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Showing papers on "Helicopter rotor published in 1994"


01 Sep 1994
TL;DR: In this article, a comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted Testing the wing in the nonrotating condition isolates the three-dimensional (3D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment.
Abstract: A comprehensive experimental investigation of the pressure distribution over a semispan wing undergoing pitching motions representative of a helicopter rotor blade was conducted Testing the wing in the nonrotating condition isolates the three-dimensional (3-D) blade aerodynamic and dynamic stall characteristics from the complications of the rotor blade environment The test has generated a very complete, detailed, and accurate body of data These data include static and dynamic pressure distributions, surface flow visualizations, two-dimensional (2-D) airfoil data from the same model and installation, and important supporting blockage and wall pressure distributions This body of data is sufficiently comprehensive and accurate that it can be used for the validation of rotor blade aerodynamic models over a broad range of the important parameters including 3-D dynamic stall This data report presents all the cycle-averaged lift, drag, and pitching moment coefficient data versus angle of attack obtained from the instantaneous pressure data for the 3-D wing and the 2-D airfoil Also presented are examples of the following: cycle-to-cycle variations occurring for incipient or lightly stalled conditions; 3-D surface flow visualizations; supporting blockage and wall pressure distributions; and underlying detailed pressure results

125 citations


01 Jun 1994
TL;DR: This report describes an analytical study of vibration reduction in a four-bladed helicopter rotor using an actively controlled, partial span, trailing edge flap located on the blade, which clearly demonstrates the feasibility of this new approach to vibration reduction.
Abstract: This report describes an analytical study of vibration reduction in a four-bladed helicopter rotor using an actively controlled, partial span, trailing edge flap located on the blade. The vibration reduction produced by the actively controlled flap (ACF) is compared with that obtained using individual blade control (IBC), in which the entire blade is oscillated in pitch. For both cases a deterministic feedback controller is implemented to reduce the 4/rev hub loads. For all cases considered, the ACF produced vibration reduction comparable with that obtained using IBC, but consumed only 10-30% of the power required to implement IBC. A careful parametric study is conducted to determine the influence of blade torsional stiffness, spanwise location of the control flap, and hinge moment correction on the vibration reduction characteristics of the ACF. The results clearly demonstrate the feasibility of this new approach to vibration reduction. It should be emphasized than the ACF, used together with a conventional swashplate, is completely decoupled from the primary flight control system and thus it has no influence on the airworthiness of the helicopter. This attribute is potentially a significant advantage when compared to IBC.

109 citations


01 Jan 1994
TL;DR: A wind tunnel test was conducted with a full-scale BO 105 helicopter rotor to evaluate the potential of open-loop individual blade control (IBC) to improve rotor performance, to reduce blade vortex interaction (BVI) noise, and to alleviate helicopter vibrations as mentioned in this paper.
Abstract: A wind tunnel test was conducted with a full-scale BO 105 helicopter rotor to evaluate the potential of open-loop individual blade control (IBC) to improve rotor performance, to reduce blade vortex interaction (BVI) noise, and to alleviate helicopter vibrations. The wind tunnel test was an international collaborative effort between NASA/U.S. Army AFDD, ZF Luftfahrttechnik, Eurocopter Deutschland, and the German Aerospace Laboratory (DLR) and was conducted under the auspices of the U.S./German MOU on Rotorcraft Aeromechanics. In this test the normal blade pitch links of the rotor were replaced by servo-actuators so that the pitch of each blade could be controlled independently of the other blades. The specially designed servoactuators and IBC control system were designed and manufactured by ZF Luftfahrttechnik, GmbH. The wind tunnel test was conducted in the 40- by 80-Foot Wind Tunnel at the NASA Ames Research Center. An extensive amount of measurement information was acquired for each IBC data point. These data include rotor performance, static and dynamic hub forces and moments, rotor loads, control loads, inboard and outboard blade pitch motion, and BVI noise data. The data indicated very significant (80 percent) simultaneous reductions in both BVI noise and hub vibrations could be obtained using multi-harmonic input at the critical descent (terminal approach) condition. The data also showed that performance improvements of up to 7 percent could be obtained using 2P input at high-speed forward flight conditions.

76 citations



Journal ArticleDOI
TL;DR: In this paper, the transonic blade vortex interaction (BVI) near-field flow is simulated using the rotating Kirchhoff method for the extension to the acoustic far field.
Abstract: The unsteady transonic full-potential rotor (FPR) code is used for the simulation of the transonic blade vortex interaction (BVI) near-field flow. The rotating Kirchhoff method is used for the extension to the acoustic far field. Two rotating Kirchhoff formulations are developed for the three-dimensional BVI far-field noise prediction. The first formulation (Morino's method) is for an observer rotating with the blade. This allows the direct comparison with computational fluid dynamics results. The second formulation (Farassat's method) is for a stationary observer and allows a direct comparison with acoustic experiments

43 citations


Journal ArticleDOI
TL;DR: In this paper, a method for solving the minimum-induced-loss (MIL) rotor design problem is described. But the necessary conditions are extended to include constraints on the lift.
Abstract: A method is described for solving the minimum-induced loss (MIL) rotor design problem. First, the generalized Betz condition for MIL rotors is developed. Because the resulting lift distributions would generally exceed the maximum blade lift coefficient on the retreating side of the rotor, the necessary conditions are extended to include constraints on the lift. A method for solving for the optimum lift distribution using finite elements is described. Numerical results are presented for a typical rotor in forward flight. The MIL rotor may have on the order of 10% less induced power loss than a typical unoptimized rotor.

36 citations


Patent
21 Oct 1994
TL;DR: In this article, an end plate (29) on an upper damper with a lug (36) positioned adjacent to a lug(37) on the torque tube was used to adjust the end plate of the damper relative to the longitudinal axis of the blade.
Abstract: A helicopter rotor (1) has two or more blades which extend from a hub (5) and has adjustment means for adjusting the lead/lag variations of the blades which typically occur due to manufacturing tolerance variations or to changes which occur through operation such as softening of the supporting elastomer dampers. One embodiment of the invention incorporates an end plate (29) on an upper damper (23) with a lug (36) positioned adjacent to a lug (37) on the torque tube (10) such that the end plate of the damper can be changed relative to the longitudinal axis of the blade. For example, a +/- 1/2 degree adjustment can be made in the blade angle relative to the hub to insure that lead/lag errors are reduced. In another embodiment, the damper set includes extended metal shims (49a, 49b) adjacent to the end most laminate plys of the upper and lower dampers such that these shims can be bolted to the end plates of the snubber assemblies for step increasing the stiffness of the assemblies. Utilizing the apparatus of the invention, each blade of the rotor can be adjusted to balance the rotor assembly to minimize vibration.

32 citations


Proceedings ArticleDOI
06 May 1994
TL;DR: In this article, a rotor blade with embedded piezoceramic elements as sensors and actuators to control blade vibrations was developed. But the results were limited to a 6-ft-diameter two-bladed bearingless rotor model.
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 two-bladed bearingless rotor model was built, where each blade is embedded with banks of piezoelectric actuators at +/- 45-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 piezoactuation of blade twist. Although the magnitudes of blade twist attained in this experiment were small, it is expected that future models can be built with improved performance.© (1994) COPYRIGHT SPIE--The International Society for Optical Engineering. Downloading of the abstract is permitted for personal use only.

28 citations


Journal ArticleDOI
TL;DR: In this paper, the authors considered the possibility of identifying changes in the residual unbalance of a multi-bearing rotor system, where the relative motion of the rotor journals with respect to the bearings can be measured before and after the change in unbalance take place.

27 citations


Journal ArticleDOI
TL;DR: In this paper, a noise study using an aeroelastically scaled BO-105 rotor was conducted in the German-Dutch wind tunnel to examine the use of higher harmonic control (HHC) of blade pitch to reduce impulsive blade-vortex interaction (BVI) noise.
Abstract: A noise study using an aeroelastically scaled BO-105 rotor was conducted in the German-Dutch Wind Tunnel to examine the use of higher harmonic control (HHC) of blade pitch to reduce impulsive blade-vortex interaction (BVI) noise. The noise directivity was measured over a large plane underneath the rotor using a traversing inflow microphone array. Noise and vibration measurements were made for a range of matched rotor operating conditions where prescribed (or open loop) HHC pitch, at various amplitudes and phases, was superimposed on normal (baseline) collective and cyclic trim pitch. Acoustic data are presented for 3, 4, and 5P HHC applied to a typical landing approach rotor operating condition where BVI noise is normally intense. Noise reductions of up to 6 dB were found for the advancing side BVI noise radiating upstream of the rotor, and also for the retreating side BVI noise radiating below and downstream of the rotor. The relative levels between the sides were modified by HHC control phase. To help give insight to the physics of the HHC/BVI noise problem, highresolution loading and noise prediction results are presented for comparison to the data. The predictions are based on a new high-resolution version of the CAMRAD rotor performance program under development at Langley, called HIRES.

26 citations


Journal ArticleDOI
TL;DR: In this article, a theoretical solution to the problem of determining the radar cross section and Doppler spectrum of a helicopter rotor as presented to a radar operating in the decametric wave band (3-30 MHz) is given.
Abstract: The paper gives a theoretical solution to the problem of determining the radar cross section and Doppler spectrum of a helicopter rotor as presented to a radar operating in the decametric wave band (3-30 MHz). At such frequencies for all practical rotors the scattering regime is either Rayleighan or resonant. The solution proceeds by modeling the rotor as an equivalent set of radial wires on which the incident wave is assumed to excite sinusoidally distributed currents. It is shown that, subject to certain simplifying assumptions, the Doppler spectrum has a form similar to that associated with tone modulation of a frequency modulated bearer. The theoretical work is confirmed by experiment. >

01 Oct 1994
TL;DR: In this article, an advanced higher harmonic control (HHC) analysis was developed and applied to investigate its effect on vibration reduction levels, blade and control system fatigue loads, rotor performance, and power requirements of servo-actuators.
Abstract: An advanced higher harmonic control (HHC) analysis has been developed and applied to investigate its effect on vibration reduction levels, blade and control system fatigue loads, rotor performance, and power requirements of servo-actuators. The analysis is based on a finite element method in space and time. A nonlinear time domain unsteady aerodynamic model, based on the indicial response formulation, is used to calculate the airloads. The rotor induced inflow is computed using a free wake model. The vehicle trim controls and blade steady responses are solved as one coupled solution using a modified Newton method. A linear frequency-domain quasi-steady transfer matrix is used to relate the harmonics of the vibratory hub loads to the harmonics of the HHC inputs. Optimal HHC is calculated from the minimization of the vibratory hub loads expressed in term of a quadratic performance index. Predicted vibratory hub shears are correlated with wind tunnel data. The fixed-gain HHC controller suppresses completely the vibratory hub shears for most of steady or quasi-steady flight conditions. HHC actuator amplitudes and power increase significantly at high forward speeds (above 100 knots). Due to the applied HHC, the blade torsional stresses and control loads are increased substantially. For flight conditions where the blades are stalled considerably, the HHC input-output model is quite nonlinear. For such cases, the adaptive-gain controller is effective in suppressing vibratory hub loads, even though HHC may actually increase stall areas on the rotor disk. The fixed-gain controller performs poorly for such flight conditions. Comparison study of different rotor systems indicates that a soft-inplane hingeless rotor requires less actuator power at high speeds (above 130 knots) than an articulated rotor, and a stiff-inplane hingeless rotor generally requires more actuator power than an articulated or a soft-inplane hingeless rotor. Parametric studies for a hingeless rotor operating in a transition flight regime and for an articulated rotor operating at the level-flight boundary (high speed and high thrust conditions) indicate that blade parameters including flap, lag, torsion stiffness distributions, linear pretwist, chordwise offset of center-of-mass from elastic axis and chordwise offset of elastic axis from aerodynamic center can be selected to minimize the actuator power requirements for HHC.


Journal ArticleDOI
TL;DR: In this article, a Navier-Stokes inner region near the blade and wake surrounded by a full-potential outer region are coupled through boundary conditions, and the two regions are applied to a model UH-60A forward flight case.

Dissertation
01 Jan 1994
TL;DR: Hall et al. as mentioned in this paper presented an aeroelastic model of a smart rotor system, which incorporates blade-mounted trailing-edge flap actuators and conventional root pitch actuation in linear time invariant state space form.
Abstract: An aeroelastic model of a smart rotor system (which incorporates blade-mounted trailing-edge flap actuators and conventional root pitch actuation) is presented in linear time invariant state space form. The servo-flap deflections are modeled as producing incremental lift and moment variations, so that any linear aerodynamic actuator can be evaluated. This smart rotor model is used to conduct parametric studies involving rotor blade torsional stiffness, center-of-gravity offset, additional actuator mass, and actuator placement. Results on the effects of collective root pitch and servo-flap actuation on rotor thrust response are presented. Active rotor vibration reduction is demonstrated by applying higher harmonic control algorithms to the state space rotor model. Using reasonable servo-flap deflections, a rotor equipped with trailing-edge flaps can provide enough authority to cancel higher harmonic vibration. Thesis Supervisor: Steven R. Hall, Sc.D. Title: Associate Professor of Aeronautics and Astronautics

Journal ArticleDOI
TL;DR: In this paper, the effects of collective and cyclic pitch, as well as coning, cyclic flapping, and blade loading inputs are examined to determine the necessary spatial and temporal resolution.
Abstract: Rotor noise prediction codes predict the thickness and loading noise produced by a helicopter rotor, given the blade motion, rotor operating conditions, and fluctuating force distribution over the blade surface. However, the criticality of these various inputs, and their respective effects on the predicted acoustic field, have never been fully addressed. This paper examines the importance of these inputs, and the sensitivity of the acoustic predicitions to a variation of each parameter. The effects of collective and cyclic pitch, as well as coning and cyclic flapping, are presented. Blade loading inputs are examined to determine the necessary spatial and temporal resolution, as well as the importance of the chordwise distribution. The acoustic predictions show regions in the acoustic field where significant errors occur when simplified blade motions or blade loadings are used. An assessment of the variation in the predicted acoustic field is balanced by a consideration of Central Processing Unit (CPU) time necessary for the various approximations.

Journal ArticleDOI
TL;DR: In this article, a set of composite model rotor blades was manufactured from existing blade molds for a low-twist metal helicopter rotor blade, with a view toward establishing a preliminary proof concept for extensiontwist-coupled rotor blades.
Abstract: The purpose of this Note is to present results from an analytic/experimental study that investigated the potential for passively changing blade twist through the use of extension-twist coupling. A set of composite model rotor blades was manufactured from existing blade molds for a low-twist metal helicopter rotor blade, with a view toward establishing a preliminary proof concept for extension-twist-coupled rotor blades. Data were obtained in hover for both a ballasted and unballasted blade configuration in sea-level atmospheric conditions. Test data were compared with results obtained from a geometrically nonlinear analysis of a detailed finite element model of the rotor blade developed in MSC/NASTRAN.

Journal ArticleDOI
TL;DR: In this paper, the feasibility of employing adaptive material to build both sensors and actuators to attenuate the higher harmonic loads developed at the helicopter rotor blades using the individual blade control (IBC) concept is investigated.
Abstract: This article investigates the feasibility of employing adaptive material to build both sensors and actuators to attenuate the higher harmonic loads developed at the helicopter rotor blades using the individual blade control (IBC) concept. Both the first elastic flatwise bending (second for hingeless rotors) and the first elastic torsion modes of a single blade deserve special attention in the vibration control. Theoretical investigations, supported by wind-tunnel and flight tests, confirmed that these modes are responsible for the larger amplitude loads at 3/rev in four-blade hingeless rotors. This is a situation for which IBC, based on a collocated actuator-sensor arrangement along the blade, and tailored to act specifically on the bending and the torsion modes, is expected to bring further improvements to the reduction of the overall dynamic response of rotary wings. The results indicate that there are already real situations for which the adaptive material has enough power to accomplish the task without saturation of the applied electrical field.

01 Jan 1994
TL;DR: In 1989, a heavily instrumented sub-scale model of a helicopter main rotor was tested in the NASA LeRC Icing Research Tunnel (IRT) and the results of this series of tunnel tests were published previously.
Abstract: During two entries in late 1989, a heavily instrumented sub-scale model of a helicopter main rotor was tested in the NASA LeRC Icing Research Tunnel (IRT). The results of this series of tunnel tests were published previously. After studying the results from the 1989 test and comparing them to predictions, it became clear that certain test conditions still needed investigation. Therefore, a re-entry of the Sikorsky Aircraft Powered Force Model (PFM) in the IRT was instituted in order to expand upon the current rotor craft sub-scale model experimental database. The major areas of interest included expansion of the test matrix to include a larger number of points in the FAA AC 29-2 icing envelope, inclusion of a number of high power rotor performance points, close examination of warm temperature operations, operation of the model in constant lift mode, and testing for conditions for icing test points in the full scale helicopter database. The expanded database will allow further and more detailed examination and comparison with analytical models. Participants in the test were NASA LeRC, the U.S. Army Vehicle Propulsion Directorate based at LeRC, and Sikorsky Aircraft. The model rotor was exposed to a range of icing conditions (temperature, liquid water content, median droplet diameter) and was operated over ranges of shaft angle, rotor tip speed, advance ratio, and rotor lift. The data taken included blade strain gage and balance data, as well as still photography, video, ice profile tracings, and ice molds. A discussion of the details of the test is given herein. Also, a brief examination of a subset of the data taken is also given.

Journal ArticleDOI
TL;DR: In this article, a parametric study was performed to quantitatively describe how the vortex core size, the vortex location, and the hover-tip Mach number MH affect blade-vortex interaction (BVI) noise.
Abstract: A parametric study was performed to quantitatively describe how the vortex core size, the vortex location, and the hover-tip Mach number MH affect blade-vortex interaction (BVI) noise. The effects of these parameters on BVI noise are determined using a rotor acoustic prediction program. The acoustic prediction program is based on the well-known Ffowcs Williams and Hawkings equation for acoustic pressure, wherein noncompact monopole terms model rotor blade thickness and distributed dipoles model local blade surface pressure. The dipole strengths are determined by an unsteady, three-dimensional, full-potentia l rotor code that models the aerodynamic interactions between a nonlifting rotor and a tip vortex generated by an upstream wing. The acoustic pressures were calculated for several observer positions in regions of intense acoustic radiation for a variety of blade-vortex proximities, orientations, hover-tip Mach numbers, and vortex-core sizes. This study has quantified the sensitivity of BVI noise to the dominant vortex and aerodynamic parameters. The sound pressure level (SPL) falls off as the inverse of the square of the miss distance between the vortex core and the rotor blade when the miss distance is greater than the core radius. Increasing the core radius is not as effective as increasing the miss distance when attempting to reduce BVI noise. As expected, this study shows that SPL decreases with an increase in the obliqueness of interaction. The calculations performed in this work indicate that SPL increases approximately as Af|. These results can be used to guide future research of BVI noise reduction.

Journal ArticleDOI
TL;DR: In this article, the effects of blade and root-flexure elasticity and dynamic stall on the stability of hingeless rotor blades are investigated, and the dynamic stall description is based on the ONERA models of lift, drag, and pitching moment.
Abstract: The effects of blade and root-flexure elasticity and dynamic stall on the stability of hingeless rotor blades are investigated. The dynamic stall description is based on the ONERA models of lift, drag, and pitching moment. The structural analysis is based on three blade models that range from a rigid flap-lag model to two elastic flap-lag-torsion models, which differ in representing root-flexure elasticity. The predictions are correlated with the measured lag damping of an experimental isolated three-blade rotor; the correlation covers rotor operations from near-zero-thrust conditions in hover to highly stalled, high-thrust conditions in foward flight. That correlation shows sensitivity of lag-damping predictions to structural refinements in blade and root-flexure modeling. Moreover, this sensitivity increases with increasing control pitch angle and advance ratio. For high-advance-ratio and high-thrust conditions, inclusion of dynamic stall generally improves the correlation.

Journal ArticleDOI
TL;DR: This paper addresses the first significant step in the ongoing research towards the overall goal of developing an integrated dynamic, aerodynamic, and structural optimization methodology for helicopter rotor blades made of composite materials by investigating the feasibility and the utility of the structural-dynamic phase of the optimization task with realistic blade cross-sections.


Dissertation
24 Apr 1994
TL;DR: Hall et al. as discussed by the authors used a piezoelectric bender to fit a trailing edge servoap for a helicopter rotor blade, which was designed, built, and tested.
Abstract: An actuator using a piezoelectric bender to de ect a trailing edge servoap for use on a helicopter rotor blade was designed, built, and tested. This actuator is an improvement over one developed previously at MIT. The design utilizes a new exure mechanism to connect the piezoelectric bender to the control surface. The e ciency of the bender was improved by tapering its thickness properties with length. Also, implementation of a nonlinear circuit allowing the application of a greater range of actuator voltages increased the resultant strain levels. Experiments were carried out on the bench top to determine the frequency response of the actuator, as well as hinge moment and displacement capabilities. Flap de ections of 11.5 deg were demonstrated while operating under no load conditions at 10 Hz. Excessive creep at low frequencies precluded the measurement of achievable hinge moments, but extrapolation from de ection and voltage characteristics indicate that if properly scaled, the present actuator will produce ap de ections greater than 5 deg at the 90% span location on an operational helicopter. In addition, the rst mode of the actuator was at seven times the rotational frequency (7/rev) of the target model scale rotor. Proper inertial scaling of this actuator could raise this modal frequency to 10/rev 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 seen between the experiment and analysis indicate that the model is a valid tool for predicting actuator response. Thesis Supervisor: Steven R. Hall, Sc.D. Title: Associate Professor of Aeronautics and Astronautics

Journal ArticleDOI
TL;DR: In this paper, the authors describe measurements of chordwise distributions of unsteady pressure at three radial locations on a two-bladed model helicopter rotor in hover, with the blades executing simple harmonic pitch oscillations about the quarter-chord pitch axis.
Abstract: This paper describes measurements of chordwise distributions of unsteady pressure at three radial locations on a stiff two-bladed teetering model helicopter rotor in hover, with the blades executing simple harmonic pitch oscillations about the quarter-chord pitch axis. The objective of the present work is to provide a data base to correlate the measured unsteady blade pressure distributions with dynamic inflow to establish the validity or deficiencies of available analytical methods for predictions of unsteady aerodynamic phenomena. The effect of dynamic inflow on rotor unsteady surface pressure has been


Journal ArticleDOI
TL;DR: In this article, a low aspect ratio rectangular wing was positioned at different locations in a rotor flowfield to simulate the aerodynamic environment encountered by the wings of tilt-rotors or the empennage of helicopters.
Abstract: Experiments were conducted to study the aerodynamic interactions between a rotor and a fixed lifting surface. A low aspect ratio rectangular wing was positioned at different locations in a rotor flowfield to simulate the aerodynamic environment encountered by the wings of tilt-rotors or the empennage of helicopters. Steady and unsteady pressure measurements were made on the wing at various chordwise and spanwise stations for different combinations of rotor thrust and advance ratio. Flow visualization was performed using the wide-field shadowgraph method, which helped to identify the locations of the rotor wake relative to the rotor and wing. The results have shown that the lifting surface operates in a highly unsteady three-dimensional flow environment with regions of partial or complete flow separation. In addition, large unsteady loads were induced on the wing by the rotor and its wake.


01 Jan 1994
TL;DR: In this paper, the results from a recently completed full-scale rotor test in the NASA Ames 40- by 80-Foot Wind Tunnel and a recent flight test of a BO-105 helicopter at the Deutsche Forschungsanstalt fur Luft-und Raumfahrt e.V. (DLR), Braunschweig, Germany are presented.
Abstract: Correlation of wind tunnel results of a full-scale fourbladed hingeless rotor system with flight test measurements is presented. The results presented are from a recently completed full-scale rotor test in the NASA Ames 40- by 80-Foot Wind Tunnel and a recently completed flight test of a BO-105 helicopter at the Deutsche Forschungsanstalt fur Luft-und Raumfahrt e.V. (DLR), Braunschweig, Germany. Rotor hub and blade loads are shown over a range of advance ratios from 0.1 to 0.34 for C T /σ=0.07. Good correlation of measured blade root flap bending moments in the wind tunnel with flight test measurements is shown when the rotor is trimmed to measured flight test hub pitching and rolling moments. Analytical results using the CAMRAD/JA (Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics, Johnson Aeronautics) analysis shows good correlation when the analysis is trimmed to isolated rotor once-per-revolution root flap moments.

Proceedings ArticleDOI
21 Mar 1994
TL;DR: In this article, a study of helicopter blade vibration control is presented, where the rotor is modelled by the finite element method and it is considered as a rotating beam undergoing the coupling motions of flapping, lead-lagging, axial stretching and torsion.
Abstract: A study of helicopter blade vibration control is presented in this work. The blade is modelled by the finite element method and it is considered as a rotating beam undergoing the coupling motions of flapping, lead-lagging, axial stretching and torsion. The blade model also considers a pretwist angle, offset between mass and elastic axes and isotropic material. The finite element matrices are obtained by energy methods and a linearization procedure is applied to the resulting expressions. The linearized aerodynamic loading is calculated for hover and the state-space approach is used to design the control system. The eigenstructure assignment by output feedback is used in the blade reduced model resulting from the application of the expansion method by partial fractions. The simulations for open and closed-loop systems are presented, having exhibited good response qualities, and they show that output feedback is a good alternative for helicopter vibration attenuation. >