- Email: [email protected]
CONCEPT FOR THE DESIGN OF FLIGHT CONTROL ACTUATION AND LAWS FOR FUTURE ALL-ELECTRIC AIRCRAFT
CONCEPT FOR THE DESIGN OF FLIGHT CONTROL ACTUATION AND LAWS FOR FUTURE ALL-ELECTRIC AIRCRAFT
Christian Schallert German Aerospace Centre (DLR) Institute of Robotics and Mechatronics 82234 Wessling, Germany
Abstract: This paper consolidates advanced techniques for flight control law design with the sizing issues of new technology electrical flight control actuators. An automatic optimisation tool is adapted to modify the control laws for minimum actuator deflection rates, which can be turned into weight savings for the overall aircraft. Thus, the paper outlines a new and up-to-date trade-off study, which is motivated by the recent trend towards the more-electric aircraft concept. Copyright © 2007 IFAC Keywords: flight control, actuators, electrical power system, flight dynamics simulation, control laws, automatic multi-objective parameter optimisation
1. INTRODUCTION In future aircraft, hydraulic actuators for primary flight control surfaces will be increasingly replaced with electrically powered actuators (EHA type, Fig. 1), according to the trend towards a greater electrification of aircraft systems. The actuator performance requirements are affected by the flight control laws, which must fulfil a wide range of flying quality requirements. As compared with traditional hydraulic actuators, the weight and size penalty for an EHA dependent on its power capability, namely the deflection rate, is considerably higher. Thus, it becomes crucial to consider the required actuator power directly in the flight control law design, so that acceptable trade-offs with respect to flying qualities, stability margins and system weight can be searched. The method of multi-criteria optimisation allows to find such trade-offs in an objective and automated way. First, the paper introduces the more-electric aircraft conecpt, its motivation and the related actuator technology. Then, it explains how reducing the actuator rate requirements will eventually allow to fit electriFig. 2: Flight Control System Sketch
Fig. 1: Electro-Hydrostatic Actuator (EHA)
cally powered actuators on all control surfaces, and how this can be translated into weight savings for the overall aircraft. Next, a process is outlined to modify the flight control laws for reduced actuator rates. It will be demonstrated for a generic flight dynamics simulation model of a large transport aircraft. An automatic optimisation tool is being adapted such that it interfaces with the flight dynamics simulation, in order to tune the flight control laws. Multiple, partly conflicting flying qualities criteria are implemented that the optimisation tool must balance and trade off. Complementary, a desktop flight simulator is used for real-time and interactive testing of the adapted control laws and actuator dynamics. Thus, the paper consolidates advanced techniques and tools for flight control law design with the sizing issues of new technology electrical actuators. It outlines a novel and up-to-date trade-off study, which is motivated by the recent trend towards the moreelectric aircraft concept and the aim to save weight. Currently, the concept for the studies is depicted, whereas results will be available at a later stage.
2. THE MORE-/ ALL-ELECTRIC AIRCRAFT CONCEPTS Today, transport aircraft systems are powered by hydraulic, electric and pneumatic supplies. The next generation of transport aircraft, namely the A380, A400M and B787, is evolved towards the moreelectric aircraft (MEA) concept: The conventional pneumatic and hydraulic power supplies for the consumers, such as the environmental control system, flight control actuation, landing gears and brakes, are replaced step-by-step by electrical power supplies and associated new developed consumers.
and pneumatic – each one has to be sized for the peak demands of its individual consumers. In the AEA concept, only one though highly redundant electrical power system has to be sized for the consumer peak demands. The fact that all electrical consumers are collected on one (but redundant) power system, instead of three different, offers more flexibility in distributing the consumers and applying ‘load management’: Large but non-critical consumers may be intermittently reduced in power or even suspended. This, together with the capability of electrical generators to overload for a short period of time, allows to flatten out the power peaks. That way, the optimum use is made of the electrical power supply system capability. Accordingly, the component (generators, contactors, wiring, transformers etc.) sizes can be minimised and, in turn, the overall system weight. Although electrical technologies tend to be heavier, the AEA concept may translate this way into a weight benefit for the aircraft, because the pneumatic and hydraulic supplies are replaced by one enlarged but optimised electrical power system (Schallert, et al., 2006).
Fig. 3: Present Aircraft Systems Architecture Development of electrical power supply and consumer systems, as well as investigations of the moreelectric and all-electric aircraft (AEA) concepts have been performed in the European project Power Optimised Aircraft (POA), which is currently being continued by the More-Open Electrical Technologies (MOET) project. The following trade-offs associated with the electrification of aircraft systems were identified in POA: Electrical systems are more energy efficient, leading to reduced secondary off-takes from the engines. On the other hand, electrical systems tend to be heavier than their conventional equivalents. Consequently, the MEA or AEA concept leads to a higher total aircraft weight than for the conventional one. Although heavier, the new aircraft concepts still enable a reduction of about 2% in fuel consumption, thanks to the better energy efficiency of electrical systems. An economical assessment of the MEA and AEA concepts has shown that further benefit can be expected from improved system reliability and facilitated maintenance, which complement to the decreased fuel burn in leading to reduced aircraft operating cost (Faleiro, 2006). Moving to the AEA concept has another advantage in terms of sizing the on-board power supply system. For a conventional aircraft equipped with three different types of power supplies – electric, hydraulic
Fig. 4: A Potential Future All-Electric Aircraft Systems Architecture
3. FLIGHT CONTROL ACTUATORS FOR THE MORE-/ ALL-ELECTRIC AIRCRAFT Regarding the flight controls of a more- or allelectric aircraft, the control surfaces are equipped with electrically powered actuators to an increasing degree, to replace the conventional hydraulic actuators. In comparison, however, the so called electrohydrostatic actuators (EHAs) have a larger space envelope and weight as a drawback. This is caused by the higher complexity associated with EHAs (see Fig. 5 and Fig. 6) and the comparatively smaller power density of electric motor drives.
The weight of a hydraulical actuator depends mainly on its maximum load and the supply pressure. The weight of an EHA is linked also to the required power, i.e. the product of the maximum load and deflection rate. Thus, reducing the rate of electrically powered actuators permits to save weight and to scale down the actuator space envelope. Eventually, this will enable to fit EHAs into all of the confined cross sections on the outer wing and tailplane, which currently are fitted mostly with hydraulic actuators.
aircraft or entering turbulence. In the case of an EHA, this kind of actuator power behaviour goes well with the characteristics of an electrical power supply system, that can be overloaded intermittently to cover short peak demands (refer to chapter 2). to hydraulic supply servovalve
to electric supply power electronics motor and pump
reservoir hydraulic cylinder
Fig. 6: Hydraulic Actuator Functional Scheme
hydraulic cylinder
Fig. 5: Electro-Hydrostatic Actuator (EHA) Functional Scheme The introduction of EHAs on some flight control surfaces of the A380 aircraft has allowed to eliminate one of the three usually installed central hydraulic systems. Although an EHA is about twice as heavy as a hydraulic actuator of comparable performance, the enabled elimination of a hydraulic system still results in a very significant weight saving for the aircraft (van den Bossche, 2006). If all control surfaces are equipped with EHAs, allowed by reduced rate requirements, then the central hydraulic supplies can be removed from the aircraft completely. This also requires that the other former hydraulic consumers, such as the high lift system, the landing gears and brakes are equipped with local hydraulics or electro-mechanical drives as well (refer to chapter 2). Other than an EHA, the weight and size of a hydraulic actuator cannot be scaled by a changed rate requirement, since the only component affected is the servovalve, which is relatively small and light anyway. Reducing the rate requirements would decrease the hydraulic flow demands, still enabling a slight down scaling of the hydraulic piping to the wings and empennage. Nevertheless, the sizes of the bigger part of the hydraulic systems on a conventional aircraft are determined by the short-term but large peak demands caused by the landing gears, brakes and high lift systems, which exceed the demands of hydraulic flight control actuators. In terms of power demand, both types of flight control actuators (hydraulic and EHA) have a constant and relatively low consumption caused by internal leakage, for holding a surface in position against load. Short and intermittent power demand peaks occur due to surface motion when manoeuvring the
In conclusion, if electrically powered flight control actuators are introduced in the transition over to the MEA or AEA concept, it is more beneficial and yet necessary to challenge the actuator rates, than it is for a conventional aircraft equipped with hydraulic flight control actuators and central hydraulic supplies.
4. FLIGHT AND CONTROL SYSTEM DYNAMICS MODELLING A generic model of a large transport aircraft is adopted that contains the physical nonlinear rigid body equations of motion and aerodynamics (Moormann, 2001). It was developed in the object-oriented modelling language Modelica at the German Aerospace Centre, Institute of Robotics and Mechatronics (DLR-RM). In Fig. 7 showing a simplified control structure for the pitch axis, this model is represented by the block A/C. The flight control system model is composed of the control laws, a software rate limitation and a delay. The delay represents the inevitable reaction time caused by the electronic flight control system (EFCS). The purpose of the software rate limitation is to prevent the actuators from reaching their hardware rate limitation. The actuator model (EHA) contains the physics of the hydro-mechanical and electrical components (cylinder, piston, fluid, motor and pump) and the position controller. Thus, the actuator behaviour relevant to flight dynamics is represented (response to commands, load and rate performance, rate limit). All models mentioned above are merged such that a comprehensive flight dynamics model is created. This integrated model is then used for investigating the influence on the flying qualities caused by modifications of the actuator dynamics (rate limitation), control laws and EFCS reaction time (delay).
The investigation has been started for a manual control of the pitch axis in the landing configuration, and it will be continued to include as well the roll and yaw axes of the aircraft. As depicted, the pilot may achieve a closure of the outer loop for control of the pitch angle θ (attitude). The behaviour in the landing configuration is of particular interest, since manual flight at low airspeed will require the highest flight control surface and actuator deflection rates.
Exemplarily, the bandwidth is designated as the maximum frequency, up to which the pilot can achieve closed loop control of the aircraft without compromising stability. At the bandwidth frequency, the gain margin GM and phase margin ϕM of the stick input to pitch angle dynamics Fδθes ( jω) are at least 6dB and 45°, respectively. This is depicted in Fig. 8. time-domain criteria • rise time and settling time
Fig. 7: Control of the Pitch Axis The a/c behaviour due to non-harmonic commands, such as step- or block-shaped stick inputs, is assessed by time simulation of the complete non-linear flight dynamics model and suitable time-domain criteria. 5. FLYING QUALITIES EVALUATION
The flying and handling qualities of the modelled aircraft, including the flight control laws, are assessed quantitatively by means of suitable and representative criteria. Many different criteria can be found the literature. In the following, a couple of the criteria selected and implemented for the described investigation are introduced.
Exemplarily, the rise time and settling time criterion was conceived by Mooij (1984) especially for the longitudinal dynamics of transport aircraft in landing configuration. The rise time Trise designates the time span until the normalised pitch rate q/qSS is greater than 0.9⋅qSS, following a step-shaped stick input and qSS being the quasi-stationary pitch rate. Then, the settling time Tsettle designates the time span until the normalised pitch rate q/qSS remains within a zone from 0.9⋅qSS to 1.1⋅qSS. Trise and Tsettle have to be shorter than specific values, as denoted in Fig. 9.
frequency-domain criteria • bandwidth • gain margin and phase margin • neal-smith
For further relevant aspects, such as flight in turbulence or cross winds, accordant criteria on actuator control activity and the rejection of attitude disturbances will be added.
The listed frequency-domain criteria, which can also be referred to in MIL-HDBK-1797 (1997), are concerned with the ability of the pilot and associated workload to achieve manual control of the aircraft, as well as the stability margins. Naturally, the frequency-domain flying qualities analysis only draws upon the linearised flight dynamics and the response to small harmonic input signals. Non-linear effects, such as actuator rate limitation, have to be investigated separately.
non-linear stability The effect of nonlinearities on the stability of the pilot-vehicle control loop has to be analysed. Besides other factors, actuator rate limitation is associated with the possible occurrence of unacceptable limit cycles of the pilot-vehicle control loop, known as pilot-induced oscillations (PIO) or, in more proper terms, aircraft-pilot coupling (APC).
5.1 Criteria for Quantitative Flying Qualities Assessment
A procedure suited for the prediction of APC tendencies has been compiled by Duda (1997). Again flight at low airspeed, i.e. on landing approach, must be studied carefully, since the low dynamic pressure leads to large surface deflections and rates, as well as the precision tracking performed by the pilot. The analysis of past APC incidents has shown, that this combination of influences often was the trigger (not the cause) for such events on an APC-prone aircraft.
RM. The real-time capability of the entire flight dynamics model must be ensured for using it with the interactive flight simulation. By the use of interactive flight simulation and visualisation, the aircraft and flight path responses due to manual commands can be viewed and assessed on the engineer’s desktop. This way, the understanding of the influence of flight control law changes is improved considerably.
6. FLIGHT CONTROL LAW ADAPTATION PROCESS AND TOOLS
Fig. 8: Frequency response of the of the linearised flight dynamics from stick input δes to pitch angle θ. Bandwidth ωBW is displayed.
For the process of adapting the flight control laws for reduced actuator rates, the following items (refer to Fig. 7) are varied: • flight control law structure and parameters (gains, time constants) • S/W rate limiter structure and parameters • EFCS reaction time Changes of the EFCS reaction time have to be considered carefully, since it is related to the EFCS architecture, which has to comply with vital safety and redundancy requirements. Therefore, a part of the sensitivity study is performed with an unchanged EFCS reaction time. In the adaptation process, the flight control law structure, the actuator rate limit and EFCS reaction time are selected by the user. For each selection, the control law parameters have to be optimised newly. This is performed automatically by means of a Matlabbased tool named MOPS (Multi-Objective Parameter Synthesis), which interfaces with the flight dynamics model implemented in Modelica.
Fig. 9: Normalised pitch rate q/qSS response following a step-shaped stick input. Rise time and settling time are displayed. 5.2 Interactive Desktop Flight Simulation Complementary to the evaluation by analytical flying qualities criteria, an interactive desktop flight simulator is used for qualitative checks of the flight dynamics after control law modifications (Looye, 2006). A visualisation interface and a joystick for manual inputs are added to the flight dynamics simulation model (refer to chapter 4). The visualisation interface allows various views on and out of the aircraft during the simulation, as well as the display of cockpit instruments (indication of speed, altitude, heading etc.). The 3-D desktop visualisation tool and interface are developed by the company AeroLabs AG (www. aerolabs.de) based on specifications of DLR-
Fig. 10: Parallel coordinates plot of flying qualities criteria, several iterations performed by MOPS The optimisation of the flight control laws is a multiparameter and multi-criteria problem: Several gains and time constants have to be tuned, in order to fulfill and trade off several conflicting flying qualities requirements (see Fig. 10 and chapter 5.1). MOPS has
been developed at DLR-RM for the automated solving of multi-parameter, multi-criteria and multi-case optimisation problems (Joos, et al., 2002). In each iteration of the automatic loop, MOPS simulates and linearises the flight dynamics model, then evaluates the implemented flying qualities criteria and changes the gains and time constants of the selected flight control law, as shown by Fig. 11. The process is continued, as long as the automatic optimisation algorithm finds better solutions. The achievable level of flying qualities is compared for different combinations of actuator rate limits and control law structures, on the basis of optimised control law parameters for each selected combination.
Proper flying qualities and stability margins have to be maintained, which restricts the decrease of the actuator rate requirements. Complementary to the flying qualities evaluation by analytical criteria, the effect of actuator dynamics and control law changes is assessed by interactive testing on a real-time desktop flight simulator. Reducing the flight control actuator rate requirements minimises the sizes and weight of EHAs, which will eventually permit to equip all primary flight control surfaces with EHAs and to completely remove the hydraulic systems. This will lead to weight savings at aircraft level, although an EHA is clearly heavier than its equivalent hydraulic actuator. The enabled weight savings will be determined for the overall aircraft, to account for all actuation systems affected by the removal of the hydraulic systems. In conclusion, a contribution is made on the way towards the all-electric aircraft.
REFERENCES
Fig. 11: Flight Control Law Adaptation Process
7. SUMMARY The investigation outlined in this paper has been motivated by the recent trend towards the moreelectric aircraft concept. Other than for conventional hydraulic actuators, it is beneficial and yet necessary to challenge the actuator rate requirements for electrically powered flight control actuators (e.g. EHAs), due to the noticeable impact on the actuator sizing. The paper described the process and tools arranged to study the possible trade-offs enabled by adapting the flight control system and laws for lower actuator rates. The investigation draws upon a physical flight dynamics and aerodynamics model of a large transport aircraft, as well as models of the flight control system and actuators. Whereas the flight control law structure is created manually, the respective parameters are optimised by means of an automated setup of simulation and optimisation tools. This combined tool setup uses the flight dynamics model and implementations of the time domain and frequency domain flying qualities criteria, which must be evaluated and balanced in the automatic optimisation, in searching for the best set of control law parameters.
Faleiro L. (2006). Summary of the European Power Optimised Aircraft (POA) Project. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 4th, 2006. van den Bossche D. (2006). The A380 Flight Control Electrohydrostatic Actuators, Achievements and Lessons Learnt. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 5th, 2006. Schallert C., Pfeiffer A. and Bals J. (2006). Generator Power Optimisation for a More-Electric Aircraft by Use of a Virtual Iron Bird. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 5th, 2006. Moormann D. (2001). Automatisierte Modellbildung der Flugsystemdynamik. VDI Fortschrittberichte Reihe 8, Nr. 931, ISBN 3-18-393108-7. Unknown (1997). Flying Qualities of Piloted Aircraft. MIL-HDBK-1797, U. S. Department of Defense. Mooij H. A. (1984). Criteria for low-speed longitudinal handling qualities of transport aircraft with closed-loop flight control systems. National Aerospace Laboratory NLR, The Netherlands. Duda H. (1997). Fliegbarkeitskriterien bei begrenzter Stellgeschwindigkeit. Deutsche Forschungsanstalt für Luft- u. Raumfahrt DLR e. V. Looye G. (2006). Advanced Modelling and Simulation Techniques in the Flight Control Law Design Process. 17th IFAC Symposium, June 2007. Joos H.-D., Bals J., Looye G., Schnepper K., Varga A. (2002). A multi-objective optimisation-based software environment for control systems design. IEEE International Conference on Control Applications and International Symposium on Computer Aided Control Systems Design, Glasgow, Scotland, UK, September 18-20, 2002, Proc. of CCA/CACSD 2002, P. 7-14, 2002.
Christian Schallert German Aerospace Centre (DLR) Institute of Robotics and Mechatronics 82234 Wessling, Germany
Abstract: This paper consolidates advanced techniques for flight control law design with the sizing issues of new technology electrical flight control actuators. An automatic optimisation tool is adapted to modify the control laws for minimum actuator deflection rates, which can be turned into weight savings for the overall aircraft. Thus, the paper outlines a new and up-to-date trade-off study, which is motivated by the recent trend towards the more-electric aircraft concept. Copyright © 2007 IFAC Keywords: flight control, actuators, electrical power system, flight dynamics simulation, control laws, automatic multi-objective parameter optimisation
1. INTRODUCTION In future aircraft, hydraulic actuators for primary flight control surfaces will be increasingly replaced with electrically powered actuators (EHA type, Fig. 1), according to the trend towards a greater electrification of aircraft systems. The actuator performance requirements are affected by the flight control laws, which must fulfil a wide range of flying quality requirements. As compared with traditional hydraulic actuators, the weight and size penalty for an EHA dependent on its power capability, namely the deflection rate, is considerably higher. Thus, it becomes crucial to consider the required actuator power directly in the flight control law design, so that acceptable trade-offs with respect to flying qualities, stability margins and system weight can be searched. The method of multi-criteria optimisation allows to find such trade-offs in an objective and automated way. First, the paper introduces the more-electric aircraft conecpt, its motivation and the related actuator technology. Then, it explains how reducing the actuator rate requirements will eventually allow to fit electriFig. 2: Flight Control System Sketch
Fig. 1: Electro-Hydrostatic Actuator (EHA)
cally powered actuators on all control surfaces, and how this can be translated into weight savings for the overall aircraft. Next, a process is outlined to modify the flight control laws for reduced actuator rates. It will be demonstrated for a generic flight dynamics simulation model of a large transport aircraft. An automatic optimisation tool is being adapted such that it interfaces with the flight dynamics simulation, in order to tune the flight control laws. Multiple, partly conflicting flying qualities criteria are implemented that the optimisation tool must balance and trade off. Complementary, a desktop flight simulator is used for real-time and interactive testing of the adapted control laws and actuator dynamics. Thus, the paper consolidates advanced techniques and tools for flight control law design with the sizing issues of new technology electrical actuators. It outlines a novel and up-to-date trade-off study, which is motivated by the recent trend towards the moreelectric aircraft concept and the aim to save weight. Currently, the concept for the studies is depicted, whereas results will be available at a later stage.
2. THE MORE-/ ALL-ELECTRIC AIRCRAFT CONCEPTS Today, transport aircraft systems are powered by hydraulic, electric and pneumatic supplies. The next generation of transport aircraft, namely the A380, A400M and B787, is evolved towards the moreelectric aircraft (MEA) concept: The conventional pneumatic and hydraulic power supplies for the consumers, such as the environmental control system, flight control actuation, landing gears and brakes, are replaced step-by-step by electrical power supplies and associated new developed consumers.
and pneumatic – each one has to be sized for the peak demands of its individual consumers. In the AEA concept, only one though highly redundant electrical power system has to be sized for the consumer peak demands. The fact that all electrical consumers are collected on one (but redundant) power system, instead of three different, offers more flexibility in distributing the consumers and applying ‘load management’: Large but non-critical consumers may be intermittently reduced in power or even suspended. This, together with the capability of electrical generators to overload for a short period of time, allows to flatten out the power peaks. That way, the optimum use is made of the electrical power supply system capability. Accordingly, the component (generators, contactors, wiring, transformers etc.) sizes can be minimised and, in turn, the overall system weight. Although electrical technologies tend to be heavier, the AEA concept may translate this way into a weight benefit for the aircraft, because the pneumatic and hydraulic supplies are replaced by one enlarged but optimised electrical power system (Schallert, et al., 2006).
Fig. 3: Present Aircraft Systems Architecture Development of electrical power supply and consumer systems, as well as investigations of the moreelectric and all-electric aircraft (AEA) concepts have been performed in the European project Power Optimised Aircraft (POA), which is currently being continued by the More-Open Electrical Technologies (MOET) project. The following trade-offs associated with the electrification of aircraft systems were identified in POA: Electrical systems are more energy efficient, leading to reduced secondary off-takes from the engines. On the other hand, electrical systems tend to be heavier than their conventional equivalents. Consequently, the MEA or AEA concept leads to a higher total aircraft weight than for the conventional one. Although heavier, the new aircraft concepts still enable a reduction of about 2% in fuel consumption, thanks to the better energy efficiency of electrical systems. An economical assessment of the MEA and AEA concepts has shown that further benefit can be expected from improved system reliability and facilitated maintenance, which complement to the decreased fuel burn in leading to reduced aircraft operating cost (Faleiro, 2006). Moving to the AEA concept has another advantage in terms of sizing the on-board power supply system. For a conventional aircraft equipped with three different types of power supplies – electric, hydraulic
Fig. 4: A Potential Future All-Electric Aircraft Systems Architecture
3. FLIGHT CONTROL ACTUATORS FOR THE MORE-/ ALL-ELECTRIC AIRCRAFT Regarding the flight controls of a more- or allelectric aircraft, the control surfaces are equipped with electrically powered actuators to an increasing degree, to replace the conventional hydraulic actuators. In comparison, however, the so called electrohydrostatic actuators (EHAs) have a larger space envelope and weight as a drawback. This is caused by the higher complexity associated with EHAs (see Fig. 5 and Fig. 6) and the comparatively smaller power density of electric motor drives.
The weight of a hydraulical actuator depends mainly on its maximum load and the supply pressure. The weight of an EHA is linked also to the required power, i.e. the product of the maximum load and deflection rate. Thus, reducing the rate of electrically powered actuators permits to save weight and to scale down the actuator space envelope. Eventually, this will enable to fit EHAs into all of the confined cross sections on the outer wing and tailplane, which currently are fitted mostly with hydraulic actuators.
aircraft or entering turbulence. In the case of an EHA, this kind of actuator power behaviour goes well with the characteristics of an electrical power supply system, that can be overloaded intermittently to cover short peak demands (refer to chapter 2). to hydraulic supply servovalve
to electric supply power electronics motor and pump
reservoir hydraulic cylinder
Fig. 6: Hydraulic Actuator Functional Scheme
hydraulic cylinder
Fig. 5: Electro-Hydrostatic Actuator (EHA) Functional Scheme The introduction of EHAs on some flight control surfaces of the A380 aircraft has allowed to eliminate one of the three usually installed central hydraulic systems. Although an EHA is about twice as heavy as a hydraulic actuator of comparable performance, the enabled elimination of a hydraulic system still results in a very significant weight saving for the aircraft (van den Bossche, 2006). If all control surfaces are equipped with EHAs, allowed by reduced rate requirements, then the central hydraulic supplies can be removed from the aircraft completely. This also requires that the other former hydraulic consumers, such as the high lift system, the landing gears and brakes are equipped with local hydraulics or electro-mechanical drives as well (refer to chapter 2). Other than an EHA, the weight and size of a hydraulic actuator cannot be scaled by a changed rate requirement, since the only component affected is the servovalve, which is relatively small and light anyway. Reducing the rate requirements would decrease the hydraulic flow demands, still enabling a slight down scaling of the hydraulic piping to the wings and empennage. Nevertheless, the sizes of the bigger part of the hydraulic systems on a conventional aircraft are determined by the short-term but large peak demands caused by the landing gears, brakes and high lift systems, which exceed the demands of hydraulic flight control actuators. In terms of power demand, both types of flight control actuators (hydraulic and EHA) have a constant and relatively low consumption caused by internal leakage, for holding a surface in position against load. Short and intermittent power demand peaks occur due to surface motion when manoeuvring the
In conclusion, if electrically powered flight control actuators are introduced in the transition over to the MEA or AEA concept, it is more beneficial and yet necessary to challenge the actuator rates, than it is for a conventional aircraft equipped with hydraulic flight control actuators and central hydraulic supplies.
4. FLIGHT AND CONTROL SYSTEM DYNAMICS MODELLING A generic model of a large transport aircraft is adopted that contains the physical nonlinear rigid body equations of motion and aerodynamics (Moormann, 2001). It was developed in the object-oriented modelling language Modelica at the German Aerospace Centre, Institute of Robotics and Mechatronics (DLR-RM). In Fig. 7 showing a simplified control structure for the pitch axis, this model is represented by the block A/C. The flight control system model is composed of the control laws, a software rate limitation and a delay. The delay represents the inevitable reaction time caused by the electronic flight control system (EFCS). The purpose of the software rate limitation is to prevent the actuators from reaching their hardware rate limitation. The actuator model (EHA) contains the physics of the hydro-mechanical and electrical components (cylinder, piston, fluid, motor and pump) and the position controller. Thus, the actuator behaviour relevant to flight dynamics is represented (response to commands, load and rate performance, rate limit). All models mentioned above are merged such that a comprehensive flight dynamics model is created. This integrated model is then used for investigating the influence on the flying qualities caused by modifications of the actuator dynamics (rate limitation), control laws and EFCS reaction time (delay).
The investigation has been started for a manual control of the pitch axis in the landing configuration, and it will be continued to include as well the roll and yaw axes of the aircraft. As depicted, the pilot may achieve a closure of the outer loop for control of the pitch angle θ (attitude). The behaviour in the landing configuration is of particular interest, since manual flight at low airspeed will require the highest flight control surface and actuator deflection rates.
Exemplarily, the bandwidth is designated as the maximum frequency, up to which the pilot can achieve closed loop control of the aircraft without compromising stability. At the bandwidth frequency, the gain margin GM and phase margin ϕM of the stick input to pitch angle dynamics Fδθes ( jω) are at least 6dB and 45°, respectively. This is depicted in Fig. 8. time-domain criteria • rise time and settling time
Fig. 7: Control of the Pitch Axis The a/c behaviour due to non-harmonic commands, such as step- or block-shaped stick inputs, is assessed by time simulation of the complete non-linear flight dynamics model and suitable time-domain criteria. 5. FLYING QUALITIES EVALUATION
The flying and handling qualities of the modelled aircraft, including the flight control laws, are assessed quantitatively by means of suitable and representative criteria. Many different criteria can be found the literature. In the following, a couple of the criteria selected and implemented for the described investigation are introduced.
Exemplarily, the rise time and settling time criterion was conceived by Mooij (1984) especially for the longitudinal dynamics of transport aircraft in landing configuration. The rise time Trise designates the time span until the normalised pitch rate q/qSS is greater than 0.9⋅qSS, following a step-shaped stick input and qSS being the quasi-stationary pitch rate. Then, the settling time Tsettle designates the time span until the normalised pitch rate q/qSS remains within a zone from 0.9⋅qSS to 1.1⋅qSS. Trise and Tsettle have to be shorter than specific values, as denoted in Fig. 9.
frequency-domain criteria • bandwidth • gain margin and phase margin • neal-smith
For further relevant aspects, such as flight in turbulence or cross winds, accordant criteria on actuator control activity and the rejection of attitude disturbances will be added.
The listed frequency-domain criteria, which can also be referred to in MIL-HDBK-1797 (1997), are concerned with the ability of the pilot and associated workload to achieve manual control of the aircraft, as well as the stability margins. Naturally, the frequency-domain flying qualities analysis only draws upon the linearised flight dynamics and the response to small harmonic input signals. Non-linear effects, such as actuator rate limitation, have to be investigated separately.
non-linear stability The effect of nonlinearities on the stability of the pilot-vehicle control loop has to be analysed. Besides other factors, actuator rate limitation is associated with the possible occurrence of unacceptable limit cycles of the pilot-vehicle control loop, known as pilot-induced oscillations (PIO) or, in more proper terms, aircraft-pilot coupling (APC).
5.1 Criteria for Quantitative Flying Qualities Assessment
A procedure suited for the prediction of APC tendencies has been compiled by Duda (1997). Again flight at low airspeed, i.e. on landing approach, must be studied carefully, since the low dynamic pressure leads to large surface deflections and rates, as well as the precision tracking performed by the pilot. The analysis of past APC incidents has shown, that this combination of influences often was the trigger (not the cause) for such events on an APC-prone aircraft.
RM. The real-time capability of the entire flight dynamics model must be ensured for using it with the interactive flight simulation. By the use of interactive flight simulation and visualisation, the aircraft and flight path responses due to manual commands can be viewed and assessed on the engineer’s desktop. This way, the understanding of the influence of flight control law changes is improved considerably.
6. FLIGHT CONTROL LAW ADAPTATION PROCESS AND TOOLS
Fig. 8: Frequency response of the of the linearised flight dynamics from stick input δes to pitch angle θ. Bandwidth ωBW is displayed.
For the process of adapting the flight control laws for reduced actuator rates, the following items (refer to Fig. 7) are varied: • flight control law structure and parameters (gains, time constants) • S/W rate limiter structure and parameters • EFCS reaction time Changes of the EFCS reaction time have to be considered carefully, since it is related to the EFCS architecture, which has to comply with vital safety and redundancy requirements. Therefore, a part of the sensitivity study is performed with an unchanged EFCS reaction time. In the adaptation process, the flight control law structure, the actuator rate limit and EFCS reaction time are selected by the user. For each selection, the control law parameters have to be optimised newly. This is performed automatically by means of a Matlabbased tool named MOPS (Multi-Objective Parameter Synthesis), which interfaces with the flight dynamics model implemented in Modelica.
Fig. 9: Normalised pitch rate q/qSS response following a step-shaped stick input. Rise time and settling time are displayed. 5.2 Interactive Desktop Flight Simulation Complementary to the evaluation by analytical flying qualities criteria, an interactive desktop flight simulator is used for qualitative checks of the flight dynamics after control law modifications (Looye, 2006). A visualisation interface and a joystick for manual inputs are added to the flight dynamics simulation model (refer to chapter 4). The visualisation interface allows various views on and out of the aircraft during the simulation, as well as the display of cockpit instruments (indication of speed, altitude, heading etc.). The 3-D desktop visualisation tool and interface are developed by the company AeroLabs AG (www. aerolabs.de) based on specifications of DLR-
Fig. 10: Parallel coordinates plot of flying qualities criteria, several iterations performed by MOPS The optimisation of the flight control laws is a multiparameter and multi-criteria problem: Several gains and time constants have to be tuned, in order to fulfill and trade off several conflicting flying qualities requirements (see Fig. 10 and chapter 5.1). MOPS has
been developed at DLR-RM for the automated solving of multi-parameter, multi-criteria and multi-case optimisation problems (Joos, et al., 2002). In each iteration of the automatic loop, MOPS simulates and linearises the flight dynamics model, then evaluates the implemented flying qualities criteria and changes the gains and time constants of the selected flight control law, as shown by Fig. 11. The process is continued, as long as the automatic optimisation algorithm finds better solutions. The achievable level of flying qualities is compared for different combinations of actuator rate limits and control law structures, on the basis of optimised control law parameters for each selected combination.
Proper flying qualities and stability margins have to be maintained, which restricts the decrease of the actuator rate requirements. Complementary to the flying qualities evaluation by analytical criteria, the effect of actuator dynamics and control law changes is assessed by interactive testing on a real-time desktop flight simulator. Reducing the flight control actuator rate requirements minimises the sizes and weight of EHAs, which will eventually permit to equip all primary flight control surfaces with EHAs and to completely remove the hydraulic systems. This will lead to weight savings at aircraft level, although an EHA is clearly heavier than its equivalent hydraulic actuator. The enabled weight savings will be determined for the overall aircraft, to account for all actuation systems affected by the removal of the hydraulic systems. In conclusion, a contribution is made on the way towards the all-electric aircraft.
REFERENCES
Fig. 11: Flight Control Law Adaptation Process
7. SUMMARY The investigation outlined in this paper has been motivated by the recent trend towards the moreelectric aircraft concept. Other than for conventional hydraulic actuators, it is beneficial and yet necessary to challenge the actuator rate requirements for electrically powered flight control actuators (e.g. EHAs), due to the noticeable impact on the actuator sizing. The paper described the process and tools arranged to study the possible trade-offs enabled by adapting the flight control system and laws for lower actuator rates. The investigation draws upon a physical flight dynamics and aerodynamics model of a large transport aircraft, as well as models of the flight control system and actuators. Whereas the flight control law structure is created manually, the respective parameters are optimised by means of an automated setup of simulation and optimisation tools. This combined tool setup uses the flight dynamics model and implementations of the time domain and frequency domain flying qualities criteria, which must be evaluated and balanced in the automatic optimisation, in searching for the best set of control law parameters.
Faleiro L. (2006). Summary of the European Power Optimised Aircraft (POA) Project. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 4th, 2006. van den Bossche D. (2006). The A380 Flight Control Electrohydrostatic Actuators, Achievements and Lessons Learnt. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 5th, 2006. Schallert C., Pfeiffer A. and Bals J. (2006). Generator Power Optimisation for a More-Electric Aircraft by Use of a Virtual Iron Bird. 25th International Congress of the Aeronautical Sciences (ICAS), Sep. 5th, 2006. Moormann D. (2001). Automatisierte Modellbildung der Flugsystemdynamik. VDI Fortschrittberichte Reihe 8, Nr. 931, ISBN 3-18-393108-7. Unknown (1997). Flying Qualities of Piloted Aircraft. MIL-HDBK-1797, U. S. Department of Defense. Mooij H. A. (1984). Criteria for low-speed longitudinal handling qualities of transport aircraft with closed-loop flight control systems. National Aerospace Laboratory NLR, The Netherlands. Duda H. (1997). Fliegbarkeitskriterien bei begrenzter Stellgeschwindigkeit. Deutsche Forschungsanstalt für Luft- u. Raumfahrt DLR e. V. Looye G. (2006). Advanced Modelling and Simulation Techniques in the Flight Control Law Design Process. 17th IFAC Symposium, June 2007. Joos H.-D., Bals J., Looye G., Schnepper K., Varga A. (2002). A multi-objective optimisation-based software environment for control systems design. IEEE International Conference on Control Applications and International Symposium on Computer Aided Control Systems Design, Glasgow, Scotland, UK, September 18-20, 2002, Proc. of CCA/CACSD 2002, P. 7-14, 2002.