JP6585673B2 - Aircraft attitude control method - Google Patents

Description

  Aircraft such as unmanned aerial vehicles (UAVs) can be used to perform surveillance, reconnaissance, and exploration operations for military and civilian applications. Such aircraft can carry loads that perform specific functions. The aircraft can be a multiple rotorcraft.

A typical flight control method for multiple rotorcraft utilizes cascaded proportional integral derivative (PID) control. In this case, the attitude control is cascaded with the angular velocity control. Based on the conventional PID adjustment method, the parameters of the inner loop of control (angular velocity loop) and the outer loop of control (angular loop) are sequentially adjusted. The calibration results of the inner loop, there is a dependency strong. If the tracking performance of the inner loop is not accurate, it directly affects the overall result. However, the conventional PID adjustment process is complex and redundant, and system divergence and instability problems are likely to occur during the process. Furthermore, the conventional control method adjusts only after the aircraft has already generated an angular velocity due to the disturbance, and under certain circumstances, the disturbance removal performance cannot achieve the optimum state.

  In some cases, it may be desirable to control the flight of an aircraft such as an unmanned aerial vehicle (UAV). It may be further desirable to control the attitude of an aircraft in flight using techniques that are less complex and redundant than conventional attitude control methods. There is a need to control the aircraft so that it can fly stably.

  An embodiment of the present invention is directed to a method for controlling the attitude of an aircraft, and the method includes the following steps (a) to (e): (A) calculating one or more aircraft configuration parameters based on one or more physical characteristics of the aircraft; (b) receiving at the processor a signal indicative of the target attitude of the aircraft; A command signal delivered to at least one actuator of the aircraft operatively connected to one or more propulsion units, (1) a signal indicating the target attitude of (b), and (2) one of (a) Generating using a processor based on one or more aircraft configuration parameters and further using a feedback control scheme; (d) enabling aircraft to operate aircraft dynamics resulting from the operation of one or more propulsion units; Measuring with one or more sensors connected to (e) supplying the dynamics to the processor and (c) adjusting or verifying the command signal. Steps leading to feedback control system.

  The aircraft may be an unmanned aircraft. The aircraft may include a plurality of actuators operably connected to the plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection.

  In some embodiments, the user may enter one or more physical characteristics. One or more physical characteristics of the aircraft may include physical dimensions and weight. The method may include calculating an aerodynamic center and center of gravity of the aircraft. The method may also include calculating the moment of inertia of the aircraft. Calculations using the feedback control system may include feedforward calculations using aircraft moments of inertia.

  Calculations using the feedback control scheme may be performed on aircraft attitudes about the pitch axis, roll axis, and yaw axis. The method further includes using the mixer to calculate the command signal delivered to the at least one actuator by combining the results of the calculations for the pitch, roll, and yaw axes and the aircraft configuration parameters. obtain. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  Aircraft dynamics may include aircraft attitude with respect to at least one axis, angular velocity with respect to at least one axis, and angular acceleration with respect to at least one axis.

  Further embodiments of the present invention include: (a) calculating one or more aircraft configuration parameters based on one or more physical characteristics of the aircraft; (b) receiving a signal indicative of the target attitude of the aircraft; (C) comprising one or more processors configured individually or collectively for generating command signals to be delivered to at least one actuator of the aircraft operably connected to one or more propulsion units of the aircraft; An aircraft attitude control system may be targeted. Here, the command signal is generated using a feedback control method based on the signal indicating the target attitude of (1) and (b) and one or more aircraft configuration parameters of (2) and (a). One or more sensors are operably connected to the aircraft and measure aircraft dynamics resulting from operation of the one or more propulsion units. The measured dynamics are fed to one or more processors, resulting in a feedback control scheme that adjusts or validates the command signal of (c).

  The aircraft may be an unmanned aircraft. The aircraft may include a plurality of actuators operably connected to the plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection.

  The user may enter one or more physical characteristics. One or more physical characteristics of an aircraft include physical dimensions and weight. One or more processors may be configured individually or collectively to calculate the aerodynamic center and center of gravity of the aircraft. One or more processors may be configured individually or collectively to calculate the moment of inertia of the aircraft. Calculations using the feedback control system may include feedforward calculations using aircraft moments of inertia.

  Further, the calculation using the feedback control method may be performed with respect to the aircraft attitude around the pitch axis, the roll axis, and the yaw axis. The system may include a mixer that combines the results of the calculations for the pitch, roll, and yaw axes, and aircraft configuration parameters to calculate a command signal that is delivered to at least one actuator. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  Aircraft dynamics may include aircraft attitude with respect to at least one axis, angular velocity with respect to at least one axis, and angular acceleration with respect to at least one axis.

  A method for controlling aircraft attitude may be provided in accordance with another embodiment of the present invention. The method may include the following steps (a) to (e). (A) evaluating a non-linear relationship between actuator thrust and actuator output using a processor; (b) receiving a signal indicative of a target attitude of the aircraft; (c) one or more of the aircraft A command signal delivered to at least one actuator of the aircraft operatively connected to the propulsion unit of the aircraft, based on (1) a signal indicating the target attitude of (b), and (2) based on the non-linear relationship of (a) And using a feedback control scheme to generate with the processor, (d) one or more sensors operatively connected to the aircraft, the aircraft dynamics resulting from the operation of one or more propulsion units (E) providing feedback to the processor and adjusting or verifying the command signal of (c) Steps leading to the expression.

  In some embodiments, the aircraft is an unmanned aerial vehicle. The aircraft can include a plurality of actuators operably connected to a plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection. The non-linear relationship may be entered by the user. The non-linear relationship can be calculated during calibration of one or more actuators in the aircraft. The method may include calculating an aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft. The method may further include calculating an aircraft moment of inertia based on aircraft physical characteristics. Calculations using the feedback control system may include feedforward calculations using aircraft moments of inertia.

  Calculations using the feedback control scheme may be performed on aircraft attitudes about the pitch axis, roll axis, and yaw axis. The method further includes using the mixer to calculate the command signal delivered to the at least one actuator by combining the results of the calculations for the pitch, roll, and yaw axes and the aircraft configuration parameters. obtain. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  Aircraft dynamics may include aircraft attitude with respect to at least one axis, angular velocity with respect to at least one axis, and angular acceleration with respect to at least one axis.

  Furthermore, as an embodiment, the present invention may provide an aircraft attitude control system that is configured individually or collectively, and includes one or more processors that execute the following (a) to (c). (A) evaluate a non-linear relationship between actuator thrust and actuator output; (b) receive a signal indicative of the target attitude of the aircraft; and (c) be operatively connected to one or more propulsion units of the aircraft. Generating a command signal to be delivered to at least one actuator of the aircraft. Here, the command signal is generated using a feedback control method based on the signal indicating the target posture of (1) and (b) and the nonlinear relationship of (2) and (a). One or more sensors are operably connected to the aircraft and measure aircraft dynamics resulting from operation of the one or more propulsion units. The measured dynamics are fed to one or more processors, resulting in a feedback control scheme that adjusts or validates the command signal of (c).

  In certain embodiments, the aircraft may be an unmanned aerial vehicle. The aircraft can include a plurality of actuators operably connected to a plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection. The user may enter a non-linear relationship. The non-linear relationship can be calculated during calibration of one or more actuators in the aircraft. One or more processors are configured individually or collectively to calculate the aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft. One or more processors may be configured individually or collectively to calculate the moment of inertia of the aircraft based on the physical characteristics of the aircraft. Calculations using the feedback control system may include feedforward calculations using aircraft moments of inertia.

  Calculations using the feedback control scheme may be performed on aircraft attitudes about the pitch axis, roll axis, and yaw axis. The system may also include a mixer that combines the results of the calculations for the pitch, roll, and yaw axes, and aircraft configuration parameters to calculate a command signal that is delivered to at least one actuator. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  Aircraft dynamics may include aircraft attitude with respect to at least one axis, angular velocity with respect to at least one axis, and angular acceleration with respect to at least one axis.

The present invention can further provide a method for controlling an aircraft attitude as an embodiment thereof. The method may include the following steps (a) to (d). (A) receiving a signal indicative of a target attitude of the aircraft with a processor; (b) a command signal delivered to at least one actuator of the aircraft operatively connected to one or more propulsion units of the aircraft; Using a feedback control scheme based on (a) a signal indicating the target attitude and including (1) an angular acceleration loop with angular acceleration feedback, and (2) a direct feedforward calculation based on the target angular acceleration, Generating with a processor, (c) measuring aircraft dynamics resulting from the operation of one or more propulsion units using one or more sensors operably connected to the aircraft, (d) Providing dynamics to the processor to provide a feedback control scheme to adjust or validate the command signal of (b).

  The aircraft may be an unmanned aircraft. The aircraft can include a plurality of actuators operably connected to a plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection. The method may further include calculating an aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft. The method may further include the step of calculating the moment of inertia of the aircraft based on the physical characteristics of the aircraft. The feed forward calculation may use the moment of inertia of the aircraft. Calculations using the feedback control scheme may be performed on aircraft attitudes about the pitch axis, roll axis, and yaw axis.

  The method further includes using the mixer to calculate the command signal delivered to the at least one actuator by combining the results of the calculations for the pitch, roll, and yaw axes and the aircraft configuration parameters. obtain. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  Aircraft dynamics may include aircraft attitude with respect to at least one axis, angular velocity with respect to at least one axis, and angular acceleration with respect to at least one axis.

  Embodiments of the present invention may be directed to an aircraft attitude control system that is configured individually or collectively and includes one or more processors that perform the following operations (a) to (b). (A) receiving a signal indicative of the target attitude of the aircraft; and (b) generating a command signal that is delivered to at least one actuator of the aircraft operatively connected to one or more propulsion units of the aircraft. Here, the command signal is generated using a feedback control method based on the signal indicating the target posture of (1) and (b) and the nonlinear relationship of (2) and (a). One or more sensors are operably connected to the aircraft and measure aircraft dynamics resulting from operation of the one or more propulsion units. The measured dynamics are fed to one or more processors, resulting in a feedback control scheme that adjusts or validates the command signal of (c).

  In certain embodiments, the aircraft may be an unmanned aerial vehicle. The aircraft can include a plurality of actuators operably connected to a plurality of propulsion units. The propulsion unit may include a rotor wing that generates the lift of the aircraft.

  A signal indicating the target attitude of the aircraft can be received from a remote controller over a wireless connection. One or more processors may be configured individually or collectively to calculate the aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft. One or more processors may be configured individually or collectively to calculate the moment of inertia of the aircraft based on the physical characteristics of the aircraft. The feed forward calculation may use the moment of inertia of the aircraft. Further, the calculation using the feedback control method may be performed with respect to the aircraft attitude around the pitch axis, the roll axis, and the yaw axis.

  The system may include a mixer configured to combine the results of the calculations for the pitch, roll, and yaw axes, and aircraft configuration parameters to calculate a command signal that is delivered to at least one actuator. The aircraft configuration parameter may be the distance from the actuator to the aerodynamic center of the aircraft. The one or more sensors may be inertial sensors.

  According to certain embodiments, aircraft dynamics may include an aircraft attitude with respect to at least one axis, an angular velocity with respect to at least one axis, and an angular acceleration with respect to at least one axis.

  It should be understood that the various embodiments of the invention may be recognized individually, collectively, or in combination with each other. The various embodiments of the present invention described herein may be applied to any particular application shown below, or any other type of movable object. Any description of the invention relating to an aircraft, such as an unmanned aerial vehicle, can be applied and used for any movable object, such as any aircraft. In addition, the systems, devices, and methods described herein in the context of aerial motion (eg, flight) may also include other types of motion, such as ground or water motion, underwater motion, or space motion. Can be applied in the context of

  Other objects and features of the present invention will become apparent upon review of the specification, claims and appended drawings.

(Incorporation by reference)
All publications, patents, and patent applications mentioned in this specification are referenced to the same extent as if each individual publication, patent, or patent application was specifically and individually indicated and incorporated by reference. Is incorporated herein by reference.

  The novel features of the invention are set forth with particularity in the appended claims. A better understanding of the features and advantages of the present invention will be obtained by reference to the following detailed description that sets forth illustrative embodiments, in which the principles of the invention are utilized, and the accompanying drawings of which: I will.

It is a figure which shows the multistage attitude | position control method to which this invention is applied. FIG. 2 illustrates how aircraft physical parameters can be associated with an aircraft attitude control method in accordance with an embodiment of the present invention. FIGS. 4A to 4C are diagrams illustrating various physical characteristics that can be considered for one or more physical parameters of an aircraft, according to embodiments of the present invention. FIG. 6 illustrates various physical characteristics that can be considered for one or more physical parameters of an aircraft, according to embodiments of the present invention. FIG. 2 illustrates how an aircraft model can be used to determine one or more parameters in an aircraft control method according to an embodiment of the present invention. It is a figure which shows the aircraft with a flight controller to which this invention is applied. It is a figure which shows the attitude | position control system which can implement the aircraft in connection with embodiment of this invention. It is a figure which shows the attitude | position control system which can implement the aircraft in connection with embodiment of this invention. FIG. 3 is a diagram showing a part of a control inner loop according to an embodiment of the present invention. It is a figure which shows the attitude | position control system in connection with embodiment of this invention. It is a figure which shows the tracking error in connection with embodiment of this invention. FIG. 6 is a further diagram illustrating a tracking error according to an embodiment of the present invention. FIG. 4 shows a comparison between the response of a controller to which the present invention is applied compared to a conventional controller. It is a figure showing an unmanned aerial vehicle in connection with an embodiment of the present invention. It is a figure which shows the movable object in connection with embodiment of this invention including a support mechanism and a load. It is a block diagram of the system which controls a movable object in connection with embodiment of this invention.

  The present invention provides systems, devices, and methods for performing attitude control of aircraft flight. The aircraft may be an unmanned aerial vehicle (UAV) or any other type of movable object. The aircraft may be a multi-rotor wing aircraft.

  A typical attitude control method for a multi-rotor aircraft control scheme utilizes cascaded proportional integral derivative (PID) control. In this case, the attitude control is cascaded with the angular velocity control. Based on a conventional PID adjustment method, the parameters of the inner loop (angular velocity loop) and the outer loop (angular loop) may be adjusted sequentially.

  For a conventional cascaded controller, it relies heavily on the inner loop calibration results, and if the inner loop tracking performance is not good, it directly affects the overall results. Therefore, the inner loop design and parameter adjustment are very important. However, the conventional PID adjustment process is complex and redundant, and system divergence and instability problems are likely to occur during the process. Control parameters are typically standard using actual adjustment results, but are highly system dependent. If the system changes (e.g., aircraft structure, rotor wing distance, weight changes), the parameters need to be adjusted again. In addition, the adjustment period is relatively long. The main reason is due to the lack of complete understanding of the system model and there is no relationship between the parameters.

  Under conventional cascaded PID control, the inner angular velocity loop is designed primarily to withstand velocity disturbances. Thus, under normal circumstances, the controller only makes adjustments after the aircraft has already generated an angular velocity due to disturbance. Under certain circumstances, disturbance rejection performance cannot achieve optimal conditions.

  Improved flight control methods and systems are provided in the present invention. For example, the flight attitude control method can take into account at least one of aircraft physical parameters or dynamic parameters. For example, according to aircraft structure and propulsion, performance limits can be automatically evaluated and used as a basis for control. An aircraft model can be used as a radix and at least one of physical parameters or dynamic parameters can be a coefficient of the controller. This can facilitate adjusting PID control for different aircraft models. By being able to perform prior assessment, the control system can be distinguished for different aircraft models and can be easily applied to different models. In one example, nonlinear parameters can be considered.

  In some embodiments, aircraft physical parameters may be evaluated to determine the moment of inertia of the entire aircraft. The physical parameters can also be evaluated to calculate the aerodynamic center of the aircraft and to determine the axial distance from the aircraft motor or propulsion unit to the aerodynamic center. The physical parameters can also be used to determine a motor thrust / lift curve for one or more motors of the aircraft. These parameters may be aircraft configuration parameters that can be used for aircraft attitude control. These parameters can be evaluated in real time or pre-evaluated. Aircraft configuration parameters may be associated with various models of aircraft. Aircraft models can be selected and appropriate aircraft configuration parameters can be applied to attitude control methods.

  Further, the improved flight control method may include adding an angular acceleration loop to the PID control scheme. The flight control method may also include performing direct control and enhancing disturbance rejection and tracking performance. Since the angular acceleration loop can be operated as a direct control, the response time can be shortened due to strong disturbance resistance characteristics. By directly suppressing disturbance, response time can be reduced.

  FIG. 1 shows a multistage attitude control method to which the present invention is applied. Multi-stage attitude control can be used to control the attitude of an aircraft, such as a UAV or any other type of aircraft. The aircraft may be manned or unmanned. The aircraft may be a multi-rotor aircraft that may include two or more rotors that can provide lift to the aircraft. The aircraft may be a single rotorcraft. Any statement herein relating to aircraft flight control is applicable to any other movable object. For example, the attitude control method described in this specification may be applied to at least one of a spacecraft or an underwater ship. One or more embodiments of attitude control may apply to an aircraft, an out-of-atmosphere aircraft, a land aircraft, or a water aircraft.

  Attitude control can be implemented around one or more axes of rotation about the aircraft. For example, the attitude control may be implemented around at least one of the pitch axis, the roll axis, and the yaw axis. Attitude control may be implemented around one of these axes, two of these axes, or all three of these axes.

  The attitude control method can be implemented in multiple stages. For example, a pre-configuration stage and a flight stage can be provided. The pre-configuration stage and the flight stage can be performed at different times or at least at different locations.

  In some embodiments, the pre-configuration phase may be performed while one or more physical parameters of the aircraft may be evaluated. One or more calculations may be based on aircraft physical parameters. Aircraft physical parameters may include spatial dimensions such as height, width, length, diameter, diagonal, or circumference. The physical parameters may also take into account the morphology of the aircraft, such as the shape of the aircraft fuselage or at least one of the elongated portions. The physical parameters may also take into account other factors such as weight, weight distribution, center of gravity, or density. Additional examples of physical parameters are described elsewhere herein.

  Dynamic parameters can be evaluated during the pre-setting phase. Alternatively, dynamic parameters may be considered at a later stage. Dynamic parameters may include battery related specifications, such as thrust or power, or other power supply specifications, or motor characteristics. Additional examples of dynamic parameters are described elsewhere herein.

  The pre-configuration phase can be performed on one or more models of the aircraft. For example, different aircraft models may have at least one of different physical parameters or dynamic parameters. For example, different aircraft models may have different shapes, sizes, weights, weight distributions, power supply characteristics, motor characteristics, or other different features or characteristics. Also, preconfiguration can be performed by the manufacturer or distributor of one or more models of the aircraft. Preconfiguration can be performed by a manufacturer or distributor of a control system that can be applied to one or more models of the aircraft. Preconfiguration can also be done by any third party that can provide information from the preconfiguration to assist in controlling the aircraft. The information can be stored in a memory that can be accessed by a flight controller that controls the aircraft. Also, preconfiguration is not performed by the end user of the aircraft. For example, the pre-configuration may be performed by a different entity than the user operating the aircraft. Preconfiguration can be performed before the end user accesses the aircraft. For example, preconfiguration may be performed hours, days, weeks, months, quarters, or years before the user accesses the aircraft or the user operates the aircraft.

  In some embodiments, the pre-configuration can be performed as an aircraft calibration. Calibration may be performed before the user accesses the aircraft or may be independent of the user's interaction with the aircraft. The aircraft may be able to access preconfigured information. The pre-configuration information can be stored on the aircraft or can be accessible via the aircraft from memory outside the aircraft.

  Also, the flight phase can be performed after the pre-configuration phase. The flight phase can be provided when the user accesses the aircraft. The flight phase can be when the user can operate the aircraft. During the flight phase, the attitude control method can be used on an aircraft. The attitude control system may be used by an aircraft to control the attitude of the aircraft during flight. The attitude control system may use pre-configuration information that can be collected early during the pre-configuration phase.

  Further calibration may or may not be performed during the flight phase. In some embodiments, several calibrations can occur each time the user turns on the aircraft. Alternatively, calibration may be performed when the user first receives the aircraft and operates it for the first time. In another embodiment, calibration may be performed on demand. The calibration information may or may not be used to determine the physical dimensions of the aircraft. The calibration information may or may not be used to determine aircraft dynamic characteristics. The flight stage calibration information may or may not be combined with pre-configuration information for determining one or more coefficients for the flight attitude control method.

  FIG. 2 shows how aircraft physical parameters can be associated with an aircraft attitude control method according to an embodiment of the present invention. In certain embodiments, these steps can be performed during the pre-configuration phase.

  One or more physical parameters of the aircraft can be evaluated. In some embodiments, these parameters may be measured manually by a human or automatically using a processor. In addition, a living organism may input data using a device and store it in a memory. For example, a human can measure the dimensions of an aircraft and enter data. In another embodiment, one or more machines may be used to determine aircraft physical parameters, and the data may be automatically provided to a memory storing information. For example, the weight of the aircraft may be measured and automatically communicated to the memory.

  Examples of physical parameters include factors related to at least one of aircraft structure or aircraft dynamics. Physical parameters can include spatial dimensions such as height, width, length, diameter, diagonal, or circumference. The physical parameters may also take into account the morphology of the aircraft, such as the shape of the aircraft fuselage or at least one of the elongated portions. The physical parameters may also take into account other factors such as weight, weight distribution, center of gravity, or density. One or more material properties of the aircraft may also be taken into account, such as density, stiffness, flexibility, or elasticity may be taken into account. Such physical parameters may be related to the aircraft structure. The physical parameters can be collected for the entire aircraft or at least one of one or more components of the aircraft. For example, the physical parameters may relate to a fuselage frame, power supply (eg, battery), avionics system, support mechanism, payload, sensor, motor, landing gear, communication unit, or any other component.

  The physical parameters may also include aircraft dynamics. This may include power supply specifications such as maximum battery current, maximum power output, energy density, battery capacity, discharge rate, rated voltage, battery life, or any other feature. In addition, information regarding battery chemistry or battery type can be specified or determined. This may also include motor characteristics such as thrust or power. The motor characteristics may include maximum reliability and maximum power. The current and power for electronic speed control can also be determined.

Any other physical parameter related to the aircraft may be evaluated. Information about one or more of the aircraft's physical parameters can be stored in memory. For example, physical parameter data may be stored in one or more databases. The database may be external to the aircraft. The database can be prepared in a cloud computing environment and / or distributed across multiple devices. Alternatively, the database may be prepared on the aircraft.

  One or more calculations can be performed on the physical parameters to determine one or more aircraft configuration parameters. A physical parameter may represent a physical characteristic of an aircraft that may be directly measurable. Aircraft configuration parameters can be calculated based on physical parameters. Aircraft configuration parameters can be calculated using a processor.

  In certain embodiments, aircraft configuration parameters include, but are not limited to, the aerodynamic center of the aircraft, the center of gravity of the aircraft, or the moment of inertia of the entire aircraft or aircraft components. The aerodynamic characteristics or stability of the aircraft may be analyzed.

  FIG. 3 illustrates various physical characteristics that can be considered for one or more physical parameters of an aircraft according to an embodiment of the present invention. FIG. 3A shows an example of how an aircraft center of gravity can be calculated. As shown, if the aircraft's center of gravity is below the lift plane and the aircraft's lateral flight can reach a constant equilibrium speed, the horizontal component of lift is at least one of drag and the vertical component is at least one of gravity. Is possible. Further, the aerodynamic center and gravity may or may not coincide. In situations where the aerodynamic center and the center of gravity do not coincide, the aircraft can experience a nose-up pitching moment due to the vertical lift component and gravity forming a couple. This can turn the aircraft horizontally, which can make the aircraft a stable system. Therefore, the stability of the aircraft can be adjusted by changing the position of the center of gravity when designing the aircraft. The center of gravity of the aircraft can be calculated based on physical parameters. Alternatively, the center of gravity of the aircraft may be determined by considering weight distribution amounts and position determination of various components of the aircraft. The center of gravity can vary from aircraft model to aircraft model.

  FIG. 3B shows an example of how an aircraft moment of inertia can be calculated. In one embodiment, the moment of inertia distribution for the entire aircraft may be analyzed. Aircraft model effects and payload configurations may be evaluated against their effects on the overall moment of inertia of the aircraft. These can be used as a reference for adjusting the overall configuration of the aircraft.

  A basic physical model that constitutes each component of the aircraft may be established. Examples of aircraft components include, but are not limited to, fuselage frames, power supplies (eg, batteries), avionics systems, support mechanisms, loads, sensors, motors, propulsion units, landing gear, or communication units. The basic model may include conventions relating to installation location, weight, and moment of inertia of each component. Next, the moment of inertia of the entire aircraft based on the moment of inertia of each component, using the equilibrium axis theorem. As shown, at least one of various components or the overall shape of the aircraft, as well as weight distribution, may be considered in the moment of inertia calculation. The moment of inertia can be calculated using a processor based on the collected physical parameters. In some embodiments, finite element analysis (FEM) may be used. The equilibrium axis theorem may be used to calculate the moment of inertia.

An example calculation with moment of inertia calculation can be provided as follows.
1) Material point:
I = M * L ^ 2
2) Regarding part of the cylinder:
I = 1/3 * M * L ^ 2
3) Regarding the center of the cylinder:
I = 1/2 * M * R ^ 2
Where I is the moment of inertia, M is the mass, L is the distance from the center of the mass or cylinder (or any other shape) to the axis of rotation, and R is the radius of the cylinder.

Further calculations may include torque and angular velocity calculations and can be provided as follows.
Torque = L * F
Angular velocity = torque / moment of inertia

  Based on theoretical analysis and calculations of a number of aircraft models, the sought-after primary moment of inertia of the aircraft may be approximately 50% assigned by the electronics and propulsion unit (eg, propeller / rotor blade). The moment of inertia and the rotor blade distance may be connected in a square relationship, and the torque increases linearly, so with the same propulsion, there can be better driving characteristics under small rotor blade distance conditions, and under a large rotor blade distance There may be better stability characteristics.

  In addition, other aircraft configuration parameters that can be considered include at least one of a power (eg, battery) parameter, an actuator (eg, motor) parameter, or an electronic speed control (ESC) parameter.

For example, these aircraft configuration parameters may be calculated using physical parameters. Maximum battery current, the maximum power output, or at least one of energy density, at least one of the maximum thrust or power of the motor, and / or the maximum current or at least hand power of ESC, can be evaluated as follows.
Maximum battery current = battery capacity x discharge rate
Maximum power output = maximum battery current x rated voltage
At least one of maximum motor thrust or power : from experiments
ESC maximum current or power : From experiment

  Furthermore, it may be desirable to evaluate at least one of safety or compliance of the propulsion system. For example, it may be desirable that the current that the battery can provide> the maximum current of the ESC> the current of the motor.

  Further calculation of aircraft configuration parameters from physical parameters may include hover performance evaluation. For example, at least one calculation of hover power output, power usage, efficiency, power, or estimated hover time can be made based on the weight and dynamics of the aircraft. The calculation can be performed using a processor.

  The motor output can be determined from accessing data about the motor in one or more databases. The motor data may also be a look-up table based on aircraft weight and motor tension curves, and efficiency may be determined from a look-up table based on motor power. The lookup table may be created based on empirical test data. Alternatively, the lookup table can be created based on simulated or predicted data. The data on the loop-up table may be input by an individual. The motor tension curve may be non-linear.

One or more hover characteristics can be calculated as follows.
Hover power = weight / efficiency hover current = hover power / voltage time = battery capacity / hover current

  In addition, physical performance can be used to evaluate operating performance. One or more aircraft configuration parameters from operational performance evaluation include thrust to weight ratio, speed parameters (eg, maximum ascent speed, maximum descent speed, upper speed limit based on the designed braking distance), angular parameters (eg, , Theoretical maximum attitude angle, compensated attitude angle, limit attitude angle), parameters related to torque (eg, triaxial torque), parameters related to angular velocity (eg, calculation of maximum angular velocity using braking angle), motor rotation Parameters, such as, but not limited to, calculating at least one of the effects of motor rotation direction on the yaw axis or the associated compensation.

FIG. 3C is a diagram illustrating various physical characteristics that can be considered for one or more physical parameters of an aircraft in accordance with an embodiment of the present invention. One or more of the aircraft configuration parameters can also be calculated as follows.
Thrust-to-weight ratio = Maximum thrust / weight Hover lift (N) = Overall aircraft weight (kg) * 9.8 (m / s ^ 2)
Hover throttle = lookup (lift), unit:%
Hover efficiency = Lookup (hover throttle), Unit: g / Watt Hover braking efficiency = Lookup (Hover throttle), Unit: g / Watt Hover current = Overall aircraft weight / efficiency Hover time = Battery capacity / Hover current Maximum attitude Angle = Arcos (hover lift / maximum lift)

  As shown in FIG. 2, aircraft configuration parameters can be calculated based on one or more physical parameters from the aircraft. Aircraft configuration parameters can be associated with an aircraft model. For example, different aircraft models (eg, UAV model, manned aircraft model) may have different physical parameters and associated aircraft configuration parameters. Aircraft configuration parameters can be collected and / or stored for each of the models. Also, some of the aircraft configuration parameters include empirically received data, data manually entered by the user, modeled or simulated data, or data calculated based on any existing data.

  One or more aircraft configuration parameters include controller parameters that can be used to adjust the aircraft flight controller. For example, aircraft parameters of inertia, motor tension curves, and dynamic acceleration / deceleration performance can be used to automatically adjust controller parameters. Such adjustment can be performed according to the tracking error. Using aircraft configuration parameters can simplify the parameter adjustment process and enhance controller performance.

  In certain embodiments, controller performance can be enhanced by providing at least one of an aircraft moment of inertia, a motor lift output curve, or an axial distance (eg, the distance from the motor to the aerodynamic center). Such parameters can be useful for distinguishing different aircraft models. Such parameters can be associated with the aircraft model. Thus, when a user receives an aircraft, the user may specify an aircraft model or the aircraft model may already be preprogrammed. Certain aircraft model parameters can be used in the flight control system to control the attitude of the aircraft.

  FIG. 3D shows a set of aircraft operating principles according to an embodiment of the present invention. The aircraft may also have three degrees of freedom. For example, an aircraft may be able to rotate about three axes of rotation and to generate lift. For example, in a rotary wing aircraft, the rotary wings can rotate to generate aircraft lift. The rotation of the rotor may allow the aircraft to rotate about one, two, or three axes of rotation at the same time. Also, the axes of rotation may be at least one of mutually orthogonal or remain orthogonal to each other during aircraft flight (eg, θ, φ, ψ). The axis of rotation may include at least one axis of rotation of pitch, yaw, or roll.

In some embodiments, the rotorcraft may have multiple rotor blades (eg, F ₁ , F ₂ , F ₃ , F ₄ ) each capable of generating lift to the rotorcraft. In some embodiments, four rotor blades can be provided. The rotors can produce the same amount of lift or different amounts of lift. The rotor blades may rotate at the same angular velocity or at different angular velocities (eg, ω ₁ , ω ₂ , ω ₃ , ω ₄ ).

  A robust adaptive control strategy may be useful for multi-rotor aircraft. Systems that manage the flight of multi-rotor aircraft can be inherently unstable and can diverge in seconds if appropriate control laws are not applied. The system may be non-linear. System non-linearities and aerodynamic complexity require improved controller design.

  The systems and methods according to the present invention can model aircraft flight dynamics, develop control strategies, and stabilize attitudes of multi-rotor aircraft where the configuration manifold can be non-linear. The dynamic and proposed control system can be represented on the special orthogonal group SO (3) to avoid singularities and ambiguities associated with other posture representations such as Euler angles and quaternions. Modeling a multi-rotor aircraft may include multi-rotor motion and dynamic analysis and system identification of an actuator (eg, at least one of a motor, rotor, or propeller).

  Finite element analysis (FEM) can be used to estimate the moment of inertia of the system. A first order inertial system with a time delay can be regarded as an approximate model for system identification. As described elsewhere herein, the control system can be classified into SO (3) proportional control and kinetic cascade PID (proportional, integral and derivative) control with feedforward compensation. For the control of SO (3), the error can be defined as the natural error of SO (3) (can be provided from a geodetic point of view and may be desirable). A proportional control can be defined on the exponential coordinates of SO (3) because it is a linear space. The control method can be verified by the Lyapunov function to ensure stability on the nonlinear manifold. For dynamic cascade PID control, the controller may construct SO (3) in exponential order, angular velocity, and angular acceleration. A technique using an incomplete differential PID controller and a Smith predictor can be used to further suppress noise and improve control quality. In addition, feedforward compensation can be added to improve transient response. Further descriptions of control schemes for multi-rotor aircraft are set forth elsewhere herein.

  FIG. 4 illustrates how an aircraft model can be used to determine one or more parameters in an aircraft control method according to an embodiment of the present invention. For example, an aircraft may have one or more onboard processors that can function as flight controllers.

  When a user receives an aircraft, the user may be able to enter an aircraft model. For example, a user may enter an aircraft model directly into the aircraft. In another example, the user may enter the aircraft model on an external device that can communicate with the aircraft. The external device may be an aircraft controller or a display device that displays data from the aircraft. As described elsewhere herein, the external device may be a computer, smartphone, tablet, or any other type of device or terminal.

  Also, the aircraft may already be preprogrammed with aircraft model information. The aircraft model information may or may not be changeable.

  In some embodiments, aircraft model information can be used to access one or more aircraft configuration parameters of a selected aircraft model. Data can also be stored in memory for one or more aircraft configuration parameters that can be associated with the aircraft model. For example, aircraft model A and aircraft model B may have one or more different physical characteristics. Different physical characteristics may result in different aircraft configuration parameters, such as at least one of different moments of inertia, motor lift output curves, or axial distances. The flight controller may use different aircraft configuration parameters to control the flight of the aircraft. The data can also be stored as a lookup table that may be accessible to various configuration parameters for different models of aircraft. For example, if requests for configuration parameters of aircraft model X are made, they can be provided from a lookup table. The look-up table can also be stored on the aircraft. Thus, a user can enter or modify an aircraft model to define or modify configuration parameters used by the flight controller. Alternatively, the look-up table may be stored outside the aircraft. The aircraft may be able to communicate with external devices and access data from the lookup table. For example, an aircraft can transmit an aircraft model indicator and an external device can transmit aircraft configuration parameters associated with the selected aircraft model.

  The aircraft may also have one or more aircraft configuration parameters preprogrammed therein. Aircraft configuration parameters may be stored on the aircraft when the user receives the aircraft. The user may or may not need to specify an aircraft model. Also, decisions about aircraft models may be made at the manufacturer's site or at other sites, and the flight controller is pre-programmed with aircraft configuration parameters that can be determined based on the aircraft's physical parameters. Also good. Aircraft configuration parameters can be accessed from a look-up table containing data for multiple aircraft models. For example, at the manufacturer's site, the manufacturer can specify that aircraft model X is being manufactured, access the associated aircraft configuration parameters, and pre-program the aircraft. The user may or may not be able to modify aircraft configuration parameters. The user may also be able to enter requests for new configuration parameters of the aircraft (either directly on the aircraft or using an external device that can communicate with the aircraft). Such a request may be made based on the aircraft model or may include the input of new physical characteristics of the aircraft. Also, aircraft configuration parameters can be pre-calculated and stored in the data so that it can be accessed on demand. Alternatively, new physical parameter data can be entered or measured and new aircraft configuration parameters can be calculated. Such a calculation can be performed in real time. For example, a user can modify an existing aircraft in a manner that can change one or more aircraft configuration parameters. For example, the user may add a new camera to the UAV that can change at least one of weight or weight distribution. At least one of moment of inertia or other aircraft configuration parameters can be recalculated to accommodate the change.

  In some embodiments, information about the aircraft may be stored in at least one of the aircraft's onboard memory or processor. The processor may evaluate one or more parameters when turning on the aircraft. The processor can also access pre-calculated parameters based on the stored aircraft model. In other examples, certain diagnostics or measurements may be made when the aircraft is turned on and may be used to generate one or more aircraft configuration parameters.

  During flight, the aircraft can be controlled by input from a flight control device. The flight control device may be an external device that is separate from the aircraft. The flight control device may also be a remote control operated by the user on the ground while the aircraft is flying. The flight control device can communicate with the aircraft wirelessly. Alternatively, the flight control device may be incorporated into the aircraft. For example, the user can operate the flight control device while the user is on board the aircraft. The user may be an aircraft pilot and can operate the flight control device from the cockpit. The flight control device may include information that may relate to at least one of aircraft direction or speed. Input from the flight control device can be used to determine the target attitude of the aircraft. The target attitude of the aircraft can be determined around one, two, or three axes of rotation. For example, the target attitude of the aircraft can be determined around at least one of the pitch axis, roll axis, or yaw axis.

  The aircraft may have a flight controller. The flight controller can include one or more processors on the aircraft. The flight controller can receive a signal indicative of input from the flight control device. The flight controller can control the flight of the aircraft in response to input from the flight control device. The flight controller can control the flight of the aircraft in response to one or more of the aircraft flight configuration parameters. The flight controller may perform aircraft attitude control based on signals from the flight control device (eg, target attitude) and aircraft configuration parameters. The flight controller may perform attitude control around the aircraft roll, yaw, and pitch axes based on the target roll, yaw, and pitch axes and aircraft configuration parameters.

  FIG. 5 shows an aircraft having a flight controller to which the present invention is applied. Aircraft 510 may have one or more onboard flight controllers 520. The flight controller may include one or more processors that can individually or collectively generate command signals for controlling aircraft flight.

  The flight controller 520 can communicate with one or more actuators 560a, 560b of the aircraft. The actuator can be a motor that can be connected to one or more propulsion units of the aircraft. The propulsion unit may include a rotating wing that rotates to generate aircraft lift. In some embodiments, the aircraft may be a multi-rotor aircraft having a plurality of rotors, each of which can generate aircraft lift. The command signal can determine the output supplied to the motor. This can determine the speed at which the rotor connected to the motor can rotate. Each rotor can be connected to a separate motor. In addition, one rotor blade may be connected to a plurality of rotor blades, or a plurality of motors may be used to drive a single rotor blade. The motors can be individually controllable. For example, one motor may have a different power output than another motor in a different environment. The propulsion units may all be the same type of propulsion unit or may include different types of propulsion units. For example, all propulsion units may include a rotor blade / propeller. In certain embodiments, at least one of the rotor blades or propellers may have at least one of the same configuration or dimensions, or at least one of different configurations or dimensions. Any number of motors or propulsion units may be provided on the aircraft. For example, one, two, three, four, five, six, seven, eight, nine, ten, eleven, twelve, or more motors or propulsion units Can be provided on an aircraft.

  Each motor can be controlled individually. For example, separate command signals may be provided for each motor. Each motor may have the same or different motor output as other motors in the aircraft. The output to each motor may vary depending on the desired target attitude of the aircraft. For example, if it is desired to adjust the attitude of the aircraft, one or more motors can operate at different outputs (eg, rotate at different speeds or rpms) to change the attitude of the aircraft.

  In some embodiments, a memory that includes one or more aircraft configuration parameters 530 may be provided on the aircraft. The aircraft configuration parameters may include at least one of an overall aircraft moment of inertia, a motor lift output curve, or an axial distance (eg, a distance from the motor to the aerodynamic center). Other aircraft configuration parameters can be stored, such as those described elsewhere. Aircraft configuration parameters may be derived from one or more physical characteristics of the aircraft or aircraft model. Aircraft configuration parameters can be preprogrammed into memory. Alternatively, aircraft configuration parameters may be downloaded to memory from an external device. Aircraft configuration parameters may be stored in memory and flight controller 520 may be used in generating command signals.

  One or more sensors 540 may also be provided on the aircraft. Examples of sensors include imaging devices (eg, cameras, vision sensors, infrared / thermal imaging devices, UV imaging devices, or other types of spectral imaging devices), inertial sensors (eg, gyroscopes, accelerometers, magnetometers). , Ultrasonic sensors, riders, sonars, or any other type of sensor, including but not limited to. Sensors can also communicate with external devices such as global positioning system (GPS) satellites. The sensor can be a GPS receiver. In other examples, the sensor can communicate with one or more towers or repeaters. The sensor can collect information about the environment surrounding the aircraft. The sensor may or may not be used to assist in aircraft navigation. Sensor 540 can also communicate with aircraft flight controller 520. In addition, the flight controller may use signals from the sensors in generating command signals for one or more actuators. The signal from the sensor may or may not be used in controlling the attitude of the aircraft about one or more axes.

  In certain embodiments, the sensors may be useful for collecting information about aircraft dynamics. For example, a sensor can be used to collect information about at least one of aircraft attitude, angular velocity, or angular acceleration about one or more axes of rotation. For example, the sensors include a gyroscope or other sensor that can collect information about at least one of aircraft attitude, angular velocity, or acceleration about at least one of the pitch, roll, or yaw axes. obtain. The sensor may be an inertial sensor or may be part of an inertial measurement unit (IMU). The IMU may include one or more accelerometers, one or more gyroscopes, one or more magnetometers, or a suitable combination thereof. For example, an IMU can measure up to three orthogonal accelerometers that measure linear acceleration of a movable object along up to three translational axes and up to three orthogonal gyroscopes that measure angular acceleration around up to three rotational axes. Scope. The IMU may be firmly connected to the aircraft such that the aircraft motion corresponds to the IMU motion. Alternatively, the IMU can move relative to the aircraft for up to 6 degrees of freedom. The IMU can be attached directly to the aircraft or can be connected to a support structure attached to the aircraft. The IMU can be provided outside or within the housing of the movable object. The IMU can be permanently or detachably attached to the movable object. An IMU is an aircraft motion, such as at least one of aircraft position, direction, velocity, or acceleration (eg, with respect to at least one of 1, 2, or 3 axes of translation, or 1, 2, or 3 axes of rotation). ) Can be provided. For example, an IMU can detect a signal representing aircraft acceleration, integrate the signal once to provide speed information, and integrate twice to provide at least one of location information or direction information. . The IMU may be able to determine at least one of aircraft acceleration, speed, or location / direction without interacting with any external environmental factors or receiving any signal from outside the aircraft .

  The IMU can provide signals to the flight controller 520, which can be useful for generating command signals to one or more motors 560 of the aircraft. In some embodiments, the flight controller may use a control feedback scheme that may utilize information from the IMU.

  Other sensors can be utilized to determine at least one of aircraft attitude, angular velocity, or angular acceleration. Other sensors may or may not be on the aircraft. Other sensors may or may not communicate with additional devices external to the aircraft. For example, the sensor may be able to determine information without receiving any external signals from the aircraft.

The external device 550 may also communicate with the flight controller. External devices can be provided outside the aircraft. The external device may be able to communicate with the aircraft wirelessly. The external device may have any information such as navigation or position information related to the aircraft. In some embodiments, the external device may include aircraft configuration parameter data. Also, external devices, the memory for storing the aircraft configuration parameter data for the aircraft may communicate aircraft configuration parameter data. Data on the aircraft can be updated with new data from external devices. The update can be done automatically or upon request from the user or from the aircraft.

  In certain embodiments, the external device may be a flight control device that can provide one or more flight instructions to the flight controller. For example, the user may use a remote controller that can communicate wirelessly with the aircraft. The user can specify different flight instructions, such as programming a predetermined route or providing instructions in real time. The flight instructions may include information about the target attitude of the aircraft about one or more axes. For example, the flight command can be a command to adjust the attitude of the aircraft to a certain degree. The instructions may also include or not include information about at least one of the target angular velocity or target angular acceleration of the aircraft.

  The flight controller 520 may generate command signals to one or more actuators 560a, 560b of the aircraft, thereby causing the operation of the propulsion unit and controlling flight of the aircraft. This may include attitude control of the aircraft around three orthogonal axes (eg, pitch, yaw, and roll). The flight controller can calculate a command signal based on one or more aircraft configuration parameters 530. Aircraft configuration parameters 530 are derived from feedback on aircraft physical characteristics, aircraft attitude (eg, information about at least one of aircraft attitude, angular velocity, or angular acceleration about three orthogonal axes). Or may represent them. Also, aircraft configuration parameters 530 may be derived from or represent one or more flight instructions from flight control device 550 that may be external to the aircraft. The flight controller may control the attitude of the aircraft using feedback control incorporating flight configuration parameters.

  FIG. 6A is a diagram illustrating an attitude control method that can be implemented by an aircraft according to an embodiment of the present invention. An attitude control scheme may be used to control the attitude of the aircraft about one, two, or three axes. For example, the attitude control method can be used to control the attitude of the aircraft around the pitch axis, roll axis, and yaw axis.

  A flight planner 610 may be provided to generate command signals that determine flight of the aircraft. The flight planner may be provided on the aircraft or may be provided outside the aircraft to communicate with the aircraft. The flight planner can include one or more memory units and one or more processors that can individually or collectively perform one or more of the steps described herein. The memory may include non-transitory computer readable media that may include code, logic, or instructions for performing one or more steps as described herein. One or more processors may perform one or more steps in accordance with a non-transitory computer readable medium.

  In accordance with embodiments of the present invention, a remote controller 605 or other type of flight control device may be provided. A remote controller can be operated by the user to control the flight of the aircraft. This may include the location of the aircraft and the angular orientation of the aircraft. In some embodiments, the user can directly enter aircraft flight instructions in real time. For example, the user can input a command to adjust the attitude of the aircraft. Alternatively, the user may give instructions to the aircraft to follow a predetermined or pre-programmed route. Also, the remote controller can be separated from the aircraft and can communicate with the aircraft via a wireless connection. In other examples, the flight control device may be incorporated into an aircraft and any description herein for a remote controller is applicable to a flight control device that is part of an aircraft. For example, a user may board an aircraft and give instructions for flight using an on-board flight control device.

  The remote controller 605 can provide a signal to the planner 610 indicating one or more target attitudes θ_Tar. The target attitude may be the target attitude of the aircraft about one, two, or three axes of rotation. For example, the target attitude may indicate the attitude of the aircraft around the pitch, roll, and yaw axes. The planner may attempt to achieve the target attitude by calculating one or more signals to provide to the aircraft motor.

  Planner 610 may also receive information about aircraft dynamics 650. In certain embodiments, information about aircraft dynamics may be provided by one or more sensors. In certain embodiments, information about aircraft dynamics may be provided from one or more inertial sensors (eg, one or more gyroscopes or accelerometers) from within the aircraft. Information about aircraft dynamics may include at least one of attitude, angular velocity, or angular acceleration about one, two, or three of the axes: pitch axis, roll axis, and yaw axis. . In some embodiments, the current attitude θ_Cur of the aircraft can be communicated to the planner. The planner can compare the target posture θ_Tar with the current posture θ_Cur. This comparison can be made for each of the pitch, roll, and yaw axes. It can be determined that the angle difference is an attitude error θ_Err.

  Although only the pitch control 620a is shown in detail, the same control scheme can be applied to the roll control 620b and the yaw control 620c. Any consideration of pitch control or any angle control in general can be applied to any or all of these axes. Any description of at least one of attitude, angular velocity, or angular acceleration is applicable to any or all of these axes. The three axes may be separated from each other.

Attitude error θ_Err can be used with fuzzy logic 621 to control aircraft angle. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control may be fuzzy proportional integral derivative (PID) control. In addition, the target attitude is, proportional-integral (PI) 622 or PID system as possible out of control. Angle control loop may be provided.

  A target angular velocity ω_Tar can occur. The target angular velocity can be compared with the measured angular velocity ω623. The measured angular velocity can be part of aircraft dynamics 650 that can be measured by one or more sensors. Comparing the target angular velocity with the measured angular velocity, the angular velocity error ω_Err can be determined.

The angular velocity error ω_Err may or may not be used with fuzzy logic to control the angular velocity of the aircraft. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control can be a proportional integral derivative (PID) control. Also, the target angular velocity can be proportional (P) control 624. Angle velocity loop may be provided.

  A target angular acceleration α_Tar can occur. The target angular acceleration can be compared with the measured angular acceleration α625. The measured angular acceleration can be part of aircraft dynamics 650 that can be measured by one or more sensors. The angular acceleration error α_Err can be determined by comparing the target angular acceleration with the measured angular acceleration.

The angular acceleration error α_Err may or may not be used with fuzzy logic to control the angular velocity of the aircraft. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control can be a proportional integral derivative (PID) control. In addition, the target angular acceleration is proportional-integral (PI) 626 or PID system as possible out of control. Angular acceleration loop may be provided.

  A feed forward loop 627 can also be provided. A feedforward loop may be provided for angular acceleration. For example, the target angular acceleration α_Tar can be used in a feedforward loop. In certain embodiments, one or more aircraft configuration parameters 660 that may be derived from one or more physical characteristics of the aircraft may be incorporated into the feed forward loop. For example, the moment of inertia I of the aircraft can be provided to the feedforward loop. In one embodiment, the aircraft torque τ can be calculated by multiplying the target angular acceleration α_Tar by the moment of inertia I. Therefore, the angular acceleration loop can directly calculate the output using the moment of inertia, and at the same time, can perform PID control as a compensation amount according to the current angular acceleration value.

  Therefore, both feedforward and feedback can be used to control angular acceleration. Feedforward model parameters can improve the response time of the control system, while feedback control can compensate for model errors and dynamic disturbances. Angular velocity control can be directly regarded as rotational torque control for the entire aircraft. Therefore, the response time to external disturbance can be further shortened and the suppressing action is better than a system that does not use the angular velocity control method. The feed forward loop allows the angular acceleration loop to operate as a direct control, so response time can be shortened. Disturbances can be directly suppressed, thereby reducing response time.

A mixer 630 can be provided in accordance with an embodiment of the present invention. The mixer may be part of a flight controller according to an embodiment of the present invention. The mixer may include one or more processors that may or may not be the same processor used for flight planner 610. The mixer can receive information related to attitude for at least one of pitch control, roll control, or yaw control. For example, data after the feedforward and feedback loops associated with the angular acceleration can provide pitch axis, yaw axis, the mixer for each of the axes of rotation of the roll shaft. All of the calculation results can be summed.

Mixer 630 can receive information regarding aircraft configuration parameters 660. Aircraft configuration parameters may be derived from the physical characteristics of the aircraft or aircraft model. In one embodiment, the mixer can receive an aircraft axial distance l . The axial distance can be the distance between the aircraft motor and the aerodynamic center. Alternatively, the axial distance may be the distance between the aircraft propulsion unit and the aerodynamic center. The aircraft may also have a plurality of motors 640a, 640b, 640c, 640d. The axial distance can be the same for each of the motors. Alternatively, different motors may have different axial distances. Further, the axial distance may be a distance between an axis penetrating at least one of the propulsion unit or the rotor blade and an aerodynamic center in a direction of thrust generated by the propulsion unit.

The mixer 630 can calculate the desired force exerted by each propulsion unit. Force, aircraft configuration parameters 660, and the pitch axis, yaw axis can be calculated based on information from the control system for each axis of rotation of the roll shaft. The torque of each motor may be calculated using the torque τ calculated in the feed forward loop 627 and the axial distance l . The force F can be calculated as τ / l . The desired additional force exerted on each motor can be calculated by the mixer. The desired further force can be transmitted as a command signal to each motor. The force on each motor may be the same or different. For example, for the first motor M1640a, the desired force may be ΔF1, and for the second motor M2640b, the desired force may be ΔF2, and the third motor M3640c it is against, for desired force may be ΔF3 fourth motor M4640d, the desired force may be Derutaefu4. In these examples, at least one or all may be such. The motor can operate at a level that produces the desired or approximate desired force. In some embodiments, a motor lift curve 670 can be used in determining the motor output. The curve may include lift generated per rate of motor operation. The motor lift curve may be the thrust per operating rate. The curve may show a non-linear relationship. The motor lift curve may be an aircraft configuration parameter that may be derived from one or more physical characteristics of the aircraft. One or more of the aircraft configuration parameters derived from the one or more physical parameters may be non-linear parameters.

  The output from motors 640a, 640b, 640c, 640d can be used to drive one or more propulsion units of the aircraft. This may determine at least one of aircraft positioning, speed, or acceleration. The output from the motor can affect at least one of aircraft attitude, angular velocity, or angular acceleration. Any number of motors or propulsion units can be provided. The command signals generated to determine the output for each motor can be determined individually to direct the aircraft to a target attitude from a remote controller.

  The dynamic 650 system can record location information related to the aircraft. For example, one or more inertial sensors can determine at least one of aircraft attitude, angular velocity, or angular acceleration and provide that information back to the control system. In some embodiments, the output to the motor or the output from the measured motor may be used to calculate at least one of aircraft attitude, angular velocity, or angular acceleration. Other in-flight or out-of-flight sensors may be used to determine aircraft dynamics.

  Any flight control steps may be performed using software that may be provided on the aircraft. The software can incorporate or accept values for aircraft configuration parameters. Aircraft configuration parameters may include or be derived from one or more physical characteristics of an aircraft or aircraft model. Thus, flight control software can be specific to the aircraft or aircraft model and can provide more accurate control.

  FIG. 6B is a diagram illustrating an attitude control method that can be implemented by an aircraft according to an embodiment of the present invention. The attitude control scheme may have one or more features or characteristics of the attitude control scheme illustrated in FIG. 6A.

The goal may be to track any posture command.

Constituent manifolds can be non-linear, which can create challenges.

A flight planner (se (3)) 610b can be provided to generate a command signal that determines the flight of the aircraft. The flight planner may be provided on the aircraft or may be provided outside the aircraft to communicate with the aircraft. In some embodiments, the exponential coordinates of SO (3) can be provided in a linear space. This may be optimal or preferred from a geodetic point of view. Further, at least one of a singular point (Euler angle) or ambiguity (quaternion, 2-> 1) may not exist.

A remote controller 605b or other type of flight control device may be provided in accordance with embodiments of the present invention. A remote controller can be operated by the user to control the flight of the aircraft. This may include the location of the aircraft and the angular orientation of the aircraft. In some embodiments, the user can directly enter aircraft flight instructions in real time. For example, the user can provide input to adjust the attitude of the aircraft. In other embodiments, the user may give instructions to the aircraft to follow a predetermined or pre-programmed route. Also, the remote controller can be separated from the aircraft and can communicate with the aircraft via a wireless connection. Also, the flight control device may be incorporated into an aircraft, and any description herein for a remote controller can also be applied to a flight control device that is part of an aircraft. For example, a user may board an aircraft and provide instructions for flight through an onboard flight control device.

  The remote controller 605b can provide a signal to the planner 610b indicating one or more target attitudes θ_Tar. The target attitude may be the target attitude of the aircraft about one, two, or three axes of rotation. For example, the target attitude may indicate the attitude of the aircraft around the pitch, roll, and yaw axes. The planner may attempt to achieve the target attitude by calculating one or more signals to provide to the aircraft motor.

  Planner 610b may also receive information about aircraft dynamics 650b. In certain embodiments, information about aircraft dynamics may be provided by one or more sensors. In certain embodiments, information about aircraft dynamics may be provided from one or more inertial sensors (eg, one or more gyroscopes or accelerometers) from within the aircraft. Information about aircraft dynamics may include at least one of attitude, angular velocity, or angular acceleration about one, two, or three of the axes: pitch axis, roll axis, and yaw axis. . In addition, the current attitude θ_Cur of the aircraft can be transmitted to the planner. The planner can compare the target posture θ_Tar with the current posture θ_Cur. This comparison can be made for each of the pitch, roll, and yaw axes. It can be determined that the angle difference is an attitude error θ_Err.

  Although only the pitch control 620d is shown in detail, the same control method can be applied to the roll control 620e and the yaw control 620f. Any consideration of pitch control or any angle control in general can be applied to any or all of these axes. Any description of at least one of attitude, angular velocity, or angular acceleration is applicable to any or all of these axes. The three axes may be separated from each other.

Attitude error θ_Err can be used with fuzzy logic 621b to control the angle of the aircraft. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control may be fuzzy proportional integral derivative (PID) control. In addition, the target attitude is, proportional (P) 622b or PID system as possible out of control. Angle control loop may be provided.

  A target angular velocity ω_Tar can occur. The target angular velocity can be compared with the measured angular velocity ω623b. The measured angular velocity can be part of aircraft dynamics 650b that can be measured by one or more sensors. Comparing the target angular velocity with the measured angular velocity, the angular velocity error ω_Err can be determined.

The angular velocity error ω_Err may or may not be used with fuzzy logic to control the angular velocity of the aircraft. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control can be a proportional integral derivative (PID) control. Further, the target angular velocity can be proportionally differentiated (PD) controlled 624b. Angular velocity loop may be provided.

  A target angular velocity change ω_Tar can occur. The change in angular velocity can be angular acceleration. The target angular acceleration can be compared with the measured angular velocity change ω 625b. The measured change in angular velocity can be part of aircraft dynamics 650b that can be measured by one or more sensors. The change in the target angular velocity can be compared with the measured change in angular velocity to determine the angular velocity change error ω_Err.

Angular velocity change error ω_Err may or may not be used with fuzzy logic to control aircraft angular velocity changes. The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control can be a proportional integral derivative (PID) control. In addition, a change in the target of the angular velocity is proportional-integral (PI) 626b or PID system as possible out of control. Angular acceleration loop may be provided.

A feed forward loop 627b can also be provided. The feed forward loop 627b may be provided for changes in angular velocity. For example, the target angular velocity change ω_Tar can be used in the feed forward loop 627b . In certain embodiments, one or more aircraft configuration parameters 660b that may be derived from one or more physical characteristics of the aircraft can be incorporated into the feedforward loop 627b . For example, the moment of inertia J of the aircraft can be provided to the feed forward loop 627b . The aircraft torque τ can be calculated as Jω + ω ^b × Jω ^b . Therefore, the loop for the change in angular velocity can calculate the output using the moment of inertia, and at the same time, can perform PID control as a compensation amount according to the current change in angular velocity value and the current angular acceleration value.

  Thus, both feedforward and feedback can be used to control the change in angular velocity (which can also be angular acceleration). Feedforward model parameters can improve the response time of the control system, while feedback control can compensate for model errors and dynamic disturbances. Since angular velocity control can be directly regarded as rotational torque control of the entire aircraft, the response time to external disturbance can be further shortened and the suppression action is better than a system that does not use this control method. The feed forward loop allows the angular acceleration loop to operate as a direct control, so response time can be shortened. Disturbances can be directly suppressed, thereby reducing response time.

The mixer 6 3 0b, can be provided in accordance with an embodiment of the present invention. The mixer may be part of a flight controller according to an embodiment of the present invention. The mixer may include one or more processors that may or may not be the same processor used for flight planner 610b. The mixer can receive information related to attitude for at least one of pitch control, roll control, or yaw control. For example, data after the feedforward and feedback loops associated with a change of the angular velocity can provide pitch axis, yaw axis, the mixer for each of the axes of rotation of the roll shaft. All of the calculation results can be summed.

Mixer 6 3 0b can receive information about the aircraft configuration parameters 660b. Aircraft configuration parameters may be derived from the physical characteristics of the aircraft or aircraft model. In one embodiment, the mixer can receive an aircraft axial distance l . The axial distance can be the distance between the aircraft motor and the aerodynamic center. Alternatively, the axial distance may be the distance between the aircraft propulsion unit and the aerodynamic center. The aircraft may also have a plurality of motors 641a, 641b, 641c, 641d. The axial distance can be the same for each of the motors. Alternatively, different motors may have different axial distances. Further, the axial distance may be a distance between an axis penetrating at least one of the propulsion unit or the rotor blade and an aerodynamic center in a direction of thrust generated by the propulsion unit.

Mixer 6 3 0b can calculate the desired force the propulsion unit on. Force, aircraft configuration parameters 660b, and pitch axis, yaw axis can be calculated based on information from the control system for each axis of rotation of the roll shaft. The torque of each motor may be calculated using the torque τ calculated in the feedforward loop 627b and the axial distance l . The force F can be calculated as τ / l . The desired additional force exerted on each motor can be calculated by the mixer. The desired further force can be transmitted as a command signal to each motor. The force on each motor may be the same or different. For example, for the first motor M1641a, the desired force may be ΔF1, and for the second motor M2641b, the desired force may be ΔF2, and the third motor M3641c On the other hand, the desired force may be ΔF3, or alternatively, for the fourth motor M4641d, the desired force may be ΔF4. The motor can operate at a level that produces the desired or approximate desired force. In some embodiments, the motor lift curve 670b can be used in determining the motor output. The curve may include lift generated per rate of motor operation. The motor lift curve may be the thrust per operating rate. The curve may show a non-linear relationship. The motor lift curve may be an aircraft configuration parameter that may be derived from one or more physical characteristics of the aircraft. One or more of the aircraft configuration parameters derived from the one or more physical parameters may be non-linear parameters.

  The output from the motors 641a, 641b, 641c, 641d can be used to drive one or more propulsion units of the aircraft. This may determine at least one of aircraft positioning, speed, or acceleration. The output from the motor can affect at least one of aircraft attitude, angular velocity, or angular acceleration. Any number of motors or propulsion units can be provided. The command signals generated to determine the output for each motor can be determined individually to direct the aircraft to a target attitude from a remote controller.

  The dynamic 650b system can record position information related to the aircraft. For example, one or more inertial sensors can determine at least one of aircraft attitude, angular velocity, or angular acceleration and provide that information back to the control system. In some embodiments, the output to the motor or the output from the measured motor may be used to calculate at least one of aircraft attitude, angular velocity, or angular acceleration. Other in-flight or out-of-flight sensors may be used to determine aircraft dynamics.

In some embodiments, the kinematics of the system can be considered. A primary fully operational system may be provided.

Where




Proportional control can be applied.

Geometric control may be provided for SO (3).
For regulations:


For tracking:

Where

g is the current direction,
g _d is the desired direction,
k _p is the gain of the controller.

The kinetics may be

Where










Linear control can be provided as already implemented in the feedforward loop of the control scheme. This can include angular velocity control.


Angular velocity control may include:

FIG. 6C is a diagram illustrating a portion of a control inner loop, according to an embodiment of the present invention. System identification of the actuator can be performed. The actuator may include a propeller, rotor, motor, or other type of actuator. The actuator model for a primary system with time delay may be:

  FIG. 6C shows an example of a Smith predictor that can be used in the control scheme. The Smith predictor can be part of the inner loop control. The Smith predictor can predict and correct aircraft dynamics. The Smith predictor can be a predictive controller that can be used in a system with a pure time delay.

  FIG. 6D is a diagram illustrating an attitude control method according to the embodiment of the present invention. A step signal (eg, step) can be provided as an input. A DC signal can be used. The upper part of the attitude control scheme shown may be a conventional controller design, while the lower part may be a proposed control design that uses cascade control with a predictor.

  A portion of the control scheme may include one or more PID controllers (eg, PID controller 4, PID controller 1). These can include an inner PID loop and an outer PID loop. Any number of PID loops can be provided (eg, 1, 2, 3, 4, 5, or more loops). In some embodiments, one or more of the PID loops can be nested within each other. A switch (eg, a switch) can be provided. The switch may change position when signal saturation occurs. The resulting signal can receive a system transfer function (eg, system transfer function 1) and interact with one or more integrators (eg, integrator, integrator 4). In the feedback process, there may be a transport delay (eg, transport delay 2, transport delay 9).

  A portion of the control scheme may be a controller used to control the angle based on at least one of the aircraft angle or angular velocity.

  Another part of the control scheme may include one or more PID controllers (eg, PID controller 2, PID controller 3, PID controller 5). This may include at least one of one or more inner PID loops or outer PID loops. Any number of PID loops can be provided to be nested within each other. A switch (eg, switch 3) can be provided. The switch may change position when signal saturation occurs. The resulting signal can receive a system transfer function (eg, system transfer function 2) and interact with one or more integrators (eg, integrator 2, integrator 5). In the feedback process, there may be a transport delay (eg, transport delay 5, transport delay 6, transport delay 8, transport delay 10). In feedback, there can be a model transfer function (eg, model transfer function 1, model transfer function 2, model transfer function 3) and associated switches (eg, switch 4, switch 5, switch).

  Another part of the control scheme may be a controller that can control the aircraft angle based on at least one of the aircraft angle, angular velocity, or angular acceleration. In addition, the Smith predictor can be included as part of the control scheme.

In some embodiments, the output may be provided to the scope. The result of using an improved control scheme by incorporating these additional features can be shown in FIG.

  FIG. 7A is a diagram illustrating a tracking error according to the embodiment of the present invention. This can be incorporated as part of attitude control, as described above. An error can be obtained by comparing the target with a feedback value. This may be done for at least one of angle, angular velocity, or angular acceleration. This may be done on one or more axes such as a pitch axis, roll axis, or yaw axis.

  The error may or may not be used with fuzzy logic to control aircraft attitude characteristics (eg, angle, angular velocity, or angular acceleration). The control can be feedback control. In some embodiments, feedback control may use at least one of proportional, integral, or derivative control schemes. The feedback control can be a proportional integral derivative (PID) control. The target posture feature can be proportionally (P) controlled 710. The attitude feature error 720 can be provided to the controller 710. Fuzzy logic 730 may be used to control aircraft attitude characteristics (eg, angle, angular velocity, angular acceleration).

  Further, a feedback value may be obtained. For example, errors in controlled pose features. An error value of P (error) 740 can be obtained. In some embodiments, fuzzy logic 750 may be used in determining the error 740 value. The error value can be determined using one or more measured aircraft dynamics. For example, aircraft attitude characteristics can be measured using one or more sensors. The gain of the controller can be adjusted dynamically by the fuzzy logic engine. This method is advantageous in that the response speed can be improved when the error is large. At the same time, this method can improve stability when the error is small.

  FIG. 7B is a further diagram illustrating the tracking error according to an embodiment of the present invention. Show example membership degrees. Furthermore, a proportional gain (kp) as a function of error can be calculated and / or determined. The proportional gain kp can be a non-linear function of the error. This nonlinear function can be used for aircraft attitude control. The nonlinear proportional gain function can be used for error tracking in aircraft control.

  FIG. 8 shows a comparison between the response of a controller to which the present invention is applied compared to a conventional controller. With this controller, it is possible to obtain faster response and smaller overshoot than conventional controllers. Moreover, a shorter time can be used to settle to a steady state.

  A target angle can be indicated. For example, it may be desirable to reach a target angle having a specific degree value at a specific time. For example, in 0.5 unit time, the command can be provided to change the target angle to 40 (eg, 40 degrees). It provides a response to the proposed controller and the conventional controller described herein. The proposed controller may implement the control scheme described elsewhere in this specification. The proposed controller can take into account the physical characteristics of the aircraft. The proposed controller can use feedforward and feedback loops for acceleration. In some embodiments, conventional controllers do not take into account the physical characteristics of the aircraft. Conventional controllers do not calculate the moment of inertia of the aircraft and do not incorporate the moment of inertia into the control scheme. Conventional controllers may also not include feedforward and feedback loops for acceleration.

  The response of the proposed controller may be faster than the response of a conventional controller, as shown. In certain embodiments, the response of the proposed controller may be about twice as fast as the response of a conventional controller. For example, with the proposed controller, the aircraft can reach the target angle approximately twice as fast as an aircraft using a conventional controller. The proposed controller is 1.1 times faster, 1.2 times faster, 1.3 times faster, 1.5 times faster, 2 times faster, 2.5 times faster than the conventional controller. The target angle may be reached twice as fast, 3.5 times faster, 4 times faster, 5 times faster, 6 times faster, 7 times faster, or 10 times faster. Thus, the proposed control scheme described herein allows the aircraft to respond more quickly to reach the target angle.

  The proposed controller may have weaker vibrations than conventional controllers. The proposed controller may have little or no vibration. Vibration can refer to variations in the attitude of the aircraft around a target angle. For example, when the aircraft is approaching the target angle, there may be at least one of some overshot or overcompensation that causes some variation before the aircraft reaches the target angle and stabilizes there.

  The methods and systems in accordance with the present invention can provide improved aircraft attitude control about one, two, or three axes of rotation. Simplified parameter adjustment can be provided. When evaluating aircraft performance parameters, the flight controller can directly perform parameter evaluation, thereby significantly reducing parameter adjustment time compared to conventional systems. Adaptable to changes and variations in aircraft dimensions and weight. The flight control parameter adjustment can be completed easily and reliably by adjusting the basic parameters of the aircraft. Thus, the control system can take into account different models of aircraft with different physical characteristics, or physical changes that may occur in existing aircraft.

  In addition, the angular acceleration loop described herein can enhance dynamic tracking performance and disturbance tolerance. Since the angular acceleration loop control can operate as a direct control, the response time can be shortened and strong anti-disturbance characteristics can be obtained as compared with the conventional control system. For example, conventional systems use angular velocity loop control, and control may be delayed when the aircraft has not yet generated a rotational speed. By using an angular acceleration loop as described, disturbance can be directly suppressed and response time can be reduced.

The system, apparatus and method according to the present invention can be applied to a wide range of movable objects. As described above, the description relating to the aircraft can be applied to and used with any movable object. Any description herein for aircraft is particularly applicable to UAVs. The movable object in the present invention may be in the air (eg, fixed wing aircraft, rotary wing aircraft, or aircraft having no fixed wing or rotary wing), underwater (eg, a ship or submarine), or on the ground (eg, a vehicle). , Trucks, buses, vans, motorcycles, bicycles, etc., movable structures or frames such as walking sticks, fishing rods, or trains), underground (eg subways), in space (eg spaceplanes, satellites, or space probes) ) In any suitable environment, or in any combination of these environments. The movable object may be a vehicle such as a vehicle described elsewhere herein. The movable object can be carried using a living body such as a human being or an animal, and can start from the living body. In that case, a suitable animal is, Ru include birds, canine, feline, equine, bovine animals, sheep, pigs, dolphins, rodents, or insects.

  The movable material may be freely movable in an environment with 6 degrees of freedom (eg, 3 degrees of freedom for translation and 3 degrees of freedom for rotation). On the other hand, movement of a movable object may be constrained with respect to one or more degrees of freedom by a predetermined path, trajectory, position, or the like. Movement can be achieved by any suitable actuation mechanism such as an engine or motor. The actuation mechanism of the movable object may be powered by any suitable energy source, such as electrical energy, magnetic energy, solar energy, wind energy, gravity energy, chemical energy, nuclear energy, or any suitable combination thereof. The movable object may be self-propelled using a propulsion system, as described elsewhere herein. The propulsion system may also be implemented with electrical energy, magnetic energy, solar energy, wind energy, gravity energy, chemical energy, nuclear energy, or any suitable combination thereof. Movable objects can also be carried by living beings.

  In some cases, the movable object may be a vehicle. Suitable vehicles may include water vehicles, aircraft, spacecraft, or ground vehicles. For example, an aircraft may be a fixed wing aircraft (eg, airplane, glider), a rotary wing aircraft (eg, helicopter, rotary wing aircraft), an aircraft with both fixed and rotary wings, or an aircraft without both (eg, small Airship, hot air balloon). The aircraft can be self-propelled, such as self-propelled through air. A self-propelled vehicle can utilize a propulsion system, such as a propulsion system that includes one or more engines, motors, wheels, axles, magnets, rotor blades, propellers, vanes, nozzles, or any suitable combination thereof. The propulsion system can also be used to allow the movable object to take off from the surface, land on the surface, maintain at least one of its current position or orientation (eg, hover), or change at least one of the orientation or position.

  The movable object can be remotely controlled by the user, or locally controlled by an occupant in or on the movable object. In some embodiments, the movable object is an unmanned movable object such as a UAV. An unmanned movable object such as a UAV may not have a passenger on the movable object. The movable object can be controlled by a human or by an autonomous control system (eg, a computer control system), or any suitable combination thereof. The movable object may be an autonomous robot or a semi-autonomous robot such as a robot configured with artificial intelligence.

  The movable object may have any suitable size or dimension. In certain embodiments, the movable object may have a size or dimension for providing a human occupant within or on the vehicle. Also, the movable object may be at least one of a size or dimension that is smaller than at least one of the sizes or dimensions that can comprise a human occupant within or on the movable object. The movable object can be at least one of a size or dimension suitable for being lifted or carried by a human. The movable object may also be larger than at least one of a size or dimension suitable for being lifted or carried by a human. In certain embodiments, the movable object may have a maximum dimension (eg, length, width, height, diameter, diagonal) of about 2 cm, 5 cm, 10 cm, 50 cm, 1 m, 2 m, 5 m, or 10 m or less. . The maximum dimension of the movable object may be about 2 cm, 5 cm, 10 cm, 50 cm, 1 m, 2 m, 5 m, or 10 m or more. For example, the distance between the axes of the rotor blades facing the movable object may be about 2 cm, 5 cm, 10 cm, 50 cm, 1 m, 2 m, 5 m, or 10 m or less. Alternatively, the distance between the axes of the opposing rotor blades may be about 2 cm, 5 cm, 10 cm, 50 cm, 1 m, 2 m, 5 m, or 10 m or more.

In certain embodiments, the movable object may have a volume of less than 100 cm × 100 cm × 100 cm, 50 cm × 50 cm × 30 cm, or 5 cm × 5 cm × 3 cm. The total volume of the movable object is about 1 cm ³ , 2 cm ³ , 5 cm ³ , 10 cm ³ , 20 cm ³ , 30 cm ³ , 40 cm ³ , 50 cm ³ , 60 cm ³ , 70 cm ³ , 80 cm ³ , 90 cm ³ , 100 cm ³ , 150 cm ³ , ^{200cm 3, 300cm 3, 500cm 3} , 750cm 3, 1000cm 3, 5000cm 3, 10,000cm 3, 100,000cm 3, 1m 3, or may be 10 m ³ or less. Conversely, the total volume of the movable object is about 1 cm ³ , 2 cm ³ , 5 cm ³ , 10 cm ³ , 20 cm ³ , 30 cm ³ , 40 cm ³ , 50 cm ³ , 60 cm ³ , 70 cm ³ , 80 cm ³ , 90 cm ³ , 100 cm ³ , ^{150cm 3, 200cm 3, 300cm 3} , 500cm 3, 750cm 3, 1000cm 3, 5000cm 3, 10,000cm 3, 100,000cm 3, 1m 3, or can also be a 10 m ³ or more.

In certain embodiments, the movable ^{object, 32,000cm 2, 20,000cm 2, 10,000cm} 2, 1,000cm 2, 500cm 2, 100cm 2, 50cm 2, 10cm 2 , or 5 cm ² or less of footprint, (or May have a lateral cross-sectional area surrounded by a movable object. Conversely, footprint is approximately ^{32,000cm 2, 20,000cm 2, 10,000cm 2} , may be ^{1,000cm 2, 500cm 2, 100cm 2} , 50cm 2, 10cm 2, or more of 5 cm ^2.

  In certain embodiments, the weight of the movable object may be 1000 kg or less. The weight of the movable object is about 1000 kg, 750 kg, 500 kg, 200 kg, 150 kg, 100 kg, 80 kg, 70 kg, 60 kg, 50 kg, 45 kg, 40 kg, 35 kg, 30 kg, 25 kg, 20 kg, 12 kg, 10 kg, 9 kg, 8 kg, 7 kg, 6 kg. It can be 5 kg, 4 kg, 3 kg, 2 kg, 1 kg, 0.5 kg, 0.1 kg, 0.05 kg, or 0.01 kg or less. On the contrary, the weight is about 1000 kg, 750 kg, 500 kg, 200 kg, 150 kg, 100 kg, 80 kg, 70 kg, 60 kg, 50 kg, 45 kg, 40 kg, 35 kg, 30 kg, 25 kg, 20 kg, 12 kg, 10 kg, 9 kg, 8 kg, 7 kg, 6 kg, It can be 5 kg, 4 kg, 3 kg, 2 kg, 1 kg, 0.5 kg, 0.1 kg, 0.05 kg, or 0.01 kg or more.

In certain embodiments, the movable object may be smaller than the load carried by the movable object. The load may include at least one of a load or a support mechanism, as will be described in further detail below. The ratio of the movable object weight to the load weight may be greater than, less than, or equal to about 1: 1. Also, the ratio of the support mechanism weight to the load weight may be greater than, less than, or equal to about 1: 1. If desired, the ratio of the moving object weight to the load weight may be 1: 2, 1: 3, 1: 4, 1: 5, 1:10 or less, or even less. Conversely, the ratio of movable object weight to load weight may be 2: 1, 3: 1, 4: 1, 5: 1, 10: 1 or more, or even greater.

  In some embodiments, the energy consumption of the movable object may be low. For example, a movable object may use about 5 W / hour, 4 W / hour, 3 W / hour, 2 W / hour, less than 1 W / hour, or less energy. In other embodiments, the energy consumption of the movable object support mechanism may be low. For example, the support mechanism may use about 5 W / hr, 4 W / hr, 3 W / hr, 2 W / hr, less than 1 W / hr, or less energy. Also, the energy consumption of the movable object load can be as low as about 5 W / hour, 4 W / hour, 3 W / hour, 2 W / hour, less than 1 W / hour, or less.

  In FIG. 9, an unmanned aerial vehicle (UAV) 900 is shown in accordance with an embodiment of the present invention. An unmanned aerial vehicle may be an example of a movable object, as described herein. UAV 900 may include a propulsion system having four rotor blades 902, 904, 906, and 908. Any number of rotor blades may be provided (eg, 1, 2, 3, 4, 5, 6 or more). The drone's rotor, rotor assembly, and other propulsion systems may allow the drone to at least one of hover or maintain position, change orientation, or change location. The distance between the axes of opposing rotor blades can be any suitable length 910. For example, the length 910 can be 2 m or less, or 5 m or less. In certain embodiments, the length 910 can be in the range of 40 cm to 1 m, 10 cm to 2 m, or 5 cm to 5 m. Any unmanned aerial vehicle description herein may be applied to movable objects, such as different types of movable objects, and vice versa. An unmanned aerial vehicle may use the ancillary takeoff system or method described herein.

  In some embodiments, the movable object can be configured to carry a load. A load may include one or more of passengers, cargo, equipment, instruments, and the like. The load may be provided inside the housing. The housing may be separate from the housing of the movable object or may be a part of the housing of the movable object. Conversely, a load may be provided with a housing, in which case the movable object does not have a housing. Also, a part of the load or the entire load can be provided without a housing. The load can be firmly fixed to the movable object. Also, the load can be movable relative to the movable object (eg, translatable or rotatable relative to the movable object). The load can include at least one of a load or a support mechanism, as described elsewhere herein.

  In certain embodiments, the terminal can control the movement of the movable object, the support mechanism, and the load relative to a fixed reference frame (eg, the surrounding environment) and / or both relative to each other. The terminal may be a remote control device at a location remote from at least one of the movable object, the support mechanism, and the load. The terminal may be placed on or attached to a support platform. Alternatively, the terminal may be a handheld or wearable device. For example, the terminal can be a smartphone, tablet, laptop, computer, glasses, gloves, helmet, microphone, or a suitable combination thereof. The terminal may include a user interface such as a keyboard, mouse, joystick, touch screen, or display. Any suitable user input can be used to interact with the terminal, such as manual input commands, voice control, gesture control, or position control (eg, via terminal movement, location, or tilt).

  The terminal can be used to control any suitable state of one or more of the movable object, the support mechanism, and the load. For example, the terminal may be used to control either or both the position and orientation of one or more of the movable object, the support mechanism, and the load with respect to at least one of the fixed reference and each other. In some embodiments, the terminal can be used to control one or more individual elements of the movable object, the support mechanism, the load, such as a support mechanism actuation assembly, a load sensor, or a load emitter. The terminal may include a wireless communication device adapted to communicate with at least one of a movable object, a support mechanism, or a load.

  The terminal may comprise a suitable display unit for viewing one or more information of a movable object, a support mechanism, a load. For example, the terminal can display one or more information of a movable object, a support mechanism, a load regarding position, translation speed, translation acceleration, orientation, angular velocity, angular acceleration, or any suitable combination thereof. In some embodiments, the terminal can display information provided by the load, such as data provided by the functional load (eg, an image recorded by a camera or other imaging device).

  In addition, at the same terminal, at least one of a movable object, a support mechanism, and a load, or control of one or more states of the movable object, the support mechanism, and the load, and the movable object, the support mechanism, the load, Both at least one of receiving and displaying information from at least one of them may be performed. For example, the terminal may control the positioning of the load with respect to the environment while displaying image data captured by the load or information regarding the position of the load. Alternatively, different terminals may be used for different functions. For example, the first terminal controls one or more movements or states of the movable object, the support mechanism, and the mounted object, while the second terminal has information on at least one of the movable object, the support mechanism, and the mounted object. Receive, display, or both. For example, a first terminal may be used to control the positioning of the load with respect to the environment, while the second terminal displays image data captured by the load. Various between a movable object and an integrated terminal that controls both the movable object and receives data, or between a movable object and multiple terminals that both control the movable object and receive data A communication mode can be used. For example, at least two different communication modes may be formed between a movable object and a terminal that performs both control of the movable object and data reception from the movable object.

  In FIG. 10, a movable object 1000 comprising a support mechanism 1002 and a load 1004 is shown according to an embodiment of the present invention. Although the movable object 1000 is shown as an aircraft, this figure is not intended to be limiting and any suitable type of movable object described herein may be used. Those skilled in the art will appreciate that any of the embodiments described herein in the context of an aircraft system can be applied to any suitable movable object (eg, UAV). Further, the mounted object 1004 may be provided on the movable object 1000 without using the support mechanism 1002. The movable object 1000 may include a propulsion mechanism 1006, a detection system 1008, and a communication system 1010.

  As described above, the propulsion mechanism 1006 may include at least one of a rotor blade, a propeller, a blade, an engine, a motor, a wheel, an axle, a magnet, or a nozzle. For example, the propulsion mechanism 1006 may be a rotary blade assembly or other rotary propulsion unit as described above. The movable object may have at least one, two or more, three or more, or four or more propulsion mechanisms. The one or more propulsion mechanisms may all be the same type, but may be different types of propulsion mechanisms. The propulsion mechanism 1006 can be attached to the movable object 1000 using any suitable means such as a support element (eg, drive shaft) as described above. The propulsion mechanism 1006 can be attached to any suitable part of the movable object 1000, such as the top, bottom, front, back, side, or any suitable combination thereof.

In some embodiments, the propulsion mechanism 1006 allows the movable object 1000 to take off from or land vertically on the surface without moving in its horizontal direction (e.g., not traveling on the runway). enable. Further, the propulsion mechanism 1006 can hover the movable object 1000 at at least one of a specific position and orientation in the air. At least one propulsion mechanism 100 6, but may be controlled independently of the other propulsion mechanism, it is possible to simultaneously control, respectively. For example, the movable object 1000 may have a plurality of horizontally oriented rotor blades that impart at least one of lift and thrust to the movable object. A plurality of horizontally oriented rotors may be activated to provide the movable object 1000 with vertical takeoff and landing capability and hovering capability. Also, at least one of the horizontally oriented rotor blades can rotate clockwise or counterclockwise. For example, the number of clockwise rotating blades may be equal to the number of counterclockwise rotating blades. When the rotational speed of each rotor blade directed in the horizontal direction is changed independently, at least one of lift and thrust provided by each rotor blade can be controlled. Thereby, at least one of the spatial arrangement, speed, and acceleration of the movable object 1000 is adjusted (for example, with respect to a translation degree of 3 or less and a rotation degree of 3 or less).

  The detection system 1008 can comprise one or more sensors that can detect at least one of the spatial arrangement, velocity, and acceleration of the movable object 1000 (eg, with respect to a translation degree of 3 or less and a degree of rotation of 3 or less). The one or more sensors may include a global positioning system (GPS) sensor, a motion sensor, an inertial sensor, a proximity sensor, or an image sensor. Using the detection data provided by the detection system 1008, the spatial arrangement, speed, and orientation of the movable object 1000 (eg, using at least one of a suitable processing unit and control module, as described below) At least one can be controlled. Alternatively, the detection system 1008 can be used to provide data regarding the surrounding environment of the movable object, such as weather conditions, proximity to potential obstacles, geographical features, locations of artificial structures, and the like.

The communication system 1010 enables communication with a terminal 1012 having a communication system 1014 via a radio signal 1016. The communication systems 1010, 1014 may comprise any number of transmitters, receivers, transceivers suitable for wireless communication. The communication may be one-way communication capable of transmitting data only in one direction. For example, one-way communication may only be performed by the movable object 1000 transmitting data to the terminal 1012 and vice versa. Also, it may send data to one or more receivers from one or more transmitters of the communication system 101 4 of the communication system 1010, and vice versa. Alternatively, the communication may be bidirectional communication in which data can be transmitted in both directions between the movable object 1000 and the terminal 1012. Bidirectional communication may transmit data from one or more transmitters of the communication system 1010 to one or more receivers of the communication system 1014, and vice versa.

In some embodiments, the terminal 1012 provides control data to at least one of the movable object 1000, the support mechanism 1002, and the load 1004. The terminal 1012 receives information from at least one of the movable object 1000, the support mechanism 1002, and the mounted object 1004 (for example, at least one of information on the position and operation of the movable object, the support mechanism, or the mounted object). Data detected by the mounted object such as image data photographed by the object camera). Also, the control data from the terminal may include one or more relative positions, movements, actuations, or control instructions for the movable object, the support mechanism, and the load. For example, one or more corrections of the location and orientation of the movable object may be made by control data (eg, via control of the propulsion mechanism 1006). Further, the mounted object may be moved with respect to the movable object (for example, through the control of the support mechanism 1002). In addition, the mounted data can be controlled by the control data from the terminal, such as control of the operation of the camera or other imaging device (for example, taking a still image or a movie, zooming in or out, turning the power on or off, image Switch modes, change image resolution, change focus, change depth of field, change exposure time, change viewing angle or field of view). Also, communication from at least one of the movable object, the support mechanism, and the load may include information from one or more sensors (eg, detection system 1 0 08 or load 1004). The communication may include detected information from one or more different types of sensors (eg, GPS sensors, motion sensors, inertial sensors, proximity sensors, or image sensors). Such information may relate to one or more positions (eg, location or orientation), movement, or acceleration of the movable object, support mechanism, load. Such information from the load may include data acquired by the load or detected state of the load. One or more states of the movable object 1000, the support mechanism 1002, or the mounted object 1004 can be controlled by the control data transmitted from the terminal 1012. In addition, the support mechanism 1002 and the mounted object 1004 may each include a communication module that communicates with the terminal 1012. The terminal can independently communicate with and control the movable object 1000, the support mechanism 1002, and the mounted object 1004.

  In some embodiments, the movable object 1000 can communicate with another remote device in addition to or in place of the terminal 1012. Terminal 1012 can also communicate with another remote device and movable object 1000. For example, either or both of the movable object 1000 and the terminal 1012 can communicate with another movable object, another movable object support mechanism or a load. If desired, the remote device can be a second terminal or other computing device (eg, a computer, laptop, tablet, smartphone, or other mobile device). The remote device may perform at least one of transmitting data to the movable object 1000, receiving data from the movable object 1000, transmitting data to the terminal 1012, and receiving data from the terminal 1012. The remote device can also be connected to the Internet or other telecommunications network. Therefore, data received from at least one of the movable object 1000 and the terminal 1012 can be uploaded to a website or a server.

FIG. 11 is a schematic block diagram of a system 1100 for controlling a movable object to which the present invention is applied. System 1100 may be used in combination with any suitable embodiment of the systems, devices, and methods disclosed herein. System 1100 can include a detection module 1102, a processing unit 1104, a non-transitory computer readable medium 110 6 , a control module 1108, and a communication module 1110.

The detection module 1102 can utilize different types of sensors that collect information about moving objects in different ways. Different types of sensors may detect different types of signals or signals from different sources. For example, the sensors may include inertial sensors, GPS sensors, proximity sensors (eg, riders), or visual and image sensors (eg, cameras). The detection module 1102 can be operatively connected to a processing unit 1104 having a plurality of processors. In certain embodiments, the detection module can be operatively connected to a transmission module 1112 (eg, a Wi-Fi image transmission module) that transmits detection data directly to a suitable external device or system. For example, an image captured by the camera of the detection module 1102 can be transmitted to the remote terminal using the transmission module 1112.

  The processing unit 1104 may have one or more processors, such as a programmable processor (eg, a central processing unit (CPU)). The processing unit 1104 can be operatively connected to a non-transitory computer readable medium 1106. Non-transitory computer readable media 1106 may store at least one of logic, code, program instructions executable by processing unit 1104 to perform one or more steps. Non-transitory computer readable media may include one or more memory units (eg, removable media such as SD card or random access memory (RAM) or external storage). Also, data from the detection module 1102 can be transmitted and stored directly in a memory unit of the non-transitory computer readable medium 1106. The memory unit of the non-transitory computer readable medium 1106 can store at least one of logic, code, program instructions executable by the processing unit 1104, and performs any suitable embodiment of the methods described herein. it can. For example, the processing unit 1104 can instruct the detection data generated by the detection module to be analyzed by one or more processors of the processing unit 1104. The memory unit can store detection data from the detection module to be processed by the processing unit 1104. Further, the processing result generated by the processing unit 1104 can be stored using the memory unit of the non-transitory computer-readable medium 1106.

  In some embodiments, the processing unit 1104 can be operatively connected to a control module 1108 that controls the state of the movable object. For example, the control module 1108 may control the propulsion mechanism of the movable object to adjust at least one of the spatial arrangement, speed, and acceleration of the movable object for six degrees of freedom. The control module 1108 can also control at least one of the state of the support mechanism, the load, or the detection module.

The processing unit 1104 can be operatively connected to the communication module 1110. The communication module 1110 transmits data to and / or receives data from one or more external devices (eg, a terminal, a display device, or other remote controller). Any suitable communication means such as wired communication or wireless communication can be used. For example, the communication module 1110 can use at least one of a local area network (LAN), wide area network (WAN), infrared, wireless, WiFi, point-to-point (P2P) network, telecommunication network, cloud communication, and the like. A relay station such as a tower, a satellite, or a mobile station may be used. Wireless communication can depend on proximity or be irrelevant. In addition, a line of sight may or may not be necessary for communication. Communication module 1110, processing unit 1104 generates the processing result, a predetermined control data, transmitting at least one of the user instruction or the like from a terminal or a remote controller, receiving, may do one or both.

  The components of system 1100 can be arranged in any suitable configuration. For example, at least one of the components of system 1100 may be located on a movable object, a support mechanism, a load, a terminal, a detection system, or an additional external device that communicates with at least one of the above. Also shown in FIG. 11 is one processing unit 1104 and one non-transitory computer readable medium 1106, which is not intended to limit the embodiments of the present invention. One skilled in the art will appreciate that the system 1100 may comprise at least one of a plurality of processing units and a non-transitory computer readable medium. In some embodiments, at least one of the plurality of processing units and the non-transitory computer readable medium is a movable object, a support mechanism, a load, a terminal, a detection module, an additional external device in communication with at least one of the above, or They can be placed in different places, such as their appropriate combinations. Thus, any suitable embodiment of processing performed by system 1100, or at least one of memory functions, may be performed at at least one of the locations described above.

While preferred embodiments of the present invention have been illustrated and described herein, it will be apparent to those skilled in the art that such embodiments are for illustrative purposes only. Many variations, modifications, and alternatives will occur to those skilled in the art without departing from the invention. It should be understood that various alternatives to the embodiments of the invention described herein may be employed in practicing the present invention. The claims define the scope of the invention, and the methods and structures within the scope of the claims and their equivalents are intended to be included within the scope of the present invention.

Claims (15)

A method for controlling the attitude of an aircraft,
(A) evaluating a non-linear relationship between the thrust of one or more propulsion units of the aircraft and the output of actuators of the aircraft operably connected to the propulsion units;
(B) receiving a signal indicating a target attitude of the aircraft;
(C) generating a command signal delivered to at least one of the actuators using a feedback control scheme based on the signal indicating the target attitude and the nonlinear relationship and including an angular acceleration loop having angular acceleration feedback; Steps,
(D) measuring dynamics of the aircraft due to operation of the one or more propulsion units;
(E) utilizing the dynamics to provide the feedback control scheme for adjusting the command signal.   The method of claim 1, wherein the aircraft is an unmanned aerial vehicle.   The method according to claim 1, wherein the aircraft includes a plurality of actuators connected to a plurality of propulsion units, and the propulsion unit includes a rotor wing that generates lift of the aircraft.   4. A method according to any one of the preceding claims, wherein a signal indicative of the target attitude of the aircraft is received from a remote controller over a wireless connection.   The method according to claim 1, wherein the non-linear relationship is input by a user.   5. A method according to any one of the preceding claims, wherein the non-linear relationship is calculated during calibration of one or more actuators of the aircraft.   The method of any one of claims 1 to 6, further comprising calculating an aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft.   The method of claim 7, further comprising calculating a moment of inertia of the aircraft based on the physical characteristics of the aircraft. The method of claim 8, wherein the calculation using the feedback control scheme comprises a feed forward calculation using the moment of inertia of the aircraft. A method for controlling the attitude of an aircraft,
(A) receiving a signal indicating a target attitude of the aircraft by a processor;
(B) using a command signal delivered to at least one actuator of the aircraft operatively connected to one or more propulsion units of the aircraft based on the signal indicative of the target attitude; and Generating using a feedback control scheme comprising: (1) an angular acceleration loop with angular acceleration feedback; and (2) a direct feedforward calculation based on a target angular acceleration;
(C) measuring the aircraft dynamics resulting from operation of the one or more propulsion units using one or more sensors operably connected to the aircraft;
(D) providing the dynamics to the processor to provide the feedback control scheme for adjusting the command signal.   11. The aircraft of claim 10, wherein the aircraft is an unmanned aerial vehicle and includes a plurality of actuators operably connected to a plurality of propulsion units, the propulsion unit including a rotor blade that generates lift of the aircraft. Method.   The method of claim 10, further comprising calculating an aerodynamic center and center of gravity of the aircraft based on one or more physical characteristics of the aircraft.   The method of claim 12, further comprising calculating a moment of inertia of the aircraft based on the physical characteristics of the aircraft, wherein the feedforward calculation uses the moment of inertia of the aircraft.   The calculation using the feedback control scheme is performed on an aircraft attitude centered on the pitch axis, roll axis, and yaw axis, and using a mixer, for the pitch axis, roll axis, and yaw axis The method of claim 10, further comprising: combining the result of the calculation of and an aircraft configuration parameter to calculate the command signal delivered to the at least one actuator. The method of claim 10, wherein the dynamics of the aircraft include the attitude of the aircraft with respect to at least one axis and the angular acceleration with respect to the at least one axis.