Power plants – Reaction motor – Method of operation
Reexamination Certificate
2000-11-13
2002-10-01
Kim, Ted (Department: 3746)
Power plants
Reaction motor
Method of operation
C060S257000, C417S016000, C417S423600, C417S362000
Reexamination Certificate
active
06457306
ABSTRACT:
FIELD OF THE INVENTION
The present invention relates in general to systems for pumping a liquid propellant from a propellant tank to a combustion chamber of a rocket engine and, in particular, to an electrical drive system for driving a liquid propellant pump for certain high energy, limited time applications. The invention is particularly applicable for pumping propellant to the combustion chamber of a launch vehicle during a launching thrust time interval.
BACKGROUND OF THE INVENTION
Rocket engines are used in a variety of applications including a number of types of spacecraft thrust systems. Such thrust systems include launch vehicle thrust systems and impulse engines. Launch vehicle thrust systems are used to provide the primary thrust required for inserting a spacecraft into orbit. The thruster systems may be provided in stages that are fired in sequence, with initial stages being discharged when they are exhausted. Each such stage may include one or more rocket engines, i.e., one or more combustion chambers associated with a nozzle or other structure for expelling a combustion product to produce thrust. Each rocket engine of a launch vehicle is generally operated at a high combustion chamber pressure and a high mass flow rate continuously for a limited time duration. In this regard, the launch thrusting period in which such a launch vehicle rocket engine is operated may last, for example, for about 1 to 8 minutes. In order to provide the necessary continuous and high mass flow of propellant from a propellant tank to the combustion chamber during this period a rotating impeller-based pump is typically employed.
By contrast, impulse engines are typically used for a variety of in-orbit, pulsed applications. For example, impulse engines may be operated in-orbit in a closely controlled manner to provide small velocity changes for orbit correction or transfer. Similarly, such impulse engines may be operated in-orbit for attitude control, e.g., to maintain a desired orientation for solar panels and remote sensing or other optical instruments. In such cases, the impulse engines are typically pulsed for shorter time intervals and at much smaller thrust levels, to provide the required impulses for control purposes. Accordingly, a much lower propellant mass flow rate to the impulse engine combustion chambers is typically required, and propellant flow may be discontinuous e.g., pulsed. A variety of propellant pumps can be used in such impulse engine applications including electrical or fueled reciprocating pumps.
Impeller-based pumps including those used for launch vehicle applications have generally been turbine driven. These turbine pumps force propellant at a high mass flow rate into the combustion chamber that generally has a pressure much higher than that of the propellant tank. For certain launch vehicle applications, these turbo pumps may be operated to provide 3,000 horsepower or greater at rotational speeds of 10,000 to 30,000 rpm. The turbine section of these turbo pumps is typically forced to rotate by the ingestion of hot gases from a gas generator system. This gas generator system is normally another small rocket engine, also consuming propellants, to produce the hot exhaust gas that is then ducted into the input of the turbine section of the turbo pump. Such turbo pumps and associated gas generators are costly, complex, massive and require significant monitoring and testing. In particular, the propellants and pressurants needed to power the gas generator contribute to expense, system complexity and overall mass. Moreover, the turbo pump system entails a significant amount of hot gas and high pressure plumbing in the engine feed line area with attendant welds and potential leak paths. Accordingly, such systems require careful construction and monitoring. In addition, complete testing of such systems on the launch pad can only be conducted with propellants onboard. Some such systems, including certain systems that employ a small rocket engine in conjunction with the gas generator, provide limited throttling ability for controlling the rate of propellant delivery. Despite these limitations, such turbo pump systems have been thought necessary for many impeller-based pump systems including those used for launch vehicle applications.
SUMMARY OF THE INVENTION
The present invention relates to an electrical drive system for continuously driving a liquid propellant pump during a thrusting time period. The electrical drive system is thus suitable for rotationally driving an impeller-based pump system and for use in delivering liquid propellant to a launch vehicle thruster during a time period. The invention reduces or substantially eliminates the need for turbine drive systems and associated gas generators, thereby simplifying drive system design, construction and testing and allowing for reduction in overall system cost and mass, while increasing system reliability. The electrical drive system also allows for convenience propellant mass flow rate throttling and various other operational advantages. The electrical drive system may also be employed in other high power, limited duration impeller applications such as torpedo engines.
According to one aspect of the present invention, an electrical drive system is provided for rotationally driving an impeller-based fluid drive unit such as a liquid propellant pump. In the case of a liquid propellant pump of a rocket engine, the pump is operatively interposed between a liquid propellant tank and a combustion chamber of the rocket engine to pump the liquid propellant from the tank to the combustion chamber, and includes a rotating impeller for pumping the propellant. The drive system includes an electric motor with an output shaft for driving the pump, an electrical power source for powering the motor, a controller for regulating current distribution, and linkage for coupling the output shaft of the motor to an impeller shaft of the pump. In the simplest case, the motor and impeller could share a common shaft that serves as the linkage. The electric motor is thereby utilized in conjunction with the pump to pump liquid propellant from the propellant tank to the combustion chamber.
Preferably, the electric motor is a brushless DC motor capable of providing a variable rotation rate to the output shaft. In addition, the electric motor preferably provides a substantially constant torque per amp output over a range of rotation rates of the output shaft for enhanced propellant mass flow rate control. In one embodiment, a number of electric motors are used cooperatively to drive the pump. For example, counter rotating motors may be utilized to reduce or eliminate precessional forces. That is, two or more motors may be used in counter rotating pairs, i.e., rotating in opposite rotational senses.
The electrical power source may include one or more of a high energy density battery, a super capacitor, a fly wheel and a gas generator. Depending on the specific application, in order to provide sufficient power for launch vehicle or other high power, limited duration thrusting applications, the electrical power source may provide at least about 1 megawatt of power during a thrusting time period. Additionally, the electric power source is preferably capable of providing a high power output continuously for a period of at least about 60 seconds. Alternate applications such as upper stages would require less power over shorter periods of time. Moreover, the power source preferably has an energy density of at least about 50 watt hours per kilogram so as to provide adequate power and reduced mass.
The coupling means may include a clutch and/or a mechanism for allowing a rotational and alignment difference as between the motor output shaft and the impeller shaft. In this regard, a clutch may be utilized to selectively disengage the impeller shaft from the motor output shaft as may be desired for testing and other purposes. A gear box, a belt and pulley system, or other suitable mechanism may be provided for stepping up or stepping down the rotation rate be
Abel Terry M.
Velez Thomas A.
Kim Ted
Lockheed Martin Corporation
Marsh & Fischmann & Breyfogle LLP
LandOfFree
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