Patent Document

RELATED APPLICATIONS 
     This application claims the benefit of U.S. Provisional Patent Application Ser. No. 61/381,347, filed on Sep. 9, 2010, U.S. Provisional Patent Application Ser. No. 61/403,136, filed on Sep. 9, 2010, 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, and U.S. Provisional Patent Application Ser. No. 61/499,996, filed on Jun. 22, 2011. 
     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. 
     BACKGROUND 
     1. The Field of the Invention 
     This invention relates to rotating wing aircraft, and, more particularly to rotating wing aircraft relying on autorotation of a rotor to provide lift. 
     2. The Background Art 
     Rotating wing aircraft 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 rotating wing aircraft in use today are helicopters. A helicopter typically includes a 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 rotating wing aircraft 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 rotating wing aircraft 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 rotating wing aircraft are unable to achieve high flight speed is a phenomenon known as “retreating blade stall.” As the fuselage of the rotating wing aircraft 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 solutions 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 rotating wing aircraft 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 rotating wing aircraft 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 rotating wing aircraft 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 rotating wing aircraft to the maximum air speed of the tips of the rotor blades is known as the “advance ratio. The maximum advance ratio of rotary wing aircraft available today is less than 0.5, which generally limits the top flight speed of rotary wing aircraft 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. 
     A helicopter rotor can be operated at controlled rotational speeds by external airflows only. For example, without the additional power added to the shaft rotating the rotor blades, the rotor blades or rotary wings can autorotate, operating like a windmill. However, autogyros typically, a helicopter uses a power rotor, which therefor has the rearward portion of its operating disk (the theoretical plane in which the blades rotate) upward, with the front portion relatively downward in order to both lift the aircraft up and draw it forward. In contrast, autogyros typically operate with the rotor disk in opposite configuration with the upper front edge relatively higher and the trailing edge of the rotor disk relatively lower in order that relative airflow past the rotor tends to windmill or autorotate the rotor. Thus, the rotary wing provides both windmill autorotation to rotate itself, as well as providing the Bernoulli effect of lift over the airfoil shape of each rotor blade. Thus, at least a portion of the blade or airfoil is dedicated to or responsible for providing autorotation, and at least a portion of the blade is providing airfoil lifting force. Therefore, in an autogyro, forward propulsion is provided typically by a propellor or other device separate from the rotary wing. In contrast, helicopters provide both forward propulsion and lift through the rotary wing. 
     While rotor speed control at advanced ratios substantially below 1 is straightforward with conventional rotor controls, as the aircraft speed increases, the retreating blades are increasingly exposed to the relative airspeed of the vehicle fuselage. At an advance ratio of 1, the tip speed, the relative velocity of the retreating rotor blade at its extreme end is effectively stationary. That is, for example, the fuselage is traveling forward at a velocity, into the air, while the retreating blade tip is rotating in the opposite direction, at the same speed relative to the aircraft fuselage. Accordingly, the blade tip is effectively stationary. At advance ratios greater than 1, the relative airspeed of the fuselage is such that the trailing edge of the retreating blade is actually exposed to airflow in a reverse direction, that is, from trailing edge toward leading edge. 
     It would be an advantage to provide some additional, even a relatively small value, of controllable power in order to maintain desired rotor speed at all times. For example, increased lift, typically comes from a combination of collective pitch or the blade angle of attack and a control plane angle of attack. The direction for increasing lift is a function of collective pitch for a rotor trimmed to 0 flapping motion begins to reverse at an advance ratio of 1. The amount of flapping per degree of control angle of attack becomes extremely sensitive above advance ratios of 1. Thus, control may be particularly sensitive at advance ratios that are substantially higher than 1, if the rotor is to be kept in autorotation. 
     There is a substantial contribution to drag resulting when a rotor is driven in conventional autorotating mode. Therefore, it would be an advantage at higher speeds, particularly where a fixed wing portion of an aircraft may be more effective as a lift mechanism, to still maintain the rotor blades in autorotation at a sufficient speed or angular velocity to maintain their stiffness due to centrifugal forces. Therefore, it would be an advance in the art to provide some mechanism whereby additional autorotating force may be applied to a rotary wing, specifically to an autogyro or autorotating wing, without substantially increasing drag on the aircraft or on the wing, and without increasing fuel consumption, such as would result from powering the rotor or the like. 
     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 accordance with the foregoing, an apparatus and method are devised to provide a fully unloaded rotor operating in a high forward speed and advance ratios greater than 1, and particularly much greater than 1, with autorotating power. In certain embodiments, a rotor is unloaded by a fixed wing of an aircraft. Thus, the aircraft may have a rotary wing and a fixed wing. 
     The fixed wing may be optimized for high speed cruise condition, and thus can be extremely efficient from an energy and lift point of view, with minimum drag. With the rotor fully unloaded, it is not easily possible to maintain conventional autorotation. In fact, with the blade operating completely flat, with no lift, then the angle of attack of the rotor disk, which requires air flow in order to maintain autorotation, would be unavailable to drive rotation. 
     Accordingly, a variable geometry device may be located near the tip of each blade. The device may be configured to exhibit substantially no additional drag over the baseline airfoil. That is, a matched set of flaps may be formed at or near the trailing edge, and near the tip of each blade. These flaps may be sized to close, forming a conventional trailing edge under circumstances under which the rotor blade is advancing into the airstream with the aircraft. 
     However, when the blade is on the retreating side of the mast, the flaps may be opened by suitable actuators, thus greatly increasing their drag, and providing force and resulting torque, rotating the blade in the desired direction of rotation. Once the blade has reached an advancing position in which the flow is once again positive, that is the flow is passing from the leading edge to the trailing edge, the flaps may be closed. The flaps thereby provide the typical, decreasing thickness of the blade or rotary wing, ending with a very narrow trailing edge. 
     In certain embodiments, the flap system may be operated to move both upper and lower flap portions simultaneously to a downward position. In such a configuration, the two flap portions act as a flap, such as in a conventional aircraft. The flap becomes a single large extension to the wing of an aircraft, thereby creating high drag with high lift, when needed. For example, moving the upper and lower portions of the flap in closed mode, and both in the same direction would change the camber of the airfoil. This may be used to transform the rotor into a smart rotor in which the camber control device may be used to reduce vibration, enhance performance in the flight regimes where the rotor is being operated in conventional autorotation, or the like. 
     In general, maximum fluid drag occurs when an open cavity is presented into an oncoming airstream. Less drag occurs when the frontal or convex aspect is presented to the incoming air, leaving the concave aspect in a trailing position. The minimum drag occurs when the leading edge is closed, and the trailing edge is closed, coming to a gradual close in order to improve boundary layer effects and minimize drag. 
    
    
     
       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 accordance with 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. 4  is a front elevation view, schematically rendered, of a rotor blade on a mast; 
         FIG. 5  is an end cross-sectional view of an airfoil such as may be used in the rotor system of  FIG. 4 ; 
         FIG. 6  is an end, cross-sectional elevation view of the airfoil of  FIG. 5 , illustrating the position of the flap system in an open position, and designating by dotted lines, the closed position thereof; 
         FIG. 7  is a top plan view, schematically rendered, of a rotor disk for a rotor system of  FIG. 4 , illustrating the direction of flight, the direction of rotation of the rotary wing, and the resulting positive and reversed flow regions experienced by the rotor; 
         FIG. 8  is an end, cross-sectional, elevation view of the airfoil of  FIGS. 4-7 , with the upper and lower flaps of the flap system both positioned to move in the same direction, and thus modify the camber of the airfoil; 
         FIG. 9  is an end, cross-sectional view of the airfoil of  FIG. 8 , having the flaps in the closed and camber-neutral position; 
         FIG. 10  is an end, cross-sectional, elevation view of the airfoil of  FIGS. 8-9  having the upper flap in the raised position and the lower flap in the lowered position, thus maximizing drag during reverse flow conditions, in order to provide additional auto rotating force; and 
         FIG. 11  is an end, cross-sectional, elevation view of the airfoil of  FIGS. 8-10 , showing the difference in axis of rotation and the resulting circumference of the radius of the trailing edge of each flap portion, in a camber-controlled condition with both flap portions deflecting downward. 
     
    
    
     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. 
     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. 
     Referring to  FIG. 1 , an aircraft  10  includes a fuselage  12  defining a cabin for carrying an operator, passengers, cargo, or the like. The fuselage  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 fuselage  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 fuselage 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 fuselage  12 . The rotor  40  may be selectively powered by one or more engines  48  housed in the fuselage  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 fairing  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 fairing  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 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 fuselage  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. 4 and 5 , while continuing to refer generally to  FIGS. 1-11 , a system for autorotation may include a rotor system  70  having a mast  72  about which, or with which, rotor blades  74  rotate. Each of the blades  74  may have a leading edge  75   a  and trailing edge  75   b , on an airfoil  76  or wing  76  portion. For example, a blade  74  may include various attachment mechanisms, individual blade pitch controls, and so forth. Ultimately, however, the autorotating of the blades  74  depends on the airfoil  76 , a portion of which operates as a windmill. 
     Moreover, the aircraft load  77  or weight  77  of the fuselage and cargo of an aircraft must be opposed by a lift force  78 , commonly simply referred to as lift  78 . Accordingly, the airfoil  76  operates according to the Bernoulli principles. Thus, in flight, the load  77  or weight  77  represented by an aircraft must be opposed by the lift  78  provided by the wings  76  or the rotor blades  74 . 
     In one embodiment of an aircraft in accordance with the invention, the aircraft may be augmented with fixed wings that provide lift in a more aerodynamic and efficient manner at high speeds, and particularly at high advance ratios. In such an embodiment, the rotor system  70  may actually be unloaded such that it does not provide any substantial lift. One benefit for the rotor system  70  not providing lift is that the drag that would have to be sustained in order to provide lift may be eliminated. Thus, the blades  74  may be turned to be effectively flat, and not supporting any of the load  77 . However, in order to maintain autorotation in such an embodiment, a drive system  80  is needed to maintain rotation. 
     In accordance with the invention, a drive system  80  may include a top flap  82  and a bottom flap  84 . The top flap and bottom flap may extend along as much of the length of the blades  74  as is necessary to provide sufficient area to provide the autorotating power to drive the blades  74  in autorotation. Thus, the top flap  82  and bottom flap  84  may be a matched set that selectively move between closed positions  86 ,  88 , respectively, in which the two flaps  82 ,  84  provide a suitable trailing edge  75   b  for the airfoil  76 . 
     Thus, in the closed position  86 , the top flap  82  is in contact with the bottom flap  84 , also in its closed position  88 . More correctly, these closed positions  86 , 88  may be thought of as the closed, camber-neutral positions. For example, each of the top flap  82  and the bottom flap  84  may conceivably be moved in the same direction, and come to a position of closure adding camber to the airfoil  76 . 
     Referring to  FIGS. 6 and 7 , while continuing to refer to  FIGS. 4-5  and to  FIGS. 1-11  generally, a drive system  80  on an airfoil  76  of a blade  74  of a rotor system  70  may encounter a reverse direction  90  of airflow. This is explained hereinabove with respect to retreating blades  74  at high advance ratios. 
     The direction  92  of flight of an aircraft may be thought of as moving into still air, relative velocity rendering it an airstream. Accordingly, the direction of flight  92  results in certain anomalies with respect to the shape of the airfoil  76 . In general, the direction of airflow  90  is reversed when the aircraft is traveling at a suitable speed in the direction of flight  92 , and the rotation of the blade  74  or blades  74  that are on the retreating side, moving in the reverse flow direction  90  are moving at a speed that is effectively less than the speed in the direction  92  of flight. In such an environment, a retreating blade experiences a reversed flow direction  90  in which incoming airflow passes from the trailing edge  75   b  toward the leading edge  75   a.    
     In flight, an autorotating aircraft has a rotor system  70  that rotates the blade  74  about an axis  94  of rotation. Typically, for the descriptions herein, the direction  96  of rotation of the blades  74  will be counterclockwise as illustrated. In this configuration, a region  98  of reversed flow exists, for any trailing edge  75   b  or any portion thereof along the length of the blade  74  at which the net forward speed in the direction of flight  92  exceeds the retreating speed of that portion of the plane in the reverse direction  90 . Thus, the shape of the reverse flow region  98  varies somewhat with the speed of the aircraft in the direction of flight  92 , and the net linear speed of a trailing edge  75   b  of a blade  74  opposite thereto. 
     The rotor disk  100  represents the sweep  100  of the blade  74  of a rotor system  70 . Accordingly, the region  102  of positive flow or positive leading edge flow may be thought of as the conventional experience of a blade  74 , or the airfoil  76  of such a blade  74 , advancing into the airstream by its leading edge first. 
     However, the region  98  represents that portion where the trailing edge of a retreating blade  74  is first to encounter the airstream, and the flow is in the reverse direction  90 . Thus, actuators operating to pivot the flaps  82 ,  84  about their respective pivots  104 , may alter the effective drag near the trailing edge  75   b  of the airfoil  76 . 
     Referring to  FIGS. 8-11 , in one embodiment, the flaps  82 ,  84  may act in concert both moving in the same direction. For example, in  FIG. 8 , both the flaps  82 ,  84  are positioned in a downward orientation. Accordingly, the camber provides more lift, and more drag. Likewise, the upper  82  and the lower flap  84  may be positioned both in the upper position in order to provide an opposite effect. 
     Referring to  FIGS. 9-11 , in general, an upper surface  106  and lower surface  108  of an airfoil  76  may define a chord  110  or effective airfoil length  110 . Similarly, the relative thickness  112  of the airfoil  74  will have a direct effect on drag. The chord  110  and thickness  112  may be designed according to suitable practice as engineered in the art. 
     Meanwhile, implementation of the pivots  104  in which the flaps  82 ,  84  pivot results in respective radii  114 ,  116  for the two flaps  82 ,  84 . Thus, the trailing edge  75   b  of each of the flaps  82 ,  84  need not necessarily align. In the camber position of  FIG. 11 , a slight mismatch in the contact area would result in the trailing edge  75   b  on the lower flap  84  representing the actual final trailing edge of the airfoil  76 . Where the camber is reversed, then the trailing edge  75   b  of the upper flap  82  would represent the trailing edge of the airfoil  76 . 
     Meanwhile, the configuration of  FIG. 9  represents a camber-neutral configuration of the airfoil  76 . The configuration of  FIG. 10  represents the maximum drag position in the reverse flow direction  90 . Here, the trailing edges  75   b  of the upper  82  and lower flap  84  are opened in opposite directions in order to maximize drag on the retreating blade  74  in reverse flow. Thus, maximum power is imparted to the blades  74  of the rotor system  70  by the reverse air flow  90 . 
     Some of the benefits of the system are that upon rotation into the region  102  of positive edge flow, the flaps  82 ,  84  may be moved to the camber neutral position of  FIG. 9 , and thus minimize drag of the airfoil  76  advancing into the air in the direction  92  of flight. Thus, during a rotation, the operation of the upper flap  82  and lower flap  84  may be optimized in order to provide the appropriate drag for autorotative loading on retreating blades  74  power, and the appropriate, minimized drag on the advancing blades  74  of the rotor system  70 . 
     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.

Technology Category: 7