Office Action 1: Notice of Allowance (2025-05-07)

CLAIMS:
CLAIMS
1\. A system for detecting the position and/or the operating state of a movable component of an aircraft, wherein the system has one movable component, two actuators, two motion sensors, one asymmetry sensor, one drive train, and one monitoring device; wherein the component can be actuated by the at least two mutually offset actuators; wherein at least one motion sensor is provided for one respective actuator; wherein the motion sensor can monitor a mechanical movement of the drive train; and wherein the monitoring device is connected to the motion sensors and to the asymmetry sensor, wherein the system is configured such that an initialization of the monitoring device and/or of the motion sensors and/or of the asymmetry sensor can take place after a booting of the system and/or when the component has adopted a fully retracted and/or a fully extended position.
2\. A system in accordance with claim 1, wherein the monitoring device is configured such that a limit value and/or a limiting curve are calculated by means of the asymmetry sensor while taking account of correction values for kinematic conditions and/or tolerances and/or operational conditions, with the limit value and/or the limiting curve being able to be used for the recognition of skew of the component and/or for the recognition of the disconnect of the component from an actuator.
3\. A system in accordance with claim 1, wherein the monitoring device is configured such that the initialization takes place by an initialization angle of the component calculated by means of the asymmetry sensor and a correlation curve.
4\. A system in accordance with claim 3, wherein the monitoring device is configured such that an angle of the component takes place one the basis of the correlation curve by adding or subtracting a converted differential flap angle to/from the initialization angle.
5\. A system in accordance with claim 3, wherein the monitoring device is configured such that an absolute difference of the angles of the component from a limit value can be compared, with an exceeding of the limit value being able to result in a switching off of the system and a setting of a brake.
6\. A system in accordance with claim 1, wherein the monitoring device is configured such that a value of the asymmetry sensor can be converted into a theoretical angle of the component.
7\. A system in accordance with claim 6, wherein the monitoring device is configured such that a difference between an angle of the component and the theoretical angle of the component can be formed, with the difference from a limit value being able to be compared, with an exceeding of the limit value being able to result in a report.
8\. A system in accordance with claim 1, wherein the monitoring device is configured such that an isolation of a cause of skew of the component and/or of a disconnect of the component from an actuator can take place.
9\. A system in accordance with claim 1, wherein the component is an actuable lift aid.
10\. A system in accordance with claim 1, wherein the motion sensors and/or the asymmetry sensor is/are an incremental sensor.
11\. An aircraft, having a system in accordance with claim 1.
12\. A method of detecting the position and/or the operating state of a movable component of an aircraft having a system in accordance with claim 1, wherein an initialization of the monitoring device and/or of the motion sensors and/or of the asymmetry sensor takes place after a booting of the system and/or when the component has adopted a fully retracted and/or a fully extended position.

OFFICE ACTION:
DETAILED ACTION
Claim Status
1.	The present application, filed on or after March 16, 2013, is being examined under the first inventor to file provisions of the AIA .
Examiner’s amendment/Comment
2.	Claim 1-12 are allowed.
Allowable Subject Matter
3.	The following is a statement of reasons for the indication of allowable subject matter:  The following is an examiner’s statement of reasons for allowance: the related art of record fails to disclose or suggest a combination of a plurality of system for detecting position and/or operating state of a movable component of an aircraft (claimed). Uses include but are not limited to a state in, which an aircraft is on ground or in flight and/or a direction of travel of system, fully retracted and/or fully extended position of component and/or a flight speed, angle of attack, load status, outside temperature, and/or altitude of aircraft.  The system has high robustness in operation, and improves recognition of disconnect faults, where simple failure isolation can take place after the occurrence of skew and/or a disconnect. The different kinematic sensor connections within a flap or drive system are made possible, where the use of less expensive incremental sensors can take place. The mechanical to electrical rigging can be bypassed, so that the maintenance effort on the replacement of a sensor or the monitoring device can be reduced. The resolution of the system is increased, and a recognizability of a possible disconnect fault is increased as the recognized movement difference is larger. The system reduces maintenance effort on an replacement of a sensor and/or computer, ensures faster replacement of affected units due to failure isolation, better robustness of the skew detection, robustness and/or resolution of the disconnect detection and use of different kinematically linked sensors within a flap for skew/disconnect detection implementation possible with incremental sensors.  The system has a movable component actuated by two mutually offset actuators. A motion sensor is provided for respective actuator, and monitors a mechanical movement of a drive train. A monitoring device is connected to the motion sensors and the asymmetry sensor. The system is configured such that an initialization of the monitoring device and/or the motion sensors and/or of the asymmetry sensor takes place after a booting of the system and/or when the component has adopted a fully retracted and/or a fully extended position. The component is an actuable lift aid such as slat or a landing flap. The motion sensors and/or the asymmetry sensor are selected as an incremental sensor.  Moreover, the related art indicates that this particular device and method is novel and has not been published or patented by other entities.  This along with the rest of the claimed limitations is not shown by the related art.
Consider claim 1, the best reference found during the process of examination,Ricci Moretti (US 12110098 B1), discloses an electromechanical boost actuator for aircraft flight control is attached to a mechanical input by a mechanical linkage. The mechanical linkage is also attached to the output shaft of the electromechanical boost actuator. When a pilot or control system manually moves the mechanical input, a position sensor attached to the linkage activates electric motor(s), extending or retracting the output shaft in the direction commanded by the pilot or control system. The motion of the output shaft, acting through the linkage, moves the position sensor to a neutral position, stopping the electric motor(s) and the output shaft when the output shaft reaches the commanded position. Additional linear actuators may inject another control signal in conjunction with the mechanical input.
Consider claim 1, another best reference found during the process of examination,Richter (US 11810478 B1), discloses a system and method for an electric aircraft coaching simulator is illustrated. The simulator includes a pilot control that is configured to generate a pilot command. The simulator includes a processor that is configured to receive optimized flight data, simulate a battery performance of an electric aircraft as a function of the received pilot command and generate an optimal flight recommendation based off the optimized flight data. The simulation model is configured to communicatively connect with the flight simulator to mimic a real flight situation. Optimized flight data from other pilots can be used to coach pilots how to fly with more energy efficiency.
Claims 2-11 depend from claim 1.  Therefore, these claims are patentable over in view of related arts for the reasons given with respect to claims 1.
Citation of pertinent Prior Arts
4.	The prior art made of record and not relied upon is considered pertinent to applicant's disclosure: see PTO-892 Notice of References Cited.
Conclusion
5.	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.”



CITED REFERENCE SUMMARIES: Ref US20240359821A1: Claims:
System for monitoring and/or detecting an operational state of a movable component of an aircraft, wherein the system has a movable component and a monitoring device, wherein the system also has a combination sensor connected to the monitoring device, which is designed and arranged to detect a kinematic and/or kinetic state of the component, wherein the monitoring device is designed to monitor and/or detect an operational state of the component.
System according to claim 1 , wherein the operational state comprises or is a skew of the component and/or a disconnect of the component from a structure.
System according to claim 1 , wherein the monitoring device is designed to monitor the functionality of the combination sensor.
System according to claim 3 , wherein the monitoring device is designed such that the functionality can be monitored by comparing a kinematic state of the component detected by the combination sensor with a kinetic state of the component detected by the combination sensor.
System according to claim 1 , wherein the combination sensor comprises a kinematic sensor and a kinetic sensor.
System according to claim 5 , wherein the kinematic sensor is a discrete and/or incremental sensor and/or wherein the kinetic sensor is designed to detect a direction of a kinetic state of the component.
System according to claim 1 , wherein the system comprises an actuator, wherein the combination sensor is arranged in or on the actuator and/or wherein there are no combination sensors and/or sensors arranged in the component.
System according to claim 1 , wherein the component comprises or consists of a lift aid.
System according to claim 1 , wherein the system comprises a further combination sensor, wherein the monitoring device is designed such that monitoring and/or detecting the operational state of the component and/or monitoring the functionality of the combination sensor and/or the further combination sensor can be performed by comparing a kinematic and/or kinetic state of the component detected by the combination sensor with a kinematic and/or kinetic state of the component detected by the further combination sensor.
System according to claim 9 , wherein the comparison of a kinematic or kinetic state of the component detected by the combination sensor with a further kinematic or kinetic state of the component detected by the further combination sensor can be performed by converting the kinematic or kinetic state into a kinematic or kinetic state.
Aircraft, having a system according to claim 1 .
Method for monitoring and/or detecting an operational state of a movable component of an aircraft having a system according to claim 1 , wherein the method comprises the following steps: detecting a kinematic and/or kinetic state of a component using a combination sensor, monitoring and/or detecting an operational state of the component.
detecting a kinematic and/or kinetic state of a component using a combination sensor,
monitoring and/or detecting an operational state of the component.
Method according to claim 12 , wherein the operational state comprises or is a skew of the component and/or a disconnect of the component from a structure.
Method according to claim 12 , wherein a disconnect of the component that takes place in front of an area in which the combination sensor is arranged, takes place by detecting a kinematic and kinetic state of the component and a disconnect of the component that takes place behind an area in which the combination sensor is arranged, takes place only by detecting a kinetic state of the component.
Method according to claim 12 , wherein monitoring and/or detecting an operational state of the component is performed by detecting a kinematic state of the component after a kinetic state of the component reaches or exceeds a limit value and/or in that monitoring and/or detecting an operational state of the component is performed by detecting a kinetic state of the component after a kinematic state of the component reaches or exceeds a limit value.
System according to claim, wherein the kinematic sensor and the kinetic sensor share an element.
System according to claim 16 , wherein the kinematic sensor and the kinetic sensor share exactly, a measuring coil.
System according to claim 17 , wherein the kinematic sensor and the kinetic sensor have a common housing at least in certain areas.
System according to claim 18 , where the certain areas include a leading-edge slat or a landing flap, an actuator and/or a drive train.
Ref US12110098B1: Claims:
A control apparatus for a boost actuator having an output shaft, the control apparatus comprising: a) an output shaft bar, the output shaft bar having an output shaft bar first end and an output shaft bar second end, the output shaft bar first end being rotatably attached to the output shaft of the boost actuator, the output shaft having a longitudinal axis, the boost actuator having at least one electric motor having a configuration to selectably move the output shaft between a retracted position, an extended position and an intermediate position along the output shaft longitudinal axis; b) a floating bar having a floating bar first end and a floating bar second end, the output shaft bar second end being rotatably attached to the floating bar first end; c) a mechanical input, the mechanical input having a configuration for attachment to the floating bar and to selectably move the floating bar second end in a first or a second direction, the first and second directions being generally parallel to the output shaft longitudinal axis, the first and second directions being substantially opposing; d) a pivoting bar, the pivoting bar having a pivoting bar first end and a pivoting bar second end, the pivoting bar first end being attached to and configured for rotation about a fixed pivot location, the fixed pivot location being stationary with respect to the boost actuator, the pivoting bar second end being rotatably attached to the floating bar in a spaced-apart relation to the floating bar first end; e) a position detector, the position detector being configured to detect an angular position of the pivoting about the fixed pivot location; f) a motor controller, the position detector being operably attached to the motor controller, the motor controller being operably attached to the at least one electric motor.
a) an output shaft bar, the output shaft bar having an output shaft bar first end and an output shaft bar second end, the output shaft bar first end being rotatably attached to the output shaft of the boost actuator, the output shaft having a longitudinal axis, the boost actuator having at least one electric motor having a configuration to selectably move the output shaft between a retracted position, an extended position and an intermediate position along the output shaft longitudinal axis;
b) a floating bar having a floating bar first end and a floating bar second end, the output shaft bar second end being rotatably attached to the floating bar first end;
c) a mechanical input, the mechanical input having a configuration for attachment to the floating bar and to selectably move the floating bar second end in a first or a second direction, the first and second directions being generally parallel to the output shaft longitudinal axis, the first and second directions being substantially opposing;
d) a pivoting bar, the pivoting bar having a pivoting bar first end and a pivoting bar second end, the pivoting bar first end being attached to and configured for rotation about a fixed pivot location, the fixed pivot location being stationary with respect to the boost actuator, the pivoting bar second end being rotatably attached to the floating bar in a spaced-apart relation to the floating bar first end;
e) a position detector, the position detector being configured to detect an angular position of the pivoting about the fixed pivot location;
f) a motor controller, the position detector being operably attached to the motor controller, the motor controller being operably attached to the at least one electric motor.
The control apparatus of claim 1 wherein the configuration of the electric motor to selectably move the output shaft comprising: the pivoting bar having a neutral location in which the motor controller does not cause the at least one electric motor to move the output shaft, the position sensor being configured to detect whether the pivoting bar is in the neutral location.
The control apparatus claim 2 wherein the configuration of the electric motor to selectably move the output shaft further comprising: the pivoting bar having a first displaced position in which the pivoting bar is displaced in the first direction from the neutral position, the pivoting bar having a second displaced position in which the pivoting bar is displaced in the second direction from the neutral position, the motor controller being configured to cause the motor to move the output shaft in the second direction when the pivoting bar is in the first displaced position, the motor controller being configured to cause the motor to move the output shaft in the first direction when the pivoting bar is in the second displaced position.
The control apparatus of claim 3 in which each of the following rotatable connections is a pin connection having a pin axis of rotation and in which all of the pin axes of rotation are substantially parallel: a) between the output shaft bar first end to the output shaft; b) between the output shaft bar second end to the floating bar first end; c) between the floating bar and the pivoting bar second end; and d) between the pivoting bar first end and the fixed pivot location is a first pin connection.
a) between the output shaft bar first end to the output shaft;
b) between the output shaft bar second end to the floating bar first end;
c) between the floating bar and the pivoting bar second end; and
d) between the pivoting bar first end and the fixed pivot location is a first pin connection.
The control apparatus of claim 1 wherein the mechanical input comprises a mechanical linkage or a mechanical cable.
The control apparatus of claim 1 wherein the at least one electric motor comprises a pair of redundant electric motors.
The control apparatus of claim 1 wherein the apparatus further comprises: a redundant mechanical load path between the output shaft and the mechanical input.
The control apparatus of claim 7 wherein the redundant mechanical load path comprises: a first output shaft bar and a second output shaft bar.
The control apparatus of claim 8 wherein the first and the second output shafts are separate.
The control apparatus of claim 9 wherein the first output shaft bar comprises: a first linear actuator, the first linear actuator being configured to selectably change a distance between the first output shaft bar first end and the first output shaft bar second end.
The control apparatus of claim 10 wherein the second output shaft bar comprises: a second linear actuator, the second linear actuator being configured to selectably change a distance between the second output shaft bar first end and the second output shaft bar second end.
The control apparatus of claim 1 wherein the output shaft bar comprises: a linear actuator, the linear actuator being configured to selectably change a distance between the output shaft bar first end and the output shaft bar second end.
Ref US20240084742A1: Claims:
An aircraft comprising: a multi-engine configuration with multiple engines; and a flight control system coupled to the multiple engines with a multi-engine interface, wherein the flight control system is configured to reduce fuel consumption of at least one engine of the multiple engines during reduced-engine operation by pulse modulating fuel delivery to the at least one engine of the multiple engines.
a multi-engine configuration with multiple engines; and
a flight control system coupled to the multiple engines with a multi-engine interface,
wherein the flight control system is configured to reduce fuel consumption of at least one engine of the multiple engines during reduced-engine operation by pulse modulating fuel delivery to the at least one engine of the multiple engines.
The aircraft of claim 1 , wherein: upon selective activation of the reduced-engine operation by a pilot or automated means, the flight control system is configured to adjust operation of the at least one engine of the multiple engines by throttling the at least one engine to IDLE or sub-IDLE.
upon selective activation of the reduced-engine operation by a pilot or automated means, the flight control system is configured to adjust operation of the at least one engine of the multiple engines by throttling the at least one engine to IDLE or sub-IDLE.
The aircraft of claim 2 , wherein: upon selective activation of the reduced-engine operation by the pilot or automated means, the flight control system is configured to maintain IDLE or sub-IDLE via oscillating fuel flow to the at least one engine during reduced-engine operation.
upon selective activation of the reduced-engine operation by the pilot or automated means, the flight control system is configured to maintain IDLE or sub-IDLE via oscillating fuel flow to the at least one engine during reduced-engine operation.
The aircraft of claim 3 , wherein: by maintaining an IDLE or sub-IDLE operation condition, the at least one engine has a reduced time required to produce and contribute additional engine power available.
by maintaining an IDLE or sub-IDLE operation condition, the at least one engine has a reduced time required to produce and contribute additional engine power available.
The aircraft of claim 1 , wherein: while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to command at least one reduced engine to produce and contribute additional engine power available and provide corresponding advisory messages to a pilot.
while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to command at least one reduced engine to produce and contribute additional engine power available and provide corresponding advisory messages to a pilot.
The aircraft of claim 1 , wherein: the flight control system is configured to selectively reduce engine power at least one other engine of the multiple engines based on input from a pilot, and the flight control system is configured to provide visual indication of the at least one other engine selectively operating with reduced engine power by the pilot.
the flight control system is configured to selectively reduce engine power at least one other engine of the multiple engines based on input from a pilot, and
the flight control system is configured to provide visual indication of the at least one other engine selectively operating with reduced engine power by the pilot.
The aircraft of claim 6 , wherein: upon selective reduction in engine power at least one other engine by the pilot or by automated means, the flight control system is configured to adjust operation of the at least one other engine of the multiple engines by throttling the at least one other engine to IDLE, sub-IDLE, or completely OFF.
upon selective reduction in engine power at least one other engine by the pilot or by automated means, the flight control system is configured to adjust operation of the at least one other engine of the multiple engines by throttling the at least one other engine to IDLE, sub-IDLE, or completely OFF.
The aircraft of claim 6 , wherein: while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to automatically restart the at least one other engine and provide corresponding advisory messages to the pilot.
while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to automatically restart the at least one other engine and provide corresponding advisory messages to the pilot.
The aircraft of claim 6 , wherein the flight control system is configured to selectively reduce engine power of the at least one other engine of the multiple engines by continuously calculating altitude for the reduced-engine operation based on an aircraft descent rate of the at least one other engine.
The aircraft of claim 9 , wherein: the aircraft descent rate is based on current ambient conditions related to the altitude including outside ambient temperature during autorotation of the aircraft with the at least one engine of the multiple engines operating with reduced engine power.
the aircraft descent rate is based on current ambient conditions related to the altitude including outside ambient temperature during autorotation of the aircraft with the at least one engine of the multiple engines operating with reduced engine power.
The aircraft of claim 1 , wherein: the aircraft comprises a rotorcraft, and the multiple engines comprise multiple turbo-shaft engines.
the aircraft comprises a rotorcraft, and
the multiple engines comprise multiple turbo-shaft engines.
An aircraft comprising: a multi-engine configuration with multiple engines of three or more engines; and a flight control system coupled to the multiple engines with a multi-engine interface, wherein the flight control system is configured to shutdown at least one engine of the multiple engines during reduced engine operation, and wherein the flight control system is configured to reduce fuel consumption of at least one other engine of the multiple engines during the reduced engine operation by pulse modulating fuel delivery to the at least one other engine of the multiple engines.
a multi-engine configuration with multiple engines of three or more engines; and
wherein the flight control system is configured to shutdown at least one engine of the multiple engines during reduced engine operation, and
wherein the flight control system is configured to reduce fuel consumption of at least one other engine of the multiple engines during the reduced engine operation by pulse modulating fuel delivery to the at least one other engine of the multiple engines.
The aircraft of claim 12 , wherein: upon selective activation of the reduced-engine operation by a pilot or automated means, the flight control system is configured to adjust operation of the at least one other engine by throttling the at least one other engine to IDLE or sub-IDLE.
upon selective activation of the reduced-engine operation by a pilot or automated means, the flight control system is configured to adjust operation of the at least one other engine by throttling the at least one other engine to IDLE or sub-IDLE.
The aircraft of claim 13 , wherein: upon selective activation of the reduced-engine operation by the pilot or automated means, the flight control system is configured to maintain IDLE or sub-IDLE via oscillating fuel flow to the at least one other engine during reduced-engine operation.
upon selective activation of the reduced-engine operation by the pilot or automated means, the flight control system is configured to maintain IDLE or sub-IDLE via oscillating fuel flow to the at least one other engine during reduced-engine operation.
The aircraft of claim 12 , wherein: while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to increase fuel delivery to the at least one other engine and provide corresponding advisory messages to a pilot.
while in the reduced-engine operation during flight, if altitude drops significantly, then the flight control system is configured to increase fuel delivery to the at least one other engine and provide corresponding advisory messages to a pilot.
The aircraft of claim 12 , wherein: wherein the 
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