Patent Description:
Various applications use launchable payloads that are launched from a 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 such as a missile having a rocket motor.

Propulsion systems that use multiple pulses may be particularly suitable for use in hypersonic applications. A multi-pulse propulsion system may include a multi-pulse rocket motor, which generates multiple discrete thrust events. An exemplary use would be accelerating a payload with an initial pulse and reaccelerating the payload with one or more additional pulses, achieving increased range of the vehicle and/or achieving control authority as the vehicle approaches critical proximity of a desired target. Using a multi-pulse propulsion system is advantageous in enabling increased range, maneuverability and efficiency in launching a payload such as a warhead.

<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, includes 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 a super-staged rocket including at least approximately <NUM> rocket engines, where the engines are distributed according to at least one of: at least five multi-engine stages connected in series, each stage including at least ten engines connected in parallel; and at least five multi-stage units connected in parallel, each unit including at least five engines connected in series.

<CIT> discloses a guided projectile, especially a propelled or ballistic missile, that has its trajectory corrected by gas jets from pulse thrusters disposed in at least on axial plane of the missile symmetrically on opposite sides of the center of gravity thereof and whose thrusts are countered, when no longer needed, by the operation of diametrically opposite pulse thrusters in the same plane and at the same side of the center of gravity. The pulse thrusters are formed as gas generators which can be triggered to feed respective nozzles. The projectile is also roll stabilized, e.g. by a rotatable empennage. The transverse thrusts produced by the pulse thrusters are controlled by a sensor which responds to deviations from the correct orientation of the missile.

Conventional multi-pulse propulsion systems typically use igniters that are controlled by electronics from the mission computer (e.g., guidance system) to actively initiate each pulse ignition. Such conventional multi-pulse propulsion systems may be disadvantageous in that the duplication of system electronics and igniters for initiating each pulse of the multi-pulse motor accommodates a large volume in a volume-constrained system and adds overall cost and complexity to system. This is often at the expense of propellant volume and mass. Additionally, retrofitting an existing vehicle designed and deployed with a single pulse motor requires extensive changes to guidance system hardware and software to control the additional pulses.

An aspect of the present disclosure provides a multi-pulse rocket motor having a passive system for initiating one or more pulses of the rocket motor. This eliminates or reduces the complexity and volume constraints associated with conventional multi-pulse propulsion systems. Such a passive system, therefore, may be incorporated into smaller vehicle platforms than is currently practical with conventional multi-pulse systems, and would allow retrofit activities as described above to be implemented more cost efficiently.

According to an aspect, a multi-pulse propulsion system comprising: at least one pulse chamber containing at least one propellant for igniting during at least one pulse of the multi-pulse propulsion system; a first igniter configured to initiate the at least one pulse by igniting the at least one propellant contained in the at least one pulse chamber, the first igniter being controlled by a main system controller; at least one additional pulse chamber containing at least one additional propellant for igniting during at least one additional pulse of the multi-pulse propulsion system, wherein the at least one additional pulse is subsequent to and separate from the at least one pulse, wherein the at least one additional propellant is not ignited until the at least one additional pulse; and at least one passive fuzing system configured to initiate the at least one additional pulse, the at least one passive fuzing system comprising a sensor and a second igniter, the sensor being configured to sense an environmental condition and/or a ballistic condition, and the second igniter being configured to provide a stimulus that causes ignition of the at least one additional propellant in response to the sensor sensing that the environmental condition and/or the ballistic condition has reached or exceeded one or more threshold values, wherein the passive fuzing system is operable independent of a connection to the main system controller.

According to an embodiment of any paragraph(s) of this summary, the sensor is a passive electronic device.

According to an embodiment of any paragraph(s) of this summary, the sensor is a piezoelectric device.

According to an embodiment of any paragraph(s) of this summary, the sensor is a micro-electro-mechanical system device, a semiconductor device, a shape memory material device, or a Peltier device.

According to an embodiment of any paragraph(s) of this summary, the environmental condition sensed by the sensor includes temperature or pressure, and when a threshold temperature or pressure value is sensed by the sensor, the igniter is activated to ignite the at least one additional propellant, thereby initiating the at least one additional pulse.

According to an embodiment of any paragraph(s) of this summary, the sensor is a temperature sensor selected from: a thermistor, a thermocouple, a resistance temperature detector, or a piezoelectric temperature sensor.

According to an embodiment of any paragraph(s) of this summary, the sensor is a piezoelectric pressure sensor or a micro-electro-mechanical pressure sensor.

According to an embodiment of any paragraph(s) of this summary, the ballistic condition sensed by the sensor includes velocity and/or acceleration, and when a threshold velocity and/or acceleration value is sensed by the sensor, the igniter is activated to ignite the at least one additional propellant, thereby initiating the at least one additional pulse.

According to an embodiment of any paragraph(s) of this summary, the sensor is located in a skin of a flight vehicle that contains the multi-pulse propulsion system; or the sensor is located in a nose cone of the flight vehicle containing the multi-pulse propulsion system.

According to an embodiment of any paragraph(s) of this summary, the passive fuzing system further includes an energetic transfer part that is configured to transmit and/or transform energy from the sensor and transfer the energy to the igniter.

According to an embodiment of any paragraph(s) of this summary, the energetic transfer part transmits electrical energy from the sensor to the igniter.

According to an embodiment of any paragraph(s) of this summary, the energetic transfer part is a pyrotechnic device or pyrotechnic material that transmits thermal energy to the igniter.

According to an embodiment of any paragraph(s) of this summary, the igniter includes an explosive material.

According to an embodiment of any paragraph(s) of this summary, the igniter includes a pyrotechnic initiator having a pyrotechnic material composition.

According to an embodiment of any paragraph(s) of this summary, the at least one pulse chamber is separated from the at least one additional pulse chamber by a barrier, and the igniter is disposed in the at least one additional pulse chamber and coupled to a surface of the barrier.

According to an embodiment of any paragraph(s) of this summary, the at least one pulse chamber is separated from the at least one additional pulse chamber by a barrier, and the igniter is located in the at least one additional pulse chamber at an opposite end from the barrier.

According to another aspect, a flight vehicle includes: a payload; and a multi-pulse propulsion system according to the previous aspect.

According to an embodiment of any paragraph(s) of this summary, the flight vehicle has a diameter in the range from <NUM> to <NUM>.

According to another aspect, a method of operating a multi-pulse propulsion system of a flight vehicle, comprising: sensing an environmental condition and/or a ballistic condition of the flight vehicle with a sensor of a passive fuzing system of the multi-pulse rocket motor; determining with the passive fuzing system when the environmental condition and/or the ballistic condition has reached or exceed a threshold environmental condition value and/or a threshold ballistic condition value; and when the environmental condition and/or the ballistic condition has reached or exceed the threshold environmental condition value, initiating a second pulse of the multi-pulse propulsion system by igniting a propellant contained in a second pulse chamber of the multi-pulse propulsion system, wherein the multi-pulse propulsion further includes; a first pulse chamber containing at least one first propellant for igniting during a first pulse of the multi-pulse propulsion system; and a first igniter configured to initiate the first pulse by igniting the first propellant contained in the first pulse chamber, wherein the first igniter being controlled by a main system controller; wherein the passive fuzing system is operable independent of a correction to the main system controller, wherein the first pulse is prior to and separate from the second pulse, wherein the at propellant in the second pulse chamber is not ignited until the second pulse.

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 and aspects described herein have application in defense applications, such as in a hypersonic vehicle or any flight vehicle where space may be constrained. The multi-pulse propulsion system described herein may be implemented in a rocket that includes a multi-pulse rocket motor and carries a warhead. Other suitable applications may include different launching platforms or vehicles that include multi-pulse propulsion systems for launching a payload.

Referring initially to <FIG>, a propulsion system <NUM> is shown arranged in a flight vehicle <NUM>, such as a hypersonic vehicle or a rocket. The flight vehicle <NUM> includes a payload module <NUM> having at least one launchable payload. Any suitable payload may be arranged in the payload module <NUM> and the payload module <NUM> may include a plurality of payloads. Exemplary payloads include satellites, space probes, cargo, or warheads.

The propulsion system <NUM> may be arranged in a separable stage of the flight vehicle <NUM>. The flight vehicle <NUM> may have any suitable number of separable stages. For example, the flight vehicle <NUM> may include between two and five separable stages that are separable from the flight vehicle <NUM> at pre-determined times during travel of the flight vehicle <NUM>. In an exemplary application, the flight vehicle <NUM> may include at least a first stage and a second stage. The first stage may include a first-stage propulsion device <NUM> that is arranged opposite the payload module <NUM>. The first-stage propulsion device <NUM> may include engines, boosters, tail fins, other thrusters, or any other suitable propulsion devices.

Referring now particularly to <FIG>, embodiments of the exemplary propulsion system <NUM> are shown in further detail. As shown, the propulsion system <NUM> includes a multi-pulse rocket motor <NUM> for providing at least two distinct propulsive pulses. In exemplary embodiments, the multi-pulse rocket motor <NUM> may be a dual-pulse rocket motor that burns in two segments such that the rocket motor has a first pulse state and an additional pulse state that is initiated after the first pulse state. It is understood that the multi-pulse rocket motor <NUM> may include any number of pulse states as may be desired for a particular application.

As shown in the illustrated embodiment, the rocket motor includes at least a first pulse chamber <NUM> containing a first burnable propellant <NUM> for burning during the first pulse state, and a second pulse chamber <NUM> containing a second burnable propellant <NUM> for burning during a second pulse state. A first igniter <NUM> is located in the first pulse chamber <NUM> and is configured to initiate the first pulse state by igniting the first burnable propellant, which causes pressurization in the first pulse chamber <NUM>. The first pulse chamber <NUM> is configured for pressurization prior to pressurization of the second pulse chamber <NUM>, such as by separating the pulse chambers <NUM>, <NUM> via a barrier <NUM>. The pressurized gas in the first pulse chamber <NUM> is exhausted through nozzle assembly <NUM>, thereby generating thrust. As is well-known in the art, the first igniter <NUM> is controlled by an electronic controller <NUM>, such as the main vehicle computer <NUM> (e.g., guidance system) to actively initiate the first pulse ignition.

The first burn propellant <NUM> contained in the first pulse chamber <NUM> may have different characteristics as compared with the second burnable propellant <NUM> contained in the second pulse chamber <NUM>. For example, the propellants <NUM>, <NUM> may be configured to provide different burning rates relative to each other. The pulse chambers <NUM>, <NUM> may be formed to have different sizes such that different amounts of the propellants <NUM>, <NUM> may be provided. The sizes and burn rates of the propellants <NUM>, <NUM> and pulse chambers <NUM>, <NUM> may be dependent on the desired operation for a particular application of the flight vehicle <NUM>.

In exemplary embodiments, the propellants <NUM>, <NUM> are solid propellant grains that are configured to burn when ignited to produce exhaust gas in the corresponding pulse chambers <NUM>, <NUM>. The exhaust gas is directed through the nozzle assembly <NUM> to produce thrust for the flight vehicle <NUM>. The shape and size of the propellant grains is predetermined to achieve a specific burn time, amount of exhaust gas, and a thrust level. As shown, the pulse chambers <NUM>, <NUM> and thus the propellants <NUM>, <NUM> are separated by the barrier <NUM>, which may be any suitable separation device configured such that during the first pulse state of the flight vehicle <NUM>, the first propellant <NUM> burns separately relative to the second propellant <NUM> which burns during the second pulse state of the flight vehicle <NUM>. The barrier <NUM> may include a bulkhead of the flight vehicle <NUM> and/or the barrier <NUM> may include a suitable thermal barrier, for example.

Referring particularly to <FIG> and <FIG>, exemplary embodiments of a passive fuzing system <NUM> of the multi-pulse rocket motor <NUM> are shown in further detail. The passive fuzing system <NUM> is configured to initiate the second pulse state of the motor via ignition of the second propellant <NUM>. Any number of passive fuzing systems <NUM> may be provided for initiating the second pulse state, or any subsequent pulse state via ignition of additional propellant in subsequent pulse chambers. As discussed in further detail below, the passive fuzing system <NUM> may include one or more passive components or parts that are operable without external triggering logic. In this manner, initiation of the second pulse state and/or subsequent pulse states of the rocket motor <NUM> may be accomplished through activation of the passive fuzing system <NUM> independently of initiation by an electronic controller, and more particularly independent of the main vehicle electronics (e.g., main controller <NUM>) that initiates the first pulse state of the rocket motor <NUM>. This eliminates or reduces the complexity and volume constraints associated with conventional multi-pulse propulsion systems. Such a passive fuzing system <NUM>, therefore, may be incorporated into smaller vehicle platforms than is currently practical with conventional multi-pulse systems.

As shown in the illustrated embodiments, the passive fuzing system <NUM> generally includes at least one sensor <NUM> and at least one igniter assembly <NUM>. The sensor(s) (or sensing part(s)) <NUM> is/are configured to sense an environmental condition and/or ballistic condition of the flight vehicle <NUM> during flight. As shown, the igniter assembly <NUM> may include an energetic transfer device <NUM> (or part) and an igniter <NUM> (or initiator part). As described in further detail below, the igniter assembly <NUM>, and more particularly the igniter <NUM>, is configured to provide a stimulus that causes ignition of the second burnable propellant <NUM> in response to the sensor <NUM> sensing that the environmental condition and/or ballistic condition has reached or exceeded one or more threshold values.

The sensor <NUM>, or sensing part, of the passive fuzing system <NUM> may be any suitable sensor or sensing part that is capable of sensing or detecting an environmental condition and/or ballistic condition of the flight vehicle <NUM>. Multiple sensors or sensing parts may be utilized for achieving such functionality, which may be singly or collectively referred to hereinafter as sensor <NUM> for simplicity. The sensor <NUM> may have any suitable construction for sensing the desired condition. In exemplary embodiments, one or more of the sensor(s) <NUM> is/are passive devices, such as passive electronic devices, that are operable to sense the desired condition without an external logic controlling the firing current by way of an external electrical signal. Examples of such passive electronic devices include, but are not limited to, one or more of a piezoelectric device, a micro-electro-mechanical system (MEMS) device, a semiconductor device, a shape memory material device, a Peltier device, or the like.

The environmental and/or ballistic condition sensed by the sensor <NUM> may be any desired environmental and/or ballistic condition. In exemplary embodiments, the environmental condition sensed by the sensor <NUM> is temperature and/or pressure. By way of example, and not limitation, a suitable passive temperature sensor may include a thermistor, a thermocouple, a resistance temperature detector, a piezoelectric temperature sensor, or any other suitable passive temperature sensing device or part. By way of example, and not limitation, a suitable passive pressure sensor may include a piezoelectric pressure sensor, a MEMS device, pitot tube, pressure transducer, or any other suitable passive pressure sensing device or part. In exemplary embodiments, the ballistic condition sensed by the sensor <NUM> is velocity (e.g., linear or angular velocity) and/or acceleration/deceleration. By way of example, and not limitation, a suitable passive ballistic condition sensor may include a MEMS device, such as an accelerometer, or any other suitable passive velocity/acceleration sensing device or part, such as a gyroscope (e.g., a passive resonator gyroscope).

The sensor(s) <NUM> of the passive fuzing system <NUM> may be located at any suitable part of the flight vehicle <NUM> for sensing the desired condition. For example, the sensor(s) <NUM> may have an external face that is exposed to the external environment of the flight vehicle <NUM>, which may be atmospheric (e.g., troposphere, stratosphere, mesosphere, thermosphere, exosphere) or outer space. As shown in the illustrated embodiments, for example, the sensor(s) <NUM> may be located on a skin <NUM> of the flight vehicle <NUM>. Alternatively or additionally, the sensor(s) <NUM> may be located in a nose cone <NUM> (see <FIG>) of the flight vehicle <NUM>, for example.

The energetic transfer device <NUM> (or energetic transfer part) of the passive fuzing system <NUM> is configured to transmit energy from the sensor <NUM> to the igniter <NUM>. As such, the energetic transfer device <NUM> is operably coupled to both the sensor <NUM> and the igniter <NUM>, which may include other components or parts interposed there between for facilitating operation of the passive fuzing system <NUM>. The energetic transfer device <NUM> may be located at any suitable part of the flight vehicle <NUM> for providing such operable connection between the sensor <NUM> and igniter <NUM>, which such placement may be dependent on the type of energetic transfer device <NUM>. One or more energetic transfer devices <NUM> may be provided. The energy transferred by the energetic transfer device <NUM> may be any form of energy dependent on the particular type of sensor <NUM>, and the energy from the sensor <NUM> may be transformed to another type of energy by the energetic transfer device, or other connected component, for transferring to the igniter <NUM>.

By way of example, and not limitation, the energetic transfer device <NUM> may transmit electrical energy in the form of current provided by the sensor <NUM> to the igniter <NUM>, which may activate the igniter <NUM> to cause ignition of the second burnable propellant <NUM>. In such a scenario, the energetic transfer device <NUM> may include an electrical wire, such as a copper wire, for transmitting the current from the sensor <NUM> to the igniter <NUM>.

Alternatively or additionally, the energetic transfer device <NUM> may include a pyrotechnic device or material that transmits thermal energy to the igniter <NUM>. By way of example, and not limitation, the pyrotechnic device may include a detonation cord, such as a mild detonation cord, for example shielded mild detonation cord, flexible detonation cord, or the like. The pyrotechnic device may include a detonation cord assembly, for example, an RDX filled transfer line with insulative and structural overwraps. The pyrotechnic device may include a suitable blasting cap. The pyrotechnic device may be any other suitable type of fuse.

Alternatively or additionally, the energetic transfer device <NUM> may provide a combination of electrical and thermal impulses for causing the igniter <NUM> to ignite the propellant <NUM>. For example, an electrical wire may be provided for transmitting an electrical stimulus from the sensor <NUM> to a blasting cap, which in response to a threshold electrical stimulus by the sensor <NUM>, may detonate the blasting cap to thereby transmit thermal energy that ignites a detonation cord, which thereby transmits thermal energy to activate the igniter <NUM>. Alternatively, for example, a bridgewire may be provided with a pyrotechnic material that is activated when the bridgewire is heated by resistance beyond a threshold level in response to an electrical current from the sensor <NUM>.

The igniter <NUM> may be any suitable device or part that provides a sufficient stimulus to cause ignition of the propellant <NUM> when the igniter <NUM> has been activated or triggered. For example, the igniter <NUM> may include an explosive material, such as a shape charge, which is triggered to explode by energy transferred by the energetic transfer device <NUM> (e.g., wire or detonation cord). The heat from the explosion of the shape charge will initiate the second pulse via burning of propellant <NUM>.

Alternatively or additionally, the igniter <NUM> may include a pyrotechnic initiator, which may have a pyrotechnic material composition that causes a self-sustained exothermic chemical reaction when activated to make heat sufficient to burn the propellant <NUM>. The composition of the pyrotechnic initiator may be adjusted to tune the activation temperature of the igniter <NUM>. Examples of such pyrotechnic initiators include, but are not limited to: metal-oxidizers (e.g., zirconium - potassium perchlorate, boron - potassium nitrate, aluminum-potassium perchlorate, or titanium-aluminum-potassium perchlorate); metal hydride-oxidizer (e.g., zirconium hydride - potassium perchlorate, titanium hydride potassium perchlorate); intermetallics (e.g., titanium-boron, nickel-aluminum, palladium-aluminum); or the like.

The igniter <NUM> may be located at any suitable part of the flight vehicle <NUM> for providing the stimulus that ignites the second propellant <NUM>. In exemplary embodiments, the igniter <NUM> is located in the second pulse chamber <NUM> with sufficient proximity to the propellant <NUM> to cause the propellant to burn when the igniter <NUM> is activated or triggered.

Referring briefly to <FIG>, for example, the igniter <NUM> may be located proximally to, or coupled to, a surface of the barrier <NUM>. When activated, the igniter <NUM> will break down the barrier <NUM> to open the pulse chamber <NUM> to the nozzle assembly <NUM>. The igniter <NUM> may break down the barrier <NUM> via explosive force (e.g., when the igniter <NUM> includes a shape charge) and/or the igniter <NUM> may break down the barrier <NUM> via thermal energy (e.g., when the igniter <NUM> includes a pyrotechnic material). Also when the igniter <NUM> is activated, the igniter <NUM> will provide sufficient energetic stimulus to ignite the propellant <NUM>, thereby causing the propellant <NUM> to burn and produce exhaust gas in the chamber <NUM>. With the barrier <NUM> destroyed, the exhaust gas is directed through the nozzle assembly <NUM> to produce thrust for the flight vehicle <NUM>.

Referring briefly to <FIG>, for example, the passive fuzing system <NUM> is shown at a different location of the flight vehicle <NUM> as compared to <FIG>. As shown, the igniter <NUM> is located at an opposite (e.g., forward) end of the chamber <NUM> away from the barrier <NUM>. The igniter <NUM> may be the same as that described above in connection with <FIG>. In the embodiment illustrated in <FIG>, when the igniter <NUM> is activated, the igniter <NUM> will provide sufficient energetic stimulus to ignite the propellant <NUM>, thereby causing the propellant <NUM> to burn and produce exhaust gas in the chamber <NUM>. The pressure and/or temperature generated by burning the propellant <NUM> will break down barrier <NUM> to open the chamber <NUM> to the nozzle assembly <NUM>. With the barrier <NUM> destroyed, the exhaust gas is directed through the nozzle assembly <NUM> to produce thrust for the flight vehicle <NUM>.

Turning now to <FIG>, an exemplary method <NUM> of operating the multi-pulse rocket motor <NUM> is shown. As shown, the method <NUM> starts at step <NUM> with initiating the first pulse state of the rocket motor <NUM>, which may be accomplished in a conventional manner as described above.

At step <NUM>, the sensor(s) <NUM> of the passive fuzing system <NUM> sense the environmental and/or ballistic condition of the flight vehicle <NUM>. This is accomplished as described above, in which the sensor(s) <NUM> may be configured to sense temperature, pressure, acceleration, velocity, or any other suitable environmental and/or ballistic condition.

At step <NUM>, the passive fuzing system <NUM> is configured to determine whether the environmental and/or ballistic condition sensed by the sensor(s) <NUM> has reached or exceeded a threshold value. As shown at step <NUM>, when the sensed condition has reached or exceeded the threshold value, the second pulse state is initiated. The threshold value may be any suitable value which may be adjusted depending on the configuration of the passive fuzing system <NUM>. By way of example, and not limitation, a threshold temperature value that causes initiation of the second pulse state may be in a range from <NUM> - <NUM> °F (<NUM> - <NUM>). By way of example, and not limitation, a threshold pressure value that causes initiation of the second pulse state may be in a range from <NUM> - <NUM> psi (<NUM> - <NUM> Pa). By way of example, and not limitation, a threshold acceleration value that causes initiation of the second pulse state may be in a range from <NUM> - <NUM> ft/sec<NUM> (<NUM> -<NUM>/s<NUM>).

The determination at step <NUM> may be made in any suitable manner dependent on the particular configuration of the passive fuzing system <NUM>. For example, the energy output from the sensor <NUM> may vary proportionally to the input condition sensed by the sensor <NUM>. At step <NUM>, when the input condition reaches the threshold value, the energy output from the sensor <NUM> may reach a threshold value that causes, directly or indirectly, activation or triggering of the igniter <NUM> which initiates the second pulse state of the motor <NUM>, as described above. On the other hand, if the input condition sensed by the sensor <NUM> does not reach or exceed the threshold value, then the proportional output from the sensor <NUM> may be insufficient to activate, directly or indirectly, the igniter <NUM>. In this scenario, the sensor(s) <NUM> will continue to sense the environmental and/or ballistic conditions, as shown in the method <NUM> by looping back to step <NUM>.

As discussed above, multiple sensors <NUM> may be used to sense the same or different environmental and/or ballistic conditions, and the respective outputs of these sensors <NUM> may be combined to reach the activation threshold that causes, directly or indirectly, activation of the igniter <NUM>. Such combination of sensor outputs may enable the passive fuzing system <NUM> to require multiple threshold values to be met prior to initiating the second pulse state. As noted above, such energy output from the sensor(s) <NUM> may be transferred directly to the igniter <NUM> in the same form, such as via the energetic transfer device <NUM> (e.g., electrical wire); or the energy output from the sensor(s) <NUM> may be transferred indirectly by the energetic transfer device <NUM> transforming the energy to another form (e.g., blasting cap and/or detonation cord, for example).

A non-limiting example of a passive fuzing system <NUM> according to the foregoing will now be described in further detail. In an exemplary system, the sensor <NUM> is a passive piezoelectric pressure sensor that generates an electrical output signal in response to strain applied to the sensor. The output from the piezoelectric sensor is a charge proportional to the sensed pressure. A small battery may be connected to amplify this charge and/or provide sufficient impetus to ignite a propellant grain. The energetic transfer device <NUM> is an electric wire that transmits this electrical charge to the igniter <NUM>. The igniter <NUM> is a shape charge explosive that is configured to detonate when the charge transmitted from the pressure sensor reaches a threshold activation value. This threshold activation value may be achieved when the strain applied to the sensor exceeds a threshold value in response to the environmental pressure sensed reaching a threshold value.

Another non-limiting example of a passive fuzing system <NUM> may include a sensor <NUM> in the form of a passive piezoelectric pressure sensor, an energetic transfer device <NUM> in the form of a blasting cap and detonation cord operably coupled to the piezoelectric sensor, and an igniter <NUM> in the form of an intermetallic pyrotechnic material (e.g., Pd/Al) that is operably coupled to the detonation cord. In an exemplary system, the sensor <NUM> is a passive piezoelectric pressure sensor that generates an electrical output signal in response to strain applied to the sensor. The output from the piezoelectric sensor is a command charge triggered by the sensed pressure. A battery may be connected to amplify this charge. The energetic transfer device <NUM> is blasting cap, the output of which triggers a detonation cord that transmits a combination of heat and/or shock to an igniter. The igniter <NUM> contains intermetallic pyrotechnic material that is initiated when the charge output is transmitted from the detonation cord.

Turning to <FIG>, trajectory shaping of the flight vehicle <NUM> may be utilized for optimally activating the passive fuzing system <NUM> and thereby initiating the second pulse state of the rocket motor <NUM>. In <FIG>, a long range trajectory of the flight vehicle <NUM> is shown, in which the flight vehicle <NUM> rockets to a high elevation (such as via the first pulse), and when the threshold condition is met, the second pulse is initiated. In such a scenario, the threshold condition for initiating the second pulse may be a reduction in atmospheric pressure as the vehicle <NUM> increases in altitude, an increase in temperature as the vehicle <NUM> re-enters denser atmosphere, or reduction in acceleration as the vehicle <NUM> loses velocity, for example. In <FIG>, a short range trajectory of the flight vehicle <NUM> is shown, in which higher density air may be utilized for reaching the threshold condition (e.g., via heating as the vehicle <NUM> rockets through the atmosphere).

Claim 1:
A multi-pulse propulsion system (<NUM>) comprising:
at least one pulse chamber (<NUM>) containing at least one propellant (<NUM>) for igniting during at least one pulse of the multi-pulse propulsion system;
a first igniter (<NUM>) configured to initiate the at least one pulse by igniting the at least one propellant contained in the at least one pulse chamber, the first igniter being controlled by a main system controller (<NUM>);
at least one additional pulse chamber (<NUM>) containing at least one additional propellant (<NUM>) for igniting during at least one additional pulse of the multi-pulse propulsion system, wherein the at least one additional pulse is subsequent to and separate from the at least one pulse, wherein the at least one additional propellant is not ignited until the at least one additional pulse; and
at least one passive fuzing system (<NUM>) configured to initiate the at least one additional pulse, the at least one passive fuzing system comprising a sensor (<NUM>) and a second igniter (<NUM>), the sensor being configured to sense an environmental condition and/or a ballistic condition, and the second igniter being configured to provide a stimulus that causes ignition of the at least one additional propellant in response to the sensor sensing that the environmental condition and/or the ballistic condition has reached or exceeded one or more threshold values, wherein the passive fuzing system is operable independent of a connection to the main system controller.