Patent Description:
Various applications require payloads that are launched from a mission suitable platform, such as a land, sea, air, or space vehicle. The payload to be launched is dependent on the application. Military applications that use land vehicles, aircrafts, surface ships, or underwater vehicles may use deployable munitions as payloads. The payloads may be carried by a flight vehicle including a rocket motor having multiple pulses which enables non-continuous propulsion, different propulsive forces, and backup pulses without requiring completely separate rocket motor stages.

The propellant grains for each pulse may be separated by an inert barrier that prevents the subsequent pulses from igniting due to contact with combustion gases formed from another burning propellant grain. When the propellant grain is ignited, the force of the generated combustion gases breaks through the barrier to expel thrust gases from the projectile until the propellant of the pulse burns out. At least one igniter is used to ignite each propellant pulse.

Conventional multi-pulse rocket motors can be deficient in that the rocket motors may have complicated barrier and igniter systems between the pulses that are difficult and costly to manufacture. One possible deficiency of conventional barriers is that the barrier may be ejected or inverted which may lead to clogging of the nozzle in the rocket motor or combustion instability. Still another disadvantage of conventional multi-pulse rocket motors is that they may be limited to two pulses.

<CIT> discloses a gas generation system for generating gases, such as for use as or as part of a rocket motor in propelling a projectile, including two or more propellant charges and electrically operated propellant initiators operatively coupled to respective of the propellant charges, to initiate combustion in the propellant charges, wherein the propellant charges are operatively isolated from one another such that the propellant charges can be individually initiated and are not ignited due to gases generated from other of the propellant charges being combusted.

<CIT> discloses that electrically operated propellant is used to supplement the thrust provided by solid rocket motor (SRM) propellant to manage thrust produced by a SRM. The gas produced by burning the electrically operated propellant may be injected upstream of the nozzle to add mass and increase chamber pressure Pc, injected at the throat of the nozzle to reduce the effect throat area At to increase chamber pressure Pc or injected downstream of the throat to provide thrust vector control or a combination thereof. Certain types of electrically operated propellants can be turned on and off provided the chamber pressure Pc does not exceed a self-sustaining threshold pressure eliminating the requirement for physical control valves.

The present application provides a rocket motor having an electrically operated propellant initiator for a propellant grain that includes an electrode arrangement configured to concentrate an electric field at an ignition electrode for igniting an electrically operated propellant.

According to an aspect of the invention, a rocket motor comprising: a pulse chamber containing at least one propellant grain; and an electrically operated propellant initiator operatively coupled to the at least one propellant grain to initiate combustion of the at least one propellant grain, the electrically operated propellant initiator including an electrically operated propellant and at least one pair of electrodes arranged to ignite the electrically operated propellant, the at least one pair of electrodes including a ground plane electrode and an ignition electrode at which an electric field is concentrated to ignite the electrically operated propellant; wherein the ground plane electrode extends along a first surface area of the at least one propellant grain that is larger than a second surface area of the at least one propellant grain along which the ignition electrode extends, wherein a ratio of the first surface area to the second surface area is at least <NUM>:<NUM>.

According to an embodiment of any paragraph(s) of this summary, the ignition electrode has a greater current density than the ground plane electrode.

According to an embodiment of any paragraph(s) of this summary, the ground plane electrode is formed of wires having a first diameter that is larger than a second diameter of wires forming the ignition electrode.

According to an embodiment of any paragraph(s) of this summary, the at least one pair of electrodes are formed of a plurality of wires having a zig-zag, diagonal, crisscross, rectangular, parallel, non-parallel, or serpentine arrangement.

According to an embodiment of any paragraph(s) of this summary, the at least one pair of electrodes are separately arranged without crossing each other.

According to an embodiment of any paragraph(s) of this summary, the at least one propellant grain has at least one cavity in which at least one of the ignition electrode and the ground plane electrode is arranged.

According to an embodiment of any paragraph(s) of this summary, the ignition electrode and the ground plane electrode are arranged in a same plane that is parallel with an outer surface of the at least one propellant grain.

According to an embodiment of any paragraph(s) of this summary, the ignition electrode and the ground plane electrode are arranged in different planes.

According to an embodiment of any paragraph(s) of this summary, the ignition electrode is formed of a refractory metal or refractory alloy.

According to an embodiment of any paragraph(s) of this summary, the rocket motor is a multi-pulse rocket motor and the at least one propellant grain includes a first propellant grain and a second propellant grain that is burned during a second pulse of the multi-pulse rocket motor, the second propellant grain being operatively isolated from the first propellant grain via the electrically operated propellant initiator whereby the first propellant grain and the second propellant grain are individually initiated.

According to an embodiment of any paragraph(s) of this summary, the at least one propellant grain includes a third propellant grain that is burned during a third pulse of the multi-pulse rocket motor, the third propellant grain being operatively isolated from the second propellant grain via another electrically operated propellant initiator.

According to an embodiment of any paragraph(s) of this summary, the electrically operated propellant is configured to transition from an unignited state to an ignited state when electrical input is applied across the electrically operated propellant initiator, and is configured to maintain the unignited state when the electrical input is not applied.

According to an embodiment of any paragraph(s) of this summary, the rocket motor includes a power source and a pair of leads extending from the power source to the electrically operated propellant initiator.

According to an embodiment the at least one propellant grain comprises: a first pulse containing a first propellant grain; and a second pulse containing a second propellant grain operatively isolated from the first propellant grain; wherein, the electrically operated propellant initiator comprises at least one electrically operated propellant initiator configured to isolate the first propellant grain and the second propellant grain, the at least one electrically operated propellant initiator being configured to initiate combustion of the first propellant grain or the second propellant grain; and wherein the first surface area and the second surface area are surface areas of the second propellant grain.

According to an embodiment of any paragraph(s) of this summary, the multi-pulse rocket motor includes a third pulse containing a third propellant grain operatively isolated from the first propellant grain and the second propellant grain, wherein the at least one electrically operated propellant initiator includes a first electrically operated propellant initiator configured to isolate the first propellant grain and the second propellant grain and a second electrically operated propellant initiator configured to isolate the second propellant grain and the third propellant grain.

According to an embodiment of any paragraph(s) of this summary, a ratio of the first surface area to the second surface area is at least <NUM>:<NUM>.

According to another aspect of the invention, a method of operating a rocket motor comprising: applying an electrical input across an electrically operated propellant initiator that is operatively coupled to a propellant grain and includes an electrically operated propellant and at least one pair of electrodes, the at least one pair of electrodes including a ground plane electrode and an ignition electrode, wherein the ground plane electrode extends along a first surface area of the at least one propellant grain that is larger than a second surface area of the at least one propellant grain along which the ignition electrode extends, wherein a ratio of the first surface area to the second surface area is at least <NUM>:<NUM>; igniting the electrically operated propellant by an electric field being concentrated at the ignition electrode and initiating combustion of the propellant grain via igniting the electrically operated propellant.

According to an embodiment of any paragraph(s) of this summary, the method includes maintaining operative isolation of the propellant grain from a second propellant grain of a second propulsion pulse for the rocket motor, providing a second propulsion pulse for the rocket motor by applying another electrical input across a second electrically operated propellant initiator subsequent to the electrically operated propellant initiator that is operatively coupled to the propellant grain, the second electrically operated propellant initiator being operatively coupled to the second propellant grain, igniting a second electrically operated propellant of the second electrically operated propellant initiator, and initiating combustion of the second propellant grain via igniting the second electrically operated propellant.

To the accomplishment of the foregoing and related ends, the invention comprises the features hereinafter fully described and particularly pointed out in the claims. The following description and the annexed drawings set forth in detail certain illustrative embodiments of the invention. These embodiments are indicative, however, of but a few of the various ways in which the principles of the invention may be employed. Other objects, advantages and novel features of the invention will become apparent from the following detailed description of the invention when considered in conjunction with the drawings.

The principles described herein have application in defense applications, such as in a hypersonic vehicle or in any flight vehicle where space may be constrained. The rocket motor described herein may be implemented in any suitable flight vehicle or projectile. Single-pulse or multi-pulse rocket motors may be suitable. The rocket motor may be part of a missile that is suitable for carrying a payload. For example, the missile may include a payload module for carrying a munition for a military application. The rocket motor includes a gas generation system described herein for generating combustion gases that propel the munition. In other exemplary embodiments, the gas generation system may be used for other purposes, such as to drive a turbine, to operate a pressure driven mechanical device, to provide tank gas pressurization, etc..

Referring first to <FIG>, a projectile <NUM> includes a rocket motor <NUM> for generating gases to drive movement of the projectile <NUM>. The projectile <NUM> includes a body having a chamber or casing <NUM> for the rocket motor <NUM>. The rocket motor <NUM> includes a nozzle assembly <NUM> having a nozzle opening <NUM> for expelling propellant gases generated from the rocket motor <NUM>. The rocket motor <NUM> is a solid rocket motor including the casing <NUM>, at least one pulse <NUM>, <NUM>, <NUM> containing a propellant charge or grain 24a, 26a, 28a, and at least one initiator <NUM>, <NUM>. The initiator <NUM>, <NUM> is configured to ignite at least one of the propellant grains 24a, 26a, 28a to produce thrust during the corresponding pulse <NUM>, <NUM>, <NUM>.

The rocket motor <NUM> may be a single pulse motor having only one pulse or a multi-pulse rocket motor having two, three, or more pulses <NUM>, <NUM>, <NUM>. In an exemplary embodiment, the pulses <NUM>, <NUM>, <NUM> may be formed in a single chamber or casing <NUM>. Other configurations of the pulses <NUM>, <NUM>, <NUM> and propellant grains 24a, 26a, 28a may be suitable, as required for a particular application. In addition to initiating the propellant grain 24a, 26a, 28a, the initiators <NUM>, <NUM> are also configured to act as a barrier between the propellant grain 24a, 26a, 28a. Accordingly, isolated initiation of the propellant grain 24a, 26a, 28a is enabled to provide pulses that may be fired separately from one another, meaning that an adjacent propellant grain 24a, 26a, 28a cannot be ignited due to gases generated from the combustion of other propellant grain 24a, 26a, 28a.

The casing <NUM> is configured to enable expulsion of combustion gases from the casing interior. The casing <NUM> may have any suitable shape, such as a cylindrical shape that is useful in a projectile application. The casing <NUM> may be formed from any suitable material for containing burning combustion gases at high pressures and high temperatures. For example, the casing may be formed of an inert material that is not ignitable during normal use of the rocket motor <NUM>.

The propellant grains 24a, 26a, 28a are provided for being ignited and combusted to generate high pressure gases for being used to propel or move an object or to pressurize a container, for example. The propellant grain 24a, 26a, 28a may be solid, single pieces having any shape or form, such as core burning, slotted core, rod and tubes, pellets, grains, etc. The propellant grain 24a, 26a, 28a may be made from any suitable material or materials, including fuels, oxidizers, binders, plasticizers, metallic additives, etc. Each of the propellant grain 24a, 26a, 28a may be made of different materials, have different form, or have different shape and/or size than the other propellant grain. End burning grain and center-perforated grain may be suitable. Other fluted grain configurations are also suitable. The propellant grain 24a, 26a, 28a may be isolated from one another. In other exemplary embodiments, additional grain or pulses may or may not be separately isolated.

As shown in <FIG>, the propellant grain 24a, 26a for the first and second pulses <NUM>, <NUM> are isolated from one another via a first initiator <NUM> of the two initiators <NUM>, <NUM>, and the propellant grain 26a, 28a for the second and third pulses <NUM>, <NUM> are isolated from one another via a second initiator <NUM>. In exemplary embodiments, only one initiator may be provided, and in still other exemplary embodiment, more than two initiators <NUM>, <NUM> may be provided. For example, the initiation of an initiator may cause combustion of more than one propellant grain. The initiators <NUM>, <NUM> are operatively coupled to respective ones of the propellant grain 26a, 28a, to initiate combustion in the propellant grain 26a, 28a. Coupling may include direct coupling, such as via an initiator <NUM>, <NUM> being proximate or contiguous with the propellant grain 26a, 28a. In addition to the isolation provided via the initiators <NUM>, <NUM>, in other embodiments, one or more propellant grain may be isolated from other propellant grain via a casing wall or another inert material that is not ignitable during normal use of the rocket motor <NUM>. The propellant grain 24a for the first pulse <NUM> may also be initiated by another initiator in other exemplary embodiments.

Initiation, such as ignition, of the initiator <NUM>, <NUM> will generate heat and/or a flame front that is great enough to ignite the corresponding propellant grain 26a, 28a. The initiator <NUM>, <NUM> includes at least one pair of electrodes <NUM> configured to ignite an electrically operated propellant <NUM>. The electrically operated propellant <NUM> may be applied over the propellant grain 26a, 28a. One or more pairs of electrodes <NUM> may be used. In exemplary embodiments, the pair of electrodes <NUM> may be embedded in the electrically operated propellant <NUM>. The initiator <NUM>, <NUM> may extend radially outwardly from a center longitudinal axis of the casing <NUM> to seal against the inner surface of the casing <NUM>. Accordingly, the initiator <NUM>, <NUM> serves as a barrier between the propellant grain 26a, 28a and an environment external to the rocket motor <NUM>, while also providing structural support to retain the propellant grain 26a, 28a in the casing <NUM>. In other embodiments, the initiator <NUM>, <NUM> may have any other suitable shape.

The initiators <NUM>, <NUM> may have a circular or cylindrical shape to fit within the cylindrical casing <NUM>. This circular shape may be provided in the form of thin sheets of material. The thin sheets may have sufficient thickness to account for degradation of a surface of the initiator <NUM>, <NUM> when not ignited but still exposed to high heat and combustion environments within the casing <NUM>. For example, combustion of the propellant grain 26a, 28a may cause high heat in the casing <NUM>, which will not ignite the electrically operated propellant of the auxiliary initiator <NUM> when an electrical input is not applied across the auxiliary initiator <NUM>. A thin portion of the auxiliary or secondary initiator <NUM> exposed to the high heat and combustion environment may be configured to degrade. The initiators <NUM>, <NUM> may have any suitable shape to provide operative isolation of the propellant grain 26a, 28a from one another.

The electrically operated propellant <NUM>, and thus the initiators <NUM>, <NUM>, are configured to ignite in response to an electrical input and to generate gas when ignited, such that the initiators <NUM>, <NUM> form a gas generation system for the rocket motor <NUM>. The electrically operated propellant <NUM> is configured to transition from an unignited state to an ignited state when a respective electrical input is applied across the initiator <NUM>, <NUM> between the pair of electrodes <NUM>. The electrically operated propellant <NUM> is also configured to maintain the unignited state when the electrical input is not applied.

An exemplary electrically operated propellant <NUM> ignites with the application of electricity and correspondingly extinguishes with the cessation of electricity, even when exposed to high pressures, though below a high pressure threshold. For example, when exposed to ambient or high pressures within the casing <NUM>, such as atmospheric pressure, pressures greater than <NUM> psi, <NUM> psi, <NUM> psi, <NUM> psi and up to <NUM> psi, the electrically operated propellant <NUM> is extinguished with the interruption of electricity (e.g., voltage or current) applied across the electrically operated propellant.

Exemplary embodiments of the electrically operated propellant <NUM> include a propellant having a plurality of components such as an oxidizer, a fuel, and a binder. The electrically operated propellant <NUM> may include between approximately <NUM> and <NUM> percent by mass of an oxidizer, such as a perchlorate based oxidizer. The perchlorate based oxidizer may include perchlorate based oxidizers such as aluminum perchlorate, barium perchlorate, calcium perchlorate, lithium perchlorate, magnesium perchlorate, perchlorate acid, strontium perchlorate, sodium perchlorate and the like. The electrically operated propellant <NUM> may include approximately <NUM> to <NUM> percent by mass of fuel, such as a metal based fuel that assists in the ignition and extinguishing of the electrically operated propellant <NUM>. The metal based fuel may include tungsten, magnesium, copper oxide, copper, titanium and aluminum. Many other configurations of the electrically operated propellant <NUM> may be suitable.

The electrically operated propellant <NUM> is also configured to maintain its shape when exposed to dynamic kinematic conditions in the rocket motor <NUM>. The electrically operated propellant <NUM> may be formable, e.g. cast or molded, into any number of grain configurations. Accordingly, the burn rate and other performance characteristics of the electrically operated propellant <NUM> are maintained throughout the operation of the rocket motor <NUM>. Exemplary performance characteristics include total impulse value, ignition rise time, peak pressure, and weight propellant density.

As shown in <FIG>, a power source <NUM> is provided to generate the electrical input for igniting the initiators <NUM>, <NUM>. A pair of leads <NUM> electrically couples each of the initiators <NUM>, <NUM> to the power source <NUM>. The leads <NUM> extend between the respective initiators <NUM> and <NUM> and the power source <NUM>, which may be a battery or any other suitable device capable of generating electrical input. The leads <NUM> may be wires, such as insulated wires having materials capable of withstanding the high heat generated in the rocket motor <NUM>. In other embodiments, more than one power source <NUM> may be included and more than one pair of leads <NUM> may provide electrical input across one or more of the initiators <NUM>, <NUM>.

For each of the electrically operated propellant initiators <NUM>, <NUM>, the pair of electrodes <NUM> couples the respective initiator to the respective pair of leads <NUM>. As shown in <FIG>, one of the electrodes of the pair of electrodes <NUM> is coupled between a lead <NUM> of the pair of electrodes <NUM> and the initiator <NUM>, <NUM>. The electrodes <NUM> are configured to allow control of the particular location or portion of the initiator <NUM>, <NUM> that is ignited, the speed at which the initiator <NUM>, <NUM> burns, and the intensity of the burn.

Referring now to <FIG>, one of the pair of electrodes <NUM> is a ground plane electrode <NUM> and the other of the pair of electrodes <NUM> is an ignition electrode <NUM> that are configured such that an electric field <NUM> of the electrical input is concentrated at the ignition electrode <NUM> to ignite the electrically operated propellant <NUM> at the location where the ignition electrode <NUM> is arranged. The ground plane electrode <NUM> extends along a first surface area of the propellant grain 26a, 28a that is greater than a second surface area of the propellant grain 26a, 28a along which the ignition electrode <NUM> extends. The ignition electrode <NUM> has a greater current density as compared with the ground plane electrode <NUM>, which is defined as the amount of electric current per unit area of cross section, as compared with the ground plane electrode <NUM>. The ratio of the covered surface area of the propellant grain 26a, 28a from the ground plane electrode <NUM> to the ignition electrode <NUM> may be at least <NUM>:<NUM>.

The ground plane electrode <NUM> may be formed of wires having a diameter that is larger than the diameter of the wires of the ignition electrode <NUM>. For example, the diameter of the wires of the ground plane electrode <NUM> may be between <NUM> and <NUM> millimeters (between <NUM> and <NUM> inches), and the wires of the ignition electrode <NUM> may have a diameter that is between <NUM> and <NUM> millimeters (between <NUM> and <NUM> inches. The wires may have any other suitable dimensions and may be sized up or down depending on the application. During ignition, the electrical input will flow between the electrodes <NUM>, <NUM> and the electric field <NUM> will be concentrated at the ignition electrode <NUM>. An electro-chemical reaction ignition <NUM> occurs at the highest current density, as schematically shown in <FIG>. Consequently, burning of the electrically operated propellant <NUM> is initiated.

Location of burn may be controlled via the location of the electrodes <NUM>, <NUM>, e.g. where the wires forming the ignition electrode <NUM> are arranged. The electrodes <NUM>, <NUM> may be arranged adjacent the propellant grain 26a, 28a, as shown in <FIG>. The electrodes <NUM>, <NUM> may be embedded in the electrically operated propellant <NUM>, such that the electrodes <NUM>, <NUM> are contiguous with an outer surface <NUM> of the propellant grain 26a, 28a. The outer surface <NUM> of the propellant grain 26a, 28a extends along and is flush with the electrically operated propellant <NUM>, such that the outer surface <NUM> directly touches the electrically operated propellant <NUM>. Only the ignition electrode <NUM> may be embedded in the electrically operated propellant <NUM>.

The location and intensity of a pulse ignition is also controlled via the shape and form of the electrodes <NUM>, <NUM>. For example, the intensity and temperature of the burn is greater adjacent the ignition electrode <NUM> since the ignition electrode <NUM> has a greater current density, which is defined as the amount of electric current per unit area of cross section. If there is a large enough disparity between current density provided by the electrodes <NUM>, <NUM>, burning, such as the initial burning upon initiation of electrical input, may only take place adjacent the electrode providing the highest current density, such as the ignition electrode <NUM> shown in <FIG>.

The electrodes <NUM>, <NUM> may be formed of wires having many different configurations. At least the ignition electrode <NUM> is formed of a refractory metal, such as tungsten. Other suitable materials include metals and alloys having melting points that are greater than <NUM> degrees Celsius, such as molybdenum, tantalum, niobium, chromium and rhenium. Metals and alloys having melting points above <NUM> degrees Celsius may be suitable, including vanadium, hafnium, titanium, zirconium, ruthenium, osmium, rhodium, and iridium. Still other materials and alloys thereof may be suitable. In exemplary embodiments, the ground plane electrode <NUM> may also be formed of a refractory metal or refractory alloy, or other metal, such as copper. If the ground plane electrode <NUM> and the ignition electrode <NUM> are both formed of a refractory metal or refractory alloy, the diameter of the electrode wires may be the same.

As shown in <FIG>, one of the electrodes <NUM>, <NUM> may be arranged in a cavity or groove <NUM> formed in the propellant grain 26a, 28a. For example, the ignition electrode <NUM> may be positioned in the groove <NUM> such that the electric field <NUM> is directed into the groove <NUM> for ignition in the groove <NUM>. More than one groove <NUM> may be formed in the propellant grain 26a, 28a and the groove <NUM> may be formed to have any suitable shape, such as a tapering or triangular shape. In other exemplary embodiments, the groove <NUM> may be formed to have a planar support surface for the electrodes <NUM>, <NUM>. More than one electrode <NUM>, <NUM> may be accommodated in the groove <NUM>. For example, both the ground plane electrode <NUM> and the ignition electrode <NUM> may be accommodated in the groove <NUM>. In still other embodiments, the groove <NUM> may have another shape, such as a circular shape. The groove <NUM> may also instead be formed in the propellant grain 26a, 28a.

The ground plane electrode <NUM> and the ignition electrode <NUM> may be arranged in a same plane, such as in the groove <NUM> or along the outer surface <NUM> of the propellant grain 26a, 28a, as shown in <FIG>. The electrodes <NUM>, <NUM> may be arranged in a plane that is parallel with the outer surface <NUM> of the propellant grain 26a, 28a. As shown in <FIG>, the electrodes <NUM>, <NUM> may be arranged in different planes. For example, the ground plane electrode <NUM> may be arranged at the outer surface <NUM> of the propellant grain 26a, 28a whereas the ignition electrode <NUM> is arranged in the groove <NUM> such that the ignition electrode <NUM> is in a different plane. If the ground plane electrode <NUM> and the ignition electrode <NUM> are arranged in the same plane, such as in <FIG>, the electric field <NUM> may extend laterally across the propellant grain 26a, 28a.

In operation, the initial burning takes place only adjacent the ignition electrode <NUM>, while initial burning may not take place at the ground plane electrode <NUM>. The electrical input may be supplied to the ground plane electrode <NUM>. Once the electrically operated propellant <NUM> is ignited, burning may take place adjacent both electrodes <NUM>, <NUM>. Using the disparity between the current densities provided by the electrodes <NUM>, <NUM>, electrical input is provided in one direction across the electrodes <NUM>, <NUM> to vary the speed and intensity of the burn.

Referring back to <FIG>, the electrical input for the electrodes <NUM>, <NUM> may be controlled via a controller <NUM> controlling the timing and direction of the electrical input across the electrically operated propellant initiators <NUM>, <NUM> from the power source <NUM>. The controller <NUM> may be any suitable device, such as a processor having an algorithm suitable for controlling the power source <NUM>. The timing and direction of the electrical input across the electrically operated propellant initiators may thus be controlled. The controller <NUM> may act autonomously or may be directed, such as wirelessly, via an operator. The electrically operated propellant initiator may be extinguished via stopping the application of electrical current across the initiator to ignite less than <NUM>% of the initiator to produce a desired propulsion pulse.

The electrodes <NUM>, <NUM> may be configured differently for each of the propellant grains 26a, 28a and the configuration will be dependent on the desired operation for the corresponding pulse <NUM>, <NUM>. The first pulse <NUM> may occur during which the propellant grain 24a is initially burned without using an electrically operated propellant. After the propellant grain 24a is burned up or exhausted, an electrical input is applied to the electrodes <NUM>, <NUM> of the initiator <NUM> to ignite the electrically operated propellant <NUM> and subsequently the propellant grain 26a in the second pulse <NUM>. After the propellant grain 26a is burned, another electrical input is applied to the electrodes to the electrodes <NUM>, <NUM> of the initiator <NUM> to ignite the electrically operated propellant <NUM> and subsequently the propellant grain 28a in the third pulse <NUM>. In an exemplary operation, the duration of the pulses may occur between <NUM> and <NUM> seconds and inter-pulse delays may occur between <NUM> and <NUM> seconds. Both of the initiators <NUM>, <NUM> may be burned up during the corresponding pulse.

Advantageously, the gas generation system using the electrode configuration described herein may be used in a multi-pulse rocket motor in which the propellant grain is operatively isolated and ignited by the initiator <NUM>, <NUM>. In contrast, conventional multi-pulse rocket motors may use complicated and expensive barrier/ignitor systems. The system and method described herein are further advantageous in that the electrode configuration enables expeditious and uniform grain ignition by directing the electric field to the location at which the ignition electrode is arranged.

Referring now to <FIG>, exemplary configurations of the electrodes <NUM>, <NUM> are shown. Many other configurations are possible. Each of the electrodes <NUM>, <NUM> may be formed of a plurality of wires that are arranged in a particular pattern. The pattern may be an ordered and/or symmetrical pattern, or in other exemplary embodiments, the pattern may not be symmetrical. The pattern may be selected to achieve a particular burn profile. As shown in <FIG>, the ignition electrode <NUM> may include a plurality of wires that intersect and are spaced from the wires of the ground plane electrode <NUM> such that the wires are separated. The wires may extend radially along the electrically operated propellant <NUM>. The surface area of the propellant grain 26a, 28a that is covered by the ground plane electrode <NUM> may be at least twice as large as the surface area covered by the ignition electrode <NUM>.

As shown in <FIG>, the wires of the ignition electrode <NUM> do not intersect with the wires of the ground plane electrode <NUM>. The wires may be arranged in zig-zag, diagonal, crisscross, rectangular, parallel, non-parallel, and/or serpentine configurations. The wires may be formed to extend in directions that are transverse to each other, as shown in <FIG>, or in directions that are parallel to each other, as shown in <FIG>. The wires may be arranged to extend along an entire surface of the propellant grain 26a, 28a, such that the electric current may move across an entirety of the propellant grain 26a, 28a. In other exemplary embodiments, the wires may extend along a portion of the propellant grain 26a, 28a that is less than an entire length. Many different configurations and patterns of the wires may be possible and the configuration of the wires is dependent on the application.

Referring now to <FIG>, a method <NUM> of operating a rocket motor, such as the rocket motor <NUM> of <FIG>, is shown in a flowchart. Step <NUM> of the method <NUM> includes applying an electrical input across an electrically operated propellant initiator <NUM> that is operatively coupled to a propellant grain 26a and includes an electrically operated propellant <NUM> and at least one pair of electrodes <NUM> (shown in <FIG>). The at least one pair of electrodes <NUM> includes a ground plane electrode <NUM> and an ignition electrode <NUM> at which the electric field <NUM> is concentrated (shown in <FIG>) for igniting the electrically operated propellant <NUM> at the location where the ignition electrode <NUM> is located. In exemplary embodiments, the ground plane electrode <NUM> extends along a first surface area of the propellant grain 26a, 28a that is larger than a second surface area that the ignition electrode <NUM> extends along. The ignition electrode <NUM> has a greater current density than the ground plane electrode <NUM>.

Step <NUM> of the method <NUM> includes igniting the electrically operated propellant <NUM> by the electric field <NUM> being concentrated at the ignition electrode <NUM> (shown in <FIG>). Step <NUM> of the method <NUM> includes initiating combustion of the propellant grain 26a via igniting the electrically operated propellant <NUM>. Step <NUM> of the method <NUM> includes maintaining operative isolation of the propellant grain 26a from a second propellant grain 28a of an auxiliary or second propulsion pulse <NUM>, <NUM> for the rocket motor <NUM>. The electrically operated propellant initiator <NUM> may be completely burned during the pulse.

Claim 1:
A rocket motor (<NUM>) comprising:
a combustion chamber (<NUM>) containing at least one propellant grain (24a, 26a, 28a); and
an electrically operated propellant initiator (<NUM>, <NUM>) operatively coupled to the at least one propellant grain to initiate combustion of the at least one propellant grain, the electrically operated propellant initiator including an electrically operated propellant (<NUM>) and at least one pair of electrodes (<NUM>) arranged to ignite the electrically operated propellant, the at least one pair of electrodes including a ground plane electrode (<NUM>) and an ignition electrode (<NUM>) at which an electric field is concentrated to ignite the electrically operated propellant;
wherein the ground plane electrode extends along a first surface area of the at least one propellant grain that is larger than a second surface area of the at least one propellant grain along which the ignition electrode extends, wherein a ratio of the first surface area to the second surface area is at least <NUM>:<NUM>.