Office Action 3: Notice of Allowance (2025-03-25)

CLAIMS:
 AMENDMENTS TO THE CLAIMSThis listing of claims replaces are prior listings.

1\. (Currently Amended) A control system for an aircraft, the control system comprising:
a pilot input device configured to receive a pilot input;
a first sensor positioned on a landing gear of the aircraft, the first sensor configured to sense a force on the landing gear; and
a controller in communication with the pilot input device and the first sensor, the controller configured to receive the pilot input via the pilot input device,
receive a sensed force on the landing gear via the first sensor,
calculate an output command based at least on the pilot input and the force on the landing gear, wherein the controller calculates the output command by calculating a correction factor, wherein the pilot input and the sensed force on the landing gear are inputs used to calculate the correction factor, and
transmit the output command to control the aircraft,
wherein the controller determines the correction factor using a PID control mode when the aircraft is in a landed state.
2\. (Canceled)
3\. (Canceled)
4\. (Original) The control system of claim 1, wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed force on the landing gear.
5\. (Original) The control system of claim 1, wherein the output command controls a rotor of the aircraft.

6\. (Original) The control system of claim 1, further comprising a second sensor positioned on the landing gear, the second sensor being a weight on wheels sensor.
7\. (Currently Amended) The control system of claim 6, wherein the controller is further configured to determine when the aircraft is in \[\[a\]\]the landed flight state based on a signal received from the second sensor.
8\. (Currently Amended) The control system of claim 1, wherein the first sensor is active when the aircraft is in \[\[a\]\]the landed flight state, and wherein the first sensor is inactive or zeroed-out when the aircraft is in an in-air flight state.
9\. (Original) The control system of claim 1, further comprising a second sensor positioned on a body of the aircraft, wherein the second sensor is configured to sense a parameter associated with an attitude of the aircraft.

10\. (Previously Presented) The control system of claim 9, wherein the controller is configured to calculate the output command based on the pilot input, the force on the landing gear, and the parameter sensed by the second sensor.

11\. (Currently Amended) A method of controlling an aircraft, the method comprising:receiving, by an electronic processor, a pilot input including a setpoint;
receiving, by the electronic processor, a first signal from a first sensor positioned on a landing gear of the aircraft;
calculating, by the electronic processor, a first correction factor wherein the setpoint and the first signal are inputs used to calculate the first correction factor;
determining, by the electronic processor, an output command based on the first correction factor, the output command including instructions for controlling the aircraft; and
transmitting, by the electronic processor, the output command to control the aircraft,
wherein, when the aircraft is in a landed state, determining the output command includes determining the output command using a PID control mode.

12\. (Original) The method of claim 11, wherein receiving the first signal from the first sensor includes receiving at least one of a force on the landing gear and a pressure on the landing gear.

13\. (Canceled)

14\. (Original) The method of claim 11, wherein the output command controls a rotor interface.

15\. (Original) The method of claim 11, wherein determining the output command includes determining the moment of the landing gear.

16\. (Original) The method of claim 11, further comprising receiving, by the electronic processor, a second signal from a second sensor positioned on a body of the aircraft, the second sensor configured to sense a parameter associated with an attitude of the aircraft, wherein the output command is further based on the signal from the second sensor.

17\. (Currently Amended) A control system for an aircraft, the control system comprising:a pilot input device configured to receive a pilot input;
a plurality of sensors, each of the plurality of sensors positioned on a corresponding landing gear of the aircraft and configured to sense a parameter on the corresponding landing gear; and
a controller in communication with the plurality of sensors, the controller configured to calculate an output command based on the pilot input and the sensed parameters of the landing gear, the output command including instructions for controlling a rotor of the aircraft,
wherein the controller calculates the output command by calculating a correction factor wherein the pilot input and the sensed parameters on the landing gear are inputs used to calculate the correction factor, and
wherein, when the aircraft is in a landed state, the controller is configured to determine the correction factor using a PID control mode.

18\. (Canceled)

19\. (Canceled)

20\. (Original) The control system of claim 17, wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed parameters on the landing gear.

21\. (Previously Presented) The control system of claim 3, wherein the correction factor includes a first component and a second component, wherein the first component is calculated using the PID control mode, and wherein the second component is calculated using a PD control mode.

22\. (Previously Presented) The control system of claim 10, wherein the correction factor includes a first component and a second component, wherein the first component is calculated based on the force on the landing gear sensed by the first sensor, and wherein the second component is calculated based on the parameter sensed by the second sensor.

OFFICE ACTION:
EXAMINER’S AMENDMENT
An examiner’s amendment to the record appears below. Should the changes and/or additions be unacceptable to applicant, an amendment may be filed as provided by 37 CFR 1.312. To ensure consideration of such an amendment, it MUST be submitted no later than the payment of the issue fee.
The application has been amended as follows:
1. (Currently Amended) A control system for an aircraft, the control system comprising:
a pilot input device configured to receive a pilot input;
a first sensor positioned on a landing gear of the aircraft, the first sensor configured to sense a force on the landing gear; and
a controller in communication with the pilot input device and the first sensor, the controller configured to
receive the pilot input via the pilot input device,
receive a sensed force on the landing gear via the first sensor,
calculate an output command based at least on the pilot input and the force on the landing gear, wherein the controller calculates the output command by calculating a correction factor, wherein the pilot input and the sensed force on the landing gear are inputs used to calculate the correction factor, and
transmit the output command to control the aircraft,
wherein the controller determines the correction factor using a PID control mode when the aircraft is in a landed flight state.
11. (Currently Amended) A method of controlling an aircraft, the method comprising:
receiving, by an electronic processor, a pilot input including a setpoint;
receiving, by the electronic processor, a first signal from a first sensor positioned on a landing gear of the aircraft;
calculating, by the electronic processor, a first correction factor wherein the setpoint and the first signal are inputs used to calculate the first correction factor;
determining, by the electronic processor, an output command based on the first correction factor, the output command including instructions for controlling the aircraft; and
transmitting, by the electronic processor, the output command to control the aircraft,
wherein, when the aircraft is in a landed flight state, determining the output command includes determining the output command using a PID control mode.
17. (Currently Amended) A control system for an aircraft, the control system comprising:
a pilot input device configured to receive a pilot input;
a plurality of sensors, each of the plurality of sensors positioned on a corresponding landing gear of the aircraft and configured to sense a parameter on the corresponding landing gear; and
a controller in communication with the plurality of sensors, the controller configured to calculate an output command based on the pilot input and the sensed parameters of the landing gear, the output command including instructions for controlling a rotor of the aircraft,
wherein the controller calculates the output command by calculating a correction factor wherein the pilot input and the sensed parameters on the landing gear are inputs used to calculate the correction factor, and
wherein, when the aircraft is in a landed flight state, the controller is configured to determine the correction factor using a PID control mode.
21. (Currently Amended) The control system of claim [[3]]1, wherein the correction factor includes a first component and a second component, wherein the first component is calculated using the PID control mode, and wherein the second component is calculated using a PD control mode.
DETAILED ACTION
Notice of Pre-AIA  or AIA  Status
The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Prosecution History
Claims 1-20 were originally filed 09/29/2021.
Claims 21 and 22 have since been added to the claim set by applicant.
Claims 1, 7-8, 10-11, and 17 have since been amended by applicant.
Claims 2-3, 13, and 18-19 have since been canceled by applicant.
Claims 1, 4-12, 14-17, and 20-22 are pending and allowed.
Allowable Subject Matter
Claims 1, 4-12, 14-17, and 20-22 are allowed over the prior art of record.
The following is an Examiner’s statement of reasons for allowance:
The closest prior art of record is Vallart et al. US 20160327958 (“Vallart”), Romana et al. US 20170008619 (“Romana”), White EP 3208190 (“White”), Waltner et al. US 20180305038 (“Waltner”), Seydoux et al. US 20110049290 (“Seydoux”), and Thomas US 20210271267 (“Thomas”).
Vallart discloses a control system for a rotorcraft that uses inputs provided by a pilot and detected forces applied to landing gear units to determine PID control adjusters to implement control laws for driving the rotors of the rotorcraft when the rotorcraft is in an in flight state or a touching state, however, when the rotorcraft is in a landed state, the control system implements a new control law that is different from those applied by the PID adjusters with a direct or proportional law that may not include an integrating adjuster.
Romana teaches a controller for an aircraft steering system that adjusts a steering command based on undesired moments detected on the nose landing gear.
White teaches a rotor moment control system that uses weight on wheel sensors to detect whether an aircraft is resting on the ground.
Waltner teaches a touchdown control system for an aircraft that includes an attitude sensor positioned on the aircraft.
Seydoux teaches a system for controlling flight of an aircraft using a proportional derivative controller.
Thomas teaches a system for controlling an aerial vehicle that is capable of measuring an attitude orientation of the aerial vehicle and determining compensations/corrections to be forwarded to an attitude controller in order to achieve a desired attitude.
As to independent claims 1, 11, and 17, the prior art of record, taken individually or in combination, fails to teach or suggest the following claimed subject matter:
“wherein the controller determines the correction factor using a PID control mode when the aircraft is in a landed state”
Claims 4-10 and 21-22 depend from claim 1, Claims 12 and 14-16 depend from claim 11, and Claim 20 depends from claim 17, and are therefore allowable.
Any comments considered necessary by Applicant must be submitted no later than the payment of the issue fee and, to avoid processing delays, should preferably accompany the issue fee. Such submissions should be clearly labeled “Comments on Statement of Reasons for Allowance.”
Conclusion

APPLICANT REBUTTAL: REMARKS
By this response, claims 1, 4-12, 14-17, and 20-22 are pending. Claims 1, 7, 8, 11, and 17 are amended. Claims 3, 13, and 19 are canceled. No new subject matter is added.
I.
Rejections Under 35 U.S.C.
102 and 103
In the final Office action, claims 1-3, 5, 11-14, and 17-19 are rejected are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Vallart et al. US 20160327958 ("Vallart"). Claims 4, 15, and 20 are rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of Romana et al. US 20170008619 ("Romana"). Claims 6-8 are rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of White EP 3208190 ("White"). Claims 9-10 and 16 are rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of Waltner et al. US 20180305038 ("Waltner").
In the Advisory Action, the Examiner also noted an additional reference (US 2017/0008619 - Romana) deemed to be relevant to the claims.
a.
Inde
p
endent Claim 1
Amended independent claim 1 recites a control system for an aircraft including, among other things, a controller configured to receive the pilot input via the pilot input device, receive a sensed force on the landing gear via the first sensor, calculate an output command based at least on the pilot input and the force on the landing gear, wherein the controller calculates the output command by calculating a correction factor, wherein the pilot input and the sensed force on the landing gear are inputs used to calculate the correction factor, and transmit the output command to control the aircraft,
wherein the controller determines the correction factor usin
g
a PID control mode when the aircraft is in a landed state.
The Examiner rejects independent claim 1 as being anticipated by Vallart. The Examiner also notes that Romana is relevant to claim 1. Claim 1 is amended to provide further clarity on calculating the correction factor using the pilot input and the sensed force on the landing gear as inputs. Additionally, claim 1 is amended to specify that the controller determines the correction factor using a PID control mode
when the aircraft is in a landed state.
Both Vallart and Romana fail to teach or suggest determining a correction factor using a PID control mode when the
aircraft is in a landed state. Vallart teaches away from using a PID control during the landed state. Romana does not discuss the use of PID controls.
As brief summary, Vallart explains that the control laws generally make use of an
adjuster defined by either a proportional gain or an integral time constant, but that there are many dangers and problems associated with using these types of adjusters if used during the landed state. See Vallart [0015]-[0035. Accordingly, the proportional gain and integral time constant are typically reduced or canceled at the first sign of the rotorcraft making contact with the ground. Vallart [0023]-[0026]. Vallart's goal is to identify a solution for using these types of adjusters during an intermediate state called the "touching state" in which one the rotorcraft is between the in-flight state and the fully landed state. However, once the rotorcraft is in a fully landed state, Vallart no longer uses an adjuster with either proportional gain or an integral time constant. This is because the initial dangers and concerns expressed by Vallart still remain. Therefore, Vallart teaches away from the use of a PID control mode (which uses proportional gains and an integral time constant) during the landed state.
In order to fully appreciate why Vallart teaches away from using a PID control during the landed state, it is important to understand the problem and solution addressed by Vallart. In the background section, Vallart explains that the topography of the landing sites may be extremely varied. Vallart [0009]. Flight crews may face complex phase of landing, particularly between flight and the fully landed state of the rotorcraft. Vallart [0014]. This is typical when personnel, such as physicians or rescue workers, are disembarked on sloping terrain. Vallart [0014]. Landings on sloping terrain are some of the most delicate maneuvers that rotorcraft crews are required to perform. Vallart [0014]. In these situations, the flight crew performs a maneuver on the slopped terrain by positioning the nose of the rotorcraft facing the rising terrain, with the forward landing gear in contact with the ground and with the rear landing gear off the ground. Vallart [0014].
The problem Vallart seeks to address is that the control law used to operate the rotorcraft is dramatically different depending on whether the rotorcraft is in flight or in contact with the ground. This can be problematic when doing complex maneuvers like the one described above, in which the rotorcraft is essentially in a
transition state between flight and fully landed state,
with one landing gear is on the ground and the other two are off the ground. Therefore, Vallart seeks to identify a solution for operating the rotorcraft when it is in a
"touching" state.
Vallart explains that there are significantly different goals and risk associated with using the appropriate control logic during each of the flight states:
In point of fact, the control of a rotorcraft differs depending on whether the rotorcraft is in flight or in contact with the ground. Accordingly, different control laws are traditionally used to control a rotor, depending on whether the rotorcraft is or is not in contact with the ground. If the in-flight control law is retained when the rotorcraft is in contact with the ground, the preservation of the control goal would be negatively affected by this contact with the ground. The use of such a control law when the rotorcraft is in contact with the ground could, for example, cause the tilting of the rotor or even an
amplification of ground resonance. Vallart [001S]
Therefore, it is customary to use a ground control law that is very different from the in-flight control law. Thus, the structure and the benefits of the ground law usually differ greatly from those of the in-flight control law. The in-flight control law favors the maintenance of flight parameters, whereas the goal of the ground control law consists of directly controlling the position of the rotor. Therefore, control of the helicopter must be adapted between these two operating modes, while maintaining a maximum level of assistance and the controllability of the aircraft during the transition phase between on-the-ground and in-flight situations. This adaptation of the control modes assumes the availability of information about the status of the aircraft in relation to the ground. This information is usually referred to collectively as the "ground/flight logic".
Vallart [0016]
Vallart explains that the control laws generally make use of an adjuster defined by either a
proportional
g
ain
or an
integral time constant.
Vallart [0017]. The integral time constant makes it possible to ensure the long-term stability of a parameter, such as attitude or the traveling speed of the rotorcraft. Vallart [0017]. Meanwhile, the proportional gain contributes towards the short-term stability of the parameter. Vallart [0018]. However, Vallart warns that when these types of adjusters are activated on the ground, the proportional gains and the integral time constant can have dangerous effects, such as tilting the helicopter on the ground (i.e., "roll over"). Vallart [0019]. Therefore, Vallart explains that "it is known that in order to avoid these risks, simpler control law can be implemented, and the proportional gains of the control laws can be reduced when contact with the ground is detected." Vallart [0023]. Simplification
of the control law can also be accomplished "by canceling the integral adjustment at the first sign of contact with the ground." Vallart [0025]. Alternatively, the integral adjustment may be retained via an active adjuster during the landing phase, but the integrator performs an integration until the static error becomes null. Vallart [0029].
Vallart sets out to establish a control system in which the integrating adjuster can remain active while the rotorcraft is in the "touching" state so that the pilot can retain full control authority during this flight state. Vallart [0057] (see also, Vallart [0056]-[0060]). However, once the rotorcraft is in the landed state, the integrated adjuster is not in use. Vallart [0065]. Vallart uses sensors to identify which state the rotorcraft is in as well as which control law should be utilized, as follows:
"  The "in flight" state is identified when each of the reaction forces F1, F2, and F3
are simultaneously lower than the low threshold (Ss) value. Vallart [0044]. In
this flight state, the first control law is used. Vallart [0046].
"  The "touching" state is identified when the following two conditions are met: 1)
the reaction force F 1 is greater than the low threshold (Ss) value but lower than
the high threshold (SH); and 2) each of the reaction forces F2 and F3 are lower
than the low threshold (Ss) value. Vallart [0041]-[0043]. In this flight state, the
first control law is used. Vallart [0047].
"  The "landed" state is identified when the force on the landing gear is greater than
the high threshold (SH). Vallart [0065J. In this state, a third/different control law
is used which does not use an integrated adjuster. Vallart [0065].
Significantly, when the landed state of the rotorcraft is identified, "a new control law is implemented with a different structure than the first and second control laws. In particular, such a new law may be a direct or proportional law, and may not included an integrated adjuster." Vallart [0065J. Therefore, Vallart does not use a PID control (i.e., with an integrated adjuster) when in the landed state. This is likely due to the number of dangers and concerns described by Vallart throughout the specification. Accordingly, Vallart teaches away from using a PID control mode during a landed state.
Romana fails to cure the deficiencies of Vallart. Romana does not discuss the use of
adjusters such as a direct, proportional, or integral adjusters. Rather Romana merely discusses the use of differential breaking to correct the direction of travel of the aircraft. Therefore, neither Vallart nor Romana, together or alone, teaches or suggests determining a correction factor using a PID control mode when the aircraft is in a landed state, as recited in claim 1. Therefore, Applicant respectfully contends that claim 1 is allowable for at least these reasons, in addition to other reasons not explicitly discussed herein. Claims 4-10 ultimately depend from claim 1, and therefore, are allowable for at least the same reasons, in addition to independent reasons not explicitly discussed herein.
b.
Inde
p
endent Claim 11
Amended independent claim 11 recites a method of controlling an aircraft including, among other things, receiving a pilot input including a setpoint, receiving a first signal from a first sensor positioned on a landing gear of the aircraft, and calculating a first correction factor, wherein the setpoint and the first signal are inputs used to calculate the first correction factor, wherein,
when the aircraft is in a landed state, the controller is con
fig
ured to determine the correction factor usin
g
a PID control mode.
The Examiner rejects independent claim 11 as being anticipated by Vallart and notes the relevancy of Romana to claim 11. As explained in further detail above with respect to claim 1, Vallart fails to teach or suggest that when the aircraft is in a landed state, the controller is configured to determine the correction factor using a PID control mode. Rather, Vallart teaches against the use of a PID control during a landed state due to the dangers and complications of using such a control mode in the landed state. While Vallart discusses the use of integrating adjusters during the flight state and the touching state, Vallart explains that once the landed state is identified, a new control law is implemented in which direct and proportional control law is used and the integrating adjuster is not included. Therefore, Vallart fail to disclose the use of an integrated adjuster during a landed state. In fact, Vallart teaches away from using an integrated adjuster during the landed state due to the dangers and problems associated with using such a control law during the landed state. Therefore, Vallart does not disclose each and every element of claim 11.
Romana fails to cure the deficiencies of Vallart. Romana does not discuss the use of
adjusters such as a direct, proportional, or integral adjusters. Rather Romana merely discusses the use of differential breaking to correct the direction of travel of the aircraft. Therefore, neither Vallart nor Romana, together or alone, teaches or suggests that when the aircraft is in a landed state, the controller is configured to determine the correction factor using a PID control mode, as recited in claim 11. Therefore, Applicant respectfully contends that claim 11 is allowable for at least these reasons, in addition to other reasons not explicitly discussed herein. Claims 12 and 14-16 ultimately depend from claim 11, and therefore, are allowable for at least the same reasons, in addition to independent reasons not explicitly discussed herein.
c.
Inde
p
endent Claim 17
Amended independent claim 17 recites a control system for an aircraft including, among other things, a plurality of sensors positioned on a corresponding landing gear of the aircraft and configured to sense a parameter on the corresponding landing gear, and a configured to calculate an output command based on the pilot input and the sensed parameters of the landing gear, wherein the controller calculates the output command by calculating a correction factor, wherein the setpoint and the first signal are inputs used to calculate the first correction factor, and
wherein, when the aircraft is in a landed state, the controller is configured to determine the correction factor usin
g
a PID control mode.
The Examiner rejects independent claim 17 as being anticipated by Vallart and notes the relevancy of Romana to claim 17. As explained in further detail above with respect to claim 1, Vallart fails to teach or suggest that when the aircraft is in a landed state, the controller is configured to determine the correction factor using a PID control mode. Rather, Vallart teaches against the use of a PID control during a landed state due to the dangers and complications of using such a control mode in the landed state. While Vallart discusses the use of integrating adjusters during the flight state and the touching state, Vallart explains that once the landed state is identified, a new control law is implemented in which direct and proportional control law is used and the integrating adjuster is not included. Therefore, Vallart fail to disclose the use of an integrated adjuster during a landed state. In fact, Vallart teaches away from using an integrated adjuster during the landed state due to the dangers and problems associated with using such a
control law during the landed state. Therefore, Vallart does not disclose each and every element of claim 11.
Romana fails to cure the deficiencies of Vallart. Romana does not discuss the use of
adjusters such as a direct, proportional, or integral adjusters. Rather Romana merely discusses the use of differential breaking to correct the direction of travel of the aircraft. Therefore, neither Vallart nor Romana, together or alone, teaches or suggests that when the aircraft is in a landed state, the controller is configured to determine the correction factor using a PID control mode, as recited in claim 17. Therefore, Applicant respectfully contends that claim 17 is allowable for at least these reasons, in addition to other reasons not explicitly discussed herein. Claims 20-22 ultimately depend from claim 17, and therefore, are allowable for at least the same reasons, in addition to independent reasons not explicitly discussed herein.
II.
Conclusion


CITED REFERENCE SUMMARIES: Ref US20160327958: Claims:
1. A control system, which system includes at least one control unit allowing a rotor of a rotorcraft to be driven, with the rotorcraft including at least three independent landing-gear units, with each landing-gear unit including means for detecting a ground reaction force F1, F2, F3 exerted on the landing-gear unit when the rotorcraft is in contact with the ground, and with the control system being suitable for receiving information from the detection means, and with the control system including:
identification means allowing the identification of a current state from among at least two different states of the rotorcraft, namely, a state of “touching” the ground and an “in flight” state, by comparing the reaction forces F1, F2, F3, as measured by the detection means, against predetermined low (SB) and high (SH) threshold values; with the “touching” state being identified when the following two conditions have been confirmed:
the reaction force F1, as measured by one of the detection means, is greater than the low threshold (SB) value but lower than the high threshold (SH) value; and
each of the reaction forces F2, F3, as measured by at least two other detection means, is lower than the low threshold (SB) value;
and with the “in flight” state being identified when each of the reaction forces F1, F2, F3, as measured simultaneously by the at least three detection means, is lower than the low threshold (SB) value;
control means for controlling the rotor of the rotorcraft according to at least two control laws that are separate from one another, depending on the current state of the rotorcraft as identified by the identification means, namely:
a first control law, implemented by the control means when the current state has been identified as the “in flight” state; and
a second control law, implemented by the control means when the current state has been identified as the “touching” state;
with the control system being characterized in that:
the low threshold (SB) value is greater than zero;
the control means include a computer allowing the implementation of the first and second control laws by means of integrating adjusters that stay active, with the integrating adjusters enabling at least two different settings of an integral adjustment dynamic, with a first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law and a second setting of the integral adjustment dynamic of the integrating adjuster corresponding to the second control law, and with the second setting enabling the continuous active state of an integrated component located at the output of the integrating adjuster corresponding to the second control law; and
the second control law makes it possible to maintain full control authority over the control of the rotor of the rotorcraft.
2. A control system according to claim 1, wherein the first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law generates control changes that are faster than the ones obtained with the second setting of the integral adjustment dynamic of the integrating adjuster corresponding to the second control law.
3. A control system according to claim 1, wherein the first setting of the dynamic of the integrating adjuster corresponding to the first control law generates control changes that are twice as fast as the ones obtained with the second setting of the dynamic of the integrating adjuster corresponding to the second control law.
4. A control system according to claim 1, wherein the high threshold (SH) value is between 2,500 Newtons and 20,000 Newtons.
5. A control system according to claim 1, wherein the low threshold (SB) value is between 1,000 Newtons and 10,000 Newtons.
6. A rotorcraft including at least one control unit allowing a rotor of the rotorcraft to be driven, with the rotorcraft including at least three independent landing-gear units, with each landing-gear unit including means for detecting a ground reaction force F1, F2, F3 exerted on the landing-gear unit when the rotorcraft is in contact with the ground, wherein the rotorcraft includes a control system according to claim 1.
7. A method for controlling a rotorcraft by means of at least one control unit allowing a rotor of the rotorcraft to be driven, with the rotorcraft including at least three independent landing-gear units, with each landing-gear unit including means for detecting a ground reaction force F1, F2, F3 exerted on the landing-gear unit when the rotorcraft is in contact with the ground, with the method including stages consisting of:
storing, in memory, predetermined low (SB) and high (SH) threshold values; measuring the ground reaction forces F1, F2, F3 on each landing-gear unit; comparing the reaction forces F1, F2, F3 against the predetermined low (SB) and high (SH) threshold values; identifying a current state from among at least two different states of the rotorcraft, namely, a state of “touching” the ground and an “in flight” state; with the “touching” state being identified when the following two conditions have been confirmed:
the reaction force F1, as measured by one of the detection means, is greater than the low threshold (SB) value but lower than the high threshold (SH) value; and
each of the reaction forces F2, F3, as measured by at least two other detection means (8), is lower than the low threshold (SB) value; and
 with the “in flight” state being identified when each of the reaction forces F1, F2, F3, as measured simultaneously by the at least three detection means, is lower than the low threshold (SB) value; controlling the rotor of the rotorcraft according to at least two control laws that are separate from one another, depending on the identified current state of the rotorcraft; a first control law, implemented when the current state has been identified as the “in flight” state; and a second control law, implemented when the current state has been identified as the “touching” state;
the low threshold (SB) value being greater than zero;
the first and second control laws are implemented by means of integrating adjusters that stay active, with the integrating adjusters enabling at least two different settings of an integral adjustment dynamic, with a first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law and a second setting of the integral adjustment dynamic of the integrating adjuster corresponding to the second control law, and with the second setting enabling the continuous active state of an integrated component located at the output of the integrating adjuster corresponding to the second control law; and
the second control law makes it possible to maintain full control authority over the control of the rotor of the rotorcraft.
8. A control method according to claim 7, wherein the first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law is selected so as to generate control changes that are faster than the ones obtained with the second setting of the integral adjustment dynamic of the integrating adjuster corresponding to the second control law.
Ref US20170008619: Claims:
1. A controller for an aircraft steering system, the controller being configured to receive a steering input representative of a desired direction of travel of a steerable nose landing gear, and to receive one or more force-based inputs representative of lateral forces acting upon the nose landing gear, wherein the controller is adapted to automatically adjust the steering input based upon the one or more force-based inputs and output an adjusted steering command for a steering actuator of the nose landing gear.
2. The controller according to claim 1, wherein the one or more force-based inputs include a closed-loop feedback signal representative of actual lateral force acting upon the nose landing gear.
3. The controller according to claim 1 wherein the one or more force-based inputs include an open-loop estimate of lateral force acting upon the nose landing gear.
4. The controller according to claim 3, wherein the controller is configured to calculate the estimated lateral force estimate based upon one or more aircraft parameters and a pre-defined model stored in the controller.
5. The controller according to claim 2 wherein the one or more force-based inputs include an open-loop estimate of lateral force acting upon the nose landing gear and wherein the controller is configured to automatically revert to the open-loop estimate of lateral force when the closed-loop signal of actual lateral force is unavailable or is unreliable.
6. The controller according to claim 1, wherein the controller is configured to calculate a maximum nose landing gear lateral force for maintaining the lateral force within operational boundaries.
7. The controller according to claim 6, wherein the controller is configured to calculate the maximum nose landing gear lateral force based upon the one or more force-based inputs representative of lateral forces acting upon the nose landing gear.
8. The controller according to claim 6, wherein the controller is configured to limit the adjustment to the steering input to ensure that the maximum nose landing gear lateral force is not exceeded.
9. The controller according to claim 1, wherein the controller is configured to calculate a desired nose landing gear lateral force to be generated based upon the steering input.
10. The controller according to claim 9, wherein the controller is configured to compare the desired nose landing gear lateral force with a signal representative of actual or estimated lateral force acting on the nose landing gear, and to automatically adjust the steering input based upon any difference between these signals.
11. The controller according to claim 1, wherein the controller is configured to calculate an instantaneous maximum nose wheel slip angle based upon the one or more force-based inputs representative of lateral forces acting upon the nose landing gear.
12. The controller according to claim 1, wherein the controller is configured to calculate a maximum nose wheel slip angle for maintaining the slip angle within operational boundaries.
13. The controller according to claim 12, wherein the controller is configured to limit the adjustment to the steering input to ensure that the maximum nose wheel slip angle is not exceeded.
14. The controller according to claim 1, wherein the controller is configured to output a signal representative of a braking force to be achieved by a braking system of the same aircraft so as to provide differential braking.
15. The controller according to claim 14, wherein the controller is configured to split the steering input into a steering component to be accomplished by the steering system and a differential braking component to be accomplished by the braking system.
16. The controller according to claim 14, wherein the controller is configured to output the differential braking signal when the controller determines that the steering actuator is unable to accomplish the steering command.
17. An aircraft steering system including the controller of claim 1.
18. The aircraft steering system according to claim 17, wherein the automatic steering input adjustment function is selectively activated.
19. The aircraft steering system according to claim 18, further comprising a steering input device for receiving the steering input, and a steering actuator for directing a rotary steering angle of a steerable nose landing gear, wherein the steering command to the steering actuator remains invariant of lateral forces acting on the nose landing gear when the automatic steering input adjustment function is de-activated.
20. An aircraft including the steering system of claim 17.
21. The aircraft according to claim 20, further comprising a braking system for decelerating the aircraft and operably coupled to the steering system for performing differential braking.
22. A method for steering an aircraft, the method comprising:
automatically adjusting a steering input representative of a desired direction of travel of a steerable nose landing gear based upon one or more force-based inputs representative of lateral forces acting upon the nose landing gear, and controlling a steering actuator of the nose landing gear using the adjusted steering command.
Ref EP3208190: Claims:
A method of controlling rotor moments comprising:
receiving, in a flight control computer (FCC), a rotor moment reference value based on pilot inceptor inputs;
sensing rotor moment from one or more sensors;
receiving, in the FCC, a rotary wing aircraft condition parameter; and
establishing, through the FCC, a rotor blade pitch angle for one or more of a plurality of rotor blades that counteracts external forces acting upon the rotary wing aircraft.
The method of claim 1, wherein receiving the rotor moment reference value includes receiving a cyclic inceptor position.
The method of claim 1 or 2, wherein receiving the rotary wing aircraft condition parameter includes receiving signals from one or more weight-on-wheels sensors.
The method of claim 3, wherein receiving signals from one or more weight-on-wheels sensors includes detecting a transition from a proportional ground control mode to a model following controls mode.
The method of one of the claims 1 to 4, wherein sensing rotor moment includes receiving signals from at least one sensor mounted to one of a rotor shaft, a rotor hub arm, and along a span of a rotor blade.
The method of one of the claims 1 to 5, wherein establishing, through the FCC, the rotor blade pitch angle includes comparing rotor moment inputs from at least two flight control computers.
The method of one of the claims 1 to 6, wherein establishing, through the FCC, the rotor blade pitch angle includes determining a rotor moment error and providing rotor moment feedback commands to one or more rotor blades.
The method of one of the claims 1 to 7, wherein establishing, through the FCC, the rotor blade pitch angle includes determining a rotor moment error and providing rotor moment feed forward commands to one or more rotor blades
The method of one of the claims 1 to 8, wherein establishing, through the FCC, the rotor blade pitch angle includes combining rotor moment feedback commands and rotor moment feed forward commands to one or more rotor blades.
A rotor moment control system for a rotary wing aircraft comprising:
a flight control computer (FCC);
a rotor moment control system configured to analyze rotor moments and reduce rotary wing aircraft transients during a transition from a proportional ground control mode and a model following controls mode, the rotor moment control system including computer readable program code embodied therewith, the computer readable program code, when executed by the FCC, causes the FCC to:
receive a rotor moment reference value based on pilot inceptor inputs;
sense rotor moment from one or more sensors;
receive a rotary wing aircraft condition parameter; and
establish a rotor blade pitch angle for one or more of a plurality of rotor blades that counteracts external forces acting upon the rotary wing aircraft.
The system of claim 10, wherein the computer readable program code, when executed by the FCC, causes the FCC to: receive a cyclic position when receiving the rotor moment reference value includes receiving a cyclic inceptor position.
The system of claim 10 or 11, wherein the computer readable program code, when executed by the FCC, causes the FCC to: receive signals from one or more weight-on-wheels sensors indicating a transition from a proportional ground control mode to a model following controls mode when receiving the rotary wing aircraft condition parameter.
The system of one of the claims 10 to 12, wherein establishing, through the processor, the rotor blade pitch angle includes comparing rotor moment inputs from at least two flight control computers.
The system of one of the claims 10 to 13, wherein the computer readable program code, when executed by the FCC, causes the FCC to: determine a rotor moment error and provide rotor moment feedback commands to one or more rotor blades when establishing the rotor blade pitch angle.
The system of one of the claims 10 to 14, wherein the computer readable program code, when executed by the FCC, causes the FCC to: combining rotor moment feedback commands and rotor moment feed forward commands to one or more rotor blades when establishing the rotor blade pitch angle.
Ref US20180305038: Claims:
1. A touchdown control system for a vertical take-off and landing (VTOL) aircraft comprising:
a detection controller receptive to aircraft state signals including one or more of an altitude signal, an aircraft attitude signal, an external environment signal, an aircraft velocity signal, an attitude rate signal, and a proximity signal; an enable controller operatively connected to the detection controller, the enable controller selectively providing a touchdown control signal based on one or more of the altitude signal, the aircraft attitude signal, the external environment signal, the aircraft velocity signal, the attitude rate signal, and the proximity signal; and a regulation controller operatively connected to the enable controller, the regulation controller selectively adjusting aircraft control surfaces based on the touchdown control signal to facilitate final touchdown of the VTOL aircraft.
2. The touchdown control system according to claim 1, wherein the detection controller determines a proximity of a landing surface and an undesirable aircraft condition based on signals from the one or more of the altitude signal, the aircraft attitude signal, the external environment signal, the aircraft velocity signal, the attitude rate signal, and the proximity signal.
3. The touchdown control system according to claim 2, wherein the undesirable aircraft condition includes one or more of an aircraft attitude exceeding a pre-selected aircraft attitude threshold, aircraft velocity exceeding a pre-selected velocity threshold, aircraft attitude exceeding a pre-selected attitude threshold, wind exceeding a pre-selected wind speed threshold, and an attitude of a landing surface exceeding a pre-selected landing surface attitude threshold.
4. The touchdown control system according to claim 2, wherein the undesirable aircraft condition includes at least of a one engine inoperative (OEI) condition, a multiple engine inoperative (MEI) condition, and an all engine inoperative (AEI) condition.
5. The touchdown control system according to claim 1, wherein the regulation controller selectively controls one or more of a yaw command, a pitch command, a roll command, a lift command, a velocity command, an attitude command, and a thrust command of the VTOL aircraft based on the touchdown control signal from the enable controller.
6. The touchdown control system according to claim 1, wherein the enable controller issues the touchdown control signal based on an operating mode of the VTOL aircraft.
7. A vertical take-off and landing (VTOL) aircraft comprising:
an airframe; a main rotor assembly supported by the airframe; at least one engine operatively connected to the main rotor assembly; a plurality of aircraft control surfaces associated with at least one of the main rotor assembly and the airframe; and a touchdown control system for a vertical take-off and landing (VTOL) aircraft comprising:
a detection controller receptive to aircraft state signals including one or more of an altitude signal, an aircraft attitude signal, an external environment signal, a velocity signal, an attitude rate signal, and a proximity signal;
an enable controller operatively connected to the detection controller, the enable controller selectively providing a touchdown control signal base
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