Aeronautics and astronautics – Spacecraft – Spacecraft formation – orbit – or interplanetary path
Reexamination Certificate
2001-03-16
2002-12-31
Jordan, Charles T. (Department: 3644)
Aeronautics and astronautics
Spacecraft
Spacecraft formation, orbit, or interplanetary path
C244S172200
Reexamination Certificate
active
06499696
ABSTRACT:
FIELD OF THE INVENTION
The present invention generally relates to high pressure systems in rockets (e.g., rocket engines, oxidizer systems, propellant tanks) and, more particularly, to a vent system for such high pressure systems.
BACKGROUND OF THE INVENTION
Multi-stage rockets have long been used to propel various types a spacecraft into outer space, including satellites. One known multi-stage rocket configuration has an upper stage whose forward end is attached to the spacecraft and whose aft end is interconnected with a lower stage by an interstage adapter. Prior to separation of the lower stage from the upper stage during flight and also prior to ignition of the upper stage engine, a flow of rocket fuel into the upper stage engine is initiated. This flow of rocket fuel is directed out of the upper stage engine through an upper stage rocket engine vent system. The upper stage rocket engine vent system directs this rocket fuel overboard via an appropriate conduit that passes through an exterior wall of the rocket (e.g., on the interstage adapter).
The lower stage is separated from the upper stage rocket engine at the appropriate time during flight, and the upper stage rocket engine is also ignited at the appropriate time. Further travel will then be affected by the thrust provided by the upper stage rocket engine. One prior art protocol that ultimately leads to the separation of the upper stage from the spacecraft entails shutting down the upper stage rocket engine a number of times. That is, the upper stage rocket engine is shut down for a predetermined period of time, is thereafter ignited and run for a predetermined period of time, and is thereafter shut down once again. Fuel is discharged from the upper stage rocket engine through the upper stage rocket engine vent system under very high pressure each time that the upper stage rocket engine is shut down. One prior art configuration for this upper stage rocket engine fuel vent system uses a pair of about 2 inch diameter cylindrical ducts that are interconnected with the upper stage rocket engine. Both of these ducts extend out from the rocket engine at least generally perpendicular to the primary or longitudinal axis of the upper stage, and thereafter are directed rearwardly (e.g., in a generally L-shaped configuration). The thrust that is generated out of each these cylindrical vent ducts is on the order of about 600 pounds in one prior art embodiment. Thrusts of this magnitude cause relatively significant movement of the ducts relative to the upper stage rocket engine and introduces significant structural stresses at the interconnection with the upper stage rocket engine. There is a strong potential for damage to the upper stage rocket engine because of the stresses.
BRIEF SUMMARY OF THE INVENTION
The present invention generally relates to a system for reducing axial thrust for typically high pressure discharges. One particularly desirable application of the present invention is in a rocket engine vent system.
A first aspect of the present invention is embodied by a rocket that includes a spacecraft, as well a first or an upper stage that is interconnected with this spacecraft and that includes at least one rocket engine. A first or an upper stage rocket engine vent system is fluidly interconnected with this upper stage rocket engine (or any other rocket engine(s) associated with any other stage used by the rocket) and includes at least one first or upper stage rocket engine vent system conduit (hereafter a “first rocket engine vent system conduit”). A pair of vent apertures extend entirely through this first rocket engine vent system conduit. A first flow diverter that includes first and second vanes is disposed within this first rocket engine vent system conduit such that each vane projects at least generally toward its own vent aperture. Therefore, at least two separate and discrete fluid flows may be directed out of the first rocket engine vent system conduit.
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. The static pressure profile throughout the first rocket engine vent system conduit may be at least substantially unchanged (in relation to such a conduit that does not include the first flow diverter and vent apertures) by having the collective area of the two vent apertures (i.e., the sum of the surface area occupied by each of these vent apertures) be greater than the cross-sectional area of the first rocket engine vent system conduit (i.e., the area of the first rocket engine vent system conduit, taken perpendicular to a center reference axis about which the first rocket engine vent system conduit may be disposed). If a significant back pressure did develop in the first rocket engine vent system conduit, that is if the venting did not occur quickly enough such that this conduit (as well as the upper stage rocket engine) became overpressurized or such that the static pressure therein significantly increased, the upper stage rocket engine could be damaged. The pair of vent apertures are also preferably disposed in at least substantially opposing relation (e.g., 180 degrees apart), or stated another way on opposite sides of the first rocket vent system conduit. As such, two fluid flows may be directed out of the same first rocket engine vent system conduit in at least substantially opposite directions to at least substantially cancel any lateral thrust vectors that may be exerted on the first rocket engine vent system conduit. This also allows the fluid flows to be directed from the first rocket engine vent system conduit away from sensitive components or so as to otherwise reduce the potential for damaging other components of the rocket.
Axial thrusts to which the upper stage rocket engine is exposed during venting may be significantly reduced. In one embodiment, the end of the first rocket engine vent system conduit that is opposite that which interfaces with the upper stage rocket engine is closed. Stated another way, axial thrusts may be significantly reduced by having the projected area of the first and second vanes onto a plane perpendicular to a center reference axis about which the first rocket engine vent system conduit may be disposed, equal to the area of the first rocket engine vent system conduit defined by its diameter.
Multiple components may collectively define the first rocket engine vent system conduit associated with the subject first aspect. For instance, this first rocket engine vent system conduit may include a first conduit section having a first end that interfaces with the upper stage rocket engine. A separate first flow adapter may be separately interconnected with this first conduit section. The noted pair of vent apertures and the first flow diverter may be part of this first flow adapter. In one embodiment, the first flow adapter is attached to and extends beyond a second end of the first conduit section that is opposite the above-noted first end that interfaces with the upper stage rocket engine.
The first rocket engine vent system conduit may be disposed about a center reference axis and may include a pair of beveled surfaces. Each beveled surface may include one of the noted vent apertures. These beveled surfaces may be oriented on the first rocket engine vent system conduit so as to extend at least generally toward the noted center reference when progressing in a direction of a downstream end of the first rocket engine vent system conduit. In one embodiment, the pair of beveled surfaces are disposed in opposing relation, or stated another way, the beveled surfaces are spaced about 180 degrees apart. The beveled surfaces allow for increasing the surface area of the vent apertures for purposes of allowing adequate flow out of the first rocket engine vent system conduit.
The pair of vanes of the first flow diverter may direct two
Dinh T.
Jordan Charles T.
Lockheed Martin Corporation
Marsh & Fischmann & Breyfogle LLP
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