Browse Topic: Pitch
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
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
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
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
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
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
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 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
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
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
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
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
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
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
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
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
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.
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
In any human space flight program, safety of the crew is of utmost priority. In case of exigency in atmospheric flight, the crew is safely and quickly rescued from the launch vehicle using Crew Escape System (CES). CES is a critical part of the Human Space Flight which carries the crew module away from the ascending launch vehicle by firing its rocket motors (Pitch Motor (PM), Low altitude Escape Motor (LEM) and High altitude Escape Motor (HEM)). The structural loads experienced by the CES during the mission abort are severe as the propulsive, aerodynamic and inertial forces on the vehicle are significantly high. Since the mission abort can occur at anytime during the ascent phase of the launch vehicle, trajectory profiles are generated for abort at every one second interval of ascent flight period considering several combinations of dispersions on various propulsive parameters of abort motors and aero parameters. Depending on the time of abort, the ignition delay of PM, LEM and HEM
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 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
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
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
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
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
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
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
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