Browse Topic: Thrust
This paper presents results of flight tests conducted on a coaxial ultralight helicopter. An automated flight test evaluation method is presented and exemplified through its application to steady horizontal flight. The results shown include pilot controls, helicopter attitude angles, power, thrust and torque distribution between the rotors, rotor harmonic thrust components, and teeter angles, along with their rotor harmonic components across varying flight speeds. This study focuses on the dependencies of these parameters on center of gravity position and sideslip angle.
An experimental investigation was conducted to characterize the effects of partial-ground on the aerodynamics of a hovering rotor. A model-scale rotor was tested at a range of heights above ground and under partial-ground coverage, and rotor hub forces and moments were measured using a six-axis force/torque transducer during constant-power operation. The measurements were used to develop a semi-empirical thrust ratio model that accurately captures trends from out-of-ground effect to full-ground effect conditions. This model predicts realistic thrust behavior at low ground-coverage conditions, exhibiting high adjusted R2 and minimal root mean square error. Time-resolved particle image velocimetry was conducted for selected cases to examine induced flow features and to qualitatively assess changes in the downwash and edge-driven crossflow associated with partial-ground interactions. A geometric rotor-ground interaction area based on a circular-segment formulation was correlated to the thrust coefficient over the interacting region. Results show that thrust increases with ground coverage and decreases with increasing out-of-ground height, whereas moment coefficients increase with decreasing height. The increases in thrust and moment can be attributed to a reduction in the induced velocity above ground, as observed in the measured flow field, thereby increasing the effective ground-induced pressure beneath the rotor.
The induced and profile power of a hovering rotor was evaluated using experimental and computational methods. Momentum theory principles were coupled with experimental measurements over a range of thrust conditions to characterize the induced and profile power consumption at low Reynolds number conditions ∼ 105. An empirical induced power factor, κi, was extracted to quantify the non-ideal losses. Results show that these losses increase as the Reynolds number reduces, and nearly twice the power is required at Retip = 0.27×105 than the ideal momentum theory prediction. These results were compared with high-fidelity computational fluid dynamics simulations using the partial-pressure field (PPF) force/power decomposition to extract the induced and profile power contributions of the rotor. The PPF method decomposes the static pressure field of a numerical Reynolds-averaged Navier-Stokes solution into Euler and dissipative partial pressure fields. Simulations were performed across a range of thrust conditions, from which the induced power factor and profile drag coefficient,Cd0 , were computed for each simulation from the extracted power contributors.
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.
During helicopter air-to-air refueling the rotor of the helicopter might enter the slipstream of the tanker aircraft's propeller. Based on blade element momentum theory, the impact of the accelerated air within the propeller slipstream on rotor blade aerodynamics (thrust, rolling and pitching moments) can be solved analytically. Also, DLR's comprehensive rotorcraft code has been used with the Pitt-Peters induced inflow plus rotor-rotor interference model. Additionally, DLR's free-wake code was used for both the propeller and the helicopter main rotor, including mutual wake-wake-interactions. The helicopter rotor's collective and cyclic controls needed for disturbance rejection are computed with all these models for a typical air-to-air refueling scenario without and with blade flapping motion. A propeller wake affecting the retreating side of the rotor requires much larger control inputs to retrim than an impingement on the advancing side. The results of all modelling approaches are compared to each other and generally a reasonably good agreement is found, despite their very different levels of complexity and computational effort.
The development of a coupled computational structural dynamics (CSD) and electrodynamic suspension (EDS) system was critical in modeling and predicting the aeromechanics of MagLev Aero's (MLA) propulsion system, ensuring safe testing and proving viability of levitated rotors for vertical lift systems. This advancement validates the feasibility of this enabling technology in applications of uncrewed aerial systems (UAS) with high hover lift efficiencies. This paper explores the implementation of an electromagnetic motor hub on a large-root-cutout, slowed rotor system with a specific focus on the impacts on aeromechanics: loads, performance, vibrations, and aeroelastic stability. The performance benefits of a large-root-cutout system, with an external or internal rotor, are well known; however, the mechanisms to implement such a design have been impractical. The development of an EDS motor bearing enables previously unattainable configurations like large-root-cutout and tip-driven ducted systems that have advantageous loading, acoustic, and vibration performance; such benefits thus require characterization in dynamics and aeroelastic stability to prove the feasibility of these complex concepts. Furthermore, the ability to investigate a multidisciplinary, multiphysics-based system is nontrivial; customization and advancement of existing comprehensive analysis systems is required to solve these complex problems. The investigation of performance, dynamics, and aeroelastic stability for the EDS rotor on a subscale thrust test system, EXO600, is being substantiated through MagLev Aero's test programs. The results presented herein are a subset of the validations in aeromechanics and electrodynamics.
Developed in the frame of the European Clean Sky 2 program, the RACER High Speed Helicopter Demonstrator of Airbus performed its maiden flight on April 25th, 2024. In the continuity of the previous high-speed demonstrator X3 (1st flight in 2010) the RACER is a 7/8t (15000 / 18000 lb) class compound helicopter powered by two SHE Aneto-1X engines, including a wing and two propellers. The tail rotor is removed as the two propellers control the yaw axis by differential thrust. At flight 07, with its initial default settings, it reached a true airspeed of 227 kts in level flight, exceeding its objective of 220 kts.
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.
The influence of ground, wall, and corner boundaries on multirotor vehicle performance was investigated through a series of controlled flight tests. Changes in rotor inflow profiles were represented by near-field rotor pressure measurements captured by a custom Kiel probe wake rake. Ground effect was characterized by reduced thrust and power requirements, primarily driven by the vehicle fuselage, which induced regions of reduced pressure and increased flow unsteadiness around the airframe. Operating near a wall boundary was found to restrict airflow into the portion of the rotor disk closest to the wall, leading to increased power requirements to maintain hover and a consequent reduction in performance. While vehicle orientation had minimal impact on overall rotor performance, it did influence local rotor inflow behavior near the wall, depending on the relative position of the interaction region formed with adjacent rotors. As the vehicle descends from the isolated wall effect into corner effect, created by the intersection of the wall and ground, an exchange between the dominating ground and wall effects is observed, with corresponding ground heights and wall distances identified.
This paper explores a significant step forward, regarding the further detailed understanding of the Fenestron®. Since its patent in 1968 – for the Gazelle helicopter –, the shrouded tail rotor has been resized, inclined, modulated, etc. and has thus been continuously enhanced on different rotorcraft. Half a century after its invention, Airbus is once again exploring in more detail the magic of the Fenestron®, with the objective of optimizing it even further, for future helicopter applications. To grasp and observe properly some specific phenomena, a model (scaled to one third) capable of both unprecedented functions and modularities, was developed. The present paper will describe in detail the novel model and the related challenges and solutions. This model is capable of high rotor speed and dynamic pitch inputs, delivering power levels high enough to reach stall effects, while allowing the measurement of propulsive efficiency and to differentiate rotor vs fairing thrust. Furthermore, the model had to provide aerodynamic-shape modularity, both in the shroud, on the covers and on the tail gearbox supports. The first test campaign performed on this model allowed us to define design drivers and aerodynamic preferences. And since one idea often leads to another, another challenge is being addressed on this scaled mock-up: Reduction of the tail rotor noise level.
This paper 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 presents experimental research aimed at developing novel low lubrication methods for rotorcraft and jet engines, focusing on sustaining minimal lubrication to prevent catastrophic bearing failure during loss of lubrication (LoL) events or to increase fuel consumption performance on once-through, fuel-oil bearing lubrication engines. Utilizing two high-speed bearing test rigs simulating low and high thrust class engine conditions, the study establishes lower bounds for oil flow rates necessary to maintain thermal stability and prevent thermal runaway in hybrid ball bearings. These findings inform the design of the Zulu Pod (ZPod), a passively driven, self-contained oil delivery system that uses engine compressor bleed air to precisely meter lubricant flow. Engine test stand results demonstrate that replacing traditional fuel-oil lubrication with the ZPod system reduces thrust specific fuel consumption (TSFC) by an average of 7%, with up to 11% savings, without compromising engine thrust or bearing health. The ZPod offers a simplified, efficient alternative to fuel-lubricated systems by eliminating fuel diversion for lubrication, enhancing fuel efficiency, and maintaining bearing performance in attritable or single-use engines. Additionally, the study highlights the potential of minimal lubrication supplied by the ZPod to extend operational life during LoL scenarios, enabling safer aircraft recovery. Future work will focus on extending testing to higher thrust classes and optimizing ZPod designs for broader applications.
Rotor performance in a Martian environment was analyzed with an objective of increasing thrust with minimal impact on efficiency. The Sample Recovery Helicopter (SRH) and Rotorcraft Optimization for the Advancement of Mars Exploration (ROAMX) rotors were studied by varying solidity, blade count, and chord distribution to determine which configuration delivered the most desirable performance. For all configurations, the ROAMX rotor displayed better performance than the SRH rotor. It was observed that increasing solidity reduced the blade loading required to achieve the peak figure of merit, and beyond a solidity ratio of 0.3 the figure of merit was negatively impacted. For both rotors a 6-bladed configuration with a solidity ratio of 0.3 delivered the optimal figure of merit.
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.
The Shake-The-Box technique was applied to experimentally quantify the time-resolved volumetric flow field around a free-flying quadcopter UAV with an overall span of about 0.5 m. State-of-the-art LED illumination and high-speed camera equipment was combined with modern Lagrangian tracer particle tracking and data assimilation techniques, facilitating a measurement volume larger than 1.5m3. The setup allowed for both hover and limited maneuvering of the quadcopter, while resolving even small details of the complex interactional aerodynamics. In hover out of ground effect, the four individual rotor wakes merged into a single jet within a few rotor radii below the rotor planes. Evaluating the mass and momentum fluxes over suitable control volumes yields accurate estimates for the quadcopter's total thrust, the asymmetric thrust distribution between front and back rotors, and the entrainment of external flow through turbulent mixing. Hover in ground effect decreases the power requirement and induces recirculating flow in the center of the four rotors. The outwash pattern is non-uniform with jets developing between the rotors and pointing in radially outward directions. Forward flight cases result in a skewed, rapidly merging wake flanked by the roll-up of two "super-vortices" similar to the wingtip vortices of fixed-wing vehicles.
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.
Rotorcraft responses to idealized disturbances are examined to gain insights into model fidelity requirements for flight simulations of the ship-rotorcraft dynamic interface. Two disturbance fields are considered: an isolated straight vortex that represents the canonical vortex that results from the corners of flat top ships in oblique wind-over-deck conditions and a horseshoe vortex derived from a nondimensional characterization of the time-averaged flow observed aft of a simplified ship superstructure. Rotorcraft models considered include: an analytical blade element theory-based rotor model, where the disturbance velocities are integrated over the rotor, and a coupled blade element / free wake flight dynamic model of the full UH-60 aircraft, which is used to perform time-marching simulations with the disturbances modeled as a frozen field that is fixed in space and not interacting with the aircraft (one-way coupling), and as a distorting field (two-way coupling). Analytical thrust responses of four idealized rotors to the disturbance associated with the horseshoe vortex suggest the existence of a scaling relationship between nondimensional blade loading and the product of nondimensional vortex width and strength. Time-averaged rotor thrust responses to the disturbance induced by the isolated straight vortex indicate significant distortion of the rotor wake and disturbance field at low forward speeds, with the frozen field assumption indicating greater disk tilt responses. Modeling the horseshoe vortex disturbance as a frozen field generates a greater rotor response than in the distorting case. Inclusion of full aircraft dynamics attenuates rotor responses to the disturbances when compared to isolated rotor responses.
The effectiveness of the counter-torque capabilities of rotorcraft tail rotors has undergone extensive study and development due to its critical role in balance and maneuverability. The Fenestron provides an interesting alternative to the conventional open tail rotor, improving upon several key elements, but still demonstrates drawbacks in others. This paper seeks to provide a side-by-side comparison of Fenestron and open tail rotor performance, in terms of thrust and power, using the mid-fidelity surface-vorticity panel-method flow solver FlightStream® to simulate Fenestron and open rotor configurations in hover and cruise conditions. While this work is not a comprehensive analysis, as it does not compare to experimental data, nor implements exact geometrical representation, it does propose trends between the Fenestron and open tail rotor cases.
Installation effects of the Volocopter 2-X beam structures are studied by performing high-fidelity CFD simulations of a single and three-rotor configurations in hover. The studied cases are compared with simulations without airframe to investigate the installation effects. In addition, the noise emission of the configurations is simulated by using a Ffowcs Williams-Hawkings based CAA code. Scattering effects are also included by using a BEM code. The rotors are simulated at an identical RPM and are placed in their mounting position. Furthermore, an additional setup with individual rotor RPMs is simulated for the three-rotor configuration. The installation mainly affects the rotor wake, thrust and pressure fluctuations on the rotor, while the integral aerodynamic quantities remain almost unchanged. This resulted in additional oscillations in the acoustic pressure signal. Overall, the installation increases the OSPL by about 1.5 dB, but has a greater effect on the 3-20 harmonics. The simulation data were compared with the Volocopter measurements and showed good agreement. For the three-rotor configurations, the rotor-rotor interactions were found to be dominant for the aerodynamic and acoustic performance, while the installation effects only locally influenced the noise emission by 3-4dB.
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.
This study uses a mid-fidelity, aeromechanics coupled framework using the Lattice-Boltzmann Method (LBM) fluid solver to investigate an experimental coaxial rotor system. Co-rotating and counter-rotating rotor operation scenarios in hover are studied. The rotors are represented as an actuator line in the LBM fluid simulation. Simulation results are compared to experimental data, mid-fidelity and CFD simulation results available in literature. Results indicate that the framework can accurately predict thrust and thrust variations for both upper and lower rotors. Power prediction has deficiencies compared to experimental and CFD results, but is in line with mid-fidelity simulation results in literature. Flow field results are also compared qualitatively with CFD results. Results are sensitive to the actuator line representation of the blade, the inflow sampling, and tip corrections.
ABSTRACT A proof of concept test to measure the unsteady boundary layer transition locations on the lower surface of a Machscaled rotor in forward flight was performed during the Summer of 2017 in the NASA Langley 14- by 22-Foot Subsonic Tunnel. The transition locations were measured using high-speed infrared thermography with a rotating mirror assembly that could be remotely actuated to acquire data at several rotor azimuths. Data were acquired for eight unique rotor flight conditions for a range of advance ratios (μ=0:10 : 0:38), thrust coefficients (CT/α =0:04 : 0:12) and rotor shaft angles (αs = -6 deg : 0 deg). This paper presents the transition locations as a function of azimuth and radius for an advance ratio of, μ, of 0.30, and thrust coefficent, CT/α, of 0.08. At this condition, the lower surface is fully laminar on the retreating side and mostly turbulent on the advancing side except near the tip. The tip airfoils were greater than 60 percent laminar on the lower surface advancing side. Capturing the location of natural transition on a rotating blade in forward flight represents a new advancement toward understanding the boundary layer state and its important contribution to rotor aerodynamics. Documentation of the boundary layer transition location during testing is critical to understanding scaling model to full-scale performance data, validation of newly developed turbulence models, and the design of the next generation of high performance rotor blades.
ABSTRACT Within the framework of NACOR project in CleanSky 2 AIRFRAME ITD, ONERA and DLR performed parallel investigations dealing with the RACER high-speed demonstrator, and especially with its tail parts, each partner respectively focusing on vertical fins (ONERA) and horizontal stabilizer (DLR). During this design phase, most of the CFD simulations were steady-state and neglected the effect of the rotor (or rotor-head) and of the propellers. It however turned out that the rotor-head had a significant effect on the vertical fins and that it was essential to take into account its rotation in time-accurate simulations: the wake from the rotor-head, the upper deck and the engine cowlings indeed strongly impacts the left vertical fin because of the clockwise rotation of the rotor-head. It induces strong oscillations on the tail unit loads, and the mean tail unit lateral thrust is also significantly increased. Moreover the main conclusions of this 'aerodynamic interactions' investigation are almost identical, no matter what the computed configuration: rotating rotor-head, rotating rotor-head with actuator-disk, rotating full-rotor or rotating full-rotor with propellers effect.
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.
ABSTRACT Phase-resolved particle image velocity measurements were taken to document the wake generated by a rotor operating in ground effect above inclined surfaces. In particular, the current work focused on the average wake structure and axial velocity distribution through the rotor. A two-bladed rotor was operated at a height of one rotor radius above a ground plane, and ground plane angles from 0° to 30° were investigated. Rotor performance measurements were also taken, using a six-axis load cell, to examine the effect ground plane angle had on the thrust produced and power required. The wake structure was found to be very sensitive to ground plane angle causing the radial distribution of axial velocity through the rotor to increase inboard and decrease outboard with increasing ground plane angle. The peak figure of merit of the rotor decreased with increasing ground plane angle.
ABSTRACT This paper presents the preliminary results from the experimental study of the aerodynamic characterization of a novel dissimilar coaxial rotor concept with reduced rotor-rotor interaction. The performance of dissimilar coaxial rotor concept is compared with existing rotor configurations such as conventional single main-rotor tail-rotor and regular coaxial rotor. The performance measurement is carried out using a small scale coaxial hover test stand facility setup for this purpose. The hover performance of different rotor concepts namely conventional, regular coaxial and dissimilar coaxial rotors are compared as anti-torque mechanisms with same baseline rotor by comparing their power loading. The inter-rotor spacing between coaxial rotors is kept constant at 15% of the rotor radius. The dissimilar coaxial rotor appears to be a promising solution for development of high efficiency hovering rotor system for high thrust condition. From the tests conducted on a main rotor of 0.415 m radius operated at 1000 RPM, it is established that the dissimilar coaxial rotor system consistently shows higher power loading than the conventional rotor (especially at high thrust condition) as well as the coaxial rotor system. To achieve maximum benefits from the dissimilar coaxial rotor system, the anti-torque rotor in the dissimilar rotor pair must be spun at 50% or 40% of the main rotor RPM. The effect of the cross-section shape of the blade-hub attachment (root cutout) plays a significant role in reducing the profile power loss of the anti-torque rotor of the dissimilar coaxial rotor. In the current study, three different shapes were compared and the elliptical cross-section shape was observed to give best performance.
A towing tank investigation of a single rotor blade operating at hovering and high advance ratio conditions is presented. A custom blade was manufactured and instrumented with fully bridged axial strain gauges to monitor the flap bending strain at three radial locations. Measurements of rotor thrust and torque were obtained to characterise the rotor aerodynamic environment for advance ratios ranging from 0.4 to 1.00 and to identify the presence of stalled and reverse flow. Strain measurements obtained at three locations across the blade span show minima and maxima at approximately the same azimuthal location as the load data. Moreover, the strain distribution shows a growth in strain magnitude with increasing advance ratio. Spectra of strain shows a dominant 1/rev signal and for the ∅ = 25° collective, non-harmonic frequencies are observed due to aperiodic vortex shedding from the presence of stalled flow.
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.
The flow behavior of the two-blade MERIT rotor in hover, focusing on both pre-stall and stall regimes, is investigated through a comprehensive numerical-experimental approach. The study leverages unsteady RANS simulations to compute rotor thrust and power polars and validates them against experimental measurements. Valuable insights are provided into the capabilities of unsteady RANS methods and modern turbulence models for predicting rotor performance across these critical operating conditions. Furthermore, the numerical model incorporates blade deformations by implementing the experimentally measured flap and torsion displacements. A more realistic depiction of the rotor's aerodynamics is provided accounting for the structural deformations of the blades under aerodynamic loads. Highfidelity simulations closely predict the experiments in pre-stall conditions while discrepancies are present when the flow exhibits extended stalled regions. Blade deformations demonstrated to have only a minor effect on thrust and power polars due to the high stiffness of the system.
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.
A new framework for performing high-fidelity computational aeromechanics simulations of the V-22 tiltrotor aircraft in hover mode has been developed. It is built on the HPCMP CREATE-AV Helios tool and utilizes scripted input generation and automatic replacement of modular model components. This new framework has been used to investigate the impact of various approaches to modeling the rotor aerodynamics, airframe aerodynamics, and periodic blade motion on predictions of aircraft hover performance in and out of ground effect. The findings indicate that an actuator line method can provide rotor performance predictions with comparable accuracy to a meshed-blade approach. However, body-fitted meshes are required to compute accurate airframe download. Furthermore, active trimming of rotor collective to a target thrust provides more representative aircraft aerodynamic performance than directly applying a collective angle as measured during flight test. The computational framework can accurately capture trends in rotor loads and airframe download as height-above-ground is varied relative to measured flight-test data.
This study models flow around isolated and side-by-side three-bladed propellers in (IGE) and out of ground effect (OGE) using actuator-based techniques of varying fidelity. Actuator techniques model propellers using momentum sources distributed over the disk in actuator disk method (ADM) or distributed over moving lines in actuator line method (ALM) to reduce computational cost compared to blade-resolved DDES simulations. The lowest fidelity ADM method is observed to reasonably predict thrust with the use of a tip loss model to control runaway thrust at the tip while not resolving flow features such as blade-bound vortices and helical tip vortices at a fraction of the cost of BR-DDES (1/100). The coarser ALM model resolves these features but still requires a tip loss model to control runaway thrust at 1/10th the cost of BR-DDES. Finally, the finer ALM model used in this study accurately captures blade-related features and further predicts the tip loss trend from first principles at 1/3rd the cost of BRDDES. Demonstrating the efficacy of these techniques for a commonly encountered flow scenario - side-by-side rotors at 2.5R hub separation are simulated where turbulent fountaining flow is observed between the rotors which eliminates the thrust increase normally seen IGE, a feature captured by both ADM and ALM techniques. However, only ALM captures the impulsive 3-per-rev thrust loading seen in BR-DDES.
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.
ABSTRACT
ABSTRACT
Emerging vertical flight concepts being proffered for solutions to the Future Vertical Lift (FVL) mission set such as compound high speed rotorcraft can be designed with multiple, coupled control effectors thus creating redundant systems in one or two more axes to generate control forces and moments which allow for a range of trim states. In the FVL mission area future rotorcraft will be asked to fly into high threat environments where potential failure modes can be encountered due to enemy fire or mechanical failure causing reduction of the safe flight envelope. Fault detection creates options to increase the survivability of the crew and passengers allowing an emergency flight envelope to be proposed. One of the more serious potential failures due to enemy fire is a loss of yaw control. Faults in yaw control can be detected in a compound rotorcraft with a vectored thrust ducted propeller (VTDP) or similar anti-torque thruster. An online Kalman filter (KF) for a dimensional yaw moment coeff icient model will be used to estimate vehicle yaw coeff icients. Deviation from the nominal coefficients will be monitored based on the KF statistics in the case of both rudder and tail rotor failure at 60, 40, and 20 ft/s in forward flight. Both frozen zero rudder and ganged sector faults as well as failed tail rotor faults were successfully detected at all airspeeds except the failed tail rotor at 60 ft/s. For the yaw control faults considered, post fault excitation appears airspeed dependent. An online KF estimator for yaw control fault detection could successfully be integrated into the design of a compound rotorcraft with VTDP thereby increasing system safety.
In this work, a contract-based reasoning approach is developed for obstacle avoidance in unmanned aerial vehicles (UAV's) under evolving subsystem performance. This approach is built on an assume-guarantee framework, where each subsystem (guidance, navigation, control and the environment) assumes a certain level of performance from other subsystems and in turn provides a guarantee of its own performance. The assume-guarantee construct then assures the performance of the overall system (in this case, safe obstacle avoidance). The implementation of the assume-guarantee framework is done through a set of contracts that are encoded into the guidance subsystem, in the form of a set of inequality constraints in the trajectory planner. The inequalities encode the relationships between subsystem performance and operational limits that ensure safe and robust operation as the performance of the control and navigation subsystems and environment evolve over time. The contract inequalities can be obtained analytically or numerically using an optimization based path planner and UAV simulation. The methodology is evaluated in the context of head-on obstacle avoidance, where the contracts are constructed in terms of (1) minimum obstacle detection range, (2) expected obstacle size, (3) maximum allowed cruise velocity, (4) maximum allowable thrust, roll and pitch angles, and (5) inner-loop tracking performance. Numerical and analytical generation of these contracts for this scenario is demonstrated. Finally, in-flight contract enforcement is illustrated for typical scenarios.
The paper discusses the application of the Array Controlled Turn-less Structures (ACTS) motor for VTOL application. The motor enhances the three main competing characteristics of electric motors; namely specific power, efficiency and reliability. The motor arrays an ensemble of elemental turn-less motors which include turn-less elements each with their dedicated inverters which are operated in synchronism. The resulting small pole size enhances the power density, the enhanced conductor packing enhances the efficiency, and the massive parallelism enhance the reliability. Vertical takeoff requires much higher thrust compared to wing assisted takeoff. With limited on-board power, this higher thrust is presently provided by in ordinary larger propulsion disk area which reduces the craft aerodynamics, and the cruising Lift-to-Drag (L/D) ratio and accordingly the flight efficiency and range. The high specific power of the ACTS motor allows for a different scenario and thus craft architecture. By substantially increasing the takeoff power which is now possible with the higher specific power of the ACTS motor, the propulsion disk area can be substantially reduced, with the resulting greater streamlined, high L/D craft, and thus longer range. The paper discusses key aspect of the motor and inverter architecture and technology. Furthermore, it discusses its application to a high L/D VTOL and prospective performance.
Acoustic characteristics of two- and four-bladed rotors with a 1.108 m radius in hover were measured experimentally at rotor speeds up to 1200 RPM and tip Mach number up to 0.41. The Rotorcraft Comprehensive Analysis System (RCAS) using a viscous vortex particle method (VVPM) coupled with PSU-WOPWOP were used to simulate the acoustic characteristics of the same rotors. Simulations were also conducted for a stacked rotor over a range of azimuthal spacings with zero axial spacing. PSU-WOPWOP predictions included thickness, loading, and broadband noise. Both experiments and simulations showed that broadband noise was the dominant contributor to A-weighted sound pressure level at a distance of 8.25 radii from the rotor center. Simulations showed a sharp increase in the broadband and overall noise at higher thrust, but the experiments did not show such an increase. Higher harmonic components of the tonal noise were nearly as large as the blade passage frequency noise in the experiments, but were underpredicted in the simulations by up to 30 dB. Simulations incorporating wind showed closer agreement in the higher harmonic tonal components. The unweighted overall sound pressure level in the stacked rotor simulations showed 5 dB to 10 dB variations with azimuthal spacing, but the A-weighted tonal and overall SPL was insensitive to azimuthal spacing.
As rotor diameter and inertia increases, the quickness of the thrust response to pilot inputs slows, yielding negative implications to handling qualities and limitations on scaling electric propulsion. This study presents a novel approach to alleviating these scaling effects by introducing a pitch-lag coupled hinge to the root of 40" and 50" diameter props. The impacts of chordwise hinge placement and hinge angle are examined and compared to a baseline rigid rotor to provide physical understanding of the rotor dynamics. It is shown that a coupled hinge can be designed to maintain propeller efficiency for a design thrust, while increasing the sensitivity of thrust to rotor speed and the maximum thrust of an RPM-limited rotor. Finally, the dynamic implications of this are tested using a first-order motor model. When a 40" diameter trimmed rotor is set to max throttle, rotors with a coupled hinge angle achieve a 6% higher thrust in 9% less time. The dynamic response improvement scales favorably when the rotor diameter is increased. For the 50" diameter rotor, the introduction of pitch-lag coupling reduces the time constant of the rotor’s thrust response by 32%.
The British Experimental Rotor Programme (BERP) tip design is well known for its superior performance for high speed flight. This paper revisits the BERP design by presenting a parameterized model of the planform design based on published data and investigating its performance on the Apache rotor blade. The underlying airfoil sections, HH02 and NACA64A006, are retained in a new Apache BERP-shape rotor blade. The performance of the Apache BERP-shape rotor blade is evaluated for hover and forward flight by using US Army CREATE™-AV Helios software and compared with the Apache baseline rotor blade. The results are presented in the form of rotor thrust, rotor torque, figure of merit, trim condition, sectional blade loading and pressure distribution. A visualization of complex vortex structures is provided to offer insight into the airflow characteristics at different flight conditions.
This paper presents a new blade element momentum theory (BEMT) for a coaxial rotor in hover. The new BEMT solves the upper and lower rotor induced velocities iteratively so that the mutual rotor-to-rotor interaction is accounted for. The upper rotor induced velocity is affected by the lower rotor thrust and induced velocity, while the lower rotor induced velocity is affected by the upper rotor thrust and induced velocity. Two empirical constants are included in each rotor calculation. This new BEMT provides the performance of each rotor as a function of the rotor separation distance. The new BEMT is validated with measurement data for two coaxial rotor experiments. The first experiment validates the thrust to power coefficients at a given separation distance. The second experiment validates each rotor's figure of merit, thrust, power, interference loss factors, etc., as a function of the rotor separation distance. It is shown that the BEMT captures the trends and magnitudes of the performance as a function of the rotor separation distance compared to the measurement data. Detailed radial distributions of aerodynamic properties are also presented at several separation distances.
The purpose of this study is to determine if phase synchronization can be used to reduce the net radiated sound power from two rotors. Phase synchronization implies that the rotors have the same rotational speed with a fixed relative azimuthal blade position, or phase. The concept is evaluated both experimentally and numerically. Measurements of source directivity and thrust are initially compared with predictions to confirm that the model accurately captures the relevant trends. The model is then used to explore the design space and identify relevant parameters. Both experimental and numerical results show that the radiated sound power at the blade passage frequency can be reduced by appropriately controlling the relative azimuthal phase of the rotors. Vehicle level predictions are also provided for a notional octocopter, comparing two different modes of operation. Predictions show that phase synchronization can be used to achieve a 4-5 dB reduction of the sound pressure level at the blade passage frequency nearly everywhere on the ground plane beneath the vehicle.
This paper discusses an effort to determine useful scaling laws for predicting the sound pressure levels at various observer positions for small scale rotors in hover. Surveys of rotor thrust, rotor torque and near-field acoustics are acquired using a series of identically shaped two-bladed rotors ranging between 8 and 18 inches in diameter. Rotor speeds were varied between 50 and 90 revolutions per second which equate to blade tip Mach numbers as high as 0.56. Common sound metrics like overall sound pressure level, sound pressure level and blade-pass frequency harmonic noise were correlated with both dimensional variables (rotor thrust, torque, power and blade tip speed) and non-dimensional variables (thrust and power coefficients, Reynolds number, blade tip Mach number and figure of merit) pertaining to rotor performance. The findings revealed a new correlation parameter for predicting rotor noise based on the product of figure of merit with blade tip Mach number. It is postulated that an increase in rotor collective translates to an increase in loading noise. So long as the flow remains attached, the increased loading noise is expressed by an increase in figure of merit. This new correlation parameter is shown to provide a reasonable collapse of all rotor noise data. In order to account for the directivity of the sound field, a second order response surface model of this new correlation parameter was developed and validated.
The present study proposes and explores a new autonomous morphing concept, whereby an increase in helicopter rotor blade camber of the order of 12-13° is realized over the inboard section of the blade with increase in ambient temperature. The camber change is achieved through a proper integration of Shape Memory Alloys (SMAs) on the lower surface of the blade aft of the leading-edge spar. For a reference rotor (no-SMA) generating 21,000 lbs thrust, operation in hot conditions resulted in a 2,590lb loss in lift. When the SMA camber morphing section extends from the blade root to 50% span, the rotor recovered up to 43% of the lift loss at high temperature. If the camber-morphing section is further extended to 75% span, up to 82% of the lost lift can be recovered.
Slowing down the rotor in forward flight is a viable means of extending the cruise speed of a rotorcraft by alleviating compressibility effects at the advancing side blade tip. It was shown by previous wind tunnel tests that an articulated rotor trimmed to zero hub moment generates limited thrust at high advance ratios, because the advancing side of the rotor needs to be trimmed against the retreating side in the reverse flow state, where the rotor is ineffective in generating thrust. Therefore, a hingeless rotor that allows the advancing side to generate more thrust can be rewarding in overall thrust potential. At the University of Maryland, a rotor test stand was modified for hingeless rotors and two wind tunnel tests were conducted to investigate the behavior of hingeless rotors at high advance ratios. The experimental results, including performance and control, hub vibratory loads and blade structural loads, are presented in this paper and compared with predictions of the in-house comprehensive analysis, UMARC. The performance results demonstrate that a hingeless rotor with lift offset is capable in generating more thrust at high advance ratios, and the blade structural load results reveal that the predominant structural constraint on the rotor is in 2/rev flap bending moment, combined with 1/rev flap bending moment from the lift offset.
Hub load measurements were taken on a rotor in ground effect above a ground plane undergoing single degree-of-freedom motion. A two-bladed teetering hub rotor was operated in a hovering condition at multiple heights above a sinusoidal pitching or heaving ground plane. A six-axis load cell was used to measure the load generated by a rotor at multiple blade collective pitch angles, as well as ground plane amplitudes and frequencies. The magnitude and frequency of the ground plane motion was determined based off a ratio of rotor blade passage frequency (for a representative maritime helicopter) to ground motion frequency (for a DDG-51 class ship). A distinct and measurable fluctuation on the rotor pitching moment was observed in the presence of ground plane pitching motion, and variations in thrust as high as twenty percent were observed in the presence of ground plane heaving motion. An appropriate scaling parameter for the problem was investigated and determined to be the ratio of the rotor inflow velocity to the peak ground motion velocity. This parameter showed good agreement over the complete range of measurement conditions, including when the rotor was operated at or near zero thrust.
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