Aircraft Attitude Determination Using Paired Accelerometer Integration
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Solution Overview
Problem
Existing emergency navigation systems for aircraft rely on methods similar to operational channels, making them vulnerable to common failure modes and costly, bulky, and prone to measurement biases, with none effectively compensating for accelerometer calibration errors.
Innovation Solution
An inertial system using paired accelerometers to determine aircraft attitude angles by successive double integration of their second derivatives, with calibration errors corrected by differential bias compensation, and long-term drifts addressed through gravimetric acceleration comparison.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Measurement precision
If conventional inertial systems use three accelerometers and three gyrometers to determine attitude angles, then measurement accuracy is improved, but equipment cost and system bulk increase significantly
Solution Approach 1:
The invention extracts and removes the gyrometers from the conventional inertial system, retaining only the accelerometers. This extraction principle reduces system complexity and cost while maintaining attitude determination capability through a different mathematical approach using paired accelerometers and double integration of their second derivatives.
Solution Approach 2:
The invention uses multiple paired accelerometers positioned at different locations to copy the functional capability of gyrometers. By measuring acceleration differences between paired sensors and performing double integration, the system replicates the attitude angle measurement function previously requiring expensive gyrometric sensors.
2Device complexity
If emergency navigation systems use the same measurement methods as operational channels, then system simplicity is improved, but reliability decreases due to common failure modes
Solution Approach 1:
Instead of using the conventional approach of Pitot probes and static pressure probes for emergency navigation, the invention inverts the approach by using accelerometers with a different measurement paradigm (double integration of second derivatives). This inverted methodology provides independent measurement principles that are not susceptible to the same failure modes as the operational anemo-barometric channel.
3Reliability
If satellite radio-navigation signals are used for attitude measurement, then independence from operational channel is improved, but system size and synchronization complexity increase
Solution Approach 1:
The invention replaces expensive and complex satellite navigation equipment (multiple antennas and synchronization systems) with simpler, more robust accelerometers. The accelerometer-based system provides independent navigation capability without requiring the bulky antenna arrays and complex synchronization infrastructure needed for satellite radio-navigation attitude determination.
4Device complexity
If accelerometers are used without calibration error compensation, then system simplicity is improved, but measurement accuracy deteriorates due to calibration errors and long-term drift
Solution Approach 1:
The invention implements feedback mechanisms to compensate for accelerometer calibration errors and long-term drift. By continuously monitoring and correcting measurement deviations, the system maintains high accuracy over time without requiring complex pre-calibration procedures or replacing the simple accelerometer-based architecture.
Applied Scientific Principles
This section explains which scientific principles are used to turn an abstract innovation direction into a practical engineering solution.
Function Achieved in This Case
This approach enables a compact, cost-effective emergency navigation system with improved reliability and accuracy by independent measurement and reduced long-term drift, enhancing aircraft safety and navigation integrity.
Implementation Method 1
determining the pitch angle θ and/or the heading angle ψ and /or the roll angle φ of said aircraft, each of said attitude angles being determined by successive double integration of their second derivative, said second derivative being determined as the difference between the acceleration measurements delivered by two paired accelerometers
Implementation Method 2
compensating for the calibration errors of said accelerometers by correcting the second derivative of the attitude angle(s) θ, ψ, φ, by a differential bias Δ x21 divided by the sum of the respective distances between the accelerometers and the center of gravity G
Implementation Method 3
correcting the long-term drifts impacting said attitude measurements by determining at least one compensation δ θ of the pitch angle and/or one compensation δφ of the roll angle, said compensations δ θ ,δ φ being determined from the comparison of the vector B , orthogonal to the attitude plane defined by the pitch θ and roll φ angles and the acceleration vector g of the aircraft subjected to terrestrial gravity
Data Source
Figure 1~2
Figure 3
AI summary
A method for determining, by an inertial system, the attitude of an aircraft (100) characterized in that it consists at least in determining the pitch angle θ and/or the heading angle ψ and/or the roll angle ϕ of said aircraft (100), each of said attitude angles being determined by successive double integration of their second derivative, said second derivative being determined as the difference between the acceleration measurements delivered by two matched accelerometers ((A1,A2),(A3,A4),(A5,A6)) divided by the sum of the respective distances ((x1,x2),(z1,z2),(y1,y2)) between said accelerometers and the center of gravity G of said inertial system, the pair of accelerometers (A1,A2) used for determining the pitch angle θ being arranged on either side of the center of gravity G along an axis x substantially coincident with the longitudinal axis of the aircraft (100), the pair of accelerometers (A5,A6) used for determining the heading angle ψ being arranged on either side of the center of gravity G along an axis y substantially coinciding with the transverse axis of the aircraft (100), the pair of accelerometers (A3,A4) used for determining the roll angle ϕ being arranged on either side of the center of gravity G along a vertical axis z perpendicular to the plane formed by the x and y axes.