Browse Topic: Yaw
This study examines the ability of a large (1200 lb gross weight) hexacopter with collective pitch controlled rotors to tolerate single motor failure. The hexacopter is considered in various orientations, and the vehicle is trimmed with one motor inoperative (OMI). Unlike RPM-controlled hexacopters, which were trimmable but uncontrollable in hover, and were untrimmable in cruise with an aft-rotor failure; with pitch-control the hexacopter is controllable in hover as well as trimmable for failure of any rotor in cruise (including an aft rotor failure). The study examines how pitch controls, and thrust are redistributed amongst the operational rotors, post-failure, for the different hexacopter orientations. For each case, the maximum thrust and torque increases on any individual rotor, and the total power increase, post-failure is examined. It is found that the hardest to trim cases are those where the hub torque and the hub drag induced yaw moment of the failed rotor add, and fault compensation for these cases usually comes at a high cost in terms of torque penalty and power requirement (necessitating use of more powerful motors and batteries). The results from the study indicated that operating the pitch-controlled hexacopter in an edge-first configuration with a clockwise spinning rotor 1 is the best choice from a fault tolerance perspective.
Developed in the frame of the European Clean Sky 2 program, the RACER High Speed Helicopter Demonstrator of Airbus performed its maiden flight on April 25th, 2024. In the continuity of the previous high-speed demonstrator X3 (1st flight in 2010) the RACER is a 7/8t (15000 / 18000 lb) class compound helicopter powered by two SHE Aneto-1X engines, including a wing and two propellers. The tail rotor is removed as the two propellers control the yaw axis by differential thrust. At flight 07, with its initial default settings, it reached a true airspeed of 227 kts in level flight, exceeding its objective of 220 kts.
This paper investigates the use of multi-modal cueing through full-body haptic feedback to enhance pilot-vehicle system (PVS) performance, reduce mental workload (MWL), and increase situational awareness (SA) in both good and degraded visual environments (GVE/DVE). Piloted simulations were conducted using an H-60-like flight dynamics model in a virtual reality (VR) motion-based simulator, evaluating two ADS-33-like mission task elements (MTEs) – precision hover and slalom – under visual-only and combined visual and haptic feedback conditions in both GVE and DVE. The H-60 flight dynamics were augmented with a dynamic inversion (DI)- based stability augmentation system (SAS), implementing rate-command/attitude hold (RCAH) response type on the roll, pitch, and yaw axes and altitude hold response type on the vertical axis. The SAS was designed to achieve Level 1 handling qualities per ADS-33 standards. The full-body haptic cueing strategy leveraged an outer-loop DI control law, which provided vibrotactile feedback to cue desired roll, pitch, and yaw attitudes to the pilot. Roll cues were delivered via tactors mounted on the upper arms, pitch cues via tactors on the chest and back, and yaw cues via tactors on the calves. Eight test subjects participated in the piloted simulations, including three U.S. Navy test pilots and five subjects with different flying experiences. Results indicated that haptic feedback significantly improved hover performance, reducing MWL and enhancing SA, particularly in DVE. However, in the slalom task, predefined haptic guidance misaligned with pilots’ individual control strategies, leading to performance degradation. This finding highlights the need for pilot-specific adaptive haptic feedback to mitigate inconsistencies in dynamic maneuvering tasks.
This paper provides an overview on the contributing phenomena to unanticipated yaw described in the FAA Helicopter Flying Handbook. Trimmed aerodynamic - flight-mechanic - coupled simulations with a validated model of the BK117 C-2 capture the relevant interactions for weathervaning, main rotor-to-tail rotor interactions and vortex ring state effects at the tail rotor. An investigation of the impact of the main rotor downwash on the vortex ring state at the tail rotor in sideward flight and yaw turn is provided, concluding that the presence of the main rotor effectively inhibits the occurrence of a fully developed deep vortex ring state at the tail rotor. The consequent limited impact of the incipient tail rotor vortex ring state on the helicopter trim is estimated. Further, maneuver simulations of the BK117 C-2 are provided, describing the typical entry in unanticipated yaw turn and the exit to stop the yaw motion by means of pedal inputs of different magnitude and input speeds.
A piloted simulation experiment was conducted in the NASA Ames Vertical Motion Simulator to investigate the effects of bandwidth, phase delay, attitude quickness, and maximum achievable rate on yaw-axis handling qualities in hover and forward flight. Two different aircraft were tested, representative of advanced scout-class rotorcraft. Five target acquisition and tracking Mission Task Elements were used in the study. Two of the tasks were modified versions of tasks used to determine the ADS-33E target acquisition and tracking yaw attitude quickness boundaries. Two of the tasks were modified versions of attitude capture and hold and sum-of-sines tracking previously used to evaluate pitch and roll axis handling qualities. The final task was a forward flight target acquisition task developed for this study based on a ground attack or strafing maneuver. Eight Army pilots participated in the study and evaluated 60 yaw-axis configurations. The results of the study suggest that the current yaw-axis hover/low-speed and forward flight bandwidth and hover/low-speed attitude quickness requirements for target acquisition and tracking are too low. Updated boundaries for these requirements are recommended.
A quadrotor was modified by adding wings to the frame to directly compare the flight dynamics characteristics as well as the stability and control derivatives of the quadrotor and its biplane tailsitter variant. The on axis response of the quadrotor and a biplane tailsitter variant were measured through flight test and frequency domain system identification was used for non-parametric and parametric model identification. Identification of the full vehicle dynamics demonstrated that also identifying the motor torque and back-EMF constants from no-load measurements and the remaining motor parameters from a rotor-motor test stand provided the most accurate identified full vehicle model. The motor dynamics were shown to add a pole to the thrust-based responses (roll, pitch, and heave), while the torque based response (yaw) included a pole and a zero. This approach was then used to identify and compare the quadrotor dynamics, tailsitter dynamics, and the total impact of canting the motors. It was found that the presence of the wing added pitch damping to the dynamics and pitch stability became negative. The yaw axis saw an increase in yaw damping derivative, and a reduction in the yaw control derivative to the point where it became difficult to control the aircraft. By introducing cant, both the quadrotor and tailsitter saw large increases in the yaw control derivative. Further, the rotor thrust based moment generation due to cant resulted in the yaw response zero being canceled by the motor dynamics, resulting in a purely first order yaw response. Neither the wing nor cant produced any change in the lateral and heave axes.
ABSTRACT
Emerging vertical flight concepts being proffered for solutions to the Future Vertical Lift (FVL) mission set such as compound high speed rotorcraft can be designed with multiple, coupled control effectors thus creating redundant systems in one or two more axes to generate control forces and moments which allow for a range of trim states. In the FVL mission area future rotorcraft will be asked to fly into high threat environments where potential failure modes can be encountered due to enemy fire or mechanical failure causing reduction of the safe flight envelope. Fault detection creates options to increase the survivability of the crew and passengers allowing an emergency flight envelope to be proposed. One of the more serious potential failures due to enemy fire is a loss of yaw control. Faults in yaw control can be detected in a compound rotorcraft with a vectored thrust ducted propeller (VTDP) or similar anti-torque thruster. An online Kalman filter (KF) for a dimensional yaw moment coeff icient model will be used to estimate vehicle yaw coeff icients. Deviation from the nominal coefficients will be monitored based on the KF statistics in the case of both rudder and tail rotor failure at 60, 40, and 20 ft/s in forward flight. Both frozen zero rudder and ganged sector faults as well as failed tail rotor faults were successfully detected at all airspeeds except the failed tail rotor at 60 ft/s. For the yaw control faults considered, post fault excitation appears airspeed dependent. An online KF estimator for yaw control fault detection could successfully be integrated into the design of a compound rotorcraft with VTDP thereby increasing system safety.
Modern system identification techniques were used to identify a linear model based on a nonlinear simulation of a concept Urban Air Mobility quadcopter, and compared to a perturbation-based model. These models were used to develop feedback controllers for both variable-pitch and variable-RPM variants of the quadcopter, with the handling qualities requirements determining current requirements for the electric motors. To have sufficient stability margins and bandwidth, the motor time constant for the variable-RPM system must be no greater than 0.122s. Both variable- RPM and variable-pitch systems were limited by the yaw axis, which relies on differential motor torque for control. The introduction of rotor cant alleviated this problem for the variable-pitch vehicle, allowing a 47% reduction in motor weight, relative to the uncanted variable-pitch system.
The objective of this investigation is three-fold. First, to assess the flight dynamics of an electric Vertical Take-Off and Landing (eVTOL) concept aircraft with a propeller-driven rotor. Second, to develop a Stability and Control Augmentation System (SCAS) for this concept aircraft. Third, to verify the potential safety benefits of the concept aircraft by analyzing the autorotation performance following a total loss of power. The paper begins with a description of the simulation model, including a detailed discussion on the inflow model of the propellers that drive the main rotor. Next, the flight dynamics are assessed at hover and in forward flight. A SCAS based on Dynamic Inversion (DI) is developed to provide stability and desired response characteristics about the roll, pitch, yaw, and heave axes for speeds ranging from hover to 80 kts. Additionally, an RPM governor is implemented to hold the main rotor angular speed constant at its nominal value. Finally, simulations that make use of the SCAS are performed to analyze the autorotation performance following total loss of power.
Optimization-based control design techniques are applied to multicopters with variable-RPM rotors. The handling qualities and motor current requirements of a quadcopter, hexacopter, and octocopter with equal gross weights (1200 lb) and total disk areas (producing a 6 lb/ft2 disk loading) are compared to one another in hover. For axes that rely on the rotor thrust (all except yaw), the increased inertia of the larger rotors on the quadcopter increase the current requirement, relative to vehicles with fewer, smaller rotors. Both the quadcopter and hexacopter have maximum current margin requirements (relative to hover) during a step command in longitudinal velocity. In yaw, rotor inertia is irrelevant, as the reaction torque of the motor is the same whether the rotor is accelerating or overcoming drag. This, combined with the octocopter's greater inertia as well as the fact that it requires 30% less current to drive its motors in hover, results in the octocopter requiring the greatest current margin, relative to hover conditions. To meet handling qualities requirements, the total weight of the motors of the octocopter and hexacopter is comparable at 15% weight fraction, but the quadcopter’s motors are heavier, requiring 17% weight fraction. If the longitudinal and lateral axes were flown in ACAH mode, rather than TRC mode, the total motor weight of all configurations would be nearly identical, requiring about 14.6% weight fraction for motors (compared to 8-9% weight fraction from hover torque requirements).
Optimization-based control design techniques are applied to multicopters with variable-RPM rotors. The handling qualities and motor current requirements of a quadcopter, hexacopter, and octocopter with equal gross weights (1200 lb) and total disk areas (producing a 6 lb/ft2 disk loading) are compared to one another in hover. For axes that rely on the rotor thrust (all except yaw), the increased inertia of the larger rotors on the quadcopter increase the current requirement, relative to vehicles with fewer, smaller rotors. Both the quadcopter and hexacopter have maximum current margin requirements (relative to hover) during a step command in longitudinal velocity. In yaw, rotor inertia is irrelevant, as the reaction torque of the motor is the same whether the rotor is accelerating or overcoming drag. This, combined with the octocopter’s greater inertia as well as the fact that it requires 30% less current to drive its motors in hover, results in the octocopter requiring the greatest current margin, relative to hover conditions. To meet handling qualities requirements, the total weight of the motors of the octocopter and hexacopter is comparable at 15% weight fraction, but the quadcopter’s motors are heavier, requiring 17% weight fraction. If the longitudinal and lateral axes were flown in ACAH mode, rather than TRC mode, the total motor weight of all configurations would be nearly identical, requiring about 14.6% weight fraction for motors (compared to 8-9% weight fraction from hover torque requirements).
The paper investigates structural coupling problems in tiltrotor aircraft. A detailed tiltrotor model, representative of the Bell XV-15, has been built. The airframe model has been modified with a thinner wing to better reveal structural coupling proneness. A linearized FCS has been introduced to analyze the overall stability on an extended frequency band, ranging from the flight mechanics up to the aeroelastic modes. In addition to the FCS, biomechanical models of the pilot, acting on the power-lever and on the center stick, are included in feedback loop. Overall stability analyses demonstrate that the FCS improves handling qualities although several structural coupling mechanisms arise, in combination with the involuntary pilot's response, reducing flutter clearance. A modified version of the XV-15, using differential collective pitch for yaw control in airplane mode, has been also investigated. This configuration reduces costs and weights although the FCS destabilizes the antisymmetric wing chord mode at low speed flight, severely limiting the flight envelope. Means of prevention, based on notch filters, are implemented and discussed.
This paper describes the development of a compact and re-configurable rotary-wing micro air vehicle (MAV) that is capable of sustained hover and could potentially be launched from a 40 mm grenade launcher in the future. Launching the vehicle as a projectile up to the point of operation could significantly improve the mission range for these energy constrained platforms. The MAV design used coaxial rotors with foldable blades, a thrust-vectoring mechanism for pitch and roll control, and a strict constraint on the outer diameter, which was relaxed to 52 mm for this study. Yaw control was accomplished by using a specialized counter-rotating motor that is composed of two independently controlled motors. Passive unfolding of the coaxial rotor blades utilizing centrifugal force was demonstrated. The vehicle attitude was stabilized in hover using a closed-loop proportional-derivative controller implemented on a 1.7 gram custom autopilot. Through systematic trimming and tuning of the feedback gains, the vehicle was able to achieve stable hover. When the vehicle was subjected to large impulsive pitch and roll perturbations, the feedback controller was able to successfully reject the disturbance and return the vehicle to a stable hover within a second. In parallel, an analogue of the flying vehicle or a "dummy" was built and launched using a pneumatic canon to understand the dynamics of the vehicle during the projectile phase without risking the actual flying vehicle. The launch demonstrated that with the right center of gravity location, the present vehicle configuration could be stable during the projectile flight even without fins.
This paper presents aircraft concepts and designs which demonstrate that distributed electric propulsion can enable another paradigm in aircraft design: asymmetry. This attribute is leveraged upon to address operational issues relating to single motor failure. It is shown that the unique combination of minimum number of motors and a corresponding placement for which any one of the motors could fail, and full flight control in roll/pitch/yaw throughout VTOL and airplane modes can still be maintained, requires an asymmetric arrangement of six motors and their proprotors. This all-round redundancy is particularly important in applications where the aircraft, in the event of single motor failure during airplane mode cruise, needs to continue to be recoverable by VTOL mode landing in geometrically constrained environments (e.g. forested areas, small ships, urban locations etc.). In addition, the mechanical simplicity of the asymmetric arrangement enables the motors to be installed with a cant angle, thereby leveraging moment arm effects to enhance yaw control power. This is a much-needed enhancement for certain challenging operating conditions such as hovering or vertical landing in a crosswind. A small-scale prototype is designed, built and flight tested. Also presented is an example aircraft sizing of a larger scale variant to FAA Part 107 requirements, and it is shown that one area that can yield substantial improvement in loiter endurance is variable pitch proprotors.
An examination is conducted into the effects of increasing rotor diameter on the handling qualities of a quadcopter with fixed-pitch, variable-RPM rotors. Five aircraft are simulated, with rotors ranging from 1 to 8 feet in diameter. The flight characteristics of the aircraft are quantified using Froude-scaled handling qualities metrics. Several scaled ADS-33E-PRF handling qualities metrics are evaluated, including response to a collective controller, disturbance rejection, and bandwidth in roll, pitch, and yaw. It is concluded that aircraft performance is limited by disturbance rejection requirements in yaw as well as actuator saturation limitations that are present in other control channels, and a quadcopter with rotors over 2 feet in diameter will need greater installed power than what is currently estimated in order to meet handling qualities metrics without violating actuator constraints.
ABSTRACT This work investigates rotorcraft-pilot coupling phenomena in tiltrotors. A detailed tiltrotor model, representative of the Bell-Boeing XV-15, has been built. Biomechanical models of the pilot, acting on the power lever and on the centre stick, are included in feedback loop to define the Pilot-Vehicle System. Pilot-Assisted Oscillation phenomena are investigated on the overall conversion corridor using Nyquist's criterion. Pilot-in-the-loop analyses demonstrate that a critical parameter is detected in the vertical fins geometry. Due to an asymmetric flaperons deflection the wing's wake impacts on the vertical fins, producing a side force. The pulsating tail-side-force makes the fuselage to yaw and excites the asymmetric wing chord mode coupled with the lateral pilot's biomechanics, leading to a reduction, or even a loss, of stability. No unstable event is detected about the longitudinal direction. Conversely, a resonance between the pilot's biomechanics and the aircraft poorly damped symmetric wing bending mode is predicted about the vertical axis. The instability is found on the whole conversion corridor, although the source of excitation changes with reference to the nacelle angle. Means of prevention are implemented and discussed.
ABSTRACT Rotor cant is simulated on an SUI Endurance quadcopter. Two types of rotor cant, flapwise and torsional cant, are defined, and multirotor coordinates are used to define four aircraft-level modes of cant for each type. Collective flapwise cant causes an increase in collective control and power required, and a positive correlation exists between collective flapwise cant and pitch control. It also causes the longitudinal and lateral poles to retreat from the origin. Postive longitudinal flapwise and negative lateral torsional cant cause a reduction in nose-down attitude in forward flight, reducing drag and negative lift on the fuselage by 13% and 31% at 15 m/s, which reduces power required by 6% while increasing hover power by only 0.5%. Lateral flapwise cant and longitudinal torsional cant affect the roll attitude, though no power savings is available. Differential flapwise cant causes forward speed to impose a net rolling moment, which is compensated by roll control. Differential torsional cant is positively correlated with roll control. Both differential cant modes cause some poles to move toward the origin while others move away, but differential torsional cant can increase yaw authority by up to 325%.
ABSTRACT A quadrotor was assembled with commercial off the-shelf (COTS) components readily available on the market as a platform for future research at Penn State. As a first step in this research, a model of the quadrotor is identified from flight data. Given the largely decoupled dynamics at low speed, frequency sweeps in different channels are performed separately on the roll, pitch, yaw and heave axes. A frequency-domain approach is used to perform system identification. First, frequency responses of the aircraft output are extracted from frequency-sweep flight data. Next, state-space models are fit to the frequency response data. Overall the identified model matched flight data well in both the frequency and time domain. Dynamic Inversion (DI) and Explicit Model Following (EMF) with LQR disturbance rejection control laws are developed for both an inner attitude loop and outer velocity loop. The control laws were developed to meet similar requirements, and have similar performance and robustness.
ABSTRACT A feedback controller is designed and implemented for a regular hexacopter based on the AeroQuad Cyclone ARF kit. This controller is designed with an inner loop control law as a set of parallel PID controllers for aircraft altitude, pitch, roll, and yaw attitudes, as well as an outer loop for control over aircraft body velocities. Rotor failure is modeled in the dynamic simulation by setting the rotor force and moment output to be zero regardless of the commanded control input to that rotor, the feedback controller utilizes no knowledge of this fault during simulation. Various trajectories are commanded to examine the performance of the baseline feedback controller in the event of forward rotor failure, including hover, forward flight, and more complex maneuvers. The controller is demonstrated to recover the aircraft states after the transient effects of the rotor failure, as well as complete the defined state trajectory, demonstrating tolerance to single rotor failure.
ABSTRACT This paper presents a concept of a multi-copter that can be reconfigured between a quadcopter, hexacopter, octocopter and decacopter. The maximum useful weights of the octocopter, hexacopter and quadcopter, were 77%, 55% and 32% that of the decacopter. The controls for each of the configurations are identified and for the configurations with control redundancy, the power optimal controls are presented. A dynamic simulation model is implemented and used to compare the various configurations. Over a range of aircraft useful weights, it was observed that the decacopter required minimum power when the useful weight was greater than around 30% of its maximum, due to lower induced and profile power requirements of the lighter loaded slower spinning rotors. At lower useful weights, the hexacopter required less power due to its significantly lower empty weight. The octocopter and quadcopter did not emerge as low-power configurations of choice. Increasing the number of rotors increased the maximum hover endurance. At a useful weight corresponding to 53% of the decacopter's maximum useful weight, the decacopter's hover endurance time was 9% higher than that of the hexacopter. At low useful weights, the maximum pitch and yaw accelerations generated by the hexacopter were around 80-95% higher than the decacopter, but at higher useful weights the decacopter's accelerations were around 60% higher. All multi-copter configurations displayed a longitudinal and lateral phugoid mode in hover. Compared to the hexacopter, the phugoid modes for the decacopter had significantly higher components of surge rate compared to pitch rate, or sway rate compared to roll rate, while the modal damping ratios were 20-40% lower.
ABSTRACT This study examines the effect of rotor blade elastic deformations on a quadcopter in forward flight conditions. The blade equations are discretized using the Galerkin method and the blade periodic response is calculated using the harmonic balance method. Simulations are conducted on a 2 kg quadcopter with 12 inch diameter two-bladed rotors. The blade root vertical shear, flap bending moment and drag shear showed a strong 1/rev variation due to the azimuthal variation in aerodynamic loads. Elastic blade deformations did not affect the aerodynamic loads but the addition of in-phase 1/rev inertial loads resulted in net increases of 28%, 36% and 48% in the 1/rev blade root vertical shear, flap bending moment, and drag shear, respectively, at a forward flight speed of 10 m/s. The in-plane elastic deformations further introduced a 1/rev blade root radial shear due to the radial Coriolis force. At 10 m/s forward flight speed, accounting for elastic blade deformations resulted in a 7.6% reduction in the steady hub drag force, a 37% increase in steady hub roll moment, and a 15% increase in steady hub pitching moment that requires a 12% higher pitch RPM control input to trim the quadcopter. Increases of 103% in the 2/rev rotor hub in-plane loads (drag and side force), of 12% in the 2/rev hub vertical force, and of 25% in the 2/rev hub pitching and rolling moments were also observed. Since the forward rotor, lateral rotors, and the rear rotor operate at distinct rotational speeds, the aircraft is subject to vibratory loading at three different frequencies resulting in a beating phenomenon. The maximum amplitude of these oscillations equals the sum of the amplitudes of the three constituent vibrations, but reaches its maximum only once every twelve cycles. The 2/rev rotor hub vibratory moments are the dominant contributors to the aircraft pitch and roll vibrations, while the 2/rev rotor hub in-plane forces are the dominant contributors to the aircraft vibratory yaw moments. The blade phasing between the two lateral rotors has an effect on the magnitude of the total roll vibrations on the aircraft.
ABSTRACT This paper presents a turbulence model with accurate spatial correlations for helicopter flight simulation and handling-quality analysis. First, digital filters with longitudinal correlations of the von Karman turbulence are developed to generate discrete turbulence velocity components. Turbulence transverse correlations are considered by relating the filters in different positions with spatial correlations of the von Karman theory. Then, the distributions of both the related filters in front of helicopter and their velocity components in the longitudinal direction of airspeed, as well as turbulence models for helicopter aerodynamic surfaces are established. On this basis, a flight dynamics model coupled with the turbulence model is developed and validated against the flight test data. The contribution of each aerodynamic surface to the helicopter handling qualities is analyzed. Finally, the helicopter handling qualities in turbulent atmospheric environment are discussed. The results show that turbulence transverse correlations have important impact on the handling qualities of both the roll and heave motions, while there is a little impact on the handling quality of pitch motion and nearly no impact on that of yaw motion.
ABSTRACT Helicopter Sling Load is the most accurate form of aerial delivery in the military due to the ability to air land materiel in an exact location; however, some missions have a tendency for the payloads to become unstable due to both pilot-in-the-loop and aerodynamic effects. Past research demonstrated that allowing the container to rotate freely in yaw stabilizes pendulum motions. Other research utilized rigid fins affixed to the rear of the container. These methods work during tests; however, they become difficult to use in an operational environment. This paper discusses tests using a flexible fabric stabilizer that can be temporarily added to any payload. The flight tests were conducted Moffett Field, CA using the same payload as the previously mentioned research. Tests showed the flexible stabilizer provided an intermediate level of performance eliminating sling wind up and stabilizing pendulum motions out to the aircrafts’ power limit in exchange for very little operational overhead.
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