The present disclosure relates generally to thruster devices for controlling the attitude of a flying body. More specifically, various embodiments of the present disclosure relate to vector thrust devices and methods of making vector thrust devices for use in thrust vector control systems of flying bodies.
Various self-propelled flying bodies, such as rockets and missiles, are typically employed for a variety of uses, such as military and scientific. One of the basic goals of the technology of flight bodies is to improve the maneuverability of the body. The maneuverability of a flight body is related to its ability to change its flight path. Since lateral forces may cause a flight body to change its flight path, the maneuverability of a flight body is related to its ability to develop lateral forces. Various approaches are conventionally applied to develop lateral forces for controlling the attitude and direction of the flight body (e.g., controlling the pitch, yaw, and roll of the flight body).
One conventional means for controlling the attitude and direction of flight bodies includes the use of thrust motors positioned to generate a transverse thrust which provide lateral forces on the flight body. A lateral thrust motor is typically employed in combination with one or more other lateral thrust motors to form a lateral thrust module, which may also be characterized as a divert propulsion system. Generally, a lateral thrust motor is mounted on the flight body to generate thrust in a transverse direction during deployment. The thrust is conventionally generated by injecting high pressure gas, or by combusting a propellant, such as a solid propellant.
The solid propellant used in conventional lateral thrust motors is typically formed into a single unit that is referred to as a grain. The conventional single grain of solid propellant is typically formed large enough to at least substantially fill a chamber in the lateral thrust motor, resulting in a substantially thick grain. With conventional lateral thrust motors, it is typically desired to combust all the propellant material of the grain within a specified period of time in order to achieve a desired net force from the thrust motor.
In manufacturing such propellant grains having the required short action time, it becomes a trade-off between two options, faster-burning propellants and slower-burning propellants. When using a relatively faster-burning propellant (e.g., high burn-rate propellant), the creation of the grain becomes easier, as the burn web (minimum distance between two surfaces of the grain) can be large. However, when creating the chemistry for such high burn-rate propellant material, it is more difficult to keep the burn rate consistent from part to part. For example, fast-burning solid propellant materials typically employ very fine powders that are costly to produce, and that can vary substantially in particle size between manufactured lots when mass produced, resulting in inconsistent burn rates, variable thrust values, and less predictability from part to part.
On the other hand, when using a relatively slower-burning propellant (e.g., low burn-rate propellant), it is easier to produce grains having more consistent burn-rates from part to part, but it becomes more difficult to create grain geometry that has a thin enough burn web for the short action time needed.
In accordance with one or more aspects of the present disclosure, thruster devices and/or lateral thrust modules are provided for use in a flight body, such as a rocket or missile, which thruster devices are adapted to facilitate use of propellant materials having relatively low burning rates, while still achieving relatively short action times. Such thruster devices may reduce the costs and danger involved in manufacturing and handling the propellant materials, and may improve the consistency of thrust values between thruster devices and between manufactured lots of thruster devices, resulting in greater predictability in thrust forces that will result when initiated.
Various embodiments of the present disclosure comprise thruster devices employable in a flight body for generating a transverse thrust. In one or more embodiments, a thruster device may comprise a combustion chamber with a plurality of propellant grains disposed therein. At least some of the plurality of propellant grains are formed into at least one selected shape. An igniter is located in relation to the plurality of propellant grains to initiate combustion of the plurality of propellant grains when the thruster device is deployed.
Other embodiments of the present disclosure include lateral thrust modules employable in a flight body for adjusting attitude and direction of the flight body. In one or more embodiments, a lateral thrust module may comprise a plurality of thruster devices. Each thruster device is oriented to direct a thrust in one of a plurality of different directions. Each thruster device may include a housing defining a combustion chamber and including an injection nozzle at a first longitudinal end thereof. The injection nozzle can be adapted to be joined to an aperture of a flight body. A quantity of propellant material can be disposed within the combustion chamber, where the quantity of propellant material comprises a plurality of propellant grains that are each formed with a selected shape. An igniter can be coupled to the housing at a second longitudinal end thereof. The igniter can be adapted to initiate a combustion of the quantity of propellant material when the thruster device is deployed.
Additional embodiments of the present disclosure include methods for making a thruster device that is capable of being employed in a flight body. One or more implementations of such methods may include forming a housing that comprises a first longitudinal end and an opposing second longitudinal end, where the housing defines a combustion chamber and includes an injection nozzle at the first longitudinal end. A plurality of propellant grains can be disposed in the combustion chamber of the housing. Each propellant grain comprises a selected shape. An igniter may be coupled to the second longitudinal end of the housing. The igniter can be adapted to initiate a combustion of the plurality of propellant grains during deployment of the thruster device.
Exemplary embodiments of the disclosure will become more fully apparent from the following description and appended claims, taken in conjunction with the accompanying drawings. Understanding that these drawings depict only exemplary embodiments and are, therefore, not to be considered limiting of the disclosure's scope, the exemplary embodiments of the disclosure will be described with additional specificity and detail through use of the accompanying drawings in which:
The illustrations presented herein are, in some instances, not actual views of any particular thruster devices, lateral thrust modules or flight bodies, but are merely idealized representations which are employed to describe the present disclosure. Additionally, elements common between figures may retain the same numerical reference designation.
Various embodiments of the present disclosure include thruster devices and thrust modules for use in various flight bodies.
Within the combustion chamber 304, a conduit 312 may be disposed. As shown herein, the conduit 312 comprises a cylindrical tube having a plurality of holes 314 formed in the sidewall of the conduit 312. The illustrated conduit 312 includes two portions shown separated by a wall, an igniter portion 316 and a propellant portion 318.
An igniter 320 is coupled to the housing 302 at the top longitudinal end 308. The igniter 320 may generally include a squib 322 coupled to one or more wires 324 for creating an initial reaction upon receipt of a current and/or electrical charge via the one or more wires 324. In addition, the igniter 320 may include a quantity of combustible material (not shown) capable of being combusted upon deployment of the squib 322. Upon ignition of the squib 322 and/or the combustible material of the igniter 320, the hot gases generated may flow through the igniter portion 316 of the conduit 312 and out through the holes 314 to ignite a quantity of solid propellant material 326 disposed within the combustion chamber 304 and generally positioned around an outer surface of the conduit 312. Hot gases generated by combustion of the propellant material 326 enter into the propellant portion 318 of the conduit 312 through holes 314, increasing the internal pressure and causing the burst disk 310 to rupture. After the burst disk 310 ruptures, the thrust gases exit through the injection nozzle 204.
The propellant material 326 may comprise any conventional propellant material comprising a relatively normal or even slow burn rate. By way of example and not limitation, the propellant material may be selected to comprise a burn rate between about 0.5 in/sec. and 2 in/sec. (about 12.7 mm/sec. and 50.8 mm/sec.). In at least some implementations, the propellant material 326 may comprise a composite propellant material. In general, composite propellants typically comprise a metallic fuel, such as aluminum and/or magnesium, mixed with an oxidizer and immobilized with a rubbery binder such as synthetic rubber. Composite propellants may comprise an ammonium nitrate-based composite propellant (ANCP) or an ammonium perchlorate-based composite propellant (APCP). In at least some implementations, other propellant materials 326 may be employed such as, by way of example and not limitation, variations of boron potassium nitrate (BKNO3) or basic copper nitrate (BCN), and/or guanidine nitrate (GuNO3)-based gas generating materials, as well as any other propellant formulations including fuels and oxidizers.
The propellant material 326 is typically employed in forms called grains. A grain generally comprises an individual unit of propellant, no matter the size. Conventionally, the propellant grain is formed by casting the propellant material into a single grain that is sized and shaped to fill substantially all of the combustion chamber of a solid propellant motor. Cast grains, however, can vary significantly from part to part and cannot be easily or accurately adjusted prior to loading.
According to at least one feature of the present disclosure, the thruster 202 employs a quantity of solid propellant material 326 that comprises a plurality of discrete grains, which are formed into one or more selected shapes. In at least some implementations, the discrete grains may be formed by subjecting the propellant material 326 to high pressure to press the propellant material into the selected shape for each grain. A binder, such as an organic or non-organic binder, can be employed when pressing the propellant material into the selected shape. By way of example only, the binder may comprise a rubber binder such as Hydroxyl-terminated polybutadiene, or the binder may comprise a guanidine nitrate or similar material given the forces encountered during pressing operations. In at least some other implementations, the discrete grains may be formed by extruding the propellant material 326 to form the discrete grains with the desired shape.
The discrete grains comprising the quantity of solid propellant material 326 can be shaped and sized according to a plurality of different embodiments. For example, the various discrete grains employed as some or all of the quantity of solid propellant material 326 may comprise one or more configurations of pellet-shaped grains. Such pellet-shaped grains 402 can have any of a plurality of general shapes. By way of example and not limitation, the pellet-shaped grains can be generally spherical, elliptical, ovoid, cylindrical, toroidal and/or tablet-shaped.
Various embodiments of the wafers 502 may comprise one or more surface irregularities, such as grooves, indentations, slots, channels, and the like of different shapes formed in one or more surfaces of the wafer 502 so that the generally flat sides of each wafer 502 do not abut or seat in abuting relationship against any adjacent wafer 502. For example, one or more irregularities may be formed in a top and/or bottom surface (as oriented in
By using a plurality of discrete grains, such as any of the examples of pellet-shaped grains just described, the ignitable surface area for the quantity of solid propellant material 326 within the combustion chamber 304 is substantially increased, while maintaining a relatively smaller burning thickness or web. As a result, a thruster design can be achieved which exhibits reasonable combustion pressure (e.g, between about 2,000 psi (about 13.79 MPa) and about 10,000 psi (about 68.95 MPa)) while using a relatively lower burning rate propellant material 326 that is still capable of exhibiting a relatively short action time. Such slower burning propellant materials are generally more stable and more predictable than the conventional high burning-rate materials used in conventional thrusters exhibiting short action times. In addition, propellant materials exhibiting relatively lower burning rates are typically easier and cheaper to manufacture, without substantial variations between manufacturing lots. Employing a slower burning propellant material can result in more repeatable (i.e., less variable) thrust forces produced by thrusters 202 of the present disclosure.
The various embodiments of discrete grains of solid propellant material 326 provided above are merely some examples of suitable sizes and/or shapes for discrete grains employed in one or more thrusters 202 of the present disclosure. Other sizes and/or shapes of discrete grains may also be employed in one or more thrusters 202 according to other various embodiments of the present disclosure. Furthermore, various implementations of the current thrusters 202 may employ a combination of more than one size and/or shape of discrete grains for the solid propellant material 326. That is, in some embodiments, a thruster 202 may employ a combination of two or more different embodiments of discrete grains of solid propellant material 326, for example pellet-shaped grains formed as both tablets and hollow cylinders.
Additional embodiments of the present disclosure relate to methods of forming thrusters, such as thrusters 202.
At step 704, a plurality of propellant grains are formed to comprise at least one selected shape. For example, the plurality of propellant grains can comprise pellet-shaped grains (see, e.g.,
At step 706, the plurality of propellant grains are disposed in the combustion chamber 304 of the housing 302. For example, the plurality of propellant grains can be randomly packed into the combustion chamber 304 in some implementations (e.g., tablet-shaped grains, hollow cylinder grains), while the propellant grains also can be selectively positioned in the combustion chamber 304 in other implementations (e.g., wafer-shaped grains).
At step 708, an igniter 320 can be coupled to the second longitudinal end 308 of the housing 302. The igniter 320 can be adapted to initiate a combustion of the plurality or propellant grains during deployment of the thruster device.
The various embodiments and implementations of the present disclosure result in thruster devices and lateral thrust modules with relatively low performance variability. In particular, the use of the discrete propellant grains, as described herein, provides for a more consistent thrust value over conventional devices, at least in part as a result of the ability to more accurately load the propellant grains by weight from thruster device to thruster device.
The present invention may be embodied in other specific forms without departing from its structures, methods, or other essential characteristics as broadly described herein and claimed hereinafter. The described embodiments are to be considered in all respects only as illustrative, and not restrictive. The scope of the invention is, therefore, indicated by the appended claims, rather than by the foregoing description. All changes that come within the meaning and range of equivalency of the claims are to be embraced within their scope.