Data processing: vehicles – navigation – and relative location – Vehicle control – guidance – operation – or indication – Aeronautical vehicle
Reexamination Certificate
2002-01-17
2003-10-07
Nguyen, Tan Q. (Department: 3661)
Data processing: vehicles, navigation, and relative location
Vehicle control, guidance, operation, or indication
Aeronautical vehicle
C701S003000, C244S172200, C060S228000, C060S250000
Reexamination Certificate
active
06631314
ABSTRACT:
FIELD OF THE INVENTION
The invention is related to the field of space vehicles for use in launching a payload from a stationary ground-based position into orbit, and specifically to a propellant utilization system for controlling depletion of multiple propellant sources connected to multiple thrust sources on the space vehicle.
BACKGROUND OF THE INVENTION
Rocket powered space vehicles used in space missions can generally be categorized into launch vehicles and payloads. Launch vehicles provide the primary thrust for launching and delivering a payload from the earth's surface into orbit. Launch vehicles generally include one or more rocket engines arranged to fire at different times, or stages, as the launch vehicle travels from the earth's surface into orbit. The different stages are fired sequentially and typically include at least a first stage or booster stage and a second stage or upper stage. The booster stage is designed to launch and deliver the payload a pre-determined distance above the earth before exhaustion. Upon exhaustion of the booster stage, the upper stage is fired to deliver the payload the remainder of the distance into a desired orbit.
The launch vehicle rockets may use either a solid propellant or a liquid propellant and typically include at least one propellant tank, a combustion chamber, and a nozzle for accelerating/discharging the combustion product. Liquid propellant rockets are known in the art as bi-propellant rocket systems because a liquid fuel and liquid oxidizer are stored in separate propellant tanks and brought into contact in the combustion chamber to provide the thrust. Such bi-propellant rocket systems have gained favor for many applications because of performance, economics, safety, throttling capabilities and flexible mission design.
Propellant utilization systems are used on a “per stage basis” (i.e., each stage has its own independent propellant utilization system) to maximize the efficiency of its corresponding bi-propellant rocket system by controlling the mixture ratio at which the liquid fuel and oxidizer are combined. These systems calculate mixture ratios according to desired exhaustion characteristics for the liquid fuel and oxidizer propellants. For example, in some cases, it is desirable to simultaneously exhaust both the liquid fuel and the oxidizer propellants to minimize the amount of one propellant remaining at the actual depletion of the other propellant. In other cases, however, it is desirable to exhaust one of the propellants before the other one of the propellants.
Propellant utilization systems include sensors located in the liquid fuel and oxidizer propellant tanks as well as software to compute a remaining amount of propellant in each tank and an engine mixture ratio for the engine controls that achieves the desired exhaustion rate. Unfortunately, however, bi-propellant rocket engines have proprietary operational characteristics that require the use of different flight parameters to calculate remaining propellant, mixture ratios, and exhaustion rates. As a result, each rocket engine on a space vehicle is equipped with its own proprietary propellant utilization system that is programmed with the flight parameters for that engine. Additionally, the propellant utilization for one engine cannot be used on another engine.
SUMMARY OF THE INVENTION
The present invention generally relates to rocket powered space vehicles that have bi-propellant thrust sources or rocket engines. The propellant utilization system described herein, however, may be appropriate for use on any bi-propellant engine that includes mixture ratio control capability. In the context of the present invention, the term “bi-propellant” is defined as having at least two sources of a fuel or fuel component, including without limitation a liquid fuel, an oxidizer fuel, and/or multiple liquid, solid, or gaseous fuels.
In view of the foregoing, a primary object of the present invention is to provide a single propellant utilization system that accommodates one or more bi-propellant thrust sources or rocket engines. Another object of the present invention is to provide a propellant utilization system that is easily adaptable for operation with any bi-propellant thrust source. Yet another object of the present invention is to provide a propellant utilization system for space vehicles equipped with multiple rocket engines that sequences with the firing of the different engines to generate and provide mixture ratios for the active engine.
One or more of the above-noted objectives, as well as additional advantages, are provided by the present invention, which includes a propellant utilization system for a space vehicle having at least a first thrust source and a second thrust source. The present propellant utilization system utilizes a common set of algorithms to generate mixture ratios for the individual thrust sources as the thrust source becomes active during a flight.
According to a first aspect of the invention, a propellant utilization system is provided that includes a processing system comprising sequencer logic, propellant logic, and mixture ratio logic. The sequencer logic determines when a thrust source is active, e.g. one of the first and second thrust sources, and provides flight parameters for the active thrust source to the propellant logic and the mixture ratio logic. The propellant logic processes information from a pair of propellant sources connected to the active thrust source, using the flight parameters for that thrust source, to determine an amount of propellant in each source. The mixture ratio logic generates a mixture ratio for the active thrust source using the flight parameters for that thrust source and information on the amount of propellant in the pair of propellant sources connected to the active thrust source.
Various refinements exist of the features noted in relation to the subject first aspect of the present invention. Further features may also be incorporated in the subject first aspect of the present invention as well. These refinements and additional features may exist individually or in any combination. For example, the above-noted propellants for each thrust source are typically a liquid fuel propellant and a liquid oxidizer propellant, and the above-noted thrust sources are each typically rocket engines. The mixture ratios for the active thrust source, e.g. rocket engine, may result in the substantial simultaneous depletion of the liquid fuel and oxidizer propellants for the active thrust source. Alternatively, the mixture ratios for the active thrust source may result in the depletion of one of the liquid fuel or oxidizer propellants before the other one of the liquid fuel or oxidizer propellants.
The information provided to the propellant logic from the propellant sources connected to the active thrust source in the case of the first aspect could be any information that is usable to determine an amount of propellant in the propellant sources. In one embodiment, the information from the propellant sources is the pressure information of the liquid fuel propellant in the liquid fuel propellant source and the pressure information of the oxidizer propellant in the oxidizer propellant source. The propellant logic uses the pressure information to generate a mass of propellant in each of the propellant sources. In some embodiments of the present invention, the propellant logic may also use the mass information to generate a difference error representative of the difference between the amount of liquid fuel propellant and the amount of oxidizer propellant relative to a depletion rate of the liquid fuel and oxidizer propellant. In another embodiment of the present invention, the mixture ratio logic may use the mass information to generate the difference error between the amount of liquid fuel propellant and the amount of oxidizer propellant.
According to a second aspect of the present invention, a software product for a propellant utilization system is provided. The software product includes sequencer logic instructions, p
Fisher Jay Andrew
Wurst Joseph Patrick
Lockheed Martin Corporation
Marsh & Fischmann & Breyfogle LLP
Nguyen Tan Q.
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