Browse Topic: Pitch
This paper proposes a nonlinear and robust State-Dependent Riccati Equation (SDRE) combined with H∞ control architecture for brake- by-wire systems, specifically designed to handle severe tire-road friction variations and μ-split scenarios. The primary objective is to maximize deceleration capabilities while rigorously maintaining yaw stability, trajectory tracking, and passenger comfort through jerk limitation. Situated within the domain of active safety, this research addresses robustness against real-world uncertainties by utilizing a high-fidelity 14-degree-of-freedom vehicle model that accounts for longitudinal, lateral, and yaw dynamics, suspension-induced pitch and roll effects, and nonlinear tire behavior with explicit load transfer. To ensure near-optimal slip tracking under variable surface conditions, the system employs online friction estimation via Extended and Unscented Kalman Filters (EKF/UKF) fusing wheel and IMU data to adaptively adjust slip targets. The control strategy is bifurcated: the SDRE component manages dominant nonlinearities through state-dependent gains to prevent wheel lock-up, while the H∞ component provides robust disturbance rejection against parametric uncertainties such as mass variations and sensor noise. Control efforts are distributed via a Quadratic Programming (QP) torque allocator featuring anti-windup mechanisms and explicit saturation handling to compensate for lateral drift during μ-split braking. Validation is conducted through a Model- in-the-Loop (MIL) to Software-in-the-Loop (SIL) pipeline using scenarios including wet surfaces and panic braking. Simulation results demonstrate enhanced yaw stability and controlled deceleration profiles compared to conventional baselines, ensuring computational feasibility for automotive Electronic Control Units (ECUs).
A multi-objective optimization of a rotor blade airfoil is presented using compressible unsteady Reynolds-averaged Navier-Stokes simulations directly within the optimization loop. The baseline SC1095 airfoil is optimized using NSGA-II with two objectives: pre-stall aerodynamic efficiency representing hover performance, and lift hysteresis loop area representing dynamic stall severity. The optimized airfoil exhibits increased maximum thickness with an aft-shifted crest and substantially higher camber. Static polars show improved lift-to-drag ratio at $Ma = 0.5$ and $0.6$. Hover performance is essentially unchanged relative to the baseline. In forward flight, a progressive power penalty is incurred above $\mu = 0.2$, attributed to higher profile drag at advancing blade Mach numbers. Dynamic stall simulations show an 80% reduction in peak drag and a 50% reduction in peak pitching moment excursion relative to the SC1095, demonstrating the effectiveness of the optimization for retreating blade conditions.
A new approach that enables the synthesis of fully coupled system dynamics is described in this paper. The approach facilitates collaboration between solver developers by explicitly avoiding inter-code coupling and instead uses a generic interface that enables inputs to be set and flags outputs available to other solvers. The assembly of a fully coupled linearized system matrix is obtained entirely from the existence of coupling maps, and does not rely on user intervention. Tiltrotor whirl flutter in cruise conditions is systematically investigated by careful examination of results obtained through various combinations of domain synthesis and obtained from the Hermes coupling framework. Investigated solver domains include nonlinear rotor dynamics, nonlinear aerodynamics, and linear structural dynamics. Utilized software modules include RCAS; Project Chrono, a purpose-built lifting line aerodynamics solver; and FuselageSolver for linear structural dynamics. Aeroelastic predictions are synthesized and verified for a pitch/plunge airfoil and a low-speed wing. Rotor aeroelastic effects are verified by coupling rotor dynamics with rotor aerodynamics. Structure-to-structure results are synthesized by coupling both RCAS and Project Chrono rotor dynamics models to a pylon structure modeled in FuselageSolver. Finally, full whirl flutter predictions are synthesized by coupling Project Chrono, lifting line aerodynamics, and FuselageSolver, as well as an RCAS model without the wing/pylon structure to FuselageSolver. Results indicate that the new synthesis approach is viable and accurate.
Shrouded rotor configurations provide aerodynamic efficiency benefits that are valuable for aerial vehicles requiring high endurance in hovering flight. The effect of the shroud on the loads produced by a vehicle in response to cyclic pitch are not well known in literature. This paper aims to understand the magnitude and phasing of the steady hub loads produced by shrouded coaxial hingeless rotors in hover due to cyclic pitch. A shrouded coaxial rotor vehicle weighing 0.934 kg and with a rotor radius 0.138 m is used as the test platform. The rotors are operated at a tip Reynolds number of 70,000 and a tip Mach number of 0.18. First, the origins of steady hub loads due to cyclic pitch on a hingeless rotor are described using theory. To quantify the effect of the shroud, different configurations of rotors were tested by giving cyclic pitch to only one rotor. For the open hingeless rotor, cyclic variation in drag force at blade sections contributed to 55% of the steady lateral force. In response to cyclic pitch input, the ducted configurations produced 35% larger lateral forces and 60% (coaxial rotors) to 100% (single rotor) larger lateral hub moments than those produced by the unshrouded configurations. The presence of the shroud also changed the phase angle between the forces and moments. While the lateral force and moment acted in similar directions for the open configurations, the angle between them increased by more than 60° when enclosed in a shroud.
Wake measurements were performed on a 2-m diameter rotor in forward flight at an advance ratio of 0.2 undergoing sinusoidal collective and cyclic inputs using 2D-3C phase-resolved PIV. Input frequencies of 0.05/rev, 0.1/rev, 0.2/rev, and 0.4/rev were tested. The goal of this study was to characterize the time-varying rotor wake and extract Pitt-Peters dynamic inflow model parameters for the lateral cyclic inflow state. In both input cases, the effects of the pitch inputs manifested as modulation of the local upwash and downwash of the trailing tip vortices near the tip path plane. It was found that additional azimuthal measurements are necessary to improve the extracted value of the steady Pitt-Peters term. However, the extracted mass term was within 12% of the Pitt-Peters value, demonstrating the ability of the presented analysis to resolve the dynamics of the wake with a limited number of azimuthal measurements.
Full state feedback offers theoretically guaranteed multi-axis stability, making it superior to conventional PID controllers. There is however one drawback, a full state controller has a mathematical difficulty if the B matrix is not square and thus not invertible. This is the case for helicopters with 6 degrees of freedom and 4 inceptors. Variations of linear quadratic regulators are a work around, however complexity dramatically increases. Best would be a direct solution to the original problem. This is the breakthrough result of this paper. This paper documents an approach which removes the analysis roadblock by partitioning the 6 x 6 system "A" matrix into two groups of 4 x 4 matrices. The 4x4 matrices are individually stabilized with full state gain matrices. One matrix is designated “Driver Matrix” which provides actuator commands. The other matrix is designated "Reference Matrix" which provides references. The two matrices are coupled together by requiring that the driver matrix follow references generated by the reference matrix. With each matrix individually stabilized, the coupled combination is also stabilized. Computation of flight dynamics states (u, v, w, p, q, r) is shared between the matrices. Initial results are very encouraging, showing an originally sluggish, heavy lift helicopter having now concise decoupled responses to pitch and roll commands. Stability derivatives are recomputed during flight allowing coverage over the whole flight envelope. A handling qualities task has been defined to relocate a 40 ft standard seaborne container directed by a pilot in a ground control station. Cooper Harper ratings of this task have demonstrated favorable Level 1 handling qualities if use is made of an automated lateral repositioning command.
This study investigates the dynamics and associated vibratory loads of an underactuated swashplate-less rotor and its impact on the flight dynamics of a small-scale helicopter powered by this rotor via a combined experimental and computational approach. Unlike prescribing cyclic pitch using a swashplate, here the pitch is a response to the 1/rev cyclic rotor speed input. This is enabled on the current two-bladed rotor using a skewed lag hinge that utilizes the cyclic speed variation to produce lagging motion and subsequently pitching the blades in a cyclic fashion (ƍ4 coupling) for generating the pitch and roll control moments. One of the key dynamic characteristics that distinguishes this rotor from a conventional swashplate-controlled rotor is that the two blades have dissimilar pitch and flap responses leading to high fixed-frame vibratory loads. Results show that a large 1/rev vertical shear force is transferred to the fuselage resulting in half-peak-to-peak loads of +/−0.67g. The flap responses of the two blades being out of phase caused a net inertial force, which was the dominant contributor to the vibratory vertical shear force. Upon removal of the flap hinge, the vertical shear force was reduced by almost 85% while increasing the control moment authority by 50%. The inertial moment about the rotor axis generated by cyclic lagging nearly balances the moment due to the angular acceleration (Ω ) of the rotor, which significantly reduced the dynamic torque requirement from the motor.
This paper presents the investigation of experimental data belonging to main rotor loads during Never-Exceed-Speed demonstration of T625 Gökbey helicopter. Load data from the critical flight conditions in the VVNNNN envelope including cold-weather testing are collected. Maximum advancing tip Mach number demonstration, power-on and power-off flight conditions are investigated in terms of pitch link loads and blade loads. Blade loads including flapwise and chordwise bending moments, torsional moments and pitch link loads are examined to assess any divergence due to compressibility effects and the onset of stall. Load trends that are correlated with the tip Mach number are isolated from the effect of increasing dynamic pressure. Compressibility effects are observed to be the most dominant factor on the blade torsional moment and pitch link loads in advancing blade. The retreating blade stall phenomenon is apparent cases with a high advance ratio and mainly leads to dynamic stall cycles on the retreating blade, resulting in torsional moments and pitch link loads. Experimental results are compared with blade-resolved Unsteady Reynolds-Averaged Navier-Stokes simulations.
Achieving noise reduction in rotorcraft requires an analysis of various design parameters and flight conditions. However, high-fidelity methods are computationally expensive. To overcome this limitation, reduced order model (ROM)-based surrogate models have been applied to aerodynamics and aeroacoustics prediction. This study proposes a ROM-based surrogate model employing a variational autoencoder (VAE) to predict rotor aerodynamic loads and associated noise. Train and test datasets were generated using reformulated vortex particle method across a wide range of flight conditions. The proposed framework was applied to a single rotor, and its performance was evaluated qualitatively and quantitively in comparison with proper orthogonal decomposition (POD)-based surrogate model. The results show that VAE-based model consistently outperformed the POD model in noise prediction. These results demonstrate that the proposed framework enables accurate rotor noise prediction under various flight conditions and provides a promising approach for low-noise rotorcraft design and operational planning.
This study presents a comprehensive analysis of single-rotor failure tolerance for a classical octocopter configuration, examining both hover and forward flight at the best range speed. Using a state-of-the-art eVTOL comprehensive analysis to retrim the octocopter post-failure, the redistribution of rotor thrust, torque, and power following individual rotor failures was quantified, along with resulting aircraft-level power penalties. In hover, orthogonal rotors to the failed rotor provide primary lift compensation, the opposing rotor operates mostly unchanged, and the four opposite spinning rotors primarily provide pitch/roll moment compensation. This results in a total aircraft level power increase of approximately 10.4%, roughly half that of comparable hexacopters. In forward flight, at best range cruise speed, load redistributions were again calculated for various individual rotor failures. In the worst case, a maximum individual rotor torque increase of 62% and power increase of 108% was observed, while total aircraft power requirements increased between 7-12%. These results demonstrate the fault-tolerant capabilities of octocopters and provide practical guidance for propulsion system sizing, energy management, and failure-case assessment on classical octocopters.
An experimental investigation was conducted to explore the loads, acoustics, and tip vortex trajectories of coaxial counter-rotating (CCR) rotor with unequal upper and lower radii. The upper and lower rotor radii were tested both at the nominal radius of 1.108 m, and also with a lower rotor radius of 90% nominal radius, for a constant rotor speed of 1180 RPM and a constant inter-rotor spacing of z/R = 0.108. Rotors were torque balanced and tested for a range of upper rotor collective pitch from -2◦ to 10◦ . The power required for both CCR systems was within 0.9% for most trim conditions, and equal thrust was produced at upper rotor collectives of 6◦ and 8◦ (within 1.0%). At low loading conditions the unequal radii configuration produced more thrust for the same power due to a reduction in profile drag. The overall sound pressure level (OASPL) was lower for the CCR rotor with shortened lower rotor blades at all angles of elevation. Larger reductions in A-weighted OASPL(A) were observed, due to a larger contribution of broadband noise to the total OASPL(A).
Characterization of rotor–rotor wake interactions and their influence on flight dynamics is an important step toward advancing control system design and evaluating the performance of next-generation Mars multirotors. In this work, a Viscous Vortex Particle Method (VVPM) is utilized to generate rotor–rotor interference data for the Chopper Mars Helicopter platform, a large-scale hexacopter concept designed to be capable of carrying payload and pursuing independent science tasks. A reduced-order model compatible with finite state dynamic inflow is derived from the database. Interpolation strategies for continuous look-up are evaluated, with Gaussian Process Regression providing up to 20% improvement in prediction accuracy over linear interpolation of the interference data, although its scalability is limited by the large number of output channels. The interference model is implemented in HeliCAT, the flight dynamics analysis framework used for the Ingenuity Mars Helicopter, to assess the impact of rotor–rotor interference on trim, open-loop response and key stability derivatives. For Chopper, the first-order impact of rotor–rotor interaction on flight dynamics in hover is small, but with an observed increased pitch and roll damping. In forward flight, on-axis bare airframe responses exhibit attenuation at the lower frequencies due to interference, and increased sensitivity of the pitch response to vertical perturbations in cruise. Finally, a real-time closed-loop simulation is performed with and without the interference model to assess the impact on rotor power during a representative science mission flight trajectory on Mars, showing that total increase in mechanical shaft power due to multirotor wake interactions does not exceed 5% throughout the flight.
Tire noise reduction is important for improving ride comfort, especially in electric vehicle due to lack of engine noise and majority of the noise generated in-cabin is from tire-road interaction. Therefore, the tire tread pattern contribution is one of the important criteria for NVH performance apart from other structurally generated noise and vibration. In this work a GUI-based pitch sequence optimization tool is developed to support tire design engineers in generating acoustically optimized tread sequences. The tool operates in two modes: without constraints, where the pitch sequence is optimized freely to reduce tonal noise levels; and with constraints, where specific design rules are applied to preserve pattern consistency and manufacturability. The key point to be considered in this pitch sequence is that it should be reducing the tonal sound and equally spread i.e., the same pitch cannot be concentrated on one side which may lead to non-uniformity. So, the restriction is that the highest and lowest pitch types cannot occur adjacent to one another. This design rule helps in reducing undesirable pattern non-uniformity and improves both acoustic and structural performance. This tool helps in faster design iteration and integration with downstream development processes. This tool is also validated in current OE projects showing promising improvements in tire noise behavior while maintaining realistic design feasibility.
This paper explores a significant step forward, regarding the further detailed understanding of the Fenestron®. Since its patent in 1968 – for the Gazelle helicopter –, the shrouded tail rotor has been resized, inclined, modulated, etc. and has thus been continuously enhanced on different rotorcraft. Half a century after its invention, Airbus is once again exploring in more detail the magic of the Fenestron®, with the objective of optimizing it even further, for future helicopter applications. To grasp and observe properly some specific phenomena, a model (scaled to one third) capable of both unprecedented functions and modularities, was developed. The present paper will describe in detail the novel model and the related challenges and solutions. This model is capable of high rotor speed and dynamic pitch inputs, delivering power levels high enough to reach stall effects, while allowing the measurement of propulsive efficiency and to differentiate rotor vs fairing thrust. Furthermore, the model had to provide aerodynamic-shape modularity, both in the shroud, on the covers and on the tail gearbox supports. The first test campaign performed on this model allowed us to define design drivers and aerodynamic preferences. And since one idea often leads to another, another challenge is being addressed on this scaled mock-up: Reduction of the tail rotor noise level.
This paper demonstrates extraction of linear models from a state-space free wake model by applying analytical linearization, extending the research presented in (Ref. 1). Two distinct Linear Time Invariant (LTI) models are developed: the first is a high-order LTI model derived from the direct conversion of the analytical Linear Time Periodic (LTP) model, and the second is a reduced-order LTI model generated by first applying the Proper Orthogonal Decomposition (POD) model order reduction technique to the LTP model, followed by conversion. In both cases, the LTP-to-LTI conversion is achieved using harmonic decomposition. A substantial reduction in the number of wake states, from 15552 to 4050, is accomplished while maintaining a similar degree of accuracy. The time domain responses of step and doublet inputs for rotor collective and cyclic pitch are analyzed by comparing the GENHEL rotor model coupled with the LTI wake against the non-linear free wake model. Good agreement is observed in the rotor forces and hub moments. The paper investigates the feasibility of converting the linearized wake model into a linear parameter-varying inflow model with radial and azimuthal basis functions representative of a Peters-He inflow model. A least squares mapping technique is used to derive equivalent Peters-He inflow coefficients which are shown to accurately model the inflow distribution over the rotor disk and the time variation in induced velocities when compared to the non-linear free wake model.
This study investigates the effects of chord-to-radius ratio (c/R) and blade count on the aerodynamic and aeroacoustic performance of cyclorotors through experimental testing and a low-fidelity streamtube model. Cyclorotors with c/R ratios between 0.3 to 0.75 and blade counts ranging from 5 to 9 were tested across pitch amplitudes up to 51°. For a 5-bladed configuration, the pitch amplitude that maximizes the force-to-power coefficient (CF/CP) increases with c/R from approximately 32° at low c/R to around 51° at high c/R. However, the peak attainable CF/CP decreases with increasing c/R, indicating a trade-off between optimal pitch amplitude and aerodynamic efficiency. Increasing blade count enhances the generated force but reduces efficiency in all cases except for the lowest c/R configuration (0.3). Aeroacoustic analysis shows that tonal noise is primarily driven by pitch amplitude and intensifies with increasing c/R, while additional blades effectively mitigate it. In contrast, broadband noise is less sensitive to variations in pitch amplitude, c/R, and blade count. The streamtube model captures key aerodynamic trends, particularly at moderate pitch amplitudes. Scaling studies identify optimal configurations for a given disk loading, balancing power consumption and noise levels and highlighting key trade-offs critical to urban air mobility applications.
This paper explores the effect of addition of a horizontal tail on the longitudinal stability and performance of a Biplane Tailsitter Unmanned Aerial Vehicle (UAV). Biplane tailsitters a type of hybrid UAVs, often exhibits poor longitudinal stability during forward flight, necessitating continuous active control through application of differential motor thrust to maintain attitude. To address this challenge, this work proposes the integration of a horizontal tail on a quadrotor biplane tailsitter UAV, aiming to improve pitch stability and control authority during critical flight phases. Experimental flight data was utilized to determine the appropriate sizing of the elevator. A detailed flight dynamics model validated the effectiveness of the elevator control. The design was validated through outdoor flight testing, comparing the performance of tail-less and tail-attached configurations. The results demonstrate that the modified design results in a reduction control power requirement for pitch control on the motors, enhancing longitudinal stability, and improved ascent and descent rates. This study provides a systematic approach to improving the operational capabilities of tailsitter UAVs, contributing valuable insights for the design of future UAVs requiring enhanced maneuverability and efficiency in VTOL and forward flight modes.
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 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.
Active vibration damping by rotor torque modulation has been demonstrated for vibratory modes in the rotor disk plane. In this study, we introduce a simple, first-principles model, which includes kinematic coupling between lag movement and blade pitch, in order to extend damping authority to strut vibratory modes normal to the rotor disk plane. Using a medium-sized (12kg) quadcopter drone model, we demonstrate the capability to excite strut vibrations normal to the rotor disk plane, indicating control authority for vibration damping. For this vehicle model, a steady state strut deflection of over 12% is obtained using a 15% voltage perturbation, with under 2% rotor speed change. Redesign of the vehicle to have lower and/or co-located lag and structural frequencies increases the control authority of rotor torque actuation with pitch-lag coupling.
A wind tunnel investigation to characterise the aerodynamic performance and aeroelastic response of a tiltrotor blade set operating in propeller mode is presented. A custom blade set was instrumented with fully bridged axial strain gauges to monitor the flap bending and torsional strain at several radial locations. Propeller thrust and torque measurements were acquired using a custom six component Rotating Shaft Balance. Measurements of blade tip deflection were obtained via stereoscopic Digital Image Correlation. Testing was performed at a range of rotational frequencies, blade pitch angles and advance ratios to assess the blade aerodynamic performance and aeroelastic response in both attached and stalled operating conditions. Strain measurements were shown to identify stall and blade eigenmode frequencies, where flap bending bridges show a more reliable capture of stalled flow than torsional bridges. Furthermore, blade tip deflection measurements were shown to reduce with increased advance ratio and excessive values of advance ratio resulted in negative thrusting at the blade tip.
Aeroelastic stability prediction is critical to the successful design, development and flight testing of rotorcraft. As configurations reach higher speeds, new challenges in high Mach number unsteady aerodynamic modeling need to be addressed, especially for higher frequency aeroelastic modes with significant coupling. In this paper, Linear Unsteady aerodynamics and Leishman-Beddoes attached flow models are applied and compared to 2D CFD (airfoil) and 3D CFD/CSD (rotor) analysis for operating conditions of interest. The Leishman-Beddoes model demonstrates improved agreement with CFD data. In the 2D assessment, RCAS is used to model a representative airfoil undergoing prescribed pitch and heave oscillations. CFD results are presented to compare each model (Linear Unsteady and Leishman-Beddoes). In the 3D assessment, a full rotor CFD/CSD test case is evaluated for aeroelastic stability and compared to RCAS standalone analysis. The RCAS rotor structural model is coupled with the HELIOS CFD code and a swashplate cyclic pitch input is used to excite a lightly-damped rotor mode. The transient response based on RCASHELIOS is compared to the standalone RCAS internal aerodynamic result for both Linear Unsteady and Leishman- Beddoes unsteady aerodynamics. This study demonstrates that the Leishman-Beddoes model can produce similar stability results to the computationally expensive coupled CFD/CSD approach of RCAS-HELIOS, but at a lower computational cost, even for critical high-speed conditions.
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 study characterizes the dynamics of a novel lag-pitch-coupled underactuated rotor design that can be incorporated into rotary-wing unmanned aerial vehicles (UAVs) to provide pitch and roll control with effectiveness comparable to that of a conventional swashplate albeit with significantly lower mechanical complexity and weight. The concept integrates a single lag hinge tilted at a 45-degree angle located at the center of the rotor hub with independent flap hinges for each of the two blades. This idea relies on the ability to cyclically vary the angular velocity of the rotor in a 1/rev fashion via motor torque modulation, which induces a cyclic lag resulting in a cyclic pitch variation due to the tilted lag hinge (lag-pitch coupling) and causes the tip path plane (TPP) to tilt in a desired direction for pitch and roll control. To understand this concept, simulations using the Rotorcraft Comprehensive Analysis System (RCAS) were performed to capture the 1/rev response in lag, pitch, flap and moments under steady flight conditions. The predictions were validated using hover and wind tunnel experiments. The results show that the maximum lag and pitch response are delayed by 180° relative to the maximum input torque. Additionally, Blade-1 and Blade-2 have different flap responses, which lag the peak pitch response by phase delays of approximately 15° and 40°, respectively. Wind tunnel data indicated an inherent coupling between pitch and roll moments as a function of the advance ratio. Finally, a single main rotor helicopter UAV was built and flight-tested to demonstrate the effectiveness of the concept in both hover and high-speed forward flight.
This paper investigates the relationship between broadband noise behavior and helical wake structure in coaxial corotating rotors. Experimental measurements were conducted across variations in collective pitch (9.4°, 12.5°, and 15.0°) and rotor speeds (1500–4500 RPM). The inflow ratio (λ) was shown to govern the slope of broadband noise trends mapped in phase offset versus separation distance space, with experimental and theoretical λ values agreeing within 1%. Tip vortex core growth was estimated using the Ramasamy-Leishman model and normalized by the blade tip chord, reflecting the location of tip vortex formation. Across collective pitch variations, initial vortex core radii ranged between 7.5% and 9.1% and across rotor speeds, it ranged between 7.5% to 8.5% of the blade tip chord. When broadband noise trends became less coherent across phase offset angles, the corresponding vortex core radii were observed to approach or exceed 10% of the tip chord. At 4500 and 3500 RPM, vortex cores radii remained near or slightly above 10% at diffusion and depletion of broadband noise coherence. At 1500 RPM, initial vortex core radius is significantly beyond 10% of blade tip chord, leading to early diffusion, loss of organized wake structures, and broadband noise saturation across all phase offsets. If an inboard chord length were used instead for normalization, the specific normalized values would differ, but the overall trends in vortex diffusion and broadband noise coherence would remain similar. These results define a practical RPM threshold below which coherent wake structures and slope-based inflow prediction are no longer valid. Overall, the findings offer a framework for anticipating wake breakdown and broadband noise saturation, providing guidance for rotor design and operation to preserve low-noise performance.
Axial velocity measurements were performed in the wake of a hovering rotor with constant and sinusoidal cyclic pitch inputs ranging from 0.05/rev to 0.4/rev using a fixed, 2D-3C PIV system. Measurements were taken at 36 azimuths of the rotor with a constant cyclic input producing a pitching moment of CM = -0.00037. Using a Pitt-Peters definition, a longitudinal inflow state of λ1c = 0.0059 was extracted from the velocity measurements. A phase-resolved, undersampling approach was used to reconstruct the time history of the wake for the dynamic inputs. Simultaneous rotor hub loads measurements were used to obtain the frequency response of the longitudinal inflow state to pitching moment perturbations. The pitching moment perturbations ranged from ΔCM = 0.00027 at f=0.05/rev to 0.00046 at f=0.4/rev. The inflow perturbations ranged from Δλ1c = 0.0085 at f=0.1/rev to 0.0085 at f=0.4/rev. A first order transfer function was fit to the frequency response to compute Pitt-Peters dynamic inflow model parameters, but the low signal strength of λ1c produced high uncertainties in the model parameters.
Inspecting the interiors of tanks and ships for defects involves accessing confined and elevated spaces. This can be difficult and hazardous for a person. Ducted aerial vehicles that can hover close to the object of interest can achieve this in a safer and more efficient manner. Such a vehicle is desired to be compact, to have a high hover endurance and to be protected from impact. This paper describes a design concept comprising ducted coaxial counter-rotating rotors with a compact swashplate mechanism for cyclic pitch input to the lower rotor. An experimental setup was used to investigate the effect of the duct. A numerical Blade Element Momentum Theory model was developed and validated to inform rotor selection. A prototype was designed and built with a hover thrust of 9.17 N, outer diameter of 350 mm, and height 173 mm. The duct provided a thrust benefit of 32% for this configuration for a given power. The prototype achieved stable controlled flight in hover and in passing near walls and surrounding obstacles. Some oscillations in attitude were seen in flight. Future work includes improving the control algorithm of the vehicle.
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