Military Aircraft

Emerging technologies in electrical drives and power distribution systems in future aircrafts

Emerging technologies in electrical drives and power distribution systems in future aircrafts

                       u.sureshkumar*

                 

             *  professor in EEE department

                                      Mohamed sathak enginnering college

                                      Kilakkarai

                                      E mail :uskrk@sify.com

 

 

 

 

 

 

Abstract:

 

It is projected that in future aircraft, all power, except propulsion, will be distributed and processed electrically. In other words, electrical power will be utilized for driving aircraft subsystems currently powered by hydraulic, pneumatic or mechanical means including utility and flight control actuation, environmental control system, lubrication and fuel pumps, and numerous other utility functions. These concepts are embraced by what is known as the “More Electric Aircraft (MEA)” initiative. The MEA emphasizes the utilization of electrical power as opposed to hydraulic, pneumatic, and mechanical power for optimizing aircraft performance and life cycle cost. It would eliminate the need for gearboxes and transmissions since the power transmission is through electrical rather than mechanical means, which reduces the weight of the aircraft and increases the fuel efficiency. Detailed analysis of

interaction between an Electro Mechanical Actuator (EMA) connected to the DC bus of the power distribution system in a next generation transport aircraft with the bus regulator is presented. Development of reliable power-by-wire actuation systems for both aeronautical and space applications has been sought recently to eliminate hydraulic systems from aircraft. Aircraft

engineers have tested electrohydrostatic actuators (EHAs), which combine electrical and hydraulic power, hence the

evolutionary "more electric aircraft" idea. Efforts are being made to replace

all the hydraulic systems with electrical systems, which will lead to a new technology called “All Electric Aircraft”.

 

Introduction:

Flight Control System

A flight control system consists of the flight control surfaces, the respective cockpit controls, connecting linkage, and necessary operating mechanism to cont4rol aircraft in flight.

Classification

Flight control systems (FCS) are classified as follows:

• Mechanical FCS
• Hydro mechanical FCS (powered flight control units (PFCU))
• Fly-by-wire FCS

Mechanical FCS:

The mechanical FCS is the most basic designs. They were used in early aircraft and currently in small aeroplanes where the aerodynamic forces are not excessive. The FCS uses a collection of mechanical parts such as rods, cables, pulleys and sometimes chains to transmit the forces of the cockpit controls to the control surfaces.

 

Hydro mechanical FCS (powered flight control units (PFCU)):

The complexity and weight of a mechanical FCS increases considerably with size and performance of the airplane. Hydraulic power overcomes these limitations

A hydraulic FCS has 2 parts:

• The mechanical circuit
• The hydraulic circuit

The mechanical circuit links the cockpit controls with the hydraulic circuits. Like the mechanical FCS, it is made of rods, cables, pulleys, and sometimes chains. The hydraulic circuit has hydraulic pumps, pipes, valves and actuators. The hydraulic pressure generated by the pumps in the hydraulic circuit powers the actuators. The actuators convert hydraulic pressure into control surface movements. The servo valves control the movement of the actuators. The above two-control system has a major drawback that it contains heavy mechanical circuitry, which increases the weight of the system. To overcome this drawback a new technology “MORE ELECTRIC TECHNOLOGY IN AIRCRAFT” was developed. The aircraft in which this technology was used was called “MORE ELECTRIC AIRCRAFT”(MEA).

 

When describing the MEA, flight control actuation systems can be considered to involve two main technological areas: fly-by-wire (FBW) and power-by-wire (PBW). FBW technology comprises the design, development and implementation of electronics for flight control systems. Electronic control provides flight control and actuator control functionality implemented using either centralized or distributed architectures. Distributed control systems reduce the processing load on centralized flight control computers, and offer more flexibility during system architecture development. A further benefit is the reduction in weight achieved by reducing harness size and component quantity. In recent years, technological advancement has centered on the FBW field, to the extent that FBW control systems are now the standard in today’s commercial and military aircraft. Power-by-wire (PBW) actuation is the next major breakthrough in aircraft control. Just as the fly-by-wire flight control system eliminated the need for mechanical interfaces, power-by-wire actuators eliminate the need for central hydraulic systems. Control power comes directly from the aircraft electrical system. This has several advantages. Central hydraulic systems are complicated and difficult to maintain. Removing these systems would greatly reduce the amount of support equipment and personnel required to maintain and operate current air and space vehicles. In addition, PBW actuators have the potential to be more efficient than their hydraulic counterparts. A central hydraulic system must generate and sustain significant hydraulic pressure (3,000 to 6,000 pounds per square inch) at all times, regardless of demand. PBW actuators only use electrical power when needed. Finally, PBW actuation systems can be made far more fault tolerant than those depending on a central hydraulic supply. Once a hydraulic line is compromised, it usually leads to the loss of that entire hydraulic circuit. As a result, multiple hydraulic circuits are required to maintain some level of redundancy. With a PBW system, a failed actuator can simply be switched off, isolating the problem to a single surface.

 

Types of PBW Actuators

 

There are several different types of PBW actuators, including electrohydrostatic actuators (EHA) and electromechanical actuators (EMA). EHAs use a reversible, electrically driven pumpmotor to directly pump self-contained hydraulic fluid to a piston. This drives the ram in the same fashion as a standard hydraulic actuator (Figure 1(a)). An EMA has no internal hydraulic fluid, instead using electric motors to directly drive the ram through a mechanical gearbox (Figure 1(b)). Compared to an EHA, the EMA has certain advantages. It is lighter, smaller, and less complex than an equivalent EHA because of the absence of an internal hydraulic system. Since there is no hydraulic fluid in the load path, the EMA tends to be stiffer than an equivalent EHA. The EMA tends to be more efficient because there are no windage losses or pump inefficiencies. Finally, since there is no leak potential with an EMA, it is better suited to long term storage or space applications.

 

 

 

 

Electromechanical Actuation (EMA)

 

An EMA uses mechanical gearing to couple an electric motor to a flight control surface. This is achieved using a rotary gearbox, and depending on the actuation method required, can include some form of rotary-to-linear conversion, such as a ball screw. Electric motors requiring a DC electrical supply are typically used, although the addition of a diode rectification stage will also allow them to operate from an AC electrical supply. Motor speed, direction, and torque translate directly to speed, direction, and load in the actuator. Figure 1 shows an EMA currently being developed by TRW for a high-power flight control application. In its basic form, the EMA is susceptible to certain single-point failures that can lead to a mechanical jam, and consequently presents complications for flight certification on certain surfaces. Additional devices can be used to mitigate against this failure mode, but in doing so, complexity, cost, and weight are increased. For these reasons, the basic EMA is not suited for primary flight control applications. However, spoiler systems and secondary actuation systems could accommodate EMA technology.

 

 

 

 

 

 

EMA system layout

 

 

 

Large EMA for High-Power flight controls

 

 

 

 

 

Baseline Power System Architecture

 

The proposed power distribution system is built around a 270V DC distribution bus. The typical baseline power system architecture for a next generation aircraft is shown in Fig. 1. It can be seen that the key components that control the power are the bidirectional power converters (BDCs). A bus regulator provides an interface between the starter/generator and the distribution bus. Most of the loads, including the actuators, are regulated using bidirectional power converters, which control and condition the power from the DC bus.

 

 

 

 

 

 

 

 

 

 

 

 

 

 

With the proliferation of bidirectional power converters and advanced actuators in the power distribution system, it is important to develop methods to analyze the interaction between the different subsystems. Due to the complexity of the baseline power system and the large number of subsystems, a sample power distribution system, which captures the essential features of the baseline system but is not as complicated, is introduced. The sample power system is represented as a interconnection of a source and load subsystem.

 

Sample Power Distribution System

The sample power distribution system is shown in Fig. 2. The source subsystem represented by subsystem 1 consists of an ideal three phase voltage source, a three-phase boost rectifier to provide the regulated 270V DC required by the DC bus. The load subsystem represented by Subsystem 2 is an electromechanical actuator used to control the secondary flight control surfaces on the aircraft. The other loads on the DC bus are modeled by a current source, or a simple resistance.

 

 

 

The EMA model shown in Fig. 5 is shown to include a DC motor with constant field, a ball screw transmission between the motor and the control surface, and a model of the surface dynamics. The motor voltage is controlled by a PWM bidirectional buck converter with an input filter. The EMA is controlled by a multi-loop controller, which includes a motor current, motor speed, and the ball screw position feedback loops.All of the other loads on the bus are modeled by a resistor or a current source.

 

 

 

 

 

Electro hydrostatic Actuation (EHA)

 

1. In contrast to EMA, EHA (Figure 2) uses fluidic gearing between the electric motor and the surface actuator. Hydraulic fluid provides an intermediate means of transmitting power to the surface. Here, a variable-speed electric motor (typically DC) is used to drive a fixed-displacement hydraulic pump, which in turn, powers a conventional hydraulic piston jack. Change in direction is achieved by the use of a bi-directional motor. A major advantage to this approach is that the EHA operating mode can be managed like a conventional hydraulic actuator. This approach is achieved using standard hydraulic bypass or damping valves (Figure 3); thus traditional active-standby, or active-active, actuator configurations can be readily adopted. This capability makes the EHA more suitable for primary flight control applications than the EMA. Although EHA technology reintroduces hydraulic components and fluid, it is totally self-contained          within   the        actuator assembly. Compared to traditional hydraulic actuator systems, the inconvenience   of         hydraulic disconnection from aircraft supplies and the complications of bleeding the system during reinstallation are not encountered during     maintenance.

 

 

 

 

 

 

 

Electrohydrostatic Actuators (EHA)

 

 

 

 

 

Large EHA

 

EHA Control Schematic

Benefits of electrically powered Actuators:

 

The potential benefits of electric actuation at a system level have been well publicized.

Electric actuation can offer:

• Improved aircraft maintainability:
• Fewer hydraulic components are required,
• Faster aircraft turnaround,
• Fewer spares and tools are needed,
• Improved fault-diagnosis through        built-in test (BIT).
  • Improved system availability and reliability:
  • Electrical distribution is more practical and offers system flexibility with respect to reconfiguration Ñ a capability previously difficult to achieve using hydraulics,
  • Improved mean-time-between-failures (MTBFs) through removal (electromechanical actuation or EMA) or on-demand usage   (electrohydrostatic actuation or EHA) of hydraulic components.
  • Improved flight safety Ñ in the MEA configuration, improved system safety is achieved through dissimilar actuator power supplies and subsequent avoidance of common mode failures.
  • Reduced system weight Ñ weight saving, achieved through the replacement of entire hydraulic systems, including pumps, distribution networks (pipes and fluid), and valve blocks, by electric systems.

The main benefit is the reduction of aircraft operating costs, for example, reduced fuel cost (as a result of reduced weight), and lower maintenance costs (quicker turnaround). However, before such benefits can be realized, additional work is required to improve the technology and provide the appropriate application platforms to introduce the technology into service.

Furthermore, the aircraft maintenance industry must realign its infrastructure so that it can reap the benefits of electric technologies.

 

Some additional benefits of both EMA and EHA actuators are:

• Low quiescent power consumption during standby operation,
• Rapid start-up response,
• Can be easily adapted for use with AC or DC electric supplies,
• Insensitive to supply frequency variation of AC electric supplies.

 

EHA versus EMA?

 

An alternative to EHAs, are ‘electromechanical actuators’ (EMAs), in which the motor torque is mechanically amplified and transmitted to the control surface using a gear set, screw or other mechanical transmission device, can be seen as an alternative. Indeed, as far as complexity, weight, reliability and maintenance requirement are concerned, EMAs are potentially more attractive than EHAs, at least for low power applications. In particular, all hydraulic technology relevant problems are obviously eliminated from the EHA configuration. However, in the three following areas EHAs are still preferable to EMAs:

?The jamming probability of an EMA used in a primary flight control application is difficult to predict and substantiate from existing in-service experience. Jamming probability of an EHA, can be directly assessed from the current servo control experience, and shown as ‘extremely improbable’ if properly bypassed. In contrast, the jamming probability of mechanical systems incorporating hundreds of gear teeth and screw mechanisms is questionable and present-day experience in secondary flight control applications may not be directly transferable to primary flight controls, due to very different duty cycles in particular

Wear of the mechanical transmissions components may result in control surface ‘free-play’ or other non-linearities, which may generate unacceptable limit cycles

?The introduction of an EHA in parallel with regular servo control in the basic more-electric architecture described above is easier than an EMA. EHAs can easily be made reversible in standby mode, they can incorporate identical damping devices to those currently used for flutter protection, and they can be built with many components common with the adjacent servo control such as the piston, cylinder, associated position transducer or the accumulator. In an obvious move to spread the technical as well as financial risk, Airbus has called on the talents of several companies for the design, production and supply of the many actuators on this mammoth aircraft. Specifically, the A380 aileron and elevator EHAs, as well as rudder EBHAs are purchased from Goodrich, while Messier-Bugatti will supply the associated EHA pumps. Meanwhile, the spoiler EBHAs are from Liebherr, which supplies its own pumps. Phil Hudson, Goodrich VP engineering for actuation systems notes: “The electronic EHA concept can also be designed to serve more functions than simply motor control. It can serve as a smart actuator controller in its own right and be part of a distributed control system or to control a set of multiple actuators. Another benefit is that this distributed technology puts intelligence local to the actuation elements in a control system and can substantially reduce harness weight and improve fault detection and isolation.”

Maintenance benefits are also substantial. Power-by-wire EHA actuation units are line-removable with only mechanical and electrical connections to the aircraft, which eliminates the need to refill or bleed systems of hydraulic fluids as is required with central hydraulics. Since power-by-wire actuators are self contained and remotely located at the surfaces, the area exposed to damage is greatly reduced. Additionally, power-by-wire actuators can be designed as position sensitive, which means that the actuators provide only the flow and pressure necessary to move and hold the actuator in a desired

position. Conventional central-hydraulic systems are configured to produce continuous pressure. Flow is metered at each actuator, which can lead to a large consumption of power and generate unwanted heat. William Schley, R&D supervisor, Parker Aerospace, Controls Systems Division explains that EHAs only consume power on demand.  Specifically, they consume power in proportion to the power delivered to the load. In contrast, a conventional EVSV-equipped hydraulic servoactuator consumes power in proportion to output speed, allocating power to output load as needed, with the remainder of the power being dissipated through pressure drop (heat) across the main control valve. Whilst hydraulic actuators become more efficient the more they are loaded, loads are typically low during most of a flight.” Another important advantage of electric actuators is survivability. Ballistic or explosive damage to an electric power distribution system or actuator usually does not cause loss of function of that entire channel, particularly if the damage is peripheral. In a hydraulic system, depending on its design, even a small leak can cause a major loss of function and/or fire. Although some electric actuators contain

hydraulic fluid, the system as a whole is still usually more survivable. For now, these more advanced failure management functions are being provided by the EHA and its variants. EHA combines the best of electric actuation and conventional hydraulics for a hybrid design approach, which is more fault tolerant than most current EMAs. Moreover, EHAs are mechanically simple, and immune to gear train jams. The typical long-term storage capability for EHA is 10 years plus.

 

 

 

Next-Generation---All-Electric Aircraft:

 

The “All-Electric” aircraft is a concept that emerged in the 1970s and has engendered a large amount of research activity. An all-electric engine, which could replace current aero gas turbines, would drive all accessories electrically, via a distribution network, from motor/generators embedded in the engine spools. Extending the function of the motor/generators to include service as active magnetic bearings would facilitate deletion of the oil system. The all-electric concept thus offers a huge scope for both engine and airframe reconfiguration and operational improvements, with studies indicating benefits of overall weight reduction, increased reliability, easier maintainability, reduced operating costs (including reduced fuel burn), and enhanced safety.

 

 

 

 

 

Conclusion:

 

Beginning with the scenario of a single hydraulic power supply replaced by an electric one, it is possible to establish the relativity and scale for the changes required in the migration toward the “All-Electric” aircraft concept. On a small civil airliner, typically a minimum of five electric actuators would be needed to provide one lane of electrical control for the primary flight control surfaces. If all hydraulic systems were converted to electric, in excess of 20 electric actuators would be needed to provide complete control of all primary and secondary flight control surfaces. The consequential increase in electrical power demand has major implications for electrical power generation and distribution systems. Thus, a significant amount of work is still needed to address the consequences of distributing many electrical actuators around an aircraft, and the consequential start-up, steady state, and peak demands required of aircraft electrical power supplies.

It is clear that the migration to electric actuation systems is affecting both civil and military markets. As described previously, the replacement of a single hydraulic system by an electric substitute is a major step in the transition to all-electric technologies. It is quite evident that the demands being made on aircraft generators and distribution architectures will increase considerably to meet the needs of this migration. A company named TRW has already developed products to meet the current demands envisioned by PBW and has programs to ensure that it will meet any future demands required by the all-electric aircraft. Finally, it is envisioned that once in service, electric actuator technology and electrical system architectures will improve the commercial viability and in-service reliability of the airframes to which they are fitted. These improvements will undoubtedly drive the adoption of greater levels of electric actuation on future aircraft.

 

References:

 

• Weimer J. A, “Power management and distribution for the More Electric Aircraft”, Proceedings of the30thIntersocietyEnergyConversion Engineering Conference, vol. 1, July 1995, pp. 273-277

 

 

 

• Technology Review Journal — Millennium Issue • Fall/Winter 2000

 

• ACTUATOR DEVELOPMENT OVERVIEW

   D. Tesar, UT Austin, Robotics Research Group April 1, 2006

 

 

 

About the Author

military aircraft mishaps