Distributed Pressure Sensing for Supersonic Missile Flow Parameters
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Solution Overview
Problem
Existing methods for determining flow parameters of supersonic missiles are limited by the need for sensors in critical areas like the nose or wing, which can interfere with the flow and are unsuitable for reverse flight, and lack efficiency in determining parameters during both forward and reverse flight.
Innovation Solution
The method involves arranging pressure sensors on various components of the missile, such as fins and the base body, with different surface normals, allowing for simultaneous pressure measurements that are evaluated using a database, neural network, or formulas to determine flow parameters like Mach number, angle of attack, and sideslip angle, enabling accurate determination during both forward and reverse flight.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Measurement precision
If multi-hole probes or vane sensors are positioned near the stagnation point (nose or wing front surface), then flow parameters can be measured, but the sensors interfere with the flow around the missile and increase thermal stress
Solution Approach 1:
The patent extracts the pressure measurement function from the traditional stagnation point location and redistributes it across multiple sensors positioned at different locations on the missile body. This extraction allows the sensors to be removed from the high-stress stagnation region while maintaining measurement capability through distributed pressure taps that sample flow parameters from multiple positions simultaneously.
Solution Approach 2:
The patent segments the pressure measurement function into multiple distributed pressure sensors rather than using a single multi-hole probe or vane sensor at the stagnation point. Each sensor measures pressure at its specific location, and the combined data from all sensors provides comprehensive flow parameter information while distributing the harmful effects of flow interference and thermal stress across multiple locations.
2Adaptability or versatility
If sensors are arranged on the nose or front surface to measure flow parameters, then measurements can be obtained during forward flight, but the system cannot accurately measure during backward flight
Solution Approach 1:
The patent creates a universal measurement system by positioning pressure sensors at multiple locations around the missile body rather than concentrating them only on the nose or front surface. This distributed arrangement ensures that sensors remain effective regardless of flight direction, as each sensor measures pressure at its specific location relative to the oncoming flow, enabling accurate flow parameter determination during both forward and backward flight.
3Measurement precision
If pressure sensors are positioned on the nose or wing front end, then flow parameters can be measured, but other components like radomes or nose-integrated systems are interfered with
Solution Approach 1:
The patent extracts the pressure measurement function from the nose and front surface regions where it would interfere with radomes and other nose-integrated systems. By redistributing pressure sensors to locations on the missile body away from these sensitive areas, the measurement capability is maintained while eliminating the interference with other components.
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 minimizes flow interference, reduces thermal stress, and improves the accuracy and reliability of flow parameter determination, allowing for effective flight control during all flight phases without the need for additional sensors.
Implementation Method 1
pressures are measured at different positions at the stagnation point of an aircraft's nose or wing
Data Source
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AI summary
The invention relates to a method for determining the flow parameters of a supersonic aircraft, namely the Mach number, static atmospheric pressure, global angle of attack, and sideslip angle. In the method according to the invention, pressure is measured using at least four pressure sensors located away from the nose or forward face of the aircraft and having different surface normals. The aforementioned flow parameters are then determined using these at least four pressure measurements and a database, a formula, or a neural network.