Cryogenic propellant depletion system for a launch vehicle

Details Cryogenic propellant depletion system for a launch vehicle Cryogenic propellant depletion system for a launch vehicle

2002-01-25

2004-02-24

Barefoot, Galen L. (Department: 3644)

Aeronautics and astronautics

Aircraft structure

Fuel supply

C702S055000, C073S29000R

Reexamination Certificate

active

06695256

ABSTRACT:

FIELD OF THE INVENTION
The invention is related to the field of space launch vehicles for use in launching a payload from a stationary ground-based position into orbit, and specifically to a cryogenic propellant depletion system for maximizing the utilization of cryogenic propellant by the booster stage of the space launch vehicle.
BACKGROUND OF THE INVENTION
Launch vehicles are generally used to launch payloads, such as satellites or scientific equipment, from the Earth's surface into space. 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 upper stage. The booster stage is designed to launch and deliver the payload a predetermined distance above the earth before exhaustion. Upon exhaustion, the booster stage and upper stage separate whereupon the upper stage is fired the to deliver the payload the remainder of the distance into a desired orbit. In the case where the booster stage is a reusable component, the booster stage controllably falls back to the earth's surface upon separation for retrieval, refitting, and future use.
The booster stage's rocket engine(s) typically utilize liquid propellants and in the case of bi-propellant rockets generally include two or more propellant tanks, booster pumps, a combustion chamber, plumbing interconnecting the various components, and a nozzle for accelerating and/or discharging the combustion product. Liquid propellant rockets generally utilize a liquid fuel such as RP-1 (i.e., kerosene) and an oxidizer such as liquid oxygen (LOX), which are stored in separate propellant tanks and brought into contact in the combustion chamber to provide thrust.
In order to preserve the booster stage rocket engine for re-use, the engine must be properly shut down at or near the end of the launch boost stage. That is, the engine must be shut down prior to violating any engine requirements that may result in some sort of permanent engine or vehicle damage. For, example, reusable rocket engines generally require that the propellant(s) be supplied with a minimum “head” pressure in order for the engine to properly function. As the propellant in one or both of the propellant tanks nears exhaustion, the head pressure generally drops. This head pressure drop may potentially result in engine and/or booster pump damage if the head pressure drops below a minimum allowable threshold (i.e., engine requirement). Further, proper engine shut down generally requires a predetermined mass of one or both of the propellants to prevent engine and/or booster pump damage which my result in catastrophic failure. Therefore, it is important to initiate engine shut down prior to the propellant(s) dropping below any minimum allowable threshold, in both reusable and expendable booster rockets. However, initiating engine shut down to prevent engine damage and/or rocket failure must be balanced with the desire to fully utilize the available propellant to maximize the launch vehicle's booster stage performance. As will be appreciated, it is desirable to utilize the propellant right up until the last possible moment prior to exceeding a minimum allowable threshold in order to maximize boost. Accordingly, it is desirable to continuously monitor the amount of the propellant remaining such that engine shut down may be initiated just prior to the propellant dropping below any minimum allowable threshold.
Existing propellant monitoring systems generally utilize hot point sensors which indicate a transition between liquid to gas in the propellant supply system through, for example, a change in a monitored capacitance. In these systems booster engine cut off (BECO) is initiated a predetermined time after the propellant level drops below one of the hot point sensors. Unfortunately, it is often difficult to precisely determine the amount of the remaining propellant using hot point sensors when a two-phase mixture of the propellant and/or an ullage gas exist in the system. This is especially true with cryogenic propellants which are susceptible to “boiling off” (i.e., liquid oxygen to gaseous oxygen) and which may introduce a two-phase mixture into the cryogenic storage tank and the plumbing interconnecting the storage tank to the rest of the system. Two-phase mixtures of the cryogenic propellants make determination of the remaining propellant parameters difficult as hot point sensors may prematurely or belatedly indicate the remaining level/pressure of propellant or otherwise provide erratic signals. As a result of inaccurate propellant parameter measurements the booster stage engine may prematurely shut down and fail to utilize all available propellant, thus, reducing booster stage performance. Alternatively, the propellant may be depleted beyond a minimum allowable threshold required for proper engine shutdown prior to initiating engine shut down which may result in violation of the engine requirements and/or engine damage.
SUMMARY OF THE INVENTION
In view of the foregoing, a primary object of the present invention is to provide a system for monitoring a cryogenic propellant, irrespective of that propellant being a pure liquid. Another objective of the present invention is to provide a cryogenic depletion monitoring system that is easily adaptable for operation with current launch vehicles.
One or more of the above-noted objectives, as well as additional advantages, are provided by the present invention, which includes a cryogenic depletion monitoring system for use in monitoring a cryogenic propellant in a feed line between a cryogenic storage tank and a booster engine in a space launch vehicle. The cryogenic depletion monitoring system utilizes a processing system, memory, supported logic, and one or more pressure sensor readings to generate one or more parameters related to the propellant in the feed line. The propellant parameters may then be used to determine when to initiate booster engine cut off such that booster stage performance is maximized while no engine requirements are violated.
According to a first aspect of the present invention, a cryogenic propellant depletion monitoring system is provided that includes a processing system containing logic to determine and monitor at least a first propellant parameter associated with the cryogenic propellant in the launch vehicle's cryogenic feed line. Additionally, the system contains control means to initiate engine shut off once at least one of the determined propellant parameters falls below a minimum allowable threshold. The system is operable to determine these propellant parameters and initiate booster engine shutdown irrespective of the cryogenic propellant being a pure liquid.
Various refinements exist of the features noted in relation to the subject first aspect of the present invention. Further features may also be incorporated into 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 processing system and the supported logic may utilize any available information for determining the propellant parameters associated with the propellant in the feed line. This information may include, inter alia, algorithms, material constants and/or prior test data that may be stored in an accessible memory structure.
In a preferred embodiment of the present invention, the processing system utilizes at least two pressure measurements associated with the cryogenic system to determine the one or more parameters related to the propellant in the feed line. These pressure measurements may include a first pressure measurement of the propellant from a point along the length of the feed line and a second pressure measurement from the cryogenic storage tank. In this regard, the processing system may utilize the pressure measurements to determine a pressure differential betwee

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