Variable burn-rate propellant

Details Variable burn-rate propellant Variable burn-rate propellant

1999-11-23

2003-01-07

Carone, Michael J. (Department: 3641)

Explosive and thermic compositions or charges

Metal or alloy or metalloid, each in particulate form, with...

C149S076000, C149S109600

Reexamination Certificate

active

06503350

ABSTRACT:

FIELD OF THE INVENTION
This invention relates to propellants such as may be used in solid rocket motors. In preferred embodiments, the propellant comprises one high energy propellant composition comprising a homogeneous mixture of fuel and oxidizer present in a predetermined ratio, wherein individual fuel particles are generally uniformly distributed throughout a matrix of solid oxidizer, and a low energy propellant composition comprising a fuel and oxidizer. The amounts of the two propellants are present in amounts which achieve a preselected burn rate.
BACKGROUND OF THE INVENTION
Solid rocket motor propellants are widely used in a variety of aerospace applications, such as launch vehicles for satellites and spacecraft. Solid propellants have many advantages over liquid propellants for these applications because of their good performance characteristics, ease of formulation, ease and safety of use, and the simplicity of design of the solid fueled rocket motor when compared to the liquid fueled rocket motor.
The conventional solid propellant typically consists of an organic or inorganic solid oxidizing agent, a solid metallic fuel, a liquid polymeric binder, and a curing agent for the binder. Additional components for improving the properties of the propellant, i.e., processability, curability, mechanical strength, stability, and burning characteristics, may also be present. These additives may include bonding agents, plasticizers, cure catalysts, burn rate catalysts, and other similar materials. The solid propellant is typically prepared by mechanical mixing of the oxidizer and metallic fuel particles, followed by addition of the binder and curing agent with additional mixing. The resulting mixture is then poured or vacuum cast into the motor casing and cured to a solid mass.
The solid propellant formulations most widely used today in such applications as the Space Shuttle solid rocket booster and Delta rockets contain as key ingredients aluminum (Al) particles as the metal fuel and ammonium perchlorate (AP) particles as the oxidizer. The Al and AP particles are held together by a binder, which is also a fuel, albeit one of substantially less energetic content than the metal. The most commonly used binder comprises hydroxy-terminated polybutadiene (HTPB). This particular type of propellant formulation is favored for its ease of manufacture and handling, good performance characteristics, reliability and cost-effectiveness.
A typical Al+AP solid rocket propellant formulation consists of 68 wt. % AP (trimodal particle size distribution, i.e., 24 wt. % 200 &mgr;m, 17 wt. % 20 &mgr;m, 27 wt. % 3 &mgr;m), 19 wt. % Al (30 &mgr;m average particle diameter), 12 wt. % binder (HTPB) and isophorone diisocyanate (IPDI) curing agent), and 1 wt. % burn rate catalyst (e.g., Fe2O3 powder).
The relative amounts of the components in this formulation are chemically stoichiometric. In other words, there should be just enough oxidizer molecules present in the formulation to completely react with all the fuel molecules that are present, with no excess of either oxidizer or fuel. This formulation contains one oxidizer (AP) and two distinct fuels, i.e., Al and binder. The weight ratio of AP to Al for a stoichiometric mixture, i.e., no excess oxidizer or fuel, is 42:19. The weight ratio of ammonium perchlorate to binder for a stoichiometric mixture is 26:12. These ratios are the same regardless of any other components that may be present in the mixture.
Because of their burn characteristics, conventional Al/AP propellants are most suitable for use in conjunction with a particular motor design. This design is the hollow core or center perforated (CP) core motor design in which the propellant grain is formed with its outer surface bonded to the inside of the rocket motor's casing with a hollow core extending through most or all of the length of the grain. The burning front progresses radially outwardly from the core to the case. This motor design is by far the most common design for solid fuel motors. One example of a current application utilizing this design is the Space Shuttle, which uses solid motors which are 150 ft. long and 12 ft. in diameter with a 4 ft. hollow core.
The propellant grain in a CP design must have substantial structural integrity to keep the grain intact during operation. A binder is therefore used to “glue” the particulate components of the propellant together. During the initial mixing of the propellant, the percentage of the binder, initially in the form of a liquid resin, is high enough to maintain a relatively low viscosity, such that the propellant is in a slurry form, allowing the propellant mixture to be poured or injected into the motor casing. A mandrel is placed in the middle of the motor casing to create the hollow core (typically before the propellant is poured into the core) and is removed once the propellant has cured.
Propellants comprising a metal fuel in combination with a solid oxidizer may be used in other applications outside of aerospace, including gas generators. Solid propellants are also used in launch vehicles, e.g., NASA rockets, Space Shuttle, French Ariane rockets. Virtually all launch vehicles use a combination of liquid fuel motors with solid fuel boosters. Both the Delta III and the Space Shuttle are examples having combined liquid and solid motors. The Delta rocket has a main liquid motor with nine smaller strap-on solid boosters, while the shuttle has three onboard liquid motors with two strap-on solid boosters.
Although enormous innovations have occurred in guidance, electronics and virtually every part of spacecraft to date, the propulsion systems have remained essentially the same for decades. Boeing's Delta III, introduced in 1998, utilizes a liquid engine that was designed in the 1960's and is fueled by kerosene and oxidized by liquid oxygen. The solid boosters were designed in 1961 and are virtually unchanged since then, except for an epoxy motor casing. Additionally, over the past decade, almost every system on the Shuttle has been replaced or upgraded, except for its propellant. It is therefore desirable to provide a novel solid rocket propellant that affords superior performance to the conventional propellants in current use today.
SUMMARY OF THE INVENTION
A propellant is a composition of matter comprising at least one fuel and at least one oxidizer. The reduction/oxidation (redox) reaction between the fuel and oxidizer provides energy, frequently in the form of evolved gas, which is useful in providing an impulse to move a projectile such as a rocket or spacecraft. The present invention provides propellant compositions capable of achieving very high burn rates. The propellant compositions of the present invention may comprise a single fuel and oxidizer. In some embodiments, the propellants are mixed propellants. A mixed propellant is a mixture of at least two propellants. The two component propellants may have the same fuel and/or oxidizer, but there should be some difference, such as a different fuel particle size, additional or different catalyst, etc.
The present invention also provides methods of reducing the burn rates of the high burn rate propellants by varying their composition. Such methods include addition of lower burn rate materials and/or propellants, and altering the particle size of one or more components of a propellant as disclosed below. In preferred embodiments, the propellants disclosed are of the type which may be used in solid rocket motors such as are found in launch vehicles. Other embodiments may be used in other applications for propellants as may be known in the art.
In accordance with one aspect of the present invention there is provided a mixed solid propellant. The propellant comprises a first propellant composition comprising a substantially homogeneous mixture of fuel particles distributed throughout a matrix of a first oxidizer, and a second propellant composition comprising a fuel and a second oxidizer. In preferred embodiments, the second propellant is present in a quantity sufficie

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