Method for controlling east/west motion of a geostationary satellite

Abstract: A method for East/West stationkeeping of a geostationary satellite maintains the osculating value of longitude from exceeding a specified deadband for a specified drift period between maneuvers. In the planning, the mean longitude, the mean drift rate and the mean eccentricity vector are calculated and then maneuvers are executed to maintain the values below a magnitude such that the osculating longitude will be within the specified deadband over a specified drift period. The target conditions are achieved through a plurality of maneuvers which initiate a period of free drift which lasts for a specified number of days. During the free-drift period, longitude remains within its deadband, and no additional maneuvers are needed or are performed. The method has the advantage that it can take into account limitations on thruster on-time by allowing for a generalized number of stationkeeping maneuvers.

Description: BACKGROUND OF THE INVENTION This invention relates to a method for simultaneously controlling the mean eccentricity and the longitudinal (East/West) motion of a geostationary satellite. A method for controlling the East/West motion of geostationary satellites is of interest due to the stringent constraints imposed upon a satellite's longitudinal motion by ground-based antennas. The two methods of orbital control are referred to as North/South stationkeeping, which controls latitude excursions, and East/West stationkeeping, which controls longitude excursions. Though it is understood that there can be a cross-coupling effect during North/South stationkeeping maneuvers, attention herein is directed to the East/West control problem and neglects the effect of North/South maneuvers in the East/West direction. With the increase of capability of the on-board payloads on satellites, there is a proportional increase in the power requirement of the payloads. The conventional solution is to enlarge the surface area of solar panels attached to the satellite in order to increase reception of available solar power. However, increasing the surface area of solar arrays enhances the perturbing force due to solar radiation pressure. Specifically, the solar panels act as sails driven by the solar wind with a force directly proportional to the area-to-mass ratio of the satellite. Perturbations caused by solar radiation pressure result in the change in the magnitude of orbital effects of solar radiation pressure cause eccentricity magnitude to vary with a yearly period during which time the eccentricity magnitude will grow to a value that may cause apparent longitudinal motion (East/West motion) to exceed the limits of the permissible deadband. Eccentricity appears as daily longitudinal oscillations about the nominal location of the geostationary satellite as observed from a position on the Earth. For geostationary satellites, the mean longitudinal acceleration is governed by the non-spherical effects of the Earth and the gravitational effects of the Sun and the Moon. In general, the Earth's tesseral harmonics are the dominant contribution to the longitudinal behavior, and their effects can be considered constant over longitudinal variations of less than 1 degree (< 1°). However, for satellites near the Earth's gravitational equilibrium points, the longitudinal acceleration due to the Earth becomes small in comparison to that due to the Sun and the Moon. It is of interest that there are two stable equilibrium locations and two unstable equilibrium locations for satellites at geostationary altitude. The longitudinal acceleration caused by luni-solar effects dominate the mean longitudinal behavior near the equilibrium points and have significant effects away from the equilibrium points (Figure 1). Therefore, an accurate model of the Sun's and Moon's effects on the orbital motion of the satellite must be used in order to target the longitude and longitudinal drift rate needed to initiate a proper free-drift cycle for any arbitrary satellite station location. It is necessary to account for a longitudinal acceleration that is not constant, but which is time varying in magnitude and, for station locations near an equilibrium point, in direction as well. In order to appreciate the variability of longitudinal acceleration, reference is made to Figure 1 which illustrates the long-periodic behavior of the longitudinal acceleration over a period of one month for various station longitudes of an object in geostationary orbit. The curves represent longitudinal accelerations in degrees per days squared (°/day 2 ). The longitudinal acceleration is time-varying with a period of approximately 13.6 days. This graph illustrates the complexity of East/West stationkeeping, especially near an equilibrium point in that at the longitudinal acceleration can change direction. Moreover, due to variations in the gravity at different positions on the Earth, the value of the longitudinal acceleration will vary with the Earth's longitude. An additional factor to consider is the solar radiation pressure effects on a spacecraft. Effects due to solar radiation pressure will perturb an orbit in such a way as to move the eccentricity vector perpendicular to the Earth-Sun line and in the direction of the orbital motion. If allowed to proceed naturally, this would result in an eccentricity vector roughly describing an ellipse with a peak magnitude possible greater than the allowable deadband and with a period of approximately one year. Superimposing the luni-solar effects onto this behavior serves to distort the ellipse and to cause local variations in the magnitude and rotation rate of the eccentricity vector. Depending upon the area-to-mass ratio of the satellite, the daily excursions of longitude could exceed the deadband by a significant factor. The eccentricity vector extends from the Earth to the perigee. The value e y is equal to [e sin(ω)], and the value e x is equal to [e cos(ω)], where ω is the argument of perigee. Figure 2 is a graph illustrating the mean eccentricity motion e x versus e y over a one-year period t 0 to t T due to luni-solar gravitational effects and solar radiation pressure. The solar radiation pressure causes the annual elliptical motion, and the solar gravitation effects result in an ellipse which does not close over a period of one year. The smaller oscillations are primarily due to the Moon's gravitational effects on the satellite and have a period of approximately 13.6 days. In order to meet the constraints established for controlling East/West stationkeeping of the satellite, the mean values of longitude and drift rate must be maintained to prevent any secular or long-periodic trends caused by the Earth's and the luni-solar gravitational effects from allowing the daily longitude variations to exceed the deadbands for a specified drift period between maneuvers. Furthermore, the mean eccentricity vector (which is manifested as the daily longitude variation) must be maintained below a preselected magnitude so that when superimposed upon the mean longitude, the net, or osculating, longitude will not exceed the specified deadband for the same drift period. Both of these conditions must be met simultaneously using the same set of maneuvers while making efficient use of propellant. A search of prior art literature uncovered a number of references, only one of which appears to address the effects of the Sun and Moon on eccentricity targeting, although a few address the luni-solar effects on longitude and drift-rate targeting. The following is a summary of the references uncovered and their relevance. M. C. Eckstein, "Station Keeping Strategy Test, Design and Optimization by Computer Simulation," Space Dynamics for Geostationary Satellites , October 1985, Toulouse, France (CEPAD): This paper discusses stationkeeping with small longitudinal deadbands. The luni-solar effects are included by removing some of the deadband, that is, the effects due to the Sun and Moon are considered small, and a certain amount of the deadband is budgeted to these effects. This paper includes a study on the result of a simulation that simultaneously targets eccentricity and longitude to initiate a free-drift cycle and uses the Sun-Pointing Perigee Strategy (SPPS) to target eccentricity. (SPPS is a method for controlling the eccentricity vector by maintaining it in the general direction of the Sun.) The targeting scheme used therein excludes the Sun's and the Moon's gravitational effects on eccentricity. This paper does not address the possibility of a variable longitudinal acceleration due to luni-solar effects, and it assumes that acceleration will be constant in direction and magnitude. This results in a need for an iterative control scheme requiring some operator intervention. This problem is addressed by the present invention. A. Kamel and C. Wagner, "On the Orbital Eccentricity Control of Synchronous Satellites," The Journal of the Astronautical Sciences , Vol. 30, No. 1 (January-March 1982), pp. 61-73: This paper contains many of the fundamental ideas behind the current strategy. The eccentricity targeting strategy therein used the Sun Pointing Perigee Strategy (SPPS) but only included the effects of the solar radiation pressure on eccentricity, thus excluding the luni-solar effects. The calculation of the longitudinal acceleration took into account only the Earth's gravitational effects (tesseral harmonics). The paper also suggests allotment of the luni-solar effects on eccentricity into the deadband budget, thereby tightening the usable deadband and costing more propellant. The present work represents an advancement over this simplification, particularly where the limited deadband constraints renders it unfavorable to include the luni-solar effects by removing part of the available deadband. C. F. Gartrell, "Simultaneous Eccentricity and Drift Rate Control," Journal of Guidance and Control , Vol. 4 , No. 3 (May-June 1981), pp. 310-315: This paper describes a method for simultaneous control of eccentricity and drift rate using equations which only take into account solar pressure on orbit eccentricity. The spacecraft considered in this analysis is quite small, so the effect of solar radiation pressure on eccentricity is not large. Thus, the maneuvers required to control drift rate are not required to control eccentricity magnitude, therefore, only one maneuver is required. The maneuver strategy is not generalized to N-maneuvers (where N is an arbitrary number of maneuvers), and the equations do not consider the gravitational effects of the Sun and the Moon in the targeting strategy. E. M. Soop, Introduction to Geostationary Orbits , European Space Agency (November 1983), Publication SP-1053 , pp. 1-143: This work provides a fairly broad explanation of orbital control of geostationary satellites. Particular attention is directed to Chapter 7 relating to longitude stationkeeping. It discusses the Sun-Pointing Perigee Strategy to target eccentricity and the use of two maneuvers to simultaneously target both drift rate and eccentricity in order to maintain the longitudinal deadband. A. Kamel et al ., "East-West Stationkeeping Requirements of Nearly Synchronous Satellites Due to Earth's Triaxiality and Luni-Solar Effects," Celestial Mechanics , Vol. 8 (1973), (D. Reidel Publishing Co., Dordrecht-Holland), pp. 129-148: This paper describes the equations for the disturbing function (which contains Earth's gravitational effects, as well as the luni-solar gravitational effects). This papers indicates that by not including the perturbations in the orbit caused by the Moon's and Sun's gravitational effects, longitude and drift rate will be incorrectly targeted. Eccentricity targeting is not mentioned in this paper, nor was any strategy suggested here which would achieve the proper target drift and longitude. Kamel has written several papers which contain a description of an analytic solution addressing the geostationary stationkeeping problem. Not all of those are discussed herein. D. J. Gantous, "Eccentricity Control Strategy for Geosynchronous Communications Satellites," Telesat Canada , May 1987: This paper describes an amended Sun-Pointing Perigee Strategy based on earlier work of Kamel and Wagner. The author observes that the natural motion of the eccentricity vector caused by the solar radiation pressure and the luni-solar gravitational effects will have a linear trend, i.e., the eccentricity plot does not achieve a closed ellipse after a year (as can be seen in Figure 2). The targeting strategy described therein accounts for this phenomenon. However, the targeting strategy does not consider the local variations due to the gravitational effects of the Sun and the Moon. J. A. Kechichian, "A Split ΔV for Drift Control of Geosynchronous Spacecraft," AIAA/AAS 21st Aerospace Sciences Meeting , (January 10-13, 1983), AIAA-83-0017: This paper discusses the use of a two-maneuver technique over a 24-hour period in order to maintain a longitude deadband of ±0.05° while simultaneously controlling longitude and eccentricity. The author discusses targeting osculating effects, whereas the present invention addresses targeting mean values. The author concludes that his technique will deviate from a specified deadband limit after only a few cycles. S. Parvez et al ., "East-West Stationkeeping Near an Equilibrium Longitude at 105 Degrees West," AAS/AIAA Astrodynamics Specialist Conference , AAS-87-545 (August 10-13, 1987), pp. 1-14: This paper describes stationkeeping near a stable equilibrium point. This paper commented on the problems of predicting spacecraft longitudinal motion after a change in orbital velocity due to a required attitude adjustment. This paper did not consider the luni-solar effects and seemed mainly a plea for an explanation for the apparently high sensitivity of longitude motion to attitude maneuvers and the lack of a predictable drift cycle. Various other papers were uncovered of less relevance than those mentioned above. Further, a search of the Patent Office records relating to this invention uncovered the following patents, none of which are particularly relevant to the problem at hand. U.S. Patent No. 4,767,084 to Chan et al. uses momentum unloads to control stationkeeping. There are no instructions on how to control the orbit, and this is a scheme that may not be appropriate for spacecraft with large area-to-mass ratios because inadequate thruster power is available. U.S. Patent No. 4,521,855 to Lehner et al. relates to physically controlling the attitude of the satellite and does not relate to stationkeeping. U.S. Patent No. 4,599,697 to Chan et al . also relates to spacecraft attitude control. U.S. Patent No. 4,537,375 to Chan also relates to attitude control. U.S. Patent No. 4,776,540 to Westerlund relates to attitude control of a satellite to compensate for changes in latitude. U.S. Patent No. 4,837,699 to Smay et al. also relates to attitude control, but in the context of a spinning satellite. In view of the foregoing, what is needed is a generalized method for East/West stationkeeping and targeting of geosynchronous satellites that uses an arbitrary number of maneuvers and which takes into account luni-solar effects on both eccentricity and longitudinal motion. According to the present invention there is provided a method for controlling east/west motion of a geostationary satellite in an orbit characterized by a mean eccentricity vector, said method comprising the steps of: targeting mean eccentricity of said orbit by calculating a first change in velocity of said satellite that would place a target mean eccentricity vector at a first state which will result in a subsequent mean eccentricity vector being within a specified tolerance during a period of free drift; thereafter targeting mean longitude and mean longitudinal drift rate of said orbit by calculating a second change of velocity as a function of mean longitudinal acceleration of said satellite that would place a target mean longitude and a target mean longitudinal drift rate at second states that will result in the total longitude of said satellite east/west motion relative to a station on the Earth to be constrained within a specified deadband; apportioning the larger of said first change in velocity and said second change in velocity into N changes in velocity to simultaneously control mean eccentricity, longitude and longitudinal drift rate, where N is an arbitrary integer greater than 1; and imparting said changes in velocity to said orbit in N manoeuvres and providing specified wait times between said manoeuvres in integral multiples of one-half day to attain said target mean eccentricity vector being at said first state and said target mean longitude and target mean longitudinal drift rate being at said second states; characterised in that each of said mean longitude and longitudinal drift rate targeting steps take Earth. moon, and sun gravitational effects into account. The invention will be better understood by reference to the following description in connection with the accompanying drawings. BRIEF DESCRIPTION OF DRAWINGS Figure 1 is a graph of long-periodic behavior of the longitudinal acceleration of an object in geostationary orbit about various station longitudes. Figure 2 is a graph illustrating the mean eccentricity motion due to luni-solar effects and solar radiation pressure. Figure 3 is a vector diagram for illustrating a Sun-pointing perigee strategy according to the invention. Figure 4 is a diagram for illustrating change in eccentricity in the y -direction corresponding to a change in rotation. Figure 5 is a diagram for illustrating a change in eccentricity in the x -direction corresponding to the change in magnitude. Figure 6 is a drift rate versus longitude diagram for a five-maneuver, seven-day East/West drift cycle. Figure 7 is a drift rate versus longitude diagram for a four-maneuver, seven-day East/West drift cycle. Figure 8 is a flowchart illustrating a method for controlling eccentricity in accordance with the invention. DESCRIPTION OF SPECIFIC EMBODIMENTS To achieve desired conditions according to the invention, the total change in velocity vector ΔV required for eccentricity control is first determined. The longitude and drift rate targeting scheme then splits the total change in velocity vector into a plurality of or "N" maneuvers, each of which is sufficiently small so that the daily or short-periodic spacecraft longitudinal oscillations about the designated station do not exceed a predetermined deadband. In the particular scheme adopted, special attention is given to the local variations due to luni-solar effects. After the "N" maneuvers are completed, the free-drift period begins which is based on the long-period oscillations of the spacecraft, herein referred to as the mean parameters. Especially during the free-drift period, the daily oscillations of longitude do not exceed the deadband. It is desired to keep the entire East/West stationkeeping cycle duration to a minimum in order to save propellent, but the cycle should not be so short as to require daily maneuvers. A typical minimum cycle is one week as an operational convenience. This cycle allows up to seven maneuvers and still permits an adequate free-drift period in which to verify orbit parameters through the orbit determination process. Figure 3 is a phase space diagram for illustrating eccentricity control based on the Sun Pointing Perigee Strategy (SPPS). The Earth 10 is located at the center of a constraint circle 12 having an eccentricity magnitude E c specifying the desired stationkeeping constraint. Eccentricity is initially referenced to an angle α o between an orbital eccentricity vector 13 and a line 14 drawn between the center of the seasonal eccentricity ellipse 20 and the Sun 16. The line 14 moves with the annual progression of the Sun 16 to a final position of the Sun 16". Therefore, the targeting scheme as explained herein is based on a predicted mean position 18 of the Sun 16' defined as a midpoint in time of the free-drift period following the last maneuver. The targeting equations are referenced to an average Earth-Sun line 14'. Relative movement of the Sun 16 with time is West to East (or counterclockwise in Figure 3) during its annual progression. Because of solar radiation pressure on the satellite, the eccentricity vector, e, of the orbit is naturally driven perpendicular to the Earth-Sun line and rotates in the same direction as the orbital motion. The eccentricity magnitude will grow to a seasonal value e r at a rate dependent on the area-to-mass ratio and the reflectivity characteristics of the satellite. It is desired to control the eccentricity of the orbit. In Figure 3, three states of the eccentricity vector e of the orbit are shown: an initial eccentricity vector e o , a target eccentricity e t , and a final eccentricity e f , each of which is referenced to the Earth as an origin. The initial eccentricity vector e o 13 leads the Earth-Sun line 14. The target eccentricity vector e t is driven in accordance with the invention by a composite vector Δe (which is greatly exaggerated in this diagram) to lag the Earth-Sun line 14 and to bring the target eccentricity vector e t within the constraint circle 12 (if not already within the constraint circle). The final eccentricity vector e f is the result of disturbances caused by solar radiation pressure and luni-solar gravitational effects. These effects are represented by a seasonal eccentricity vector e r (t), shown in various forms as e ro (t), e rt (t), and e rf (t). It is displaced from an origin 20 to an origin 20' by the composite change in the eccentricity vector Δe. Its track is represented as the original perturbed ellipse 22 and as the final perturbed ellipse 22' as displaced by the target eccentricity vector, e t . Ellipse 22 is the natural path of the seasonal eccentricity vector e r . Ellipse 22' is the path e r will take after the Nth maneuver is completed. A cusp 24 is displaced by the forced disturbance, as indicated by corresponding cusp 24'. It is the object of the invention to maintain the track of the final ellipse 22' within the constraint circle 12 at all times during the East/West stationkeeping cycle. Among the parameters which are monitored during the control of the eccentricity are the magnitude and direction of the angle between the eccentricity vector, e , and the current line 14, 14' or 14". The initial angle is α o , the target angle is α t , and the final angle is α f . The final angle α f is referenced to the final eccentricity vector and line 14". The sign of the angle indicates whether the eccentricity vector leads (positive) or lags (negative) line 14. Maneuvers executed near the crossing of the semi-latus rectum (i.e., when true anomaly is at ±90°) will maximize the rotation of the eccentricity vector, e , relative to the Sun while minimizing the change in magnitude of the eccentricity vector. Maneuvers executed near or on the line of apsides (i.e., when true anomaly is at 0° or at 180°) will maximize the change in magnitude of e while minimizing the change in rotation. According to the invention, the maneuver strategy starts with perigee pointing in a direction generally towards the Sun in order to keep corrections to the magnitude of the eccentricity vector e small. Therefore, East/West stationkeeping maneuvers mainly take place near the crossing of the semi-latus rectum. In this way, the eccentricity vector, e, roughly tracks the Sun, and is made to librate about the Earth-Sun line, which is the basic idea behind the SPPS. The maneuver sequence is initiated with the eccentricity vector, e o , (Figure 3) leading line 14 by angle α o . The total maneuvers comprising Δe are performed to reset the target eccentricity vector e t to lag line 14 by angle α t . The effects of the solar radiation pressure and the luni-solar gravitational effects will naturally drive the eccentricity vector to e f at angle α f line 14". The process is repeated to maintain the eccentricity vector e within the bounds E c of the circle of constraint 12. If eccentricity drifts outside of the circle of constraint 12, the daily longitude librations may exceed the deadband requirements. In accordance with the invention, the maneuvers take into account the time-dependent variations (24) on the eccentricity ellipse. In order to prevent any one maneuver sequence from becoming too large, limits are placed on the amount of permissible change in the eccentricity vector. A coordinate system is selected which will rotate with the average Earth-Sun line (i.e., it will have a period of approximately one year). For a Sun Pointing Perigee Strategy, the coordinate system may be a Cartesian (x-y) coordinate system wherein the x direction (magnitude) is along the average Earth-Sun line (Figure 4 and Figure 5) and the y direction is in the orbit plane perpendicular to the Earth-Sun line. Changes in the x direction affect the approximate magnitude of the eccentricity vector, and changes in the y-direction affect the approximate rotation of the eccentricity vector. This coordinate system is interpreted in the above fashion only for non-zero eccentricity vectors that point in the general vicinity of the Sun; these conditions are appropriate for normal operational conditions. Referring to Figure 4, the two parameters that control the rotation of the eccentricity vector are E cu and E cl , which relate to the eccentricity constraint value E C . The parameter E cu places an upper limit on the amount that the eccentricity can be translated down relative to the selected coordinate system to the average Earth-Sun line and establishes the maximum rotation allowed. The parameter E cl is a bias off of the average Earth-Sun line which allows the method to slightly over target. The average change in eccentricity needed to target the y-direction only is: Δe yav = 1/2 (δe yt + δe yf ) Thus, Δe yav is the change in eccentricity in the y direction from the initial ellipse 22 to the target ellipse projected in the y direction. By placing a limit on the amount the eccentricity is allowed to rotate in order to meet the target conditions, the expression for the y direction of the required change in eccentricity is: Δe y = MIN[MAX(-E cl , Δe yav ), E cu ] + E cl This expression is illustrated graphically in Figure 4. Figure 5 illustrates the change in eccentricity in the x-direction, corresponding to a change in magnitude. Shown are the eccentricity value at the midpoint of the free drift cycle, e mp , the change in eccentricity vector at the beginning of the free drift cycle, δe xt , the change in eccentricity vector at the final position δe xf , and the scaled efficiency parameters η 1 and η 2 The change needed in the x-component of eccentricity is taken as the difference between the x-component of the eccentricity vector at the beginning of the free drift cycle or the x-component for the eccentricity vector at the end of the free drift cycle and the constraining value for eccentricity E c whichever is larger: Δe x = MAX(δe xt - E c , δe xf - E c ) The two efficiency parameters are introduced to control the change in magnitude of the eccentricity vector, these parameters are η 1 and η 2 . Since it is more efficient to locate the maneuvers close to the semi-latus crossing, these two parameters bind the location of the firings near the perpendicular to the average Earth-Sun line. Therefore, the allowed change in the x-component of the eccentricity vector is: Δe x = MAX(-η 1 * Δe y , MIN(η 2 * Δe y , Δe x )) Note that the allowed change in the x-component of the eccentricity vector (measured in the rotating coordinate system which rotates with the average Earth-Sun line) is expressed as a function of the change in the y-component of eccentricity. This is to restrict the amount of ΔV that goes into changing the magnitude of eccentricity as opposed to changing its direction. Magnitude changes to eccentricity are inefficient, and therefore, are scaled by the rotation change. Given the components for the change required in eccentricity that achieve the proper target conditions which will initiate a free drift cycle, the corresponding value of the change in velocity ΔV needed can be calculated by the square root of the sum of the squares of eccentricity components: ΔV ec = 1/2*V s *SQRT(Δe x 2 + Δe y 2 ) where V s is the orbital velocity at synchronous altitude. Finally, the location of the true anomaly for the first maneuver (with all subsequent maneuvers located no less than half a sidereal day apart) can be determined by locating the inertial angle for the change in eccentricity vector. The true anomaly, or angle in the orbit plane from the perigee, for the maneuver is located on the line parallel to the Δe vector, noting that there are two solutions 180° apart. Care must be taken that the first maneuver is placed in the vicinity of the initial true anomaly provided to the East/West stationkeeping algorithm. The true anomaly ν 0 can be taken as: ν 0 = arctan (Δe y /Δe x ) - E SRA where E SRA is the angle between the average Earth-Sun line and the initial eccentricity vector. The change in velocity ΔV ec is subdivided in order to simultaneously control eccentricity and longitude. Since in general the change in velocity needed to control the eccentricity vector is not the same as the change in velocity needed to control longitude and drift rate, two or three maneuvers are typically required. If it is then desired to perform the stationkeeping maneuvers in the on-orbit, or momentum wheel, control mode, then the total change in velocity ΔV ec required to control eccentricity may be further subdivided into a set of smaller ΔV i 's. These smaller ΔV i 's are chosen with attention to the sign (indicative of direction of thrust) relative to the orbital velocity. This is done in such a manner that total change in velocity allows eccentricity to achieve its target conditions. The other portion of the problem is determination of the target conditions for longitude and drift rate. It is essential to control longitude within the constraints of acceleration; it is important to recognize that acceleration will change in magnitude and may even change direction. Figures 6 and 7 are examples which illustrate the effects of ΔV maneuvers in the longitude and drift rate phase plane for two different initial conditions and different directions of longitudinal acceleration. Figure 6 illustrates an odd-number of maneuvers, namely, a 5-maneuver, 7-day cycle, wherein the ordinate axis is drift rate in degrees (East) per day and the abscissa is longitude in degree (East) relative to a station location at the origin, and where the maneuvers are symmetric about each axis. Referring to Figure 6, the time from point 0 to point F is the East/West stationkeeping period. The free drift period P D is the period from the completion of the last maneuver (point 5) to the final point F. In this example, the longitudinal acceleration is negative. The sum of the change in velocities for the five maneuver example is given by: ΔV ec = -ΔV 1 + ΔV 2 - ΔV 3 + ΔV 4 - ΔV 5 where each ΔV i carries its own sign. The maneuvers are shown for the case where the total change in velocity to control eccentricity ΔV ec is greater than the change in velocity to target longitude and drift rate ΔV D , as indicated by the alternating directions for the maneuvers shown in both Figure 6 and Figure 7. Otherwise, the maneuvers will not alternate direction if performed each 1/2 day. The kinematicequation relating to the maneuvers of Figure 6 is given by: λ N = λ 0 + Σ dλ i /dt*Δt i + Σ d 2 λ i /dt 2 *(Δt i 2 )/2 where the summations Σ are from i= 0 to i= N-1 except as noted otherwise hereinafter; for example N=5; and λ is longitude, its first derivative with respect to time dλ/dt is the longitudinal rate and its second derivative d 2 λ/dt 2 is the drift acceleration. The time Δt i is nominally 1/2 day between maneuvers, but may be integer multiples of 1/2 day. The above equation can be rearranged to place all unknowns on one side of the equation and to allow the equation to be expressed more conveniently in matrix form: -λ N + Σ dλ i /dt*Δt i = -λ 0 - Σ d 2 λ i /dt 2 *(Δt i 2 )/2 where the unknowns are all placed on the left-hand side of the equation. The unknowns in matrix form for the N=5 example are thus: X = [ λ 5 , dλ 1 /dt, dλ 2 /dt, dλ 3 /dt, dλ 4 /dt, dλ 5 /dt ] where the target longitude is λ 5 , the target drift rate is dλ 5 /dt and the intermediate drift rates are dλ 1 /dt, dλ 2 /dt, dλ 3 /dt and dλ 4 /dt. Shaping parameters may be introduced to control the relative magnitudes of the intermediate longitudes λ i and drift rates dλ i /dt between successive maneuvers. Rewriting Equation (9) yields: Σ dλ i /dt*Δt i = Δλ D Σ γ i - Σ d 2 λ i /dt 2 *(Δt i 2 )/2 where: Σ γ i = 1, for i=1 to N-1 which represents the shaping parameters. Re-expressing Equation 7 in terms of sums of drift rates and rearranging to place all unknowns on one side of the equation yields: 2 Σ (-1) i dλ i /dt + (-1) N dλ N /dt = = 6πβ/ (T s *V s )¦ΔV ec ¦ -Σ (-1) i d 2 λ i /dt 2 *(Δt i ) where T s is the ratio of the sidereal day to the solar day; V s is the synchronous (two body) velocity; and β is the sign (+/-) used to keep track of the direction of the change of velocity. The foregoing equations may be expressed in matrix form for any number of maneuvers. In a system having dual axis symmetry as in Figure 6, an odd or even number of maneuvers can be required. An odd number of maneuvers assures dual axis symmetry. The foregoing equations may be expressed in matrix form as follows: Ax = B Figure 7 shows a diagram of maneuvers in the longitude and drift rate phase plane for a 4-maneuver example, 7-day cycle. The ordinate is drift rate in degrees (East) per day and the abscissa is longitude is in degree (East) relative to a station longitude located at the origin. Note, the maneuvers are symmetric about only the drift axis example and the longitudinal acceleration is positive. Equation (18) guarantees symmetry about the drift rate axis (λ TA = -λ F ): 2λ N + dλ N /dt*(P D +τ)=-dλ N 2 /dt 2 *(P D 2 + τ 2 )/2 where P D is the free drift period, τ is the time required to transition from target conditions to turnaround conditions (λ TA ), dλ N /dt is the target drift rate and dλ N 2 /dt 2 is the mean daily acceleration averaged over the free drift cycle. The free drift period P D is the period from the completion of the last maneuver (point 4) to the final point F. The sum of the change in velocities for the four maneuvers is given by: ΔV ec = -ΔV 1 + ΔV 2 - ΔV 3 + ΔV 4 where each ΔV i carries its own sign. In this case with single axis symmetry, the matrix equations for the target longitude λ F are given as follows: Ax = B where In order to execute the desired maneuvers, the following inputs are needed: Cycle length (in days, such as 7 or 14); Station longitude; A table of unperturbed eccentricity components (e x, e y ) subject to solar radiation pressure, Earth gravitational effects and the luni-solar gravitational effects; A table of unperturbed longitudes and drift rates (in the absence of maneuvers), subject to the Earth's gravitational effects and the luni-solar effects; Initial true anomaly in the approximate location of the first maneuver (ν 0 ); Initial mean daily longitude and mean daily drift rate. Longitudinal acceleration is not constant over the free drift period so that symmetry about the drift rate axis (Figure 6 or Figure 7) cannot be achieved in reality. Because the longitudinal acceleration is not constant it is averaged again over the period of free-drift cycle. By doing so, off-line iterations by an operator to account for luni-solar effects in the targeting scheme are not required, as the technique described above produces very accurate results without operator intervention. These functions, called eccentricity targeting and longitude and drift rate targeting, are implemented by way of example in a digital computer executing the programs EWMNVR and ECNSTR disclosed in Appendix A attached hereto. The subroutine ECNSTR is called by EWMNVR. It determines the change in velocity required for eccentricity to achieve its target conditions which will maintain eccentricity below its constraint value for the entire East/West cycle time and calculates the exact time of the first maneuver. It also provides the true anomaly for the first maneuver (which is converted into time of the first maneuver). Figure 8 is a flow chart for ECNSTR. Called from EWMNVR (Step A), ECNSTR calculates the unperturbed eccentricity magnitude at the beginning and end of the East/West cycle (EO, EF) (Step B), then calculates the Average Right Ascension of the Sun over the free drift cycle (AVSRA) (Step C). Thereafter it calculates the right ascension of the eccentricity vector at the beginning of the free drift cycle and at the end of the free drift cycle (EORA, EFRA) (Step D), then it calculates the projection of the eccentricity vector onto the Average Sun-Line at the beginning of the free drift cycle and at the end of the free drift cycle (DEOX, DEOY, DEFX, DEFY) (Step E), and then calculates the change in the eccentricity vector required to meet targeting constraints (DELEY and DELEX), restricting the change to values dictated by η 1 and η 2 (EFF1 and EFF2) in the x-direction and EUPL and ELWL in the y direction (Step F). It then calculates the change in velocity (DVEC), determining it from the required change in eccentricity (Step G). Finally, it calculates the true anomaly of the first maneuver (F1) based on the location of the initial eccentricity, located in the same half plane (Step H) before returning to the calling program (Step I). The above technique is intended to provide East/West stationkeeping within the maximum value of its constraint with a minimal expenditure of propellant and no off-line operator intervention. It is not suitable for station acquisition. It takes into account the constraints on eccentricity and size of the longitude deadband. Whereas previous techniques have attempted to lump many of the uncertainties into the deadband, the present invention takes some of these into account, including luni-solar effects, in order to reduce fuel usage. The invention has now been explained with reference to specific embodiments. Other embodiments will be apparent to those of ordinary skill in the art. It is therefore not intended that the invention be limited except as indicated by the appended claims.

Claims: A method for controlling east/west motion of a geostationary satellite in an orbit characterized by a mean eccentricity vector, said method comprising the steps of: targeting mean eccentricity of said orbit by calculating a first change in velocity (Δv ec ) of said satellite that would place a target mean eccentricity vector (e t ) at a first state which will result in a subsequent mean eccentricity vector being within a specified tolerance during a period of free drift; thereafter targeting mean longitude and mean longitudinal drift rate of said orbit by calculating a second change of velocity (ΔV D ) as a function of mean longitudinal acceleration of said satellite that would place a target mean longitude (λ N ) and a target mean longitudinal drift rate (dλ N ,/dt ) at second states that will result in the total longitude of said satellite east/west motion relative to a station on the Earth to be constrained within a specified deadband; apportioning the larger of said first change in velocity and said second change in velocity into N changes (Δv i ) in velocity to simultaneously control mean eccentricity, longitude and longitudinal drift rate, where N is an arbitrary integer greater than 1; and imparting said changes in velocity to said orbit in N manoeuvres and providing specified wait times between said manoeuvres in integral multiples of one-half day to attain said target mean eccentricity vector being at said first state and said target mean longitude and target mean longitudinal drift rate being at said second states; characterised in that each of said mean longitude and longitudinal drift rate targeting steps take Earth, moon, and sun gravitational effects into account. The method according to claim 1 wherein said imparting step comprises at least three manoeuvres and wherein said manoeuvres guarantee single-axis symmetry in a phase plane of longitude (λ) versus longitudinal drift rate, (dλ/dt) to maximize use of said deadband. The method according claim 1 wherein said imparting step comprises at least three manoeuvres and there is an even number of manoeuvres in a case where there is a reversal in direction of said mean longitudinal acceleration near an equilibrium point in an orbit due to said luni-solar effects, despite the fact that the station longitude of the satellite has not changed, and in order to guarantee single-axis symmetry in a phase plane of longitude versus longitudinal drift rate, to maximize use of said deadband. The method according to claim 1 wherein said imparting step comprises at least two manoeuvres in a case where there is a reversal in direction of said mean longitudinal acceleration near an equilibrium point in an orbit due to said luni-solar effects, despite the fact that the station longitude of the satellite has not changed. The method according to claim 1 wherein said imparting step comprises at least two manoeuvres in order to guarantee that the longitude at the end of the stationkeeping cycle and the longitude at turn around are equidistant either side of the centre of the required station. The method according to claim 1 wherein said mean eccentricity targeting step and said mean longitude and mean longitudinal drift rate targeting step each further comprises selecting a stationkeeping period of arbitrary duration as an input, said stationkeeping period being a function of the total longitudinal deadband and the longitude of the satellite. The method according to claim 1 wherein said method is performed several times, and for each sequence the change in the eccentricity vector is limited to a maximum value therefore limiting the maximum change in velocity per sequence. The method according to claim 1 wherein said eccentricity targeting step comprises determining said target mean eccentricity vector (e t ) lag an average Earth-Sun line, said average Earth-Sun line being the location of the Earth-Sun line at a midpoint in time over said period of free drift. The method according to claim 1 wherein said mean eccentricity targeting step comprises calculating an orbital ephemeris for a future East/West stationkeeping period to obtain a table of orbital elements versus time, said orbital ephemeris includes a force model of solar radiation pressure on said satellite, gravitational effects of the Earth, gravitational effects of the Moon and gravitational effects of the Sun under conditions where luni-solar effects dominate said deadband in order to preserve propellant.

Inventors: BINGAMAN RONALD H, GAMBLE DONALD W, KELLY THOMAS J, WHITE LISA K, BINGAMAN, RONALD H., GAMBLE, DONALD W., KELLY, THOMAS J., WHITE, LISA K.,
Applicants: LORAL SPACE SYSTEMS INC, SPACE SYSTEMS / LORAL INC.
Application year: 1991
Main classification: G05D1
IPC codes: G05D1, B64G1