[특허]Attitude determination system for yaw-steering spacecraft

Lockheed Martin Corporation

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내보내기 페이스북 공유 트위터 공유 카카오톡 공유

Attitude determination system for yaw-steering spacecraft

특허상세정보
국가/구분	United States(US) Patent 등록
국제특허분류(IPC7판)	B64G-001/24
미국특허분류(USC)	701/004; 244/164
출원번호	UP-0485974 (2006-07-14)
등록번호	US-7835826 (2011-01-16)
발명자 / 주소	Weigl, Harald J. Goodzeit, Neil E. Ratan, Santosh
출원인 / 주소	Lockheed Martin Corporation
대리인 / 주소	McDermott Will & Emery LLP
인용정보	피인용 횟수 : 8 인용 특허 : 85

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An attitude determination system is provided for determining attitude values of a yaw-steering spacecraft, and includes a rate sensor, at least one attitude sensor, and a processor operable to receive measured spacecraft body rates from the rate sensor, to receive measured spacecraft attitude values from the at least one attitude sensor, and to calculate estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error ...

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What is claimed is: 1. An attitude determination system for determining attitude values of a yaw-steering spacecraft, the attitude determination system comprising: a rate sensor; at least one attitude sensor; and a processor operable to receive measured spacecraft body rates from the rate sensor, to receive measured spacecraft attitude values from the at least one attitude sensor, and to calculate estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error states, a plurality of gyro bias states, and a plurality of commanded rate dependent gyro error states. 2. The attitude determination system of claim 1, wherein, in the Kalman filter, the plurality of attitude estimate error states comprises a pitch attitude estimate error state, a roll attitude estimate error state and a yaw attitude estimate error state, and the plurality of gyro bias states including a pitch rate gyro bias state, a roll rate gyro bias state and a yaw rate gyro bias state. 3. The attitude determination system of claim 2, wherein, in the Kalman filter, the plurality of commanded rate dependent gyro error states correspond to a set of nine time-dependant error coefficients, each error coefficient representing a relationship between a spacecraft body rate component and a spacecraft attitude component. 4. The attitude determination system of claim 3, wherein, in the Kalman filter, derivative functions of the attitude estimate error states, the gyro bias states, and the commanded rate dependent gyro error states are calculated according to: [ e . ⁡ ( t ) ω . bias ⁡ ( t ) k . ⁡ ( t ) ] = [ W ⁡ ( t ) I 3 Ω ⁡ ( t ) 0 3 × 3 0 3 × 3 0 3 × 9 0 9 × 3 0 9 × 3 0 9 × 9 ] ⁡ [ e ⁡ ( t ) ω bias ⁡ ( t ) k ⁡ ( t ) ] + [ η 1 ⁡ ( t ) η 2 ⁡ ( t ) η 3 ⁡ ( t ) ] in which ė(t) is the derivative function of the attitude estimate error states e(t), {dot over (ω)}bias(t) is a derivative function of the gyro bias states ωbias(t), {dot over (k)}(t) is a derivative function of the commanded rate dependent gyro error states k(t) according to: k ⁡ ( t ) = [ k 11 ⁡ ( t ) k 12 ⁡ ( t ) k 13 ⁡ ( t ) k 21 ⁡ ( t ) k 22 ⁡ ( t ) k 23 ⁡ ( t ) k 31 ⁡ ( t ) k 32 ⁡ ( t ) k 33 ⁡ ( t ) ] in which k(t) is comprised of the nine time-dependant error coefficients, and in which W(t) is a matrix of spacecraft body rate components of (ω(t)) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which I3 is a 3×3 identity matrix, Ω(t) is a matrix of commanded spacecraft body rate components according to: Ω ⁡ ( t ) = [ ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c ] in which ωc=[ωcx(t), ωcy(t), ωcz(t)], and in which η1(t), η2(t) and η3(t) are noise functions corresponding to the attitude estimate error states e(t), the gyro bias states ωbias(t), and the commanded rate dependent gyro error states k(t). 5. The attitude determination system of claim 4, wherein the commanded spacecraft body rate components included in the matrix Ω(t) are output from a commanded rate generator on the spacecraft. 6. The attitude determination system of claim 1, wherein the calculation of the estimated spacecraft attitude values is based on an attitude estimate error model according to: {dot over (e)}(t)=W(t)e(t)+ωbias+Kωtrue(t) in which K is an error coefficient matrix according to: K = [ k 11 k 12 k 13 k 21 k 22 k 23 k 31 k 32 k 33 ] and in which ė(t) is the derivative function of attitude estimate error e(t), W is a matrix of spacecraft body rates (ω) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which ωbias is a set of gyro rate bias components, and ωtrue(t) is a time-based set of actual spacecraft body rate components. 7. The attitude determination system of claim 1, wherein the rate sensor is an inertial measurement unit that outputs the measured spacecraft body rates. 8. The attitude determination system of claim 1, wherein the at least one attitude sensor is at least one of an earth sensor assembly, a star tracker assembly and a sun sensor assembly. 9. A method for determining attitude values of a yaw-steering spacecraft, the method comprising the steps of: receiving measured spacecraft body rates a rate sensor; receiving measured spacecraft attitude values from the at least one attitude sensor; and calculating estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error states, a plurality of gyro bias states, and a plurality of commanded rate dependent gyro error states. 10. The method according to claim 9, wherein, in the Kalman filter, the plurality of attitude estimate error states comprises a pitch attitude estimate error state, a roll attitude estimate error state and a yaw attitude estimate error state, and the plurality of gyro bias states including a pitch rate gyro bias state, a roll rate gyro bias state and a yaw rate gyro bias state. 11. The method according to claim 10, wherein, in the Kalman filter, the plurality of commanded rate dependent gyro error states correspond to a set of nine time-dependant error coefficients, each error coefficient representing a relationship between a spacecraft body rate component and a spacecraft attitude component. 12. The method according to claim 11, wherein, in the Kalman filter, derivative functions of the attitude estimate error states, the gyro bias states, and the commanded rate dependent gyro error states are calculated according to: [ e . ⁡ ( t ) ω . bias ⁡ ( t ) k . ⁡ ( t ) ] = [ w ⁡ ( t ) I 3 Ω ⁡ ( t ) 0 3 ⁢ x ⁢ ⁢ 3 0 3 ⁢ x ⁢ ⁢ 3 0 3 ⁢ x ⁢ ⁢ 9 0 9 ⁢ x ⁢ ⁢ 3 0 9 ⁢ x ⁢ ⁢ 3 0 9 ⁢ x ⁢ ⁢ 9 ] ⁡ [ e ⁡ ( t ) ω bias ⁡ ( t ) k ⁡ ( t ) ] + [ η 1 ⁡ ( t ) η 2 ⁡ ( t ) η 3 ⁡ ( t ) ] in which ė(t) is the derivative function of the attitude estimate error states e(t), {dot over (ω)}bias(t) is a derivative function of the gyro bias states ωbias(t), {dot over (k)}(t) is a derivative function of the commanded rate dependent gyro error states k(t) according to: k ⁡ ( t ) = [ k 11 ⁡ ( t ) k 12 ⁡ ( t ) k 13 ⁡ ( t ) k 21 ⁡ ( t ) k 22 ⁡ ( t ) k 23 ⁡ ( t ) k 31 ⁡ ( t ) k 32 ⁡ ( t ) k 33 ⁡ ( t ) ] in which k(t) is comprised of the nine time-dependant error coefficients, and in which W(t) is a matrix of spacecraft body rate components of (ω(t)) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which I3 is a 3×3 identity matrix, Ω(t) is a matrix of commanded spacecraft body rate components according to: Ω ⁡ ( t ) = [ ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c ] in which ωc=[ωcx(t), ωcy(t), ωcz(t)], and in which η1(t), η2(t) and η3(t) are noise functions corresponding to the attitude estimate error states e(t), the gyro bias states ωbias(t), and the commanded rate dependent gyro error states k(t). 13. The method according to claim 12, wherein the commanded spacecraft body rate components included in the matrix Ω(t) are output from a commanded rate generator on the spacecraft. 14. The method according to claim 9, wherein the calculation of the estimated spacecraft attitude values is based on an attitude estimate error model according to: {dot over (e)}(t)=W(t)e(t)+ωbias+Kωtrue(t) in which K is an error coefficient matrix according to: K = [ k 11 k 12 k 13 k 21 k 22 k 23 k 31 k 32 k 33 ] and in which ė(t) is the derivative function of attitude estimate error e(t), W is a matrix of spacecraft body rates (ω) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which ωbias is a set of gyro rate bias components, and ωtrue(t) is a time-based set of actual spacecraft body rate components. 15. The method according to claim 9, wherein the rate sensor is an inertial measurement unit that outputs the measured spacecraft body rates. 16. The method according to claim 9, wherein the at least one attitude sensor is at least one of an earth sensor assembly, a star tracker assembly and a sun sensor assembly. 17. A spacecraft comprising: a rate sensor; at least one attitude sensor; and a processor operable to receive measured spacecraft body rates from the rate sensor, to receive measured spacecraft attitude values from the at least one attitude sensor, and to calculate estimated spacecraft attitude values based on the measured spacecraft body rates and the measured spacecraft attitude values by using a Kalman filter that includes a plurality of attitude estimate error states, a plurality of gyro bias states, and a plurality of commanded rate dependent gyro error states. 18. The spacecraft of claim 17, wherein, in the Kalman filter, the plurality of attitude estimate error states comprises a pitch attitude estimate error state, a roll attitude estimate error state and a yaw attitude estimate error state, and the plurality of gyro bias states including a pitch rate gyro bias state, a roll rate gyro bias state and a yaw rate gyro bias state. 19. The spacecraft of claim 18, wherein, in the Kalman filter, the plurality of commanded rate dependent gyro error states correspond to a set of nine time-dependant error coefficients, each error coefficient representing a relationship between a spacecraft body rate component and a spacecraft attitude component. 20. The spacecraft of claim 19, wherein, in the Kalman filter, derivative functions of the attitude estimate error states, the gyro bias states, and the commanded rate dependent gyro error states are calculated according to: [ e . ⁡ ( t ) ω . bias ⁡ ( t ) k . ⁡ ( t ) ] = [ w ⁡ ( t ) I 3 Ω ⁡ ( t ) 0 3 ⁢ x ⁢ ⁢ 3 0 3 ⁢ x ⁢ ⁢ 3 0 3 ⁢ x ⁢ ⁢ 9 0 9 ⁢ x ⁢ ⁢ 3 0 9 ⁢ x ⁢ ⁢ 3 0 9 ⁢ x ⁢ ⁢ 9 ] ⁡ [ e ⁡ ( t ) ω bias ⁡ ( t ) k ⁡ ( t ) ] + [ η 1 ⁡ ( t ) η 2 ⁡ ( t ) η 3 ⁡ ( t ) ] in which ė(t) is the derivative function of the attitude estimate error states e(t), {dot over (ω)}bias(t) is a derivative function of the gyro bias states ωbias(t), {dot over (k)}(t) is a derivative function of the commanded rate dependent gyro error states k(t) according to: k ⁡ ( t ) = [ k 11 ⁡ ( t ) k 12 ⁡ ( t ) k 13 ⁡ ( t ) k 21 ⁡ ( t ) k 22 ⁡ ( t ) k 23 ⁡ ( t ) k 31 ⁡ ( t ) k 32 ⁡ ( t ) k 33 ⁡ ( t ) ] in which k(t) is comprised of the nine time-dependant error coefficients, and in which W(t) is a matrix of spacecraft body rate components of (ω(t)) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which I3 is a 3×3 identity matrix, Ω(t) is a matrix of commanded spacecraft body rate components according to: Ω ⁡ ( t ) = [ ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 0 1 ⁢ x ⁢ ⁢ 3 ω c ] in which ωc=[ωcx(t), ωcy(t), ωcz(t)], and in which η1(t), η2(t) and η3(t) are noise functions corresponding to the attitude estimate error states e(t), the gyro bias states ωbias(t), and the commanded rate dependent gyro error states k(t). 21. The spacecraft of claim 20, wherein the commanded spacecraft body rate components included in the matrix Ω(t) are output from a commanded rate generator on the spacecraft. 22. The spacecraft of claim 17, wherein the calculation of the estimated spacecraft attitude values is based on an attitude estimate error model according to: {dot over (e)}(t)=W(t)e(t)+ωbias+Kωtrue(t) in which K is an error coefficient matrix according to: K = [ k 11 k 12 k 13 k 21 k 22 k 23 k 31 k 32 k 33 ] and in which ė(t) is the derivative function of attitude estimate error e(t), W is a matrix of spacecraft body rates (ω) according to: W ⁡ ( t ) = - [ 0 - ω z ⁡ ( t ) ω y ⁡ ( t ) ω z ⁡ ( t ) 0 - ω x ⁡ ( t ) - ω y ⁡ ( t ) ω x ⁡ ( t ) 0 ] and in which ωbias is a set of gyro rate bias components, and ωtrue(t) is a time-based set of actual spacecraft body rate components. 23. The spacecraft of claim 17, wherein the rate sensor is an inertial measurement unit that outputs the measured spacecraft body rates. 24. The spacecraft of claim 17, wherein the at least one attitude sensor is at least one of an earth sensor assembly, a star tracker assembly and a sun sensor assembly.

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