Abstract: An electronic initiation system for use with a firing panel may include an input connector, a plurality of electronic ignition circuits (EICs) operably coupled together in series, and an indicator operably coupled to an output of each EIC of the plurality of EICs. A first EIC of the plurality of EICs may be operably coupled to the input connector. The indicator is configured to generate an indication in response to an output of an EIC of the plurality of EICs satisfying a predetermined condition.

Abstract: A universal tube launched munitions system has a launcher base having an array of a plurality of initiator stations, first and second cassettes each having a mounting facility operable for releasable connection to the launcher base, each of the cassettes having a plurality of barrels, the barrels of the first cassette been axially registered with the first quantity of the initiator stations, and the barrels of the second cassette been axially registered with a different second quantity of the initiator stations. The barrels of the first cassette may be axially registered with all of the initiator stations. The barrels of the second cassette may be axially registered with a subset of the initiator stations the barrels of the first cassette are axially registered with. The barrels of the first cassette may have a different diameter than the barrels of the second cassette.

Abstract: An unmanned aerial vehicle (UAV) system provides for UAV deployment and remote, unattended operation with reduced logistics requirements. The system includes a launcher that can include one or more launch tubes, each launch tube configured to house a UAV in a canister and one or more gas generators operatively connected to each canister and configured to push the UAV out of the launch tube by releasing gas into the canister. A controller for activating the gas generators can sequentially, and with a predetermined time delay, expel the UAV with a desired velocity and acceleration. The system further includes a UAV recovery device, a power supply, a security subsystem, a command and control subsystem and a communications subsystem. Command, control and communications can be provided between a remote station and the UAV.

Abstract: Provided are a controlling apparatus for ejecting a guided missile and a method for ejecting the guided missile inserted into a launching tube. More particularly, The present invention provides a controlling apparatus for ejecting a guided missile additionally including a power supply controller which applies a signal for ejecting a guided weapon to a fuse during separating the fuse used for igniting a propulsion engine in the guided weapon and an ejection controlling method using the same.

Abstract: A technique for launching a missile that avoids some of the costs and disadvantages for doing so in the prior art. In particular, the illustrative embodiment of the present invention uses an electromagnetic catapult to throw the missile clear of the launch platform—with sufficient velocity to attain aerodynamic flight—before the missile's engine is ignited.

Abstract: A power source is provided for a resistive heating element that heats a catalyst bed of a reactor that receives a monopropellant fuel and reacts the fuel in the presence of the catalyst bed to provide hot gas driving a turbine. The power source is a capacitor that discharges its charge across the resistive heating element so that its temperature is elevated to a point where the reaction of the monopropellant fuel occurs. The power source may be used in airborne applications, such as emergency power systems for aircraft, where it is desirable to minimize the weight of the unit and to allow the unit to use safer monopropellant fuels.

Abstract: A launch system for use as a standalone munition launcher or as a guest launcher within a main battery host launcher.

Abstract: A method and apparatus for thrust vectoring a flight vehicle during homing are disclosed. The flight vehicle includes a body; a plurality of attitude control mechanisms on the body; a multi-pulse motor housed at least partially in the body; and a control unit housed in the body. The control unit generates signals controlling an actuation of the attitude control mechanisms and the multi-pulse motor to initiate a follow-on burn of the multi-pulse motor to effect a maneuver directed by the actuation of the attitude control mechanisms. The method includes conserving a second burn of a dual pulse motor until target acquisition; acquiring a target; actuating a attitude control mechanism for the flight vehicle to alter the flight vehicle's heading; and initiating the second burn.

Abstract: A rocket assisted payload launch system includes a heavy gauge metal shipping container that houses an array of containerized concentric launch tubes, or canisters. Spacers on each containerized concentric launch tube contact the adjacent tubes and retain the centers of the tubes a predetermined distance apart, and in a vertical orientation. A missile including rocket assisted payload (RAP) can be inserted into the open upper end of each tube, which also has a curved, closed bottom end. Umbilical cords are connected to each tube for providing target location information to the missile placed therein. A sequence controller selects the rocket assisted payload to be fired, and initiates the firing sequence by delivering power, via the umbilical cord, to the selected missile. The sequence controller may be manually operated, or may be operated, automatically, via a datalink, in the preferred mode of operation.

Abstract: The present invention relates to a device for deploying ammunition, wherein a recess (2) provided in the body shell (1) of a vehicle, in particular of a ship, is covered by cover means (4), whereby a heightening of the radar signature caused by this recess (2) and having a negative effect is diminished.

Abstract: A launcher (10) is provided having a separation unit (15) and a launch tube (16) in fluid communication with the separation unit (15). The separation unit (15) has a pressure sensing chamber (23) for sensing the pressure within the separation unit (15), a separation chamber (24) in fluid communication with the pressure sensing chamber (23), and a combustion chamber (26) in fluid communication with the separation chamber (24). The combustion chamber (26) is in fluid communication with the launch tube (16) through opening (28). The rocket launcher (10) also includes a hydrogen separation circuit (21) and an ignition circuit (22). The hydrogen separation circuit (21) includes a battery pack (44) electrically coupled to a fuel cell (45) which when powered separates water into hydrogen gas and oxygen gas. The ignition circuit (22) includes a piezoelectric spark generator (61) for igniting the hydrogen gas produced by the fuel cell and contained within the combustion.

Abstract: A lightweight multiple missile launch assembly (10) adapted to carry newer heavier Maverick missiles (AGM-65D, E, F, G or K). The invention is illustrated with a BRU-33, 55 or 57 bomb rack (12) adapted to carry two LAU-117 missile launchers (24, 26). The missile launchers (24, 26) are secured to the bomb rack (12) with a suspension unit (20, 22) of unique and novel design. The suspension units (20, 22) allow for releasable engagement of the missile launchers (24, 26). This allows for different missile packages to be deployed in the field.

Abstract: A missile system includes a launcher having a built-in pulse power source and an electrical connection to a missile. The pulse power source is used to send activation and firing pulses to the missile. The pulse power source utilizes a piezoelectric crystal which produces a pulse of energy when struck by a spring-loaded hammer that is released by an operator. In one version, a first piezoelectric crystal produces a first pulse to activate the system when struck by a first hammer, and a second piezoelectric crystal produces a second pulse to fire the missile when struck by a second hammer. The energy pulses are conducted from the launcher to the missile through an inductive coupling between a primary induction coil in the launcher and a facing secondary induction coil in the missile.