Patent Application
20220269291
The present invention generally relates to a flight control system for an aircraft, and an aircraft, that may have such a flight control system. In particular, the invention relates to an aircraft to the canard type having multiple lift/thrust units distributed along the front or canard wings and along the aft or main wings.
Aircrafts maybe generally classified into fixed wing and rotating wing types. Fixed wing aircrafts typically include a plurality of flight control surfaces that, when controllably positioned, guide the movement of the aircraft from one destination to another destination. The number and type of flight control surfaces included in an aircraft may vary. Primary flight control surfaces are typically those that are used for control the aircraft movement with respect to the pitch, yaw and roll axes. Secondary flight control surface are typically those that are used to influence the lift or drag (or both) of the aircraft. Typical primary flight control surfaces include elevators, ailerons and rudder, and typical secondary flight control surfaces include a plurality of flaps, slats speed brakes and spoilers.
Rotary-wing aircraft, such as e.g. helicopters, typically do not have flight control surfaces that are separated from the airfoils that produce lift, but the airfoils that constitute the rotating wing have a cyclic control for pitch and roll, and a collective control for lift.
Further, aircrafts are known, which have a vertical take-off and landing capability based on propulsion engines, which are mounted rotatably with respect to a transvers axis or pitch axis of the aircraft. The propulsion engines are controllably movable between a cruising flight position and a take-off/landing position. In the cruising position the engines provide forward thrust and the movement of the aircraft through the air is controlled by means of suitable flight control surfaces. In the take-off/landing position, the propulsion engines are angled downwards for allowing a vertical take-off or landing based on the thrust provided by the engines.
Such types of aircraft with vertical take-off and landing capability and having electrically driven ducted propellers as propulsion engines have been proposed by the present applicant Lilium GmbH according to the publications US 2016/0023754 A1 and US 2016/0311522 A1 and further publications of the same patent families. The present applicant meanwhile has developed an aircraft denoted as Lilium jet, which is a canard type aircraft and has plural left front engines, plural right front engines, plural left aft engines and plural right aft engines in the form of ducted propellers operated electrically, which are mounted to respective flaps of the front left and right canard wing and the left and right aft or main wing of the canard type aircraft. The first test flight of this Lilium jet was performed on Oct. 1, 2019.
For any such kind of aircrafts and any other type of aircraft, resilience against technical failures is one of the most important aspects. The flight control system is a critical system, for which redundancy must be provided, also in view of regulatory requirements, which have to be met.
According to a conventional approach, typically a well-known (traditional) triplex architecture having three independent flight control computers and three independent networks is used.
However, traditional triplex architectures generally have a substantial weight, because of a higher number of busses, connectors and the necessary physical path segregation.
State of the art technology solutions for bus communication generally have the following disadvantages: rather slow communication (e.g. ARINC-429), need for a critical bus controller or router (e.g. MIL-STD-1553, AFDX), high costs (e.g. MIL-STD-1554, AFDX), proprietary single supplier (monopoly and therewith resulting high costs and unwanted dependency from a single source) or use of complex technology (e.g. TTE, TTP, . . . ), are rather heavy because of many wires (e.g. RS-422 full duplex) and have limitations in the network configuration, e.g. not allowing N-to-N communication.
Based on conventional approaches, the use of other communication busses and in particular the use of CAN bus architecture for fly-by-wire flight control systems is limited to unmanned operations.
A known conventional flight control system for an aircraft, in general terms, comprises a flight computer system, which is connected via an electronic or optoelectronic bus system with a plurality of bus nodes, which each are configured to at least one of controlling an associated aircraft device based on command messages received from the flight control computer system via the bus system and sending information messages to the flight control computer system via the bus system.
In view of the foregoing, it is an object of the present invention, according to a first aspect thereof, to provide such a flight control system for an aircraft, on basis of which a substantial resiliency against technical failures can be achieved in an efficient manner.
According to a second aspect, the invention shall provide an aircraft having such a flight control system, which is configured to achieve such a resiliency.
According to a third aspect, it is an object of the present invention to provide an aircraft of the canard type, which achieves technical advantages over aircrafts of the prior art and may be provided to advantage with a flight control system according to the invention.
For achieving the firstly mentioned object, the invention, according to the first aspect, provides a flight control system for an aircraft, comprising: a flight control computer system, which is connected via an electronic or optoelectronic bus system with a plurality of bus nodes, which each are configured to at least one of controlling an associated aircraft device based on command messages received from the flight control computer system via the bus system and sending information messages to the flight control computer system via the bus system.
According to the invention (first aspect), the electronic or optoelectronic bus system is a redundant electronic or optoelectronic bus system comprising plural independent bus sub-systems, wherein each bus node is configured to communicate with the flight control computer system via two different bus sub-systems of the plural independent bus sub-systems, wherein each bus node further is configured to communicate with the flight control computer system on basis of an associated predetermined bus communication protocol via a first bus sub-system of the respective two different bus sub-systems and on basis of an associated predetermined bus communication protocol via a second bus sub-system of the respective two different bus sub-systems.
The provision of two different bus sub-systems is the basis for achieving integrity, availability and dissimilarity, as is appropriate for a critical system like the flight control system. Therewith tolerance to failures in the bus communication can be achieved with minimal impact in the flight mission.
According to a favorable approach, different bus communication protocols may be used. In this respect, it is proposed that each bus node of a first group of the bus nodes is configured to communicate with the flight control computer system via the first bus sub-system on basis of a first bus communication protocol, which is the associated predetermined bus communication protocol of the bus nodes of the first group for the first bus sub-system, and is configured to communicate with the flight control computer system via the second bus sub-system on basis of a second bus communication protocol, which is the associated predetermined bus communication protocol of the bus nodes of the first group for the second bus sub-system and differs from the first bus communication protocol.
Further, it is proposed that each bus node of a second group of the bus nodes is configured to communicate with the flight control computer system via the first bus sub-system on basis of the second bus communication protocol, which is the associated predetermined bus communication protocol of the bus nodes of the second group for the first bus sub-system, and is configured to communicate with the flight control computer system via the second bus sub-system on basis of the first bus communication protocol, which is the associated predetermined bus communication protocol of the bus nodes of the second group for the second bus sub-system.
Therewith favorable dissimilarity is achieved also with respect to the bus communication protocols, resulting in a particular good resiliency against failures.
According to a further favorable approach, each of the plural independent bus sub-systems may be composed of plural independent communication busses. In this respect, it is proposed further that each bus node is connected via one associated communication bus of the first bus sub-system with the flight control computer system and is configured to communicate via this communication bus of the first bus sub-system with the flight control computer system on basis of the associated predetermined bus communication protocol and that each bus node is connected via one associated communication bus of the second bus sub-system with the flight control computer system and is configured to communicate via this communication bus of the second bus sub-system with the flight control computer system on basis of the associated predetermined bus communication protocol.
The provision of plural independent communication busses within each bus sub-system allows achieving a high degree of redundancy and therewith good resiliency against failures, in particular if the various aircraft devices and their bus nodes are associated to the independent communication busses in an intelligent manner. In this context, it is assumed and suggested that plural bus nodes are associated to each of a plurality or most or all of the plural independent communication busses of the first bus sub-system, wherein the bus nodes being associated to the same independent communication bus of the first bus sub-system are configured to communicate via this common independent communication bus with the flight control computer system, and that plural bus nodes are associated to each of a plurality or most or all of the plural independent communication busses of the second bus sub-system, wherein the bus nodes being associated to the same independent communication bus of the second bus sub-system are configured to communicate via this common independent communication bus with the flight control computer system.
Particularly preferred embodiments of the flight control system of the invention are based on a CAN bus system. To this end, it is proposed that the bus nodes are CAN bus nodes, the plurality independent bus sub-systems are realized as independent CAN bus sub-systems, and the predetermined bus communication protocols each are one of plural different CAN bus protocols according to a respective CAN standard, including a first CAN bus protocol according to a first CAN standard and a second CAN bus protocol according to a second CAN standard differing from the first CAN standard.
According to the secondly mentioned favorable approach, each of the independent CAN bus sub-systems may be composed of plural independent CAN busses realizing the plural independent communication busses, so that each CAN bus node is connected via one associated CAN bus of a first CAN bus sub-system with the flight control computer system and each bus node is connected via one associated CAN bus of a second bus CAN sub-system with the flight control computer system.
In this respect, it is further proposed that each of plural or most or all of the CAN busses of the first CAN bus sub-system is connected with plural associated of the CAN bus nodes and that each of plural or most or all of the CAN busses of the second CAN bus sub-system is connected with plural associated of the CAN bus nodes, wherein the CAN bus nodes being connected with the same CAN bus are configured to communicate via this common CAN bus with the flight control computer system.
According to the firstly mentioned favorable approach to use plural bus communication protocols, each CAN bus node of a/the first group of the bus nodes is configured to communicate with the flight control computer system via the associated CAN bus of the first CAN bus sub-system on basis of the first CAN bus protocol, and is configured to communicate with the flight control computer system via the associated CAN bus of the second CAN bus sub-system on basis of the second CAN bus protocol, and each CAN bus node of a/the second group of the bus nodes is configured to communicate with the flight control computer system via the associated CAN bus of the first CAN bus sub-system on basis of the second CAN bus protocol, and is configured to communicate with the flight control computer system via the second CAN bus sub-system on basis of the first CAN bus protocol.
To advantage, one of the first and second CAN bus protocols may follow one of the ISO 11888 standard referred to as CAN STANDARD and the SAE J2284-5:2016 standard referred to as CAN FD.
Further, it is proposed that the other of the first and second CAN bus protocols follows the other of the ISO 11888 standard referred to as CAN STANDARD and the SAE J2284-5:2016 standard referred to as CAN FD.
Both protocols are conventionally known and have favorable characteristics. Typically a high speed CAN according to ISO 11898-2 (CAN STANDARD) allows up to 1 Mbit/s bus communications. High speed CAN according to SAE J2284-5:2016 (CAN FD) typically allows up to 5 Mbit/s. Therewith bus communication according to the CAN FD standard may to advantage be used also for other bus communications than the usual transmission of command messages and information messages for controlling the aircraft. For example, firmware or software updates of various bus nodes and other components connected to the bus system could be done in a very efficient manner by using the CAN FD standard.
Preferably, the flight control computer system is a redundant flight control computer system that comprises plural independent flight control computers. Preferably, three independent flight control computers are provided. The plural independent flight control computers may be dissimilar flight control computers differing in at least one of the flight control computer hardware and the flight control computer software, preferably differing in both thereof.
Each flight control computer may be connected via one of the independent bus sub-systems or independent CAN bus sub-systems with each of the bus nodes or CAN bus nodes. Preferably, at least one of the flight control computers is connected via the first bus sub-system or first CAN bus sub-system with each of the bus nodes or CAN bus nodes and at least one other of the flight control computers is connected via the second bus sub-system or second CAN bus sub-system with each of the bus nodes or CAN bus nodes. This substantially contributes to the desired resiliency against failures.
In this respect, it is further proposed that at least one of the flight control computers is configured to communicate with the bus nodes or CAN bus nodes of a/the first group on basis of the first bus communication protocol or first CAN bus protocol and with the bus nodes or CAN bus nodes of a/the second group on basis of the second bus communication protocol or second first CAN bus protocol, and that at least one other of the flight control computers is configured to communicate with the bus nodes or CAN bus nodes of the first group on basis of the second bus communication protocol or second CAN bus protocol and with the bus nodes or CAN bus nodes of the second group on basis of the first bus communication protocol or first CAN bus protocol.
A particularly preferred embodiment has three flight control computers, as already mentioned. In this respect, it is proposed further that a first flight control computer and a second flight control computer are connected via the first bus sub-system or first CAN bus sub-system with each of the bus nodes or CAN bus nodes, and that a third flight control computer is connected via the second bus sub-system or second CAN bus sub-system with each of the bus nodes or CAN bus nodes.
With reference to the mentioned bus communication protocols and CAN bus protocols, it is proposed further that the first flight control computer and the second flight control computer are configured to communicate with the bus nodes or CAN bus nodes of the first group on basis of the first bus communication protocol or first CAN bus protocol and are configured to communicate with the bus nodes or CAN bus nodes of the second group on basis of the second bus communication protocol or second CAN bus protocol, and that the third flight control computer is configured to communicate with the bus nodes or CAN bus nodes of the first group on basis of the second bus communication protocol or second CAN bus protocol and is configured to communicate with the bus nodes or CAN bus nodes of the second group on basis of the first bus communication protocol or first CAN bus protocol.
If three flight control computers are provided, a kind of majority voting or negotiating between the flight control computers may be realized. In this respect, it is further proposed that the flight control computers are configured to elect one of the flight control computers to be the flight control computer in control and therewith to elect the other two flight control computers to be a supervising flight control computer. Each flight control computer may be configured to operate as flight control computer in control and to control the aircraft based on command messages sent to bus nodes or CAN bus nodes via the respective independent bus sub-system or independent CAN bus sub-system and possibly based on information messages received from bus nodes or CAN bus nodes via the respective independent bus sub-system or independent CAN bus sub-system. At least two, preferably all three flight control computers may be configured to operate as a supervising flight control computer and to monitor at least one of the operation of the flight control computer currently being the flight control computer in control and messages sent via the respective independent bus sub-system or independent CAN bus sub-system.
To advantage, the flight control computers or at least the flight control computers configured to be a supervising flight control computer may be configured to elect a different flight control computer than the flight control computer currently operating as flight control computer in control as new flight control computer in control, based on the monitoring done by the flight control computers operating as supervising flight control computers.
According to the second aspect of the invention, an aircraft comprising a flight control system according to the invention as described in the foregoing is provided. The aircraft may be at least one of a single pilot aircraft, an aircraft having a vertical take-off and landing capability and an aircraft of the canard type, for example. A particular preferred embodiment is an aircraft of the canard type having a vertical take-off and landing capability. However, the flight control system of the invention can favorably be implemented in any other kind of aircrafts, including all types of aircrafts mentioned in the foregoing description of the technical background.
The aircraft may have plural aircraft devices of a common type, which each have associated a respective bus node or CAN bus node of the flight control system. In this respect, it is proposed that the aircraft devices are arranged in a number and configuration at one or both of the fuselage of the aircraft and wings of the aircraft to achieve a resiliency against failures, such that various subgroups of the plural aircraft devices each comprising at least two of the aircraft devices of the common type may fail without endangering the flight capability and the controllability of the aircraft.
Concerning the mentioned preferred realization of the flight control system having independent communication busses or CAN busses, it is proposed that the bus nodes or CAN bus nodes of the aircraft devices of the common type are associated in such a number and manner to a respective independent communication bus or CAN bus of the first bus sub-system or first CAN bus sub-system and are associated in such a number and manner to a respective independent communication bus or CAN bus of the second bus sub-system or second CAN bus sub-system, that any combination of two independent communication buses or CAN busses of the flight control system may fail without substantially compromising the flight capability and the controllability of the aircraft.
The aircraft devices of the common type or of a first common type may be flaps, which have air control surfaces. The flaps may be mounted in a moveable, e.g. pivotable manner to wings of the aircraft, and each flap may have associated at least one flap actuator and a bus node or CAN bus node, which is configured to control a deflection angle of the flap by controlling the at least one flap actuator based on command messages received from the flight control computer system.
The aircraft devices of the common type or of a second common type may be propulsion engines. Each propulsion engine has associated a bus node or CAN bus node, which is configured to control the operation of the propulsion engine based on command messages received from the flight control computer system.
According to a particular preferred embodiment, flaps as well as propulsion engines are provided, which are associated to each other. In this respect, it is proposed that the propulsion engines are mounted to or integrated with an associated one of the flaps, so that a thrust direction of the propulsion engines can be controlled by controlling the deflection angle of the respective flap by means of the respective at least one flap actuator and the respective bus node or CAN bus node associated thereto.
According to one approach, plural or all flaps which serve to control a thrust direction each have associated only one of the propulsion engines, which is mounted to or integrated with the respective flap.
According to another approach, plural or all flaps which serve to control a thrust direction each have associated plural of the propulsion engines, which are mounted to or integrated with the respective flap. In this respect, it is further proposed that a propulsion module comprising plural propulsion engines is mounted to or integrated with the respective flap.
According to a favorable realization, flap actuators and propulsion engines associated thereto have associated common bus nodes or CAN bus nodes. In this respect, it proposed that for each flap having associated one propulsion engine or plural propulsion engines the at least one flap actuator of the flap and the associated propulsion engine or plural propulsion engines have associated a common bus node or CAN bus node, which is configured to control the propulsion engine or propulsion engines and the at least one flap actuator based on command messages received from the flight control computer system.
According to the third aspect of the invention, which is an independent aspect (independent of the first and second aspects), an aircraft of the canard type is provided, which comprises: a fuselage, two main wings extending transversely from the fuselage, two canard wings extending transversely from the fuselage and being located forward of the main wings, flap mounted to the wings, a flight control system, and associated aircraft devices comprising flap actuators and propulsion engines, which can be controlled by of the flight control system. Each of the main wings and the canard wings is provided with plural flaps having air control surfaces, wherein each flap has associated at least one flap actuator, which serves to adjust a deflection angle of the flap as commanded by the flight control system. Each of the main wings and the canard wings is provided with plural propulsion engines, which can be operated with variable thrust as commanded by the flight control system and which each are mounted to or integrated with an associated one of the flaps, so that a thrust direction of the propulsion engines can be controlled by controlling the deflection angle of the respective flap my means of the respective at least one flap actuator. The flight control system comprises a flight control computer system, an electronic or optoelectronic transmission system and plural control nodes which are connected via the electronic or optoelectronic transmission system with the flight control computer system and which each are associated to at least one of the aircraft devices, wherein the control nodes are configured to control the associated at least one aircraft device based on commands received via the electronic or optoelectronic transmission system from the flight control computer system.
According to one favorable approach, plural or all flaps which serve to control a thrust direction each may have associated only one of the propulsion engines, which is mounted to or integrated with the respective flap.
According to another favorable approach, plural or all flaps which serve to control a thrust direction each have associated plural of the propulsion engines, which are mounted to or integrated with the respective flap. For achieving this mounting or integration in a cost effective manner, a propulsion module comprising plural propulsion engines may be mounted to or integrated with the respective flap.
According to a preferred embodiment, common control nodes serving to control one or plural propulsion engines and one or plural flap actuators may be provided. In this respect, it is proposed that for each flap having associated one propulsion engine or plural propulsion engines the at least one flap actuator of the flap and the associated propulsion engine or plural propulsion engines have associated a common control node, which is configured to control the propulsion engine or propulsion engines and the at least one flap actuator based on commands received from the flight control computer system via the electronic or optoelectronic transmission system.
The flight control system of the aircraft may be characterized by the features of the flight control system according to the first aspect of the invention. To this end, the control nodes may be realized as bus nodes or CAN bus nodes and the electronic or optoelectronic transmission system may be realized as redundant electronic or optoelectronic bus system or CAN bus system as proposed according to the first aspect of the invention.
The aircraft according to the third aspect may be further characterized by features of the aircraft as proposed according to the second aspect of the invention.
Under the first and second aspects of the invention, a flight control system architecture with triple dissimilar flight control computers and an aircraft having a flight control system according to this architecture are provided.
Each flight control computer, FCC, may be implemented by a different CPU and may communicate with each other FCC via a dedicated bus. During operation, one FCC is elected to effectively control the aircraft (referred as “FCC In Control”), based on its I/O capability, while the others monitor the operation and the messages propagated by the “FCC In Control” via the dedicated bus. If the monitoring FCCs agree that the “FCC In Control” is misbehaving, one of them will be elected the “new FCC In Control” and this decision will be propagated to the other systems of the flight control system.
The FCCs may be connected to the actuators and other aircraft devices of the aircraft via a Dual Fly-By-Wire distribution of the CAN type referred to as CAN A and CAN B. FCC1 and FCC2 may be connected to CAN A and FCC3 may be connected to CAN B. Each CAN network A/B may be composed by plural independent full-duplex CAN buses which connect the FCCs to all flap actuators of plural flaps and all propulsion engines of the aircraft. The flap actuators preferably are realized as electro-mechanical actuators, and the propulsion engines preferably are realized as electrically driven propulsion engines. The propulsion engines favorably may be integrated into lift/thrust units comprising one electrically driven propulsion engine or plural electrically driven propulsion engines, a flap and the respective at least one electromechanical flap actuator. The plural independent full-duplex CAN buses provide direct control to all lift/thrust units for each individual FCC.
For example, a canard type aircraft of the invention may have thirty six (36) lift/thrust units, preferably six (6) lift/thrust units per canard wing and twelve (12) per main wing. Each lift/thrust units comprises one flap and one propulsion engine mounted to or integrated with the flap. The lift/thrust units may be connected via six (6) independent full-duplex CAN buses of CAN network A and via six (6) independent full-duplex CAN buses of CAN network B with the FCCs. Each network provides enough throughput for high update rate commands (up to 200 Hz) and is still compliant with ARINC-825 throughput limitation of 50% bus utilization.
Each lift/thrust unit may have two interfaces, one connected to CAN A and the other connected to CAN B. Each interface of the two interfaces may implement a different type of CAN protocol: e.g. one following the ISO 11898 referred as CAN STANDARD and the other the SAE J2284-5:2016 referred as CAN FD, providing a level of dissimilarity on the lift/thrust unit side.
Favorably, the connection between each lift/thrust unit and the individual CAN bus in each network may be arranged in such a manner, that full fault tolerance to single failures is obtained and impact of dual bus failures is minimized. In a particularly preferred arrangement, dual CAN failure will cause, in the worst case, loss of two lift/thrust units: one in the canard and one in the opposite side wing; or two lift/thrust unit in the same wing close to the fuselage. In both cases, there would only be a rather low impact on the lateral balancing of the aircraft, which can easily be handled by the pilot, if needed. According to a preferred embodiment, the unbalancing caused by the failure is handled by the flight control system itself, redistributing part of the command that was supposed to be performed by the failed units to the remaining healthy units. In this case, no pilot action is needed to compensate the unbalancing.
Some advantages of the flight control system architecture of the invention are the following:
According to the invention, advantages of a triplex architecture on the flight control computer side, providing a system tolerant to any single FCC failure, may favorably be combined with the advantages of a dual CAN network, providing a lighter and simpler solution than a conventional triple network architecture, but keeping the required availability and dissimilarity requirements.
Of particular technical value are the following advantages and achievements enabled by the invention according to its first and second aspects:
In
The flight control system 10 further comprises a pilot user interface, which may include a left sidestick apparatus 30a and a right sidestick apparatus 30b, the left sidestick apparatus having a left sidestick 32a and the right sidestick apparatus having a right sidestick 32b. Both sidesticks can be pivoted in a left-right direction, about a first maneuvering axis extending at least roughly in a longitudinal direction of the aircraft and in a forward-backward direction, about a second maneuvering axis extending at least roughly in a traverse direction of the aircraft, preferably orthogonal to the first maneuvering axis.
Corresponding multiple degree of freedom assemblies and sensor assemblies 38a, 38b sensitive to the pivoting movements of the sidesticks or/and pivoting forces acting via the sidesticks as conventionally known may be provided for the two sidesticks.
Electronic flight control signals or electronic flight control commands generated by the sensor assembly 38a and sensor assembly 38b are transmitted via electronic or optical connection links 42a and 42b to the flight control computer system 12.
The shown embodiment of
These flaps are mounted pivotably or moveably to the respective wing and can be pivoted about a pivoting axis or moved with a pivoting movement component by a respective electric actuator arrangement, preferably independently of each other for each flap. Each flap can be pivoted between an upper first operational position and a lower second operational position. Each flap may assume a position of minimum or vanishing inclination with respect to a longitudinal axis of the aircraft, possibly the upper first operational position, and a position of maximum downward inclination with respect to the longitudinal axis of the aircraft, possibly the lower second operational position. However, if the position of maximum downward inclination corresponds to a vertical orientation of the flap, the lower second operational position may alternatively be a position beyond the position of maximum downward inclination, so that the flap points slightly forward.
To each of these flaps at least one propulsion engine in the form of a ducted propeller, which is operated electrically, is mounted. The ducted propellers preferably are mounted to an upper surface of the respective flap. Alternatively, the propulsion engines may be integrated into a respective flap in a manner, that an air channel of the respective propulsion engine, in which the respective ducted propeller rotates, is located above and aligned with an upper surface of the respective front wing or aft wing.
Preferably, the flaps may assume a position corresponding to the lower second operational position or another operational position between the first and the second operational positions, in which the ducted propellers provide only vertical thrust downwardly, which provides the aircraft with a vertical take-off and landing (VTOL) capability. In the upper first operational position or another operational position between the first and the second operational positions, in which the flaps extend in the longitudinal direction or at a minimum angle with respect to the longitudinal direction of the aircraft, the operating ducted propellers provide maximum forward thrust for the aircraft. The flaps operate not only for controlling the thrust direction of the propulsion engines or propulsion modules, but also as flight control surfaces influencing the movement of the aircraft in the air based on the usual aerodynamic principles.
In the shown embodiment of
In the shown embodiment of
The intermediate deflection angles of the flap as illustrated in
According to preferred embodiments, in case of the approach of
Analogously, in case of the approach of
In the following, such combinations of one flap 234 with at least one propulsion engine 232 or 232a, 232b and 232c, the flap actuator or flap actuator arrangement 240 and the control node 250, as illustrated in
The bus nodes 250 are bus nodes such as the bus nodes 14, 16, 18 and 20 shown in
As mentioned, the canard-type aircraft 200 may be provided with a flight control system shown schematically in
According to the invention, the bus system or CAN bus system 22 comprises a first independent bus sub-system 22a or independent CAN bus sub-system 22a and a second independent bus sub-system 22b or second independent CAN bus sub-system 22b, as illustrated in
The bus nodes 14, 16, 18 and 20, possibly each representing one of plural bus nodes 250 belonging to a respective flap and propulsion engine combination as illustrated in
As illustrated in
For example, only a minor effect on the lateral balancing of the aircraft may result, if two lift/thrust units, e.g. of the aircraft shown in
Even more resiliency against failures in the bus communications is achieved, if each of the bus nodes 14, 16, 18 and 20, in particular bus nodes like the bus nodes 250, are configured to communicate via one of the independent bus sub-systems 22a and 22b according to a first bus communication protocol and are configured to communicate via the other independent bus sub-system of the bus sub-systems 22a and 22b according to a second bus communication protocol, which differs from the first bus communication protocol. In case of CAN busses, it is preferred that each of the bus nodes is configured to communicate with the flight control computer system 12 on basis of a first CAN bus protocol via one of the two CAN bus sub-systems 22a and 22b and on basis of a second CAN bus protocol different from the first CAN bus protocol via the other of the two CAN bus sub-systems 22a and 22b. Correspondingly, the flight control computer system, more particular its flight control computers, is/are configured to communicate with the bus nodes via the respective CAN bus sub-system based on the two different CAN bus protocols.
Preferably, one of the first and second CAN bus protocols follows one of the ISO 11888 standard referred to as CAN STANDARD and the SAE J2284-5:2016 standard referred to as CAN FD, and the other of the first and second CAN bus protocols follows the other of the ISO 11888 standard referred to as CAN STANDARD and the SAE J2284-5:2016 standard referred to as CAN FD.
Providing a CAN bus system 22 with independent CAN bus sub-systems 22a and 22b and respectively plural independent CAN busses 24a1, 24a2, 24a3 and 24b1, 24b2, 24b3, respectively, is indeed highly preferred in the context of the present invention. The CAN bus system is a mature system providing many advantages. The bus participants or bus node decide based on an object identifier of a respective bus message, whether the bus message is relevant. The bus participants are not individually addressed, in agreement with the so-called “ATM” or “anyone-to-many” principle. The bus access is automatically controlled based on the object identifier by bitwise arbitration. However, the ARINC-825 specification even provides additionally for “peer-to-peer” or “PTP” communications, which could be used for certain functions, if desired. A typical CAN bus requires only two signal lines typically provided as twisted pair, namely CAN-Low and CAN-High. In practice, there is normally also a GND (ground) and a CAN V+ (power) line. If desired, CAN bus data may be sent via optical single mode or multimode fibers, using suitable CAN to fiber optic convertors and the like.
In the following, a more detailed embodiment of the flight control system of the canard type aircraft of
The six flaps or lift/thrust units 214 of the canard wing 206 have assigned the identification numbers 1.1 to 1.6. The six flaps or lift/thrust units 214 of the canard wing 208 have assigned the identification numbers 2.1 to 2.6. The twelve flaps or lift/thrust units 210 of the main wing 202 have assigned the identification numbers 3.1 to 3.12. The twelve flaps or lift/thrust units 212 of the main wing 204 have assigned the identification numbers 4.1 to 4.12. These identification numbers are also included in
The identification numbers 1.1, 2.1, 3.1 and 4.1 identify the respectively most inner flap or lift/thrust unit adjacent or near the fuselage 203, and identification numbers 1.6, 2.6, 3.12 and 4.12 identify the outmost flap or lift/thrust unit having a maximum distance from the fuselage 203, and the other flaps or lift/thrust units and their positions along the respective wing or canard are correspondingly identified by the four identification number inserts in
The bus communication via the CAN busses identified by the identification numbers 1, 2 and 3 follows the ISO 11888 standard, to which it was referred as CAN STANDARD, and is referred to in
The independent CAN busses identified by the identification numbers 1, 2 and 3 of the first CAN bus sub-system 24a and the independent CAN busses identified by the identification numbers 4, 5 and 6 of the second CAN bus sub-system 24b are arranged in an upper portion of the drawing of
For the first flight control computer 12a or FCC1, the second flight control computer 12b or FCC2 and the third flight control computer 12c or FCC3, the bus nodes or lift/thrust units 1.1, 1.2, 4.3, 4.4, 4.7, 4.8, 1.3, 1.4, 4.2, 4.5, 4.9, 4.10, 1.5, 1.6, 4.1, 4.6, 4.11, 4.12 (upper portion of the drawing) belong to a first group of bus nodes and the bus nodes or lift/thrust units 2.1, 2.2, 3.3, 3.4, 3.7, 3.8, 2.3, 2.4, 3.2, 3.5, 3.9, 3.10, 2.5, 2.6, 3.1, 3.6, 3.11 und 3.12 (lower portion of the drawing) belong to a second group of bus nodes. The flight control computers 12a and 12b communicate with the bus nodes of the first group according to the ISO 11888 standard referred to as CAN STANDARD or CAN 2.0B and with the bus nodes of the second group according to SAE J2284-5:2016 standard referred to as CAN FD. The third flight control computer 12c communicates with the bus nodes of the first group according to the SAE J2284-5:2016 standard referred to as CAN FD and with the bus nodes of the second group according to the ISO 11888 standard referred to as CAN STANDARD or CAN 2.0B.
As already mentioned, resiliency against failures is achieved. A twofold bus communication failure in CAN bus 6 of the first independent CAN bus sub-system 24a denoted as CAN A and in CAN bus 1 of the second independent CAN bus sub-system 24b denoted as CAN B will result in a failure of the lift/thrust units 3.1 and 3.6 of the left main wing 202. Analogously, a twofold failure in the CAN busses 3 and 4 of the CAN A and CAN B bus sub-systems will result in a failure of the lift/thrust units 4.1 and 4.6 in the right main wing 204. Therewith a lift/thrust unit adjacent to the fuselage and a lift/thrust unit still rather close to the fuselage would be affected, so that only a minor impact of the lateral balancing would occur.
Further, a bus communication error in the independent CAN busses 3 and 6 of the CAN A and CAN B bus sub-systems would affect the outermost lift/thrust unit 1.6 of the left canard wing 206 and the outermost lift/thrust unit 4.12 of the right main wing 204, and a bus failure in the CAN busses 6 and 3 of the CAN A and CAN B bus sub-systems would affect the outermost lift/thrust unit 2.6 of the right canard wing 208 and the outermost lift/thrust unit 3.12 of the left main wing 202. Again, the lateral balancing would not be affected very much. The lift/thrust units are assigned to the individual independent CAN busses of both CAN bus sub-systems in a manner that any combination of two bus failures in any pair of these busses would have no major impact on the lateral balancing, so that the flight capability and the controllability of the aircraft is not substantially compromised.
Of course, other assignments of the various bus nodes of the lift/thrust units to the individual CAN busses of the two independent CAN bus sub-systems than shown in
The illustrated principle of achieving resiliency against failures based on the present invention can of course also be applied to other kind of aircraft than the aircraft shown in
The aircraft of the type and variants illustrated in
Herewith also the following items are disclosed:
| Number | Date | Country | Kind |
|---|---|---|---|
| 21158264.8 | Feb 2021 | EP | regional |
