HYBRID SPACECRAFT CONTROL SYSTEM

Abstract: A spacecraft control system (32) having monopropellant and bipropellant thrusters (54, 50), both types of thrusters operating from the same fuel. The selected fuel may be catalytically reacted in the monopropellant thrusters (54), or may be oxidized with a separately provided oxidizer in the bipropellant thrusters (50), in each case producing a hot gas which is expelled from the thrusters to create thrust. The fuel is provided in an excess relative to the amount required to react with the available oxidizer, so that the monopropellant thrusters (54) can continue to operate even after the oxidizer supply for the bipropellant thrusters (50) is exhausted. There are no unusable propellants, and the higher efficiency of the bipropellant thrusters is still utilized fully. The preferred fuel is hydrazine and the preferred oxidizer is nitrogen tetroxide.

Description: -I- " - HYBRID SPACECRAFT CONTROL SYSTEM BACKGROUND OF THE INVENTION This invention relates to spacecraft, and, more particularly, to the control system used in spacecraft. A spacecraft such as a satellite depends upon several types of propulsion and control systems during its life. For example, in the case of a spacecraft launched by the space shuttle, a powerful combination of main and booster engines lifts the entire shuttle orbiter into a low orbit around the earth. A spacecraft or satellite to be placed into a permanent or semi-permanent orbit is carried aloft in the payload bay of the orbiter, and then is launched into its own orbit from the orbiter with additional rocket thrusters. One important type of spacecraft is the communications satellite, which acts as a relay station for transmitted signals. An earth station transmits a television, telephone, data or other signal to the satellite, which in turn relays the signal to- another earth station, possibly ' after amplifying or otherwise modifying it. The earth stations are a key part of the overall system, and it is commercially important to minimize their cost. One approach to minimizing cost is to position the communications satellite at a fixed location in the sky relative to the earth stations, so that the sending and receiving antennas of the earth - stations can remain stationary and pointed at the fixed location, avoiding the need for tracking electronics and mechanisms that would otherwise be necessary for the ground antennas to continuously point at a satellite that moves relative to the earth stations. While this approach minimizes the cost of the earth station, it adds complexity to the requirements of the satellite control system. There is a unique orbit of the satellite that maintains the satellite in a position above a selected position on the earth. In this orbit,f ' termed a geosynchronous or geostationary orbit, the satellite is moving at the correct combination of altitude and . velocity to remain at a location that is fixed relative- to the earth stations, as the earth- rotates. The satellite would therefore appear to remain motionless above the earth, allowing fixed earth station antennas .to be used, were it not for external forces that are imposed on the satellite and tend to move it from its fixed location in geosynchronous orbit. For example, external forces arising from the gravity of the sun and moon cause the satellite to drift from its geosynchonous fixed position. The force - of the solar wind.acting on the asymetrical areas of the satellite cause it to rotate and point in directions other than that initiaϊly established. Consequently, even though the satellite is nominally in a stationary geosynchronous orbit, there must be a control system on the satellite to compensate for the external forces and to maintain the satellite in exactly the right position or station at which the fixed ground antennas are pointed. After the satellite is ' transported into space by the space shuttle or an expendable rocket, it is then boosted into a highly elliptical orbit by a thruster termed a perigee motor. When the elliptical orbit intercepts the altitude required for " the geosynchronous orbit, another thruster termed the apogee motor is fired to cause the satellite to enter the circular geosynchronous orbit. Once the satellite is in geosynchronous orbit, smaller thrusters are used . to position it precisely at the desired orbital station and with the proper orientation. This description has been presented for a communications satellite in geosynchronous orbit, but the same principles apply for other types of satellites in low-earth orbit, polar orbit, inclined orbit, etc. Whatever the orbit chosen, the spacecraft must be provided with the means to control its attitude and orbital position relative to the intended position in orbit, also termed the station, so that minor deviations from the intended position can be corrected. As discussed, the attitude and station of a communications satellite in geosynchronous orbit must be maintained precisely, so that communications links with the ground are not broken. The spacecraft is therefore provided with its own built-in control system, including thrust.ers which can propel the spacecraft in various modes. There are two, and possibly three, control modes of the thrusters in the control system. The thrusters control the attitude or orientation of the spacecraft with respect to the earth or other reference points. They also shift the position of the spacecraft in a north-south or east-west direction, should the spacecraft stray from its intended station in orbit, in maneuvers termed stationkeeping. The control system may also provide propulsion to move the spacecraft to an orbit of slightly different altitude above the earth. In current spacecraft, both attitude control and stationkeeping functions are usually provided by a single type of thruster, although there are numerous thrusters of that single type located at various positions on the spacecraft to allow combinations of thrusters to be fired to achieve the desired control movements. Attitude control usually requires more frequent, shorter pulses of thrust, while north-south and east-west stationkeeping and orbital changes require less frequent, longer pulses of thrust. Two types of liquid-propellant thrusters have been developed for spacecraft applications in attitude control and stationkeeping. Monopropellant thrusters generate hot gas by passing a fuel such as hydrazine over a catalyst such as iridium to decompose the fuel into a hot gaseous mixture, ana then direct the gas through a nozzle to provide thrust. Bipropellant thrusters generate gas by reacting a fuel such as hydrazine or the less-energetic monomethylhydrazine with an oxidizer such as nitrogen tetroxide, and then direct the hot gas through a nozzle to provide thrust. Considering both the energy potential of t-he fuel and the. thruster. design, monopropellant thrusters are most favored for small, inexpensive satellites because of their simplicity. Bipropellant thrusters are more favored when, used for the control of larger satellites and consequently greater propellant loads because of their greater efficiency. The selection of the particular thrusters for the control system depends upon the exact mission of the spacecraft. The selection of a control system utilizing bipropellant thrusters creates some problems in the use of the spacecraft. Attitude control firings and east-west stationkeeping are less efficient than steady state firings, but these maneuvers usually utilize a small fraction of the total propellants, on the order of about 1056. It is nearly always the case that the fuel and the oxidizer cannot be fully expended at the same time, leaving an unusable excess of one of the propellants. While the amount remaining of one of the propellants may not be more than a few percent, even this amount represents an important problem because of the high cost of boosting any weight into orbit, and for large spacecraft may represent several hundred pounds of propellant boosted into orbit and then unused. Attempts to predict with high precision the rate of expenditure of the propellants have not been successful because of the varied nature of the firing types and durations. Consequently, there usually remains either fuel or oxidizer that cannot be utilized. Selection of monopropellant thrusters avoids this problem of unused propellant, but causes the majority of thruster use to be less efficiently conducted. If only monopropellant thrusters are provided, much larger amounts of fuel must be carried, and again there is an excessive weight that must be lifted into space. ' There is therefore a need for an improved spacecraft control system that achieves .the necessary attitude control and stationkeeping functions, while at the same time operating efficiently and with full effective expenditure of the propellants. The control system must be compatible with the other spacecraft systems, and should not introduce new problems with safety of the spacecraft, launch preparation, or the incorporation of radically different structure or fluids into spacecraft design. The present invention fulfills this need, and further provides related advantages. SUMMARY OF THE INVENTION The present invention resides in a spacecraft control system which utilizes a hybrid monopropellant and bipropellant approach to providing .thrust for attitude control and stationkeeping, and for apogee boost. The control system achieves significantly improved spacecraft efficiency, without requiring major modifications to spacecraft procedures. In accordance with the " invention, a spacecraft control system comprises a bipropellant thruster operating from a fuel and an oxidizer; a monopropellant thruster operating from the same type of fuel as said bipropellant thruster; a supply of oxidizer communicating with the bipropellant thruster; and a supply of fuel communicating with the bipropellant thruster and the monopropellant thruster, the amount of fuel being greater than the amount required to react with all of the oxidizer. The fuel is preferably hydrazine and the oxidizer is preferably nitrogen tetroxide. The apogee motor can alsq be operated as a bipropellant thruster using the common fuel and oxidizer. The monopropellant thruster is preferably initially used for maneuvers requiring short pulses of thrust which require a small percentage of the total propellant and consequently its lower performance 'is of little signi icance. The bipropellant thruster is preferably used for maneuvers requiring long pulses of thrust, which require the majority of the total propellant. Because the bipropellant thruster only has to operate in the steady state mode, its design can be optimized for that single type of operation, so that it is more efficient than the conventional thrusters that must operate in both the steady state and pulsing modes. • The fuel is provided in excess, so that the oxidizer is exhausted first. When the oxidizer is depleted, the bipropellant thruster no longer functions. The monopropellant thruster can then be used for . both short and long duration firings, even though the monopropellant thruster is somewhat less efficient than the bipropellant thruster. An overall advantage is gained by using the otherwise unusable residual propellant. In the preferred configuration used for geosynchronous satellites, two sets of monopropellant thrusters and a bipropellant thruster are provided for attitude control and stationkeeping maneuvers. Radial monopropellant thrusters are positioned for performing attitude control and east-west stationkeeping, and the bipropellant thrusters are positioned for north-south stationkeeping in their normal operation. These thrusters are used in the early part of the life of the satellite, while both fuel and oxidizer are available. Axial monopropellant thrusters are . positioned for north-south stationkeeping when the oxidizer is depleted. After the oxidizer is exhausted and the bipropellant thrusters can no longer be used, then both the radial and axial monopropellant thrusters are used to control the satellite, expending the remainder of the fuel excess completely. The amount of fuel carried is therefore greater than the sum of the amount required to react with all of the oxidizer, plus the expected consumption of fuel by the monopropellant thrusters prior to the time that the oxidizer is expended. With this approach, all of the propellant " (fuel and oxidizer) is used fully in a useful manner. Alternatively stated, a method for controlling a spacecraft comprises providing the spacecraft with at least one bipropellant thruster and at least one monopropellant thruster; providing the spacecraft with a fuel storage tank communicating with the bipropellant thruster and the monopropellant thruster; providing the spacecraft with am oxidizer storage tank communicating with the bipropellant thruster; loading the oxidizer storage tank with an amount f 5" 5 of oxidizer; and loading the fuel storage tank with an amount of fuel in excess of the amount required to react with the available oxidizer in the oxidizer storage tank. Once the spacecraft is thus prepared it may be used by launching it into 10 space, operating the monopropellant and bipropellant thrusters until the oxidizer is exhausted, and then continuing to operate the monopropellant thruster after the oxidizer is exhausted. 15 It will now be appreciated that the system and method of the present invention represent an important advance in the control of spacecraft. Bipropellant and monopropellant thrusters can *>e designed^and used optimally, and 20 the onboard fuel and oxidizer may be pr.ovided. in quantities that will allow the spacecraft to function for the greatest possible time before exhaustion of these consumables. Both monopropellant and bipropellant thruster and 25 propellant technologies are separately known and developed for use in space, and therefore the potential problems in the technical, support and safety areas of the hybrid control system are minimized. Other features and advantages of the 30 invention will become apparent from the following more detailed discussion, taken in conjunction with the accompanying drawings, which illustrate, by way of example, the principles of the invention. 35 BRIEF DESCRIPTION OF THE DRAWINGS Figure 1 is a schematic view of a typical communications satellite in relation to the earth and two earth stations; Figure 2 is a schematic view of three types of satellite orbits in relation to the earth; Figure 3 is a schematic view of a communications satellite in geosynchronous orbit, illustrating different types of movement induced by operation of the thrusters; Figure 4 is a sectional elevational view of a communications satellite, illustrating the control system; and Figure 5 is a schematic diagram of the control system. DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT Spacecraft are often placed into orbit as satellites in approximately the desired a -titude and orbit by a launch vehicle such as a dedicated booster rocket or the reusable space shuttle. The ' attitude and orbit must often then be refined by the control system of the spacecraft, using the propulsion thrusters located on the spacecraft. With the passage of time, the attitude of the spacecraft can change slightly, but even a few degrees of misorientation may be sufficient to render the spacecraft ineffective in performing its mission unless the attitude is corrected. Similarly, small changes in the orbital station must also be corrected. Both attitude and station corrections are achieved using the control system of the spacecraft. Additionally, major changes in the orbit, such as moving to a higher orbit, must be achieved in some satellites through the use of an apogee motor having much greater thrust than the thrusters used for attitude control and stationkeeping. As the term is used herein, the control system of a spacecraft includes its propellant supply, the thrusters, and the associated propellant lines to provide propellant to the thrusters. It does not include the control apparatus required to command firings of the thrusters. Figure 1 depicts a satellite 10 in orbit about the earth, acting as a relay station. A first earth station 12 transmits a signal to the satellite 10, where the signal may be amplified and modified for transmission to the second earth station 14. To minimize the cost of the ground antennas 16 located at the earth stations 12 and 14, it is desirable to keep the satellite 10 at exactly a fixed point in space, termed its station, and with each of the satellite antennas 18 pointed exactly at the corresponding ground antennas 16. If the satellite 10 is allowed to move more than a. short ^distance from its fixed station, then the ground antennas 16 must be controllably movable in order to track the satellite, with associated cost increases. The orbit of the satellite 10 is one key factor in the ability to position the satell ' ite so that it appears to be fixed in space at its station. As illustrated in Figure 2, a satellite 10 may be launched into an equatorial orbit 20 lying in the equatorial plane of the earth, an inclined orbit 22, or a polar orbit 24 that takes it over the poles of the earth. A satellite 10 in an equatorial orbit 20 that is circular at an altitude of about 22,235 statute miles moves at just the right velocity to maintain an essentially stationary position above a selected point on the earth as the earth rotates, in a geosynchronous orbit. To a ground observer, such a satellite appears to be nearly motionless, and ground antennas 16 can be pointed permanently at the satellite .10 without the need for continual redirection. Because of the curvature of the earth, a satellite in geosynchronous orbit cannot communicate with points on the earth at latitudes greater than about 80 degrees, but this shortcoming is more than overcome by the reduced complexity of ground stations for the majority of communications. Although the use of a geosynchronous orbit provides the capability of placing a satellite at a fixed station in space, in practice a satellite in geosynchronous orbit tends to move slightly from its assigned station due to various forces. Geosynchronous satellites must therefore have control systems to maintain them continuously on station, and such a control system is the subject of the present invention. Other types of .satellites such as those in low earth orbit or in inclined orbits can also make use of the control system of the present invention, but the presently preferred embodiment is used in conjunction with geosynchronous communications satellites. Figure 3 depicts a satellite 10 in orbit about the earth, with superimposed coordinates corresponding to control functions. Three coordinates are shown, corresponding to an above/below direction 26, an east/west direction 28, and a north/south direction 30. The satellite 10 must maintain the proper attitude or angular orientation with respect to these coordinates, with respect to reference points such as the earth, certain stars, or other bodies, this angular orientation being termed its attitude. The satellite 10 must also stay fixed with respect to a desired location in the orbit, termed the station X. The spacecraft 10 may deviate from the desired station by being north or south (relative to the earth) of the desired station along axis 30, by being east or west (relative to the earth) of the desired station along axis 28, or by being at a greater or lesser altitude (relative to the earth) than the desired station along the axis 26. Attitude control is therefore understood to relate to the angular orientation of the satellite relative to the proper station X, while stationkeeping is understood to relate to a movement of the entire satellite 10 away from its proper station. Deviations of attitude or station can arise for a number of reasons, including solar pressure, gravitational forces of other bodies such as the sun or moon, drag, thermal effects, electromagnetic effects, imprecision in prior corrective maneuvers, or other external and internal factors. The use of spacecraft as satellites is closely linked to their economics, particularly the weight and . size of the spacecraft when launched and the expected, effective life of the satellite in its mission. It is desirable to reduce size and weight by reducing the amount of propellant carried on the spacecraft, but a reduction in the amount of onboard propellant used for attitude control and stationkeeping maneuvers limits the number of corrective maneuvers and thence the useful life of the satellite. It is therefore highly important to optimize the control system to avoid wasted propellant and minimize inefficiencies in the operation of the thrusters used in the control system. Figure 4 illustrates a satellite 10 in greater detail, emphasizing elements of its control system 32. The illustrated satellite 10 includes a body 34 that supports a platform 36. The body 34 is generally cylindrical and rotates about its cylindrical axis at a rate of about 60 revolutions per minute. This rotation causes the entire satellite 10 to behave somewhat as a large gyroscope and maintain a generally constant attitude or angular orientation relative to the axes 26, 28 and 30. The platform 36 does not rotate at the same rate as the body 34, but instead remains relatively fixed so that the satellite antennas 18 can continuously point at the ground antennas 16. The platform 36 is thus said to be despun to have only that rotation necessary to permit the satellite antennas to be properly pointed. The control system 32 is positioned within the body 34, and includes a propellant supply 38, thrusters 40, and propellant lines 42 that deliver propellant from the supply 38 to the thrusters 40. In Figure 4 the control system 32 is indicated generally, and Figure 5 presents a detailed schematic illustration of the manner of conducting propellant to the thrusters. Figure 5 . schematically illustrates the control . system 32 in accordance with the present invention. The control system 32 utilizes two propellants, contained in a fuel tank 44 and an oxidizer tank 46. The fuel presently preferred is hydrazine, having the chemical compo ' sition N2H4, which can be used either as a bipropellant fuel or a monopropellant. That is, hydrazine can be used to generate gas in a thruster by reacting the hydrazine with a proper oxidizer, or to generate hot gas in a thruster by passing the hydrazine over a catalyst to decompose it to nitrogen and hydrogen gasses. See U.S. Patents 3,871,828; 4,069,664; 4,324,096; and 4,490,972 by Ellion, et al. , whose disclosures are herein incorporated by reference. The oxidizer presently preferred is nitrogen tetroxide, having the chemical composition N2O4. Hydrazine and nitrogen tetroxide react spontaneously to generate hot gas, in the bipropellant thruster. Hydrazine decomposes exothermally in a catalytic bed such as iridium coated alumina oxide to produce hot gas, in a monopropellant thruster. The technology for using these propellants in thrusters is known, both for the fuel as a monopropellant and the fuel and oxidizer as bipropellants. The fuel from the tank 44 and the oxidizer from the tank 46 are delivered to the thrusters 40 through piping collectively termed the propellant lines 42. Lines are provided to conduct both fuel and oxidizer to the bipropellant thrusters, and fuel to the monopropellant thrusters. In the simplest system, the tanks 44 and 46 are typically pressurized with a sufficient pressure to deliver the fuel and oxidizer as needed, without the need for pumps. Optionally, propellant pumps can be provided for specific applications. Fuel only is delivered to a monopropellant thruster 48._ Fuel and oxidizer are delivered to a bipropellant thruster 50. In actual practice, there will be a plurality of monopropellant thrusters 48 and a plurality of bipropellant thrusters 50, located at selected locations on the satellite 10 to perform their intended functions of attitude control and stationkeeping. In a presently preferred embodiment, the fuel and oxidizer are also provided to an apogee motor 52. The apogee motor 52 is typically a much larger rocket engine than thrusters 48 and 50. It is sometimes provided as a solid rocket motor, or a cryogenic rocket motor that is fired only to establish the orbit at the beginning of the mission and then not used again. The use of the hybrid monopropellant and -bipropellant approach of the present invention is therefore optional as applied to- the apogee motor 52, and its use depends upon the mission and type of , spacecraft. In a preferred approach wherein the present invention is applied to a spacecraft in geosynchronous . orbit as a satellite, the monopropellant thrusters 48 are positioned for attitude control, for east-west stationkeeping, and for north-south stationkeeping. The bipropellant thrusters 50 are positioned only for north-south stationkeeping. The mode of using the thrusters varies with the stage in the life of the satellite, and the amount of oxidizer remaining. In initial operation of the control system 32 at an early stage of the life of the satellite, a set of radial monopropellant thrusters 54 are operated for attitude control and for east-west stationkeeping, for which a relatively small proportion of the total propellant is used. The bipropellant thrusters 50 are operated for north-south stationkeeping, for which the majority of the propellant. is used. The amount of fuel initially loaded into the fuel tank 44 is greater than the sum of the amount required to react with the oxidizer in the oxidizer tank 46, plus the "expected consumption" of fuel required to perform the attitude control and east-west stationkeeping functions of the radial monopropellant thrusters 54 during the period prior to exhaustion of the oxidizer in the tank 46. The "expected consumption" is defined as the amount of fuel required to operate the monopropellant thrusters during the period prior to exhaustion of the oxidizer initially loaded into the spacecraft. This expected consumption of fuel is a value which can be estimated during design of the satellite, and is usually based on data from prior operations of other satellites used for similar missions. The expected consumption cannot be determined exactly beforehand, but an advantage of the present approach is that, as long as an excess of fuel is provided, small errors _in estimating expected consumption do not result in supplies of unconsumed and unusable propellants. Of course, it is still desirable to estimate the expected consumption as closely as possible to maximize the period during which the efficient bipropellant thrusters are used for north-south stationkeeping maneuvers. At a later stage of the life of the satellite, after the oxidizer is depleted, the bipropellant thrusters 50 no longer can be operated for north-south stationkeeping, and a set of axial monopropellant thrusters 56 assume this function. The propulsion functions of the control system 32 therefore are accomplished jointly by the monopropellant thrusters 54 and the bipropellant thrusters 50 early in the life of the satellite ' , and solely by the monopropellant thrusters 54 and 56 during .the later stages of the life of the satellite 10. By this approach, all of the propellant is used to advantage, wasting none and leaving no potentially dangerous propellants in storage. In a most preferred version of the system just described, the apogee motor 52 is also a liquid fueled thruster of the bipropellant type, and operates from the common fuel tank 44 and oxidizer tank 46 prior to exhaustion of the oxidizer, according to the principles just described. In the preferred approach, the bipropellant thrusters 50 operate in their most efficient mode of long duration firings for north-south stationkeeping, the firings typically lasting 50-200 seconds.- The monopropellant thrusters 48 operate in pulsed or short duration firings over most of the life of the satellite, and consume a small fraction of the propellant. After the oxidizer used in the bipropellant thrusters 50 is expended and the axial monopropellant thrusters 56 assume the north-south stationkeeping functions, they are used in place of the more efficient bipropellant thrusters, for long duration firings for north-south stationkeeping. The overall control system efficiency is thereby improved through full use of the available propellants and extended life of the satellite. Thus, the amount of fuel initially loaded into the fuel tank 44 and provided for consumption by the hybrid monopropellant and bipropellant control system 32 is in all cases greater than the amount required to react with the oxidizer contained within the oxidizer tank 46. In the preferred approach, the amount of fuel is greater than the sum of the amount required to react with the oxidizer plus the amount required for the expected consumption of fuel, by the monopropellant thrusters . prior to the exhaustion of the oxidizer. It is known from prior geosynchronous communications satellites that approximately 1056 of the propellant is expended in pulsed discharges * for attitude control and east-west stationkeeping, while 9056 is expended on north-south stationkeeping, primarily in long bursts of 50 to 200 seconds duration for the correction of north-south drift. For a satellite expected to have this type of experience, and using the preferred embodiment of the invention wherein bipropellant thrusters are used for correction of north-south drift, the amount of fuel initially loaded into the fuel tank of the satellite is determined to be at least about 1056 (that is, an expected consumption value of 1056) greater ' than the amount of fuel required to react with all of the oxidizer initially loaded into the oxidizer tank of the satellite. Other arrangements are possible, and the precise fuel load would be chosen using a similar of calculation to that just discussed. It is projected that use of the present hybrid monopropellant and bipropellant control system can increase the efficiency of spacecraft significantly, thereby increasing their payloads. As an example, replacing an all-monopropellant system with the present hybrid control system results in a net reduction of the end-of-life weight of a Hughes Aircraft Co. type HS 376 class geosynchronous satellite, normally having a payload of 350 pounds, by as much as about 106 pounds. This weight savings allows an equal increase in the payload, thus achieving a 3056 increase in payload through improvement of the control system. As is now. apparent, the present invention represents an important ' advance in spacecraft * control systems such as used in geosynchronous communications satellites. Efficiency of the control system is improved, resulting in increased payload for the spacecraft. The propellants are fully consumed, leaving no wasted residual propellant that causes a weight penalty and might be dangerous. No major changes in spacecraft operating and preparation procedures are required. Although a particular embodiment of the invention has . been described in detail for purposes of illustration, various modifications may be made without departing fro -the spirit and scope of the invention. Accordingly, the invention is not to be limited except as by the appended claims.

Claims: CLAIMS What is claimed is: 1. A spacecraft control system, comprising: a bipropellant thruster operating from -a fuel and an oxidizer; a monopropellant thruster operating from the same type of fuel as said bipropellant thruster; a supply of oxidizer communicating with said bipropellant thruster; and a supply of fuel communicating with said bipropellant thruster and said monopropellant thruster, the amount of fuel being greater than the amount required to react with all of the oxidizer. 2. The control system of claim 1, wherein the fuel is hydrazine. 3. The control system of claim 1, wherein the oxidizer is nitrogen tetroxide. 4. The control system of claim 1, wherein the bipropellant thruster is a stationkeeping thruster. 5. The control system of claim 1, wherein the amount of fuel is greater than the sum of the amount required to react with all of the oxidizer plus the expected consumption of fuel in said monopropellant thruster prior to the exhaustion of all of the oxidizer. 6. The control system of claim 1, wherein the amount of fuel is greater than 110 percent of the amount required to react with all of the oxidizer. 7. The control system of claim 1, wherein the bipropellant thruster is an apogee motor. 8. The control system of claim 1, wherein at least two monopropellant thrusters are provided, and wherein one of the monopropellant thrusters is positioned for attitude control and east-west stationkeeping maneuvers, and one of the monopropellant thrusters is positioned for north-south stationkeeping maneuvers. 9. A method for controlling a spacecraft, comprising: providing the spacecraft with at least one bipropellant thruster and at least one monopropellant thruster; providing the spacecraft with a fuel storage tank communicating with the bipropellant thruster and the monopropellant thruster; providing the spacecraft with an oxidizer storage . tank communicating with the bipropellant thruster; loading the oxidizer storage tank with an amount of oxidizer; and loading the fuel storage tank with an amount of fuel in excess of the amount required to react with the available oxidizer in the oxidizer storage tank. 10. The method of claim 9, wherein the fuel is hydrazine. 11. The method of claim 9, wherein the oxidizer is nitrogen tetraoxide. 12. The method of claim 9, including the additional steps of: launching the spacecraft; operating the monopropellant and bipropellant thruster until the oxidizer is exhausted; and operating the monopropellant thruster after the oxidizer is exhausted. 13. The method of claim 9, wherein at least two monopropellant thrusters are provided, and wherein one of the monopropellant thrusters positioned for attitude control and east-west stationkeeping maneuvers, and one of the monopropellant thrusters is positioned for north-south stationkeeping maneuvers. 14. The method of claim-9, wherein the bipropellant thruster is provided in a position for north-south stationkeeping only. 15. The method of claim 9, wherein the amount of fuel is greater than the sum of the amount required to react with all of the oxidizer plus the expected consumption of fuel in the monopropellant thruster prior to the exhaustion of all of the oxidizer.

Inventors: DONATELLI PHILIP A, ELLION M EDMUND, DONATELLI, PHILIP, A., ELLION, M., EDMUND
Applicants: HUGHES AIRCRAFT CO, HUGHES AIRCRAFT COMPANY
Application year: 1987
Main classification: B64G1
IPC codes: B64G1, F02K9