Office Action 2: Final Rejection (2024-06-11)
Rejection types: 102 - Lack of Novelty, 103 - Obviousness

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 based on the pilot input and the sensed force on the landing gear, and
transmit the output command to control the aircraft.
2\. (Canceled)
3\. (Currently Amended) The control system of claim \[\[2\]\]1, wherein the controller determines the correction factor using a PID control mode.
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\. (Original) The control system of claim 6, wherein the controller is further configured to determine when the aircraft is in a landed flight state based on a signal received from the second sensor.
8\. (Original) The control system of claim 1, wherein the first sensor is active when the aircraft is in a 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\. (Currently Amended) The control system of claim 9, wherein the controller is configured to calculate \[\[an\]\]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 based on the setpoint and the first signal;
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.

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\. (Original) The method of claim 11, wherein determining the output command includes determining the output command using a PID control mode.

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 based on the pilot input and the sensed parameters on the landing gear.

18\. (Canceled)

19\. (Original) The control system of claim 18, wherein the controller is configured to determine the correction factor using a PID control mode.

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\. (New) 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\. (New) 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:
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 .
Status of Claims
Claims 1-20 filed on 09/29/2021 have been examined.
This Office Action is in response to the Applicant’s amendments and remarks filed on 02/22/2024. Claims 1, 3, 10, 11, and 17 have been amended.  Claims 2 and 18 are canceled. Claims 21 and 22 are new claims that have been added to the claim set.  Claims 1, 3-17, and 19-22 are currently pending and addressed below.
Response to Remarks/Arguments
Applicant’s accompanying amendments and arguments, on pages 6-9 of the Applicant Arguments/Remarks (hereinafter referred to as the “Remarks”), filed 02/22/2024, with respect to the rejection of claims 1, 11, and 17 under 35 U.S.C. 102 stating “… Claim 1 is amended to include the subject matter of dependent claim 2 in order to clarify that the controller calculates the output command by calculating a correction factor based on the pilot input and the sensed force on the landing gear… Accordingly, Vallart discloses the use of sensors on the landing gear solely for the purpose of determining a flight state of the rotorcraft.  However, Vallart fails to teach or suggest utilizing sensors on the landing gear to calculate a correction factor…  Applicant respectfully submits that claim 1 is allowable over the prior art… As explained in further detail above with respect to claim 1… Applicant respectfully submits that claim 11 is allowable over the prior art… Claim 17 is amended to include the subject matter of dependent claim 18… As explained in further detail above with respect to claim 1… Applicant respectfully submits that claim 17 is allowable over the prior art…” have been considered and are not persuasive because Vallart discloses calculating a correction factor based on the pilot input/setpoint and the sensed force on the landing gear/first signal (See at least [0054]-[0060] – “The detection means that make it possible to measure the reaction forces F1, F2, F3 that are applied to the various landing-gear units… The signals from these sensors are then transmitted to a computer and processed by ground/flight software that allows the various states of the rotorcraft to be identified… The in-flight state is defined by a lack of contact between the rotorcraft and the ground, within the limits of accuracy of the sensors. Meanwhile, the “touching” state is defined by the detection, by at least one of the landing-gear units, of a load that is between a low threshold (S.sub.B) value and a high threshold (S.sub.H) value… as soon as the identification means identify a change between these two states, the control means modify the control law, allowing the rotor to be driven, in particular, by means of the integrating adjusters… the pilot retains full control authority during this critical phase…  the first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law can 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… when the “touching” state is detected, the computer for the control means makes it possible to limit the integral adjustment dynamic of the integrating adjusters, by clipping the error generated by a comparator between the control setpoint at the computer input and the control setpoint adjusted by the computer in order to control the rotor…” and [0111] of Vallart – “As shown in FIG. 3, the control means 10 may include a computer 11 that includes integrating adjusters 12, 22 that are suitable for implementing the first and second control laws for the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).  Examiner notes that Vallart discloses a rotorcraft control system that measures forces acting on the landing gear to identify states of the rotorcraft and then calculates adjusters (i.e., correction factors) based on the sensed forces acting on the landing gear and determined states as a result of forces on the landing gear (i.e., in flight vs touching) to implement a first control law and second control law via adjusters 22 and 12, respectively, for the servocontrols corresponding the rotors of the rotorcraft, with the adjuster corresponding to the first control law generating control changes that are faster than the ones obtained with the adjuster corresponding to the second control law as specified in at least [0059] of Vallart.  Therefore, the rejection of claims 1, 11, 17 are under 35 U.S.C. 102 are maintained.
Claim Interpretation
The following is a quotation of 35 U.S.C. 112(f):
(f) Element in Claim for a Combination. – An element in a claim for a combination may be expressed as a means or step for performing a specified function without the recital of structure, material, or acts in support thereof, and such claim shall be construed to cover the corresponding structure, material, or acts described in the specification and equivalents thereof.
The following is a quotation of pre-AIA  35 U.S.C. 112, sixth paragraph:
An element in a claim for a combination may be expressed as a means or step for performing a specified function without the recital of structure, material, or acts in support thereof, and such claim shall be construed to cover the corresponding structure, material, or acts described in the specification and equivalents thereof.
The claims in this application are given their broadest reasonable interpretation using the plain meaning of the claim language in light of the specification as it would be understood by one of ordinary skill in the art.  The broadest reasonable interpretation of a claim element (also commonly referred to as a claim limitation) is limited by the description in the specification when 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, is invoked.
As explained in MPEP § 2181, subsection I, claim limitations that meet the following three-prong test will be interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph:
(A)	the claim limitation uses the term “means” or “step” or a term used as a substitute for “means” that is a generic placeholder (also called a nonce term or a non-structural term having no specific structural meaning) for performing the claimed function;
(B)	the term “means” or “step” or the generic placeholder is modified by functional language, typically, but not always linked by the transition word “for” (e.g., “means for”) or another linking word or phrase, such as “configured to” or “so that”; and
(C)	the term “means” or “step” or the generic placeholder is not modified by sufficient structure, material, or acts for performing the claimed function.
Use of the word “means” (or “step”) in a claim with functional language creates a rebuttable presumption that the claim limitation is to be treated in accordance with 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph. The presumption that the claim limitation is interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, is rebutted when the claim limitation recites sufficient structure, material, or acts to entirely perform the recited function.
Absence of the word “means” (or “step”) in a claim creates a rebuttable presumption that the claim limitation is not to be treated in accordance with 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph. The presumption that the claim limitation is not interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, is rebutted when the claim limitation recites function without reciting sufficient structure, material or acts to entirely perform the recited function.
Claim limitations in this application that use the word “means” (or “step”) are being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, except as otherwise indicated in an Office action. Conversely, claim limitations in this application that do not use the word “means” (or “step”) are not being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, except as otherwise indicated in an Office action.
This application includes one or more claim limitations that do not use the word “means,” but are nonetheless being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, because the claim limitation(s) uses a generic placeholder that is coupled with functional language without reciting sufficient structure to perform the recited function and the generic placeholder is not preceded by a structural modifier.  Such claim limitation(s) is/are:
“a pilot input device” provided in claim 1
The specification and drawings were used to define the generic placeholders specified:
Specification [0042] – The inputs 126 from the pilot may be initiated through various input devices, such as control actuators and buttons, collective stick, cyclic controls, and the like.
Because this/these claim limitation(s) is/are being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, it/they is/are being interpreted to cover the corresponding structure described in the specification as performing the claimed function, and equivalents thereof.
If applicant does not intend to have this/these limitation(s) interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph, applicant may:  (1) amend the claim limitation(s) to avoid it/them being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph (e.g., by reciting sufficient structure to perform the claimed function); or (2) present a sufficient showing that the claim limitation(s) recite(s) sufficient structure to perform the claimed function so as to avoid it/them being interpreted under 35 U.S.C. 112(f) or pre-AIA  35 U.S.C. 112, sixth paragraph.
Claim Rejections - 35 USC § 102
In the event the determination of the status of the application as subject to AIA  35 U.S.C. 102 and 103 (or as subject to pre-AIA  35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of the appropriate paragraphs of 35 U.S.C. 102 that form the basis for the rejections under this section made in this Office action:
A person shall be entitled to a patent unless –
(a)(1) the claimed invention was patented, described in a printed publication, or in public use, on sale, or otherwise available to the public before the effective filing date of the claimed invention.
(a)(2) the claimed invention was described in a patent issued under section 151, or in an application for patent published or deemed published under section 122(b), in which the patent or application, as the case may be, names another inventor and was effectively filed before the effective filing date of the claimed invention.
Claims 1, 3, 5, 11-14, 17, and 19 are rejected are rejected under 35 U.S.C. 102(a)(1) as being anticipated by Vallart et al. US 20160327958 (“Vallart”).
For claim 1, Vallart discloses a control system for an aircraft (See at least [0003] of Vallart – “The present invention relates to a rotorcraft control system…”), the control system comprising:
a pilot input device configured to receive a pilot input (See at least [0003] – “…the collective stick and the pitch lever respectively control the collective pitch and the cyclic pitch of the blades of a main rotor of a rotorcraft” and Fig. 1 – reference character 2);
a first sensor positioned on a landing gear of the aircraft, the first sensor configured to sense a force on the landing gear (See at least [0038] – “…a control system that includes at least one control unit that allows a rotorcraft rotor to be driven. Such a rotorcraft also includes 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. The control system is thus suitable for receiving information from the detection means” and [0098] of Vallart – “Such a rotorcraft 4 is also equipped with at least three landing-gear units 5, 6, 7, and each of these landing-gear units 5, 6, 7 includes detection means 8, such as sensors…”); and
a controller in communication with the pilot input device and the first sensor (See at least [0106] – “As shown in FIG. 2… an analytical unit 18, such as a computer or a microprocessor that receives the signals from the detection means 8 that represent the reaction forces F1, F2, and F3 between the landing-gear units 5, 6, 7…” and [0107] of Vallart – “Once the comparison has been made by the analytical unit 18, the current state is identified and then transmitted to the control means 10 that include a computer 11, so that one of the different control laws for the rotor 3, 13 can be generated. Thus, the control means 10 also receive a control setpoint generated by the control unit 2…”), the controller configured to
receive the pilot input via the pilot input device (See at least [0107] of Vallart – “Thus, the control means 10 also receive a control setpoint generated by the control unit 2…”),
receive a sensed force on the landing gear via the first sensor (See at least [0106] of Vallart – “…a computer or a microprocessor that receives the signals from the detection means 8 that represent the reaction forces F1, F2, and F3 between the landing-gear units 5, 6, 7 and the ground…”),
calculate an output command based at least on the pilot input and the force on the landing gear (See at least [0107] of Vallart – “Once the comparison has been made by the analytical unit 18, the current state is identified and then transmitted to the control means 10…. Thus, the control means 10 also receive a control setpoint generated by the control unit 2, allowing, as an output, the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”), wherein the controller calculates the output command by calculating a correction factor based on the pilot input and the sensed force on the landing gear (See at least [0054]-[0060] – “The detection means that make it possible to measure the reaction forces F1, F2, F3 that are applied to the various landing-gear units… The signals from these sensors are then transmitted to a computer and processed by ground/flight software that allows the various states of the rotorcraft to be identified… The in-flight state is defined by a lack of contact between the rotorcraft and the ground, within the limits of accuracy of the sensors. Meanwhile, the “touching” state is defined by the detection, by at least one of the landing-gear units, of a load that is between a low threshold (S.sub.B) value and a high threshold (S.sub.H) value… as soon as the identification means identify a change between these two states, the control means modify the control law, allowing the rotor to be driven, in particular, by means of the integrating adjusters… the pilot retains full control authority during this critical phase…  the first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law can 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… when the “touching” state is detected, the computer for the control means makes it possible to limit the integral adjustment dynamic of the integrating adjusters, by clipping the error generated by a comparator between the control setpoint at the computer input and the control setpoint adjusted by the computer in order to control the rotor…” and [0111] of Vallart – “As shown in FIG. 3, the control means 10 may include a computer 11 that includes integrating adjusters 12, 22 that are suitable for implementing the first and second control laws for the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”), and
transmit the output command to control the aircraft (See at least [0107] of Vallart – “….the control means 10 … allowing… the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).
For claim 3, Vallart discloses wherein the controller determines the correction factor using a PID control mode gear (See at least [0062] of Vallart – “the integrating adjusters are advantageously selected from among the group including, in particular, pure integrating adjusters and so-called Proportional Integral Derivative (PID) adjusters”).
For claim 5, Vallart discloses wherein the output command controls a rotor of the aircraft (See at least [0107] of Vallart – “….the control means 10… allowing, as an output, the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).
For claim 11, Vallart discloses a method of controlling an aircraft (See at least [0003] of Vallart – “The present invention relates to a rotorcraft control system…”), the method comprising:
receiving, by an electronic processor, a pilot input including a setpoint (See at least [0107] of Vallart – “Thus, the control means 10 also receive a control setpoint generated by the control unit 2…”;
receiving, by the electronic processor, a first signal from a first sensor positioned on a landing gear of the aircraft (See at least [0106] – “As shown in FIG. 2… an analytical unit 18, such as a computer or a microprocessor that receives the signals from the detection means 8 that represent the reaction forces F1, F2, and F3 between the landing-gear units 5, 6, 7…” and [0098] of Vallart – “Such a rotorcraft 4 is also equipped with at least three landing-gear units 5, 6, 7, and each of these landing-gear units 5, 6, 7 includes detection means 8, such as sensors that are suitable for measuring a relative displacement in a landing-gear unit 5, 6, 7”);
calculating, by the electronic processor, a first correction factor based on the setpoint and the first signal (See at least [0068] – “Such a rotorcraft also includes 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 [0060] of Vallart – “… when the “touching” state is detected, the computer for the control means makes it possible to limit the integral adjustment dynamic of the integrating adjusters, by clipping the error generated by a comparator between the control setpoint at the computer input and the control setpoint adjusted by the computer in order to control the rotor…”);
determining, by the electronic processor, an output command based on the first correction factor, the output command including instructions for controlling the aircraft (See at least [0111] of Vallart – “As shown in FIG. 3, the control means 10 may include a computer 11 that includes integrating adjusters 12, 22 that are suitable for implementing the first and second control laws for the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4”); and
transmitting, by the electronic processor, the output command to control the aircraft (See at least [0108] of Vallart – “Furthermore, the control means 10 make it possible to control the rotor 3, 13 of the rotorcraft 4 (or, more accurately, the servocontrols of the actuators of the rotor 3, 13), according to at least two control laws that are separate from one another, depending on the current state of the rotorcraft 4 as identified by the identification means 9…”).
For claim 12, Vallart discloses 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 (See at least [0098] – “Such a rotorcraft 4 is also equipped with at least three landing-gear units 5, 6, 7, and each of these landing-gear units 5, 6, 7 includes detection means 8 such, as sensors…” and [0068] of Vallart – “Such a rotorcraft also includes 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”).
For claim 13, Vallart discloses wherein determining the output command includes determining the output command using a PID control mode (See at least [0062] of Vallart – “Therefore, such integrating adjusters …. may be implemented by means of the same algorithm, some of whose calculation parameters may be modified between the two control laws). The integrating adjusters are advantageously selected from among the group including, in particular, pure integrating adjusters and so-called Proportional Integral Derivative (PID) adjusters”).
For claim 14, Vallart discloses wherein the output command controls a rotor interface (See at least [0107] of Vallart – “….the control means 10… allowing, as an output, the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).
For claim 17, Vallart discloses a control system for an aircraft (See at least [0003] of Vallart – “The present invention relates to a rotorcraft control system…”), the control system comprising:
a pilot input device configured to receive a pilot input (See at least [0003] – “…the collective stick and the pitch lever respectively control the collective pitch and the cyclic pitch of the blades of a main rotor of a rotorcraft” and Fig. 1 – reference character 2);
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 (See at least [0038] – “…a control system that includes at least one control unit that allows a rotorcraft rotor to be driven. Such a rotorcraft also includes 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. The control system is thus suitable for receiving information from the detection means” and [0098] of Vallart – “Such a rotorcraft 4 is also equipped with at least three landing-gear units 5, 6, 7, and each of these landing-gear units 5, 6, 7 includes detection means 8, such as sensors that are suitable for measuring a relative displacement in a landing-gear unit 5, 6, 7”); and
a controller in communication with the plurality of sensors (See at least [0106] – “As shown in FIG. 2… an analytical unit 18, such as a computer or a microprocessor that receives the signals from the detection means 8 that represent the reaction forces F1, F2, and F3 between the landing-gear units 5, 6, 7…” and [0107] of Vallart – “Once the comparison has been made by the analytical unit 18, the current state is identified and then transmitted to the control means 10 that include a computer 11, so that one of the different control laws for the rotor 3, 13 can be generated. Thus, the control means 10 also receive a control setpoint generated by the control unit 2…”), the controller configured to calculate an output command based on the pilot input and the sensed parameters of the landing gear (See at least [0107] of Vallart – “Once the comparison has been made by the analytical unit 18, the current state is identified and then transmitted to the control means 10…. Thus, the control means 10 also receive a control setpoint generated by the control unit 2, allowing, as an output, the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”), the output command including instructions for controlling a rotor of the aircraft (See at least [0107] of Vallart – “….the control means 10 … allowing… the driving of the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”),
wherein the controller calculates the output command by calculating a correction factor based on the pilot input and the sensed parameters on the landing gear (See at least [0054]-[0060] – “The detection means that make it possible to measure the reaction forces F1, F2, F3 that are applied to the various landing-gear units… The signals from these sensors are then transmitted to a computer and processed by ground/flight software that allows the various states of the rotorcraft to be identified… The in-flight state is defined by a lack of contact between the rotorcraft and the ground, within the limits of accuracy of the sensors. Meanwhile, the “touching” state is defined by the detection, by at least one of the landing-gear units, of a load that is between a low threshold (S.sub.B) value and a high threshold (S.sub.H) value… as soon as the identification means identify a change between these two states, the control means modify the control law, allowing the rotor to be driven, in particular, by means of the integrating adjusters… the pilot retains full control authority during this critical phase…  the first setting of the integral adjustment dynamic of the integrating adjuster corresponding to the first control law can 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… when the “touching” state is detected, the computer for the control means makes it possible to limit the integral adjustment dynamic of the integrating adjusters, by clipping the error generated by a comparator between the control setpoint at the computer input and the control setpoint adjusted by the computer in order to control the rotor…” and [0111] of Vallart – “As shown in FIG. 3, the control means 10 may include a computer 11 that includes integrating adjusters 12, 22 that are suitable for implementing the first and second control laws for the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).
For claim 19, Vallart discloses wherein the controller is configured to determine the correction factor using a PID control mode (See at least [0062] of Vallart – “Therefore, such integrating adjusters …. may be implemented by means of the same algorithm, some of whose calculation parameters may be modified between the two control laws). The integrating adjusters are advantageously selected from among the group including, in particular, pure integrating adjusters and so-called Proportional Integral Derivative (PID) adjusters”).
Claim Rejections - 35 USC § 103
In the event the determination of the status of the application as subject to AIA  35 U.S.C. 102 and 103 (or as subject to pre-AIA  35 U.S.C. 102 and 103) is incorrect, any correction of the statutory basis for the rejection will not be considered a new ground of rejection if the prior art relied upon, and the rationale supporting the rejection, would be the same under either status.
The following is a quotation of 35 U.S.C. 103 which forms the basis for all obviousness rejections set forth in this Office action:
A patent for a claimed invention may not be obtained, notwithstanding that the claimed invention is not identically disclosed as set forth in section 102, if the differences between the claimed invention and the prior art are such that the claimed invention as a whole would have been obvious before the effective filing date of the claimed invention to a person having ordinary skill in the art to which the claimed invention pertains. Patentability shall not be negated by the manner in which the invention was made.
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”).
For claim 4, Vallart fails to specifically disclose wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed force on the landing gear.
However, Romana, in the same field of endeavor teaches wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed force on the landing gear (See at least [0087] of Romana – “The NLG force controller 115 also uses the current lateral force Fy.sub.N generated in the nose landing gear to detect unwanted yaw moments causing changes in heading which do not correspond to a change in a steering input command… If an unwanted change in heading is detected, the NLG force controller 115 responds by adjusting the slip angle command to control the lateral force generated by the steering wheels 13 to counteract the undesired yaw moment”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Romana teaches a controller for an aircraft steering system that adjusts a steering command based on undesired moments detected on the nose landing gear.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of calculating an output command by determining the moment of the landing gear based on the sensed force on the landing gear, as taught by Romana, in order to counteract unwanted yaw moments as a result of lateral forces acting on the nose landing gear by transmitting a control signal that adjusts the steering system to maintain control of the aircraft as specified in [0087] of Romana.
For claim 15, Vallart fails to specifically disclose wherein determining the output command includes determining the moment of the landing gear
However, Romana, in the same field of endeavor teaches wherein determining the output command includes determining the moment of the landing gear (See at least [0087] of Romana – “The NLG force controller 115 also uses the current lateral force Fy.sub.N generated in the nose landing gear to detect unwanted yaw moments causing changes in heading which do not correspond to a change in a steering input command… If an unwanted change in heading is detected, the NLG force controller 115 responds by adjusting the slip angle command to control the lateral force generated by the steering wheels 13 to counteract the undesired yaw moment”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Romana teaches a controller for an aircraft steering system that adjusts a steering command based on undesired moments detected on the nose landing gear.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of determining an output command while determining the moment of the landing gear, as taught by Romana, in order to counteract unwanted yaw moments as a result of lateral forces acting on the nose landing gear by transmitting a control signal that adjusts the steering system to maintain control of the aircraft as specified in [0087] of Romana.
For claim 20, Vallart fails to specifically disclose wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed parameters on the landing gear.
However, Romana, in the same field of endeavor teaches wherein the controller calculates the output command by determining the moment of the landing gear based on the sensed parameters on the landing gear (See at least [0087] of Romana – “The NLG force controller 115 also uses the current lateral force Fy.sub.N generated in the nose landing gear to detect unwanted yaw moments causing changes in heading which do not correspond to a change in a steering input command… If an unwanted change in heading is detected, the NLG force controller 115 responds by adjusting the slip angle command to control the lateral force generated by the steering wheels 13 to counteract the undesired yaw moment”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Romana teaches a controller for an aircraft steering system that adjusts a steering command based on undesired moments detected on the nose landing gear.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a controller calculating an output command by determining the moment of the landing gear based on the sensed parameters on the landing gear, as taught by Romana, in order to counteract unwanted yaw moments as a result of lateral forces acting on the nose landing gear by transmitting a control signal that adjusts the steering system to maintain control of the aircraft as specified in [0087] of Romana.
Claims 6-8 are rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of White EP 3208190 (“White”).
For claim 6, Vallart fails to specifically disclose further comprising a second sensor positioned on the landing gear, the second sensor being a weight on wheels sensor.
However, White, in the same field of endeavor teaches further comprising a second sensor positioned on the landing gear, the second sensor being a weight on wheels sensor (See at least [0028] of White – “Rotor moment control system 64 may also receive input from an aircraft condition parameter sensor system 80. Aircraft condition parameter sensor system 80 may include a first weight-on-wheels (WOW) sensor 82, a second WOW sensor 83, and a third WOW sensor 84…”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while White teaches a rotor moment control system that uses weight on wheel sensors to detect whether an aircraft is resting on the ground.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a second sensor positioned on the landing gear being a weigh on wheels sensor, as taught by White, in order to determine whether the aircraft is contacting the ground as specified in [0028] of White.
For claim 7, Vallart fails to specifically disclose wherein the controller is further configured to determine when the aircraft is in a landed flight state based on a signal received from the second sensor.
However, White, in the same field of endeavor teaches wherein the controller is further configured to determine when the aircraft is in a landed flight state based on a signal received from the second sensor (See at least [0028] of White – “Rotor moment control system 64 may also receive input from an aircraft condition parameter sensor system 80. Aircraft condition parameter sensor system 80 may include a first weight-on-wheels (WOW) sensor 82, a second WOW sensor 83, and a third WOW sensor 84…”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while White teaches a rotor moment control system that uses weight on wheel sensors to detect whether an aircraft is resting on the ground.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of determining when an aircraft is in a landed flight state based on a signal received from the second sensor, as taught by White, in order to determine whether the aircraft is contacting the ground as specified in [0028] of White.
For claim 8, Vallart fails to specifically disclose wherein the first sensor is active when the aircraft is in a landed flight state, and wherein the first sensor is inactive or zeroed-out when the aircraft is in an in-air flight state.
However, White, in the same field of endeavor teaches wherein the first sensor is active when the aircraft is in a landed flight state, and wherein the first sensor is inactive or zeroed-out when the aircraft is in an in-air flight state (See at least [0028] of White – “When WOW sensors 82, 83 and/or 84 no longer sense weight on wheels, a determination may be made that rotary wing aircraft 10 is transitioning from the proportional ground control mode to the model following controls or flight mode”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while White teaches a rotor moment control system that determines transition an aircraft to flight mode when WOW sensors no longer sense weight on the wheels of an aircraft.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of zeroing out sensor readings when an aircraft is in an in-airflight state, as taught by White, in order for the aircraft control system to determine that an aircraft is transitioning to in flight mode when the WOW sensors no longer sense weight on the wheels as specified in [0028] of 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”).
For claim 9, Vallart fails to specifically disclose 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.
However, Waltner, in the same field of endeavor teaches 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 (See at least [0034] of Waltner – “… control system 70 is operatively connected to a plurality of aircraft state sensors 94 that may include … an aircraft attitude sensor 97 that may provide an aircraft attitude signal”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Waltner teaches a touchdown control system for an aircraft that includes an attitude sensor positioned on the aircraft.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a second sensor positioned on the aircraft that measures the attitude of the aircraft, as taught by Waltner, in order for the aircraft control system to facilitate touchdown operations of an aircraft as specified in [0035] of Waltner.
For claim 10, Vallart fails to specifically disclose 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.
However, Waltner, in the same field of endeavor teaches 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 (See at least [0031] – “…. Touchdown control system 70 promises a desired landing or touchdown envelope that maintains force levels on landing gear (not shown) at or below selected levels… It is also to be understood that the touchdown control system may establish a selected aircraft attitude that promotes substantially simultaneous contact of all ground contact members” and [0035] of Waltner – “As will be detailed herein, touchdown control system 70 will, through CPU 80, control one or more control surface of VTOL aircraft 10 to facilitate a touchdown within a selected touchdown envelope to reduce pilot demand and prolong landing gear, support operational life. That is, depending on external factors such as wind speed, ground angle and the like, as well as aircraft state conditions, touchdown controller 70 may selectively signal … aircraft attitude threshold, attitude rate threshold, landing surface attitude threshold, rotor speed threshold, or engine operation state during touchdown, touchdown control system 70 may provide touchdown control and/or touchdown augmentation control as discussed herein”).   Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Waltner teaches a touchdown control system for an aircraft that determines output commands to facilitate touchdown of an aircraft by considering aircraft attitude thresholds, reduced pilot demand, and maintaining forces on landing gear below selected levels.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a controller calculating the output command based on the pilot input, the force on the landing gear, and the attitude of the aircraft sensed by the second sensor, as taught by Waltner, in order for the aircraft control system to facilitate touchdown operations that promote substantially simultaneous contact of all ground contact members as specified in [0031] of Waltner.
For claim 16, Vallart fails to specifically disclose 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.
However, Waltner, in the same field of endeavor teaches 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 (See at least [0034] of Waltner – “… control system 70 is operatively connected to a plurality of aircraft state sensors 94 that may include … an aircraft attitude sensor 97 that may provide an aircraft attitude signal”), wherein the output command is further based on the signal from the second sensor (See at least [0031] – “…. Touchdown control system 70 promises a desired landing or touchdown envelope that maintains force levels on landing gear (not shown) at or below selected levels… It is also to be understood that the touchdown control system may establish a selected aircraft attitude that promotes substantially simultaneous contact of all ground contact members”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Waltner teaches a touchdown control system for an aircraft that determines output commands to facilitate touchdown of an aircraft by considering aircraft attitude thresholds.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a processor obtaining an aircraft attitude by a sensor and calculating an output command based on the attitude of the aircraft sensed by the second sensor, as taught by Waltner, in order for the aircraft control system to facilitate touchdown operations that promote substantially simultaneous contact of all ground contact members as specified in [0031] of Waltner.
Claim 21 is rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of Seydoux et al. US 20110049290 (“Seydoux”).
For claim 21, Vallart discloses wherein the correction factor includes a first component and a second component, wherein the first component is calculated using the PID control mode (See at least [0062] of Vallart – “such integrating adjusters … some of whose calculation parameters may be modified between the two control laws…  selected from among the group including, in particular, pure integrating adjusters and so-called Proportional Integral Derivative (PID) adjusters...”).
Vallart fails to specifically disclose and wherein the second component is calculated using a PD control mode.
However, Seydoux, in the same field of endeavor teaches and wherein the second component is calculated using a PD control mode (See at least [0087]-[0089] of Seydoux – “The on-board software servo-controls the altitude of the aircraft about said setpoint altitude. Servo-control is performed with the help of a proportional-derivative (PD) corrector as shown in FIG. 4… In order to obtain a response that is both fast and presents little setpoint overshoot, it is necessary to transform this equation into a second order differential equation… That is why a PD corrector is used in which the derivative component is applied directly to the measurement (by simplifying the equations) so as to avoid introducing zeros in the closed loop transfer function and avoid having two adjustment parameters…”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while Seydoux teaches a system for controlling flight of an aircraft using a proportional derivative controller.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a second component being calculated using a PD control mode, as taught by Seydoux, in order for the aircraft control system to obtain a response that is both fast and presents little setpoint overshoot as specified in at least [0088] and [0089] of Seydoux.
Claim 22 is rejected under 35 U.S.C. 103 as being unpatentable over Vallart in view of Waltner, as applied to claim 10 above, and further in view of Thomas US 20210271267 (“Thomas”).
For claim 22, Vallart discloses 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 (See at least [0054]-[0060] – “The detection means that make it possible to measure the reaction forces F1, F2, F3 that are applied to the various landing-gear units… The signals from these sensors are then transmitted to a computer and processed by ground/flight software that allows the various states of the rotorcraft to be identified… as soon as the identification means identify a change between these two states, the control means modify the control law, allowing the rotor to be driven, in particular, by means of the integrating adjusters… and [0111] of Vallart – “As shown in FIG. 3, the control means 10 may include a computer 11 that includes integrating adjusters 12, 22 that are suitable for implementing the first and second control laws for the servocontrols of the actuators of the rotors 3, 13 of such a rotorcraft 4…”).
Vallart fails to specifically disclose and wherein the second component is calculated based on the parameter sensed by the second sensor.
However, Thomas, in the same field of endeavor teaches and wherein the second component is calculated based on the parameter sensed by the second sensor (See at least [0065] of Thomas – “FIG. 5 is a diagram of a system according to an aspect of the present disclosure wherein the system may contain a Flight Controller 520 capable of accepting feed-forward torques from the Position Controller 502. Flight Controller 520 … may include the systems of Attitude Controller 504 … Position Controller 502 may be capable of determining course and/or attitude corrections based upon position data received from Plant 510 as part of a feedback loop … Position Controller 502 may then forward analytically useful signals to Compensator 508…”).  Thus, Vallart discloses a control system for a rotorcraft that detects forces on landing gear equipment, while 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.
Therefore, it would have been obvious to one of ordinary skill in the art, before the effective filing date of the claimed invention, to modify the control systems and method for controlling an aircraft of Vallart to include the feature of a second component being calculated based on the parameter sensed by the second sensor, as taught by Thomas, in order for the aircraft control system to achieve a desired attitude as specified in at least [0072] and [0073] of Thomas.
Conclusion
Applicant's amendment necessitated the new ground(s) of rejection presented in this Office action.  Accordingly, THIS ACTION IS MADE FINAL.  See MPEP § 706.07(a).  Applicant is reminded of the extension of time policy as set forth in 37 CFR 1.136(a).
A shortened statutory period for reply to this final action is set to expire THREE MONTHS from the mailing date of this action.  In the event a first reply is filed within TWO MONTHS of the mailing date of this final action and the advisory action is not mailed until after the end of the THREE-MONTH shortened statutory period, then the shortened statutory period will expire on the date the advisory action is mailed, and any extension fee pursuant to 37 CFR 1.136(a) will be calculated from the mailing date of the advisory action.  In no event, however, will the statutory period for reply expire later than SIX MONTHS from the date of this final action.

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

NEXT 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

CITED REFERENCE SUMMARIES: Ref US20110049290A1: Claims:
A method of piloting a rotary-wing drone with automatic stabilization of hovering, the method comprising the steps consisting in: fitting the drone with a telemeter and a video camera; acquiring the altitude of the drone relative to the ground by means of a telemeter; acquiring the horizontal speed of the drone; and automatically stabilizing the drone in hovering by: servo-controlling the vertical thrust force of the drone so as to stabilize the altitude acquired by the telemeter; and servo-controlling the horizontal thrust force of the drone so as to obtain zero horizontal speed; the method being characterized in that: the video camera is a front-sight camera pointing towards the front of the drone; and the horizontal speed of the drone is acquired from a plurality of video images captured by said front-sight camera.
fitting the drone with a telemeter and a video camera;
acquiring the altitude of the drone relative to the ground by means of a telemeter;
acquiring the horizontal speed of the drone; and
automatically stabilizing the drone in hovering by: servo-controlling the vertical thrust force of the drone so as to stabilize the altitude acquired by the telemeter; and servo-controlling the horizontal thrust force of the drone so as to obtain zero horizontal speed;
servo-controlling the vertical thrust force of the drone so as to stabilize the altitude acquired by the telemeter; and
servo-controlling the horizontal thrust force of the drone so as to obtain zero horizontal speed;
the method being characterized in that: the video camera is a front-sight camera pointing towards the front of the drone; and the horizontal speed of the drone is acquired from a plurality of video images captured by said front-sight camera.
the video camera is a front-sight camera pointing towards the front of the drone; and
the horizontal speed of the drone is acquired from a plurality of video images captured by said front-sight camera.
A method of piloting a drone with automatic stabilization according to claim 1 , the method being characterized in that it further comprises the operations consisting in: defining elementary piloting functions, each elementary piloting function being suitable for determining flight parameters to be executed by a set of actuators of said drone so as to perform said elementary piloting function; providing a user with activation means for activating said elementary piloting functions; and the user piloting the drone by actuating said activation means for activating elementary piloting functions, with the drone being placed automatically in stabilized hovering flight whenever no function is being activated.
defining elementary piloting functions, each elementary piloting function being suitable for determining flight parameters to be executed by a set of actuators of said drone so as to perform said elementary piloting function;
providing a user with activation means for activating said elementary piloting functions; and
the user piloting the drone by actuating said activation means for activating elementary piloting functions, with the drone being placed automatically in stabilized hovering flight whenever no function is being activated.
A method of piloting a drone with automatic stabilization according to claim 2 , wherein said elementary piloting functions comprise the following actions: move up; move down; turn right; turn left; move forwards; reverse.
A method of piloting a drone with automatic stabilization according to claim 3 , wherein said elementary piloting functions also comprise: move left in horizontal translation; move right in horizontal translation.
A method of piloting a drone with automatic stabilization according to claim 2 , wherein said activation means are constituted by keys of a piloting box.
A method of piloting a drone with automatic stabilization according to claim 2 , wherein said activation means are constituted by traces drawn by a stylus on a touch-sensitive surface of a piloting box.
A method of piloting a drone with automatic stabilization according to claim 2 , wherein said activation means are multi-action means suitable for engaging, setting, and stopping associated elementary piloting functions.
A method of piloting a drone with automatic stabilization according to claim 2 , wherein said activation means also comprise means for activating automatic sequences.
A method of piloting a drone with automatic stabilization according to claim 8 , wherein said automatic sequences comprise the drone taking off and landing.
A rotary-wing drone comprising: a telemeter and a video camera; means for acquiring the altitude of the drone relative to the ground by means of the telemeter; means for acquiring the horizontal speed of the drone; and a system for automatically stabilizing hovering, the system comprising: servo-controlling the vertical thrust force of the drone so as to stabilize the altitude acquired by the telemeter; and servo-controlling the horizontal thrust force of the drone so as to obtain zero horizontal speed; the drone being characterized in that: the video camera is a front-sight video camera pointing towards the front of the drone; and the means for acquiring the horizontal speed of the drone are means for acquiring said speed from a plurality of video images captured by said front-sight camera.
a telemeter and a video camera;
means for acquiring the altitude of the drone relative to the ground by means of the telemeter;
means for acquiring the horizontal speed of the drone; and
a system for automatically stabilizing hovering, the system comprising: servo-controlling the vertical thrust force of the drone so as to stabilize the altitude acquired by the telemeter; and servo-controlling the horizontal thrust force of the drone so as to obtain zero horizontal speed;
the drone being characterized in that: the video camera is a front-sight video camera pointing towards the front of the drone; and the means for acquiring the horizontal speed of the drone are means for acquiring said speed from a plurality of video images captured by said front-sight camera.
the video camera is a front-sight video camera pointing towards the front of the drone; and
the means for acquiring the horizontal speed of the drone are means for acquiring said speed from a plurality of video images captured by said front-sight camera.
A piloting assembly, characterized in that it comprises: a rotary-wing drone according to claim 10 ; and a piloting box comprising means for activating elementary piloting functions; each elementary piloting function being suitable for determining flight parameters to be executed by a set of actuators of said drone so as to implement said elementary piloting function; and whenever no function is being activated, the drone is automatically placed in stabilized hovering flight by means of the system for automatically stabilizing hovering flight of the drone.
a rotary-wing drone according to claim 10 ; and
a piloting box comprising means for activating elementary piloting functions;
each elementary piloting function being suitable for determining flight parameters to be executed by a set of actuators of said drone so as to implement said elementary piloting function; and
whenever no function is being activated, the drone is automatically placed in stabilized hovering flight by means of the system for automatically stabilizing hovering flight of the drone.
A piloting box for a rotary-wing drone according to claim 10 , said piloting box being characterized in that it comprises means for activating elementary piloting functions; each elementary piloting function being suitable for determining flight parameters to be executed by a set of actuators of said drone so as to implement said elementary piloting function; and whenever no function is being activated, the drone is automatically placed in stabilized hovering flight.
whenever no function is being activated, the drone is automatically placed in stabilized hovering flight.
Ref US20210271267A1: [content unavailable]
Ref US20220326704A1: Claims:
A system for flight control configured for use in an electric aircraft, the system comprising: an inertial measurement unit, the inertial measurement unit configured to: detect at least an aircraft angle; and detect at least an aircraft angle rate; a flight controller, the flight controller comprising: an outer loop controller, the outer loop controller configured to: receive at least an input datum; receive at least an aircraft angle from the inertial measurement unit; and generate a rate setpoint as a function of at least an input datum; an inner loop controller, the inner loop controller configured to: receive at least an aircraft angle rate; receive the rate setpoint from the outer loop controller; and generate a moment datum as a function of the rate setpoint; and a mixer, the mixer configured to: receive the moment datum; perform a torque allocation as a function of the moment datum; and generate at least a motor command datum as a function of the torque allocation.
an inertial measurement unit, the inertial measurement unit configured to: detect at least an aircraft angle; and detect at least an aircraft angle rate;
detect at least an aircraft angle; and
detect at least an aircraft angle rate;
a flight controller, the flight controller comprising: an outer loop controller, the outer loop controller configured to: receive at least an input datum; receive at least an aircraft angle from the inertial measurement unit; and generate a rate setpoint as a function of at least an input datum; an inner loop controller, the inner loop controller configured to: receive at least an aircraft angle rate; receive the rate setpoint from the outer loop controller; and generate a moment datum as a function of the rate setpoint; and a mixer, the mixer configured to: receive the moment datum; perform a torque allocation as a function of the moment datum; and generate at least a motor command datum as a function of the torque allocation.
an outer loop controller, the outer loop controller configured to: receive at least an input datum; receive at least an aircraft angle from the inertial measurement unit; and generate a rate setpoint as a function of at least an input datum;
receive at least an input datum;
receive at least an aircraft angle from the inertial measurement unit; and
generate a rate setpoint as a function of at least an input datum;
an inner loop controller, the inner loop controller configured to: receive at least an aircraft angle rate; receive the rate setpoint from the outer loop controller; and generate a moment datum as a function of the rate setpoint; and
receive at least an aircraft angle rate;
receive the rate setpoint from the outer loop controller; and
generate a moment datum as a function of the rate setpoint; and
a mixer, the mixer configured to: receive the moment datum; perform a torque allocation as a function of the moment datum; and generate at least a motor command datum as a function of the torque allocation.
receive the moment datum;
perform a torque allocation as a function of the moment datum; and
generate at least a motor command datum as a function of the torque allocation.
The system of claim 1 , wherein the system further comprises a remote device, wherein the remote device is configured to transmit the at least an input datum to the flight controller.
The system of claim 2 , wherein the remote device is a controller.
The system of claim 2 , wherein the remote device is a remote pilot control.
The system of claim 1 , wherein the flight controller further comprises a processor.
The system of claim 1 , wherein the mixer further comprises a logic circuit.
The system of claim 1 , wherein the mixer further comprises an inertia compensator.
The system of claim 1 , wherein the inner loop controller further comprises a lead-lag filter.
The system of claim 1 , wherein the inner loop controller further comprises an integrator.
The system of claim 1 , wherein the motor torque command is transmitted to a plurality of flight components.
A method of flight control configured for use in electric aircraft, the method comprising: detecting, at the inertial measurement unit, at least an aircraft angle; detecting, at the inertial measurement unit, at least an aircraft angle rate; receiving, at a flight controller, at least an input datum; receiving, at the flight controller, at least an aircraft angle from the inertial measurement unit; generating, at the flight controller, a rate setpoint as a function of at least an input datum; receiving, at the flight controller, at least an aircraft angle rate from the inertial measurement unit; generating, at the flight controller, a moment datum as a function of the rate setpoint; receiving, at a mixer, the moment datum; perform, at the mixer, a torque allocation as a function of the moment datum; and generating, at the mixer, at least a motor command datum as a function of the torque allocation.
detecting, at the inertial measurement unit, at least an aircraft angle;
detecting, at the inertial measurement unit, at least an aircraft angle rate;
receiving, at a flight controller, at least an input datum;
receiving, at the flight controller, at least an aircraft angle from the inertial measurement unit;
generating, at the flight controller, a rate setpoint as a function of at least an input datum;
receiving, at the flight controller, at least an aircraft angle rate from the inertial measurement unit;
generating, at the flight controller, a moment datum as a function of the rate setpoint;
receiving, at a mixer, the moment datum;
perform, at the mixer, a torque allocation as a function of the moment datum; and
generating, at the mixer, at least a motor command datum as a function of the torque allocation.
The method of claim 11 , wherein the method further comprises transmitting, by a remote device, the at least an input datum to the flight controller.
The method of claim 12 , wherein the remote device is a controller.
The method of claim 12 , wherein the remote device is a remote pilot control.
The method of claim 11 , wherein the flight controller is implemented using a processor.
The method of claim 11 , wherein the mixer is implemented using an electrical logic circuit.
The method of claim 11 , wherein the mixer comprises an inertia compensator.
The method of claim 11 , wherein flight controller includes an inner loop controller.
The method of claim 1 , wherein the flight controller includes an outer loops controller.
The method of claim 11 , wherein the motor torque command is transmitted to a plurality of flight components.
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 (S
[Truncated to input limit]