Showing papers on "Helicopter rotor published in 1996"

Development of a piezoelectric servoflap for helicopter rotor control

Abstract: A servoflap that uses a piezoelectric bender to deflect a trailing edge flap for use on a helicopter rotor blade was designed, built, and tested. This servoflap design is an improvement over a design developed previously at MIT. The design utilizes a new flexure mechanism to connect the piezoelectric bender to the control surface. The efficiency of the bender was improved by tapering its thickness with length. Also, the authority of the actuator was increased by implementing a nonlinear circuit to control the applied electric field, allowing a greater range of actuator voltages. Experiments were carried out on a bench test article to determine the frequency response of the actuator, as well as hinge moment and displacement capabilities. Flap deflections of or more were demonstrated while operating under no-load conditions at frequencies up to 100 Hz. The data indicate that, if properly scaled, the actuator will produce flap deflections greater than at the 90% span location on a full-scale helicopter. In addition, the first mode of the actuator was at frequency of the target model rotor. Proper inertial scaling of this actuator could raise this modal frequency to greater than on an operational helicopter, which is adequate for most rotor control purposes. A linear state space model of the actuator was derived. Comparisons of this model with the experimental data highlighted a number of mild nonlinearities in the actuator's response. However, the agreement between the experiment and analysis indicate that the model is a valid tool for predicting actuator performance.

Induced strain actuation of composite beams and rotor blades with embedded piezoceramic elements

Abstract: The objective of this research is to develop a dynamically-scaled (Froude scale) helicopter rotor blade with embedded piezoceramic elements as sensors and actuators to control blade vibrations. A 6 ft diameter 2-bladed bearingless rotor model was built where each blade is embedded with banks of piezoelectric actuators at degree angles with respect to the beam axis on the top and bottom surfaces. A twist distribution along the blade span is achieved through in-phase excitation of the top and bottom actuators at equal potentials, while a bending distribution is achieved through out-of-phase excitation. In order to fix design variables and to optimize blade performance, a uniform strain beam theory is formulated to analytically predict the static bending and torsional response of composite rectangular beams with embedded piezoelectric actuators. Parameters such as bond thicknesses, actuator skew angle and actuator spacing are investigated by experiments and then validated by theory. The static bending and torsional response of the rotor blades is experimentally measured and correlated with theory. Dynamic torsional and bending responses are experimentally determined for frequencies from 2 - 120 Hz to assess the viability of a vibration reduction system based on piezo-actuation of blade twist. To assess the performance of the piezo-actuators in rotation, hover tests were conducted where accelerometers embedded in the blades were used to resolve the tip twist amplitudes. Although the magnitudes of blade twist attained in this experiment were small, it is expected that future models can be built with improved performance.

Helicopter rotor blade computation in unsteady flows using moving overset grids

Abstract: An overset grid thin-layer Navier-Stokes code has been extended to include dynamic motion of helicopter rotor blades through relative grid motion. The unsteady flowfield and airloads on an AH-IG rotor in forward flight were computed to verify the methodology and to demonstrate the method's potential usefulness towards comprehensive helicopter codes. In addition, the method uses the blade's first harmonics measured in the flight test to prescribe the blade motion. The solution was impulsively started and became periodic in less than three rotor revolutions. Detailed unsteady numerical flow visualization techniques were applied to the entire unsteady data set of five rotor revolutions and exhibited flowfield features such as blade vortex interaction and wake roll-up. The unsteady blade loads and surface pressures compare well against those from flight measurements. Details of the method, a discussion of the resulting predicted flowfield, and requirements for future work are presented. Overall, given the proper blade dynamics, this method can compute the unsteady flowfield of a general helicopter rotor in forward flight.

Transient Response Technique Applied to Active Magnetic Bearing Machinery During Rotor Drop

Abstract: The active magnetic bearing (AMB) is a relatively new technology which has many advantages compared with conventional bearing design. In an AMB system. the rolling-element back-up bearings are indispensable to protect the magnetic bearing rotor and stator, and other stationary seals along the rotor shqft. In this paper, a theoretical formulation is proposed and solved numerically to examine the transient response of the flexible rotor, from the time just previous to when the AMB shuts down and including the rotor drop onto the back-up bearing. The backward whirl of the rotor, which may lead to the destructive damage of the machinery, has been analytically predicted at very light support damping and very high support damping. Also, the vibration due to the nonlinearity of the contact point geometry has been included in the analysis. The influence of the support damping on the displacement of the disk and also the contact force between the journal and the inner-race of the back-up bearing have been computed for various rotor system parameters. By comparing these results with the optimum support damping for the simple flexible rotor model, it is shown that this support damping optimization can be applicable for specifying the required optimum range of support damping for the back-up bearings of AMB systems.

Applications of torsional shape memory alloy actuators for active rotor blade control: opportunities and limitations

Abstract: A shape memory alloy rod, if it is pre-twisted to create stress induced martensitic transformation, will exhibit a torsional shape memory effect, namely, some of the pre-twist will be recovered once the SMA rod is heated above its austenitic start temperature. If the pre- twisted SMA rod is constrained, a large recovery torque will be generated. Torsional shape memory alloy actuators have broad applications. This paper focuses on their applications of active helicopter rotor blade control. In this paper, we first review the effort of active rotor blade control, introduce the basic properties of shape memory alloy actuators, present the important characteristics of torsional SMA actuators governing their applications, and address issues involved in the use of torsional SMA actuators for automatic rotor blade tracking systems.

Helicopter rotor equipped with flaps

Abstract: A blade of the rotor is attached to a main rotor shaft, and is rotated counterclockwise at a high speed when viewed from above. In the vicinity of the blade tip, a flap having a wing shape in section is pivotally supported with a hinge so as to be angularly displaceable and at the end a part of the rear end of the blade is cut out. A hinge line which indicates the center of the angular displacement of the hinge is formed so as to be parallel to a straight line passing through the rotation center of the rotor and the center of gravity of the flap. A helicopter rotor equipped with flaps is provided in which the influence of the centrifugal force acting on the center of gravity of a blade can be largely reduced, and a flap control mechanism can be reduced in size and weight.

Application of passive dampers to modern helicopters

Abstract: Boeing Helicopters has been actively testing elastomeric and dampers for application on modern rotors. Damper bench tests have been performed on various damper configurations to understand their dynamic characteristics. A recently completed series of wind tunnel tests was performed on a 1/6-Froude-scale model of the Comanche bearingless main rotor system with both elastomeric and dampers. Elastomeric dampers have also been incorporated in the full-scale Model 360 articulated rotor design. This paper summarizes elastomeric and damper activities at Boeing, and discusses issues that must be considered in the design and analysis of such systems.

Advances in the development of an intelligent helicopter rotor employing smart trailing-edge flaps

Abstract: Significant advances in the development of a Froude scaled helicopter rotor model featuring a trailing-edge flap driven by piezoceramic bimorph actuators for active vibration suppression are discussed. A quasisteady aerodynamic analysis used to determine flap size and actuator requirements is presented. The block force and stroke of the current actuators are evaluated using two theories and compared with experimental results. The dynamic performance of the actuator as well as the actuator - flap assembly is examined. Earlier hover tests showed severe degradation in flap deflections with increasing rotor speed, and flap deflections were too small to be effectively utilized for significant vibration control. To investigate the causes of the performance degradation, new blades are constructed and tested in vacuo to isolate the effects of centrifugal loading on the actuator - flap system. A beam model of the piezo bimorph including propeller moment effects is formulated to better illustrate the physical mechanisms affecting the system in a rotating environment. The cause of the reduced deflections is traced to frictional forces created at the junction where the flap is supported during rotation of the blades. The use of a thrust bearing was found to alleviate this problem and subsequent hover tests showed a dramatic increase in flap deflection at high excitation frequencies.

Aeroelastic optimization of a helicopter rotor with two-cell composite blades

Abstract: Aeroelastic and sensitivity analyses of the rotor based on a finite element in space and time are linked to an automated optimization algorithm to perform optimization studies for a four-bladed, soft in-plane composite rotor consisting of a two-cell thin-walled beam. The design variables used in this study are the ply angles of the laminated walls of the composite beam. The objective function minimizes the 4/rev hub loads, with constraints on blade frequencies and aeroelastic stability in forward flight. Optimum design solutions show a reduction in the objective function of about 20% due to elastic stiffnesses and an additional 13% due to composite couplings. Starting from an initially infeasible design, the optimum design solution with negative lag bending-torsion coupling results in an increase in lag mode damping of about 140% compared to the baseline layup. c* c* c* ^d-> ^/> ^m CT c

Half-plow vortex generators for rotorcraft blades for reducing blade-vortex interaction noise

Abstract: In one embodiment for a helicopter main rotor assembly, a half-plow vortex generator is mounted in combination with the upper aerodynamic surface of each main rotor blade and is operative to generate a primary corotating vortex of sufficient strength to interact with and accelerate the dissipation of the tip vortex generated by the same main rotor blade, thereby reducing blade-vortex interaction noise radiating from the helicopter main rotor assembly. The half-plow vortex generator has a right triangular planform configuration defined by a length, a width, and an apex angle. The three-dimensional configuration of the vortex generator is further defined by an apex height. The apex height is the primary determinant of the strength of the generated primary corotating vortex and is defined in terms of the thickness of the main rotor blade at the local chord where the vortex generator is mounted. The apex height may be approximately equal to the local thickness, but preferably has a magnitude within the range of about one-eighth to about three-quarters of the local thickness. The length, width, and apex angle are secondary determinants of the strength of the primary corotating vortex generated by the half-plow vortex generator. The length and width of the vortex generator are defined in terms of the tip chord length of the main rotor blade, the length preferably having a magnitude within the range of about one-fourth to one-half of the tip chord length and the width preferably having a value of about one-third of the length of the vortex generator. The apex angle preferably has a value within the range of about twenty to about thirty degrees. The mounting site for the half-plow vortex generator is defined in terms of the length of the local and tip chords, respectively. The vortex generator is mounted inboardly from the tip of the main rotor or blade a spanwise distance having a magnitude preferably within the range of about one-half of to about equal to the tip chord length in substantial alignment with the local chord. The apex of the vortex generator is mounted inwardly from the leading edge of the main rotor blade by a chordal distance having a magnitude of about one-quarter the local chord length.

Application of smart materials to control of a helicopter rotor

Abstract: Smart material actuator technology for operation `on the blade' is now becoming available and has the promise to overcome the size, weight and complexity issues of hydraulic and electric on-rotor actuation. However, the challenges of the limited output capability of the materials and the dynamic operating environment must be fully addressed and resolved. The present study covers the conceptual sizing and design of a full scale demonstration system to provide active control of noise and vibrations as well as inflight blade tracking for the MD-900 helicopter. Active control is achieved via a trailing edge flap and trim tab, both driven by on- blade smart material actuators. Overall, this ARPA sponsored program entails the design, development, and whirl tower testing of the full scale active control rotor system. If successful, an entry in the NASA Ames 40 X 80 foot wind tunnel and flight tests are planned for a follow on program.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Aeroelastic optimization of a helicopter rotor to reduce vibration and dynamic stresses

Abstract: Optimization studies are carried out for a four-bladed, soft in-plane hingeless rotor consisting of a twocell composite box-beam spar. The design variables are the ply angles of the box-beam walls. The objective functions are the vibratory hub loads and the vibratory blade bending moments; constraints are imposed on blade rotating frequencies and aeroelastic stability. The objective functions are first minimized individually, and then a combined optimization is performed to minimize both the objectives simultaneously. As compared to the starting design, the optimum solution results in a 15-60% reduction of the 4/rev hub loads as well as a reduction in the peak-to-peak flap and lag bending moments of 11 and 14%, respectively. Starting from an initially infeasible starting design with a 3% requirement on lag mode damping, the optimum solution with composite chordwise bending-torsion coupling results in an increase in lag mode damping of over 200% compared to the starting design.

Aeroelastic Analysis of Helicopter Rotor Blades Incorporating Anisotropic Piezoelectric Twist Actuation

Abstract: A simple aeroelastic analysis of a helicopter rotor blade incorporating embedded piezoelectric fiber composite, interdigitated electrode blade twist actuators is described. The analysis consists of a linear torsion and flapwise bending model coupled with a nonlinear ONERA based unsteady aerodynamics model. A modified Galerkin procedure is performed upon the rotor blade partial differential equations of motion to develop a system of ordinary differential equations suitable for dynamics simulation using numerical integration. The twist actuation responses for three conceptual full-scale blade designs with realistic constraints on blade mass are numerically evaluated using the analysis. Numerical results indicate that useful amplitudes of nonresonant elastic twist, on the order of one to two degrees, are achievable under one-g hovering flight conditions for interdigitated electrode poling configurations. Twist actuation for the interdigitated electrode blades is also compared with the twist actuation of a conventionally poled piezoelectric fiber composite blade. Elastic twist produced using the interdigitated electrode actuators was found to be four to five times larger than that obtained with the conventionally poled actuators.

Main rotor system for model helicopters

Abstract: A main rotor is provided for use on a model helicopter. The main rotor includes rotor blades having an inboard section and an outboard section. The inboard section includes a higher camber, thin, wide airfoils, and greater pitch and camber relative to the outboard section. The outboard section includes an average chord length that is less than the chord length of the inboard section, The rotor blades also include a blade body that is semiflexible to bend during a crash landing of the model helicopter.

An Efficient and Robust Method for Predicting Helicopter Rotor High-Speed Impulsive Noise

Abstract: A new formulation for the Ffowcs Williams-Hawkings quadrupole source, which is valid for a far-field in-plane observer, is presented. The far-field approximation is new and unique in that no further approximation of the quadrupole source strength is made and integrands with r to the -2 and r to the -3 dependence are retained. This paper focuses on the development of a retarded-time formulation in which time derivatives are analytically taken inside the integrals to avoid unnecessary computational work when the observer moves with the rotor. The new quadrupole formulation is similar to Farassat''s thickness and loading formulation 1A. Quadrupole noise prediction is carried out in two parts: a preprocessing stage in which the previously computed flow field is integrated in the direction normal to the rotor disk, and a noise computation stage in which quadrupole surface integrals are evaluated for a particular observer position. Preliminary predictions for hover and forward flight agree well with experimental data. The method is robust and requires computer resources comparable to thickness and loading noise prediction.

Aeroelastic Optimization of an Advanced Geometry Helicopter Rotor

Abstract: Sensitivity derivatives of blade loads and aeroelastic stability of a helicopter rotor in forward flight are calculated as an integral part of a basic aeroelastic analysis using a direct analytical approach. Design variables include nonstructural mass and its placement, chordwise offset of blade center of gravity and aerodynamic center from the elastic axis, blade bending stiffnesses (flap, lag, torsion), and tip geometry (sweep, anhedral, pretwist and planform taper). By means of a sensitivity study, the importance of different design variables on oscillatory hub loads and damping of blade modes is examined. Aeroelastic and sensitivity analyses of the rotor based on a finite element method in space and time are linked with automated optimization algorithms to perform optimization studies of rotor blades. Optimum design solutions, calculated for a four-bladed, soft-inplane hingeless rotor achieved a reduction of 25-60 percent of all 4/rev loads.

Design and testing of a helicopter trailing edge flap with piezoelectric stack actuators

Abstract: This paper presents the results of research in using piezoelectric stacks to power a trailing edge flap for use as an individual blade control device to suppress vibrations in helicopter rotor blades. A flap actuator was designed and constructed using two piezoelectric stacks arranged in series. Mechanical amplification was provided through a hinged L-arm. An analytical model was formulated to calculate the predicted flap deflection using quasi-steady aerodynamics. The model showed that the high force piezoelectric stacks were capable of maintaining flap deflection over a wide range of freestream velocities. An experimental model consisting of an 8 inch chord NACA 0012 airfoil was constructed incorporating a single 4-inch span, 1.6 inch chord trailing edge flap. The flap performance of the model was tested in no load conditions and in the wind tunnel under freestream velocities of up to 118 ft/sec. The flap deflections were lower for higher velocities and below analytical predictions.© (1996) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

Design of a servo-flap rotor for reduced control loads

Abstract: The results of a study to conceptually define a servo-flap rotor with on-blade smart material actuators for primary and active control are presented. Selection of the flap and rotor design parameters is directed to minimize the control effort. Simulation results for an AH-64 size aircraft show that substantial reductions in control effort as well as main rotor power can be achieved through judicious selection of the rotor and flap design parameters and introduction of adaptive rotor features.

A helicopter with adaptive rotor blades for collective control

Abstract: In this paper, we discuss the development of a helicopter rotor system which employs an adaptive rotor blade. The aspect of this rotor system which is new and unique is the replacement of the traditional swash-plate-based collective control with adaptive rotor blades which generate lift by using shape-memory-alloy-based active camber control. The initial work produced a fixed magnitude chamber change in an on/off fashion. Additional work has been done to develop a control scheme which allows a smooth variation of camber changes.

Solid-State Actuation of Rotor Blade Servo-Flap for Active Vibration Control

Abstract: The basic concepts and some innovative ideas associated with the analysis, design, and experimentation of induced-strain actuators for rotor blade aeroelastic vibration control are presented and discussed. A trailing-edge servo-flap actuated by a hydraulically-amplified large-displacement induced-strain actuator is considered. Design requirements based on extensive literature review are proposed and utilized. The principle of high-power induced-strain actuation, and the energy and energy-density of several commercially-available induced-strain actuators are presented. A full-scale proof-of-concept demonstrator, designation HAHDIS Mk. 1, was designed, built, and tested. Results from static and dynamic tests are presented and discussed. Good frequency response of the full-scale proof-of-concept demonstrator was proven in the range of 1 to 30 Hz in spite of the power and current limitations of the available electronic equipment.

Aerodynamic aspects of blade-vortex interaction (BVI)

Abstract: This paper presents a review of the blade-vortex interaction (BVI) phenomenon. It covers both analytical and experimental studies of two-dimesional parallel airfoil-vortex interaction, three dimensional BVIs, and helicopter rotor BVIs. The important parameters that affect the BVI airloads are discussed.

ACSR system for vibration suppression in coupled helicopter rotor/flexible fuselage model

Abstract: The paper presents an approach for the active control of structural response (ACSR) and its implemention in a system for vibration reduction in a helicopter fuselage. Vibrations in the fuselage are simulated analytically using a recently derived coupled rotor/flexible fuselage aeroelastic response model. Baseline fuselage vibrations, are sensed as vibratory components of the acceleration at specific fuselage locations. These quantities are used to generate signals to four servo actuators located at the four corners of the ACSR platform. The servo actuators produce controlled vibratory forces, which when combined with the vibratory rotor loads produce vibration levels at various fuselage locations which are much lower than the baseline vibrations. The vibration levels, in the presence of the controller, display very substantial vibration reduction. This vibration suppression system, operates exclusively in the hub fixed nonrotating system, and therefore has no influence on vehicle airworthiness.