  200  may engage the flange  220  and seat  222  to secure the flange  220  to the seat  222  and promote creation of a seal therebetween. In an alternative embodiment, a flange secured to, or formed monolithically with, the mast flange and projecting upwardly from the mast flange  98  may secure to the shroud  58  and the seat  22  and flange  220  may be omitted. 
       FIG. 10  is a side elevation cross-sectional view of a swashplate  104 . As noted above, the swashplate  104  includes a rotating ring  108  and a non-rotating ring  110 . The non-rotating ring  110  is mounted to a spherical bearing  106  that is slidably mounted to the mast  46 . In the illustrated embodiment, the rotating ring  108  includes pitch control rod mounts  230  secured to or monolithically formed therewith and pivotally secured to the pitch control rods  116 . Likewise, the non-rotating ring  110  includes actuator rod mounts  232  secured there to or monolithically formed therewith. The actuator rods  114  pivotally secure to the actuator rod mounts  232 . 
     An upper seal  234  and a lower seal  236  are interposed between the rotating ring  108  and the non-rotating ring  110 . One or two bearings  238 ,  240  may likewise be interposed between the rotating ring  108  and the non-rotating ring  110 . The seals  234 ,  236  may be positioned within seal seats  242 ,  244 , respectively. The seats  242 ,  244  may be embodied as separate members secured to the non-rotating ring  110  and rotating ring  108 , respectively. In the illustrated embodiments, the seat  242  also capture a portion of the bearings  238 ,  240  between itself and the non-rotating ring  110 . In a like manner, the seat  244  captures a portion of the bearings  238 ,  240  between itself and the rotating ring  108 . 
     An upper clamping ring  246  captures the seal  234  between itself and the seat  242 . Likewise, a lower clamping ring  248  captures the seal  236  between itself and the seat  244 . The upper clamping ring  246  may be secured to the seat  242  by means of fasteners  250 , such as screws, bolts, or the like. The fasteners  250  may extend through both the upper clamping ring  246  and the seat  242  and fasten to the non-rotating ring  110  thereby securing both the clamping ring  246  and the seat  242  to the non-rotating ring  110 . The upper clamping ring  246  may additionally capture the spherical bearing  106  between itself and the non-rotating ring  110 . The upper clamping ring  246  may define a spherical bearing seat  252  having a spherical contour for engaging the spherical bearing  106 . In a like manner, fasteners  254  may secure the lower clamping ring  248  to the seat  244  and may pass through both the lower clamping ring  248  and the seat  244  and secure to the rotating ring  108 , thereby securing both the lower clamping ring  248  and the seat  244  to the rotating ring  108 . A sealing material, such as a polymer gasket, may be positioned between the upper clamping ring  246  and the seat  242  and between the lower clamping ring  248  and the seat  244  to create a seal therebetween hindering the leakage of oil. 
     Referring to  FIG. 11 , the rotating ring  108  defines a bearing seat  260  sized to receive a portion of the bearings  238 ,  240 . The bearing seat  260  may be embodied as an inward facing cylindrical wall  262  and a radial wall  264  extending radially inward from one edge of the cylindrical wall  262 . The bearings  238 ,  240  may be captured within the bearing seat  260  by means of the seal seat  244  secured to the rotating ring  108 . The rotating ring  108  may additionally include a sealing surface  266  positioned to engage the upper seal  234 . In the illustrated embodiment, the sealing surface  266  is a cylindrical surface centered on the central axis  268  of the rotating ring  108 . The cylindrical wall  262  may likewise be centered on the central axis  268 . 
     The bearing seat  260  may have a plurality of grooves formed therein to facilitate the flow of oil around and through the bearings  238 ,  240 . The grooves may include circumferential grooves  270  formed in one or both of the cylindrical wall  262  and the radial wall  264 . The grooves may also include grooves  272  extending vertically along the cylindrical wall  262  and radially along the radial wall  264 . The rotating ring  108  may additionally define apertures  274 , which may be threaded, for receiving the fasteners  254 . 
     Referring to  FIG. 12 , the non-rotating ring  110  defines a bearing seat  280  sized to receive a portion of the bearings  238 ,  240 . The bearing seat  280  may be embodied as an outward facing cylindrical wall  282  and a radial wall  284  extending radially outward from one edge of the cylindrical wall  282 . The bearings  238 ,  240  may be captured within the bearing seat  280  by means of the seal seat  242  secured to the non-rotating ring  110 . The non-rotating ring  110  may additionally include a sealing surface  286  positioned to engage the lower seal  236 . In the illustrated embodiment, the sealing surface  286  is a cylindrical surface centered on the central axis  288  of the non-rotating ring  110 . The cylindrical wall  282  may likewise be centered on the central axis  288 . 
     The bearing seat  280  may have a plurality of grooves formed therein to facilitate the flow of oil around and through the bearings  238 ,  240 . The grooves may include circumferential grooves  290  formed in one or both of the cylindrical wall  282  and the radial wall  284 . The grooves may also include grooves  292  extending vertically along the cylindrical wall  282  and radially along the radial wall  284 . 
     The non-rotating ring  110  may additionally define apertures  294 , which may be threaded, for receiving the fasteners  250 . The non-rotating ring  110  may also define a spherical bearing surface  296  engaging the spherical bearing  106  in the assembled swashplate  104 . The spherical bearing surface  296  may have a spherical contour sized to mate with the spherical bearing  106 . 
     Referring to  FIGS. 13A and 13B , the non-rotating ring  110  may include a feed port  300   a  and a return port  300   b  mounted thereto. Each of the feed port  300   a  and return port  300   b  is in fluid communication with a channel  302  in fluid communication with one or both of a vertical groove  292  and a circumferential groove  290 . Oil for lubricating the bearings  238 ,  240  is pumped into the feed port  300   a  and is forced out of the return port  300   b . In addition to lubrication, the oil may be used to cool or heat the bearings  238 ,  240  in order to maintain the bearings  238 ,  240  in a preloaded condition notwithstanding temperature variations of air flow over the swashplate  104  and heat buildup due to friction. In an alternative embodiment, one or both of the feed port  300   a  and the return port  300   b  are secured to the rotating plate  108 . In such embodiments, the feed port  300   a  and/or return port  300   b  coupled to the rotating plate  108  may be connected to fluid lines  132  emanating from the hydraulic rotary union  138 . 
     Referring to  FIG. 14A , the seal seat  242  may be embodied as a ring including radial grooves  310  formed on a lower surface thereof and extending partially radially inwardly from an outer circumference of the lower surface. The grooves  310  may be positioned such that the grooves  310  are aligned with the vertical grooves  292  of the non-rotating ring  110  when in place as shown in  FIG. 10 . The seal seat  242  may likewise include apertures  312  for receiving the fasteners  250 . 
     The seal seat  242  may include an outwardly facing cylindrical wall  314  centered on the central axis  316  of the seal seat  242  and a radial wall  318  extending radially outward from the central axis  316  from an edge of the cylindrical wall  314 . The seal  234  may abut the cylindrical wall  314  and the radial wall  318  in the assembled swashplate  104 . 
     Referring to  FIG. 14B , in a like manner, the seal seat  244  may be embodied as a ring including radial grooves  320  formed on an upper surface thereof and extending partially radially outwardly from an inner circumference of the upper surface. The grooves  320  may be positioned such that the grooves  320  are aligned with the vertical grooves  272  of the rotating ring  108  when in place as shown in  FIG. 10 . The seal seat  244  may likewise include apertures  322  for receiving the fasteners  254 . 
     The seal seat  244  may include an inwardly facing cylindrical wall  324  centered on the central axis  326  of the seal seat  244  and a radial wall  328  extending radially inward from the central axis  326  from an edge of the cylindrical wall  324 . The seal  236  may abut the cylindrical wall  324  and the radial wall  328  in the assembled swashplate  104 . 
     Referring to  FIG. 15 , the radial grooves  310  of the seal seat  242 , the grooves  272  of the rotating plate  108 , the grooves  292  of the non-rotating plate  110 , and the radial grooves  320  of the seal seat  244  form a fluid path extending around the bearings  238 ,  240 . The sealing surface  266  of the rotating ring  108  engages the seal  234  and the sealing surface  286  of the non-rotating ring  110  engages the seal  236 . The seals  234 ,  236  hinder the entry of contaminants into the bearings  238 ,  240  and hinder the leakage of oil from between the rotating plate  108  and the non-rotating plate  110 . 
     The bearings  238 ,  240  may each include an outer race  330  engaging the rotating ring  108  and moving synchronously therewith and an inner race  332  engaging the non-rotating ring  110  and being fixed relative to the non-rotating ring  110 . A plurality of rolling elements  334 , such as ball bearings, are captured between the outer race  330  and the inner race  332 . A cage  336  may also be positioned between the inner and outer races  330 ,  332  to maintain the rolling elements separated from one another and evenly distributed around the races  330 ,  332 . 
     The rolling elements  334  may be preloaded such that they are deformed from an undeformed shape even in the absence of any loads on the rotating ring  108  or non-rotating ring  110 . Preloading the rolling elements  334  may eliminate slop or play between the rotating ring  108  and non-rotating ring  110  that would exist if gaps were present between the rolling elements  334  and the inner and outer races  330 ,  332 . Due to thermal contraction of the rolling elements  332 ,  334 , rotating ring  108 , and non-rotating ring  110 , the preloaded condition of the rolling elements  334  may be reduced or disappear. 
     In embodiments of the present invention, air directed through the air channel  92  to the tip jets  50  may be at an elevated temperature due to the input of energy during compression of the air. In some embodiments, the temperature of air forced through the air channel  92  may be above 300° F. when the tip jets  50  are ignited. As a result, during vertical takeoff and landing or during hover, the preload of the bearing elements  334  will be increased due to thermal expansion of the bearings elements  334 , races  330 ,  332 , rotating ring  108 , and non-rotating ring  110 . However, during sustained longitudinal flight at high speeds and altitudes, hot compressed is no longer needs to flow to the tip jets  50  and the ambient air temperature can be very low. For example, above an altitude of 8000 ft, the air temperature is typically at or below 32° F. Accordingly, the preload of the rolling elements  334  may decrease to zero and gaps may occur between the rolling elements  334  and the races  330 ,  332 , resulting in increased slop or play between the rotating plate  108  and the non-rotating plate  110 . Increases in slop or play between the rotating plate  108  and the non-rotating plate  110  may result in destructive harmonics at the frequency of cyclic loads on the blades  42  during high speed flight. 
     Accordingly, in some embodiments, oil flowing through the bearings  238 ,  240  may be selectively cooled to prevent over loading or heat related failure of the rolling elements  334  due to hot air flow and heat buildup due to friction. The oil flowing through the bearings  238 ,  240  may also be selectively heated to prevent cooling to the point that the rolling elements  334  are no longer preloaded or the preload of the rolling elements  334  is below a predetermined threshold. 
     Each of the seals  234 ,  236  may include an outer seal  340  and an inner seal  342 . In the illustrated embodiment, the outer seal  340  and inner seal  342  are mirror images of one another. The outer seal  340  may be separated from the inner seal  342  by a spacer  344 . The seals  340 ,  342  may include a sealing material  346  disposed in a ring and defining a sealing surface  348  for engaging corresponding sealing surfaces  266 ,  286  of the rotating ring  108  and non-rotating ring  110 , respectively. The sealing material  346  may define a groove  350  having a circumferential spring  352  positioned therein and biased to urge the sealing surface  348  against the sealing surface  266  or sealing surface  286 . The groove  350  of the upper seal  340  and the groove  350  of the inner seal  342  may face away from one another. The sealing material  346  may be mounted within a retainer  354  formed of a metal or other rigid material for maintaining the shape of the sealing material  346  during use and installation of the upper and lower seals  340 ,  342 . 
     The spacer  344  may include one or more grooves  356  extending radially therethrough. The grooves  356  may permit the passage of any oil leaking between the outer and inner seals  340 ,  342  to flow into a fluid path  358  formed in the upper seal seat  242  non-rotating ring  110 . The fluid path  358  may be in fluid communication with the return port  300   b . The fluid path  358  preferably connects to the fluid path between the feed port  300   a  and the return port  300   b  at a point that is at a lower pressure than oil flowing adjacent the inner seal  342  such that oil tends to flow only outwardly from the space between the seals  340 ,  342  into the fluid path  358 . 
     Referring to  FIG. 16 , the hub  44  may mount to the mast  46  by means of an upper bearing  370  and a lower bearing  372  captured between the hub  44  and the mast  46 . The bearings  370 ,  372  may be embodied as tapered roller bearings including tapered rolling elements  374 . Tapered rolling elements  374  advantageously support loads perpendicular to the axis of rotation  72  of the hub  44  and support lift loads parallel to the axis of rotation  72 . The axes  376  of the tapered rolling elements  376  form an angle  378  with the axis of rotation  72 . The angle  378  is dictated by the magnitude of radial and longitudinal forces. For example, the angle  378  may be between 3 and 20 degrees. The angle  378  of the rolling elements  374  of the upper bearing  270  may be unequal that of the rolling elements  374  of the lower bearing  372 . As is apparent from  FIG. 16 , the rolling elements  374  of the upper bearing  370  are angled away from the axis of rotation  72  with upward distance along the axis of rotation  72  whereas the rolling elements  374  of the lower bearing  372  are angled away from the axis of rotation  72  with downward distance along the axis of rotation  72 , where the downward and upward directions are in the frame of reference of the page. The opposing orientation of the angles  378  enables support of loads in both directions parallel to the axis of rotation  72 . 
     Each of the upper bearings  370 ,  372  includes a cup  380  and a cone  382  as known in the art of tapered rolling design. The cup  380  extends around the rolling elements  374  and includes a shallow channel  384 , or depression  384 , for retaining the rolling elements  374 . The cup  380  of the upper bearing  370  faces opposite the cup  380  of the lower bearing  372 . The cone  382  is located among the rolling elements  374  having the rolling elements  374  captured between the cone  382  and the cup  384 . The cone  382  includes a channel  386 , or depression  386 , for retaining the rolling elements  374 . For each of the upper bearings  370 ,  372 , the rolling elements  374  are captured between the channel  384  of the cup  380  and the channel  386  of the cone  382 . 
     Referring to  FIG. 17 , the hub  44  includes an upper bearing seat  390  and a lower bearing seat  392 . The cups  380  of the upper and lower bearings  370 ,  372  are positioned within the bearings seats  390 ,  392 , respectively. Each of the seats  390 ,  392  includes a vertical wall  394  having a cylindrical shape parallel to the axis of rotation  72  of the hub  44 . The seats  390 ,  392  further include a radial wall  396  extending from an edge of the vertical wall  394  radially inward toward the axis of rotation  72 . The radial walls  396  of the seats  390 ,  392  face outwardly in opposite directions from one another. 
     The vertical wall  394  of the seats  390 ,  392  may define one or more circumferential grooves  398 . The Seats  390 ,  392  may likewise define grooves  400  extending continuously from vertically along the vertical wall  394  to radially along the radial wall  396 . The grooves  398 ,  404  may facilitate the flow of oil around the bearings  370 ,  372 . 
     Referring to  FIG. 18 , the fluid lines  132  emanating from the hydraulic rotary union  138  may include feed oil line  410  and a return oil line  412 . The hub  44  may define a feed port  414  receiving a fitting  416  coupling the feed oil line  410  to the feed port  414 . Likewise, the hub  44  may define a return port  418  receiving a fitting  420  coupling the return oil line  412  to the return portion  418 . A feed oil passage  422  may extend through the hub  44  from the feed port  414  to adjacent the upper bearing seat  390 . In the illustrated embodiment, the feed oil passage  422  is in fluid communication with one or both of the circumferential groove  398  or the grooves  400  of the upper bearing seat  390 . In a like manner, a return oil passage  424  extends through the hub  44  from the return port  418  to adjacent the lower bearing seat  392 . In the illustrated embodiment, the return oil passage  424  is in fluid communication with one or both of the circumferential groove  398  or the grooves  400  of the lower bearing seat  392 . In the illustrated embodiment, the return oil passage  424  is formed in a ridge  426  projecting outwardly from surfaces  428  of the hub  44  having a circular cross section. A corresponding ridge  430  providing a counter balance for the ridge  426  may be formed opposite the ridge  426  on the hub  44  and the feed passage  422  may extend through a portion of the ridge  430 . Although individual components have been labeled as a “feed” and “return” type components, each of these labels could be reversed for oil flow in the opposite direction. 
     An upper seal  432  may be interposed between the hub  44  and mast  46  above the upper bearing  370  to hinder leakage of oil therefrom. In the illustrated embodiment, an upper cap  434  secures a to the hub  44  in order to prevent air flow out of the cavity  94 . In such embodiments, the upper seal  432  may engage a downwardly depending flange  436  secured to or formed monolithically with the cap  434 . The seal  432  also engages the mast  46  to create a seal between itself and the flange  436 . In some embodiments, the seal  432  may directly engage a corresponding sealing surface of the hub  44  or some other structure secured to the hub  44 . 
     Referring to  FIG. 19 , in some embodiments, the return passage  424  may be replaced with a extends from adjacent the lower bearing seat  292  to a passage  440  extending through the center of the mast  46 . The electrical lines  130  and fluid lines  132  may also pass through the passage  442 . Oil  442  passing through the return passage  424  collects within the passage  440 . In such embodiments, a return oil port  444  coupled to the hydraulic rotary union  138  may conduct oil to the hydraulic rotary union  138 , which will conduct the oil away from the passage  440 . One or more seals  446  engaging the hydraulic rotary union  138  may prevent leakage of the oil past the hydraulic rotary union  138 . 
     Referring to  FIG. 20 , in some embodiments a lower seal  450  hinders leakage of oil away from the lower bearing  372  between the hub  44  and the mast  46 . In the illustrated embodiment, the upper and lower bearings  370 ,  372  are positioned between the upper seal  432  and the lower seal  450 . In the illustrated embodiment, the seal  450  secures to a mounting ring  452  that is secured to the hub  44 . The mounting ring  452  may define a seal seat  454  for receiving the seal  450  and a seal clamp  456  selectively secured to the mounting ring  452  by means of a fastener  458 , such as a screw, such that the seal  450  is captured between the seal seat  454  and the seal clamp  456 . The mounting ring  452  additionally include a circular flange  460  secured to the hub  44  such as by means of fasteners  462 . The flange  460  may create a sealed interface between itself and the hub  44  to hinder oil leakage from between the flange  460  and the hub  44 . The mounting ring  452  may additionally include one or more flanges  464 . The flanges  464  may extend vertically downward from the flange  460 . The mast fairing  142  may secure to the flanges  464  by means of fasteners  466 , such as screws, bolts, rivets, or the like. In addition to reducing pressure losses in air flow over the transition from the mast  46  to the hub  44 , the mast fairing  142  may additionally protect the seal  450  from heated air flow thereover which could degrade the seal  450  and blow past the seal  450  and strip oil from the bearings  370 ,  372 . 
     The seal  450 , may include an inner seal  468  and an outer seal  470  that have identical configurations but mirrored about a horizontal plane. The seals  468 ,  470  may have a spacer  472  positioned therebetween. The seals  468 ,  470  and spacer  472  may have the same configuration as the seals  340 ,  342  and spacer  344  of the swashplate  104  as discussed hereinabove. 
     Referring to  FIG. 21 , the blade spars  96  may be supported within a spar bore  490  extending through the hub  44  to the cavity  94 . An inboard bearing  492  and an outboard bearing  494  may support rotation of the blade spar  96  within the spar bore  490  in order to change the pitch angle  78  of the blade  42  including the blade spar  96 . The bearings  492 ,  494  may be embodied as tapered roller bearings. The fluid lines  132  emanating from the hydraulic rotary union  138  may include a feed oil line  496  and a return oil line  498  coupled to a feed oil port  500  and a return oil port  502 , respectively. The feed oil port  500  and return oil port  502  penetrate the hub  44  and are in fluid communication with the spar bore  490 . An outboard seal  504  and an inboard seal  506  may be interposed between the blade spar  90  and the spar bore  490  having the inboard bearing  492  and the outboard bearing  494  positioned between the outboard seal  504  and the inboard seal  506  in order to prevent leakage of oil. 
     As with other bearings described herein, the bearings  492 ,  494  may include rolling elements  508  that are preloaded within a certain operating temperature range. Heating and cooling of the oil passing between the feed port  500  and the return port  502  may be used to prevent overheating of the rolling elements  508  due to heated air flow or heat buildup due to friction and to prevent over reduction or elimination of the preload due to thermal shrinkage of the bearings  492 ,  494 , hub  44 , or blade spar  96 . 
     The bearings  492 ,  494  may be protected from heated air flow thereover by the blade duct fairing  146 , which may extend from the blade duct  60  to engage the wall of the cavity  94  of the hub  44 . The blade duct fairing may describe smooth contour from the blade duct  60  to the wall of the cavity  94  such that air flow from the cavity to the blade duct experiences a smaller pressure drop, such as between five and fifty percent, preferably between twenty percent and fifty percent, lower than the pressure drop that would result if the blade duct fairing  146  were removed. 
     A more complete description of the coupling between the blade spar  96  and the hub  44  may be found in U.S. Prov. Pat. App. Ser. No. 61/403,097, filed Sep. 9, 2010 and entitled “FEATHERING-SPINDLE-BEARING LUBRICATION AND TEMPERATURE CONTROL”. 
     Referring to  FIG. 22 , the flow of oil for the lubrication and cooling or heating of the bearings within the rotor  40  may be performed by an oil distribution system  20 . The system  20  includes an oil pump  522  powered by one or both of the engines  48  or by electrical or hydraulic power generated by a generator or hydraulic pump powered by the engines  48 . The pump  522  may be constantly on or may be turned on and off according to conditions of the rotor  40 , such as the temperature of one or more bearings thereof. The low pressure port of the pump  522  draws oil from a reservoir  524 . 
     The high pressure port of the pump  522  is coupled to a thermal modulation system  526 . The thermal modulation system  526  senses and responds to the temperature within the rotor system  40 . The thermal modulation system  526  may extract thermal energy from the oil within the oil distribution system  520  in order to lower the temperature of bearings within the rotor  40  in order to avoid bearing failure due to high heat. The thermal modulation system  526  may input thermal energy to the oil in order to raise the temperature of bearings within the rotor  40  in order to avoid thermal shrinkage that will reduce the preload of rolling elements within the bearings below acceptable levels or eliminate the preload of the rolling elements within the bearings entirely. 
     The thermal modulation system  526  may include one or more radiators  528  and one or more fans  530  directing air at the radiators  528  in order to extract thermal energy from oil within the oil distribution system. The radiators  528  are located within the fluid path between the high pressure port and low pressure port of the pump  522 . 
     The thermal modulation system  526  may include one or more heating elements  532  in thermal contact with oil within the oil distribution system  520 . The heating elements  532  may be selectively activated to input heat into oil within the oil distribution system  520 . In some embodiments, a bypass valve  534  directs oil to either the radiators  528  or heating elements  532  according to the need for heat input to the oil or heat extraction from the oil. 
     A thermal valve  536  in thermal contact with oil within the oil distribution system  520  may control the fans  530 , heating elements  532 , and bypass valve  534  according to a temperature of oil within the oil distribution system  520 . The thermal valve  536  may be a simple thermostatic switch or may be a digitally programmable sensor and actuator having the capacity to independently control each of the fans  530 , heating elements  532 , and bypass valve  534  in order to modulate the temperature of oil within the oil distribution system  520 . 
     Oil within the oil distribution system  520  may flow over a swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 . The swashplate bearing set  538  may include the swashplate bearings  338 ,  340 . Oil flow through the swashplate bearing set  538  may pass through the feed port  300   a  and return port  300   b  as described hereinabove. The mast bearing set  540  may include the upper bearing  370  and lower bearing  372 . Oil flow through the mast bearing set  540  may include oil flow through the feed port  414  and the return port  418 . The spindle bearing set  542  may include the inboard bearing  492  and the outboard bearing  494 . Oil flow through the spindle bearing set  542  may pass through the feed port  500  and return port  502  as described hereinabove. 
     In the illustrated oil distribution system  520 , oil flows through the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542  in parallel through separate fluid paths  544   a ,  544   b ,  544   c . Temperature controlled valves  546   a ,  546   b ,  546   c  may control oil flow through the paths  544   a ,  544   b ,  544   c , respectively, according to the temperature of oil exiting the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 , respectively. Oil flow through the paths  544   a ,  544   b ,  544   c  may return to the reservoir  524  after exiting the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 . 
     Thermal sensors  548   a ,  548   b ,  548   c  may be in thermal contact with oil flowing through the paths  544   a ,  544   b ,  544   c  downstream from the bearing sets  538 ,  540 ,  542 , respectively. In some embodiments, an additional temperature sensor  548   d  may sense the temperature of oil flowing from the pump upstream from the radiators  528  and heating elements  532 . 
     The order of elements along the fluid path between the high pressure port and lower pressure port of the pump  522  may be different that that illustrated in  FIG. 22 . Oil distribution systems according to embodiments of the invention may also include only one of a radiator  528 , or other cooling system, and a heating element  532 . Likewise, more or fewer bearing sets than the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542  may be lubricated, heated, and/or cooled by oil flowing through the oil distribution system  520 . 
     Referring to  FIGS. 23A and 23B , preloading of the bearing sets swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 , may additionally or alternatively be maintained despite low ambient air temperatures, such as during high altitude sustained longitudinal flight, by means of heated air forced or drawn through the air channel  92  and the cavity  94 . 
     Referring specifically to  FIG. 23A , during sustained longitudinal flight the rotor hub  44  and blades  42  continue to rotate though unpowered due to autorotation. The centrifugal force exerted on air within the blade ducts  60  may draw air through the air channel  92  formed by the mast shroud  58  and the mast  46  and into the cavity  94  formed in the hub  44 . 
     In some embodiments, one or more heating elements  550   a ,  550   b ,  550   c  are positioned within one or more of the plenum  56  and ducts  54  and are selectively powered to heat air drawn into the mast shroud  58 . As noted above, compressed air from the bypass turbines  62  may be urged through the ducts  54  during takeoff, landing, and hover. However, during sustained longitudinal flight, the engines  48  may operate more efficiently by directing all bypass air rearwardly from the engines  48  rather than through the ducts  54 . Accordingly, one or more valves  552   a ,  552   b  may turn off air flow from the bypass turbine  62  to the ducts  54  during sustained longitudinal flight of the aircraft  10 . However, to permit air flow over the heating elements  550   a ,  550   b ,  550   c  as needed to heat the rotor  40 , the valves  550  may be partially opened during sustained longitudinal flight. In some embodiments, to avoid drawing power from the engines  48 , one or more valves  554   a ,  554   b  may selectively permit air flow into the ducts  54  or directly into the plenum  56 . The opening and closing of the one or more valves  554   a ,  554   b  may be controlled by temperature feedback from the rotor  40 . 
     For example, a sensor monitoring the temperature of the oil within the oil distribution system  520  may indicate when the oil temperature drops below a certain threshold such that the heating elements  532  are no longer sufficient to maintain the bearings sets  538 ,  540 ,  542  within an operating temperature range at which the rolling elements thereof are preloaded, the one or more valves  554   a ,  554   b  may be partially or completely opened and the one or more heating elements  550   a ,  550   b ,  550   c  may be activated to warm the rotor  40  to the proper operating temperature range at which preloading of the rolling elements within the bearing sets  538 ,  540 ,  542  is above a predetermined threshold. 
     Referring to  FIG. 23B , in some embodiments, in lieu of heating elements  550  bleed air from a stage of the engines  48  prior to the combustion stage  560  may be conducted by means of one or more channels  562  to one or both of the ducts  54  or the plenum  56 . The channel  562  preferably bypasses any valve  552   a ,  552   b  controlling bypass air from the bypass turbine  62 . Air flow through the channels  562  may be controlled by one or more valves  564  that may be controlled according to a temperature of one or more of the bearing sets  538 ,  540 ,  542  of the rotor  40 . In the illustrated embodiment, the engine  48  includes two compression stages  566   a ,  566   b  and the channel  562  may be positioned to draw air from either of the compression stages  566   a ,  566   b.    
     Referring to both  FIGS. 23A and 23B , either system may be used to device the blades  42  inasmuch as the heated air eventually flows through the blade ducts  60 . 
     Referring to  FIG. 24 , a flight control system  580  may include flight controls  582  providing activating signals, such as electrical, hydraulic, or mechanical inputs, to the swashplate actuators  112 , mast tilt actuator  174 , one or more throttle actuators  584  for controlling the engines  48  and tip jets  50 , and control surface actuators  586  for controlling the control surface such as the rudder  22 , elevator  30 , and ailerons  36 . The flight controls  582  may receive pilot inputs  588  from pilot controls as known in the art of fixed and rotary aircraft design such as rudder control pedals, aileron and elevator control stick, cyclic pitch control stick, throttle control lever, and cyclic pitch control knob. The flight controls  582  may additionally receive inputs from an avionic computer  590  enabling autopiloted flying of the aircraft  10 . 
     The flight control system  580  may additionally include a thermal management module  592  programmed to maintain the temperature of the rotor  40  effective to avoid bearing failure and to maintain bearings and structures in which they are mounted within an operating temperature range in which the bearings will be in a preloaded condition or a preloaded condition above a minimum preload. The thermal management module  592  may receive inputs from the temperature sensors  548   a ,  548   b ,  548   c ,  548   d  measuring the temperature of oil exiting the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 , respectively. The thermal management module  592  may be electrically, hydraulically, or mechanically coupled to the valves  552   a ,  552   b  controlling flow of bypass air from the engines  48 , the valves  554   a ,  554   b  controlling the air passively drawn into the ducts  54  or plenum  56 , the valves  564  controlling the flow of bleed air from the engines  48  into the ducts  54  or plenum  56 , the temperature modulation system  526 , including the fans  530  and the heating elements  532 , and the thermal valves  536 ,  546   a ,  546   b ,  546   c  of the oil distribution system  520 . In such embodiments, the thermal valves  536 ,  546   a ,  546   b ,  546   c  may be embodied as electrically, hydraulically, or mechanically actuated valves controlled by the thermal management module  592 . 
     The thermal management module  592  may be embodied as a digital or analog computer programmed to respond to inputs from some or all of the sensors  548   a ,  548   b ,  548   c ,  548   d  by activating one or more of the devices electrically, hydraulically, or mechanically coupled thereto. Alternatively, the thermal management module may be distributed such that each device listed in the preceding paragraph is activated or deactivated according to a sensed temperature. In particular, the thermal valves  536 ,  546   a ,  546   b ,  546   c  may respond independently to the temperature of oil flowing therethrough and open and close according to whether the temperature is within a set operating temperature range, e.g., a temperature range between the temperature at which the rolling elements of the bearing sets  538 ,  540 ,  542  will fail and the temperature at which the preloading of the rolling elements is still present or is above a proscribed threshold providing the needed rigidity of the rotor  40  against destructive harmonics. 
       FIG. 25  illustrates a method  600  for thermal management of a rotor  40 , and, in particular, for reducing probability of bearing failure and decreasing the amount of slop or play within the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing  542  by maintaining an adequate preload upon the rolling elements thereof. The method  600  may be performed by the thermal management module  592 . 
     The method  600  may include evaluating at step  602 , whether the aircraft  10  is taking off, landing, or hovering. The aircraft  10  may be capable of horizontal takeoff along a runway, in which case taking off and landing for purposes of step  602  may include evaluating whether a vertical or short landing or take off is being performed such that powered rotation of the rotor  40  by means of the tip jets  50  is needed to achieve the degree of verticality of the landing or take off. If hovering, taking off, or landing, is being performed, then at step  604  the tip jets are activated at step  604  and compressed air for driving the tip jet  50 , such as compressed air from the engines  48 , is forced through the rotor  40  to the tip jets  50  at step  606 . Steps  604  and  606  may be performed simultaneously and either step  604  or step  606  may be begun first. If the aircraft  10  is no longer taking off, landing, or hovering, then the tip jets  50  are deactivated at step  608  and at step  610 , the compressed air from the bypass turbine  62  is directed rearwardly, in embodiments having engines  48  embodied as jet engines. 
     Throughout operation of the aircraft, for both sustained longitudinal flight and vertical flight as in a vertical takeoff or landing or hovering, the method  600  may include executing some or all of steps  612  through steps  618 . 
     At step  612  the method  600  includes evaluating whether the temperature of the rotor  40  is above an upper threshold, such as a temperature within some tolerance of the temperature above which the bearings of the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing  542  will fail or have an unduly shortened useful life. If so, then at step  614 , thermal energy is extracted from the rotor  40 . Extracting energy from the rotor may include activating the fans  530  in order to increase the rate of heat transfer from the radiators  528 . 
     At step  616 , the temperature of the rotor  40  is evaluated with respect to a lower threshold equal to or some tolerance above the temperature at which the bearings of the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing  542  are no longer preloaded or have a preload below a minimum preload magnitude. If so, then at step  618 , thermal energy is added to the rotor  40  according to the functionality of the oil and air heating systems described hereinabove. Adding thermal energy to the rotor may include one or more of activating the heating elements  532 , activating the heating elements  550   a - 550   c  and opening the valves  554   a ,  554   b , and opening a valve  564  permitting flow of bleed air over the rotor  40 . The possible methods of adding heat to the rotor  40  may be performed simultaneously or may be attempted in a specified order such that one method is attempted alone, then another method is attempted simultaneously if the temperature increase is insufficient, other methods may then be attempted simultaneously if the temperature increase is again insufficient. 
     Referring to  FIG. 26 , in some embodiments, a method  620  may be used for thermal management of a rotor  40 . The method  620  may include evaluating whether the average temperature of oil flowing through the rotor  40  has an average temperature above the operating temperature range at step  622 . If so, at step  624 , heat is extracted from the oil, such as by activating the fans  530  blowing the radiators  528 . 
     At step  626 , the method  620  includes evaluating whether the temperature of oil circulating through the rotor  40  has a temperature below the operating temperature range. If so then at step  628 , heat is input to the radiator. Adding thermal energy to the rotor may include one or more of activating the heating elements  532 , activating the heating elements  550   a - 550   c  and opening the valves  554   a ,  554   b , and opening a valve  564  permitting flow of bleed air over the rotor  40 . The possible methods of adding heat to the rotor  40  may be performed simultaneously or may be attempted in a specified order such that one method is attempted alone, then another method is attempted simultaneously if the temperature increase is insufficient, other methods may then be attempted simultaneously if the temperature increase is again insufficient. 
     Steps  622  and  626  may include measuring the temperature of oil entering or exiting the pump  522  upstream of the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542 . Steps  622  and  626  may be performed by the thermal valve  536  and steps  624  may include activation of the fans  530  by the thermal valve  536 . Opening and closing of one or more of the valves  554   a ,  554   b ,  564  and activating of the heating elements  532 ,  550   a - 550   c  may also be controlled according a temperature dependant signal from the thermal valve  536 . Alternatively, or in addition, activation of the fans  530 , opening and closing of the valves  554   a ,  554   b ,  564 , and activation of the heating elements  532 ,  550   a - 550   c  may be controlled by a digital or analog computer, such as the thermal management module  592 . In such embodiments, steps  622  and  626  may include evaluating the output of the thermal sensor  548   d . The thermal sensor  548   d  preferably measures the temperature of the consolidated flow of oil from each of the paths  544   a ,  544   b ,  544   c , such as at a point between the high pressure port of the pump  522  and the radiators  528  and the heating elements  532 . 
     At steps  630 ,  632 , and  634 , the temperatures of oil flowing through the paths  544   a ,  544   b ,  544   c , respectively, downstream from the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542  are evaluated to determine whether the temperatures lie within the operating temperature range. If the temperature of oil flowing through any of the paths  544   a ,  544   b ,  544   c  downstream from the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542  is determined to lie outside of the operating temperature range, then at steps  636 ,  638 , and  640  oil flow through whichever of the paths  544   a ,  544   b ,  544   c  has a temperature outside of the operating temperature range is increased. If the temperature of oil flowing through any of the paths  544   a ,  544   b ,  544   c  downstream from the swashplate bearing set  538 , mast bearing set  540 , and spindle bearing set  542  is determined to lie within the predetermined range, then at steps  642 ,  644 , and  646 , oil flow through whichever of the paths  544   a ,  544   b ,  544   c  has a temperature outside of the operating temperature range is decreased. 
     Evaluating the temperature of oil flow through the paths  544   a ,  544   b ,  544   c  may be performed by the thermal valves  546   a ,  546   b ,  546   c , respectively. Evaluating the temperature of oil flow through the paths  544   a ,  544   b ,  544   c  may additionally or alternatively be performed by the thermal sensors  548   a ,  548   b ,  548   c , respectively and the thermal valves  554   a ,  554   b ,  554   c  may be replaced by valves opened and closed by the thermal management module  592  electrically, hydraulically coupled to the valves in order to increase or decrease the flow of oil through the paths  554   a ,  544   b ,  544   c.    
     Referring to  FIG. 27 , as noted above, the swashplate  104  may be mounted to a spherical bearing  106  to enable the orientation of the swashplate  104  to be changed in response to actuation of the rods  114 . The spherical bearing  106  may be slidably mounted to the mast  46  to enable adjustment of collective pitch by sliding the swashplate  104  relative to the mast  46 . 
     As also noted above, the swashplate  104  is positioned within an air channel defined by the shroud  58  and the mast  46 . When the tip jets  50  are in operation, compressed air from the one or more engines  48  is directed through this air channel. Compression of the air raises its temperature such that air flowing around the swashplate  104  may have a temperature of above 300° F. In contrast, during sustained translational flight at high altitudes, the tip jets  50  may be deactivated and the temperature of the swashplate  104  may drop below 32° F. Due to the high temperature of air flowing over the swashplate  104 , conventional petroleum-based lubricants are ineffective because they may be ignited by the high temperatures or be stripped away from the swashplate  104  by the hot air flow. 
     In some embodiments, one or more bearing surfaces defusing interfaces between the swashplate  104  and other structures engaging the swashplate  104  and transmitting blade pitch controls to the blades  42  may be coated with a solid lubricant. In some embodiments, both bearing surfaces at an interface between components may bear a solid lubricant in order to achieve a desired reduction in friction. The solid lubricant may include molybdenum disulphide (MoS 2 ), Boron Nitride (BN), or polytetrafluoroethylene (PTFE). In particular, MoS 2  and BN may be fused to surfaces to which they are applied, such as by means of baking at an elevated temperature. PTFE, which is a thermopolymer with a melting point of 621° F., may fused to surfaces to which it is applied by applying it in a liquid state. Hereinbelow, a “low friction coating” may refer to any one of these three coatings, or to other solid, low-friction coatings that have an operating temperature within the range above about 300° F., preferably above about 350° F., and more preferably above about 400° F. 
       FIG. 27  illustrates an interface between the mast  46  and the swashplate  104  that may bear a low friction coating. The spherical bearing  106  includes a spherical bearing surface  670  that engages the swashplate  104 . The surface  670  may bear a low friction coating. In a like manner, the spherical bearing seat  252  defined by the upper clamping ring  246  may also bear a low friction coating. A surface  672  of the non-rotating ring  110  engaging the surface  670  may also bear a low friction coating. The spherical bearing  106  may engage a surface  674  of the mast  46  by means of a surface  676 . One or both of the surfaces  674 ,  676  may include a low friction coating. 
     Referring to  FIG. 28 , while still referring to  FIG. 27 , other interfaces may also bear low friction coatings, such as the interface between the swashplate  104  and the actuator rods  114 . The actuator rods  114  may engage the swashplate  104  by means of actuator rod mounts  232 . In some embodiments, the rods  114  are coupled to the mounts  232  by means of a pin  678 . One or both of an aperture  680  defined by the actuator rod  114  and a surface  682  of the pin  678  engaging the aperture  680  may bear a low friction coating. 
     In some embodiments, the aperture  680  may be formed in a roller bearing  684  secured to the rod  114 . In such embodiments, a low friction coating of the aperture  680  and pin surface  682  may be omitted. One or more components of the roller bearing  684 , such as the inner and outer race, cage, and rolling elements, may also bear a low friction coating. The actuator rod mounts  232  may define one or more apertures  686  for receiving the pin  678 . The apertures  686  may also bear a low friction coating or be defined by roller bearings having one or more components bearing low friction coatings. 
       FIGS. 29 and 30 , in combination with  FIG. 27 , illustrate an interface between the pitch control rods  116  and the swashplate  104  bearing a low friction coating. The pitch control rods  116  may define an aperture  690 , which may itself be defined by an inner race of a roller bearing  692 . As for the roller bearing  684 , one or more of the inner race, outer race, cage, and rolling elements may bear a low friction coating. 
     A pin  694  may insert within the aperture  690  and an aperture  696  formed in the pitch control rod mount  230 . Where relative movement between the pin  694  and the aperture  690  is permitted, one or both of the aperture  690  and the surface  698  of the pin  694  may be bear a low friction coating at an interface therebetween. Where relative movement is permitted between the pin  694  and the aperture  696 , one or both of the pin surface  698  and aperture  696  may bear a low friction coating. The apertures  690  and  696  may also be defined by roller bearings having one or more components bearing low friction coatings. 
       FIGS. 31 through 33  illustrate an interface between the rotating stabilizer linkage  150  and the mast fairing, and an interface between the rotating stabilizer linkage  150  and the swashplate  104 , each of which may bear a low friction coating. As noted above, the rotating stabilizer linkage  150  includes a hinge  156  coupling legs  154   a ,  154   b  to one another. 
     A hinge  700  may likewise couple the leg  154   b  to the mast fairing  142 . The hinge  700  may include a pin  702  including a pin bearing surface  704  engaging an aperture  706  extending through the leg  154   b  and a mounting structure  708  secured to the mast fairing  142 . 
     One or more of the surface  704 , portions of the aperture  706  extending through the leg  154   b , and portions of the aperture  706  extending through the mounting structure  708  may bear a low friction coating. As with the actuator rods  114  and pitch control rods, the aperture  706  may be defined by one or more roller bearings secured to one or both of the leg  154   b  and the mounting structure  708  and having one or more components bearing low friction coatings. 
     The hinge  156  may also include a pin  710  positioned within an aperture  712  extending through the legs  154   a ,  154   b . One or both of a bearing surface  714 , portions of the aperture  712  passing through the leg  154   a , and portions of the aperture  712  passing through the leg  154   b  may bear a low friction coating. The aperture  712  may be defined by one or more roller bearings secured to one or both of the leg  154   a  and the leg  154   b  and having one or more components bearing low friction coatings. 
     Referring to  FIGS. 32 and 33 , while still referring to  FIG. 31 , the leg  154   b  may couple to the rotating ring  108  by means of a spherical bearing  716 . The spherical bearing  716  may engage a seat  718  formed in a linkage mount  160  and retained in place by a retaining ring  720 . One or more of the spherical bearing  716 , seat  718 , and a lower surface  722  of the retaining ring  720  may bear a low friction coating. 
       FIGS. 34 through 36  illustrate an interface between a non-rotating stabilizer linkage  152  and the mast  46 , and between the non-rotating stabilizer linkage  152  and the swashplate  104 , each of which may bear a low friction coating. The hinge  156  of the non-rotating stabilizer linkage  152  may include a pin  730  engaging an aperture  732  passing through the legs  154   a ,  154   b  of the linkage  152 . One or both of the bearing surface  734  of the pin  730 , portions of the aperture  732  passing through the leg  154   a , and portions of the aperture  732  passing through the leg  154   b  may bear a low friction coating. The aperture  732  may be defined by one or more roller bearings secured to one or both of the leg  154   a  or the leg  154   b  and having one or more components bearing low friction coatings. 
     The leg  154   b  may also couple to the mast  46  by means of a pin  736  passing through an aperture  738  extending through a mast mounting structure  740 . One or both of a bearing surface  742  of the pin and the aperture  738  may bear a low friction coating. The aperture  738  may be defined by a roller bearing having one or more components bearing a low friction coating. 
     Referring to  FIGS. 35 and 36 , while still referring to  FIG. 34 , the leg  154   a  may secure to the non-rotating ring  110  by means of a pin  744  having a bearing surface  746  engaging an aperture  748  extending through a linkage mount  162  and an aperture  750  passing through the leg  154   a . One or both of the apertures  748 ,  750  may be defined by a roller bearing  752  having one or more components bearing low friction coatings. 
       FIG. 37  illustrates an interface, between a pitch control rod  116  and the pitch horn  118 , that may bear a low friction coating. A pitch control rod  116  may be mounted to the pitch horn  118  by means of a pin  760  having a bearing surface  762 . The pin may insert through an aperture  764  defined by the pitch horn  118  and an aperture  690  defined by the pitch control rod  116  (See  FIG. 29 ). One or more of the aperture  764 , aperture  690 , and bearing surface  762  may bear a low friction coating. One or both of the apertures  690 ,  764  may be defined by a roller bearing  768  having one or more components bearing a low friction coating. 
     Referring to  FIG. 38 , the low friction coating may include a layer  770  of solid lubricant fused to a surface  772  of a substrate  774  composing any of the parts described hereinabove as bearing a low friction coating. The thickness  776  of the layer  770  may be effective to provide a low friction interface for a predetermined number of operating hours. For example, in some embodiments, a layer  770  may have a thickness  776  of between about 0.001 and about 0.003 inches when formed of MoS 2  or PTFE. In some embodiments a layer  770  formed of BN may have a thickness between 0.1 μm and 2 mm. These thicknesses are exemplary only. However, maintaining fit tolerances is best achieved by controlling this thickness  776 . The thickness  776  may have any value that is achievable for a given coating material and suitable for expected operating conditions. 
     The present invention may be embodied in other specific forms without departing from its spirit or essential characteristics. The described embodiments are to be considered in all respects only as illustrative, and not restrictive. The scope of the invention is, therefore, indicated by the appended claims, rather than by the foregoing description. All changes which come within the meaning and range of equivalency of the claims are to be embraced within their scope.