- Email: [email protected]
AWIATOR's Study of a Wing Load Control: Design and Flight-Test Results
ELSEVIER
IFAC PUBLICATIONS www.elsevier.comllocatelifac
AWIATOR'S STUDY OF A WING LOAD CONTROL: DESIGN AND FLIGHT-TEST RESULTS. Matthieu Jeanneau*, Nicky Aversa*, Stephane Delannoy*, Mark Hockenhull*
* Airbus - Contact adress: M014216 - 316, route de Bayonne - 31060 Toulouse cedex 03 - France Tel: +33561182511- Fax: +33561184325 - [email protected]
Abstract: Loads at the wing root sections of an aircraft are mainly induced by the vertical accelerations encountered in manoeuvre or in turbulence, but a non-negligible part comes from the wing bending deflections. In an attempt to reduce such loads and hence reduce structural weight, a load alleviation function has been studied by Airbus flight-control research-engineers. This study was conducted under the auspices of A WIA TOR, a European research program aimed at developing and testing challenging flying technologies. The load alleviation function developed in this study is divided into 3 parts, each with a dedicated objective. First, a passive control reduces the loads induced by pilot inputs or by turbulence, by deflecting ailerons and spoilers proportionally to the vertical acceleration of the aircraft. Secondly, an active control deals with wing oscillations induced by the bending structural modes. This part has been designed with modem H~ methods, to take into account and optimise various specifications: load reduction, robustness to payload, but also roll-on and roll-off criteria to avoid any interaction possibly modifying handling qualities. The third part concerns the activation !0gics of the first two parts. This last part is not discussed hereafter. Keywords: Active control, H-infinity optimization, Load frequency control, Flexible, Flight control.
I. INTRODUCTION AND CONTEXT
2. PRINCIPLE OF THE LOAD-DEDICATED LAW
This work has been conducted in the frame of A WIATOR. This European project involves more than 23 partners from Europe+Israel. Its target is the proof of concept and in-flight validation of wing (design) technologies for future aircraft application. The study presented in this paper deals with the design and flight test of a new flight control system which provides weight saving and enhanced passenger protection in exceptional gusts through an alleviation of loads encountered at the ilU1er wing. The solution presented hereafter relies on today's aircraft architecture (existing sensors and control surfaces). Its principle is explained in §2, while a more precise description of each sub-function is given in §3-4. The synthesis of the active control of structural oscillations and its performances are presented in §5. Coding, validation and flight tests results are discussed in §6-7. Conclusion and prospective works initiated by this study are then listed in §8. References of past works that have inspired this study are given in the final paragraph.
The load-dedicated function is based on the knowledge that the wing weight is directly linked to the bending moment of the inner wing. Therefore, by bringing the spanwise aerodynamic centre of pressure inboard, and by attenuating any oscillation around the static balance, one reduces the loads that drive wing weights. Pilot or Autopilot
Fig. I: Implementation of the load-dedicated law in the Flight-Control Primary Computer. This load-function neither modifies nor replaces the usual flight-control law, called normal law, flying on-board Airbus' aircraft. It is designed so as not to interfere with it in order to guarantee no modification
Copyright @ 2003 by Airbus France SAS published by IFAC.
withpermi5Sion
469
when entering a manoeuvre or turbulence, the spoiler and aileron deflections can only increase, unless Passive Control is switched inactive. This prevents from a short-time varying spoiler or aileron response in case of fast varying pilot orders or turbulence (highly varying by nature). Thus it guarantees the passiveness of the wing deflections control. Fig. 3 illustrates this description.
in the aircraft response to pilot orders. The implementation of the load-dedicated function is made via an additional feedback loop - see Fig. I . The load-dedicated function is divided into three parts designed to reduce inner wing bending moment in a given context. First part is dedicated to steady loads . It consists in deflecting both ailerons and spoilers to counter the wing root loads. This is a Passive Control efficient both in manoeuvres and turbulence. Second part, called Active Control consists in controlling the bending of the wing by deflecting the ailerons around their passive value . A dedicated Nz measure, called Nz1aw , is used for this closed-loop. Third part consists in the activation logics. They rely on the level of vertical acceleration Nz detected, then activating or deactivating the loadfunction during cruise so as to avoid continuous very small control demands . This part is not discussed in this article.
N
Fig. 3: Passive Control description. 4. ACTIVE CONTROL
Activation logic PQniv~
Control!-_ _ _ _ _ _- ,
4.1 Objectives
NZcc
AiIe ro ns flIte n Acti~
Active Control is dedicated to damping wing oscillations encountered in turbulence by active control of wings through symmetrical aileron-orders.
Activation logic
COIW,ol
Acliv~
CO,.,o{
Fig. 2: Structure of the load-dedicated law.
4.2 Active Control requirements
This load-function is active only in cruise, and deactivated during all other flight phases.
Regarding loads at the wing-root section, main contributors are the flight-mechanics modes (below 0.5 Hz) and the first wing-bending structural mode (between 1 arid 2 Hz). Flight-mechanics modes are fixed by the normal control law, and must not be modified, as they ensure handling qualities that fulfil all the regulation requirements. Thus, main objective of Active Control is to reduce loads by acting on the first bending mode only.
An additional order is sent to the elevator when the load-function is active. It is computed so as to compensate for the pitch moment induced by the ailerons and spoilers deflections. This pitch-moment compensation is not described hereafter. 3. PASSIVE CONTROL
The strong requirement not to modify handling qualities implies that the Active Control must not interfere with low frequency modes (below 0.5Hz).
3.1 Objectives and means Passive Control is dedicated to reducing high upward wing deflections induced either when pilotmanoeuvres corrunand positive Nz, or when aircraft enters some gust or wind turbulence. To counter the induced deflection, a spoiler or an upward aileron deflection is corrunanded.
Apart from the first bending mode which is wellknown, other structural modes (beginning around 2.4 Hz) exist. To simplify the validation of the control law, and particularly the aeroservoelastic justification, it has been decided that the Active Control must be as uninfluential as possible on structural modes other than wing fundamental bending.
3.2 Passive Control description When Passive Control is active, a spoiler deflection is corrunanded, proportional to the Nz at the center of gravity (CG). A symmetrical aileron deflection is also corrunanded proportional to the NZcG too. The amounts of aileron and spoiler deflections commanded are different. In any case, maximum aileron deflection allows sufficient remaining rollauthority for pilot manoeuvres. Activation filters are added to smooth the deflection demands. When Passive Control becomes inactive, their timeconstant become longer to have a very smooth return to neutral. Anti-backup filters are added, so that
Regarding robustness, the first bending mode is likely to have a varying frequency depending on the aircraft payload, wing-tanks filling, altitude and Mach . It is required that the Active Control performance is not affected by any variation possibly occurring in cruise configuration. Last requirement is a limitation of the Active Control order to half the ailerons maximum Passive Control deflection. This limitation ensures to let sufficient roll authority for pilot lateral manoeuvres.
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4.3 Active Control description The Active Control is a closed loop control of the first wing bending-mode, using a sixth order filter. Input of the law is Nz1aw , a dedicated measure, based on accelerometers giving Nz on the wing, minus the Nz at CG. NZwing gives a good observability of the first wing bending-mode. Removing NZcG cancels the impact of low frequency modes (flight mechanics' modes) .
Nz,uw
=
Nz /ejI wing + Nz right wing 2 - NZCG Fig. 4: H_-schema for Active Control design, showing in red the aircraft model with aileronorder as input u, and Nz1aw as measured output y, and in black the exogenous e to z transfer function of which the H_ norm is minimized.
5. ACTIVE CONTROL SYNTHESIS
5.1 Objectives Handling qualities, loads, aeroelastics and robustness requirements, as described in previous paragraph, are explicitly taken into account in the design of the control-law. Actuator-saturation requirement is not. It is verified a posteriori. Thus, the goal of the synthesis can be described as reducing the impact of wind on wing oscillations in the frequency domain of the first wing-bending mode, while not modifying the aircraft response in other frequency domains .
5.4 Evaluated performances Fig. 5 shows the root-locus of aircraji+Active Control modelled for one flight point and mass case. The minimisation of the wind to Nz1aw transfer function on the frequency range of the first bending mode is here clearly visible. Main effect is the damping of this structural mode . One notices in green the poles of the 6 th order control law, achieving the roll-on and roll-off specifications. These specifications have been achieved, as shown by the lack of impact of the Active Control on the rigid mode (Iow frequency) and on all other structural modes (high frequency) .
5.2 H_synthesis A dedicated H_-schema was designed . It contains a reduced order aircraft model, including rigid modes and part of the aircraft flexible .structural-modes. Output of this model is Nz1aw, while input is the symmetrical aileron order. See Fig. 4. The exogenous transfer function to be minimized is composed of 4 SISO transfer functions, each dedicated to one of the 4 requirements . The first one represents the transfer function from wind to Nz1aw and includes Von Karman filters that represent wind frequency distribution, and weightings for design tuning. Second and third transfer functions are dedicated to loop-shaping. They introduce a roll-on and a roll-off specifications that guarantees gains close to zero in the frequency domain outside the range of frequency of the first wing bending-mode. A fourth transfer function models the variations of the first bendingmode frequency. An appropriate weighting ensures to cover the range of possible variations in cruise.
Fig. 6 gives the theoretical wind to Nz transfer functions at three different locations. First subplot concerns Nz at the cockpit. Second one gives Nz at the centre of gravity. Third one concerns Nz at the rear of the fuselage . Last graph gives the wind to Nz1aw transfer function. One notices that the first wing bending-mode has almost no impact on the longitudinal accelerations at the front of the aircraft: resonance peaks occur only for higher frequency modes. At the middle and back of the fuselage (also noticeable on Nz1aw), impact of the first bending mode is clearly visible with a major resonance peak at frequency around 1.2 Hz for this flight point and mass case. One visualizes the performance realised by the Active Control when comparing the aircraft without specific control (k=O) and with the Active Control (k= I). The resonance peak is reduced by a factor of about 2. Sensitivity to the gain is provided via the plots with some intermediate gains ranging from k=OA to 0.9. These plots also show that the roll-on and roll-off specifications are achieved. Open-loop plots (k=O) and closed-loop plots (k>O) are similar in the frequency domains of the flight-mechanics mode (below O.5Hz) and of the higher frequency structural modes (above 2Hz), proving that the Active Control does not modify the aircraft response for these dynamics.
5.3 Controller reduction Due to aircraft rigid and flexible dynamics, Von Karman filters and weightings, the control law th obtained by the H_-synthesis is a 40 order filter . A reduction based on classical tools allows us to finally fmd a 6th order filter which fulfils all our requirements with no significant loss of performance compared to the 40 th order filter.
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Fig. 5: Root-locus of aircraft+Active Control for a feedback gain varying from 0 to I. In green are plotted the open-loop poles of the Active Control.
Fig. 6: Theoretical wind to Nz transfer functions for different values of feedback gain. k=0 corresponds to the aircraft without Active Control. First graph gives the transfer function at the cockpil, second at the middle of the fuselage, third at the rear, and fourth is the transfer function between wind and Nz 1aw •
Fig. 7 shows the theoretical time-response in presence of turbulence representative of structural sizing conditions. Plot in dashed red is the response without control, and plot in blue the response with the Active Control. One notices that without control, main harmonic is slightly above I Hz, corresponding to the first bending mode . With the Active Control this harmonic is hardly noticeable, and most noticeable harmonic seems slightly above 2 Hz. Apart from that, one notices that the peak values are reduced with the Active Control and that the total energy of the signal is also reduced, thus leading to lower loads at the wing root section.
Fig. 7: Aircraft time-response to a sizing gust.
Fig. 8 displays the time-response of the aileronsdeflection in presence of the same sizing gust. One verifies that maximum deflections remain below half the Passive Control values (plotted in dashed-red), thus allowing for sufficient roll authority for lateral manoeuvres.
Fig. 8: Ailerons-deflection in response to a sizing gust (in blue). In dashed-red are plotted the maximum values not to be exceeded.
Figures presented in this paragraph show the performances of the Active Control for one flight point and mass case, known as one of the most severe for loads. Robustness to other flight points and mass cases has been verified. Similar results in term of performances are observed but are not presented in this paper.
6. ON BOARD IMPLEMENT AnON Coding of the whole load-function (including both passive and active functions) was done using SAO at the sampling rate of 40ms. SAO is a DOl78B qualified C-code generator which automatically produces and implements on-board flying computers from engineering system block diagrams.
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Exhaustive validations of the load-function have been performed during simulator sessions and through loads and aeroelastics analysis, in order to deliver a flight clearance prior to any flight test. 7. FLIGHT TESTS
7.1 Flight-test program Flight tests were performed in October 2003. Both active and passive part of the control law were evaluated through specific tests designed to excite the aircraft and activate the load function . To test the robustness to the aircraft mass and flight points, two flights were performed. During each flight, flight points for three Mach/speed combinations were tested. These points are 0.7/280knots, 0 .8/300knots and 0.86/330knots. For each flight point the plane was excited with sinus sweep and pulse signals directly sent to the control surfaces: ailerons (inner and outer separately) and elevators, both in a symmetrical way. The sweeps were performed in open and closed loop to assess the control law performances on the damping of the first wing bending mode and the non degradation of the others flexible modes. The excitations swept from 0.5Hz up to 6Hz and the transfer function of the wing were measured into this range. The pulse signals perturbed the flight mechanics at levelled flight and roughly emulate the aircraft dynamics as it is in a light discrete gust. This resulted in the activation of the control law and the aircraft response gave a qualitative result on the control law efficiency.
Fig. 9: Transfer functions from inner ailerons to longitudinal accelerations Nz. First graph is Nz at outer engine location, second graph at wing tip . Flight point is Mach 0.86. At low frequencies (below 0.7 Hz) corresponding to the flight mechanics modes, one can verify that the activation of the load function does not have any influence (id est the red and blue lines overlay each other). The same is true at frequencies above 1.7 Hz, demonstrating compliance with the request that the law should not be active on any other structural mode apart from the first wing bending mode for robustness reasons.
During both flights, some turbulence area were encountered, and data were recorded showing the ailerons and spoilers deflections commanded by the law. Typical pitch and roll manoeuvres were also performed by the pilots to test the activation logics and the possible interactions of the law with aircraft handling qualities.
Fig. 10: Transfer functions from outer ailerons to longitudinal accelerations Nz at Mach 0.86.
7.2 Flight-test results The next four figures present the transfer functions identified in flight at two different flight points. In each figure, the first graph gives the transfer function (TF) from ailerons (inner or outer) to Nz sensor located above the outer engine. Second graph shows TF from ailerons to Nz at the wing tip. In blue is plotted the response without the load function. In red is plotted the response with the load function active . Results are very similar for all four figures . The peak in the frequency range for which the load control law is active (around 1.2Hz) is significantly shortened both in term of area and in term of maximum value. At the wing tip, the decrease is above a factor 2. At outer engine, the decrease is around 30% and the peak with the law active occurs at a higher frequency than without the law. This also contributes to a reduction of the loads.
Fig. 11 : Transfer functions from inner ailerons to longitudinal accelerations Nz at Mach 0.82. The requirements for the Active Control are therefore fulfilled , at least at the flight points and mass cases tested in flight. But simulation results provide confidence in the fulfilment of these requirements in the whole cruise flight domain. Concerning the degree of load alleviation (1 .2 Hz peak reduction), it may appear already satisfying, but one expects even better results in future studies, as the kinematics of the plane did include a low-pass Chebychev filter
473
which limit the achievable bandwidth of the Active Control (see alleviation better at - 1.0 Hz than at - 1.5 Hz). This filter is mounted as standard on board the test aircraft, but it could be removed or reshaped for load dedicated controls for which there is a need for higher bandwidth.
LIDAR, gust estimators and new control surfaces (Trailing Edge Devices). The potential benefits will be integrated in the future versions of load-dedicated functions . 9. ACKNOWLEDGEMENTS
Tramfer flJr'lC tion for outer MerON Ve-3CO:no b . Mach.. O.B2
The authors want to thank all A WIATOR partners for their cooperation in this project. The AWIATOR project is co-financed by the European Union within the Thematic Priority Aeronautics & Space of the Sixth Research Framework Programme. Special thanks also to the ONERA-DCSD for providing the tools used for the design of the Active Control.
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REFERENCES Canavin, J.R. (1978) . The Control of Spacecraft Vibrations Using Multivariable Output AIAAJAAS Astrodynamics Feedback. Conference. Palo Alto, California. August 7-9. Gilbert, M.G., D.K. Schmidt and T.A. Weisshaar (1984) . Quadratic Synthesis of Integrated Active Controls for an Aeroelastic Forward-SweptWing Aircraji. Journal of Guidance. Vo!. 7, No.
Fig. 12: Transfer functions from outer ailerons to longitudinal accelerations Nz at Mach 0.82.
IS
10
:2
/~,
, !
I ~F'~ ,.. r 'l , 10
~
25
30
3S
2.
'." . :: /I:' !YI~~"
.:;:1. ,:'
20
-
Nzj)ltol
Honliner, Fr., R . Manser and D. Mussmann (1985) . Flutter Suppression and Gust Load Alleviation. Garteur Program AG3 . Kaynes, I.W. and D.E. Fry (1982) . Gust Load Alleviation on a Flexible Aircraft. Technical Report of the Royal Aircraft Establishment, June. Koenig, K. (f988). Experience in Application of Active Vibration Control Technology to a Wind Tunnel Model and to Flying Airbus. ICAS-883.10-3. Pages 1542-1553 . Smith, S.E. (1991). Name of book in italics or underlined, page or chapter numbers if relevant. Publisher, Place of publication. Livne, E. (1993). Integrated Structure/ Control/Aerodynamic Synthesis of Actively Controlled Composite Wings. Journal of Aircraft. Vo!. 30, No. 3, May-June. Mukhopadhyay, V. (1987). Digital Robust Active Control Law Synthesis for Large Order Flexible Structure using Parameter Optimization . AIAA Symposium on Dynamics and Control of Large Structures. Suzuki, S. (1993) . Simultaneous Structure/Control Design Optimization of a Wing Structure with a Gust Load Alleviation System. Journal of Aircraft. Vol. 30, No. 2, March-April. Takahashi, M. and G.L. Slater (1986). Design of a Flutter Mode Controller Using Positive Real Feedback. Journal of Guidance. Vol. 9, No. 3.
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Fig. 13 : Response to a pulse on the outer ailerons. Excitations by a pulse on the outer ailerons were also performed and one example is plotted in blue in the first graph of Fig. 13 . The black trace on the second graph shows how it stimulated the wing modes, (mainly first ~ing bending). This pulse activated the law from t=9s to 18s (red plot of 2 nd graph) . Response of the law on ailerons is plotted in green on the I SI graph. The static part of this response is plotted in red and the part due to the Active Control in light blue. The effect of the first order filter which smooths the ailerons deflection commanded by the Passive Control is noticeable, as is the effect of the anti-backup filter which ensures that the passive deflection remains at a constant level during the Nz oscillations. 8. CONCLUSION AND PROSPECTS The load-reduction function described in this paper relies on current aeronautical technologies. Its requirements were quite challenging in regard to loads, handling qualities, aeroe1astics and robustness specifications. Validations on models, through simulator sessions, and measured in flight show a significant alleviation of the bending oscillations of the wing. This function is the latest step in the process of synthesizing control-laws dedicated to load alleviation. AWIA TOR partners are working hard in developing new technologies such as load sensors,
474
IFAC PUBLICATIONS www.elsevier.comllocatelifac
AWIATOR'S STUDY OF A WING LOAD CONTROL: DESIGN AND FLIGHT-TEST RESULTS. Matthieu Jeanneau*, Nicky Aversa*, Stephane Delannoy*, Mark Hockenhull*
* Airbus - Contact adress: M014216 - 316, route de Bayonne - 31060 Toulouse cedex 03 - France Tel: +33561182511- Fax: +33561184325 - [email protected]
Abstract: Loads at the wing root sections of an aircraft are mainly induced by the vertical accelerations encountered in manoeuvre or in turbulence, but a non-negligible part comes from the wing bending deflections. In an attempt to reduce such loads and hence reduce structural weight, a load alleviation function has been studied by Airbus flight-control research-engineers. This study was conducted under the auspices of A WIA TOR, a European research program aimed at developing and testing challenging flying technologies. The load alleviation function developed in this study is divided into 3 parts, each with a dedicated objective. First, a passive control reduces the loads induced by pilot inputs or by turbulence, by deflecting ailerons and spoilers proportionally to the vertical acceleration of the aircraft. Secondly, an active control deals with wing oscillations induced by the bending structural modes. This part has been designed with modem H~ methods, to take into account and optimise various specifications: load reduction, robustness to payload, but also roll-on and roll-off criteria to avoid any interaction possibly modifying handling qualities. The third part concerns the activation !0gics of the first two parts. This last part is not discussed hereafter. Keywords: Active control, H-infinity optimization, Load frequency control, Flexible, Flight control.
I. INTRODUCTION AND CONTEXT
2. PRINCIPLE OF THE LOAD-DEDICATED LAW
This work has been conducted in the frame of A WIATOR. This European project involves more than 23 partners from Europe+Israel. Its target is the proof of concept and in-flight validation of wing (design) technologies for future aircraft application. The study presented in this paper deals with the design and flight test of a new flight control system which provides weight saving and enhanced passenger protection in exceptional gusts through an alleviation of loads encountered at the ilU1er wing. The solution presented hereafter relies on today's aircraft architecture (existing sensors and control surfaces). Its principle is explained in §2, while a more precise description of each sub-function is given in §3-4. The synthesis of the active control of structural oscillations and its performances are presented in §5. Coding, validation and flight tests results are discussed in §6-7. Conclusion and prospective works initiated by this study are then listed in §8. References of past works that have inspired this study are given in the final paragraph.
The load-dedicated function is based on the knowledge that the wing weight is directly linked to the bending moment of the inner wing. Therefore, by bringing the spanwise aerodynamic centre of pressure inboard, and by attenuating any oscillation around the static balance, one reduces the loads that drive wing weights. Pilot or Autopilot
Fig. I: Implementation of the load-dedicated law in the Flight-Control Primary Computer. This load-function neither modifies nor replaces the usual flight-control law, called normal law, flying on-board Airbus' aircraft. It is designed so as not to interfere with it in order to guarantee no modification
Copyright @ 2003 by Airbus France SAS published by IFAC.
withpermi5Sion
469
when entering a manoeuvre or turbulence, the spoiler and aileron deflections can only increase, unless Passive Control is switched inactive. This prevents from a short-time varying spoiler or aileron response in case of fast varying pilot orders or turbulence (highly varying by nature). Thus it guarantees the passiveness of the wing deflections control. Fig. 3 illustrates this description.
in the aircraft response to pilot orders. The implementation of the load-dedicated function is made via an additional feedback loop - see Fig. I . The load-dedicated function is divided into three parts designed to reduce inner wing bending moment in a given context. First part is dedicated to steady loads . It consists in deflecting both ailerons and spoilers to counter the wing root loads. This is a Passive Control efficient both in manoeuvres and turbulence. Second part, called Active Control consists in controlling the bending of the wing by deflecting the ailerons around their passive value . A dedicated Nz measure, called Nz1aw , is used for this closed-loop. Third part consists in the activation logics. They rely on the level of vertical acceleration Nz detected, then activating or deactivating the loadfunction during cruise so as to avoid continuous very small control demands . This part is not discussed in this article.
N
Fig. 3: Passive Control description. 4. ACTIVE CONTROL
Activation logic PQniv~
Control!-_ _ _ _ _ _- ,
4.1 Objectives
NZcc
AiIe ro ns flIte n Acti~
Active Control is dedicated to damping wing oscillations encountered in turbulence by active control of wings through symmetrical aileron-orders.
Activation logic
COIW,ol
Acliv~
CO,.,o{
Fig. 2: Structure of the load-dedicated law.
4.2 Active Control requirements
This load-function is active only in cruise, and deactivated during all other flight phases.
Regarding loads at the wing-root section, main contributors are the flight-mechanics modes (below 0.5 Hz) and the first wing-bending structural mode (between 1 arid 2 Hz). Flight-mechanics modes are fixed by the normal control law, and must not be modified, as they ensure handling qualities that fulfil all the regulation requirements. Thus, main objective of Active Control is to reduce loads by acting on the first bending mode only.
An additional order is sent to the elevator when the load-function is active. It is computed so as to compensate for the pitch moment induced by the ailerons and spoilers deflections. This pitch-moment compensation is not described hereafter. 3. PASSIVE CONTROL
The strong requirement not to modify handling qualities implies that the Active Control must not interfere with low frequency modes (below 0.5Hz).
3.1 Objectives and means Passive Control is dedicated to reducing high upward wing deflections induced either when pilotmanoeuvres corrunand positive Nz, or when aircraft enters some gust or wind turbulence. To counter the induced deflection, a spoiler or an upward aileron deflection is corrunanded.
Apart from the first bending mode which is wellknown, other structural modes (beginning around 2.4 Hz) exist. To simplify the validation of the control law, and particularly the aeroservoelastic justification, it has been decided that the Active Control must be as uninfluential as possible on structural modes other than wing fundamental bending.
3.2 Passive Control description When Passive Control is active, a spoiler deflection is corrunanded, proportional to the Nz at the center of gravity (CG). A symmetrical aileron deflection is also corrunanded proportional to the NZcG too. The amounts of aileron and spoiler deflections commanded are different. In any case, maximum aileron deflection allows sufficient remaining rollauthority for pilot manoeuvres. Activation filters are added to smooth the deflection demands. When Passive Control becomes inactive, their timeconstant become longer to have a very smooth return to neutral. Anti-backup filters are added, so that
Regarding robustness, the first bending mode is likely to have a varying frequency depending on the aircraft payload, wing-tanks filling, altitude and Mach . It is required that the Active Control performance is not affected by any variation possibly occurring in cruise configuration. Last requirement is a limitation of the Active Control order to half the ailerons maximum Passive Control deflection. This limitation ensures to let sufficient roll authority for pilot lateral manoeuvres.
470
4.3 Active Control description The Active Control is a closed loop control of the first wing bending-mode, using a sixth order filter. Input of the law is Nz1aw , a dedicated measure, based on accelerometers giving Nz on the wing, minus the Nz at CG. NZwing gives a good observability of the first wing bending-mode. Removing NZcG cancels the impact of low frequency modes (flight mechanics' modes) .
Nz,uw
=
Nz /ejI wing + Nz right wing 2 - NZCG Fig. 4: H_-schema for Active Control design, showing in red the aircraft model with aileronorder as input u, and Nz1aw as measured output y, and in black the exogenous e to z transfer function of which the H_ norm is minimized.
5. ACTIVE CONTROL SYNTHESIS
5.1 Objectives Handling qualities, loads, aeroelastics and robustness requirements, as described in previous paragraph, are explicitly taken into account in the design of the control-law. Actuator-saturation requirement is not. It is verified a posteriori. Thus, the goal of the synthesis can be described as reducing the impact of wind on wing oscillations in the frequency domain of the first wing-bending mode, while not modifying the aircraft response in other frequency domains .
5.4 Evaluated performances Fig. 5 shows the root-locus of aircraji+Active Control modelled for one flight point and mass case. The minimisation of the wind to Nz1aw transfer function on the frequency range of the first bending mode is here clearly visible. Main effect is the damping of this structural mode . One notices in green the poles of the 6 th order control law, achieving the roll-on and roll-off specifications. These specifications have been achieved, as shown by the lack of impact of the Active Control on the rigid mode (Iow frequency) and on all other structural modes (high frequency) .
5.2 H_synthesis A dedicated H_-schema was designed . It contains a reduced order aircraft model, including rigid modes and part of the aircraft flexible .structural-modes. Output of this model is Nz1aw, while input is the symmetrical aileron order. See Fig. 4. The exogenous transfer function to be minimized is composed of 4 SISO transfer functions, each dedicated to one of the 4 requirements . The first one represents the transfer function from wind to Nz1aw and includes Von Karman filters that represent wind frequency distribution, and weightings for design tuning. Second and third transfer functions are dedicated to loop-shaping. They introduce a roll-on and a roll-off specifications that guarantees gains close to zero in the frequency domain outside the range of frequency of the first wing bending-mode. A fourth transfer function models the variations of the first bendingmode frequency. An appropriate weighting ensures to cover the range of possible variations in cruise.
Fig. 6 gives the theoretical wind to Nz transfer functions at three different locations. First subplot concerns Nz at the cockpit. Second one gives Nz at the centre of gravity. Third one concerns Nz at the rear of the fuselage . Last graph gives the wind to Nz1aw transfer function. One notices that the first wing bending-mode has almost no impact on the longitudinal accelerations at the front of the aircraft: resonance peaks occur only for higher frequency modes. At the middle and back of the fuselage (also noticeable on Nz1aw), impact of the first bending mode is clearly visible with a major resonance peak at frequency around 1.2 Hz for this flight point and mass case. One visualizes the performance realised by the Active Control when comparing the aircraft without specific control (k=O) and with the Active Control (k= I). The resonance peak is reduced by a factor of about 2. Sensitivity to the gain is provided via the plots with some intermediate gains ranging from k=OA to 0.9. These plots also show that the roll-on and roll-off specifications are achieved. Open-loop plots (k=O) and closed-loop plots (k>O) are similar in the frequency domains of the flight-mechanics mode (below O.5Hz) and of the higher frequency structural modes (above 2Hz), proving that the Active Control does not modify the aircraft response for these dynamics.
5.3 Controller reduction Due to aircraft rigid and flexible dynamics, Von Karman filters and weightings, the control law th obtained by the H_-synthesis is a 40 order filter . A reduction based on classical tools allows us to finally fmd a 6th order filter which fulfils all our requirements with no significant loss of performance compared to the 40 th order filter.
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Fig. 5: Root-locus of aircraft+Active Control for a feedback gain varying from 0 to I. In green are plotted the open-loop poles of the Active Control.
Fig. 6: Theoretical wind to Nz transfer functions for different values of feedback gain. k=0 corresponds to the aircraft without Active Control. First graph gives the transfer function at the cockpil, second at the middle of the fuselage, third at the rear, and fourth is the transfer function between wind and Nz 1aw •
Fig. 7 shows the theoretical time-response in presence of turbulence representative of structural sizing conditions. Plot in dashed red is the response without control, and plot in blue the response with the Active Control. One notices that without control, main harmonic is slightly above I Hz, corresponding to the first bending mode . With the Active Control this harmonic is hardly noticeable, and most noticeable harmonic seems slightly above 2 Hz. Apart from that, one notices that the peak values are reduced with the Active Control and that the total energy of the signal is also reduced, thus leading to lower loads at the wing root section.
Fig. 7: Aircraft time-response to a sizing gust.
Fig. 8 displays the time-response of the aileronsdeflection in presence of the same sizing gust. One verifies that maximum deflections remain below half the Passive Control values (plotted in dashed-red), thus allowing for sufficient roll authority for lateral manoeuvres.
Fig. 8: Ailerons-deflection in response to a sizing gust (in blue). In dashed-red are plotted the maximum values not to be exceeded.
Figures presented in this paragraph show the performances of the Active Control for one flight point and mass case, known as one of the most severe for loads. Robustness to other flight points and mass cases has been verified. Similar results in term of performances are observed but are not presented in this paper.
6. ON BOARD IMPLEMENT AnON Coding of the whole load-function (including both passive and active functions) was done using SAO at the sampling rate of 40ms. SAO is a DOl78B qualified C-code generator which automatically produces and implements on-board flying computers from engineering system block diagrams.
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Exhaustive validations of the load-function have been performed during simulator sessions and through loads and aeroelastics analysis, in order to deliver a flight clearance prior to any flight test. 7. FLIGHT TESTS
7.1 Flight-test program Flight tests were performed in October 2003. Both active and passive part of the control law were evaluated through specific tests designed to excite the aircraft and activate the load function . To test the robustness to the aircraft mass and flight points, two flights were performed. During each flight, flight points for three Mach/speed combinations were tested. These points are 0.7/280knots, 0 .8/300knots and 0.86/330knots. For each flight point the plane was excited with sinus sweep and pulse signals directly sent to the control surfaces: ailerons (inner and outer separately) and elevators, both in a symmetrical way. The sweeps were performed in open and closed loop to assess the control law performances on the damping of the first wing bending mode and the non degradation of the others flexible modes. The excitations swept from 0.5Hz up to 6Hz and the transfer function of the wing were measured into this range. The pulse signals perturbed the flight mechanics at levelled flight and roughly emulate the aircraft dynamics as it is in a light discrete gust. This resulted in the activation of the control law and the aircraft response gave a qualitative result on the control law efficiency.
Fig. 9: Transfer functions from inner ailerons to longitudinal accelerations Nz. First graph is Nz at outer engine location, second graph at wing tip . Flight point is Mach 0.86. At low frequencies (below 0.7 Hz) corresponding to the flight mechanics modes, one can verify that the activation of the load function does not have any influence (id est the red and blue lines overlay each other). The same is true at frequencies above 1.7 Hz, demonstrating compliance with the request that the law should not be active on any other structural mode apart from the first wing bending mode for robustness reasons.
During both flights, some turbulence area were encountered, and data were recorded showing the ailerons and spoilers deflections commanded by the law. Typical pitch and roll manoeuvres were also performed by the pilots to test the activation logics and the possible interactions of the law with aircraft handling qualities.
Fig. 10: Transfer functions from outer ailerons to longitudinal accelerations Nz at Mach 0.86.
7.2 Flight-test results The next four figures present the transfer functions identified in flight at two different flight points. In each figure, the first graph gives the transfer function (TF) from ailerons (inner or outer) to Nz sensor located above the outer engine. Second graph shows TF from ailerons to Nz at the wing tip. In blue is plotted the response without the load function. In red is plotted the response with the load function active . Results are very similar for all four figures . The peak in the frequency range for which the load control law is active (around 1.2Hz) is significantly shortened both in term of area and in term of maximum value. At the wing tip, the decrease is above a factor 2. At outer engine, the decrease is around 30% and the peak with the law active occurs at a higher frequency than without the law. This also contributes to a reduction of the loads.
Fig. 11 : Transfer functions from inner ailerons to longitudinal accelerations Nz at Mach 0.82. The requirements for the Active Control are therefore fulfilled , at least at the flight points and mass cases tested in flight. But simulation results provide confidence in the fulfilment of these requirements in the whole cruise flight domain. Concerning the degree of load alleviation (1 .2 Hz peak reduction), it may appear already satisfying, but one expects even better results in future studies, as the kinematics of the plane did include a low-pass Chebychev filter
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which limit the achievable bandwidth of the Active Control (see alleviation better at - 1.0 Hz than at - 1.5 Hz). This filter is mounted as standard on board the test aircraft, but it could be removed or reshaped for load dedicated controls for which there is a need for higher bandwidth.
LIDAR, gust estimators and new control surfaces (Trailing Edge Devices). The potential benefits will be integrated in the future versions of load-dedicated functions . 9. ACKNOWLEDGEMENTS
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The authors want to thank all A WIATOR partners for their cooperation in this project. The AWIATOR project is co-financed by the European Union within the Thematic Priority Aeronautics & Space of the Sixth Research Framework Programme. Special thanks also to the ONERA-DCSD for providing the tools used for the design of the Active Control.
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REFERENCES Canavin, J.R. (1978) . The Control of Spacecraft Vibrations Using Multivariable Output AIAAJAAS Astrodynamics Feedback. Conference. Palo Alto, California. August 7-9. Gilbert, M.G., D.K. Schmidt and T.A. Weisshaar (1984) . Quadratic Synthesis of Integrated Active Controls for an Aeroelastic Forward-SweptWing Aircraji. Journal of Guidance. Vo!. 7, No.
Fig. 12: Transfer functions from outer ailerons to longitudinal accelerations Nz at Mach 0.82.
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Honliner, Fr., R . Manser and D. Mussmann (1985) . Flutter Suppression and Gust Load Alleviation. Garteur Program AG3 . Kaynes, I.W. and D.E. Fry (1982) . Gust Load Alleviation on a Flexible Aircraft. Technical Report of the Royal Aircraft Establishment, June. Koenig, K. (f988). Experience in Application of Active Vibration Control Technology to a Wind Tunnel Model and to Flying Airbus. ICAS-883.10-3. Pages 1542-1553 . Smith, S.E. (1991). Name of book in italics or underlined, page or chapter numbers if relevant. Publisher, Place of publication. Livne, E. (1993). Integrated Structure/ Control/Aerodynamic Synthesis of Actively Controlled Composite Wings. Journal of Aircraft. Vo!. 30, No. 3, May-June. Mukhopadhyay, V. (1987). Digital Robust Active Control Law Synthesis for Large Order Flexible Structure using Parameter Optimization . AIAA Symposium on Dynamics and Control of Large Structures. Suzuki, S. (1993) . Simultaneous Structure/Control Design Optimization of a Wing Structure with a Gust Load Alleviation System. Journal of Aircraft. Vol. 30, No. 2, March-April. Takahashi, M. and G.L. Slater (1986). Design of a Flutter Mode Controller Using Positive Real Feedback. Journal of Guidance. Vol. 9, No. 3.
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Fig. 13 : Response to a pulse on the outer ailerons. Excitations by a pulse on the outer ailerons were also performed and one example is plotted in blue in the first graph of Fig. 13 . The black trace on the second graph shows how it stimulated the wing modes, (mainly first ~ing bending). This pulse activated the law from t=9s to 18s (red plot of 2 nd graph) . Response of the law on ailerons is plotted in green on the I SI graph. The static part of this response is plotted in red and the part due to the Active Control in light blue. The effect of the first order filter which smooths the ailerons deflection commanded by the Passive Control is noticeable, as is the effect of the anti-backup filter which ensures that the passive deflection remains at a constant level during the Nz oscillations. 8. CONCLUSION AND PROSPECTS The load-reduction function described in this paper relies on current aeronautical technologies. Its requirements were quite challenging in regard to loads, handling qualities, aeroe1astics and robustness specifications. Validations on models, through simulator sessions, and measured in flight show a significant alleviation of the bending oscillations of the wing. This function is the latest step in the process of synthesizing control-laws dedicated to load alleviation. AWIA TOR partners are working hard in developing new technologies such as load sensors,
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