Browse Topic: Pitch
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
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
This paper investigates the use of multi-modal cueing through full-body haptic feedback to enhance pilot-vehicle system (PVS) performance, reduce mental workload (MWL), and increase situational awareness (SA) in both good and degraded visual environments (GVE/DVE). Piloted simulations were conducted using an H-60-like flight dynamics model in a virtual reality (VR) motion-based simulator, evaluating two ADS-33-like mission task elements (MTEs) – precision hover and slalom – under visual-only and combined visual and haptic feedback conditions in both GVE and DVE. The H-60 flight dynamics were augmented with a dynamic inversion (DI)- based stability augmentation system (SAS), implementing rate-command/attitude hold (RCAH) response type on the roll, pitch, and yaw axes and altitude hold response type on the vertical axis. The SAS was designed to achieve Level 1 handling qualities per ADS-33 standards. The full-body haptic cueing strategy leveraged an outer-loop DI control law, which
This paper 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
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
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
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
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
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
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
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
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
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
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
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
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
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.
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
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
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
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
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
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
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
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
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
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
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
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
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
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
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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This paper describes the development of full flight envelope dynamic inversion outer-loop control laws used to control airspeed and flight path for two Future Vertical Lift-relevant rotorcraft configurations - a lift offset coaxial helicopter with a pusher propeller and a tiltrotor. The outer-loop control laws for both aircraft include a control allocation scheme to account for redundant controls and reduce pilot workload. A piloted simulation experiment was conducted at the Penn State Flight Simulator facility using a series of high-speed handling qualities demonstration maneuvers to evaluate the handling qualities of the control laws. Overall, the outer-loop control laws for both coaxial-pusher and tiltrotor aircraft were assigned Level 1 handling qualities for the Break Turn and High-Speed Acceleration/Deceleration tasks, and reduced pilot workload over previously developed inner-loop control laws. The outer-loop control laws also improved performance and reduced pilot workload in a
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