Attitude and orbit control system and method for operating same
원문보기
IPC분류정보
국가/구분
United States(US) Patent
등록
국제특허분류(IPC7판)
B64G-001/36
B64G-001/10
B64G-001/24
B64G-001/28
G01C-021/02
출원번호
US-0454162
(2014-08-07)
등록번호
US-9296495
(2016-03-29)
우선권정보
DE-10 2013 108 711 (2013-08-12)
발명자
/ 주소
Hartmann, Rolf
Schoedlbauer, Dieter
Schmidt, Uwe
출원인 / 주소
JENA OPTRONIK GMBH
대리인 / 주소
Davidson, Davidson & Kappel, LLC
인용정보
피인용 횟수 :
0인용 특허 :
12
초록▼
A hybrid network of kinematic sensors of an AOCS, made up of a star sensor including an optical camera head, and a processing unit provided as the central master processing unit, and additional kinematic sensors, each made up of a sensor element and a processing unit connected to the central process
A hybrid network of kinematic sensors of an AOCS, made up of a star sensor including an optical camera head, and a processing unit provided as the central master processing unit, and additional kinematic sensors, each made up of a sensor element and a processing unit connected to the central processing unit via a first bus. An additional processing unit is equivalent to the processing unit and is a redundant central processing unit. The central processing units and—are connected via an additional bus of a spacecraft provided with the hybrid network with the aid of a central computer. The particular active central processing units-provide all kinematic sensors with a uniform time pulse via a synchronization line, and supply the central computer with hybridized kinematic measuring data formed according to a method for hybridization based on the synchronous kinematic measuring data of the star sensor and the measuring data of the other sensors.
대표청구항▼
1. An attitude and orbit control system of a spacecraft for a spacecraft, comprising: a star sensor including an optical camera head, and a processing unit provided as a central master processing unit;additional kinematic sensors, each including a sensor element and an additional processing unit, on
1. An attitude and orbit control system of a spacecraft for a spacecraft, comprising: a star sensor including an optical camera head, and a processing unit provided as a central master processing unit;additional kinematic sensors, each including a sensor element and an additional processing unit, one of the additional processing units being equivalent to the processing unit and being designated as a redundant central processing unit;a first bus connecting the additional kinematic sensors to the central master processing unit and the redundant central processing unit;a synchronization line for providing a uniform time pulse from the particular active central processing unit of the central master processing unit and the redundant central processing unit to the additional kinematic sensors; anda connection of the master central processing unit and the redundant central processing unit to a second bus of the spacecraft, via which the particular active central processing unit supplies a central computer of the spacecraft with hybridized kinematic measuring data generated according to a method for hybridization based on the synchronous kinematic measuring data of the star sensor and measuring data of the additional kinematic sensors. 2. The attitude and orbit control system as recited in claim 1 wherein the uniform time pulse is provided by the spacecraft via a synchronization line. 3. The attitude and orbit control system as recited in claim 1 wherein an arbitrary combination of gyroscope rotation rate sensors, acceleration sensors, magnetometers, solar sensors, earth sensors, or GNSS sensors is provided as the additional kinematic sensors. 4. The attitude and orbit control system as recited in claim 1 wherein the redundant additional processing unit takes over the role of the central master processing unit, and the central master processing unit of the star sensor is then designated as the redundant central processing unit. 5. The attitude and orbit control system as recited in claim 1 wherein the star sensor is formed from at least two camera heads and the central master processing unit. 6. The attitude and orbit control system as recited in claim 1 wherein at least one of the star sensor and the additional kinematic sensors are connected to the second bus of the spacecraft. 7. The attitude and orbit control system as recited in claim 1 wherein at least one of the star sensor and the additional kinematic sensors are directly connected to the central computer. 8. The attitude and orbit control system as recited in claim 1 wherein a camera head of the star sensor is detached from the master central processing unit and connected to the master processing unit according to the point-to-point principle. 9. A method for generating hybridized kinematic measuring data from synchronous position measuring data and position covariance data of at least one star sensor and from the synchronous measuring data of additional kinematic sensors, the method comprising the following processing steps: transforming coordinates of the synchronous position measuring data and position covariance data and the synchronous measuring data of additional kinematic sensors into a uniform reference system with the aid of a transformation processor, using transformation parameters, andparallel relaying of the position measuring data and position covariance data of the star sensor and the additional kinematic measuring data transformed into the uniform reference system, to a reduction of the errors and of the error correlation, to a cross-calibration, and to a non-nominal processor,generating decorrelated positions and rates having reduced covariance in the reduction unit by replacing strongly correlated components of the measuring data of a sensor with slightly correlated data of another sensor,parallel relaying of the decorrelated positions and rates to the cross-calibration and to a trend identification, and to a mode switch,identifying a long-term trend of the transformation parameters for transforming the measuring data and of all sensors into the uniform reference system in a trend identification, and storing the transformation parameters, updated according to the trend, in a configuration parameter memory, the transformation parameters being used from the parameter memory for subsequent transformations,generating hybrid calibration data in the cross-calibration from the measuring data transformed into the uniform reference system, and from the decorrelated positions and rates,generating non-nominal positions and rates in the non-nominal processor, the data of sensors having non-nominal status not being used, and nominally operating sensors being calibrated with the aid of the hybrid calibration data, and relaying the non-nominal positions and rates to the mode switch,selecting the decorrelated positions and rates, or of the non-nominal positions and rates, as a function of the nominal or non-nominal status of the sensors, and relaying the selected position data and rate data to a bank of six-dimensional EKF estimators defined by configurable maneuver parameters, and to an adaptive multiple model estimator controlling the bank,determining probabilities of the validity of maneuvers in the adaptive multiple model estimator, based on the selected position data and rate data and maneuver parameters from the configuration parameter memory, and relaying the probabilities of the validity of the maneuvers to a maneuver switch, andselecting a result of the six-dimensional EKF estimator having the highest probabilities of the validity of the maneuver in the maneuver switch, and relaying the selected optimally hybridized kinematic measuring data, made up of the optimal estimations for the position, the rate, and the position covariance, to the central computer of the spacecraft. 10. The method as recited in claim 9 wherein in addition to the data of the star sensor, synchronous rate measuring data from a gyroscope are used and transformed into rates in the uniform reference system. 11. The method as recited in claim 9 wherein the reduction of the errors and of the error correlation is implemented as rate integration and maximum likelihood fit between integrated rates and the transformed data of the star sensor. 12. The method as recited in claim 9 wherein the cross-calibration is carried out in the form of low pass filtering and conversion to rate-dependent bias values, with subsequent separation from the bias and the scaling factor of the gyroscope sensor. 13. The method as recited in claim 9 wherein in the event of malfunction of the star sensors, the non-nominal processor uses the values estimated in separation of the bias and the scaling factor to determine the non-nominal positions and rates solely from the rates in the uniform reference system. 14. The method as recited in claim 9 wherein the trend identification is carried out in the form of an estimation of the torsion axis of the satellite and an estimation of a delta rotation matrix of the distortion, and updating of the trend of the transformation parameters takes place by applying the delta rotation.
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이 특허에 인용된 특허 (12)
Didinsky Garry ; Wu Yeong-Wei ; Li Rongsheng ; Nayak Arunkumar ; Hein Douglas, Autonomous attitude acquisition for a stellar inertial attitude determination system.
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