Browse Topic: Pitch
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.
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.
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.
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 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 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.
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.
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.
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.
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.
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.
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.
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 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.
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.
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.
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 presents the development and application of analytical linearization of a State-Space Free Vortex Wake Model. Previous work developed a state-space free wake model that could be numerically linearized via finite differences into a Linear Time Periodic (LTP) system, but the numerical linearization process was computationally expensive. An improved method is developed that uses exact analytical linearization of the Biot-Savart Law. The analytical method is found to speed up linearization computations by O(N), where N is the number of free wake nodes. A simple decoupled wake model is used to develop and test the method, where the wake system's inputs are prescribed blade bound circulations. The state space matrices computed by the analytical linearization method are verified to match those of the numerical linearization method exactly as perturbation sizes approach zero. The analytically linearized LTP model was converted into a Linear Time Invariant (LTI) model using Harmonic Decomposition. The on-blade induced velocities predicted by the LTI wake model are found to match the non-linear wake model in both frequency response and for small amplitude step inputs to the blade circulation. The linearized wake model is then coupled with a non-linear blade element simulation of a utility helicopter rotor (using the GenHel rotor model). Step responses to rotor collective and cyclic pitch inputs are compared between the rotor model coupled with the linearized wake and with the non-linear wake model. The rotor forces, hub moments, and induced velocities of the two models are found to have good agreement.
Design modifications to a 3lb variant of DEVCOM Army Research Laboratory's Common Research Configuration (CRC-3) are assessed using simulation tools. To identify areas for improvement, the baseline CRC-3 is analyzed in hover and forward flight, and contributors to overall power consumption are identified, with the rotor drag consuming the greatest amount of power, due to the high rotational speeds required to maintain thrust in the face of the freestream velocity. Potential areas for improvement are identified as: wing airfoil, rotor blade pitch, and rotor orientation. Changing the airfoil has little to no measurable effect on the overall power consumption. Increasing the blade pitch improves cruise performance considerably, but at the cost of hover efficiency, for an overall range improvement of up to 28%. Changing the rotor orientation improves rotor efficiency as well, without substantial cost to hover power consumption, increasing the range by 37% but will require a redesign of the CRC-3 supporting structure to avoid the blades striking the structure. Experimentally determined trim sweeps of pitch and power requirement are shown for the quadrotor, tailsitter, and three different wing incidence angles (10◦, 20◦, and 40◦), showing up to 28% reductions in minimum power.
ABSTRACT The present study considers two notional rotorcraft models: a conventional utility helicopter, representative of an H-60, and a wing-only compound utility rotorcraft, representative of an H-60 with with a wing similar to the X-49A wing. An Explicit Model Following (EMF) control scheme is designed to achieve stability and desired Rate Command / Attitude Hold (RCAH) response around the roll, pitch and yaw axes, while alleviating vibratory loads through both feed-forward and feedback compensation. The harmonic decomposition methodology is extended to enable optimization of primary flight control laws that mitigate vibratory loads. Specifically, Linear Time Periodic (LTP) systems representative of the periodic rotorcraft dynamics are approximated by Linear Time Invariant (LTI) models, which are then reduced and used in LQR design to constrain the harmonics of the vibratory loads. The LQR gains are incorporated in the EMF scheme for feedback compensation. One innovative approach is the addition of rotor state feedback to standard rigid body state feedback. A Pseudo Inverse (PI) strategy is incorporated into the EMF scheme for redundant control allocation. Finally, simulation results with and without load alleviation are compared and the impact of PI feed-forward and rotor state feedback compensation on handling qualities is assessed in terms of ADS-33E specifications.
This paper focuses on an experimental investigation of rotor loads during dynamic stall on a rotating pitching blade. In particular, the effect of rotor control parameters—rotor speed, collective pitch, and cyclic pitch—on the structural load dynamics of a rotor blade are analyzed in hover. The rotor platform used is the Mach-scaled, two-bladed Munich Experimental Rotor Investigation Testbed (MERIT) rotor at the Technical University of Munich (TUM). The dynamic stall cases selected vary in collective and cyclic pitch angles: 14°±6°, 14°±10°, and 20°±6°. Static and dynamic stall data are measured at three different rotor speeds: 900, 1200, and 1500 RPM with the highest corresponding tip Mach and Reynolds numbers of Matip = 0.41 and Retip = 1.2•106. Increasing pitch and rotor speed shows a considerable positive trend in the load overshoot, and hysteresis of the blade root moments of most cases. Cycle-to-cycle variations with bifurcation occur in some load graphs of light dynamic stall cases and are not restricted to reattachment regions.
Airfoil optimization for rotor blades is a critical endeavor aimed at enhancing aerodynamic performance and reducing noise. This paper employs a Kriging surrogate model combined with a multi-objective genetic algorithm to optimize thrust, power, and broadband noise. Three airfoil parameterization methods including ParFoil, PARSEC, and CST are compared when used to generate various airfoil shapes for the surrogate model and optimization process. We utilize low-fidelity aerodynamic tools such as XFOIL and blade element momentum theory for aerodynamics. In addition, acoustic modeling is conducted using Lee's wall pressure spectrum model alongside Amiet's trailing-edge noise model. The paper focuses on small-scale rotor configurations, specifically an ideally twisted rotor using the NACA 0012 airfoil and a modified XV-15 blade. Both blades are used as baseline models for hover optimization. The optimization of the ideally twisted rotor across various parameterization methods demonstrates a significant reduction in A-weighted overall sound pressure level by approximately 4.0 dBA. The primary contributors to this noise reduction are identified as a decrease in the chordwise pressure gradient and wall shear stress, which are key factors that contribute to the wall pressure spectrum and resulting trailing-edge noise. Similarly, the optimization of the modified XV-15 blade using the ParFoil method achieves a notable decrease in noise of 3.5 dBA. Further analysis is extended to axial flight conditions, with axial velocities of 5.0, 10.0, and 20.0 m/s, to simulate vertical climb. The results show that these hover-optimized blades are capable of sustaining noise decreases of 3.0-4.0 dBA compared to the baseline, provided that the thrust coefficient is maintained through adjustments in collective pitch.
ABSTRACT This study examines the performance of a quadcopter in edgewise flight conditions with flow simulated using the commercial Navier-Stokes solver, AcuSolve, with a Detached Eddy Simulation (DES) model. The rotating volume around each rotor interfaces with the remainder of the computational domain using a sliding mesh. Simulations were conducted for an AeroQuad Cyclone quadcopter at 10 m/s forward speed, 5 deg nose-down pitch attitude, operating in both cross and plus configurations. From the results it was observed that in the cross configuration, the aft (South) rotors showed a 19% reduction in lift (relative to an isolated rotor at the same forward speed, pitch attitude and RPM), with an associated 3% reduction in torque. The loss in lift was primarily at the front of the aft rotors due to the downwash induced by the forward rotors, therefore reducing the aft rotor nose-up pitching moments by 54% (relative to operation in isolation). In the plus configuration, sections of the East and West rotors close to the aircraft center-body operate in upwash induced by the North rotor, increasing the lift generated by 5.5% and 7.6% respectively, relative to operation in isolation. The South rotor sees both upwash (and increased lift) along the advancing and retreating edges induced by the East and West rotors, as well as downwash at the front (and reduced lift) induced by the North rotor, but no significant overall changes in thrust or torque.
This paper presents the design and development of a swashplateless micro helicopter with a target endurance of more than 30 minutes using an optimized direct drive rotor connected to a unique rotor hub that has blades with a flap hinge and proprietary skewed-lag hinge with pitch-lag kinematic coupling. This obviates the need for conventional swashplate based cyclic pitch control, as the cyclic variation in control angle is achieved by cyclically varying the motor RPM. UP12 underactuated propulsion system developed by VertiQ is used for the baseline design. The blades in this propulsion system are optimized using Blade Element Momentum Theory (BEMT) analysis with lookup table to enhance its performance. BEMT is validated using experimental measurements and then used to optimize the geometry of the rotor. The optimized blades offer better performance and are 30% lighter than the original 3D-printed plastic blades. The prototyping of the Micro Aerial Vehicle (MAV) is completed by fabricating the airframe and using off-the-shelf electronic components. Tabletop tests are carried out using a yawing test stand to tune the yaw attitude gains for the MAV. A coupled flap-lag rotor dynamics is combined with the motor angular speed dynamics to simulate the dynamics of the system to carry out fundamental understanding of the control power being generated by the swashplateless system. The swashplateless rotor design is able to generate significant propulsive and side forces, however it is unable to generate control moments which may be inadequate for aggressive maneuvering warranted by small agile MAV.
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.
Winged Quadcopters are an increasingly popular UAS configuration due to their mechanical simplicity and high degree of aerodynamic efficiency, but this efficiency is highly sensitive to the chosen blade pitch and rotor orientation. In this study, a rotor-wing system representative of a winged quadcopter is simulated and a parametric sweep of blade pitch, rotor tilt, cruise speed, and weight is conducted. At the baseline 30 kts cruise speed and 3 lb vehicle weight, the optimal configuration (blade pitch: 10° – 20°, rotor tilt: 30° – 40°) is 4.4 times more efficient than the baseline Quadrotor Biplane Tailsitter (blade pitch: 0°, rotor tilt: 0°). Even if flight speed and weight is increased (up to 50 kts and 9 lb), combinations of blade pitch and rotor tilt can offer improved efficiency; and at the optimal condition, 12.5° blade pitch and 35° rotor tilt is 5.3 times more efficient than the baseline QBiT. The rotor-wing system is also simulated using CFD with the rotor at 58 different positions around the wing. Rotors mounted directly in front of the wing leading edge near the wing tip operate at the lowest power, but if the rotors are located laterally inboard on the wing, directly below or above is more favorable. While the effect of rotor position is somewhat small compared to that of blade pitch or rotor tilt, all are viable means for improving winged quadcopter aerodynamic efficiency. Proper selection of these parameters can dramatically transform the aerodynamic efficiency of these UAS, enabling profound enhancements to range, endurance, and payload capacity.
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.
Multicopters operate in environments subject to strongly gusting winds, and need good aeromechanical models to improve the aircraft. A common, convenient, assumption is that the gusting inflow is quasi-static at each instant, but this assumption has never been tested. This paper shows that there is a solid physical basis for the simplified aerodynamic models of multicopter response to gusts. Experiments and computations show that using the static relationship between thrust or power and aerodynamic angle of attack for a multicopter rotor (the quasi-static assumption) in sinusoidally pitching sideflow can be used to predict the thrust or power for unsteady variation of angle of attack if the instantaneous flow angle of the freestream is known. Vertical (angle) gusts up to 1885°/s (k=2.2 based on diameter) and with a wavelength longer than the rotor diameter were shown to be covered by this assumption.
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.
This paper investigates the role of the aerodynamic torque on propeller whirl flutter stability. The generalized force due to the torque is first computed and subsequently included in the equations of motion of a rigid propeller-pylon system. Preliminary evaluations indicate that the torque modifies the real part of the backward and forward modes, providing a stabilizing effect on powered propellers. Analyses are conducted on a 3-bladed propeller driven by an electric motor. Stability predictions are obtained with a simple analytical model and validated by multibody simulations coupled with a mid-fidelity aerodynamic solver, based on a vortex particle method. Furthermore, a simple control law acting on the propeller's collective pitch and rotational speed is presented. The control variables are modified to increase the whirl flutter stability margins, without altering the trim conditions of the aircraft. Results demonstrate the effectiveness of the proposed control strategy, although propeller efficiency is reduced. In a failure scenario, the control law can be exploited to compensate for a weak mounting stiffness, allowing the aircraft to land safely.
This paper investigates the feasibility of using machine learning to predict whirl flutter bifurcation diagrams. The machine learning techniques selected for the study are XGBoost and the long short-term memory neural network. These techniques are selected for their suitability for sequential and nonlinear data. The techniques are investigated for a propeller-nacelle test case with polynomial structural nonlinearities resulting in supercritical or subcritical whirl limit-cycle oscillations. The techniques are trained to learn the bifurcation diagram for the amplitude variation of pitch angle limit-cycle oscillations of the propeller-nacelle system as a function of the forward speed for various levels of cubic structural nonlinearity. Bifurcation diagram learning and testing data are generated using the bifurcation forecasting method. XGBoost is computationally faster to train but less accurate for low amounts of learning data, especially for the most weakly and strongly nonlinear cases. The long short-term memory neural network is more computationally expensive to train but shows a less scattered error pattern for sparse learning data. However, it is sensitive to the amplitude resolution of the bifurcation diagrams. The approach to sample the cubic nonlinearity range does not significantly impact the results once the techniques have a sufficient amount of data to learn from. The data requirements observed in the study suggest that, for these techniques, direct learning of bifurcation diagrams may not scale beyond a handful of input parameters.
A key objective of this work was to develop a quantitative rationale to explains some aspects of pilot rating variability, as this would point to the fundamental principles driving pilot response that may not be observable if averaged ratings are used as a handling qualities metric. This paper hypothesizes that the factors affecting a pilot's ability to stabilize and control an aircraft following abrupt control motion is neither the damping nor the frequency of the ensuing oscillation, but rather the length of time that the oscillation remains large enough to interfere with the task (i.e., the product of damping and frequency). A handling qualities metric is introduced called the decay rate parameter that reflects the decay rate of the closed loop dominant mode. Closed loop pilot-vehicle oscillation decay rates were generated by a pilot model employing pitch (visual channel) and pitch rate (vestibular channel) tracking strategies. These decay rates were used to predict minimum and maximum handling qualities ratings and pilot induced oscillation (PIO) ratings which closely matched actual pilot ratings from an inflight PIO study using a variable stability NT-33A aircraft. PIO frequency prediction results were excellent. Predicted handling qualities and PIO ratings from a piloted NASA Vertical Motion Simulator study also closely matched the actual ratings. The results indicate that even with the highest fidelity motion simulation, pilot control relies primarily on the visual channel and is constrained by its inherent limitations. Conversely, the dominant control strategy appears to be the vestibular channel when pilots conduct a visual task that is anchored in the physical, out-the-window environment. The vestibular channel is shown to incur effectively no time delay.
This study models the interaction of a two-bladed 14" propeller with the ground under different configurations using actuator disk method (ADM) where the rotor is modeled using unsteady momentum sources distributed over the entire disk. While ADM has been extensively used for standard rotorcraft analysis, it's performance in unconventional operating conditions remains an open question. Exhaustive experiments conducted at DEVCOM Army Research Laboratory are compared with ADM to evaluate the inexpensive method's ability to predict rotor loads for parametric variations in rotor-ground interaction scenarios. Partial ground effect (part of the rotor operating IGE), side-by-side rotors in ground effect and variation in IGE pitch attitude are specifically considered in this study. ADM generally predicts the thrust increase in partial ground effect (PGE) as the rotor goes from OGE to IGE although the increase is somewhat earlier and milder than measured in experiments. Side-by-side rotors in ground effect are predicted remarkably well by ADM which captures the fountaining flow between the rotor wakes as well as reduction in thrust seen compared to IGE in experiments. Finally, pitch attitude variation cases showed reasonably good agreement in trend at lower angles although ADM shows a more gradual drop in thrust at higher angles (beyond 15◦) compared to the sharper drop seen in experiments.
Tailsitter configurations that operate in both fixed and rotary wing flight modes are typically capable of generating large control forces and moments, making them inherently capable of rapid transitions and aggressive maneuvers. However, harnessing these capabilities requires feedback control strategies that can effectively estimate the non-linear aerodynamics loads involved to successfully exploit them. This paper describes initial steps in combining an onboard flow sensing strategy with a data-driven approach to estimating inflight air loads. A neural network is trained to use measurements from a multi-hole probe to predict the output from a set of pressure sensors embedded in a wing section undergoing a series of pitch motions in a wind tunnel. We hypothesize that this limited context of emulating a sensor network represents a focused and compartmentalized approach to applying emerging data-driven techniques to challenging aeronautical problems. We compare estimation results from a set of neural networks with varying input configurations to assess the feasibility of our approach and the significance of different sensing modalities on overall performance. Current results show that a gated recurrent network (GRU) trained with unsteady pressure measurements was able to predict the chordwise pressure distribution on a pitching NACA 2412 airfoil using probe measurements, reproducing the transient and non-linear effects observed in our dataset.
A real-time capable simulation model is developed for a 1200 lb quadcopter with hybrid variable-RPM and collective pitch control. Linear models and trim points are calculated using the Rensselaer Multicopter Analysis Code (RMAC), and controllers are designed to meet flying qualities specifications in hover and forward flight. Four control modes are flown by test pilots in a flight simulator. Three hybrid control configurations (Eco, Standard, and Sport modes) are evaluated, along with a baseline variable-RPM case. Five mission task elements (MTEs) are flown to test the handling qualities of each axis: Depart/Abort, Slalom, Hover Turn, Vertical Maneuver, and Precision Hover Task. Pilot feedback is collected in the form of handling qualities ratings (HQRs), as well as general comments. The baseline RPM control case is shown to be undesirable to pilots due to its increased delays, with the average HQR in the Level 2 region. Sport mode performs the best overall, with the average HQR being near the Level 1/2 boundary, though some pilot comments suggest that it may be over designed. The average HQR for Standard mode also falls on near the level 1/2 boundary, and pilots are generally satisfied with its performance. Eco mode is limited by its low pitch actuator margin, and has the lowest HQR due to its slow response and reduced magnitude of commands. Overall, it is shown that the hybrid control quadcopter is able to significantly outperform its variable-RPM counterpart in piloted simulation of MTEs.
A limited flight load survey was performed on a fleet representative UH-60L aircraft flown with various degrees and combinations of dynamically imbalanced blades. Blades selected for this test ranged in both positive and negative severity of pitching moment slope values to substantiate rotor smoothing efficacy and the effect on dynamic component oscillatory loads for dynamically imbalanced blades compared to a baseline of nominally balanced blades. For every component analyzed, greater than 70% of the maneuvers presented showed an increase in structural loads in the unbalanced configuration compared to the baseline balanced blades. Most components and maneuvers that did experience an increase in loads remained non-damaging. However, the damper experienced several maneuvers where an increase in loads may have indicated a change in fatigue lives. Based on a colloquial rule of thumb.
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A hovering two-bladed, 2-m-diameter rigid rotor was used to investigate the effects of dynamic pitch inputs at multiple blade loading coefficients. Unsteady rotor loads and the flow field were measured using an internal load cell and time-resolved particle image velocimetry, allowing for inflow, tip vortex trajectories and strength to be measured. Two sets of measurements were performed with an image acquisition rate of 44/rev: constant pitch angle (CT/σ = 0.06, 0.08, 0.10), and measurements with dynamic pitch ramp inputs (Δθ0 = 1° amplitude change over four rotor revolutions) at CT/σ = 0.06 and 0.08. The measurements at constant pitch angle and with dynamic pitch inputs were performed over 670 rotor revolutions, in both cases yielding statistically converged averages. In addition, the experiment was simulated using a finite-volume CFD solver. The results were correlated to the experimental data to assess the accuracy of the flow simulations and to gain further insight into the flow phenomenology. Vortex trajectories and axial velocities (inflow and wake) correlated well, and the tip vortex strength was slightly underpredicted by the CFD. From measurements and predictions, it was found that (independent of the thrust setting) the axial velocity in the rotor plane steadily increased to the value after the pitch input. Furthermore, vortex trajectories were little affected by the increased thrust and inflow. The vortex strength revealed trends similar to the axial velocity and gradually increased to the values after the pitch ramp input.
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.
Presently the fatigue lives of MH-60R dynamic components and airframe are based on a usage spectrum developed using pilot surveys. In order to better define the usage spectrum and to extend component and airframe fatigue life, the Health & Usage Spectrum (HUMS) System was installed on the U.S. Navy MH- 60R Rotorcraft. So far 207 aircraft are equipped with the HUMS systems and 121,334 flight hours of good data have been recorded. The regime recognition programs recognize 315 maneuvers, but are consolidated to 94 maneuvers of MH-60R usage spectrum, for which the component measured loads are available. To better define usage spectrum in detail and compute realistic component fatigue life, an additional maneuver of low Angle Of Bank (AOB) from 10 to 25 degrees was added, but the measured component loads were not available at this AOB to implement HUMS. Thus, measured flight loads data of level flight and AOB turns at 30, 45, and 60 degrees were utilized to derive component loads at 20 degrees by spline cubic interpolation technique. The cubic interpolation technique was applied to measured minimum, average, and maximum loads of variation at 10, 30, 45, and 60 degrees to interpolate load at 20 degree. This technique was applied to interpolate loads for pitch control rod, swash-plate, drag damper, shaft bending moments, blade cuff stresses, and flap deflections. The spline interpolation loads correlated with measured available loads of pitch control rod and blade stresses. The probabilistic fleet usage spectrum of various severities was developed using the HUMS recorded data of 121, 334 hours from 179 rotorcraft with and without low AOB usage. It is evident that fatigue life with 20 AOB split is significantly higher for all dynamic components. Thus, to implement HUMS successfully, it is necessary to compute loads that are not available in the original component fatigue life calculations. Further prorates of gross weigh (GW), velocity and altitude based on the HUMS fleet usage should be implemented to extend component fatigue lives.
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.
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