Flight control system and method for VTOL aircraft

Abstract

The present invention relates to a flight control system for a VTOL aircraft (10) that defines an aircraft reference coordinate system having a roll axis (Xb), a pitch axis (Yb), and a yaw axis (Zb), wherein the flight control system comprises first and second manual control devices for inputting control commands by an operator, and a flight control computer connected to the first and second manual control devices and configured to output flight control commands based on the pivot positions of first and second stick members with respect to first to fourth control axes, wherein the first manual control device comprises a first stick member, the first stick member being pivotable around first and second control axes with respect to a first neutral position, and the second manual control device The system comprises a second stick member, the second stick member being pivotable around third and fourth control axes with respect to a second neutral position, and the flight control computer is configured to derive and output flight control commands for longitudinal motion control based on the pivot position of the first stick member with respect to the first control axis, for lateral motion control based on the pivot position of the first stick member with respect to the second control axis, for vertical motion control based on the pivot position of the second stick member with respect to the third control axis, and for directional motion control based on the pivot position of the second stick member with respect to the fourth control axis, while at least partially eliminating the interconnection between the individual directions of motion. The present invention further relates to a VTOL aircraft (10) having such a flight control system.

Description

Detailed Description of the Invention

[0001] The present invention relates to a flight control system for a VTOL aircraft that defines a body reference coordinate system having a roll axis, a pitch axis, and a yaw axis, a VTOL aircraft equipped with such a flight control system, and a method of controlling the flight of a VTOL aircraft by the flight control system of the present invention.

[0002] Several generations of VTOL aircraft, as well as other aircraft having VTOL capabilities such as rotary-wing aircraft (helicopters) and unmanned drones, have been known for quite a long time, but development has been carried out in recent years aiming to provide a fully electric VTOL aircraft having the ability to transport passengers and / or cargo, which is controlled by a pilot on board. In this context, when developing a flight control concept for such types of aircraft, achieving intuitive pilot control that is easy and safe to operate in all flight modes and flight conditions, including vertical takeoff and landing, hovering flight, and horizontal cruising or forward flight, is one of the main challenges.

[0003] Already known VTOL aircraft have used different control strategies for vertical and forward flight for this purpose, which has increased the mental burden on the pilot. This is because the pilot had to adapt how to fly the aircraft and operate the controls depending on the airspeed regime. For example, in the well-known military aircraft Hawker Siddeley Harrier series, during hovering, the pilot controlled the aircraft's vertical movement with a dedicated throttle control member and the aircraft's horizontal movement with a stick member. In contrast, during subsequent forward flight, the stick member and throttle member were used in the same way as in conventional aircraft. That is, the rate of climb and airspeed were adjusted by a combination of stick input and throttle input, by manipulating the aircraft's pitch attitude by pushing and pulling the stick and controlling the engine thrust output, while roll or roll motion was controlled by the stick in the conventional way. Furthermore, a separate lever member had to be provided to control the engine nacelle angle, and therefore the direction of engine thrust, and this separate lever member had to be manually adjusted by the pilot when transitioning from vertical flight to forward flight and vice versa. In addition, a pair of pedals had to be provided in the above-mentioned aircraft to control the aircraft's yaw, and this pair of pedals was operated in the same manner as in conventional aircraft designs.

[0004] In certain examples of tiltrotor aircraft, a similar implementation is used, but the dedicated throttle lever member is replaced by a "collective-style lever" that has vertical displacement instead of longitudinal displacement but performs a similar function. In such examples of tiltrotor aircraft, a separate lever or switch is provided to control the nacelle angle, and therefore the orientation of the engine and rotor relative to the aircraft's fuselage and wings, and this separate lever or switch had to be adjusted by the pilot when transitioning from vertical flight to forward flight. Furthermore, a pair of pedals is provided to control the aircraft's yaw, and the pilot had to operate this pair of pedals with their feet in a manner similar to that described above for the Harrier aircraft.

[0005] In certain examples of rotary-wing aircraft, such as some types of helicopters, during vertical flight, a "cyclic stick component" controlled the aircraft's horizontal movement while a "collective stick lever," automatically coupled to the engine control device, commanded the aircraft's vertical axis movement. During the forward flight phase, in such examples, the rate of climb and airspeed were controlled by a combination of the cyclic stick and collective stick components by adjusting the aircraft's pitch attitude and rotor lift, while lateral control was performed solely by the cyclic stick component.

[0006] In this context, referring to recent developments in this field, there is a broad consensus in the aircraft industry that changes to pilot input strategies in accordance with the current airspeed regime of aircraft should be avoided in new generation VTOL aircraft, and especially in so-called Urban Air Mobility VTOL aircraft, in order to reduce pilot workload, increase safety, and reduce the amount of pilot training required. In this context, the so-called "integrated control strategy" has been proposed and adopted for certain military aircraft, such as Lockheed Martin's F-35B. However, the strategies used in conventional integrated control designs have several drawbacks; for example, in order to move horizontally during vertical flight, the pilot must combine inputs to two different control elements, which is more complex than using a single controller with only one hand.

[0007] Figure 1 provides a non-comprehensive overview of the control strategies described above for VTOL aircraft known in the art, where the control elements in each of the three proposed designs for conventional airplanes, conventional rotary-wing aircraft, the Hawker Siddeley Harrier Jet, a typical tiltrotor aircraft, and next-generation VTOL aircraft are shown in a schematic and at least partially simplified manner. All of the designs described above and shown in Figure 1 require numerous control elements to be handled by the pilot of the corresponding aircraft, such as a pair of foot pedals, nozzle or nacelle tilt angle levers on a conventional stick member, and / or exhibit suboptimal behavior in specific stages or conditions of typical VTOL flight. All of the exemplified control elements also require adaptation of different input strategies in different airspeed regimes of the corresponding aircraft.

[0008] Therefore, an object of the present invention is to provide a flight control system and a flight control method that are intuitive and easy to operate by a single pilot without requiring a substantial change in piloting technique when transitioning from vertical flight or hovering to horizontal flight or forward flight, and in particular without requiring the operation of multiple control elements to carry out the transition from vertical flight to forward flight.

[0009] For this purpose, the flight control system of the present invention comprises first and second manual control devices for an operator to input control commands, the first manual control device comprising a first stick member having a grip portion for the operator to grasp, the first stick member mounted on a first base member so as to be pivotable around first and second control axes with respect to a first neutral position, the second manual control device comprising a second stick member having a grip portion for the operator to grasp, the second stick member mounted on a second base member so as to be pivotable around third and fourth control axes with respect to a second neutral position, and the flight control system is further connected to the first and second manual control devices. The aircraft is equipped with a flight control computer configured to output flight control commands based on the pivot positions of the first and second stick members relative to the first to fourth control axes, the flight control computer being configured to derive and output flight control commands for longitudinal motion control based on the pivot position of the first stick member relative to the first control axis, for lateral motion control based on the pivot position of the first stick member relative to the second control axis, for vertical motion control based on the pivot position of the second stick member relative to the third control axis, and for directional motion control based on the pivot position of the second stick member relative to the fourth control axis, while at least partially eliminating the interconnection between the individual directions of motion. Therefore, no additional levers are required to achieve the transition from vertical flight to forward flight, nor pedals to control the yaw behavior of the aircraft.

[0010] Therefore, the flight control computer automatically controls aircraft effectors such as engines and control surfaces to achieve both the desired aircraft response that follows pilot input (particularly by eliminating the interconnection between individual motion axes or directions of motion) and the adaptation of engine regimes and control surface positions according to flight phases such as vertical flight and forward flight.

[0011] In particular, according to the present invention, each control element always commands the same direction of motion for the aircraft, regardless of its current flight phase or flight conditions. Furthermore, by evaluating the pilot's input in a suitable manner, the flight control computer further simplifies the piloting task by ensuring that the aircraft's response to the pilot's input matches a specific response type, regardless of the aerodynamic regime, that is, regardless of whether lift in the aircraft is generated by vertically directing engine thrust or by aerodynamic surfaces such as wings and / or flaps that are given shape during level flight.

[0012] In particular, by using the flight control system described here, additional control elements such as rudder pedals are not required, and the cockpit for this purpose is equipped only with two stick members, sometimes referred to as "inceptors" or "side stick controllers." Each such inceptor has two degrees of freedom and may be active (reverse-driven by an electric motor) or passive (i.e., spring-loaded). Thus, the motion of the aircraft is achieved by displacing the inceptors from their initial or neutral position, and the aircraft's response to input control commands is at least partially directionally decoupled based on the control strategy that is specifically implemented.

[0013] Therefore, the flight control computer is configured to receive corresponding signals from interceptors and convert them into commands for aircraft control effectors, such as aerodynamic control surfaces, engines, and engine positioning devices relative to the aircraft's fuselage, in order to achieve the desired aircraft response. These commands output by the flight control computer may, in some cases, depend on the aircraft's current airspeed regime, or, more generally, on the aircraft's current flight conditions. Specifically, the commands may depend on whether the aircraft is currently hovering, transitioning, or in level (forward) flight, because, for example, different measures and strategies are required to change the aircraft's altitude during different flight phases. However, by outputting appropriate commands to the aircraft control effectors for all conceivable flight stages and conditions, the flight control computer can ensure that each control element always commands the same direction of motion of the aircraft, regardless of the aircraft's current flight phase or conditions, and thus simplify the piloting task and reduce the pilot's mental burden when transitioning from vertical to forward flight and vice versa, as well as in all other conceivable flight conditions.

[0014] In this regard, it should be noted that in the flight control system of the present invention, for example, a certain degree of redundancy may be provided in the flight control computer, which may be equipped with at least two independent units.

[0015] It should be further noted that, using the control strategy implemented by the flight control system of the present invention, the aircraft's pitch is never directly controlled by the pilot, as the longitudinal and vertical motion controls of the aircraft are separated and assigned to different control axes of the first and second stick members. This is a fundamentally different approach from all past fixed-wing aircraft designs, enabling novel and improved control strategies in all flight configurations, in which the flight control computer determines which combination of angle-of-attack lift and direct lift (e.g., thrust) is required to achieve the desired aircraft response. For example, in vertical flight, to accelerate the aircraft upward, the flight control computer may increase engine thrust while maintaining a pitch angle of zero. In forward flight, to achieve the same vertical acceleration, the computer may increase the pitch attitude to correct the angle of attack, and thus generate further aerodynamic lift while simultaneously increasing direct lift, for example, by increasing thrust.

[0016] Furthermore, in hovering flight, any horizontal movement is commanded by a single stick member (the first stick member), and therefore, it is not necessary to combine inputs to two separate control elements to bring about, for example, diagonal movement of the aircraft, i.e., movement along both the longitudinal and lateral directions simultaneously.

[0017] In this context, it should be noted that while the roll, pitch, and yaw axes of an aircraft may substantially correspond to the horizontal and vertical directions relative to the ground during hovering flight, the aircraft reference coordinate system may differ substantially from the corresponding ground reference coordinate system during forward flight due to the pitching and rolling of the aircraft to control its altitude and direction or course. For this reason, further details of the present invention will be described below using two separate coordinate systems: an aircraft coordinate system, also known as the aircraft coordinate axes, having roll, pitch, and yaw axes, also called the Xb, Yb, and Zb axes, and so-called "following axes." The following axes Xf and Yf ​​are obtained by projecting the aircraft axes Xb and Yb onto a horizontal plane. The Zf axis is always perpendicular to the ground surface and points downward. In this context, it should be further noted that lateral motion control based on the second control axis of the first stick member can always result in motion around the Xb aircraft axis, and directional motion control based on the fourth control axis can always result in motion around the Zf tracking axis. It should also be further noted that partial or complete exclusion of interconnections between individual directions of motion can be limited to specific flight modes for at least some of the multiple directions. In this case, in some embodiments, in particular, the separation of vertical motion control from at least one of the longitudinal motion control, lateral motion control, and directional motion control may always be maintained.

[0018] To further improve the intuitive use of the flight control system of the present invention, the first and second stick members may be inclined relative to each other in the first and second neutral positions, respectively. Preferably, the first neutral position corresponds to an orientation of the first stick member substantially aligned with the yaw axis of the aircraft (upright orientation), and the second neutral position corresponds to an orientation of the second stick member that forms a predetermined angle with respect to both the roll axis and yaw axis of the aircraft, preferably in the range of 30° to 60°, and more preferably substantially 45°. Thus, the movement of each stick member or interceptor as input by the aircraft pilot corresponds to the resulting motion of the aircraft.

[0019] By tilting the first and second stick members relative to each other in their neutral positions, the intuitiveness of using the flight control system can be enhanced. This is because any movement of the second interceptor relative to the third control axis can correspond to the vertical, i.e., up-and-down motion of the aircraft. Specifically, by positioning the second stick member at a 45° angle pointing forward and upward, pushing the second stick member downward may instruct the aircraft to move vertically downward, while pulling the second stick member upward may result in the aircraft moving upward.

[0020] In this context, it should also be noted that the first and third control axes of the first and second stick members, respectively, may substantially correspond to the aircraft's pitch axis or Yb axis, thereby resulting in the movement of the stick along the aircraft's longitudinal direction, performed by the pilot to instruct the corresponding longitudinal and vertical motion control commands. However, it may also be conceivable to tilt the corresponding control axes with respect to the aircraft's reference coordinate system, for example, to improve the usefulness of the flight control system of the present invention with respect to the pilot's anatomical features. It should also be noted that, to facilitate access to the pilot's seat, one of the stick members, in particular the second stick member tilted forward, may be embodied as a so-called "swoop-type inceptor," which may be housed in the armrest portion of the corresponding base member when the pilot is not yet seated, and may only be moved to its operating position when the pilot is seated in the corresponding pilot's seat.

[0021] As briefly described above, the flight control computer of the flight control system of the present invention may be configured to derive flight commands for different directional motion controls based on a specific principle, which may or may not vary between different flight modes and flight conditions of the aircraft, in particular with respect to the current airspeed range of the aircraft detected by a dedicated sensor unit that provides corresponding data to the flight control computer.

[0022] For example, the flight control computer may be configured to derive and output flight commands for vertical motion control as altitude change rate command commands, preferably across the entire airspeed range of the aircraft. In this specification, altitude change rate is with respect to a ground reference coordinate system, and the corresponding operation of the second stick member with respect to the third control axis can be translated into the aircraft's descent rate or climb rate in some or all of the flight modes and airspeed ranges that the aircraft can take. Alternatively, in the aircraft's high airspeed range, flight commands for vertical motion control may be output as load factor change rate commands or FPA (flight path angle) change rate commands.

[0023] The flight control computer may be configured to set the rate of altitude change to zero when the second stick member is in the neutral position with respect to the third control axis. In such an example, the neutral position corresponds to the state of the manual control device where the rate of altitude change remains constant at zero. Alternatively, the neutral position may refer to an operating condition where the corresponding rate of change remains constant at a non-zero value.

[0024] Additionally or alternatively, the flight control computer may be configured to derive and output flight commands for longitudinal motion control as translational velocity command instructions (TRCs) over at least a portion of the aircraft's airspeed range, particularly over at least the low airspeed range. In this context as well, the aircraft's translational velocity is with respect to a ground reference coordinate system, and in certain embodiments, the corresponding operation of the first stick member with respect to a first control axis leads to a change in the aircraft's translational velocity. In particular, the TRCs may be controlled by first and second control axes of the first stick member, so that the stick displacement with respect to the first axis is proportional to the X component of the ground velocity vector at the follow axis, and the stick displacement with respect to the second axis is proportional to the Y component of the ground velocity vector at the follow axis.

[0025] Additionally or alternatively, the flight control computer may be configured to derive and output flight commands for longitudinal motion control as linear (kinematic) acceleration in the x-following axis. Furthermore, the flight control computer may be configured to derive and output flight commands ("speed hold") to maintain a constant airspeed even in the presence of atmospheric disturbances when the stick is in the neutral position, over at least a portion of the aircraft's airspeed range, particularly over the high airspeed range.

[0026] It is also conceivable that the flight control computer may be configured to derive and output flight commands for longitudinal motion control as airspeed rate change command instructions over at least a portion of the aircraft's airspeed range, particularly over the high airspeed range.

[0027] Additionally or alternatively, the flight control computer may be configured to derive and output flight commands for lateral motion control, i.e., around the Xb aircraft axis, as bank angle command instructions, which can be used for all aircraft speeds. Alternatively, a roll rate command around the Xb aircraft axis can be used, but this is not very practical for vertical flight close to the ground and may only be used at high speeds.

[0028] The flight computer may also be configured to derive and output flight commands for directional motion control as heading change rate command commands over at least a portion of the aircraft's airspeed range, particularly over the low airspeed range. For example, a heading change rate command around the Z-following axis may be used at low speeds, while at high speeds, there may be three possible options for the command: delta heading change rate (i.e., the difference between the actual heading change rate and the kinematic heading change rate related to the bank angle in a coordinated turn), ny (lateral acceleration along the aircraft axis), and sideslip angle. All three of these options ensure coordinated turns. That is, if the pilot commands a bank angle or roll rate on the lateral axis while keeping the directional interceptor neutral, the resulting bank turn will be coordinated in that it results in zero sideslip and / or zero ny. It is clear that by appropriately combining different control strategies for individual motion controls, advanced flight maneuvers such as declabing can be easily performed with only two stick members without requiring additional input elements. Furthermore, the flight control computer may be configured to derive and output a flight command ("heading hold") to maintain a constant heading during hovering flight, even in the presence of atmospheric disturbances, when the stick is in the neutral position.

[0029] It should also be noted that in this context, the flight control computer may further be configured to evaluate a transition speed value between a low airspeed range and a high airspeed range based on the ground speed or calibrated airspeed of the aircraft. By determining such a transition speed value, different behaviors of the aircraft regarding a set of control commands may be defined, and different aircraft control effectors suitable for achieving a desired aircraft response may be identified for different airspeed ranges. In this specification, the term "transition speed value" may also refer to a transition speed range over which a stepwise or continuous transition is performed for control commands output by the flight control computer. In particular, a so-called "blending" of different control commands may be performed over the transition speed range, for example, as a linear function of speed, in which case the low-speed control strategy is continuously blended into the high-speed control strategy.

[0030] As a further safety measure, the flight control computer may further be configured to perform at least one of flight envelope protection for vertical motion control by setting upper and lower limits for angle of attack, rate of climb, rate of descent, load factor, pitch angle, and / or flight path angle, flight envelope protection for longitudinal motion control by setting upper and lower limits for maximum forward airspeed and maximum reverse airspeed or ground speed (when measurement or estimation of reverse airspeed is not reliable), and flight envelope protection for lateral and directional motion control by setting upper and lower limits for bank angle, lateral acceleration, and sideslip angle. Such means for flight envelope protection of an aircraft function to prevent the aircraft from reaching an unstable or unsafe state, such as a stall of the aircraft during forward flight. For this purpose, and to achieve the desired effect, the input range of the stick member may be dynamically limited in the case of an active inceptor, and / or other means may be taken to ensure compliance with the limits of the safe flight envelope at all times.

[0031] As a further safety measure, if a failure of a control effector or loss of sensor data is detected, the flight control computer may automatically reconfigure to adapt to a specific failure condition. In some cases, this may cause some of the features described above to become unavailable. For example, if the angle of attack data is lost, angle of attack protection may no longer be available, and if the ground speed data is lost, the translational speed command may also be lost.

[0032] Furthermore, the flight control system of the present invention may include at least one sensor unit. The at least one sensor unit is operably connected to the flight control computer and is configured to output data representing at least one motion parameter of the aircraft. The flight control computer is further configured to modify flight control commands based on the received sensor output data. This data can include, for example, the speed, acceleration, and other motion parameters of the aircraft, and thus enables the formation of a control feedback loop for optimizing the control and behavior of the aircraft while taking into account internal and external factors such as component performance and weather conditions.

[0033] According to a second aspect, the present invention relates to a VTOL aircraft having a body reference coordinate system with a roll axis or Xb, a pitch axis or Yb, and a yaw axis or Zb. The VTOL aircraft includes a fuselage, a pair of main wings, a pair of canard wings, a plurality of propulsion units (particularly, electric ducted fan engines) distributed on the main wings and canard wings, and the flight control system of the present invention described above. The propulsion units are arranged to be pivotable about at least one axis along a predefined angular position range, and both the angular position and RPM of the propulsion units are individually controllable by the flight control system.

[0034] In this specification, one or more propulsion units may be mechanically connected to their respective main wings or canard wings by flaps, which are pivotable relative to their respective wings and also function as interfaces between the corresponding wing pairs, the propulsion units, and the aerodynamic surfaces. Meanwhile, the propulsion units are individually controllable with respect to their RPM. Thus, the aircraft of the present invention has the ability to perform a wide variety of flight maneuvers by using thrust vectoring by controlling both the flap angle and the engine RPM. For this purpose, the aircraft of the present invention may be equipped with a fly-by-wire system, in which the aforementioned flight control computer functions as a central processing unit that receives sensor inputs and control inputs and outputs suitable control commands to aircraft control effectors, including engine and flap angle adjustment motors.

[0035] In particular, the aircraft of the present invention may be configured to transition between a hovering flight mode in which the required lift is mainly generated by the vertical thrust of the propulsion unit and a forward flight mode in which the required lift is mainly generated aerodynamically by the main wings and canard wings. In this specification, the hovering flight mode may also be referred to as the vertical flight mode, and the forward flight mode may also be referred to as the horizontal flight mode or cruising flight mode.

[0036] In this context, the flight control system for a VTOL aircraft of the present invention may further be configured to keep the aircraft level at a substantially constant pitch angle in hovering flight mode and to correct the pitch angle of the aircraft in forward flight mode.

[0037] According to a third aspect, the present invention relates to a method for controlling the flight of a VTOL aircraft of the present invention using the flight control system of the present invention, the method comprising: evaluating the current pivot positions of first and second stick members with respect to first to fourth control axes; deriving a flight control strategy with respect to longitudinal motion control, lateral motion control, vertical motion control, and directional motion control based on the pivot positions of the first and second stick members; and controlling at least the propulsion unit of the aircraft, in particular the angular position and RPM of the propulsion unit, preferably controlling only the propulsion unit in such a manner, according to the flight control strategy.

[0038] Therefore, in particular, in embodiments of the aircraft of the present invention in which the flaps supporting the propulsion unit also function as aerodynamic surfaces, additional control surfaces may be redundant, and all necessary flight control may be performed by controlling only the propulsion unit with respect to its orientation and thrust output in the manner just described. The features of the flight control system described above may also be interpreted as referring to methods that may be performed by the corresponding components of the flight control system, and it should be noted that in the context of this application, the features of dependent device and method claims are considered to be interchangeable.

[0039] Further features and advantages of the present invention will become even clearer from the following description of embodiments of the invention, in conjunction with the accompanying drawings. [Brief explanation of the drawing]

[0040] [Figure 1] This is a schematic diagram of different types of control elements for aircraft known in the art. [Figure 2] This is a schematic isometric view of the electric propulsion VTOL aircraft of the present invention. [Figure 3a] Figure 2 shows schematic diagrams illustrating different motion types of the aircraft under hovering and low-speed flight conditions. [Figure 3b] Figure 2 shows schematic diagrams illustrating different types of aircraft motion under cruising flight conditions. [Figure 4] Figure 2 is a schematic side view of the pilot's seat and its respective control elements in the aircraft. [Figure 5] Figure 4 is a schematic diagram of the control elements. [Figure 6] Figure 2 is a schematic diagram of the aircraft's flight control system. [Figure 7] Figure 6 is a schematic diagram of flight control commands in the flight control system. [Figure 8] Figure 7 is a schematic diagram of the aircraft's response to longitudinal and transverse flight control commands. [Figure 9] Figure 7 is a schematic diagram of the aircraft's response to vertical flight control commands and directional flight control commands. [Modes for carrying out the invention]

[0041] In Figure 2, the electric propulsion VTOL aircraft of the present invention is shown in a schematic isometric view, with the entire structure indicated by reference numeral 10. The aircraft 10 comprises a cockpit for a single pilot, and a fuselage 12 that houses a passenger cabin and various components and systems necessary for the operation of the aircraft 10, such as a high-capacity rechargeable battery and flight control and avionics systems.

[0042] The aircraft 10 further comprises a pair of main wings 14 and a pair of canard wings 16 positioned in front of the main wings 14 with respect to the longitudinal direction of the aircraft 10. This longitudinal direction corresponds to the aircraft's roll axis or Xb axis and, together with the pitch axis or Yb axis and the yaw axis or Zb axis, forms the aircraft's reference coordinate system with respect to the center of mass CM. Note that the main wings 14 and the canard wings 16 extend substantially parallel to the Yb axis, and that the main wings 14 have a longer wingspan than the canard wings 16 and are provided with winglets 14a at their wingtips, while the winglets 16a of the canard wings 16 are provided on the outermost flaps 20, which are described below, and are therefore pivotable relative to each canard wing 16 itself.

[0043] In substantially forward flight along the Xb axis of the aircraft 10, the main wings 14 and canard wings 16 are shaped such that they can provide aerodynamic lift, and thus the thrust of the aircraft 10 is oriented primarily along the Xb axis, enabling energy-efficient horizontal or cruising flight. To provide this thrust, both the main wings 14 and canard wings 16 are each equipped with a plurality of electrically powered ducted fan engines 18 at their respective trailing edges, which function as the propulsion units of the aircraft 10. The engines 18 are mounted on flaps 20 so as to be pivotable relative to the corresponding main wings 14 or canard wings 16, around a pivot axis extending substantially parallel to the Yb axis of the aircraft 10. In this case, each individual engine 18 may be provided on an individual flap 20, or multiple engines 18 may be provided on a single flap 20, for example, two or three groups of engines 18 may be provided on each flap 20.

[0044] Furthermore, by pivoting the flaps 20 using flap actuators 20a, which are provided as interfaces between each flap 20 and the corresponding main wing 14 or canard wing 16, so that the engine thrust is directed downward along the Zb axis of the aircraft 10, hovering, low-speed flight, and vertical takeoff and landing become possible. To transition between hovering flight and forward flight, the aircraft must go through a transition flight phase in which the aircraft moves between a state in which lift is mainly generated by downward engine thrust and a state in which lift is mainly generated by the aerodynamic action of the main wing 14 and canard wing 16. It should also be noted that the flaps 20 function not only as means for interface connecting the engine 18 to the main wing 14 and canard wing 16, but also as aerodynamic control surfaces that contribute to the controllability of the aircraft 10 in addition to the thrust of the engine 18.

[0045] Referring here to Figures 3a and 3b, different types of aircraft motion based on the differential thrust of engine 18 are schematically shown for hovering and low-speed conditions, and cruising or high-speed conditions, respectively. In hovering flight, it can be seen that the pitch, roll, and yaw of aircraft 10 can be controlled by pivoting the flaps of the main wings 14 and canard wings 16, respectively, and controlling engine 18 to provide different absolute thrust values ​​(for example, by commanding different RPMs). Furthermore, it can be seen that forward motion of aircraft 10 along the Xb axis under low-speed conditions can be achieved by suitably directing the thrust of engine 18 to the main wings 14 and canard wings, rather than requiring a change in pitch angle. Similarly, in order to change the altitude of the aircraft, i.e., to move the aircraft up and down, the thrust of engine 18 must be directed downward along the Zb axis and adjusted to an appropriate absolute value, while aircraft 10 can remain nearly horizontal in the horizontal plane in the ground reference coordinate system.

[0046] In contrast, and as can be understood from Figure 3b, during level or forward flight, i.e., in the high airspeed range, the flaps 20 are substantially aligned with the profiles of the main wing 14 and the canard wing 16, and the lift is provided by the aerodynamic forces of the main wing 14, the canard wing 16, and to some extent the fuselage. Under such conditions, the pitch and roll of the aircraft 10 are controlled by adjusting the flap position, as in the case of a conventional canard wing / main wing aircraft. However, the yaw of the aircraft may still be controlled through differential thrust by providing a higher absolute thrust value by either the right-hand or left-hand engine 18.

[0047] Referring here to Figure 4, a schematic side view of the pilot's seat 22 of the aircraft 10 and its respective control elements is provided. The pilot P is seated in the pilot's seat 22 which is oriented forward along the Xb axis of the aircraft 10, and the pilot's seat 22 is equipped with two armrests 22a in an ergonomically suitable manner. Both armrests 22a are equipped with stick members 24 and 26, each having a grip portion for the pilot to grasp. The stick members 24 and 26, together with their respective first and second base members 22b, form first and second manual control devices 23a, 23b, schematically shown in Figure 5, which are for inputting control commands to enable the aircraft 10 to be controlled to perform flight maneuvers such as those shown in Figures 3a and 3b.

[0048] In this case, the right-hand stick member will be referred to as the first stick member 24, and the left-hand stick member will be referred to as the second stick member 26. Figure 4 shows that the first stick member 24, in its neutral position, is oriented approximately upward or along the Zb axis of the aircraft 10, while the second stick member 26 is inclined relative to the first stick member 24, for example, at an angle of approximately 45° forward. It should be further noted that the second stick member 26 may be embodied as a so-called "swoop-type inceptor," which may be housed in the corresponding armrest 22a when the pilot is not yet in the pilot's seat 20, or it may simply move to the operating position shown in Figure 4 once the pilot is seated.

[0049] Referring to Figure 5, the control elements 24 and 26 of the aircraft 10 described above, as well as possible control inputs, are schematically shown. Both stick members 24 and 26 are mounted on their respective base members 22b, which may be integrated with, for example, an armrest 22a so as to be able to swing around their respective two control axes, and each base member 22b may further include electronic components for evaluating the current pivot position of the stick members 24 and 26 and outputting corresponding data to the flight control computer 30, which will be described below.

[0050] In the case of the first stick member 24, forward and backward movement corresponds to oscillation around the first control axis 24a, and left and right movement corresponds to oscillation around the second control axis 24b. On the other hand, in the case of the second stick member 26, forward and backward movement corresponds to oscillation around the third control axis 26a, and left and right movement corresponds to oscillation around the fourth control axis 26b. When the second stick member 26, which is in the neutral position, is tilted forward and upward, it can be understood that the oscillation of the second stick member 26 around the third control axis 26a by the pilot may also be perceived as pushing down and pulling up the second stick member 26 in forward and backward movement of the second stick member 26.

[0051] Subsequently, and with reference to Figures 7-9, possible flight control commands that can be performed by the stick members 24 and 26, and the corresponding aircraft responses will be described in detail. Referring first to Figure 6, a schematic diagram of the flight control system 28 of the aircraft 10 is shown. The flight control system 28 comprises first and second stick members 24 and 26 as manual control devices, a flight control computer 30, and a number of sensor units 32, the flight control computer 30 being configured to output flight control commands to aircraft control effectors, in particular to the aircraft's engines 18, and to electric motors that function as flap actuators 20a for pivoting the flaps 20 relative to the main wings 14 and canard wings 16.

[0052] During the flight operation of the aircraft 10, the pivot positions of the first and second stick members 24 and 26 with respect to the first to fourth control axes 24a, 24b, 26a, and 26b are transmitted to the flight control computer 30 and evaluated by the flight control computer 30. Specifically, the evaluation takes into account the current airspeed, represented by the output data from the sensor unit 32, as well as the attitude of the aircraft 10, the angular velocity of the aircraft 10, the angular acceleration, and the linear acceleration, so that the flight control computer can output flight control commands to the engine 18 and the flap actuator 20a. In this case, different levels of redundancy can be foreseen for both hardware and software.

[0053] For example, in response to airspeed data provided by a sensor unit 32, which may include an INS / GNSS sensor, the flight control computer 30 will instruct the aircraft control effector to perform flight maneuvers as shown in Figures 3a and 3b, for example. In this case, a feedback loop may be provided in the flight control computer 30 based on the sensor output, in which case the flight control computer 30 will evaluate the output data from the sensor unit 32 to determine whether the motion of the aircraft 10 has changed as expected and instructed, or whether additional flight control commands are needed to achieve a desired state of motion for the aircraft 10.

[0054] Referring now to Figures 7-9, schematic diagrams of the flight control commands in the flight control system 28 of Figure 6 and the aircraft's response to those flight control commands are shown.

[0055] In this case, Figure 7 shows an exemplary implementation of the flight control commands in the present invention as a graph, and the corresponding flight commands output by the flight control computer are shown relative to the aircraft's blend speed, which is determined based on appropriate sensor output data provided to the flight control computer. In this specification, the blend speed, which functions as a specific measure of the aircraft's speed, is defined as the ground speed itself in the low ground speed range, while in the intermediate to high speed range, the blend speed is defined as the calibrated airspeed. This choice of blend speed as a functional parameter for the flight control computer is due to the possibility that measurement of the calibrated airspeed may not be available at low airspeeds, and in this speed range, using ground speed may provide greater robustness. In some cases, it may be more reasonable to specify changes between commands based on ground speed rather than airspeed. For example, the introduction of a coordinated turn should occur at a given ground speed, not airspeed, because, for example, in hovering flight with a headwind, it is not desirable for the coordinated turn to be active. However, it should be noted that in other modifications of the present invention, each command may be configured as a function of ground speed only or, alternatively, of calibrated airspeed only.

[0056] In the given example in Figure 7, flight control commands for longitudinal motion control across the entire speed range are output as linear acceleration commands, airspeed rate change commands, or translational velocity commands, depending on the current pivot position of the first stick member relative to the first control axis. For lateral motion control, and therefore control input via the first stick member relative to the second control axis, bank angle flight control commands are output by the flight control computer. In a further possible modification of the present invention, it should be noted that when the blend speed is high, the output command may refer to the roll rate instead of the bank angle.

[0057] Furthermore, the flight control commands for vertical motion control in the example of Figure 7, based on the pivot position of the second stick member relative to the third control axis, are given as altitude change rate command instructions across the entire blend speed range, whereas in an alternative embodiment, at high blend speeds, a transition may occur from the altitude change rate command instruction to a load factor FPA rate instruction. The transition value between the two control strategies may be in the blend speed range of approximately 20 to 60 knots, and at low blend speeds, the change in the aircraft's altitude change rate may be performed at a nearly constant pitch angle of the aircraft.

[0058] Finally, the directional motion control based on the pivot position of the second stick member relative to the fourth control axis in the example of Figure 7 is performed by a command for the rate of change of heading at low blend speeds, and in the blend speed range between 20 knots and 60 knots, the transition to a different strategy is performed when there are three possible options for the command. The three possible options are the delta rate of change of heading, which is the difference between the actual rate of change of heading and the kinematic rate of change of heading related to the bank angle in a coordinated turn, ny (e.g., lateral acceleration along the aircraft axis), or the sideslip angle. All three of these options ensure that in a turn coordination, i.e., when the pilot commands a bank angle or roll rate along the lateral axis while simultaneously holding the directional interceptor neutral, the resulting bank turn is a coordinated turn with zero sideslip and zero ny.

[0059] Based on the control strategies for each of the four control axes, the flight control computer generates a suitable response from the aircraft's control effectors to achieve the motion control instructed by the pilot. Several illustrative schematics of the aircraft's response to flight control commands are shown in Figures 8 and 9. In these, the first graph is shown in panel (I) of Figure 8, which represents the aircraft's response to the operation of the first stick member on the first control axis when the airspeed rate change command is active. It can be seen that the airspeed is maintained constant in calm conditions when the stick member is in the neutral position, while the aircraft's speed can be increased or decreased by pushing or pulling the first stick member. In the case of atmospheric disturbances or during dynamic maneuvers, slight changes in speed may occur, which the pilot can compensate for by applying the corresponding input to the first stick member on the first control axis. Alternatively, a “speed holding” feature may be implemented so that the aircraft maintains a constant airspeed even in the presence of atmospheric disturbances when the stick is in the neutral position.

[0060] It should also be noted that, as a means of protecting the flight envelope, overspeed protection may be implemented in longitudinal motion control, as illustrated by the second graph in panel (I) of Figure 8. In this case, if the aircraft's airspeed exceeds the upper limit for maximum forward speed, the flight control computer will limit the airspeed so that the aircraft stabilizes at VNE when the stick is pushed all the way forward and at VNO when the stick is neutral, even if corresponding input is given to further increase the airspeed.

[0061] Referring now to panel (II) of Figure 8, the difference between aircraft control based on bank angle command instructions and aircraft control based on roll rate command instructions will be explained. The first graph shows that in the airspeed range where bank angle command instructions are used, with the first stick member in a neutral position with respect to the second control axis, the aircraft automatically stabilizes horizontally in a substantially horizontal attitude, while in the airspeed range where roll rate command instructions are used, the aircraft is maintained at a constant roll angle or bank angle in the absence of control input and with the first stick member in a neutral position. Similarly, it should be noted that with respect to this lateral motion control, specific means of flight envelope protection may be provided, for example, with respect to the maximum allowable roll rate and / or bank angle, which may particularly depend on the aircraft's current airspeed.

[0062] Here, panel (I) of Figure 9 illustrates different strategies available in the present invention with respect to vertical motion control. First, for example, during hovering flight at blend speeds of less than 20 knots, only the direct lift provided by the engine's vertical thrust is used for vertical motion control, and by increasing or decreasing the engine's RPM, and therefore the engine's thrust, the aircraft's vertical acceleration and changes in the aircraft's altitude can be achieved while keeping the aircraft substantially at zero pitch angle. On the other hand, during forward flight at blend speeds faster than 20 knots, lift due to the angle of attack is gradually added. Thus, at low speeds, vertical acceleration is achieved at zero pitch angle, and at higher speeds, vertical acceleration is achieved by a combination of pitch, and therefore angle of attack, and direct lift.

[0063] The aircraft flight control computer of the present invention is configured to convert corresponding vertical motion control commands input by a second stick member into a suitable flight control command, which results in the pivoting of the engine-supporting flaps and / or a change in engine thrust, depending on the aircraft's current airspeed. For this purpose, panel (II) of Figure 9 illustrates the difference between altitude rate-based vertical motion control and load factor or FPA rate-based vertical motion control. When the flight control computer outputs an altitude rate-based command, the aircraft altitude is changed only when the second stick member is displaced from its neutral position, while when the second stick member is in the neutral position, the flight control computer always returns the aircraft to level flight. On the other hand, when the flight control computer outputs a load factor command or an FPA rate command, the displacement of the second stick member relative to the third control axis commands a change in the aircraft's rate of climb, which remains constant when the second stick member is returned to the neutral position. Therefore, in order to reduce the rate of climb again and bring the aircraft back to level flight, the pilot must manually reduce the rate of climb by reciprocating the second stick member until the rate of climb reaches zero.

[0064] In this context, different strategies for protecting the flight envelope from vertical motion control may be used, for example, upper and lower limits on the angle of attack, rate of climb, rate of descent, load factor, pitch angle, and / or flight path angle may be set by the flight control computer and implemented by pilot input commands.

[0065] Regarding directional control, a brief reference to panel (III) of Figure 9 shows that in hovering flight, directional control is performed by converting the corresponding pilot input to the fourth control axis on the second stick member into a rate of change of heading command, while the aircraft is kept in a nearly horizontal attitude. Furthermore, the "heading hold" feature may be implemented in hovering flight so that the aircraft maintains a constant heading even in the presence of atmospheric disturbances when the stick is in the neutral position. In forward flight, the control strategy may be progressively blended into ny or sideslip command instructions issued by the flight control computer.

Claims

1. A flight control system (28) for a VTOL aircraft (10) that defines an aircraft reference coordinate system having a roll axis (Xb), a pitch axis (Yb), and a yaw axis (Zb), - First and second manual control devices (23a, 23b) for inputting control commands by an operator (P), The first manual control device (23a) is ○The device comprises a first stick member (24) having a grip portion for the operator (P) to grasp, The first stick member (24) is mounted on the first base member (22b) so as to be able to swing around the first and second control axes (24a, 24b) with respect to a first neutral position. The second manual control device (23b) described above is ○The device comprises a second stick member (26) having a grip portion for the operator (P) to grasp, The second stick member (26) is mounted on the second base member (22b) so as to be able to swing around the third and fourth control axes (26a, 26b) with respect to a second neutral position. First and second manual control devices (23a, 23b), - A flight control computer (30) connected to the first and second manual control devices (23a, 23b) and configured to output flight control commands based on the pivot positions of the first and second stick members (24, 26) with respect to the first to fourth control axes (24a, 24b, 26a, 26b), ○ For vertical motion control based on the pivot position of the first stick member (24) with respect to the first control axis (24a), ○ For lateral motion control based on the pivot position of the first stick member (24) with respect to the second control axis (24b), ○ For vertical motion control based on the pivot position of the second stick member (26) with respect to the third control axis (26a), and, ○ For directional motion control of the second stick member (26) based on the pivot position of the second stick member (26) with respect to the fourth control axis (26b) A flight control computer (30) is configured to derive and output flight control commands while at least partially eliminating the interconnection between individual directions of motion, A flight control system (28) equipped with the following:

2. In the first and second neutral positions, the first and second stick members (24, 26) are inclined relative to each other. Preferably, the first neutral position corresponds to the orientation of the first stick member (24) substantially aligned with the yaw axis (Zb) of the aircraft (10), and the second neutral position corresponds to the orientation of the second stick member (26) that forms a predetermined angle with respect to both the roll axis (Xb) and the yaw axis (Zb) of the aircraft (10), preferably in the range of 30° to 60°, and more preferably approximately 45°, the flight control system (28) according to claim 1.

3. The flight control system (28) according to claim 1 or 2, wherein the first and third control axes (24a, 26a) substantially correspond to the pitch axis (Yb) of the aircraft.

4. The flight control computer (30) is preferably configured to derive and output the flight command for vertical motion control as an altitude change rate command command over the entire airspeed range of the aircraft (10), according to any one of claims 1 to 3, the flight control system (28).

5. The flight control computer (30) is configured to derive and output the flight commands for longitudinal motion control as translational speed command commands over at least a portion of the airspeed range of the aircraft (10), particularly over the low airspeed range, according to any one of claims 1 to 4, the flight control system (28).

6. The flight control system (28) according to claim 4 or 5, wherein the flight control computer (30) is configured to set the translational velocity and / or rate of altitude change to zero while the corresponding first and / or second stick members (24, 26) are in a neutral position with respect to the first or third control axis (24a, 26a).

7. The flight control computer (30) is configured to derive and output the flight commands for longitudinal motion control as airspeed rate change command commands over at least a portion of the airspeed range of the aircraft (10), particularly over the high airspeed range, according to any one of claims 1 to 6, the flight control system (28).

8. The flight control computer (30) is configured to derive and output the flight command for lateral motion control as a bank angle command, and / or to derive and output the flight command for directional motion control as a rate of change of heading command, over at least a portion of the airspeed range of the aircraft (10), particularly over the low airspeed range, the flight control system (28) according to any one of claims 1 to 7.

9. The flight control computer (30) is further configured to evaluate the transition speed value between the low airspeed range and the high airspeed range based on the aircraft's ground speed or calibrated airspeed, according to any one of claims 4 to 7, the flight control system (28).

10. The aforementioned flight control computer (30) further: - Flight envelope protection for vertical motion control by setting upper and lower limits for angle of attack, rate of climb, rate of descent, load factor, pitch angle, and / or flight path angle. - Flight envelope protection for longitudinal motion control by setting upper and lower limits for maximum forward airspeed and maximum reverse airspeed. - Flight envelope protection for lateral and directional motion control by setting upper and lower limits for bank angle, lateral acceleration, and sideslip angle. It is configured to perform at least one of the following: And / or, if a failure of a control effector or loss of sensor data is detected, the flight control computer (30) automatically reconfigures itself to adapt to a specific failure condition, the flight control system (28) according to any one of claims 1 to 9.

11. The flight control system (28) according to any one of claims 1 to 10, further comprising at least one sensor unit (32) operably connected to the flight control computer (30) and configured to output data representing at least one motion parameter of the aircraft (10), wherein the flight control computer (30) is further configured to modify the flight control command based on the received sensor output data.

12. A VTOL aircraft (10) having an aircraft reference coordinate system having a roll axis (Xb), a pitch axis (Yb), and a yaw axis (Zb), - Torso (12) and, - A pair of main wings (14), - A pair of canard wings (16), - A plurality of propulsion units (18), particularly electric ducted fan engines, are distributed on the main wing (14) and the canard wing (16). - A flight control system (28) according to any one of claims 1 to 11, Equipped with, The VTOL aircraft (10) is configured such that the propulsion unit (18) is pivotable about at least one axis along a certain angular position range, and both the angular position and thrust output of the propulsion unit (18) are individually controllable by the flight control system (28).

13. The VTOL aircraft (10) according to claim 11, wherein the aircraft is configured to transition between a hovering flight mode in which the required lift is mainly generated by the vertical thrust of the propulsion unit (18) and a forward flight mode in which the required lift is mainly generated aerodynamically by the main wing (14) and the canard wing (16).

14. The aforementioned flight control system (28) further, - The aircraft (10) is positioned horizontally at a substantially constant pitch angle. - In forward flight mode, to correct the pitch angle of the aircraft (10), A VTOL aircraft (10) as described in claim 12.

15. A method for controlling the flight of a VTOL aircraft (10) according to any one of claims 12 to 14, using a flight control system (28) according to any one of claims 1 to 11, - A step of evaluating the current pivot position of the first and second stick members (24, 26) with respect to the first to fourth control axes (24a, 24b, 26a, 26b), - With respect to longitudinal motion control, lateral motion control, vertical motion control, and directional motion control, the steps include deriving a flight control strategy based on the pivot positions of the first and second stick members (24, 26), - A step of controlling at least the propulsion unit (18) of the aircraft (10), in particular the angular position and thrust output of the propulsion unit (18), preferably a step of controlling only the propulsion unit (18), in accordance with the flight control strategy, A method that includes this.

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