Patent Publication Number: US-2021163128-A1

Title: Electronic control of blade pitch on a tiltrotor

Description:
CROSS-REFERENCE TO RELATED APPLICATIONS 
     This application claims priority to U.S. Provisional Patent Application No. 62/942,709, filed on 2 Dec. 2019 by Frank Bradley Stamps, et al., and titled “Electronic Control of Blade Pitch on a Tiltrotor,” the disclosure of which is incorporated by reference in its entirety. 
    
    
     STATEMENT REGARDING FEDERALLY SPONSORED RESEARCH OR DEVELOPMENT 
     Not applicable. 
     BACKGROUND 
     The control systems for helicopters and tiltrotor aircraft are complex electrical and/or mechanical systems. The control systems respond to the pilot&#39;s input, but also must accommodate forces acting upon rotor assemblies and generally outside the control of the pilot. Mechanical control systems typically include a swashplate arrangement which consists of an inboard non-rotating plate and an outboard rotating plate. Typically, the swashplate assembly has the ability to translate axially along a mast and/or tilt relative to the mast. The rotating plate is free to rotate relative to the non-rotating portion. Pilot inputs alter the axial position of the swashplate assembly through the collective control and the tilt of the swashplate assembly through the cyclic control. 
     In the prior art, the rotating plate is typically connected mechanically to each individual rotor blade. For example, in one type of control system, pitch links are connected to pitch horns carried by the rotor blade, thus allowing the rotating plate to control the blade angle of each rotor blade. However, it is necessary to include in control systems a subsystem which reduces the degree of flapping as much as possible. In the prior art, there are two basic approaches: one is to utilize a delta-3 hinge; the other is to utilize offset pitch horns. In tiltrotor aircraft, it is especially important to counteract the detrimental effects of flapping, especially because the aircraft is capable of very high speed travel, particularly in the airplane mode of flight. A feedback control system is disclosed in U.S. Pat. No. 6,616,095, titled “Coupled Aircraft Rotor System,” and the entire disclosure of which is incorporated herein. 
       FIGS. 1, 2, and 3  illustrate a tiltrotor aircraft in flight.  FIG. 1  depicts aircraft  11  in an airplane mode of flight operation, tiltrotor having a fuselage  13  and wings  15 ,  17  for providing lift during forward flight. Rotors  19 ,  21  are composed of a plurality of rotor blades and rotated by engines carried in pylons  23 ,  25  rotatable relative to wings  15 ,  17 , and the thrust from rotors  19 ,  21  is directed aft in airplane mode.  FIG. 3  depicts aircraft  11  in a helicopter mode of flight, with thrust from rotors  19 ,  21  directed downward. Rotation of pylons  23 ,  25  allows for switching between the aircraft and helicopter modes of flight.  FIG. 2  depicts aircraft  11  in a transition mode. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
         FIGS. 1 through 3  depict a tiltrotor aircraft according to this disclosure. 
         FIG. 4  is an oblique view of a portion of a rotor and control system according to this disclosure and configured for use on the aircraft of  FIG. 1 . 
         FIG. 5  is an oblique view of a portion of another embodiment of a rotor and control system according to this disclosure and configured for use on the aircraft of  FIG. 1 . 
         FIG. 6  is a schematic top view of a portion of the aircraft of  FIG. 1  showing example sensor locations. 
         FIG. 7  is a top view of a portion of another embodiment of a rotor and control system according to this disclosure and configured for use on the aircraft of  FIG. 1 . 
     
    
    
     DETAILED DESCRIPTION 
     In this disclosure, reference may be made to the spatial relationships between various components and to the spatial orientation of various aspects of components as the devices are depicted in the attached drawings. However, as will be recognized by those skilled in the art after a complete reading of this disclosure, the devices, members, apparatuses, etc. described herein may be positioned in any desired orientation. Thus, the use of terms such as “above,” “below,” “upper,” “lower,” or other like terms to describe a spatial relationship between various components or to describe the spatial orientation of aspects of such components should be understood to describe a relative relationship between the components or a spatial orientation of aspects of such components, respectively, as the device described herein may be oriented in any desired direction. 
     This disclosure divulges control systems for rotors of a tiltrotor aircraft and methods for controlling the rotors. Each system allows for a control-system configuration providing for electronic control of blade pitch, thereby allowing for mitigation or prevention of undesirable rotor behaviors that may occur during flight in airplane mode. 
       FIGS. 4 and 5  illustrate example embodiments of a rotor and control system assembly according to this disclosure. Assemblies  111  and  211  each have a rotor  113 , comprising a central hub  115  and blades  117  coupled to hub  115  for rotation with a mast  119  about a mast axis  121 . Each blade  117  is coupled to hub  115  by arms  123 , which may be discrete members or portions of a yoke. In the embodiment shown, hub  115  is configured to gimbal relative to mast  119 , allowing for out-of-plane flapping motions of blades  117 , in which blades  117  on opposite sides of hub  115  move in opposite directions. To allow for out-of-plane coning motion of each blade  117 , in which blades  117  on opposite sides of hub  115  move in the same direction, each blade  117  is coupled to the associated arm  123  with a coning hinge  125  having a coning-hinge axis  127 . Alternatively, each arm  123  may comprise a flexure portion to allow for coning motion about a virtual coning hinge. Each blade  117  is also configured for rotation relative to hub  115  about a pitch axis  129 . 
     A swashplate  131  comprises a nonrotating plate  133  and a rotating plate  135 , with plate  135  being free to rotate in one degree of freedom relative to plate  133  while motions of plate  133  are transferred to plate  135 . A pitch rod  137  couples rotating plate  135  to a pitch horn  139  of each blade  117 , allowing for swashplate  131  to control the pitch of blades  117 . Swashplate  131  is configured for translation along and relative to mast  119  for collective control of blade pitch, and swashplate  131  is configured for tilting relative to mast  119  for cyclic control of blade pitch. Clevises  141  on nonrotating plate  133  are coupled with control rods  143  to an actuation system configured for causing motion of swashplate  131  in response to inputs by a pilot and/or flight-control system (FCS)  145 . Rods  147  may optionally be used to transfer flapping motion of hub  115  to sensors for measuring and communicating to FCS  145  the magnitude and direction of flapping motions, or flapping motion may be measured and communicated to FCS  145  by sensors within hub  115 . While shown as being behind axis  129  relative to the direction of rotation of rotor  113 , pitch horns  139  may alternatively be located forward of axis  129 . 
     To allow for electronic control of the pitch angle of blades  117 , assembly  111  of  FIG. 4  and assembly  211  of  FIG. 5  show embodiments incorporating linear actuators  149 , which may be electric, hydraulic, or electrohydraulic. In assembly  111 , an actuator  149  is installed within each pitch rod  137 , allowing for FCS  145  to operate actuator  149  to control the length of each rod  137 . In assembly  211 , an actuator  149  is installed within each control rod  143 , allowing for FCS  145  to operate actuator  149  to alter the length of each rod  143 . An advantage to having actuators in the non-rotating portion is that no hydraulic or electrical slip ring is needed to cross the gap to the rotating system. In alternative embodiments, each linear actuator  149  may be located at one end of the associated pitch rod  137  or control rod  143  for altering the axial position of pitch rod  137  relative to swashplate  131  or pitch horn  139  or control rod  143  relative to swashplate  131 . 
     In assemblies  111 ,  211 , the location of the connection of each pitch horn  139  to the associated pitch rod  137  provides for minimized delta-0 (pitch-cone coupling), as the connection is located on or near coning hinge  125 . However, in other embodiments, pitch horns  139  may be located elsewhere, have a greater length, or have a configuration that places the pitch-rod connection at a location producing unfavorable delta-0 or both unfavorable delta-3 (pitch-flap coupling) and unfavorable delta-0. Using the systems and methods disclosed herein allows for compensation for these unfavorable couplings. 
     Electronic control of delta-3 can be used on tiltrotors, such as aircraft  11 , having 3 or more blades and may allow for placement of pitch horns in locations otherwise unavailable for use due to unfavorable delta-3 and/or delta-0 coupling. One example, as shown in  FIGS. 4 and 5 , is positioning the connections of pitch horns  139  to pitch rods  137  at or near coning hinge  125 , thereby producing a favorable delta-0 for flight in airplane mode. This might be necessary with rotors having four or more blades  117 . However, this may produce unfavorable delta-3, but FCS  145  can use electronic control of delta-3 in response to parameters of the rotor, drive system, or aircraft that are measured during flight. For example, the start of whirl flutter may be detected using flapping as one of the symptoms and then operating actuators  149  to damp the whirl flutter. However, flapping is not always present, and other parameters may be more useful, such as looking for effects in the pylon. 
     In the rotating system, flapping, shear forces on the hub, and mast bending may be measured by sensors and communicated to FCS  145  for use in determining when and how much delta-3 should be added to the system. Flapping may be measured using, for example:
         a. hub-mounted sensors that measure tilting of the hub relative to the mast due to gimbaling; and/or   b. rod-actuated sensors measuring tilting of the hub, as illustrated in the figures; and/or   c. accelerometers mounted on the hub, yoke, or blades.
 
Hub shears may be measured using, for example, hub-mounted accelerometers, whereas mast bending may be measured using, for example, mast-mounted strain gauges.
       

     In the nonrotating system, pylon bending and wing bending may be measured by sensors and communicated to FCS  145  for use in determining when and how much delta-3 should be added to the system. For example, pylon bending, such as bending in the transmission case, may be measured using, for example, accelerometers and/or strain gauges in various locations. Likewise, wing bending may also be measured using, for example, wing-mounted accelerometers and/or strain gauges. 
     Electronic control of delta-0 can be used on tiltrotors having 3 or more blades and may allow for placement of pitch horns in locations otherwise unavailable for use due to unfavorable delta-3 or delta-0 coupling. Though the connections of pitch horns  139  to pitch rods  137  shown in  FIGS. 4 and 5  is at or near coning hinge  125 , the number of blades or other criteria may require a pitch horn configuration that produces both unfavorable delta-3 and unfavorable delta-0. FCS  145  may be configured to use electronic control of both delta-3 and delta-0 in response to parameters of the rotor, drive system, or aircraft that are measured during flight. To provide for proper control of delta-0, coning motion of the blades may be measured using, for example, angle sensors and/or accelerometers. In addition, torque sensors can be used to measure torque at the engine, mast, transmission, and/or interconnect driveline, which extends between the pylons and couples the rotors for continued operation using one engine, allowing for measuring a differential between the rotors. The control system can be used to alleviate these loads in the interconnect driveline by adjusting the collective and correcting errors in the flight path by adjusting collective and/or cyclic. 
       FIG. 6  schematically illustrates example sensor locations on one side of aircraft  11  for measuring parameters described above, though it may not be necessary to have sensors in all locations for every application. Sensor locations would typically be duplicated on the other side of aircraft  11 , and each sensor is in wired or wireless communication with FCS  145  for transmitting and receiving signals therebetween. Sensors  151  measure parameters related to an associated blade of rotor  21 , and sensors  153  and  155  measure tilting of the hub and forces acting on the hub. Sensor  157  measures parameters related to the mast, and sensor  159  measures parameters related to transmission  161 . Sensor  163  measures parameters related to engine  165 , and sensor  167  measures parameters related to interconnect driveline  169 . Sensor  171  measures parameters related to wing  17 , and sensor  173  measures parameters related to pylon  25 . 
     A significant advantage to using the control system according to this disclosure is that tilting happens at 1/rev, which is much faster than a pilot can apply corrective inputs. Instead, the system sums changes made by the system with pilot inputs, and the relatively small changes made by the system occur while still achieving the goal of the pilot input. 
     In order to provide for robust redundancy and fault tolerance, the system preferably includes modes for voting, fault detection, and fail-safe operation. In voting and fault detection, multiple sensors are preferably used to measure each parameter, and signals associated with a specific parameter are compared by FCS  145  to determine whether a signal is an outlier and should be disregarded. For example, multiple sensors would be used to measure tilting of the hub relative to the mast, and any sensor communicating a signal that substantially differs from the other sensors will likely be removed from calculations and may be ignored until service due to a detected fault. Alternatively, or as required for certain parameters, signals from sensors measuring various parameters can be compared by FCS  145  to determine if the indicated values correspond between the arrays of sensors. In fail-safe mode, FCS  145  will command the actuators to move to a selected position for achieving compromise values of delta-0 and/or delta-3, allowing for return to base or for immediate landing. FCS  145  may preferably limit aircraft airspeed to a selected maximum (e.g., 200 kts) and automatically slow the aircraft if airspeed exceeds the allowed maximum. 
     An advantage to locating actuators  149  on pitch rods  137  within the rotating system, as in  FIG. 4 , is that the system can control for delta-0 and delta-3, as described above, and provide higher harmonic control, such as for vibration control. 
     One phenomenon that can be encountered with certain configurations of rotors is “chugging,” which can occur when a positive delta-0 coupling causes increased pitch with increased coning. The system of this disclosure can be used to provide electronic suppression of chugging by translating the swashplate in collective motion. 
     Though  FIGS. 4 and 5  illustrate gimbaled rotors, the system of this disclosure may also be used with articulated rotors, rigid rotors, and rotors with discrete flap hinges. 
     For example,  FIG. 7  illustrates a rigid rotor  311  comprising hub  313 , grips  315 , and blades  317 . Blades  317  are retained within grips  315  with pins  319  to form a blade assembly  321 , such that the root end of each blade  317  is free to rotate only about pitch axis  323 , though blades  317  may be configured to allow for flapping or lead/lag motions through deformation of blades  317 . Hub  313  is coupled to mast  325  for rotation therewith about a mast axis. In the embodiment shown, hub  313  comprises a plurality of arms  327  extending radially, one arm  327  provided for each blade assembly  321 . 
     For control of the pitch of blades  317 , rotor  311  can be configured for use with a swashplate control system and linear actuators, like that described above for assemblies  111  and  211 . In addition, electronic control of blade pitch according to this disclosure may also be used with a swashplate-less rotor design, in which rotary actuators or similar devices are actuated to control blade pitch. In the embodiment of  FIG. 6 , each arm  327  of hub  313  houses a rotary actuator  329  for causing rotation of the associated blade assembly  321  about the corresponding pitch axis  323 . Each actuator  329  is in electronic communication with FCS  331 , which provides for control of actuators  329  in response to pilot inputs or automated inputs. Sensors, such as those described above and located as described above, are used to measure parameters of rotor  311  and are in electronic communication with FCS  331 . FCS  331  controls the pitch of blades  317  in the same manner described above to minimize or eliminate unwanted behaviors of rotor  311 . 
     At least one embodiment is disclosed, and variations, combinations, and/or modifications of the embodiment(s) and/or features of the embodiment(s) made by a person having ordinary skill in the art are within the scope of this disclosure. Alternative embodiments that result from combining, integrating, and/or omitting features of the embodiment(s) are also within the scope of this disclosure. Where numerical ranges or limitations are expressly stated, such express ranges or limitations should be understood to include iterative ranges or limitations of like magnitude falling within the expressly stated ranges or limitations (e.g., from about 1 to about 10 includes, 2, 3, 4, etc.; greater than 0.10 includes 0.11, 0.12, 0.13, etc.). For example, whenever a numerical range with a lower limit, R l , and an upper limit, R u , is disclosed, any number falling within the range is specifically disclosed. In particular, the following numbers within the range are specifically disclosed: R=R l +k*(R u −R l ), wherein k is a variable ranging from 1 percent to 100 percent with a 1 percent increment, i.e., k is 1 percent, 2 percent, 3 percent, 4 percent, 5 percent, . . . 50 percent, 51 percent, 52 percent, . . . , 95 percent, 96 percent, 95 percent, 98 percent, 99 percent, or 100 percent. Moreover, any numerical range defined by two R numbers as defined in the above is also specifically disclosed. 
     Use of the term “optionally” with respect to any element of a claim means that the element is required, or alternatively, the element is not required, both alternatives being within the scope of the claim. Use of broader terms such as comprises, includes, and having should be understood to provide support for narrower terms such as consisting of, consisting essentially of, and comprised substantially of. Accordingly, the scope of protection is not limited by the description set out above but is defined by the claims that follow, that scope including all equivalents of the subject matter of the claims. Each and every claim is incorporated as further disclosure into the specification and the claims are embodiment(s) of the present invention. Also, the phrases “at least one of A, B, and C” and “A and/or B and/or C” should each be interpreted to include only A, only B, only C, or any combination of A, B, and C.