Abstract: A 6-degree-of-freedom control apparatus for a spacecraft includes a plurality of thrusters, thruster modulator, position/velocity detector, target position/velocity generator, attitude/angular velocity detector, target attitude/angular velocity generator, and 6-degree-of-freedom controller. The thrusters control three position axes and three attitude axes of a spacecraft by jet. The thruster modulator selectively drives the thrusters on the basis of a thruster control signal. The position/velocity detector measures a position and velocity of the spacecraft. The target position/velocity generator generates target position and velocity values of the spacecraft. The attitude/angular velocity detector measures an attitude and angular velocity of the spacecraft. The target attitude/angular velocity generator generates target attitude and angular velocity values of the spacecraft.

Abstract: A simple, robust algorithm for power acquisition for power, thermal, and momentum safety for high heater-power spacecraft uses current sensors rather than sun sensors and includes a wing sun search phase, an xz slew phase, and a safe hold phase. Solar wing current is continuously monitored against a high current threshold and a low current threshold. Wing sun search phase transitions either to xz slew phase or safe hold phase. When current is too low, the xz slew phase is entered and slews the spacecraft about an axis in the xz plane until current is high enough. When current is high enough, the algorithm transitions or remains in safe hold phase and the spacecraft is spun about its wing axis. A low current persistence timer is longer than a high current persistence timer in order to bias the algorithm to prefer safe hold phase over xz slew phase.

Abstract: A hand-held electronic celestial object-locating device assists in identifying a celestial object or directing a user to a desired celestial object. The device is useful for locating or identifying any celestial object including stars, constellations, planets, comets, asteroids, artificial satellites, and deep sky objects to name a few. The device utilizes sensors for 3-axis magnetic field and 3-axis gravitational field detection. The device utilizes a processor and an electronic database to perform the required calculations. The device's database may be updated through access to the Internet through which the updates may be purchased.

Abstract: A head up display system that incorporates conformal and non-conformal views and associated symbology to provide highly informative and intuitive guidance with respect to all aspects of operating an aircraft or other vehicle in a controlled geographical area, by utilizing calculated views from the pilot seat, aircraft speed, and relevant ground operation information.

Abstract: A method and apparatus for post processing a star tracker measurement to remove a systematic error characterizable at least in part by a pixel phase is disclosed. The method comprises the steps of computing the pixel phase along a first axis from a measured star position and a star tracker characteristic, computing a first axis error correction according to the computed pixel phase, and computing a compensated first axis star tracker measurement according to the measured star position and the first axis error correction.

Abstract: The invention relates to the field of radio navigation and, more specifically, it relates to methods of determining the current values of the orbital position of the artificial satellite of the Earth which is used in a satellite radio navigation system (SRNS), by the ephemeredes and can be used when performing the radio navigation measurements in the equipment of the consumers of the SRNS signals.

Abstract: A method and apparatus for calibrating rate sensor measurements to compensate for rate sensor scale factor non-linearities. An outer loop compensator compares current rate sensor scale factor estimates with the current scale factor non-linearity compensation, and deviations from the current non-linearity compensation are corrected in an updated compensation.

Abstract: A vehicular position and attitude can be accurately determined from GPS signals by positioning a master and two slave orthogonally disposed receiver antennas rigidly mounted together enabling the rotation of the slave receiver antennas about the master antenna. The slave antennas are rotated and then dither back and forth so that differential phase measurements between the slave antennas and the master antenna is nulled in carrier phase alignment of the GPS signals for determining azimuth angles and elevation angles to the GPS satellites for determining the position and attitude of the vehicle.

Abstract: A method for controlling a spacecraft comprising the steps of providing a first spacecraft in a known first predetermined orbit, and a second spacecraft in a second predetermined orbit. The distance between the first and second spacecraft is measured. The measured distance and data describing the known first predetermined orbit are used for determining an orbital error bias of the second spacecraft relative to the second predetermined orbit. The second spacecraft is maneuvered to compensate for the orbital error bias, and to maintain the second spacecraft in the second predetermined orbit.

Abstract: A method of controlling the attitude of a vehicle in space relative to three reference axes, the method consisting: initially in adjusting the attitude of the vehicle in conventional manner to desired angles on the basis of a two-axis detector which points towards a reference heavenly body; and subsequently in adjusting the attitude of the vehicle about the third reference axis by using the speeds of rotation of the reaction wheels which reflect conservation of the total angular momentum of the vehicle to perform an angle measurement about the third axis.

Abstract: A parent unit is provided having: a precedent base for accumulating data from child units; an element data determination unit for processing the data into element data; an element data classification unit for creating a classification; an element data change classification unit for determining which classification the element data belongs to, and calculates transition probabilities among the classifications, and a display unit for displaying data. The child unit collects air current data, processes the data, obtains the transition probabilities, and calculates a probability of encountering air turbulence. The child unit then displays certain data when the probability of encountering air turbulence satisfies predetermined conditions.

Abstract: A method of determining the position of a satellite in a near geosynchronous orbit includes receiving at least one main lobe signal from an antenna on a first GPS satellite. A GPS signal, including GPS time and Doppler, is received from a pseudolite positioned on the Earth and an approximate position of each of a plurality of second GPS satellites is determined from an onboard almanac. Side lobe signals and accompanying noise are received from antennas on the plurality of second GPS satellites. The GPS signal and known sequential data bits are used for sorting or integrating the side lobe signals from the accompanying noise and the position of the satellite in a near geosynchronous orbit is determined using the one or more main lobe signals and the sorted side lobe signals.

Abstract: A self-monitoring satellite system is disclosed to more quickly and accurately convey satellite signal information and verification of its reliability to users of the signal. In a first aspect, the satellite includes one or more receivers on the satellite to monitor the reliability of signals transmitted from the satellite. In a second aspect, a warning signal is generated and transmitted substantially simultaneous with an unreliable satellite signal and at the same frequency as the satellite signal. Exemplary embodiments to implement the various aspects of the self-monitoring satellite system invention are configured as global positioning system (GPS) satellites, signals, and receivers.

Abstract: A method for controlling a spacecraft comprising the steps of providing a first spacecraft in a known first predetermined orbit, and a second spacecraft in a second predetermined orbit. The distance between the first and second spacecraft is measured. The measured distance and data describing the known first predetermined orbit are used for determining an orbital error bias of the second spacecraft relative to the second predetermined orbit. The second spacecraft is maneuvered to compensate for the orbital error bias, and to maintain the second spacecraft in the second predetermined orbit.

Abstract: A spacecraft attitude control system that enables friction and stiction compensation for attitude control of a spacecraft using at least one reaction wheels that can spin through zero rpm without degradation in pointing. The system removes nuisance disturbances relating to friction or stiction that detract from the ability of the system to maintain or change spacecraft attitude. It can remove these effects before they are sensed by the systems attitude control sensors. A friction or stiction compensation buffer is arranged between the spacecraft attitude control subsystem and the torque actuators for the at least one reaction wheel. Since the compensation is localized to each wheel, it is possible to optionally: control wheel operating speeds to safe levels, provide wheel failure flags, provide accurate wheel tracking at slow or zero wheel speeds and provide full access to the wheel torques.

Abstract: A method for acquiring a satellite using satellite prediction values has a behavior analysis step for analyzing a behavior unique to a satellite orbit, a search range calculation step for calculating a search range based on the analysis, and a search step for searching ranges in accordance with the predicted values.

Abstract: A method, apparatus, article of manufacture for determining the attitude of a spacecraft having at least one star sensor. The method comprises the steps of selecting a first reference star sensor from among the star sensors; designating two stars observed by the first reference star sensor as a primary star pair; identifying a candidate star pair corresponding to the primary star pair, wherein the candidate star pair is selected from a star catalog having a plurality of entries; estimating an inertial orientation of the first star sensor at least in part from the identified candidate star pair; and determining the spacecraft attitude from the estimated inertial orientation of a reference star sensor selected from a group comprising the first star sensor.

Abstract: An attitude detection system for an artificial satellite includes an interpolator for interpolating an angular-velocity signal to generate an interpolated angular-velocity signal, a sequential Kalman filter for generating a low-frequency attitude-angle signal, and an adder for adding the low-frequency attitude-angle signal and a high-frequency attitude-angle signal generated by a high-frequency angular sensor to generate a broad-band attitude-angle signal.

Abstract: A diversity satellite system using a trailer satellite concept. The main satellite is placed into orbit and the trailer satellite, typically a much smaller satellite, is in orbit with the main satellite. The main satellite and trailer satellite are communicating with one another by crosslinks. In this way, two alternative geometric configurations are possible: a first geometric configuration direct from the ground station to the first satellite, and a second geometric configuration to the main satellite from the ground station via the trailer satellite. This enables avoiding interference.

Abstract: Apparatus and methods for performing satellite proximity operations such as inspection, recovery and life extension of a target satellite through operation of a “Satellite Inspection Recovery and Extension” (“SIRE”) spacecraft which can be operated in the following modes (teleoperated, automatic, and autonomous). The SIRE concept further consists of those methods and techniques used to perform certain (on-orbit) operations including, but not limited to, the inspection, servicing, recovery, and lifetime extension of satellites, spacecraft, space systems, space platforms, and other vehicles and objects in space, collectively defined as “target satellites”. The three basic types of SIRE proximity missions are defined as “Lifetime Extension”, “Recovery”, and “Utility”. A remote cockpit system is provided to permit human control of the SIRE spacecraft during proximity operations.

Abstract: Described is a device for determining the attitude on board of a satellite (1) with a data transfer device for carrying out optical data exchange between at least two satellites (1, 3); a device for determining one's own attitude; a device for determining the directional vectors in an inertial co-ordinate system from one's own attitude and from attitude data of at least two satellites (1, 3), said attitude data having been transmitted by means of the data transmission device; a device for determining the directional vector directed to the satellites (1, 3) of which there are at least two, in a fixed-body co-ordinate system of the satellite; and a device for attitude determination from the directional vectors determined. Also described is a method for carrying out attitude determination.

Abstract: A method for controlling spacecraft velocity. A thruster torque gimbal angle command is calculated to provide a desired thruster torque demand command including a desired momentum adjust torque and a desired attitude control torque. A velocity change error between a commanded velocity change and an actual velocity change, a commanded force vector based on the error to cause the actual velocity change to approach a commanded velocity change, and force gimbal angle commands and force thrust commands to achieve the commanded force vector are calculated; and/or an error between gimbal angles and thrust levels and reference gimbal angles and thrust levels, reference gimbal angle commands and reference thrust commands to drive actual gimbal angles and thrust levels toward reference gimbal angles and reference thrust levels are calculated. A total gimbal angle command is calculated by adding the torque gimbal command with at least one of the force gimbal angle command and the reference gimbal angle command.

Abstract: A system for analyzing aircraft data includes the steps of identifying a domain comprising sets of data, calculating ranges of typical values for components of the sets of data within the domain, and determining atypical components for each set of data within the sets of data based on the range of typical values calculated in the calculating step. A pre-cursor pattern to an exceedance event is determined by (a) identifying exceedance sets of data within the domain corresponding to an exceedance event, and (b) identifying common atypicalities of the exceedance sets of data. Mitigating factors for an exceedance event are determined by (a) identifying non-event sets of data within the domain in which atypicalities resembling the pre-cursor pattern exist but in which the exceedance event does not occur, and (b) identifying common atypicalities of the non-event sets of data as the mitigating factors.

Abstract: An attitude control system comprising a controller and a noise screen device coupled to the controller. The controller is adapted to control an attitude of a vehicle carrying an actuator system that is adapted to pulse in metered bursts in order to generate a control torque to control the attitude of the vehicle in response to a control pulse. The noise screen device is adapted to generate a noise screen signal in response to the control pulse that is generated when an input attitude error signal exceeds a predetermined deadband attitude level. The noise screen signal comprises a decaying offset signal that when combined with the attitude error input signal results in a net attitude error input signal away from the predetermined deadband level to reduce further control pulse generation.

Abstract: An attitude control system uses a first order partial derivative of an error signal, representing a difference between a commanded pitch attitude and a sensed, actual pitch attitude. The first order partial derivative forms a cross product with the Earth's magnetic field to provide a X axis torque rod dipole command, which powers the torque rod in a way that causes a satellite to move toward its commanded attitude.

Abstract: A system and method for correcting pointing errors induced by an inclined orbit (14) for a satellite (20) that is adapted to be in an orbit having zero inclination (12). The present invention generates a yaw correcting cosine rate command (34) that is integrated (36) by a steering law having input parameters dependent upon the angle of inclination, &THgr;, to derive a sinusoidal yaw offset. The present invention also generates a roll correcting cosine rate command (54) that is integrated (56) according to a steering law having input parameters dependent upon the angle of inclination, &THgr;, to derive a sinusoidal roll offset.

Abstract: A method is provided for determining position integrity in a system having a Global Navigation Satellite System (GNSS) component, such as, for example, a Global Positioning System (GPS) device. For successive alarm limits, with each alarm limit corresponding to a position integrity level, it is determined whether valid position integrity information is available. At the alarm limit at which valid position integrity information is first available, a corresponding position integrity level is determined. If no valid position integrity information is available for any of the alarm limits, a default position integrity level is then designated. An associated apparatus, system, and computer software program product are also provided.

Abstract: This invention provides an offline attitude determining apparatus capable of detecting an artificial satellite attitude at a high precision in wide band. Averaging and addition processing are carried out on a low-frequency attitude interpolation signal generated by interpolating a low-frequency attitude detection signal obtained by inputting an attitude angle determination signal generated by a sequential Kalman filter computation device and a high-frequency attitude detection signal obtained by inputting a high-frequency attitude angle sensor detection signal extracted by a high-frequency attitude angle sensor data extracting device into a band pass filter computation device, so as to generate a high-precision attitude detection signal. By combining the low-frequency attitude detection signal with the high-frequency attitude detection signal with an interpolation processing computation device and an attitude data generator, the attitude can be determined at a high precision in wide band.

Abstract: An attitude determination and alignment method and system use electro-optical sensors and global navigation satellites to determine attitude knowledge for a spacecraft, satellite, or a high-altitude aircraft. An on-board inertial navigation system uses global navigation satellite system equipment and an attitude determination system uses an electro-optical sensor. The electro-optical sensor view the navigation satellites as surrogate stellar reference sources. The electro-optical sensor replaces the function of a star sensor or tracker and associated processing required for an onboard attitude determination system. Navigation and timing information generated by the GPS/GNSS-INS is used to perform required attitude determination system functions.

Abstract: A system (30) for adjusting the orientation of a spacecraft adapted for use with a satellite (10). The system (30) includes a first control circuit (32, 38, 40) for canceling any momentum of the spacecraft via a counter-rotating spacecraft bus (16, 18). A second controller (32, 42, 44, 46, 48) orients the spacecraft via the application of internal spacecraft forces. In a specific embodiment, the spacecraft bus (16, 18) serves a dual use as storage section and includes a mass (16) having a moment of inertia on the same order as the moment of inertia of the satellite (10). The satellite (10) includes a bus section (16) and a payload section (14). The mass (16) includes the bus section (16). The first control circuit (32, 38, 40) runs software to selectively spin the mass (16) to cancel the momentum of the satellite (10). The software computes an actuator control signal, via a computer (32), that drives a first actuator (38) that spins the mass (16).

Abstract: A control system for providing attitude control in spacecraft. The control system comprising a primary attitude reference system, a secondary attitude reference system, and a hyper-complex number differencing system. The hyper-complex number differencing system is connectable to the primary attitude reference system and the secondary attitude reference system.

Abstract: Systems and methods that limit the effects of actuator saturation to certain body axes of a spacecraft. A plurality of actuators, or momentum wheels, disposed on the spacecraft are oriented non-parallel to x, y, and z axes of the spacecraft. The plurality of actuators each produce torque that is applied to the spacecraft. Outputs of the plurality of actuators or momentum wheels are processed to control the respective actuators or momentum wheels to manage actuator saturation so that saturation does not produce attitude errors about one or more predetermined axes of the spacecraft.

Abstract: A satellite positioning-based guidance system and method that utilize a simulated inertial navigation system are disclosed. The trajectory of the body of a non-thrusted flight vehicle is used instead of a conventional inertial navigation system. The attitude of the flight vehicle is determined using a satellite receiver, and the Extended Kalman filter is used to implement the navigation function.

Abstract: An easily implementable satellite constellation stationkeeping algorithm for maintaining the relative positioning among satellites of a low earth orbit constellation with minimal maneuvering while maintaining the constellation altitude over the effects of atmospheric drag. The satellite constellation stationkeeping algorithm may be implemented as a system or method that uses atmospheric drag perturbations to extend the stationkeeping cycle of the satellites in the low earth orbit constellation while simplifying the implementation algorithm and performing minimal maneuvers. In an exemplary satellite constellation stationkeeping algorithm, a plurality of satellites or a ground station are configured with a controller that implements an absolute altitude control algorithm. The plurality of satellites are launched into a respective plurality of slots of a low earth orbit.

Abstract: A controller and control method that perform satellite orbit-keeping maneuvers and proportionally scale orbit-keeping pulses to automatically minimize disturbance torques on-board a satellite. The controller and control method may also be used to remove residual momentum stored in spinning momentum wheels.

Abstract: A method generates an operational ballistic capture transfer for an object emanating substantially at earth or earth orbit to arrive at the moon or moon orbit using a computer implemented process. The method includes the steps of entering parameters including velocity magnitude VE, flight path angle &ggr;E, and implementing a forward targeting process by varying the velocity magnitude VE, and the flight path angle &ggr;E for convergence of target variables at the moon. The target variables include radial distance, rM, and inclination iM. The method also includes the step of iterating the forward targeting process until sufficient convergence to obtain the operational ballistic capture transfer from the earth or the earth orbit to the moon or the moon orbit.

Abstract: A method is provided for determining position integrity in a system having a Global Navigation Satellite System (GNSS) component, such as, for example, a Global Positioning System (GPS) device. For successive alarm limits, with each alarm limit corresponding to a position integrity level, it is determined whether valid position integrity information is available. At the alarm limit at which valid position integrity information is first available, a corresponding position integrity level is determined. If no valid position integrity information is available for any of the alarm limits, a default position integrity level is then designated. An associated apparatus, system, and computer software program product are also provided.

Abstract: A method for generating a horizontal path for the avoidance of danger zones for an aircraft. The method models the contours of each danger zone by a succession of segments demarcated by points. The method determines two homing circles and two capture circles passing respectively through the initial point and final point which are tangential respectively to the initial route and to the final route and have respectively the initial and final turning radius. The method also determines tangents both to the homing circles or the capture circles and to the contour of each danger zone. Among these tangents, a pair of tangents to a homing circle and to a capture circle are selected, defining a path skeleton connecting a homing circle to a capture circle without meeting a danger zone. And, an automatically controllable evasion path that lies on the previously defined path skeleton is then determined.

Abstract: The invention provides GPS navigational system for a satellite in space which comprises a front end section having an input for receiving GPS signals from a plurality of satellites, a digital preprocessor connected to the front end section for digitally preprocessing the GPS signals received from the front end section, and a signal processor connected to the digital preprocessor for decoding the GPS signals to determine a position of the satellite. An on-board computer is provided and a first data bus line connects the data processor and the signal processor for bi-directional data exchange therebetween and a second data bus line connects the signal processor and the on-board computer for bi-directional data exchange therebetween.

Abstract: A method and corresponding apparatus and system for determining the position of a satellite at any instant of time between two sampling instants, based on information in at least one ephemeris message provided by the satellite, the information allowing the computation of the satellite positions at the two sampling instants. The method includes the steps of: computing the satellite positions at two sampling instants using parameters provided in one or more ephemeris messages with suitable times of ephemeris; and for each of the three dimensions of motion, constructing a polynomial of at least third degree by choosing the coefficients of the polynomial so that it interpolates the satellite position at each of the two sampling instants.

Abstract: A system for determining the heading and/or attitude of a body receives radio waves from a plurality of position-fixing satellites using at least three antennas fixedly mounted at different positions of the body. To reliably obtain integer ambiguity solutions of carrier phases of the radio waves in a shorter time, the system directly determines integer ambiguities from attitude angle data obtained by an IMU attitude processing section when the integer ambiguities are to be redetermined in the event of an interruption of the received radio waves or a change in the combination of satellites to be used. This system provides a user with highly accurate uninterrupted heading and/or attitude angle information.

Abstract: A system and method for a group of artificial satellites. The artificial satellite is placed into an orbit in which an individual satellite orbits on an elliptical orbit so that at least one of the artificial satellites is always viewable within a pre-defined range of operational elevational angle in a zenith direction from a service area. The group of the artificial satellites are satellites on respective different orbits obtained by combining an inclination angle and an eccentricity squared of the elliptical orbit so that a time period during which one artificial satellite of the group of artificial satellites is viewable from ground is substantially identical to a time period during which another artificial satellite of the group of the artificial satellites is viewable from ground.

Abstract: An apparatus for regulating or controlling a satellite includes: at least one sensor for detecting reference targets such as stars, which is connected to arrangements for selecting and determining the position of reference targets, and to an arrangement for estimating model parameters of the satellite and of external interferences; a memory for storing a reference target catalog and a reference trajectory; a regulating unit for regulating the position of the satellite; and a filter that is connected to the arrangement for position determination of the reference targets, the arrangements for estimating the model parameters, the memory and the regulating unit.

Abstract: A fixed station and an on-board apparatus are constructed to carry out a short range communication. The on-board apparatus specifies an application from a command included in reception data received from the fixed station and starts a corresponding processing, when an acquired command is a menu selection command. A menu including commodities and charges are displayed. When a selected item is determined, the determined content is displayed. When an IC card unit receives an IC card for settling the charge, the charge is settled by rewriting the IC card by data exchange via wireless communication. A short range wireless unit is used to ensure individuality of the communication. This short range communication may also be used at a vehicle repair shop, at a curved corner on a travel path or the like.

Abstract: A 6-degree-of-freedom control apparatus for a spacecraft includes a plurality of thrusters, thruster modulator, position/velocity detector, target position/velocity generator, attitude/angular velocity detector, target attitude/angular velocity generator, and 6-degree-of-freedom controller. The thrusters control three position axes and three attitude axes of a spacecraft by jet. The thruster modulator selectively drives the thrusters on the basis of a thruster control signal. The position/velocity detector measures a position and velocity of the spacecraft. The target position/velocity generator generates target position and velocity values of the spacecraft. The attitude/angular velocity detector measures an attitude and angular velocity of the spacecraft. The target attitude/angular velocity generator generates target attitude and angular velocity values of the spacecraft.

Abstract: A self contained electronic system for manual or automatic control and navigation of fixed winged aircraft using electronic position sensing such as GPS, DGPS, WAAS, and the like, as the primary sensor and making use of known flight characteristics of the aircraft to determine aircraft attitude without any interaction with the aircraft, its controls, or the outside environment and without any moving mechanical devices other than switches, dials and connectors. The automatic and visual interface between the system and the pilot provides for simplified flight controls, and a new solution to the hazard of disorientation, and will reduce the time needed for a pilot to become proficient in VFR and instrument flying. A single instrument replaces many of the conventional instruments used for flight. Navigation data is provided in an easy to understand graphical format. The pilot is told explicitly where to move aircraft controls.

Abstract: Apparatuses, methods, and systems of transferring correction information are described. In certain implementations, correction factors relate to stored or calculated values, and a correspondence between a correction factor and a value is indicated by a predetermined order of the correction factors. In one application, a method according to an embodiment of the invention is used to transmit correction factors relating to the positions of physical objects. For example, such a method may be used to transmit correction factors relating to the positions of space vehicles within a Global Positioning Satellite (‘GPS’) system.

Abstract: A method and corresponding apparatus and system for determining the position of a satellite at any instant of time between two sampling instants, based on information in at least one ephemeris message provided by the satellite, the information allowing the computation of the satellite positions at the two sampling instants. The method includes the steps of: computing the satellite positions at two sampling instants using parameters provided in one or more ephemeris messages with suitable times of ephemeris; and for each of the three dimensions of motion, constructing a polynomial of at least third degree by choosing the coefficients of the polynomial so that it interpolates the satellite position at each of the two sampling instants.

Abstract: A system of inclined geosynchronous satellite orbits has a service area defined on a surface of the earth. The service area has elevation angles greater than a predetermined minimum elevation angle from the horizon. A satellite has an orbit with respect to the earth having a sky track when viewed from within said service area. An operating arc is defined by a subset of points on the sky track within the service area. The satellites operate consecutively on the operating arc.

Abstract: A technique for transferring an object such as a spacecraft to one of the stable Lagrange points, for instance L4 or L5, utilizes a substantially negligible amount of delta-V. In doing so, payload amounts may be increased. Initially, a modified weak stability boundary transfer with parameters sufficient to transfer the spacecraft from a first heavenly object or a first heavenly object orbit to a vicinity of a second heavenly object is performed. Then, the spacecraft is momentarily captured at a capture point located in the vicinity of the second heavenly object. Upon capture, a maneuver is executed at the capture point to target the stable Lagrange point utilizing a substantially negligible amount of propellant. Finally, the spacecraft arrives at the stable Lagrange point.