Powered Descent Thrust Vector Control for Accurate Spacecraft Landing
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Solution Overview
Problem
Existing spacecraft landing guidance techniques face challenges in accurately navigating to a desired target point due to initial condition errors, such as velocity and position discrepancies, and errors related to specific thrust and initial mass, which current methods struggle to effectively mitigate.
Innovation Solution
A spacecraft system that includes a thrust vector control device and a main control device, which calculates and adjusts the thrust vector based on acquired state quantities like acceleration parameters, burn time variation, and initial cross-range position and velocity errors, using polynomials to guide the spacecraft to a target point by controlling engine combustion and thrust direction.
Engineering Contradictions & Design Principles
Engineering Contradiction Analysis
1Measurement precision
If conventional guidance methods are used, then the guidance process is simple, but the guidance accuracy deteriorates due to various errors in initial conditions and spacecraft parameters
Solution Approach 1:
The invention changes the parameter representation from using raw acceleration values to using the reciprocal of acceleration (1/a). This transformation linearizes the relationship between acceleration and time, allowing polynomial fitting to accurately model the guidance trajectory despite variations in initial conditions and spacecraft parameters. The main control device calculates polynomial coefficients based on 1/a values, which compensates for errors in initial mass and specific thrust without requiring complex adaptive guidance logic.
2Measurement precision
If polynomial fitting with reciprocal acceleration is used, then the guidance accuracy improves, but the calculation complexity increases
Solution Approach 1:
The invention performs preliminary calculation of polynomial coefficients before the actual landing guidance execution. The main control device calculates the coefficients A and B of the polynomial equation 1/a = At + B based on initial measurements, and then uses these pre-calculated coefficients to determine the thrust vector throughout the descent. This preliminary action eliminates the need for complex real-time iterative calculations, significantly improving calculation efficiency while maintaining high landing accuracy.
3Reliability
If the guidance logic is designed to deal with various errors, then the error compensation improves, but the control system complexity increases
Solution Approach 1:
The invention implements a feedback mechanism where the main control device continuously monitors the actual acceleration of the spacecraft and compares it with the expected acceleration based on the polynomial model. The difference (error) is used to adjust the thrust vector in real-time, compensating for errors in initial conditions, mass, and specific thrust. This feedback approach provides robust error compensation while keeping the control logic relatively simple, as it only requires calculating the deviation from the polynomial trajectory and applying corrective thrust.
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 allows for precise guidance and correction of errors, ensuring the spacecraft lands accurately at the desired target point by effectively managing thrust and trajectory adjustments during powered descent.
Implementation Method 1
an engine (11)
Implementation Method 2
a thrust vector control device (12) configured to control a thrust vector as a direction of thrust acting on the spacecraft
Implementation Method 3
an acceleration sensor (18) configured to detect an acceleration a of the spacecraft
Implementation Method 4
making a spacecraft land on a desired point of a gravitational body
Data Source
AI summary
A spacecraft including: an engine; a thrust vector control device controlling a thrust vector as a direction of a thrust acting on the spacecraft; and a main control device configured to acquire state quantities of the spacecraft in a powered descending in which the spacecraft is guided to a target point while the engine generates the thrust, and generate a throttling command by which combustion of the engine is controlled and an operation command by which the thrust vector control device is operated. The state quantities contain a first acceleration parameter and a second acceleration parameter. The first and second acceleration parameters are calculated as coefficients A and B obtained by fitting based on acceleration of the spacecraft previously detected, supposing the following equation is satisfied between a reciprocal number 1/a of the acceleration a of the spacecraft and time t:1/a=−At+B (1).