Browse Topic: Flaps
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
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
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
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
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
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
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
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 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 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
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
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.
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
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
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
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
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
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
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
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
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
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
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 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
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
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
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
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
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
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
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
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
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
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
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
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