Patent Publication Number: US-2022227488-A1

Title: Preventing Helicopter Loss of Tail Rotor Effectiveness

Description:
BACKGROUND 
     In a conventional helicopter, Loss of Tail rotor Effectiveness (LTE) is an unanticipated high yaw rate in the direction of the advancing main rotor bade. LTE can be initiated with aggressive pedal input or if the helicopter experiences a strong tail wind, which tends to weathervane the helicopter. Once underway, the high yaw rate may result in a continued uncommanded high yaw rate that does not subside of its own accord. Thus, if not prevented by the pilot, LTE can result in the loss of aircraft control. LTE is a low-speed aerodynamic flight characteristic of helicopters and is not caused by a mechanical failure. On helicopters manufactured in the United States, the main rotor typically rotates in a counterclockwise direction when viewed from above. Torque produced by the engine is applied to the main rotor causing the main rotor to turn in one direction while the helicopter fuselage reaction torque is to rotate in the opposite direction (i.e., clockwise, or nose-right direction). On some helicopters manufactured in Europe and Russia, the main rotor rotates clockwise when viewed from above, which causes the fuselage to rotate in the counterclockwise or nose-left direction. During normal operations, the helicopter&#39;s anti-torque system, or tail rotor, provides thrust to counteract the fuselage reaction torque, which provides directional control while hovering. The continued uncommanded high yaw rate of LTE occurs when the flow of air through a tail rotor is altered due to the helicopter&#39;s yaw rate. LTE is a well-known phenomenon of single main rotor helicopters. Federal Aviation Administration (FAA) Advisory Circular AC 90-95 directs the pilot to anticipate conditions conducive to LTE and not allow a yaw rate to build. 
     SUMMARY 
     Embodiments are directed to systems and methods for incorporating an algorithm in a helicopter flight control system to limit the pilot&#39;s, or the aircraft&#39;s ability to achieve yaw rates that would cause LTE. A specific yaw rate limit is set on the helicopter to avoid LTE based on vortex ring state avoidance at a tail rotor. The LTE yaw rate limit is specified for a helicopter flight control system to alert the pilot and/or prevent the pilot or the aircraft from exceeding the limit. 
     In an example embodiment, a flight control system for a helicopter comprises a pilot interface configured to receive a control input, at least one electronically controlled actuator, and a computing device configured to translate the control input to an actuator command, wherein the computing device is further configured to apply yaw rate limits to the actuator command to avoid loss of tail rotor effectiveness. The yaw rate limits are associated with a vortex ring state (VRS) envelope for a tail rotor of the helicopter. The electronically controlled actuator comprises a tail rotor actuator. The control input is a pedal input or a main rotor collective input. 
     The flight control system may further comprise a memory device configured to store the yaw rate limits as a table indicating whether a set of yaw rates are acceptable across a set of speeds, gross weights, or atmospheric conditions, such as altitude, temperature, and wind speed. 
     The flight control system may further comprise an indicator configured to notify the pilot when a command, such as a pedal command or a main rotor collective command, is predicted to result in a yaw rate that exceeds a yaw rate limit. The flight control system may further comprise an indicator configured to notify the pilot when a current helicopter yaw rate exceeds a yaw rate limit. In some embodiments, the systems and methods disclosed herein anticipate the LTE event based, for example, on a pedal position that will eventually cause the aircraft to exceed the yaw rate limit if the pedal position is held. This allows the flight control system to alert the pilot of the LTE risk before the yaw rate has approached the yaw rate limit. 
     In another example embodiment, a flight control system for a rotorcraft comprises a flight control computer configured to control operation of a tail rotor, wherein the tail rotor is configured to control yaw of the rotorcraft, and wherein the flight control computer is configured to limit a rotorcraft yaw rate based upon a current flight speed to prevent onset of a vortex ring state at the tail rotor. The flight control computer comprises a processor executing software instructions, wherein the software instructions are configured to perform the steps of determining a commanded yaw rate for the rotorcraft, determining a current forward speed and descent rate for the rotorcraft, referencing a database to determine if the commanded yaw rate or anticipated yaw rate from the command is acceptable at the current flight speed, and if the current yaw rate or anticipated yaw rate is not acceptable at the current flight speed, controlling a tail-rotor thrust to remain below a maximum yaw rate for the current flight speed. 
     The flight control computer may comprise a processor executing software instructions, wherein the software instructions are configured to perform the steps of receiving a yaw rate and a yaw command for the rotorcraft, determining a current speed for the rotorcraft, referencing a database to determine a maximum yaw rate acceptable at the current speed, and controlling a tail-rotor thrust to yaw the rotorcraft while remaining below the maximum yaw rate for the current speed. In other embodiments, instead of referencing a database to determine a maximum acceptable yaw rate, the flight control computer may perform a calculation based on current aircraft flight conditions and command inputs to determine if there is a risk of exceeding a yaw rate limit or a risk that the tail rotor may enter a VRS condition. 
     The flight control computer may comprise a processor executing software instructions, wherein the software instructions are configured to perform the steps of determining a commanded yaw rate for the rotorcraft, determining a current forward speed for the rotorcraft, referencing a database to determine if the commanded yaw rate or anticipated yaw rate from the command is acceptable at the current forward speed, and if the commanded yaw rate or anticipated yaw rate from the command is not acceptable at the current forward speed, controlling a tail-rotor thrust to keep the rotorcraft below a maximum yaw rate for the current forward speed. The database may comprise a lookup table of yaw rates for a range of forward speeds. 
     In other embodiments, the flight control computer may not require aircraft speed to manage the yaw rate. Instead, a single, conservative nose-right yaw rate limit may be set, wherein the selected yaw rate limit can be used for preventing LTE across all flight conditions. The flight control computer manages the aircraft to the remain under the designated yaw rate limit with appropriate tail rotor commands and restrictions, for example. 
     The flight control system may further comprise a pilot warning system configured to provide pilot alerts when a current rotorcraft yaw rate or anticipated yaw rate (e.g., based on pedal position, collective input, or flight condition) is within a predetermined range of a maximum safe yaw rate for a current aircraft flight speed. The predetermined range is a percentage of the maximum safe yaw rate or a predetermined value. 
     In an example process, a method of controlling a helicopter by a flight control system comprises monitoring a commanded yaw rate for the helicopter, monitoring a current flight speed for the helicopter, performing an assessment or data lookup to determine if a commanded pedal position or other control input will result in a yaw rate that is acceptable at the current speed, wherein the assessment or data lookup identifies the onset of a vortex ring state for the helicopter&#39;s tail rotor, if the commanded yaw rate is acceptable, turning the helicopter at the commanded yaw rate by a flight control computer, and if the commanded yaw rate is not acceptable, turning the helicopter at a maximum acceptable yaw rate for the current speed by the flight control computer. The yaw rate assessment may comprise a calculation based upon current aircraft conditions and commanded inputs instead of requiring a table look up or database query. 
     The method may further comprise notifying a pilot if a commanded yaw rate or anticipated yaw rate from the command presents a risk of loss of tail rotor effectiveness due to vortex ring state. 
     The method may further comprise notifying a pilot if a yaw rate has been limited below a commanded yaw rate value when the commanded yaw rate presents a risk of loss of tail rotor effectiveness due to vortex ring state. 
     The method may further comprise identifying when a current yaw rate or anticipated yaw rate of the helicopter is above a maximum acceptable value for the current speed, and automatically applying, by the flight control computer, a maximum yaw command opposite a current yaw direction. The method may further comprise determining when an updated yaw rate is below the maximum acceptable yaw rate or anticipated yaw rate for an updated current speed and reducing the yaw command opposite the current yaw direction. 
    
    
     
       BRIEF DESCRIPTION OF THE DRAWINGS 
       Having thus described the invention in general terms, reference will now be made to the accompanying drawings, which are not necessarily drawn to scale, and wherein: 
         FIG. 1  illustrates a rotorcraft according to some embodiments. 
         FIG. 2  illustrates a fly-by-wire flight control system for a rotorcraft according to some embodiments. 
         FIG. 3A  is a top view of a helicopter according to some embodiments. 
         FIG. 3B  is a top view depicting the helicopter of  FIG. 3A  in a right turn and showing an initial position superimposed by a subsequent position. 
         FIG. 4  is a graph illustrating an example LTE VRS envelope in terms of helicopter yaw rate. 
     
    
    
     While the system of the present application is susceptible to various modifications and alternative forms, specific embodiments thereof have been shown by way of example in the drawings and are herein described in detail. It should be understood, however, that the description herein of specific embodiments is not intended to limit the system to the particular forms disclosed, but on the contrary, the intention is to cover all modifications, equivalents, and alternatives falling within the spirit and scope of the present application as defined by the appended claims. 
     DETAILED DESCRIPTION 
     Illustrative embodiments of the system of the present application are described below. In the interest of clarity, not all features of an actual implementation are described in this specification. It will of course be appreciated that in the development of any such actual embodiment, numerous implementation-specific decisions must be made to achieve the developer&#39;s specific goals, such as compliance with system-related and business-related constraints, which will vary from one implementation to another. Moreover, it will be appreciated that such a development effort might be complex and time-consuming but would nevertheless be a routine undertaking for those of ordinary skill in the art having the benefit of this disclosure. 
     In the specification, 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 the present application, 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. 
       FIG. 1  illustrates a rotorcraft  101  according to some embodiments. The rotorcraft  101  has a main rotor system  102 , which includes a plurality of main rotor blades  103 . The pitch of each main rotor blade  103  may be controlled by a swashplate  104  in order to selectively control the attitude, altitude, and movement of the rotorcraft  101 . The swashplate  104  may be used to collectively and/or cyclically change the pitch of the main rotor blades  103 . The rotorcraft  101  also has an anti-torque system, which may include a tail rotor  105 , no-tail-rotor (NOTAR), or dual main rotor system. In rotorcraft with a tail rotor  105 , the pitch of each tail rotor blade  106  is collectively changed in order to vary thrust of the anti-torque system, providing directional control of the rotorcraft  101 . The pitch of the tail rotor blades  106  is changed by one or more tail rotor actuators. In some embodiments, a fly-by-wire (FBW) system  107  sends electrical signals to the tail rotor actuators or main rotor actuators to control flight of the rotorcraft. The FBW system assists pilots in stably flying rotorcraft  101  and to reduce workload on the pilots. 
     Power is supplied to the main rotor system  102  and the anti-torque system  105  by engines  108 . There may be one or more engines  108 , which may be controlled according to signals from the FBW system  107 . The output of the engine  108  is provided to a driveshaft  109 , which is mechanically and operatively coupled to the main rotor system  102  and the anti-torque system  105  through a main rotor transmission  110  and a tail rotor transmission  111 , respectively. 
     The rotorcraft  101  further includes a fuselage  112  and tail section  113 . The tail section  113  may have other flight control devices such as horizontal or vertical stabilizers, rudder, elevators, or other control or stabilizing surfaces that are used to control or stabilize flight of the rotorcraft  101 . The fuselage  112  includes a cockpit  114 , which includes displays, flight controls (e.g., cyclic, collective, pedals, etc.), and instruments. It should be appreciated that even though rotorcraft  101  is depicted as having certain illustrated features, the rotorcraft  101  may have a variety of implementation-specific configurations. For instance, in some embodiments, cockpit  114  is configured to accommodate a pilot or a pilot and co-pilot, as illustrated. It is also contemplated, however, that rotorcraft  101  may be operated remotely, in which case the cockpit  114  could be configured as a fully functioning cockpit to accommodate a pilot/co-pilot or may have limited functionality to provide controls for a back-up pilot with the primary piloting functions being performed remotely. In yet other contemplated embodiments, rotorcraft  101  could be configured as an unmanned vehicle, in which case the cockpit  114  could be eliminated entirely in order to save space and cost. 
     The FBW system  107  may provide different control characteristics or responses for cyclic, pedal, or collective control input in the different flight regimes. FBW system  107  may be implemented in one or more flight control computers (FCCs) disposed between pilot controls in cockpit  114  and flight control systems. The FCCs provide corrections to flight controls that assist in operating the rotorcraft more efficiently or that put the rotorcraft into a stable flight. The FBW systems in rotorcraft  101  may, for example, automatically adjust power output by the engine to match a collective control input, apply collective or power correction during a cyclic control input, provide automation of one or more flight control procedures, provide for default or suggested control positioning, or the like. 
     In some embodiments, the FBW system  107  includes an autopilot function that provides automatic flight control for rotorcraft  101 . The autopilot may be capable of controlling some or all aircraft flight parameters, such as maintaining speed, velocity, heading, or altitude, or managing operations, such as hovering and navigating. The autopilot can relieve a pilot of time-intensive tasks and reduce the pilot&#39;s workload. The autopilot system provides flight commands to a flight control computer. The flight control computer interprets the flight commands and, in response to the autopilot&#39;s commands, signals actuators that control the main rotor system  102  and the anti-torque system  105  as well as the angle of attack of main rotor blades  103  and tail rotor blades  106 . Such actuators may be mechanically, electrically, pneumatically, and/or hydraulically driven and may be controlled by electrical command signals from the flight control computer. 
       FIG. 2  illustrates a fly-by-wire flight control system  201  for a rotorcraft according to some embodiments. A pilot may manipulate one or more pilot flight controls  202  in order to control flight of the rotorcraft. The pilot flight controls  202  may include manual controls such as a cyclic stick  203 , a collective stick  204 , and pedals  205 . Inputs provided by the pilot to the pilot flight controls  202  may be transmitted mechanically and/or electronically (e.g., via the FBW flight control system) to flight control elements by the flight control system  201 . Flight control elements may represent devices operable to change the flight characteristics of the rotorcraft. Flight control elements on the rotorcraft may include mechanical and/or electrical systems operable to change the positions or angle of attack of the main rotor blades  103  and the tail rotor blades  106  or to change the power output of the engines  108 , for example. Flight control elements include systems such as the swashplate  104 , tail rotor actuator  206 , and systems operable to control the engines  108 . The flight control system  201  may adjust the flight control elements independently of the flight crew in order to stabilize the rotorcraft, reduce workload of the flight crew, and the like. The flight control system  201  includes engine control computers (ECCUs)  207 , flight control computers (FCCs)  208 , and aircraft sensors  209 , which collectively adjust the flight control elements. 
     The flight control system  201  has one or more FCCs  208 . In some embodiments, multiple FCCs  208  are provided for redundancy. One or more modules within the FCCs  208  may be partially or wholly embodied as software and/or hardware for performing any functionality described herein. For example, FCCs  208  may comprise a memory  210 , including a non-transitory medium for storing software  211 , and one or more processors  212  for executing instructions of software  211 . Memory  210  in some embodiments is a memory system that includes both transitory memory such as RAM and non-transitory memory such as, ROM, EEPROM, Flash-EEPROM, magnetic media including disk drives, and optical media. Memory  210  stores software  211  as machine readable instructions executable by processor  212 . 
     In embodiments where the flight control system  201  is an FBW flight control system, the FCCs  208  may analyze pilot inputs and dispatch corresponding commands to flight control elements such as the ECCUs  207 , the tail rotor actuator  206 , actuators for the swashplate  104 , or other components used to control operation of the rotorcraft. Further, the FCCs  208  are configured and receive input commands from the pilot controls  202  through sensors associated with each of the pilot flight controls. The input commands are received by measuring the positions of the pilot controls  202 . The FCCs  208  also control tactile cueing commands to the pilot controls  202  or display information in instruments on, for example, an instrument panel  213 . 
     In some embodiments, the instrument panel  213  includes a flight management system (FMS). The FMS may include a GPS, a mapping system, a navigation system, or other types of systems. A flight director system (“FD”)  214  may further include a display that can be used to provide information to the pilot. In some embodiments, the FD  214  may be configured to send signals representing flight control targets to the FCCs  208 , and may be configured to receive coupling logic signals from the FCCs  208 . The FD  214  and FCCs  208  may be able to send or receive other signals, such as sensor signals, target error signals, command signals, or other types of signals. In some embodiments, the FD  214  is communicatively coupled to the FCCs by a communications bus such as a serial bus, a CAN bus, or another type of wired or wireless communication system. In some embodiments, the FD  214  includes an autopilot function that provides automatic flight control for the aircraft. 
     The ECCUs  207  control the engines  108 . For example, the ECCUs  207  may vary the output power of the engines  108  to control the rotational speed of the main rotor blades or the tail rotor blades. The ECCUs  207  may control the output power of the engines  108  according to commands from the FCCs  208 , or may do so based on feedback such a measured revolutions per minute (RPM) of the main rotor blades. 
     The aircraft sensors  209  are in communication with the FCCs  208 . The aircraft sensors  209  may include sensors for measuring a variety of rotorcraft systems, flight parameters, environmental conditions, and the like. For example, the aircraft sensors  209  may include sensors for measuring yaw rate, ground speed, airspeed, altitude, attitude, position, orientation, temperature, airspeed, vertical speed, and the like. Other aircraft sensors  209  could include sensors relying upon data or signals originating external to the rotorcraft, such as a global positioning system (GPS) sensor, a VHF Omnidirectional Range sensor, Instrument Landing System (ILS), and the like. In some cases, some aircraft sensors  209  are part of or connected to the FD  214 . 
     The cyclic stick  204  may be connected to one or more cyclic position sensors, one or more cyclic detent sensors, and one or more cyclic actuators or cyclic trim motors. The cyclic position sensors measure the position of the cyclic stick  204 . In some embodiments, the cyclic stick  204  is a single control stick that moves along two axes and permits a pilot to control pitch, which is the vertical angle of the nose of the rotorcraft and roll, which is the side-to-side angle of the rotorcraft. The cyclic position sensors generate roll and pitch signals that are sent to the FCCs  208 , which controls the swashplate  104 , engines  108 , tail rotor  106 , or related flight control devices. 
     The collective stick  204  has one or more collective position sensors, one or more collective detent sensors, and one or more collective actuators or collective trim motors. The collective position sensors measure the position of a collective stick  204 . In some embodiments, the collective stick  204  is a single control stick that moves along a single axis or with a lever type action. A collective position sensor detects the position of the collective stick  204  and sends a collective position signal to the FCCs  208 , which controls engines  108 , swashplate actuators  104 , or related flight control devices according to a collective position signal to control the vertical movement of the rotorcraft. In some embodiments, the FCCs  208  may send a power command signal to the ECCUs  207  and a collective command signal to the main rotor  103  or swashplate actuators  104  so that the angle of attack of the main blades is raised or lowered collectively, and the engine power is set to provide the needed power to keep the main rotor RPM substantially constant. 
     Pedals  205  have one or more pedal sensors  227  that detects the position of the pedals  205  and sends a pedal position signal to the FCCs  208 , which controls the tail rotor blades  106  to cause the rotorcraft to yaw or rotate around a vertical axis. 
     Instrument panel  213  may provide warning, caution, and advisory indications  215  to the pilot. Warning and caution indications  213  may include, for example, text alerts on a display, lights (e.g., master caution and fire lights), an annunciator panel, aural warnings (e.g., bells, chimes, and tones), spoken warnings, tactile feedback (e.g., cyclic, collective, or pedal shaker), and other indicators. FCCs  208  may determine that the aircraft is approaching or in an unsafe condition and may trigger an appropriate warning or caution indicator  215  to notify the pilot of that condition. For example, FCCs  208  may monitor a database of certain aircraft performance limitations and thresholds  216 . When an aircraft status or flight condition reaches or approaches a limitation or threshold in database  216 , FCCs  208  can provide an indication to the pilot via warnings and cautions  215 . 
     LTE is a well-known phenomenon of single main rotor helicopters. A helicopter may not recover from an LTE condition if the pilot allows excess yaw rate to develop and does not counter it immediately. In a helicopter with a FBW flight control system, LTE can be mitigated by limiting the associated yaw rate. In prior FBW systems, there has been no criteria to set such yaw rate limitations to prevent LTE. The embodiments disclosed herein use recent LTE knowledge and incorporate an algorithm into the helicopter flight control system to limit the pilot&#39;s ability to achieve yaw rates that would cause LTE under various flight conditions and aircraft states. In some embodiments, the aircraft may be under computer control versus pilot control. In either case, the main strategy employed by the disclosed embodiments is to prevent the aircraft from getting to a yaw rate that initiates VRS. 
       FIG. 3A  is a top view of a helicopter  301  having a fuselage  302  and an empennage  303 . A main rotor having a plurality of rotor blades  304  is mounted on fuselage  302  by mast  305 . Main rotor blades  304  rotate in a counterclockwise direction around mast  305  as indicated by arrow  306 . Engine torque is applied to rotate the main rotor blades  304  generating an equal and opposite torque on fuselage  302  as indicated by arrow  307 . Torque  307  induces a right yaw force on helicopter  301 . An anti-torque system, such as tail rotor  308 , is mounted at the end of empennage  303 . Tail rotor  308  generates a tail rotor thrust  309  to counteract torque  307  so that helicopter  301  can maintain directional control. Tail rotor  308  creates a wake  310  that is blown to the left side of helicopter  301 . While  FIG. 3A  illustrates a helicopter having a main rotor that rotates in a counterclockwise direction, it will be understood that the same phenomena occur in helicopters having a main rotor that rotates in a clockwise direction, but the forces and torques are reversed compared to  FIG. 3A . 
     LTE may occur in helicopter  301  under several different conditions. In some cases, tailwinds can accelerate right yaw rates and make a right turn susceptible to LTE. For example, a tailwind  311  from the right rear quarter will have the tendency to “weathervane” helicopter  301  and create forces on empennage  303  that work to turn helicopter  301  into the wind. In other cases, headwinds may create main rotor disc vortex interference the reduces the effectiveness of tail rotor blades  308 . For example, a left quartering headwind  312  can direct the main rotor vortex onto the tail rotor, which reduces the angle of attack for tail rotor blades  308 , thereby reducing the anti-torque thrust  309  and creating right yaw. The pilot must maintain control of helicopter  301  and detect any uncommanded right yaw caused by tailwinds  311  headwinds  312 , or other external conditions. Pursuant to the FAA Advisory Circular, if a sudden uncommand right yaw is detected, the pilot should apply full left pedal and move the cyclic forward to increase speed until recovery is detected. 
       FIG. 3B  is a top view depicting helicopter  301  in a right turn and showing an initial position  321  superimposed by a subsequent position  322 . Main rotor blades  304  are not shown in  FIG. 3B  to simplify the drawing. In some cases, commanded right turns at high yaw rates may cause LTE. The pilot initiates the right turn by stepping on the right pedal thereby inducing right yaw so that helicopter fuselage  302  begins to rotate clockwise  323  around mast  305 . When the pilot starts the right turn, the initial thrust value  309   a  created by tail rotor blades  308  is reduced to a lower, or negative, thrust value  309   b . Low or negative tail rotor thrust, combined with torque  307  yaw the helicopter fuselage  302  to the right. The resulting yaw rate may drive tail rotor blades  308  into the tail rotor&#39;s wake  310   a , which can build-up during the turn. As a result, the tail rotor wake cannot get away from the tail rotor. Eventually, the tail rotor&#39;s wake  310   b  can accumulate with sufficient intensity that the tail rotor is placed into a vortex ring state (VRS) condition. Once the tail rotor enters the VRS condition, the tail rotor blades  308  are immersed in a powerful flow field induced by the tail rotor&#39;s accumulated wake. The induced flow field prevents the tail rotor blades from increasing their effective angle of attack, even if the pilot attempts to increase tail rotor thrust. Thus, the helicopter remains in a spin. The tail rotor enters an equilibrium condition wherein it is unable to increase its thrust to a level that would reverse the helicopter&#39;s right turn rate, and then helicopter  301  is in LTE and will continue to yaw right uncommanded. 
     This is a pilot-induced LTE that occurs when the commanded yaw rate is sufficient to create the VRS condition on the tail rotor. The ability for this to occur depends upon, at least, the helicopter&#39;s commanded yaw rate and the flight speed. The helicopter is at more risk of LTE due to the tail-rotor-VRS condition at hover or low speeds. Once the helicopter&#39;s flight speed increases, the potential for developing an accumulated tail rotor wake is reduced along with the VRS condition. 
       FIG. 4  is a graph  400  illustrating an example LTE VRS envelope in terms of helicopter yaw rate. Having identified the cause of sustained LTE as VRS that is induced by a high yaw rate, the yaw rates that induce VRS on the tail rotor can be predicted as shown in  FIG. 4 . The x-axis represents the tail-rotor&#39;s edgewise flight speed in feet per second, and the y-axis represents the helicopter&#39;s yaw rate in degrees per second. Curves  401  and  402  define the boundaries of the tail rotor&#39;s VRS LTE envelope  403 . When the tail rotor is operating within envelope  403 , its wake is prone to accumulate, and the tail rotor may enter VRS LTE and continue an uncommanded right yaw. Curves  401  and  402  represent the yaw rate limits for the onset of tail-rotor VRS. 
     Once the tail rotor is in the VRS condition, applying full left pedal may not produce sufficient additional thrust to overcome the uncommanded right yaw. Accordingly, the pilot should avoid conditions that are conducive to LTE; however, the pilot may unintentionally induce tail rotor VRS (i.e., commanding sufficient right yaw rate to enter envelope  403 ) while performing operations, such as a rapid right turn in hover. 
     In one embodiment, knowledge of the LTE envelope  403  is incorporated as an algorithm in the helicopter&#39;s flight control system, which in turn limits the pilot&#39;s ability to achieve yaw rates that would cause LTE under the helicopter&#39;s current flight conditions. For example, a specific yaw rate limit can be set in the helicopter&#39;s flight control computers to avoid LTE based on VRS avoidance. By incorporating the specified LTE yaw rate limits into an algorithm ( 211 ) within the helicopter&#39;s FBW control system, the FCCs can alert the pilot to a pending VRS LTE based upon the current helicopter yaw rate and the pilot&#39;s pedal input. If the algorithm anticipates that VRS LTE limits will be exceeded, the system can respond with alerts such as spoken, aural, visual, or tactile warnings. The helicopter&#39;s FBW system may also prevent the pilot from exceeding the limit by directly acting upon the pedal input to prevent VRS LTE. The FCCs may also apply aircraft limits to pilot-commanded control inputs so that FBW commands to the main rotor blade, tail rotor blade, and control surface actuators are filtered or limited, based on the current operating conditions, in order to prevent the development of yaw rates that may induce tail rotor VRS. 
     Referring to  FIG. 2 , in one embodiment, a limitations and thresholds database  216  may include a table comprising yaw rate limitation data representative of graph  400 . Aircraft sensors  209  may include a yaw rate sensor. Alternatively, aircraft sensors  209  may provide other aircraft data, such as heading, that allows FCC  208  to calculate a yaw rate. FCC  208  may continuously monitor the aircraft&#39;s current yaw rate and compare the yaw rate to yaw rate limitation data in database  216 . If the aircraft&#39;s yaw rate is approaching, is at, or is above an excluded yaw rate (e.g., relative to curves  401 ,  402  and envelope  403 ), then FCC  208  may provide a warning to the pilot via warning/caution indicators  215 . 
     The FCC  208  may also monitor pilot inputs, such as yaw inputs commanded via pedals  205 . FCC  208  may calculate an estimated yaw rate for the aircraft based upon the pedal  208  input commands. If it is anticipated that the aircraft will reach or exceed yaw rate limits in database  216 , then FCCs  208  may limit or reverse commands to tail rotor actuator  206  to prevent tail rotor  106  from reaching a VRS condition and thereby prevent the aircraft from experiencing LTE. FCC  208  may also monitor current or anticipated yaw rates initiated by other command inputs, such as a main rotor collective input, in addition to pedal commands. Additionally, FCC  208  may limit or reverse commands to other actuators, such as a main rotor actuator, to prevent tail rotor  106  from reaching a VRS condition. 
     In another embodiment, an FBW system may attempt to recover the aircraft from a VRS-induced LTE condition. For example,  FIG. 4  may define a region  404  that is within the VRS LTE envelope  403 . Region  404  may represent flight conditions for which the FBW system will attempt to recover from VRS LTE. When the FBW system determines that the aircraft is approaching or is within envelope  403 , the FCCs may limit or restrict actuator commands to prevent the onset of, or further entry into, the VRS condition. In one example, a recovery of tail rotor VRS may be attempted by maneuvering the helicopter with a pitch or roll or by descending or climbing. Altitude permitting, another approach would involve a command to reduce the collective pitch of the main rotor blades. The reduction of main rotor collective relieves the tail rotor of the need to produce anti-torque thrust, allowing the tail rotor to recover from the VRS condition. Once the aircraft is operating within the VRS LTE region  403 , the pilot should execute appropriate recovery procedures. If the pilot&#39;s recovery is not initiated or is not effective, the aircraft may move “deeper” into the VRS LTE region, such as experiencing increasing unintentional/uncommanded yaw rate, and the flight conditions may reach region  404 . At that time, the FBW system may automatically execute non-intuitive LTE recovery procedure, such as commanding full right pedal followed by an abrupt full left pedal input regardless of pilot input to the pedals. The latter maneuver can allow the tail rotor to distance itself from its accumulated vortex wake by executing (temporarily) an even higher yaw rate. Once the tail rotor has cleared its wake it is able to get positive thrust that can reverse the yaw direction for a recovery. The FBW system may notify the pilot when such automatic procedures are being executed, such as by illuminating a warning or caution light, or providing some other spoken, aural, or tactile warning. The FBW system may provide an input for the pilot to deselect or override the automatic recovery procedures, such as a switch, button, trigger, or paddle on the cyclic or collective. If automatically initiated, the FBW recovery procedures may be executed until the aircraft flight conditions exit VRS LTE envelope  403  and normal flight is restored. 
     In certain flight conditions the helicopter&#39;s flight control computer or pitot static system may have difficulty getting an accurate airspeed measurement, such as during operations at low speeds when the rotor outwash may create misleading readings. In some embodiments, the flight control computer may manage yaw rates without reference to the helicopter&#39;s speed. For example, the flight control computer may establish a single yaw rate limit without regard to flight speed. Alternatively, the flight control computer may establish a single yaw rate limit for a block of flight speeds, such as all speeds under a threshold speed. The threshold speed may be associated with a minimum level of confidence in the measured airspeed. Above the threshold airspeed, the flight control computer may either calculate yaw rate limits or use a database to determine the yaw rate limits. Alternatively, yaw rate limits may be assigned all airspeeds or for blocks of airspeeds. 
     LTE is a low speed phenomenon. In certain flight conditions, particularly at higher speeds, it is difficult for a helicopter to enter LTE because the tail rotor moves away from the wake and, therefore, the wake cannot accumulate to create a VRS condition. Accordingly, the flight control computer may establish a single yaw rate limit for a block of flight speeds, such as all speeds above a designated speed, wherein the designated speed is associated with a minimum risk of LTE. 
     In an example embodiment, a flight control system for a helicopter comprises a pilot interface configured to receive a control input, at least one electronically controlled actuator, at least one yaw rate sensor, and a computing device configured to translate the control input to an actuator command, wherein the computing device is further configured to apply limits to the actuator command to prevent excessive aircraft yaw rate to avoid loss of tail rotor effectiveness. The yaw rate limits are associated with a VRS envelope for a tail rotor of the helicopter. The electronically controlled actuator may be, for example, a tail rotor actuator or a main rotor actuator. The control input may be, for example, a pedal input or a main rotor collective input. 
     The flight control system may further comprise a memory device configured to store the yaw rate limits as a table indicating whether a set of yaw rates are acceptable across a set of flight speeds. 
     The computing device may be further configured to apply an algorithm to calculate yaw rate limits and to determine whether the control input will lead to yaw rates that are acceptable for avoiding LTE. 
     The flight control system may comprise an indicator configured to notify the pilot when a current helicopter yaw rate or a commanded helicopter yaw rate exceeds a yaw rate limit. 
     In another example embodiment, a flight control system for a rotorcraft comprises a flight control computer configured to control operation of a tail rotor, wherein the tail rotor is configured to control rotorcraft heading, and wherein the flight control computer is configured to limit a rotorcraft yaw rate to prevent loss of tail rotor effectiveness (LTE). The flight control computer may comprise a processor executing software instructions, and wherein the software instructions are configured to perform the steps of determining a commanded yaw rate for the rotorcraft, determining if the commanded yaw rate is acceptable for avoiding LTE, and if the current yaw rate is not acceptable, controlling a tail-rotor thrust to remain below a maximum yaw rate. 
     The step of determining if the commanded yaw rate is acceptable for avoiding LTE may comprise comparing the commanded yaw rate to a yaw rate limit applied across the rotorcraft&#39;s entire flight envelope. 
     The step of determining if the commanded yaw rate is acceptable for avoiding LTE may comprise applying an algorithm to determine a maximum acceptable yaw rate. The algorithm may calculate acceptable yaw rates based on one or more parameters selected from: tail rotor thrust, main rotor torque, altitude, gross weight, air density, and temperature. 
     The step of determining if the commanded yaw rate is acceptable for avoiding LTE may comprise referencing a database to determine if the commanded yaw rate is acceptable at the rotorcraft&#39;s current speed. The database may comprise a lookup table of yaw rates for a range of forward speeds, altitudes, gross weights, air densities, or temperatures. 
     The flight control system may comprise a pilot warning system configured to provide pilot alerts when a current rotorcraft yaw rate or commanded yaw rate is predicted to exceed the maximum safe yaw rate. 
     In a further example embodiment, a method of controlling a helicopter by a flight control system comprises monitoring a commanded yaw rate for the helicopter, performing a calculation to determine if the commanded yaw rate is acceptable for avoiding the onset of vortex ring state (VRS) on the tail rotor, if the commanded yaw rate is acceptable, turning the helicopter at the commanded yaw rate by a flight control computer, and if the commanded yaw rate is not acceptable, turning the helicopter at a maximum acceptable yaw rate by the flight control computer. 
     The method may further comprise identifying when a current yaw rate of the helicopter exceeds a maximum acceptable yaw rate; and automatically applying, by the flight control computer, a control strategy for recovering the tail rotor authority. 
     The method may further comprise determining that the helicopter tail rotor is experiencing loss of tail rotor effectiveness (LTE) due to the tail rotor being in a VRS condition; and providing a pilot with an option to initiate, by the flight control computer, an automatic control strategy to bring the tail rotor out of VRS and to recover control. 
     The method may further comprise determining when an updated yaw rate is below the maximum acceptable yaw rate; and restoring control of the tail rotor to the pilot. 
     The example embodiments described herein refer to a flight control computer, which is intended to be broadly interpreted to mean any processor-based system that actuates flight control surfaces, such as a vehicle management system (VMC) or other system that links the aircraft&#39;s computers and networks, manages the aircraft&#39;s avionics and flight controls, supports high-speed communications, and/or enables fly-by-wire and autonomous flight capabilities. 
     The foregoing has outlined rather broadly the features and technical advantages of the present invention in order that the detailed description of the invention that follows may be better understood. Additional features and advantages of the invention will be described hereinafter which form the subject of the claims of the invention. It should be appreciated that the conception and specific embodiment disclosed may be readily utilized as a basis for modifying or designing other structures for carrying out the same purposes of the present invention. It should also be realized that such equivalent constructions do not depart from the invention as set forth in the appended claims. The novel features which are believed to be characteristic of the invention, both as to its organization and method of operation, together with further objects and advantages will be better understood from the following description when considered in connection with the accompanying figures. It is to be expressly understood, however, that each of the figures is provided for the purpose of illustration and description only and is not intended as a definition of the limits of the present invention.