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Recent advances in Eurocopter’s passive and active vibration control
Peter Konstanzer∗ Bernhard Enenkl [email protected] [email protected]
EUROCOPTER DEUTSCHLAND GmbH, Munich, Germany Pierre-Antoine Aubourg Paul Cranga [email protected] [email protected]
EUROCOPTER SAS, Marignane, France
Abstract
This paper is about recent advances in Eurocopter’s research activities on passive and active vibration control systems. Emphasis is placed on the reduction of rotor-induced vibrations which is still one of the key challenges in helicopter dynamics. Both passive and active means for the reduction of vibrations are discussed. A short review of the rotor-induced vibration problem is given with a link to typical vibration characteristics of Eurocopter helicopters. The concepts and means to influence and control vibrations are outlined. The main focus of this paper is on recent advances on i) vibration control at the rotor, ii) vibration control at the transmission and iii) vibration control at the fuselage. In the section “vibration control at the rotor”, vibrations are attacked at their source – the rotor itself. Advanced passive and active rotor dynamic layouts are of interest. Here, a 5-bladed bearingless vs. 4-bladed main rotor system on EC145 as well as an active flap rotor on the hingeless system of BK117 are discussed. For each system, key parameters of the design, data of the test configuration and test environment and in particular results on vibration reduction are presented. A substantial reduction of the exciting hub loads is achieved thus providing superior airframe vibration levels. In the section “vibration control at the transmission”, a new generation of pylon isolation system is presented. This passive system based on the SARIB principle combines advantages of efficiency, lightness, reliability and low cost design. This technology consists of a compact suspension and a flapping mass integrated in each gear box strut. It provides an important attenuation of the vibrations for all hub loads components and it was successfully tested in-flight. In the section “vibration control at the fuselage”, active anti-vibration control systems (AVCS) installed in the fuselage are presented. The systems rely on single-port active devices which are capable to generate inertia-based control forces which induce a secondary vibration field to compensate the vibration disturbance. Here, systems based on electromagnetic actuation technology for EC225 as well as Piezo-ceramic technology demonstrated on EC135 are presented.
∗ Author to whom correspondence should be addressed. Presented at the American Helicopter Society 64th Annual Forum, Montréal, Canada, April 29 – May 1, 2008. Copyright © 2008 by the American Helicopter Society International, Inc. All rights reserved.
Introduction
Rotor-induced vibration is still one of the main challenges for a passenger friendly helicopter cabin. The reason for this is threefold: Firstly, helicopters are subject to a highly asymmetric, turbulent aerodynamic environment resulting in high vibratory loads. Secondly, the requirement to design for minimum weight leads to flexible airframe structures with high modal density and considerable dynamic response. Thirdly, the passenger is in very close proximity to the disturbing sources and has a high level of perception in the frequency range of interest. Extensive research has been conducted on various concepts of vibration control, see [1]. Following the announcement of the “helicopter ride revolution”, see [2], vibration reduction from 0.4g acceleration amplitudes to values as low as 0.05g has been targeted. This stepwise reduction is associated with technological key milestones such as
- optimal tuning - dynamic absorbers - rotor isolation - active control technology
The “jet smooth ride” levels below 0.05g are today still not achieved for production type helicopters. This paper is about recent advances in Eurocopter’s research activities on vibration control. Both passive and active means are discussed. These innovative means enable Eurocopter to realize the highly desired “jet smooth ride” levels. The focus of this paper is on recent advances on vibration control
- at the rotor - at the transmission - at the fuselage
where “recent” refers to technology which has been flight tested during the last four years. In the section “vibration control at the rotor”, a 5-bladed bearingless versus a 4-bladed main rotor system on EC145 as well as the active flap rotor on the hingeless system of BK117 are discussed. In the section “vibration control at the transmission”, a new generation of a pylon isolation system is presented. In the section “vibration control at the fuselage”, active anti-
vibration control systems (AVCS) installed in the fuselage are presented based on electromagnetic actuation technology applied to EC225 and Piezo-ceramic actuation technology demonstrated on EC135. For each system, the concept, key parameters of the design, the dynamic layout, a description of the test environment as well as flight test results are presented. In this paper, techniques such as fuselage dynamics tuning, rotor blade pendulum absorbers, cabin dynamic vibration absorbers as well as “standard” anti-resonance isolation system (e.g. ARIS and SARIB®) are considered as state-of-the-art and are not addressed. Rotor-induced vibrations and means of control
Rotor-induced vibration is the oscillatory response of the airframe to periodically varying aerodynamic loads acting on the rotor blades, see [3,4]. Whereas in hover flight less oscillating aerodynamic loads are generated, the asymmetric aerodynamic environment in forward flight leads to considerable higher harmonic loads. Due to these aerodynamic loads, the rotor blades execute a forced oscillation where higher harmonic blade root loads are generated. Transmitted to the non-rotating frame at harmonics which are a multiple of the blade passage frequency, higher harmonic hub loads in their part excite the airframe structure. In Figure 1, a typical amplitude spectrum obtained from flight measurements shows the characteristics for a BO105 with a 4-bladed rotor system.
Figure 1: Vertical cabin vibration spectrum in level flight of BO105 The typical vibration characteristics at different flight speed regimes are given in Figure 2. The flight test results are obtained with a 3/rev blade pendulum absorber. There are two regimes, low speed flight and high speed flight, where the vibration levels are challenging.
Figure 2: Variation of rotor-induced cabin vibrations with flight speed of BO105
Vibration control at the rotor
In this section, vibrations are attacked at their source – the rotor itself. Advanced passive and active rotor dynamic layouts are of interest. Passive rotor Concept The bearingless main rotor is in service on EC135 as a successor of the hingeless concept of BO105, BK117, and EC145. A research project called ATR was launched to further improve the bearingless technology in terms of weight, cost, and handling qualities, see [5]. Related to vibration excitation two main targets were addressed. The blade number was increased to 5 promising a strongly decreased aerodynamic excitation in conjunction with a much easier dynamic blade tuning in the frame of an enlarged rotor speed range. Additionally, a reduction of the moment transmissibility was aspired by keeping the equivalent hinge offset as low as possible. Compared to the 4-bladed rotor a reduction of vibratory hub loads in the range of 70% was envisaged. Design The flexbeam is the key element of a bearingless rotor (see Figure 3). The flat cross section at the hub area is acting like a flapping hinge. The more outboard H-shaped beam is very soft in torsion and represents the pitch bearing and the lead-lag hinge as well. The pitching of the blade is controlled via the cuff which encloses the flexbeam. The pitch link and the shear restraint bearing are located at the inboard end in conjunction with an elastomeric
lead-lag damper providing adequate structural damping for stability reasons. The flexbeam is attached by 5 bolts at the rotor hub. Unlike to the one piece design of the EC135 the flexbeam and the rotor blade are separated in this case allowing a full folding of the 5-bladed rotor. Because only an EC145 helicopter could be taken as a test bed the main parameters of the rotor like diameter and tip speed had to be fitted. Contrarily, the initial design was based on an increased diameter and reduced tip speed.
Rotor Hub
Blade Attachment
Flexelement
Lead Lag Hinge
Control Cuff
Hub Attach-ment
Rotor Blade
Lead Lag Damper
Flapping Hinge
PitchLink
Rotor Hub
Blade Attachment
Flexelement
Lead Lag Hinge
Control Cuff
Hub Attach-ment
Rotor Blade
Lead Lag Damper
Flapping Hinge
PitchLink
Figure 3: Hub design of the ATR Layout & Validation In advance of flight testing bench tests confirmed the function and the strength of the key components. One test simulated the blade pitching in the centrifugal field. A pitch angle of 8 ± 29° was applied under the nominal longitudinal load, which overestimates the twist by 30%.The first test in the centrifugal field took place on a whirl tower. Figure 4 gives an image of the rotor on the test rig. The main reason for the whirl tower test is to check the aeroelastic properties of the isolated rotor, such as the natural frequencies and the lead-lag damping as well as the bending and control loads at different collective and cyclic angle settings.
Figure 4: Whirl tower test of the ATR
Concerning the vibration behaviour the dynamic tuning of the blades is of importance. Figure 5 shows a comparison of calculated natural frequencies and test results. The harmonic oscillations of the blades generate forces and moments at the centre of the rotor, exciting the air frame via mast and gear box. In order to keep the vibrations away from the cabin it is most beneficial to minimize the loads at the rotor hub. There are various means for reducing the loads in the rotating frame, like a thorough tuning of the blade bending natural frequencies or the use of absorbers. But the most efficient way is to increase the number of blades, because the amplitudes of the air loads significantly decrease with increasing harmonic number.
Rotor Speed - % nominal
Freq
uenc
y-H
z
0 25 50 75 1000
10
20
30
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Figure 5: Dynamic Blade tuning of the ATR Additionally, the dynamic tuning of the blades is easier and avoids the use of tuning masses. It lowers the capability of the blade to transfer the loads towards the hub. In case of a 5-bladed rotor the harmonic excitation with 4, 5, and 6/rev are responsible for the residual loads at the rotor centre. Test environment After removing the production rotor system the new rotor including mast and controls was installed on an EC145 helicopter using approximately 50 gauges for monitoring of test data. Sensors in the rotating frame measured the blade bending moments, the mast moment and the control loads, whereas in the fixed frame mainly the vibrations were analyzed. During ground tests the track and balance of the rotor was adjusted and the stability
with respect to ground resonance was confirmed. After becoming airborne the air resonance stability was checked. During 20 test flights the vibration behaviour, the handling qualities, the flight loads, the required power, and the noise emission was surveyed. The flight test vehicle is shown in Figure 6.
Figure 6: ATR research rotor on EC145 Flight test results Regarding the vibration behaviour the corresponding hub loads are of interest. Figure 7 gives a comparison of the in general most important load component in the rotating frame, which is the (N-1)/rev hub moment. N is the number of blades. The data confirm the expected reduction in the range of about 70% for the 5-bladed rotor. Absorbers as used in the production helicopter were not applied in both cases.
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(N-1
)/rev
Mas
t Ben
ding
Mom
ent
4-Bladed5-Bladed BMR
Figure 7: (N-1)/rev mast bending moments of 4- and 5-bladed rotors (w/o blade absorbers)
The vibratory response of the air frame strongly depends on the N/rev frequency and on the applied hub loads as well as on the gross mass and centre of mass location. Frequency and loads are changed with a 5-bladed system. Figure 8 and Figure 9 show the vertical and lateral 5/rev accelerations during level flight. The accelerometers were located on the cabin floor behind the crew seats and in the aft cabin. The vibration levels stay far below the commonly accepted 0.1g limit for most flight conditions.
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/rev)
Pilot CopilotAft Cabin
Figure 8: Vertical 5/rev accelerations on the cabin floor during level flight
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y - A
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/rev)
Co-/ Pilot Aft Cabin
Figure 9: Lateral 5/rev acceleration on the cabin floor during level flight
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0 20 40 60 80 100 120 140Flight Speed - kts IAS
Intr
usio
n In
dex
-
Pilot CopilotAft Cabin
Figure 10: Intrusion Index on the cabin floor
A better assessment of vibrations acting on passengers and crew members offers the so called Intrusion Index, which takes into account the frequency as well as the orientation of the applied accelerations. A level of 1 corresponds well with a good crew rating. Especially at high speeds the Intrusion Index was found far below 1, see Figure 10. Besides the in general low acceleration level the increased frequency at 32 Hz (5/rev) decreases the Index as well. Active rotor Concept The concept of the active flap rotor for vibration control is to compensate vibratory hub loads by secondary hub loads generated by aerodynamic on-blade actuation [6-12]. Although the electro-hydraulic system of the BO105 behaved well during the experimental campaign, a promising actuation concept for future applications was seen in Piezo-actuated trailing edge flaps. For the new experimental rotor system a BK117 was selected as test bed which has a Boelkow type hub as well. An important design parameter is the radial position of the flap. Parametric studies revealed that for BVI noise reduction purposes the flap should be shifted as close as possible to the blade tip. Due to the blade tip design with a swept back planform, the outboard end of the flap was limited to radius station 4.9 m (0.89R). In contrary the most beneficial location for vibration control is in the mid range span at 4 m (0.73R). Flap chord and the torsional stiffness of the blade impact the blade response as well. Lower torsional blade stiffness and smaller flap chords support the servo effect of the flap and help to limit the required actuator power [6].
Slot in theTrailing Edge
Figure 11: Installation of the flap units from the trailing edge
Design The flap actuator has to withstand high mechanical loads and should feature low volume and slim shape to fit into the blade cross section. A Piezo-ceramic system was chosen due to its good controllability and efficiency as well as its remaining inherent stiffness in case of a power loss. Basis for the implementation of the flaps was the EC145 main rotor. The design of the blade was modified to integrate the active trailing edge flaps. The flap system consists of three identical units. Dedicated studies revealed that one pair of Piezo-electric actuators are able to run a flap of 300 mm radial extension and a chord of 50 mm. The trailing edge of the blade was cut out and the foam was substituted by a flat box which is open at the aft side. After inserting the units into the blade, (see Figure 11) all the parts are screwed and sealed to ensure the stiffness and strength requirements of the blade, as well as the protection of the flap actuation system against humidity. The actuator/flap unit is self-contained and can be run on bench outside of the blade. The most important design targets were a high structural stiffness, low friction of the bearings, no mechanical play, and low mass of the unit. One pair of Piezo-electric actuators located at a most forward chordwise position act via tension rods on the flap, see Figure 12.
Flap
PiezoActuator
PreloadAdjustment
Tension Rod
Frame
Figure 12: Flap unit assembly Layout & Validation During the development process all main parts such as actuators, housing, flaps, tension rods, data pick-ups, wiring, power supply, and controllers had to undergo subsidiary test procedures. Performance tests under realistic loads demonstrated a life time adequate for an experimental system. The goal of the tests on the whirl tower was to evaluate the aeroelastic properties of the blades such as natural
frequencies and inplane damping as well as the capabilities of the active flap units. During the first tests the flaps were deflected with steady inclination to check the influence on the blade track and blade control forces. Later on the flaps were run with frequencies up to the eighth rotor harmonic.
Airframe Sensors
OutboardSensors
Transmission DeviceEnergy and Signal Trans-mission, Azimuth Sensor
Signal ProcessingTelemetry,Data Recording
Terminal(Laptop)
ControllerPower Converter, ControlInterface, I/O-Signal-Board,Shaft Encoder Interface
Flaps with Piezo ActuatorsHub Electronics
ControlDisplay
Unit
Control Panel
Hub SensorsBlade SensorsBlade Tip Sensors Flap Unit Sensors
Figure 13: System architecture of the test helicopter Test environment A sketch of the system architecture of the BK117 test helicopter is shown in Figure 13. Besides the rotor with active flaps and the applied sensors, a cylindrical compartment is mounted on the rotor hub. It houses the signal conditioning and processing as well as the power distribution to the individual flap units. The electric power is transferred by a brushless transducer and the data link is established by a bi-directional optical system [7]. Each blade is equipped with sensors monitoring important parameters like actuator forces and strokes, flap angles, accelerations at the hub, blade surface pressure, structure born noise, blade bending moments, torsional moments, and blade control forces. The airframe contains sensors for accelerations, control loads, control angles and noise. The system is completely independent from the primary flight control and thus it is not a flight safety critical item. In case of malfunction or loss of electric power of the actuation system, an uncomfortable vibration level may appear but it will not influence the controllability and safety of the helicopter. After securing a safe operation of the aircraft on ground which includes the balancing of the rotor and checking of ground resonance stability, the system was activated the first time. The helicopter became airborne with some hover flights, followed by air resonance tests and checks of the handling qualities. Although vibration suppression means were not
installed, the vibration level was rated acceptable by the crew. The official first flight of a helicopter with active trailing edge flaps took place in September 2005. The airborne demonstrator is depicted in Figure 14.
Figure 14: BK117 S7045 with active flaps rotor Individual blade control is an efficient means for reducing annoying rotor-induced vibrations. The selected control concept has the aim to eliminate the 4/rev hub loads as far as possible. The limited number of control variables restricts the design of disturbance rejection controllers to the roll and pitch moment and to the vertical force at the hub. Robust disturbance rejection control of these three parameters leads to the implementation of dynamic compensators in the feedback loop. The compensators are derived from the internal model principle and are realised as notch filters for modelling the sinusoidal nature of the disturbances at the blade passage frequency. The core of the vibration controller is formulated in the non-rotating system using multi-blade coordinates for blade sensors and actuation control data. The vibration controller consists of washout filters for pre-conditioning the sensor signals, of notch filters acting as servo compensators, and of the static gain matrix. In [9] a semi-empirical procedure has been developed for the determination of the optimal gain matrix based on flight test results. In [12] a model-based control system design based on H∞-control has been developed and tested, too. The required 4/rev hub loads are derived from shaft bending moments, as well as from flap and lead-lag bending moments. The control algorithms have been developed based on system modelling and identification, controller design and simulation, and realtime code generation.
The system modelling and identification is based on comprehensive rotor models (e.g. CAMRAD II) and Matlab scripts for open loop system transfer function identification from flight test data. The general structure of the vibration controller is shown in Figure 15.
Vibration Controller(Disturbance Rejection)
Transformation into Rotating
System
Transformation into Fixed System
2 Shaft Moments
Hub Loads
IBC
4 Flap Bending
Flap BendingMoments
Shaft Moments
IBC IBC ActuationVibration Controller(Disturbance Rejection)
Transformation into Rotating
System
Transformation into Fixed System
2 Shaft Moments
Hub Loads
IBC
4 Flap Bending
Flap BendingMoments
Shaft Moments
IBC IBC Actuation
Figure 15: Schematic view of the vibration control system Flight test results At first open loop flight tests were performed. The establishment of a reliable 4/rev transfer matrix is a crucial task for the design of the disturbance rejection controller. Figure 16 shows results in a representative manner for collective flap actuation with a control voltage of 40%.
Flap Actuation ϕIBC - Deg
Fz-k
N
0 30 60 90 120 150 180 210 240 270 300 330 3600
0.5
1
1.5
2
2.5
Figure 16: Vertical hub force vs. IBC phase in level flight at 100 kts The transfer matrix was defined assuming a linear relationship between input and output. The identified model agrees well with the appropriate test data which is shown for example in Figure 17. The controlled 4/rev hub loads during level flight conditions are plotted in Figure 18.
Reference Level
Identified
Measurement
ϕIBC Identified
Measurement
ϕIBC
Figure 17: Comparison of 4/rev pitch moments at 100 kts (identified vs. measurement) A significant reduction of up to 90% could be demonstrated, when the IBC control system was engaged. The corresponding 4/rev gearbox vibrations are depicted in Figure 19. A remarkable enhancement is obtained especially in longitudinal and lateral direction. In Figure 20 the vertical 4/rev accelerations in the cabin at the co-pilot position are shown. The vibrations drop down to values below 0.05g. The remaining accelerations may be explained by the still existing not-controlled hub loads like lateral hub forces and the moment Mz. Overall the achieved low vibration levels were confirmed by the flight test crew’s excellent rating. The robustness of the IBC controller was demonstrated in a wide range of the flight envelope. The results show that the controller efficiently rejects the vibratory hub loads (Mx, My, Fz) and reduces the cabin accelerations below the 0.1g level. The reference values are without pendulum absorbers which are used in serial aircrafts. The flight tests were aimed on testing the complete collective and cyclic disturbance rejection control for Fz, Mx, and My, simultaneously. They will be continued with an advanced controller and further optimized parameters.
Flight Speed - kts (IAS)
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omen
tMx
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m
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4/rev Roll Moment (Fixed System)
Figure 18: Controlled hub loads vs. flight speed
Flight Speed - kts (IAS)
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4/rev Gearbox Vibration (y-Dirction)
Figure 19: 4/rev gearbox vibrations vs. level flight speed
Flight Speed - kts (IAS)
VK
ZCO
-g
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ReferenceControl On
4/rev Cabin Vibration (z-Direction)
Figure 20: Vertical 4/rev cabin vibration vs. level flight speed
Vibration control at the transmission Concept The DAVI principle has been largely used over the last 30 years through numerous and various designs, see [13,14]. This isolation system is generally integrated to the pylon assembly to filter the dynamic loads transmitted to the airframe. Compared to the classic vibration absorber, DAVI interest lies in the ability to generate important inertial forces with low tuning mass thanks to an amplification of the mass movement, see Figure 21.
Fuselage
Main Gear Box
a
c Suspension
stiffness
Tuning massac
=λ
Amplification ratio:
Figure 21: DAVI principle Eurocopter has already developed and certified two concepts of isolation system: ARIS and SARIB®. ARIS system is based on hydraulic amplification by using two concentrical cylinders, see [15]. The amplification is provided by the ratio of the two cylinder section areas. The light twin multi-mission aircraft EC135 is equipped with ARIS. SARIB® system uses a mechanical lever implementation, see Figure 22. The military helicopters NH90 and Tiger are equipped with SARIB®. The system is composed of four individual units equally spaced around the main gear box (MGB), see Figure 23. Each unit is made of a leaf spring and an arm supporting a flapping mass. The leaf is connected by pivot links to the structure fitting and to the MGB strut. The leaf tip is supported by the lower part of MGB. Elastomeric bearings are mounted at each connection to permit small rotations or displacements under high static loads. A membrane permits the motion of the MGB (vertical displacement and pitching and rolling rotations) and to transmit the torque.
(y-Direction)
Figure 22: SARIB principle
Figure 23: NH90 SARIB SARIB® and ARIS are very efficient to reduce fuselage vibrations because they are able to filter all hub load components. These systems are in particular suitable for 3 or 4 bladed rotor. These two last years much effort was put to the simplification and the cost reduction of SARIB®. These investigations lead to a new concept in which the leaf spring is replaced by a torque tube spring, Figures 24 and 25. The system is so called Torque Tube SARIB®.
Figure 24: Torque Tube SARIB® principle
Design The torque tube leads to a very compact design. It simplifies the integration on the MGB upper deck and permits the compatibility with all MGB accessories. The elastomeric bearings are removed at the leaf tip and the number of metallic parts is very limited which has a significant impact on the cost and the weight of the complete system. The simplicity of the design permits to have a system which is reliable and maintenance free. The system is tuned by adjusting the number of washers.
Figure 25: Torque Tube SARIB® The Torque Tube SARIB® system is sized by using an analytical model and from in-flight identified hub loads. The model permits to perform quickly sweepings on different parameters and to optimize the system architecture. The amplification ratio is a compromise between the minimization of the tuning mass, the acceleration of the tuning mass and the anti-resonance width. It is set at about 8. A low stiffness of the suspension permits to minimize the weight of the flapping mass but the MGB displacements must remain acceptable in flight to avoid important misalignments with engine shafts and potential couplings with the flight control kinematics. The ratio between the vertical stiffness and the aircraft weight is set at 550 daN/mm/tons. Laboratory tests A first test is undertaken on an isolated SARIB® unit to validate its design. For this test the structure fitting is fixed on a bench and the action of the MGB strut is replaced by a hydraulic actuator. The Torque Tube SARIB® can be subjected simultaneously to static loads and to a dynamic excitation, see Figure 26.
Membrane
MGB struts
SARIB
Flapping mass
MGB strut
Flapping mass
Torque Tube Membrane
Fitting
MGB
Fuselage
Pivot
c
a
GB strut
Fuselage
Flapping mas
SARIB leaf
Membrane
Fitting
MGB
Pivot Support
c a
Torque tube housing
Structure fitting Bearings block
Tuning mass: Cylinder weight and washers
Support arm
Stiffener
MGB strut
Figure 26: Test rig of an isolated Torque Tube SARIB® unit The dynamic behavior of the system is analyzed by performing frequency sweeping. Loads transmitted to the structure fitting and to torque tube, and the accelerations of the flapping mass are measured. Figure 27 shows that the loads can be entirely cancelled by the inertia forces of the flapping mass.
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Frequency (Hz)
Red
uced
tran
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ad
Figure 27: Transmitted load by an isolated Torque tube SARIB unit versus excitation frequency A shake test is performed on an aircraft equipped with the complete isolation system. The main objective is to check the efficiency of the Torque Tube SARIB® in the real environment and to define a tuning for each component (moments, out and in-plane forces). Due to the layout of the 4 units the optimal tuning is not exactly the same for each component and the optimal tuning is the result of a
compromise which is easily reached. A deep and wide antiresonance is created on the complete airframe around the N/rev, see Figure 28. The comparison with a shake test reference shows that the Torque Tube SARIB® system provides an excellent attenuation of the vibration levels, see Figure 29.
Figure 28: Frequency Response of the airframe for different hub loads components
Figure 29: Frequency Response of the airframe with and without Torque Tube SARIB Flight test results During the first flights the tuning is slightly refined through an iterative process which combines flight test data and simulation. As previously observed during the shake tests, the system provides a significant vibration reduction and a high level of comfort, see Figure 30. The wide SARIB® antiresonance permits to be not too much sensitive
N/rev
N/rev
to the slight rotor speed variation and to be independent of the aircraft configuration (weight, center of gravity).
Pilot seat
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atio
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Aircraft equiped with Torque Tube SARIB
Cabin rear
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Vib
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Figure 30: Vibration level versus flight speed The low cost design, the compactness and the efficiency of the Torque Tube SARIB® makes this isolation system very attractive.
Figure 31: EC225 AVCS
Vibration control at the fuselage In this section, active anti-vibration control systems (AVCS) installed in the fuselage are presented. The systems rely on single-port active devices which are capable to generate inertia-based control forces which induce a secondary vibration field to compensate the vibration disturbance. The AVCS consists of, see Figure 31: - Accelerometers located at relevant structure areas to measure the vibration level. - A reference signal correlated with the frequency of the vibration to be minimized. - An embedded computer which determines the appropriate command for the actuators. - An amplifier of the command signal. - A set of actuators which generate dynamic forces. The potential reduction of vibration thanks to AVCS is very high because the actuators can deliver exactly the loads at the right amplitude and phase in order to counteract the primary vibrations. It is well suitable for rotor speed variation. This system is easy to maintain because it is directly integrated to the structure (not mounted on the upper deck or on the rotor) and it can be continually monitored by the crew. Electromagnetic AVCS Concept The first AVCS developed and certified by Eurocopter is based on an electromagnetic actuator. This force generator is made of a magnetic inertial mass supported by a spring and controlled by an electromagnetic field generated by a coil, see Figure 32. The stiffness of the spring elements and the magnet mass are chosen in such a way that the resonance frequency of the system is close to the frequency range of the vibrations to be controlled, see Figure 33. It permits to provide important dynamics effort with small currents and so to minimize the required power.
0.1 g
iV
M1
Aimant Permanent
Elément défo
Bobine
Circuit de contrôleControl circuit
Coil Magnet
Spring
Dynamic force
Figure 32: AVCS – electromagnetic actuator
0.001
0.01
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1
10
0 5 10 15 20 25 30 35 40
Module H0
Fréquence (Hz)
Zone d'amplification
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Amplification
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Figure 33: Transfer function of an electro-magnetic actuator Application to EC225 EC225 is the latest member of the Super Puma family and it entered in its civil and military version in 2005. This 11 tons helicopter benefits from the latest technological innovations. Thanks to electromagnetic AVCS, EC225 is a helicopter with a new generation of anti-vibration system. EC225 is also equipped with a five bladed Spheriflex®
rotor, new avionic and autopilot systems. Its power plant with two Turbomeca Makila 2A engines provides 14% more power than the previous versions and the new main gear box permits to fly 30 minutes following an accidental loss of oil [16]. Layout of the forces generators and the accelerometers In order to evaluate the number of actuators and their optimal location, a Finite Element model has been developed and validated with shake test results. Simulations in closed loop have been performed in order to evaluate accurately the dynamic forces to be provided to control the vertical airframe vibrations. Particular attention was
paid to the attachment of the force generators to transmit efficiently the loads into the structure. For the EC225 production version the system includes 3 actuators (one in the cockpit and two other ones at the cabin rear) and 4 accelerometers, see Figure 34. The AVCS system is the only anti-vibration system mounted on EC225. It represents less than 0.8% of the total mass of the aircraft. The electrical consumption remains low which makes AVCS very competitive.
RH Side LH Side
Figure 34: Layout of the electromagnetic AVCS on EC225 Flight test results Flight tests were necessary to tune the internal parameters of the control loop (convergence parameters, internal gain). Very low vibration levels are reached. With the AVCS active accelerations are divided by about 2 and are always below 0.1 g independent of flight speed, see Figure 35.
Thanks to the 5 bladed Spheriflex® rotor and the AVCS, the EC225 enjoys extra low vibration levels which do not worsen in turns or at high speed.
Figure 35: Vibration level with and without AVCS versus flight speed Piezo-AVCS on EC135 Concept The concept of the Piezo-Active Anti-Vibration Control System (Piezo-AVCS) is shown in Figure 36. Rotor induced vibrations in the cabin are actively reduced by a secondary vibration field generated through Piezo-active inertial force generators. Using acceleration sensors, this remote controlled system ensures the reduction of vibrations at desired locations, i.e. at the pilots’ and passengers’ seats as well as in the rear compartment for search & rescue tasks. The foundation for this Piezo-AVCS is an efficient force generator which is realized by innovative actuation technology, namely Piezo-ceramic actuation, see [17].
Sensor Outputs(Accelerations)Force Generator Inputs
VibrationController
PowerSupply ControllerPower
Sensors
Hub loadexcitation
Piezo-activeforce
generatorSensor Outputs(Accelerations)Force Generator Inputs
VibrationController
PowerSupply
Sensor Outputs(Accelerations)Force Generator Inputs
VibrationController
PowerSupply ControllerPower
Sensors
Hub loadexcitation
Piezo-activeforce
generator
Piezo-activeforce
generator
Figure 36: Piezo-AVCS concept Based on an active leaf-spring design, as shown in Figure 37, the force generator produces control forces by the inertia of its attached mass.
Figure 37: Piezo-active inertial force generator Design Series production Piezo-ceramic elements embedded into a GFRP structure provide direct electrical control over the control forces. Since no moving parts are involved, this smart actuation technology comprises some highly attractive features, e.g.
- no wear, no friction - high life time - silent actuation - direct electrical control - extremely accurate - very high bandwidth
AVCS OFF AVCS ON 0.1 g
Cockpit
Cabin
Flight speed (kts)
Prototype hardware is manufactured in a variety of geometrical dimensions as seen from the prototype line presented in Figure 38. The underlying technology based on Piezo-ceramic multilayer monolithic actuators has been retained unchanged for all prototypes. The prototype line ranges from preliminary test articles over endurance test articles to full-scale hardware. Single-mass as well as dual-mass inertial force generators have been manufactured and investigated.
dual
full scale hardware
endurance test articles
preliminary test articles
mass
singlemass
singlemass
dual
full scale hardware
endurance test articles
preliminary test articles
mass
singlemass
singlemass
Figure 38: Prototype line Layout & Validation The Piezo-active force generators are characterized and tested in the laboratory with respect to their frequency response characteristics and their control force output as well as their endurance capability. The schematic of the test rig is given in Figure 39.
MM
PowerAmplifier Controller
aM1 aM
2
MassAccel. 2
MassAccel. 1 FG
Fd
Hydraulics
TestPlatform
Base-Accel.
BaseVibration
Feedback
Dual Leaf Force Generator aB
28 V DC 4/rev
Force generated
Disturbanceforce
MM
PowerAmplifier Controller
aM1 aM
2
MassAccel. 2
MassAccel. 1 FG
Fd
Hydraulics
TestPlatform
Base-Accel.
BaseVibration
Feedback
Dual Leaf Force Generator aB
28 V DC 4/rev
Force generated
Disturbanceforce
Figure 39: Test rig schematic
The test rig with an installed dual-mass force generator is depicted in Figure 40. The force generator is installed on a test platform which is subjected to a disturbance induced by a hydraulic actuator. Base acceleration is fed back via a controller and a power amplifier driving the Piezo-active force generator. The control objective of this closed-loop test system is to minimize the base acceleration aB despite of the disturbance force Fd through the application of the control force FG.
Figure 40: Test rig Figure 41 shows the transfer function of the disturbance force on the base acceleration for the uncontrolled as well as the controlled case. As seen, a passive transmission zero of the test system is present at about 29.5Hz corresponding to the force generator eigenfrequency.
20 21 22 23 24 25 26 27 28 29 30 31 32 33 34-70-65-60-55-50-45-40-35-30-25-20
|aB| /
|Fd| -
dB
20 21 22 23 24 25 26 27 28 29 30 31 32 33 34-180-150-120
-90-60-30
0306090
120150180
frequency - Hz
phas
e (a
B/F
d) - d
eg
- Uncontrolled- Controlled- Simulated
TransmissionZero
26dB14dB
Frequency - Hz
Gai
n-d
BPh
ase
–de
g
20 21 22 23 24 25 26 27 28 29 30 31 32 33 34-70-65-60-55-50-45-40-35-30-25-20
|aB| /
|Fd| -
dB
20 21 22 23 24 25 26 27 28 29 30 31 32 33 34-180-150-120
-90-60-30
0306090
120150180
frequency - Hz
phas
e (a
B/F
d) - d
eg
- Uncontrolled- Controlled- Simulated
TransmissionZero
26dB14dB
Frequency - Hz
Gai
n-d
BPh
ase
–de
g
Figure 41: Transfer function (dist. force on base accel. aB/Fd)
Within the operating frequency range (25 to 27Hz), the controlled system provides about 26dB reduction with respect to the uncontrolled case. Moreover, this is about 14dB lower than the minimum at the passive transmission zero. Figure 42 shows the force generator performance at high voltage levels. Driving the dual-mass force generator in near-resonance condition leads to high control force amplitudes. This force level can be achieved down to 90% of the resonance frequency, i.e. down to 26.3Hz applying maximum input voltage. For frequencies between 26.3Hz and the resonance frequency of 29Hz the maximum force level is achieved at reduced input voltage levels.
0
200
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800
1000
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25 26 27 28 29 30
Frequency - Hz
Forc
e
Figure 42: Force amplitude vs. frequency for various input voltage levels Flight test results The Piezo-AVCS is flight-tested on an EC135 prototype aircraft, see Figure 43. The system is installed in the cabin for the reduction of lateral rotor-induced vibrations at 4/rev.
Figure 43: Piezo-AVCS prototype helicopter
As reference, a non-serial prototype of EC135 is used. Figure 44 shows the flight test results over rotor rpm, i.e. lateral vibrations in y-direction in the pilot/copilot plane, the passenger seat plane as well as the rear compartment plane.
y-vi
b-g
99 99.5 100 100.5 101 101.50
0.05
0.1
0.15
0.2
0.25
EC135 PrototypeEC135 w Piezo-AVCS
Passenger
RPM - % nom
y-vi
b-g
99 99.5 100 100.5 101 101.50
0.05
0.1
0.15
0.2
0.25Rear Compartment
y-vi
b-g
99 99.5 100 100.5 101 101.50
0.05
0.1
0.15
0.2
0.25Co/pilot
Figure 44: Vibration levels (flight measurement) over rotor speed
TAS - kts
y-vi
b-g
0 20 40 60 80 100 120 1400
0.05
0.1
0.15Rear Compartment
y-vi
b-g
0 20 40 60 80 100 120 1400
0.05
0.1
0.15EC135 PrototypeEC135 w Piezo-AVCS
Passenger
y-vi
b-g
0 20 40 60 80 100 120 1400
0.05
0.1
0.15Co/pilot
Figure 45: Vibration levels (flight measurement) over flight speed
As seen, the Piezo-AVCS ensures a considerable vibration reduction over the entire rpm range down to levels lower than 0.05g. In particular, in the pilot/copilot plane the substantial improvement with respect to the reference is clearly visible. Whereas the vibration level of the reference depends on the rotor rpm, the Piezo-AVCS ensures very low vibration levels independent of the rotor rpm. This improvement also holds when lateral vibrations are plotted over flight speed, as presented in Figure 45. Also here, vibration levels lower than 0.05g are achieved over the whole flight speed range in pilot/copilot, passenger and rear compartment plane.
Conclusions
Recent advances in Eurocopter’s research activities on passive and active vibration control systems have been presented. A wide portfolio of highly attractive vibration control means support Eurocopter’s success by innovation. All systems provide a further considerable reduction of rotor-induced vibrations down to levels of 0.05g and ensure Eurocopter to achieve the long-envisaged jet-smooth ride comfort. All means are highly efficient, minimum weight and extremely cost competitive. Which one is the best depends on whether retrofit or upgrade or entirely new solutions are required. Some of the techniques may be combined such as a 5-bladed bearingless main rotor with AVCS. Among all techniques, active rotor control is definitely pathbreaking and the road to success is multi-functionality, i.e. apart from vibration reduction, external noise reduction, rotor stability enhancement, load alleviation as well as power saving will push this technology. In future, the greening of air transport will demand for wider rotor rpm ranges in order to optimize for both low external noise levels and high fuel efficiency. In order to adapt to these rpm variations, the spread of active systems will further increase and substitute more and more passive systems with fixed frequency tuning.
References
[1] Strehlow, H., Mehlhose, R., Znika, P., Roth, D., “Review of MBB’s Passive and Active Vibration Control Activities”, The Royal Aeronautical Society in London, January 1990.
[2] Balke, H., “The Helicopter Ride Revolution”, The American Helicopter Society Northeast Region, November 1981.
[3] Reichert, G., “Helicopter Vibration Control – a Survey”, Vertica, 5(1):1 – 20, 1981.
[4] Loewy, R. G., “Helicopter Vibrations: A Technological Perspective”, Journal of the American Helicopter Society, 29:4 – 30, 1984.
[5] Bebesel, M., Schoell, E., Polz, G., “Aerodynamic and Aeroacoustic Layout of ATR (Advanced Technology Rotor)”, 55th AHS Annual Forum, Montreal, Canada, May 1999.
[6] Enenkl, B., Kloeppel, V., Preißler, D., Jänker, P., “Full Scale Rotor with Piezoelectric Actuated Blade Flaps“, 28th European Rotorcraft Forum, Bristol, United Kingdom, September 2002.
[7] Konstanzer, P., “Decentralized Vibration Control for Active Helicopter Rotor Blades“, European Rotorcraft Forum, Florence, Italy, 2005.
[8] Dieterich, O., Enenkl, B., Roth, D., “Trailing Edge Flaps for Active Rotor Control, Aeroelastic Characteristics of the ADASYS Rotor System”, 62nd AHS Annual Forum, Phoenix, AZ, USA, May 2006.
[9] Dieterich, O., “Application of Modern Control Technology for Advanced IBC Systems”, 24th European Rotorcraft Forum, Marseille, France, September 1998.
[10] Toulmay, F., Klöppel, V., Lorin, F., Enenkl, B., Gaffiero, J., “Active Blade Flaps – The Needs and Current Capabilities”, 57th Annual Forum, Washington DC, May 2001.
[11] Roth, D., “Advanced Vibration Reduction by IBC Technology”, 30th European Rotorcraft Forum, Marseille, France, September 2004.
[12] Dieterich, O., Konstanzer, P., Roth, D., Ayadi, W., Reber, D., Well, K., “Model Based H∞ Control for Helicopter Vibration Reduction - Flight Tests with Active Trailing Edge Flaps”, 33rd European Rotorcraft Forum, 2007.
[13] Jones, R., “Control of Helicopter Vibration Using the Dynamic Antiresonant Vibation Isolator”, National Aerospace Engineering and Manufacturing Meeting, Los Angeles, California, 1973.
[14] Desjardins, R. A., Hooper, W. E., “Antiresonant Rotor Isolation for Vibration Reduction”, 34th AHS Annual Forum, Washington D.C., May 1978.
[15]
Braun, D., “Development of Antiresonance Force Isolators for Helicopter Vibration Reduction”, 6th European Rotorcraft and Powered Lift Forum, September 1980.
[16] Vignal, B., “Development And Qualification Of Active Vibration Control System for the Eurocopter EC225/EC725”, 61st AHS Annual Forum, Grapevine, TX, USA, June 2005.
[17] Konstanzer, P., Grünewald, M., Jänker, P., Storm, P., “Piezo Tunable Vibration Absorber System for Aircraft Interior Noise Reduction”, Euronoise 2006, Finland, 2006.
