Browse Topic: Yaw
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
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.
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
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
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
ABSTRACT
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
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
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
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
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
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.
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
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
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
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 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
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
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
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