A storage system for a two-phase propellant fluid to provide a simple passive propellant storage device for dispensing in gas-free form the liquid propellant for use typically in an auxiliary propulsion system in a three-axis stabilized spacecraft. The internal structure of the storage tank of the system is formed into three compartments which by capillary action provides for preferrential flow of a selected one of the two fluids, comprising, typically, a pressurant gas and a two-phase propellant liquid.
A spacecraft includes bipropellant and monopropellant engines or thrusters. The oxidizer-fuel mixture ratio of the bipropellant engine is not known exactly. The spacecraft is loaded with only sufficient oxidizer to achieve the velocity for transfer from an intermediate orbit to geosynchronous orbit if the mixture ratio is nominal. Therefore, more fuel can be loaded. If the bipropellant engine burn is nominal, there is no excess oxidizer when on-orbit and more fuel is available for stationkeeping. If the burn is oxidizer-rich, there is a velocity shortfall, which is made up by firing monopropellant engines. If the burn is oxidizer-lean, the geosynchronous orbit is achieved with a load of excess oxidizer, which must be moved during each stationkeeping maneuver. A net gain of stationkeeping time results in any of the three mixture ratio cases by comparison with loading of sufficient oxidizer for a full bipropellant burn under worst-case mixture conditions.
A closed-loop two-phase thermal control loop includes an evaporator which receives subcooled coolant liquid from a condenser and heat from a source, and vaporizes the coolant to form coolant vapor. A condenser which is coupled to a thermal radiator accepts the coolant vapor and radiates thermally to space, thereby condensing the coolant vapor to liquid and subcooling the liquid. As the temperature of the environment surrounding the condenser radiator changes, the mass of vapor within the system tends to change, thereby causing undesirable pressure changes. A spherical wicked coolant reservoir is coupled to the liquid side of the closed thermal control loop. The reservoir is exposed foro thermal radiation and electrically heated, so that the temperature is thereby controlled in order to control vapor pressure within the reservoir. By controlling the reservoir vapor pressure, small pressure differentials are generated, which cause coolant liquid to be accepted or expelled while maintaining the closed loop pressure. The reservoir includes a wick for maintaining liquid coolant adjacent heat transfer areas, and also includes a standpipe supporting shaped vanes. The coolant adheres to the vanes and to the wick under zero-gravity conditions in a manner which allows at least partial simulation in an Earth gravity.
A passive propellant management system for a spacecraft liquid propellant tank (1) comprises several preferably V-shaped channels (2) which communicate liquid propellant from regions within the tank (1) to an outlet port (8), which expels liquid propellant but not pressurant gas. A liquid/bubble chamber assembly (9) couples the channels (2) with the outlet port (8). The channels (2) comprise relatively open portions (11) and relatively closed portions (10). In the relatively open portions (11), liquid is retained in a gap (12) between open ends of the V channels (2) and the inner wall of the tank (1). In the relatively closed portions (10), a screen, mesh or perforated plate (14) covers the open end of the V channels (2), intermediate the V channels (2) and the inner wall of the tank (1). The placement of the relatively open and closed portions (11, 10, respectively) is intentionally preselected based upon mission requirements. Where pressurant gas ullage is expected to be present, e.g., during periods of high g, relatively open portions (11) are used. Where liquid propellant is expected to be present, e.g., during periods of relatively low g, relatively closed portions (10) are used. The liquid/bubble chamber assembly (9) comprises a liquid trap (27) and bubble trap (28), which operate synergistically with each other and with the channels (2) to provide optimum liquid flow during all phases of the spacecraft mission.
Apparatus for gaging the amount of liquid propellant remaining in a spacecraft (38) tank (1). The tank (1) comprises several preferably V-shaped channels 2A, 2B, 2C, 2D, hereinafter collectively referred to as "2" which communicate liquid propellant from regions within the tank (1) to an outlet port (8) that expels liquid propellant but not pressurant gas. A liquid/bubble chamber assembly (9) couples the channels (2) with the outlet port (8). The channels (2) comprise relatively open portions (11) and relatively closed portions (10). A first ultrasonic transducer (31) is positioned along the outside of the tank (1) and detects a condition of near propellant depletion by measuring the amount of liquid retained in the relatively open portions (11) of the channels (2). The liquid/bubble chamber assembly (9) comprises a liquid trap (27) and a bubble trap (28). A second ultrasonic transducer (32) is positioned along the outside of the tank (1) and detects a condition of near propellant depletion by measuring the amount of liquid propellant remaining within the bubble trap (28).
A spacecraft propulsion system which integrates the function of the apogee kick motor (AKM) and reaction control system (RCS) is disclosed. In accordance with this invention, a pump-fed AKM is employed which results in lightweight main tanks and pressurization systems. The RCS thrusters are operated by small bellows tanks which are intermittently pressurized by a gas pressurization system to provide high pressure for operation of the RCS thrusters. The system according to this invention enables use of lighter weight main propellant tanks since they do not have to withstand high internal pressures and also enables realization of the numerous advantages of a pump-fed AKM. Several embodiments describe various methods for cycling the bellows tanks.
