Browse Topic: Flaps
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The paper presents a general framework for building an aeromechanic model in FLIGHTLAB, suitable for high fidelity, pilot-in-the-loop simulator. The focus is on aerodynamic modeling of AW609 tiltrotor in Airplane Mode flight regime. The framework can be extended to helicopter and conversion modes with additional considerations for rotors-airframe aerodynamic interference. It can also be adapted to different tiltrotor geometries, with some adjustments depending on their peculiarities. The model uses Blade Element Theory loads evaluation of lifting surfaces, corrected with tabulated distributed loads to tune FLIGHTLAB predictions against high-fidelity aerodynamic references. Bluff bodies are modeled using force and moment tabulated data. Verification was conducted against reference data in wind tunnel mode and against flight data in trim analysis. The proposed method allowed to match lift distribution on slender bodies, as well as lift and drag integral loads, with aerodynamic references. Trim analysis has shown satisfying accordance with experimental data for all the comparison parameters. However, rotor aerodynamic modeling is still being investigated to improve correlations in torque prediction. Additionally, installation effects, such as those related to the interaction between the wing and fuselage with the rotors, are under examination to address biases in flap motion evaluation.
Whirl testing of a full-scale rotor with positive flap-bending/twist composite coupled blades was performed to evaluate the dynamic and performance effects of the coupling. A positive flap-bending/twist coupling, in which a flap up deformation induces a nose down elastic twist, was introduced in the blades through tailoring of the laminate layups; the magnitude of the coupling was maximized through an optimization of the layup, with the intent of maximizing the potential impact of the coupling for correlation purposes. An uncoupled version of the blade using the same geometry and materials was also fabricated to provide a baseline set of measurements for comparison, with the coupled blade optimized to also minimize changes in bending and axial stiffness properties in an effort to isolate the effect of coupling by itself. Rap testing was conducted to measure blade modal frequencies and shapes in a free-free environment. Whirl testing was performed for both the coupled and baseline rotors to measure frequencies and damping for stability, load distributions, and performance characteristics. Rotorcraft Comprehensive Analysis System (RCAS) models were run for all cases and are presented along with the measurements for correlation purposes; generally good agreement was observed in trends between measurements and predictions.
Enhancing rotor efficiency has been a persistent challenge in the development of micro aerial vehicles (MAV) especially for surveillance and covert operations. This study introduces a new Hybrid Flapping Wing Rotor (Hybrid FWR) configuration inspired by insect's wing flapping mechanics to address the efficiency limitation of traditional rotor designs. Unlike traditional rotary systems that rely solely on rotational motion, the Hybrid FWR combines rotational and flapping motions to significantly enhance lift generation. A comprehensive mathematical model was developed to analyze and predict the optimal aerodynamic performance, demonstrating that the Hybrid FWR configuration achieves a substantial improvement, with a power efficiency increase of up to 2.148-fold compared to conventional micro rotorcraft. Experimental validation was conducted to confirm the theoretical predictions, identifying an optimal hybrid ratio of approximately 0.7, which effectively minimizes aerodynamic resistance during the upstroke phase while maximizing lift during the downstroke. This bio-inspired hybrid approach addresses critical limitations of existing MAV rotors, such as limited operational endurance and range. The findings of this research contribute significantly to the advancement of micro rotorcraft technology, presenting a promising direction for future MAV developments with enhanced flight performance and energy efficiency.
Structural testing of full-scale blade geometries with flap-bending/twist composite coupling was performed to evaluate the impact of coupling. Full-scale spar geometries were first fabricated with three different coupling distributions, including two with a uniform positive flap-bending/twist coupling, in which a flap up deformation induces a nose down elastic twist. The third spar geometry incorporated a mixed coupling, with a uniform positive coupling at the inboard end and a uniform negative coupling at the outboard end, where the negative flap-bending twist coupling produces a nose up elastic twist when experiencing flap up deformation. A full-scale blade was then fabricated with a positive flap-bending/twist coupling. Measurements of the structural twist distribution of the cured spars were taken to ensure the coupling did not result in any hygrothermal instabilities. Tip twist and strains were then measured under various combinations of flatwise bending and torsional bending loads and correlated with Rotorcraft Comprehensive Analysis System (RCAS) and Variational Asymptotic Beam Sectional Analysis (VABS) predictions for both the spars and blade. Using the strain measurements, stiffness properties of the test articles were also calculated to confirm the predicted coupling levels. Generally good agreement was observed between the measurements and analytical predictions for both the full-scale spars and the full-scale blade.
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 use of sub-scale vehicles as a means of predicting full-scale vehicle behavior has historically been applied to flight dynamics testing and evaluation for aircraft operating in Earth atmospheric conditions. However, the use of sub-scale testing on Earth has not been as thoroughly explored for Martian rotorcraft. In this paper, sub-scale vehicles of varying sizes were developed in simulation using Froude scaling laws to evaluate their ability to estimate fullscale linear dynamics for the Mars hexacopter, Chopper. Blade loading, Lock number, and flap frequencies were held fixed when scaling and corresponding relationships for vehicle length, mass, inertia, and rotor speed derived. Full-scale frequency response, gain margin, and instability characteristics are explored for hover and forward flight cases in a variety of Mars-to-Mars and Earth-to-Mars conditions. Mach effects are also analyzed as a consequence of Froude-scaling by comparing sub-scale vehicles that are Mach-matched to the full-scale Chopper. Finally, a first-order approach for downselecting a sub-scale vehicle based on feasibility is introduced.
This paper outlines the investigation into the effect of static stall onset in hover on the deformation of rotor blades, comparing the behaviour of a stiff blade featuring a NACA0012 aerofoil, rectangular planform and no taper, and a hingeless blade attachment; with a more flexible blade featuring a NACA23012 aerofoil, twist and taper, and a leadlag hinge. The Munich Experimental Rotor Investigation Testbed (MERIT) at the Technical University of Munich (TUM) was operated in a two-blade configuration at a variety of rotational speeds and collective pitch angles, paired with a stereooptic high speed photogrammetry system. The post-processing methodology used to extract flap and torsional deformations despite the presence of a hinge is outlined, and it was shown that the hinge affected the onset of flow separation and subsequent deformations. A comprehensive set of experimental deformation data for a repeatable setup has been generated and published.
The aeroelastic stability of rotor blades in the flap, lag, and torsion degrees of freedom is analyzed in preparation for high-advance ratio wind tunnel testing of Mach-scaled rotors. A wide range of advance ratios (0 ≤ μ ≤ 3) are evaluated for articulated and hingeless rotor configurations. Linearized equations of motion in the rotating frame are derived, which consider periodic coefficients, reverse flow, pitch-flap and pitch-lag coupling, and control inputs. The steady state trim is compared with wind tunnel data. Floquet theory is used to evaluate the stability of the equations of motion in response to perturbations from trim. Results are compared to past analyses and expanded to higher advance ratios. Damping and frequency response behavior are evaluated, and rotor stability boundaries are presented.
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.
This paper presents an open-loop hover experiment and analysis for a 4-blade Mach-scaled Seoul National University Flap (SNUF) rotor. A detailed finite element analysis is attempted to predict allowable experiment range that provides sufficient structural integrity. Multi-body dynamic analysis DYMORE and cross-sectional design program VABS are used to analyze the present trailing-edge flap rotor blade, and the flap hinge stiffness is calibrated on the static bench test. Ground testing on the present rotor shows a linear strain-displacement response, and the relevant result shows better than 80% correlation against DYMORE prediction. Appropriate test matrices are constructed and two of those are attempted herein. The first one is the baseline no-actuation collective sweep test at two different rotating speeds at tip Mach of 0.22 and 0.3, respectively. The results are utilized to correlate between the momentum theory and free wake empirical parameters. Next, a single active flap blade test is attempted to assess the trailing-edge flap driving component. It is found that the dynamic trailing-edge flap deflection result needs to be both electrically and mechanically isolated from the test stand. In the future, trailing-edge flap deflection will be recorded for 1 - 5/rev activation at tip Mach 0.5, while four-blade actuation engaged. The frequency response of the trailing-edge flap mechanism will be identified, and 1/rev tracking control will be attempted.
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.
An aeromechanics analysis of a Mach-scaled rotor with lift compounding was conducted to understand the impact of various wing configurations on performance and loads. An assessment of the single retreating side wing and dual wing configurations was conducted for advance ratios up to μ = 0.7, two wing incidence angles (4° and 8°), and three rotor shaft angles (-4°, 0°, and 4°). Aircraft performance, control angles, blade structural loads, hub vibratory loads, and aerodynamic interactions between the rotor and wing were evaluated using the University of Maryland Advanced Rotorcraft Code (UMARC). Additionally, UMARC coupled rotor-wing analysis was validated with wind tunnel data of a lift and thrust compounded rotor. The study shows that the single wing configuration is beneficial for peak vehicle performance (L/D), though the dual wing configuration minimizes blade loads. The single wing configuration observed a 7% greater wing L/D than the dual wing configuration for the same 8° wing incidence angle at μ = 0.5 and αs = 0°, however, the dual wing configuration yielded a 20% lower steady flap bending moment. The study showed that the wing wake has a negative effect on the rotor performance especially at rearward rotor shaft angles; this is overcome with the efficiency gains from lift offset present in a single wing configuration but absent in the dual wing configuration. Hence, the single wing configuration attains the highest performance while the dual wing configuration minimizes structural loads. A lift compounded rotor is limited at high advance ratios since it is almost entirely offloaded, resulting in increased total lift to drag ratio at the cost of large blade structural loads.
ABSTRACT
Presently the fatigue lives of MH-60R dynamic components and airframe are based on a usage spectrum developed using pilot surveys. In order to better define the usage spectrum and to extend component and airframe fatigue life, the Health & Usage Spectrum (HUMS) System was installed on the U.S. Navy MH- 60R Rotorcraft. So far 207 aircraft are equipped with the HUMS systems and 121,334 flight hours of good data have been recorded. The regime recognition programs recognize 315 maneuvers, but are consolidated to 94 maneuvers of MH-60R usage spectrum, for which the component measured loads are available. To better define usage spectrum in detail and compute realistic component fatigue life, an additional maneuver of low Angle Of Bank (AOB) from 10 to 25 degrees was added, but the measured component loads were not available at this AOB to implement HUMS. Thus, measured flight loads data of level flight and AOB turns at 30, 45, and 60 degrees were utilized to derive component loads at 20 degrees by spline cubic interpolation technique. The cubic interpolation technique was applied to measured minimum, average, and maximum loads of variation at 10, 30, 45, and 60 degrees to interpolate load at 20 degree. This technique was applied to interpolate loads for pitch control rod, swash-plate, drag damper, shaft bending moments, blade cuff stresses, and flap deflections. The spline interpolation loads correlated with measured available loads of pitch control rod and blade stresses. The probabilistic fleet usage spectrum of various severities was developed using the HUMS recorded data of 121, 334 hours from 179 rotorcraft with and without low AOB usage. It is evident that fatigue life with 20 AOB split is significantly higher for all dynamic components. Thus, to implement HUMS successfully, it is necessary to compute loads that are not available in the original component fatigue life calculations. Further prorates of gross weigh (GW), velocity and altitude based on the HUMS fleet usage should be implemented to extend component fatigue lives.
The scope of this paper is to assess the accuracy of the Lattice-Boltzmann/Very Large Eddy Simulation Method to predict the aerodynamics and aeroacoustics of helicopter rotors in strong Blade-Vortex Interaction conditions, and to validate a computational approach to include the effects associated to the rotor blade deflections into the numerical setup. The numerical flow solution is obtained by solving the explicit, transient and compressible Lattice-Boltzmann equation implemented in the high-fidelity CFD/CAA solver Simulia PowerFLOW R. The acoustic far-field is computed by using the Ffwocs-Williams and Hawkings integral solution applied to a permeable surface encompassing the whole helicopter geometry. The employed benchmark configuration is the 40% geometrically and aeroelastically scaled model of a BO-105 4-bladed main rotor tested in the open-jet anechoic test section of the German-Dutch wind tunnel in the framework of the HART-II project. In the present study, only the baseline operating condition of the HART-II test, without Higher-Harmonic Control enabled, is considered. Simulations are performed either assuming a fully-rigid blade motion or a computational strategy, based on a combination of a velocity wall boundary condition applied on the blade surface and fluid body-forces fields applied in proximity of the blade leading- and trailing-edge, to partially retrieve the effects related to the experimental blade flap and torsion deformations, respectively. The impact due to the inclusion of the blade elastic deformations into the computational setup on control settings, unsteady air-loads and noise footprint predictions is outlined.
The Mars Helicopter is a 1.8 kg coaxial rotorcraft designed to demonstrate aerial mobility at the surface of Mars after deployment from the Mars 2020 rover. In this paper, the authors present the development of the Mars Helicopter rotor system from preliminary design through fabrication and testing of the flight hardware. The vehicle has a 1.21 m counter-rotating coaxial rotor system which is driven by electric motors and which features collective and cyclic controls on both the upper and lower rotor sets. The rotor blade design is characterized by the low Reynolds number (∼104), high Mach number (∼0.7), high stiffness (first flap frequency ∼1.9/rev), and minimum mass. Airfoil design focused on minimizing drag at the low operating Reynolds number while maintaining sufficient spar depth for structural requirements, and the blade planform was based on a minimum induced loss profile with modifications to reduce mass of the outboard blade sections for increased flap frequency. The swashplate and servos were designed to provide actuation of the rotor at up to 12Hz bandwidth while minimizing system mass and volume, and the primary structure of the rotor blades and hubs relies heavily on molded carbon fiber composites and ceramic ball bearings.
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.
An analytical study, which is based on calculations from a rotorcraft comprehensive analysis of a notional coaxial rotor system in isolation, is presented to improve understanding of the aeromechanics of lift-offset coaxial rotors. The calculations include a trim analysis that enforces a fixed value for coaxial system lift in order to isolate the effect of lift offset. The quantities examined here are the blade pitch controls, coaxial system performance, blade airloads, blade structural loads, and coaxial system hub loads. Operating conditions examined form airspeed sweeps for four different values of lift offset. The maximum airspeed that the rotor may achieve trim is increased by increasing lift offset. For a given airspeed, increasing lift offset generally increases the coaxial system lift-to-effective-drag ratio as a result of a decrease in shaft power (despite an increase in propulsive power). Increasing lift offset decreases half peak-to-peak values of pitch link force, torsion moment, and in many cases, chord bending moment. While increasing lift offset increases half peak-to-peak flap bending moment for sufficiently low airspeeds, it may decrease this quantity for higher airspeeds. Lastly, increasing lift offset generally decreases the fixed-frame, vibratory, system hub horizontal force and pitch moment.
This paper discusses an endeavor to experimentally identify the flight dynamics of the AVFL Hummingbird, and quantify its maneuverability and gust tolerance using a control theoretic framework. The AVFL hummingbird is a 62gram, truly biomimetic robotic hummingbird developed to understand and characterize hummingbird flight. It has a pair of biologically inspired, aeroelastically tailored wings flapping at 20Hz, and is fully hover capable. Additionally, like its biological counterpart, it utilizes wing kinematic modulation techniques for control and stability. The vehicle states were measured during targeted flight tests from which a linearized, state-space model was derived. The model contained damping aerodynamic coefficients, decoupled longitudinal, lateral and directional dynamics, as well as large control coefficients. The control theoretic framework, which quantifies the maximum controllable states of the system under unit inputs, was utilized to calculate the maximum gusts tolerable by the control system. The results showed exceptional gust tolerant capabilities. The maximum gusts tolerable were (1) longitudinal gust: 21.2 ft/s (6.4 m/s); (2) lateral gust: 17.7 ft/s (5.4 m/s); (3) lateral rotational gust: 149.8 rad/s; and (4) longitudinal rotational gust: 20.5 rad/s. These are much greater than comparable rotary-wing based systems. This study represents the first time the maneuverability and gust tolerance of a hummingbird-like system has been experimentally characterized, and has shown quantitatively the exceptional flight capabilities offered by biomimetic design and control.
ABSTRACT The challenge of increasing range and speed of a rotorcraft is encountered in the scope of the European CleanSky2 "Fast Rotorcraft" project by Airbus Helicopters with the compound helicopter design RACER (RapidAndCostEfficientRotorcraft) for which the box wing and the tail parts designs are respectively protected by patent. This paper presents the DLR contributions to the RACER development. This includes the aerodynamic design of the wing and tail section as well as an overall assessment of performance and noise. In a first step the aerodynamic properties of the configuration are evaluated both isolated and with consideration of the main rotor and lateral rotor interferences by the use of actuator discs. In the second step, the investigated possibilities to improve the configurations performance are described. These include airfoil design for improved high lift performance of the wing and tail section, an optimization of the box wing circulation distribution on the upper and lower wing. Additionally, the intersection fairings were improved and the efficiency of the trim flaps was evaluated. In this regard, it could be determined for which cases an isolated approach is appropriate and when the rotor interference should be considered. At the end the evaluation of the aero acoustics of the configuration is conducted. The applied configuration shows good aerodynamic characteristics with some further cruise and off design optimization potential.
ABSTRACT Comprehensive vibration analysis of a rotor-airframe-engine-drivetrain system using a time-domain modal coupling approach was conducted. Pair-wise couplings of components (airframe and drivetrain/engine) were performed to isolate the contribution of each component to the complete coupled system, and the effect of each component on blade loads and hub loads was studied. The drivetrain model is a 6-dof model consisting of inertia and torsional spring elements, while the airframe model is a NASTRAN superelement of a detailed finite element airframe model for a medium-lift utility helicopter. Drivetrain coupling resulted in elastic twist of the rotor shaft by less than 0.02 degrees, but there were noticeable reductions in the chordwise blade bending moments as well the 8/rev hub torque. The airframe coupling produced very small hub translation amplitudes, less than 5×10⁻⁴ inches, however it had a significant impact on the higher harmonic flap bending, increasing the 9/rev flap moment by up to 60% near the root.
ABSTRACT This paper describes the architecture and performance of a 36 kg to 56 kg synchropter (intermeshing rotors) with 3.3 m rotor diameter that was specifically designed for high altitudes. The design mission of this prototype is to climb from 5000 m to 9000 m above mean sea level, descend back to 5000 m with autorotation, and cruise back to the departure point. The design is intended to be a compromise between controllability in gusty winds and low overall system power consumption through low disc loading and tip speed. The prototype features variable tip speeds from 100 m/s to 135 m/s and is able to hover with an overall system power of 80 W/kg. This paper presents flight test data of this prototype at low altitude with a special focus on rotor performance and rotor blade root flap moments in hover, forward flight, and vertical climb. The flight test data is compared to a lower order aeromechanics model using the comprehensive helicopter analysis CAMRAD II that was previously used to design the rotorcraft. The couplings between the control inputs and their corresponding axis motion are discussed. Finally, the design mission was simulated by the performance validated model to bridge the design and the real prototype.
ABSTRACT This paper describes the development of a biomimetic robotic hummingbird that utilizes biologically inspired wing kinematic modulation strategies for active stability and control. By tilting the flapping planes, varying the relative wing flapping amplitude, and shifting the mean position of the flapping stroke, the robotic hummingbird is able to modulate the magnitude, direction, and location of the lift vector of each of the wings in the same way that hummingbirds do to maneuver and stabilize themselves. In addition to the control strategies, biologically inspired, flexible, aeroelastically tailored wings were developed for use on the vehicle. Flight tests were conducted in which the vehicle was flown in a controlled hover using combinations of control techniques to quantify the effectiveness of each in stabilizing the vehicle. In the present study, emphasis was placed on pitch control, where two different control strategies were investigated, which were (1) pure tilting of the stroke plane, and (2) stroke mean shifting. Based on the flight experiments, a combination of 90% mean shifting and 10% stroke tilting proved to be significantly more effective than pure stroke plane tilting with almost 40% decrease in drift and attitude oscillations. This is the first ever study where the exact kinematic control strategies used by real hummingbirds and two-winged insects are implemented on a robotic flyer and investigated through free-flight testing.
ABSTRACT Aeroelastic stability of stiff-in-plane hingeless rotors is investigated using the comprehensive analysis RCAS. Aeroelastic stability analysis of stiff-in-plane rotors in hover is compared to experimental measurements that shows an overall fair to good agreement for various rotor parameters. The analysis reveals that blade lead-lag damping decreases sharply and the blades become aeroelastically unstable when the blades stall. Stiff-in-plane rotor aeroelastic stability analysis in forward flight is compared to a previous numerical study. Then, using the rotor models as a baseline, a parametric study is performed for various rotor parameters including aerodynamic models, rotor speed, rotor thrust, lead-lag frequency, precone, contol system flexibility, and tip sweep. The parametric study covers lead-lag frequencies of stiff-in-plane rotors from 1.1 /rev to 1.4 /rev with a flap and a torsional frequencies of 1.15 /rev and 3.0 /rev. The parametric study shows that blade lead-lag mode of the stiff-in-plane hingeless rotors tends to become aeroelastically unstable at high advance ratios. Blade lead-lag damping of the stiff-in-plane rotors decrease as the lead-lag frequencies increase from 1.1 /rev to 1.4 /rev. Aeroelastic stability boundaries (advance ratio) increase as the rotor speed increases for a fixed rotor lift. Aeroelastic stability boundaries increase as the rotor lift decreases for a fixed rotor speed. Aerodynamic stall demonstrates significant effects on aeroelastic stability of stiff-in-plane hingeless rotors.
ABSTRACT This study describes the deformation measurement and automated operational modal analysis (OMA) of a rotor blade in hover. Blade deformation of a 0.4 m-diameter two-bladed rotor was measured at two different root pitch angles and four rotational speeds up to 1500 RPM by a time-resolved digital image correlation technique (DIC). The DIC technique successfully measured the time history of 3D displacements over the entire rotor blade at approximately 900 measurement locations for the flap, lead-lag, and torsional degrees of freedom. The measured blade deformation data were then processed with the Complexity Pursuit (CP) algorithm, which is one of several Blind Source Separation (BSS) techniques, to determine the dynamic characteristics of the rotor blade without input excitation information. The modal identification process was able to reduce the number of manual steps that would have to be performed by an analyst in a conventional operational modal analysis. The identified modal parameters were compared with that obtained by another OMA technique called NExT/ERA as well as a numerical model. These results agreed quite well for modal frequency identifications within the range of 5% difference. The mode shape estimated by the CP algorithm were almost similar to the numerical results for identified modes, however, there was non-negligible discrepancy between the results of the CP and NExT/ERA approach. Modal damping was also extracted by the CP algorithm although the resultant values were less than 3% for all test cases, which is quite low under the influence of aerodynamic damping. Overall, the automated OMA with the CP algorithm was favorable to the conventional OMA approach in terms of accuracy of modal parameter identification as well as process simplification.
ABSTRACT This Paper provides the first comprehensive description of full scale demonstrator activities currently under way in Leonardo Helicopters. Two distinct technology strands are outlined, an active Gurney flap intended to address performance enhancement and active trailing edge flaps intended to address vibration control. Passive blade design constraints and the ability of active technology to overcome these constraints are discussed. The derivation of the two demonstrator rotor design's configuration and construction is reviewed. The required control algorithm development is outlined. A system hardware outline is provided together with a brief overview of actuation system options. The development of design loads for both rotor types is discussed, touching on challenging nature of minimizing conservatism in this new area. The construction of the two rotor types is reviewed, with explanations of the modifications employed to the baseline AW139 rotor blade. The Paper concludes with an outline of the flight trials content and preliminary estimates of system effectiveness in the target parameters of vibration control and performance.
ABSTRACT This paper investigates a utility helicopter rotor infield tracking and vibration control design. A Discrete Trailingedge Flap (DTEF) is employed as an actuation appliance to correct the unbalance. A rotor model with trailingedge flap is developed using Dymore - a multibody dynamic code. Two dissimilar rotors are modeled to simulate unbalanced inertia force and unbalanced aerodynamic force situations respectively. An adaptive closed-loop regulator is introduced and applied in both working conditions. Infield tracking is then conducted and control corrections are made according to different dynamics. Control process is presented and analyzed in advance ration u=0.233 in terms of controller inputs and outputs. Infield tracking and control correction at varying advance ratio is conducted in the following step and results are presented and explained.
ABSTRACT A new active flow control strategy by placing a synthetic jet actuator (SJA) and a trailing-edge flap (TEF) has been proposed, and its control effects on mitigation of large negative pitching moments and drag caused by rotor dynamic stall are numerically investigated by CFD method. A moving-embedded grid method and an unsteady Reynolds averaged Navier–Stokes (URANS) solver are established for predicting the complex flowfields of rotor and airfoil. Calculated results of VR-12 and SC1095 airfoils indicate that TEF and SJ can suppress the formation of dynamic stall vortex and postpone flow separation over rotor airfoil, resulting in much lower Cdmax and Cmmax comparing to the baseline state, and aerodynamic characteristics of airfoil could be further improved by the new control method comparing to individual control method. Furthermore, parametric analyses on dynamic stall control of airfoil by the combinational method are conducted, and it indicates that aerodynamic characteristics of an oscillatory rotor airfoil could be significantly improved when non-dimensional frequency (k*) of TEF is about 1.0, and larger angular amplitude help to increase the lift coefficient. At last, simulations are conducted for the flow control of rotor airfoil by the combinational method, and the numerical results indicate that the TEF can better improve the lift characteristics of rotor airfoil, and synthetic jet further increases the aerodynamic characteristics of rotor airfoil.
ABSTRACT Tiltrotor whirl flutter in cruise flight is investigated using comprehensive rotorcraft analysis codes CAMRAD II and RCAS. A generic tiltrotor model with a 3-bladed gimballed rotor was systematically developed starting with a simple rigid rotor mounted on a rigid pylon and a more sophisticated model was built up by adding one design variable at a time. The rotor is also coupled with a flexible wing/pylon modeled from NASTRAN for aeroelastic stability analysis. The effects of pitch-flap coupling (δ₃), blade elasticity, precone, undersling, yoke chord and flap stiffness, pitch link stiffness, rotor rotational speed, density, speed of sound, inflow modeling, unsteady aerodynamics, and realistic airfoil tables on whirl flutter speed are thoroughly examined. With careful and thorough modeling/analysis, aeroelastic stability (frequency and damping) calculated by CAMRAD II and RCAS shows consistently excellent agreement with each other for wide variations of design variables and operating conditions. For the configurations investigated in this study, blade pitch-flap coupling, rotor lag frequency, rotor rotational speed, and density play an important influence on whirl flutter speed.
ABSTRACT Instantaneous lift force and structural deformation experiments were performed on a flexible, structurally characterized, low-aspect ratio wing modeled after the wing of a hovering flapping micro air vehicle (FMAV). A six-component force balance was used to measure the variation in lift produced by the wing over the course of a flap cycle. A VICON motion capture setup tracked the passive wing deformations at select locations of the wing during experimentation. Measured lift force and wing deflections were compared against the results of a coupled computational fluid dynamics/computational structural dynamics (CFD/CSD) aeroelastic analysis. The CFD analysis was developed based on an unsteady Reynolds-averaged Navier–Stokes (URANS) solver while the CSD analysis consisted of a general purpose multibody dynamics solver capable of modeling geometrically nonlinear beam and shell elements. The CFD/CSD results were able to capture the overall trend in lift force variation and wing deformation, but the magnitude of variations were under-predicted by the coupled aeroelastic solver. The coupled CFD/CSD solver was used to evaluate the influence of increased wing flexibility on the aerodynamic forces, vortical flowfield and wing deformation. Results showed that decreasing wing stiffness increased the time-averaged aerodynamic lift with minimal effect on aerodynamic drag and prolonged the sustainment of the leading edge vortex.
ABSTRACT This paper presents a method for modeling gimbaled rotor dynamics under trim conditions in X3D, a next-generation 3-D finite element based rotor structural dynamics solver. The rotor is modeled using a free flap hinge at the hub with a single blade, the motion of which is suppressed at integer multiples of the Nb per rev, where Nb is the number of blades. This is accomplished by introducing harmonics of the joint rotation angle as additional trim variables. Rotor frequencies are examined for a three-bladed gimbaled rotor model and are found to combine the modes of both free flap hinge and fixed flap hinge (cantilevered) one-bladed rotor models, similar to the behavior of a teetering rotor, justifying the gimbal modeling methodology. Gimbal flapping is examined for a proprotor in edgewise flight: successfully suppression of steady and 3/rev flapping indicates the gimbal model is performing as intended. The 6/rev and higher harmonics are negligible to begin with for this case and are left uncontrolled, but the same method can also be applied to suppress them if required. Airloads and blade loads are examined with a free flap hinge blade with the gimbal model, revealing an increase in 3/rev sectional normal aerodynamic force on the blade but a decrease in 3/rev flap and lag bending moment when the gimbal is added. Three-dimensional axial stress distribution at the blade root is examined, revealing only minor differences between the two hub rotor models, primarily in the flexbeam.
ABSTRACT A model-scale, coaxial, counter-rotating rotor system with single-bladed rotors was tested in hover and compared to a comprehensive model developed in CAMRAD II. Measurements included vibratory hub and pitch link loads, as well as three-dimensional lower rotor blade deformations extracted using digital image correlation. The model flap dynamics were validated using a rotating frame modal extraction technique based on a modified Ibrahim Time Domain method. The CAMRAD model successfully predicted unsteady loads and deformations for the isolated lower rotor operated with cyclic pitch in hover. To investigate transient loads, measurements were taken in the coaxial configuration at a blade loading coefficient of 0.10 and for an isolated lower rotor at equivalent blade loading. The CAMRAD model accurately predicted the unsteady interaction thrust, as well as the blade flapping which was found to increase after upper-lower rotor blade passage. The CAMRAD model revealed aerodynamic forcing in the coaxial configuration consistent with vortex interaction of the upper and lower rotors, while higher harmonics observed in the experimental data were attributed to impulsive loading due to blade thickness effects.
ABSTRACT This paper focused on improving a small-scaled prototype of a helicopter rotor blade with a flap-driving mechanism termed as the Seoul National University Flap (SNUF). The design of SNUF included realizing vibratory load reduction. First, a multibody structural dynamics analysis was performed to determine the influence of the flap dimension and location within the rotor blade with respect to hub vibratory load reduction. This process selected a specific blade configuration that maximized vibration reduction capability. Then, a numerical optimization technique was applied to improve the cross-sectional design of the SNUF blade. The design optimization procedure established improved blade sectional design with desired first torsional frequency and reduced blade weight while satisfying sufficient structural integrity. Three-dimensional nonlinear static structural analysis was also performed for the optimized SNUF design. Von Mises stress distribution on the blade and components were predicted by considering external aerodynamic loads, centrifugal loads due to rotation, and contact among the internal components. Future studies will include fabricating the prototype blade based on the optimized design and performing a whirl tower test.
ABSTRACT An aeromechanics analysis was conducted of a large-winged, single main rotor, compound helicopter modified from the AH-56 Cheyenne, in cruise and high-speed flight (250 knots) at sea level and high altitude (20,000 ft.) conditions. Performance and representative loads were evaluated with the comprehensive code RCAS to show the effect of compound configuration decisions. Suitability of the analysis for high advance ratio predictions was demonstrated through comparison to the UH-60A slowed rotor test data, and validation of compound performance prediction was shown with AH-56 Cheyenne test data. An assessment of the role of compound configuration, collective setting, wing pitch, rotor speed, altitude and trim control strategy on performance and loads was made. The study shows how reducing collective, for a constant wing pitch, is beneficial for peak L=De and reducing loads. Increasing wing pitch, at a constant collective, improves peak L=De, but can reduce L=De at high airspeeds and limit maximum airspeed. A 30% reduction in rotor speed achieved significant performance increases, while larger reductions limited high airspeed and high altitude trim due to stall. High altitude flight improves L=De at all airspeeds if stall on the prop rotor can be prevented. Ailerons and horizontal stabilizer trim control does not adversely affect L=De, but free rotor cyclics to trim for reduced flap bending loads. Interference effects from the rotor on the wing reduced peak L=De, but the effect is marginal above 170 knots. The final compound configuration achieved a peak L=De = 7.6 around 140 knots and L=De > 5.0 at the maximum airspeed considered of 250 knots. The static control derivatives for the compound showed less authority than a conventional helicopter, however increasing control authority is shown across the flight envelope.
ABSTRACT This paper studies the blade loads and whirl-flutter stability of a three bladed stiff-inplane tiltrotor wind tunnel model mounted on the Wing and Rotor Aeroelastic Test System (WRATS). Proprotor loads are predicted and compared with the WRATS model at different pylon conversion angles. The tiltrotor whirl flutter stability is predicted and tested in airplane mode. The analytical models are developed with three rotorcraft codes: RCAS, CAMRAD II, and Dymore. The Dymore and RCAS models contain the structural model of the wing/pylon fixed system whereas the CAMRAD II model for this study uses an isolated proprotor. The study indicates that the analytical models capture the overall vibratory flap and chord bending moments, however, miss the higher harmonics of the blade loads. The steady flap and chord bending moments predicted by CAMRAD II and Dymore differ in the torque tube and flexbeam region but agree along the span of the blade. RCAS shows a good prediction of the damping ratio of the chord and torsion modes of the WRATS stiff-inplane tiltrotor. Dymore captures the relative trends of these modal dampings but over-predicts the wing chord mode damping and under-predicts the wing torsion mode damping. Both analyses slightly over-predict the wing chord frequency.
ABSTRACT This study examines the effect of rotor blade elastic deformations on a quadcopter in forward flight conditions. The blade equations are discretized using the Galerkin method and the blade periodic response is calculated using the harmonic balance method. Simulations are conducted on a 2 kg quadcopter with 12 inch diameter two-bladed rotors. The blade root vertical shear, flap bending moment and drag shear showed a strong 1/rev variation due to the azimuthal variation in aerodynamic loads. Elastic blade deformations did not affect the aerodynamic loads but the addition of in-phase 1/rev inertial loads resulted in net increases of 28%, 36% and 48% in the 1/rev blade root vertical shear, flap bending moment, and drag shear, respectively, at a forward flight speed of 10 m/s. The in-plane elastic deformations further introduced a 1/rev blade root radial shear due to the radial Coriolis force. At 10 m/s forward flight speed, accounting for elastic blade deformations resulted in a 7.6% reduction in the steady hub drag force, a 37% increase in steady hub roll moment, and a 15% increase in steady hub pitching moment that requires a 12% higher pitch RPM control input to trim the quadcopter. Increases of 103% in the 2/rev rotor hub in-plane loads (drag and side force), of 12% in the 2/rev hub vertical force, and of 25% in the 2/rev hub pitching and rolling moments were also observed. Since the forward rotor, lateral rotors, and the rear rotor operate at distinct rotational speeds, the aircraft is subject to vibratory loading at three different frequencies resulting in a beating phenomenon. The maximum amplitude of these oscillations equals the sum of the amplitudes of the three constituent vibrations, but reaches its maximum only once every twelve cycles. The 2/rev rotor hub vibratory moments are the dominant contributors to the aircraft pitch and roll vibrations, while the 2/rev rotor hub in-plane forces are the dominant contributors to the aircraft vibratory yaw moments. The blade phasing between the two lateral rotors has an effect on the magnitude of the total roll vibrations on the aircraft.
A suite of computational tools capable of predicting in-plane low frequency rotorcraft noise and its control using active flaps operating in closed-loop mode is developed. A combined code, consisting of AVINOR, a comprehensive code for active control of vibration and noise, and an acoustic code called HELINOIR, is first validated against wind tunnel tests and subsequently verified against computational results. The in-plane noise reduction obtained with a single 20% chord plain trailing edge flap, for a rotor configuration resembling the MBB BO-105 during level flight at a moderate advance ratio is studied. A combination of various far-field and near-field feedback microphone locations was examined. A near-field microphone located on a hypothetical left boom position was found to be the best location and produced an in-plane noise reduction of up to 6 dB. However, this noise reduction was accompanied by an out-of-plane noise increase of 18 dB and a 60% increase in the vertical hub shear. These penalties can be partially alleviated by using a dual plain flap configuration.
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