초록 ▼

The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers

The present system and methods enable simultaneous momentum dumping and orbit control of a spacecraft, such as a geostationary satellite. Control equations according to the present system and methods generate accurate station-keeping commands quickly and efficiently, reducing the number of maneuvers needed to maintain station and allowing station-keeping maneuvers to be performed with a single burn. Additional benefits include increased efficiency in propellant usage, and extension of the satellite's lifespan. The present system and methods also enable tighter orbit control, reduction in transients and number of station-keeping thrusters aboard the satellite. The present methods also eliminate the need for the thrusters to point through the center of mass of the satellite, which in turn reduces the need for dedicated station-keeping thrusters. The present methods also facilitate completely autonomous orbit control and control using Attitude Control Systems (ACS).

대표청구항 ▼

1. A method of simultaneous orbit control and momentum dumping in a spacecraft, the spacecraft including a plurality of thrusters, the method comprising the steps of: generating a first set of firing commands for the thrusters from solutions to momentum dumping and drift control equations; andfiring

1. A method of simultaneous orbit control and momentum dumping in a spacecraft, the spacecraft including a plurality of thrusters, the method comprising the steps of: generating a first set of firing commands for the thrusters from solutions to momentum dumping and drift control equations; andfiring the thrusters according to the first set of firing commands, wherein the momentum dumping and drift control equations for the first set of firing commands are defined as Δ⁢⁢H⇀=∑i⁢r⇀i⊗f⇀i⁢Δ⁢⁢ti∑i⁢fitangential⁢Δ⁢⁢ti=Δ⁢⁢PDrift;whereΔ{right arrow over (H)}=momentum dumping requirement (vector) in orbit frameΔPDrift=spacecraft mass×minimum delta velocity required to control mean Drift{right arrow over (R)}i=lever arm (vector) about the c.g. for the ith thruster in spacecraft body frame{right arrow over (F)}i=thrust vector for the ith thruster in spacecraft body frameΔti=on time for the ith thrusterCOrbit to ECI=transformation matrix from orbit to ECI frameCBody to Orbit=transformation matrix from spacecraft body to orbit frame{right arrow over (r)}i=CBody to Orbit {right arrow over (R)}i ⁢f⇀i=CBody⁢⁢to⁢⁢Orbit⁢F⇀i=[fitangentialfiradialfinormal] and wherein the method further comprises the step of: generating a second set of firing commands for the thrusters from solutions to momentum dumping/drift and eccentricity control equations, wherein the momentum dumping/drift and eccentricity control equations are defined as Ptangential=∑i⁢fitangential⁢Δ⁢⁢tiPradial=∑i⁢fitangential⁢Δ⁢⁢tiλEccentricity=tan-1⁡(2⁢⁢Ptangential⁢Δ⁢⁢PH1+Pradial⁢Δ⁢⁢VK12⁢⁢Ptangential⁢Δ⁢⁢PK1-Pradial⁢Δ⁢⁢VH1)Δ⁢⁢H⇀ECI=COrbit⁢⁢to⁢⁢ECI⁢Δ⁢⁢H⇀Δ⁢⁢H⇀=∑i⁢r⇀i⊗f⇀i⁢Δ⁢⁢ti∑i⁢fitangential⁢Δ⁢⁢ti=Δ⁢⁢PDrift; where {right arrow over (H)}=momentum dumping requirement (vector) in orbit frameΔ{right arrow over (H)}ECI=momentum dumping requirement (vector) in Earth—Centered Inertial frameΔPK1=spacecraft mass×minimum delta velocity required to control mean K1 ΔPH1=spacecraft mass×minimum delta velocity required to control mean H1 ΔPDrift=spacecraft mass×minimum delta velocity required to control mean Drift{right arrow over (R)}i=lever arm (vector) about the c.g. for the ith thruster in spacecraft body frame{right arrow over (F)}i=thrust vector for the ith thruster in spacecraft body frameΔti=on time for the ith thrusterλEccentricity=location of the maneuverCOrbit to ECI=transformation matrix from orbit to ECI frame, rotation matrix about the Z by λEccentricity CBody to Orbit=transformation matrix from spacecraft body to orbit frame{right arrow over (r)}i=CBody to Orbit{right arrow over (R)}i ⁢f⇀i=CBody⁢⁢to⁢⁢Orbit⁢F⇀i=[fitangentialfiradialfinormal]. 2. The method of claim 1, wherein the thrusters are fired according to the first and second sets of firing commands simultaneously. 3. The method of claim 1, wherein the solutions result in either under correction or over correction of the eccentricity, and wherein the method further comprises: correcting a difference in a next control cycle. 4. The method of claim 1 further comprising: firing the thrusters according to the firing commands to perform completely autonomous orbit and attitude control system control. 5. The method of claim 1 further comprising: finding a closed form solution to the momentum dumping and drift control equations by coupling momentum dumping with orbit control in specified directions. 6. The method of claim 1 further comprising: performing a momentum dump. 7. The method of claim 6 further comprising: performing the momentum dump in conjunction with drift control. 8. The method of claim 6 further comprising: performing the momentum dump in conjunction with inclination control. 9. The method of claim 6 further comprising: performing the momentum dump in conjunction with both drift control and inclination control. 10. The method of claim 1 further comprising: performing the momentum dump in conjunction with eccentricity control. 11. The method of claim 1 further comprising: performing momentum dumping and drift control independently to maintain a longitude of the spacecraft. 12. The method of claim 1 further comprising: performing station-keeping maneuvers in pulses spaced out over multiple burns, wherein an elapsed time between the pulses comprises about a second to about minutes. 13. A method of simultaneous orbit control and momentum dumping in a spacecraft, the spacecraft including a plurality of thrusters, the method comprising the steps of: generating a first set of firing commands for the thrusters from solutions to momentum dumping and drift control equations; andfiring the thrusters according to the first set of firing commands, wherein the momentum dumping and drift control equations for the first set of firing commands are defined as Δ⁢⁢H⇀=∑i⁢r⇀i⊗f⇀i⁢Δ⁢⁢ti∑i⁢fitangential⁢Δ⁢⁢ti=Δ⁢⁢PDrift;whereΔ{right arrow over (H)}=momentum dumping requirement (vector) in orbit frameΔPDrift=spacecraft mass×minimum delta velocity required to control mean Drift{right arrow over (R)}i=lever arm (vector) about the c.g. for the ith thruster in spacecraft body frame{right arrow over (F)}i=thrust vector for the ith thruster in spacecraft body frameΔti=on time for the ith thrusterCOrbit to ECI=transformation matrix from orbit to ECI frameCBody to Orbit=transformation matrix from spacecraft body to orbit frame{right arrow over (r)}i=CBody to Orbit {right arrow over (R)}i f⇀i=CBody⁢⁢to⁢⁢Orbit⁢F⇀i=[fitangentialfiradialfinormal] and wherein the method further comprises the step of: generating a second set of firing commands for the thrusters from solutions to momentum dumping/drift and eccentricity control equations wherein the momentum dumping/drift and eccentricity control equations are defined as ⁢∑j=1,2⁢⁢Pjtangential=Δ⁢⁢Pdrift⁢(2⁢⁢P1tangential⁢cos⁢⁢λ1+P1radial⁢sin⁢⁢λ1)+(2⁢⁢P2tangential⁢cos⁡(λ1-Δλ)+P2radial⁢sin⁡(λ1-Δλ))=Δ⁢⁢PK1⁢(2⁢⁢P1tangential⁢sin⁢⁢λ1-P1radial⁢cos⁢⁢λ1)+(2⁢⁢P2tangential⁢sin⁡(λ1-Δλ)-P2radial⁢cos⁡(λ1-Δλ))=Δ⁢⁢PH1⁢-2⁢⁢P1radial⁢P2radial⁢sin⁢⁢Δλ-4⁢⁢P1tangential⁢P2radial⁢cos⁢⁢Δλ-8⁢⁢P1tangential⁢P2tangential⁢sin⁢⁢Δλ+4⁢⁢P1radial⁢P2tangential⁢cos⁢⁢Δλ=0⁢λ2=λ1-Δλ⁢Δ⁢H⇀ECI=Δ⁢⁢H⇀ECI,1+Δ⁢⁢H⇀ECI,2⁢Δ⁢⁢H⇀ECI,1=COrbit⁢⁢to⁢⁢ECI⁡(λ1)⁢Δ⁢⁢H⇀1⁢Δ⁢⁢H⇀ECI,2=COrbit⁢⁢to⁢⁢ECI⁡(λ2)⁢Δ⁢⁢H⇀2⁢Pjradial=∑i⁢fi,jradial⁢Δ⁢⁢ti,j⁢Pjtangential=∑i⁢fi,jtangential⁢Δ⁢⁢ti,j⁢Δ⁢⁢H⇀j=∑i⁢r⇀i,j⊗f⇀i,j⁢Δ⁢⁢ti,j;where{right arrow over (r)}i,j=CBody to Orbit {right arrow over (R)}i,j f⇀i,j=CBody⁢⁢to⁢⁢Orbit⁢F⇀i,j=[fi,jtangentialfi,jradialfi,jnormal]j=1,2i=index⁢⁢for⁢⁢the⁢⁢ith⁢⁢thruster. 14. The method of claim 13, wherein the thrusters are fired according to the first and second sets of firing commands simultaneously. 15. A method of simultaneous orbit control and momentum dumping in a spacecraft, the spacecraft including a plurality of thrusters, the method comprising the steps of: generating a set of firing commands for the thrusters from solutions to momentum dumping/drift and eccentricity control equations; andfiring the thrusters according to the firing commands;wherein the momentum dumping/drift and eccentricity control equations are defined as Ptangential=∑i⁢fitangential⁢Δ⁢⁢tiPradial=∑i⁢firadial⁢Δ⁢⁢tiλEccentricity=tan-1⁡(2⁢⁢Ptangential⁢Δ⁢⁢PH1+Pradial⁢Δ⁢⁢VK12⁢⁢Ptangential⁢Δ⁢⁢PK1-Pradial⁢Δ⁢⁢VH1)Δ⁢⁢H⇀ECI=COrbit⁢⁢to⁢⁢ECI⁢Δ⁢⁢H⇀Δ⁢⁢H⇀=∑i⁢r⇀i⊗f⇀i⁢Δ⁢⁢ti∑i⁢fitangential⁢Δ⁢⁢ti=Δ⁢⁢PDrift;whereΔ{right arrow over (H)}=momentum dumping requirement (vector) in orbit frameΔ{right arrow over (H)}ECI=momentum dumping requirement (vector) in Earth—Centered Inertial frameΔPK1=spacecraft mass×minimum delta velocity required to control mean K1 ΔPH1=spacecraft mass×minimum delta velocity required to control mean H1 ΔPDrift=spacecraft mass×minimum delta velocity required to control mean Drift{right arrow over (R)}i=lever arm (vector) about the c.g. for the ith thruster in spacecraft body frame{right arrow over (F)}i=thrust vector for the ith thruster in spacecraft body frameΔti=on time for the ith thrusterλEccentricity=location of the maneuverCOrbit to ECI=transformation matrix from orbit to ECI frame, rotation matrix about the Z by λEccentricity CBody to Orbit=transformation matrix from spacecraft body to orbit frame{right arrow over (r)}i=CBody to Orbit {right arrow over (R)}i f⇀i=CBody⁢⁢to⁢⁢Orbit⁢F⇀i=[fitangentialfiradialfinormal]. 16. A method of simultaneous orbit control and momentum dumping in a spacecraft, the spacecraft including a plurality of thrusters, the method comprising the steps of: generating a set of firing commands for the thrusters from solutions to momentum dumping/drift and eccentricity control equations; andfiring the thrusters according to the firing commands;wherein the momentum dumping/drift and eccentricity control equations are defined as ⁢∑j=1,2⁢⁢Pjtangential=Δ⁢⁢Pdrift⁢(2⁢⁢P1tangential⁢cos⁢⁢λ1+P1radial⁢sin⁢⁢λ1)+(2⁢⁢P2tangential⁢cos⁡(λ1-Δλ)+P2radial⁢sin⁡(λ1-Δλ))=Δ⁢⁢PK1⁢(2⁢⁢P1tangential⁢sin⁢⁢λ1-P1radial⁢cos⁢⁢λ1)+(2⁢⁢P2tangential⁢sin⁡(λ1-Δλ)-P2radial⁢cos⁡(λ1-Δλ))=Δ⁢⁢PH1⁢-2⁢⁢P1radial⁢P2radial⁢sin⁢⁢Δλ-4⁢⁢P1tangential⁢P2radial⁢cos⁢⁢Δλ-8⁢⁢P1tangential⁢P2tangential⁢sin⁢⁢Δλ+4⁢⁢P1radial⁢P2tangential⁢cos⁢⁢Δλ=0⁢λ2=λ1-Δλ⁢Δ⁢H⇀ECI=Δ⁢⁢H⇀ECI,1+Δ⁢⁢H⇀ECI,2⁢Δ⁢⁢H⇀ECI,1=COrbit⁢⁢to⁢⁢ECI⁡(λ1)⁢Δ⁢⁢H⇀1⁢Δ⁢⁢H⇀ECI,2=COrbit⁢⁢to⁢⁢ECI⁡(λ2)⁢Δ⁢⁢H⇀2⁢Pjradial=∑i⁢fi,jradial⁢Δ⁢⁢ti,j⁢Pjtangential=∑i⁢fi,jtangential⁢Δ⁢⁢ti,j⁢Δ⁢⁢H⇀j=∑i⁢r⇀i,j⊗f⇀i,j⁢Δ⁢⁢ti,j;where{right arrow over (r)}i,j=CBody to Orbit {right arrow over (R)}i,j f⇀i,j=CBody⁢⁢to⁢⁢Orbit⁢F⇀i,j=[fi,jtangentialfi,jradialfi,jnormal]j=1,2i=index⁢⁢for⁢⁢the⁢⁢ith⁢⁢thruster.