Browse Topic: Roll
In this work, a vision-based solution is developed to address the challenge of landing on a ship deck with precision and accuracy. For an autonomous landing, it is important to have a fast and accurate pose estimation system along with a reliable control strategy. This research uses fractal ArUCo markers instead of multiple separate markers to allow smooth pose estimation at different heights. Pose estimates are further improved using an Extended Kalman Filter, and a tracking algorithm then uses these estimates to guide the landing. A four degree-of-freedom (roll, pitch, heave and sway) simulator platform was built and used to validate the algorithm. The accuracy of the vision system is compared against that of a motion capture system. Real-world experiments were performed on different quadrotors to demonstrate tracking and landing on the platform with sway, roll, and pitch motions. The results show that the system is efficient and reliable in achieving safe and successful landings. The proposed landing system is concluded to be applicable for landings on the deck of the ship under sea-state 4.
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.
In the last years, new rotorcraft configurations have increased the attention among industries, through which the tiltrotor one due to its capability of combining both rotorcraft and aircraft advantages. However, there are situations where the vertical take-off mode could be enhanced in hard environmental and flight conditions. Therefore, to address this challenge, this work aims to develop a methodology to characterize a roll take-off model for a general tiltrotor configuration in such situations. By combining the integration of the equation of motion and geometrical assumptions, the runway distance is determined for an acceptable range of nacelle tilting angles. The process is developed by meeting the requirements defined by the regulations, combining the aircraft certification standards (CS23 and CS25) with the available tiltrotor certification basis from the FAA project #TC3419RC-R. Following the Nominal application, a sensitivity analysis is carried out, which studies the main effects on the results by varying one variable at a time in terms of weight, wing-loading, and disk-loading.
This paper presents a path planning concept based on the Manned-Unmanned Teaming (MUM-T) between the helicopter and a drone. The drone flies ahead of the helicopter to detect possible unexpected obstacles in the mission area and sends the data to the helicopter. The path of the helicopter is automatically replanned to avoid the meteorological and physical obstacles detected by the drone. The path planning is based on the Rapidly-exploring Random Tree* (RRT*) and the Bidirectional Rapidly-exploring Random Tree (BiRRT) algorithms. The reference trajectory is planned by means of the RRT* algorithm and the replanning is performed with the BiRRT. The node connection is realized with the Dubins curves, that force the path to comply with the prescribed limitations on the helicopter's roll angle and flight path angle. The Savitzky-Golay filter is used to smooth the trajectory achieving curvature continuity. A closed-loop simulation model containing the dynamics of the pilot is used to evaluate the feasibility of a candidate trajectory basing on the pilot's workload. The pilot's dynamics is modeled through the frequency-based multi-axis Hess model. A workload assessment method is proposed basing on the aggression factor, which depends on the time histories of the control inputs. The workload quantitative assessment is carried out by comparison with the aggression factor on a "simple" manoeuvre, the spiral.
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.
This paper addresses the urgent need to enhance rotorcraft safety and performance by developing a prediction methodology for the onset of the Vortex Ring State (VRS), and therefore verifying the VRS avoidance diagram. The objectives of this research are to assess the correlation between predictions generated by a comprehensive flight dynamics code and the latest and most accurate VRS boundary models, validate the VRS avoidance diagram across diverse descending flight conditions, and identify specific parameters indicating the rotor's entry into the VRS. The methodology involves a detailed investigation of 8 descent manoeuvres using a comprehensive flight dynamics code coupled with an advanced free vortex wake model. Results show that the pitch and roll oscillations and thrust fluctuations experienced by helicopters during the VRS are also observed in the model response to steep descent maneuvers. The findings confirm the reliability and applicability of the VRS avoidance diagram, providing valuable insights for improving rotorcraft design and flight control systems. Furthermore, this study presents novel findings regarding the identification of the VRS onset boundary through the observation of pitch and roll rate fluctuations, which contribute to the systematic detection and prevention of the VRS.
ABSTRACT
In this work, a contract-based reasoning approach is developed for obstacle avoidance in unmanned aerial vehicles (UAV's) under evolving subsystem performance. This approach is built on an assume-guarantee framework, where each subsystem (guidance, navigation, control and the environment) assumes a certain level of performance from other subsystems and in turn provides a guarantee of its own performance. The assume-guarantee construct then assures the performance of the overall system (in this case, safe obstacle avoidance). The implementation of the assume-guarantee framework is done through a set of contracts that are encoded into the guidance subsystem, in the form of a set of inequality constraints in the trajectory planner. The inequalities encode the relationships between subsystem performance and operational limits that ensure safe and robust operation as the performance of the control and navigation subsystems and environment evolve over time. The contract inequalities can be obtained analytically or numerically using an optimization based path planner and UAV simulation. The methodology is evaluated in the context of head-on obstacle avoidance, where the contracts are constructed in terms of (1) minimum obstacle detection range, (2) expected obstacle size, (3) maximum allowed cruise velocity, (4) maximum allowable thrust, roll and pitch angles, and (5) inner-loop tracking performance. Numerical and analytical generation of these contracts for this scenario is demonstrated. Finally, in-flight contract enforcement is illustrated for typical scenarios.
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.
With modern aerospace vehicle configurations, highly-coupled redundant flight control surfaces are becoming standard practice. For such vehicles, traditional System Identification (SID) methods may not accurately capture the individual contributions of effectors to the vehicle bare-airframe response. A Joint Input-Output (JIO) methodology was used to estimate the control power for each highly-correlated roll effector of the Bell V-280 hover configuration. The methodology was demonstrated using flight test data, where the identification results were compared to a high-fidelity hardware-in-the-loop simulation in the V-280 System Integration Lab.
This study investigates the interactional aerodynamics for laterally and longitudinally canted two rotor systems with a front rotor and an aft rotor aligned with the flow. The 5.5 ft, 3 bladed fixed pitched rotors are simulated using CFD at a targeted 5lb/ft2 disk loading and 30 kts. Simulations are performed using the commercial Navier Stokes solver AcuSolve with a detached eddy simulation (DES) model. In addition to an uncanted case, two laterally canted cases (10° advancing sides up and 10° advancing sides down) as well as two longitudinally canted cases (10° inward and 10° outward) are simulated. Aft rotor performance is compared to isolated rotors operating at the same RPM, speed and shaft tilt angle in order to quantify the effect of rotor-rotor aerodynamic interaction. For all configurations, the aft rotors experience a lift deficit at the front of the rotor disk which also results in a nose down pitching moment relative to an isolated rotor. The lift deficit for the uncanted rotor was around 15%. Lateral canting only slightly increases the lift deficit (to 16-17%) but also produces 28-38% change in roll moments. Change in nose-up pitching moments for the uncanted and laterally canted rotors were in the 55%-64% range. Longitudinal canting produces larger changes in the magnitude of the lift deficit and pitching moment, but has minimal effect on roll moments. In particular, canting inward results in a lift deficit as high as 21% and a 94% change in pitching moment. Canting outward, on the other hand, reduces the aft rotor lift deficit to 11% and the pitching moment change to 19%. The paper explains the changes in the flow field and the underlying physics for the different cases in detail.
This study focuses on vibration reduction for quadcopters and octocopters with elastic, 2-bladed, synchronized-RPM, variable-pitch rotors through the use of relative rotor phasing. The study defines phase modes such as a pitch phase mode with relative phasing between the front and aft rotors, a roll phase mode with relative phasing between the left and right rotors, and a differential phase mode with relative phasing between the clockwise and counterclockwise spinning rotors for both the quadcopter and the octocopter, as well as additional higher harmonic phase modes for the octocopter. Parametric studies on individual phase modes indicate that for the quadcopter in forward flight the pitch and roll phase modes can almost entirely eliminate the 2/rev vibratory forces (at the aircraft level), but the 2/rev vibratory moments cannot be minimized at the same time. By simultaneously using multiple phase modes a Pareto-front can be generated and a solution selected based on the relative emphasis on force or moment vibration reduction. For the octocopter it was observed that individual higher harmonic modes (specifically the 2c or 2s modes) could almost entirely eliminate both the 2/rev vibratory forces and moments, simultaneously. Compared to vibration levels in forward flight that might, on average, be expected if the rotors were randomly phased, reductions of 62% and 96% in a composite vibration index (equally weighting 2/rev vibratory forces and moments) were calculated for the quadcopter and octocopter, respectively, with appropriate rotor phasing.
Inflow modeling is necessary for accurate predictions of performance, aeromechanics, handling qualities analyses and flight simulation of single and/or multi-rotor configurations. There are complete inflow theories, i.e., finite state dynamic wake theory, for single rotor configurations which were shown to correlate well with the test data. However, inflow models of multi-rotor configurations such as CH-47 Chinook helicopter are still heavily dependent on empirical corrections. The physical effects behind adding correction factors are reasoned to wake interference between front and rear rotors. However, effect of the interference on steady-state conditions and mechanism of how it affects the transient response are left unanswered. This paper aims to answer physical reasoning behind the correction factors. For this goal, an industry standard, empirically correlated inflow model, viz., Boeing Helicopter Simulation CH-47 Chinook Inflow Modeling Method (BHSimIMM), is compared with Pressure Potential Superposition Inflow Model (PPSIM). In addition, a high fidelity Viscous Vortex Particle Method (VVPM) is included in the study for gaining further insight into rotor-to-rotor inflow interference. In this study, it is shown that body pitch rate response is greatly affected by uniform inflow interference in hover and forward flight. The body roll rate response characteristic can be improved by having fore-to-aft inflow coupling and uniform to fore-to-aft inflow interference in hover. It is also shown that interference on the front rotor quickly diminishes as speed increases while rear rotor is always under the influence of front rotor.
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.
This paper studies aerodynamic effects of ground on regressive lag mode damping during ground resonance. The experimental investigation is performed on a scaled helicopter model built to simulate the ground resonance scenario. The study involves both stationary as well as dynamic (oscillating) ground conditions. Experiments are conducted by measuring lead-lag damping at different rotor speeds (in the ground resonance regime) for different collective inputs. Results show that ground and its dynamics have a significant effect on regressive lead-lag mode damping during ground resonance. NOTATION e Hinge offset h Distance between rotor hub center and pitch/roll axis Ib Lead-lag moment of inertia for blade kG Thrust augmentation factor Mb Mass of the blade R Rotor radius Rcg Center of gravity of blade Ω Rotor anglular speed θ0 Collective input of blade pitch IGE In ground effect IMU Inertial Measurement Unit MPM Matrix Pencil Method OGE Out of ground effect RLM Regressive lead-lag mode damping during ground resonance.
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.
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