Methods and apparatuses for aerial interception of aerial threats

Embodiments include active protection systems and methods for an aerial platform. An onboard system includes radar modules, detects aerial vehicles within a threat range of the aerial platform, and determines if any of the aerial vehicles are an aerial threat. The onboard system also determines an intercept vector to the aerial threat, communicates the intercept vector to an eject vehicle, and causes the eject vehicle to be ejected from the aerial platform to intercept the aerial threat. The eject vehicle includes alignment thrusters to rotate a longitudinal axis of the eject vehicle to substantially align with the intercept vector, a rocket motor to accelerate the eject vehicle along an intercept vector, divert thrusters to divert the eject vehicle in a direction substantially perpendicular to the intercept vector, and attitude control thrusters to make adjustments to the attitude of the eject vehicle.

TECHNICAL FIELD

Embodiments of the present disclosure relate generally to methods and apparatuses for engagement management relative to a threat and, more particularly, to aerial interception of aerial threats.

BACKGROUND

Rocket Propelled Grenades (RPGs) and other human carried projectiles such as Man-portable Air-Defense Systems (MANPADS or MPADS) and shoulder-launched Surface-to-Air Missiles (SAMs) represent serious threats to mobile land and aerial platforms. Even inexperienced RPG operators can engage a stationary target effectively from 150-300 meters, while experienced users could kill a target at up to 500 meters, and moving targets at 300 meters. One known way of protecting a platform against RPGs is often referred to as active protection and generally causes explosion or discharge of a warhead on the RPG at a safe distance away from the threatened platform. Other known protection approaches against RPGs and short range missiles are more passive and generally employ fitting the platform to be protected with armor (e.g., reactive armor, hybrid armor or slat armor).

Active Protection Systems (APS) have been proposed for ground vehicles for defense against RPGs and other rocket fired devices with a good success rate for quite some time. However, these systems are proposed to protect vehicles that are: 1) armored, 2) can carry heavy loads, and 3) have plenty of available space for incorporation of large critical systems. Currently these systems can weigh anywhere between 300 to 3000 lbs. and can protect the vehicle when intercepting incoming threats as close as 5 to 10 ft.

There is a need in the art for engagement management systems that can work in cooperation with intercept vehicles to engage and destroy aerial threats. There is also a need for such systems to be portable and lightweight enough for carrying on aerial and other mobile platforms that may have significant weight and size constraints, or on which an active protection system may be easily installed. There is also a need for such systems to coordinate with multiple engagements of aerial threats, intercept vehicles, and other nearby engagement management systems.

DETAILED DESCRIPTION

In the following description, reference is made to the accompanying drawings in which is shown, by way of illustration, specific embodiments of the present disclosure. The embodiments are intended to describe aspects of the disclosure in sufficient detail to enable those skilled in the art to practice the invention. Other embodiments may be utilized and changes may be made without departing from the scope of the disclosure. The following detailed description is not to be taken in a limiting sense, and the scope of the present invention is defined only by the appended claims.

Furthermore, specific implementations shown and described are only examples and should not be construed as the only way to implement or partition the present disclosure into functional elements unless specified otherwise herein. It will be readily apparent to one of ordinary skill in the art that the various embodiments of the present disclosure may be practiced by numerous other partitioning solutions.

In the following description, elements, circuits, and functions may be shown in block diagram form in order not to obscure the present disclosure in unnecessary detail. Additionally, block definitions and partitioning of logic between various blocks is exemplary of a specific implementation. It will be readily apparent to one of ordinary skill in the art that the present disclosure may be practiced by numerous other partitioning solutions. Those of ordinary skill in the art would understand that information and signals may be represented using any of a variety of different technologies and techniques. For example, data, instructions, commands, information, signals, bits, symbols, and chips that may be referenced throughout the description may be represented by voltages, currents, electromagnetic waves, magnetic fields or particles, optical fields or particles, or any combination thereof. Some drawings may illustrate signals as a single signal for clarity of presentation and description. It will be understood by a person of ordinary skill in the art that the signal may represent a bus of signals, wherein the bus may have a variety of bit widths and the present disclosure may be implemented on any number of data signals including a single data signal.

The various illustrative logical blocks, modules, and circuits described in connection with the embodiments disclosed herein may be implemented or performed with a general-purpose processor, a special-purpose processor, a Digital Signal Processor (DSP), an Application Specific Integrated Circuit (ASIC), a Field Programmable Gate Array (FPGA) or other programmable logic device, discrete gate or transistor logic, discrete hardware components, or any combination thereof designed to perform the functions described herein. A general-purpose processor may be a microprocessor, but in the alternative, the processor may be any conventional processor, controller, microcontroller, or state machine. A general-purpose processor may be considered a special-purpose processor while the general-purpose processor is configured to execute instructions (e.g., software code) stored on a computer-readable medium. A processor may also be implemented as a combination of computing devices, such as a combination of a DSP and a microprocessor, a plurality of microprocessors, one or more microprocessors in conjunction with a DSP core, or any other such configuration.

In addition, it is noted that the embodiments may be described in terms of a process that may be depicted as a flowchart, a flow diagram, a structure diagram, or a block diagram. Although a process may describe operational acts as a sequential process, many of these acts can be performed in another sequence, in parallel, or substantially concurrently. In addition, the order of the acts may be rearranged.

Elements described herein may include multiple instances of the same element. These elements may be generically indicated by a numerical designator (e.g.,110) and specifically indicated by the numerical indicator followed by an alphabetic designator (e.g.,110A) or a numeric indicator preceded by a “dash” (e.g.,110-1). For ease of following the description, for the most part, element number indicators begin with the number of the drawing on which the elements are introduced or most fully discussed. For example, where feasible, elements inFIG. 3are designated with a format of 3xx, where 3 indicatesFIG. 3and xx designates the unique element.

Embodiments of the present disclosure include apparatuses and methods for providing protection for mobile platforms, such as, for example, a helicopter, from an aerial threat. Some embodiments of the present disclosure may include methods and apparatuses that are portable and lightweight enough for carrying on aerial platforms that may have significant weight and size constraints. Some embodiments of the present disclosure may include methods and apparatuses that can be incorporated into existing systems already installed on aerial platforms.

FIGS. 1A and 1Billustrate a helicopter as an aerial platform100that may be under attack from an aerial threat120and coverage areas140that may be employed to sense when such a threat is present within an intercept range (may also be referred to herein as a threat range) of embodiments of the present disclosure. As shown inFIG. 1A, the aerial threat120may be shot by an attacker110toward the aerial platform100.

As used herein, “aerial threat” or “threat” are used interchangeably to refer to any threat directed toward a mobile platform, including projectiles, rockets, and missiles that may be shoulder launched or launched from other platforms. As non-limiting examples, such aerial threats include Rocket Propelled Grenades (RPGs), Man-portable Air-Defense Systems (MANPADS or MPADS), shoulder-launched Surface-to-Air Missiles (SAMs), Tube-launched, Optically tracked, Wire-guided missiles (TOWs), and other aerial weapons, having a trajectory and ordnance such that they may cause damage to the mobile platform.

The term “aerial platform” includes, but is not limited to, platforms such as helicopters, Unmanned Airborne Vehicles (UAVs), Remotely Piloted Vehicles (RPVs), light aircraft, hovering platforms, and low speed traveling platforms. The protection systems and methods of the present disclosure are particularly useful for protecting aerial platforms against many kinds of aerial threats.

While embodiments of the present disclosure may be particularly suitable for use on aerial platforms100due to the small size and weight, they may also be used in other types of mobile platforms like ground-based mobile platforms such as, for example, tanks, armored personnel carriers, personnel carriers (e.g., Humvee and Stryker vehicles) and other mobile platforms capable of bearing embodiments of the present disclosure. Moreover, embodiments of the present disclosure may be used for relatively stationary ground-based personnel protection wherein a mobile platform may not be involved. Accordingly, embodiments of the disclosure are not limited to aerial applications.

FIG. 1Billustrates coverage areas140in which one or more embodiments of the present disclosure may detect an incoming aerial threat120and perform active countermeasures using one or more embodiments of the present invention to remove the aerial threat120before it can damage the aerial platform100. Some embodiments of the present disclosure may be configured such that they can be disposed in previously existing Countermeasures Dispenser Systems (CMDS).

FIGS. 2A and 2Billustrate a dispenser200configured as a conventional CMDS (e.g., an AN/ALE-47) in which an eject vehicle400(EV) according to one or more embodiments of the present disclosure may be placed. AN/ALE-47 dispensers are conventionally used to dispense passive countermeasures, such as, for example, radar-reflecting chaff, infrared countermeasures to confuse heat-seeking missile guidance, and disposable radar transmitters. With some embodiments of the present disclosure, eject vehicles400may also be placed in the AN/ALE-47 and ejected therefrom under control of the AN/ALE-47 and other electronics on the aerial platform100(FIGS. 1A and 1B). The eject vehicle400may be configured as a substantially cylindrical vehicle to be placed in a tubular dispenser210and ejection may be controlled from control wiring220connected to the dispenser200. Moreover, the dispenser200may be configured to hold both the passive countermeasures for which it was originally designed, as well as one or more eject vehicles400according to embodiments of the present disclosure.

While some embodiments of the eject vehicle400may be configured to be disposed in an AN/ALE-47, other types of dispensers200or other types of carriers for the eject vehicle400may also be used. Moreover, the tubular dispenser210is illustrated with a circular cross section. However, other cross sections may be used, such as, for example, square, hexagonal, or octagonal.

FIG. 3illustrates systems that may be present on a helicopter frame300and that may intercommunicate according to one or more embodiments of the present disclosure. The helicopter frame300and systems described are used as specific examples to assist in giving details about embodiments of the present disclosure. In the specific example ofFIG. 3, an AAR-47 Missile Approach Warning System (MAWS) warns of threat missile approaches by detecting radiation associated with the missile. In the specific example, four MAWSs (320A,320B,320C, and320D) are disposed near four corners of the helicopter frame300. A central processor360may be used to control and coordinate the four MAWSs (320A,320B,320C, and320D).

Two AN/ALE-47 dispensers (200A and200B) are positioned on outboard sides of the helicopter frame300, each of which may contain one or more eject vehicles400. As shown inFIG. 3, there are four eject vehicles400on each side labeled EV1through EV4on one side and labeled EV5-EV8on the other side. The AN/ALE-47 dispensers (200A and200B) are each controlled by an AN/ALE-47 sequencer (350A and350B), which are, in turn, controlled by the central processer360.

According to one or more embodiments of the present disclosure four radar modules (900A,900B,900C, and900D) are included to augment and connect with the AAR-47s and communicate with the eject vehicles400. These radar modules900(seeFIG. 9A) are configured to detect and track relatively small incoming aerial threats (e.g., an RPG) as well as the outgoing eject vehicles400. Moreover, the radar modules900can send wireless communications (340A,340B,340C, and340D) to the eject vehicles400both before and after they are ejected from the dispensers (200A and200B). The radar modules900, and eject vehicles400may each include unique identifiers, such as, for example, a Media Access Control (MAC) address. The radar modules900may also be configured to detect, track, and communicate with other friendly platforms such as, for example, other helicopters flying in formation with the helicopter. Thus, all helicopters within communication range can communicate and share radar and control information to form a broad coverage area, similar to cellular telephone base station coverage. Moreover, and as explained more fully below, the helicopters may communicate to define different sector coverage areas such that one helicopter does not launch an eject vehicle400into a sector that may damage or interfere with another helicopter.

The control processors, such as the central processor360, the MAWSs320, the radar modules900, the sequencers350, and the dispensers200may be configured to form an ad hoc network and include the eject vehicles400.

The specific example ofFIG. 3is shown to illustrate how radar modules (900A-900D) and eject vehicles (EV1-EV8) of the present disclosure can be incorporated with existing systems on helicopter platforms with little change. Of course, other systems may be employed with embodiments of the present disclosure. As a non-limiting example, one radar900A may be positioned on one side of the helicopter frame300and another radar module900C may be positioned on another side of the helicopter frame. In such a case, the radar modules900would be configured to provide hemispherical coverage areas. These radar modules900may be controlled by, communicate with, or a combination thereof, a different central processor360configured specifically for embodiments of the present disclosure. Moreover, the eject vehicles400may be disposed in different carriers or different dispensers from the AN/ALE-47 dispensers (200A and200B) shown inFIG. 3.

When embodiments of the present disclosure are used as illustrated inFIG. 3, they provide an ultra-lightweight active protection system for helicopter platforms that may increase the survivability against RPG attacks to better than 90% for RPGs fired from ranges as close as about 100 meters away.

In order to satisfy the helicopter platform constraints, embodiments of the present disclosure address many significant technology areas:

1) For helicopter applications, size, weight, and power should be considered. Every pound of added airframe equipment will reduce capacity to carry personnel or cargo, and the space for adding equipment to the airframe may be at a premium. At least some embodiments of the present disclosure are configured to be less than about 50 pounds and occupy about 5.5″×5.5″ surface area at each of the four corners of a helicopter exterior shell and with minimal impact to existing wiring kits.

2) Helicopters generally do not carry armor and thus, the intercept of an incoming threat (e.g., an RPG) must occur at a range that is safe to the un-armored helicopter airframe. Using an RPG-7 as an example, to achieve a survival probability of about 99% from the blast alone, the intercept should occur at distances beyond 30 meters from the helicopter shell. This requirement significantly influences the system response time, when considering that an RPG fired at a 100-meter distance may impact the helicopter in less than about 600 milliseconds.

3) A third concern is fratricide and collateral damage to friendly forces that may be amplified by the helicopter platform deploying kinetic countermeasures in a position above ground and potentially next to a wingman helicopter or in the vicinity of civilians, friendly troops, or a combination thereof. Some embodiments of the present disclosure are configured to work in combination with embodiments on other helicopters when the helicopters are flying in formation relatively close to each other.

4) Some embodiments of the present disclosure can geo-locate the attacker110(FIG. 1A) after a few radar track frames are processed.

5) Embodiments of the present disclosure can engage multiple threats at a time. In other words, multiple incoming aerial threats120can be detected and tracked and multiple outgoing eject vehicles400can be tracked. In addition, to increase a probability of destroying an incoming aerial threat120, multiple eject vehicles400may be launched, directed toward, and detonated proximate the same aerial threat120.

6) Finally, eject vehicles400can be launched and guided to the point of attack with the same or different warheads and detonated above the threat point of origin.

To address these technology areas, some embodiments of the present disclosure include an active kinetic countermeasure projectile (i.e., the eject vehicle400ofFIG. 2B), including an ejection mechanism with an impulse charge that can fit in, and can be launched by, the AN/ALE-47 chaff/flare dispenser200. Some embodiments of the present disclosure include the radar module900covering a 90 degree sector or more (i.e., with a 90 degree sector, each helicopter platform would use four radar modules900).

When referring to the radar module900herein (e.g., as shown inFIG. 3), it should be understood that in some embodiments the radar module900may perform the operations described herein in combination with other electronics and processors on the aerial platform100. As such, the radar modules900may be used to: 1) search, acquire, and track incoming aerial threats120, 2) launch the active kinetic countermeasure (i.e., eject vehicle400), 3) track the outgoing eject vehicle400with respect to the incoming aerial threat120, 4) point and guide the eject vehicle400toward the incoming aerial threat120, 5) command detonate the eject vehicle400, and 6) geo-locate the attacker110, all in less than about one second. In one configuration, at least two AN/ALE-47 dispensers200would be used in conjunction with the four radar modules900such that each dispenser200provides hemispherical coverage.

The radar modules900may be configured as pulse Doppler radar modules900to scan the azimuth plane and the elevation plane using two orthogonal fan beams and may be configured to cover a 90 degree sector in about 20 milliseconds. Upon detecting an incoming aerial threat120, the associated radar module900may then direct the launch and guidance of an eject vehicle400from an AN/ALE-47 dispenser200that covers that sector. The eject vehicle400may be command guided to the target by the radar module900and command detonated. The radar modules900may be configured as an addition to the existing AN/AAR-47 system and may use its existing interface for launching of the eject vehicle400.

Some of the embodiments of the present disclosure may be configured to deploy an eject vehicle400that fits in a standard dispenser200but could be stabilized and pointed towards the threat after launch, in less than about 50 milliseconds, in the rotor downwash of a helicopter, and when ejected in the fixed direction dictated by the dispenser200. The radar modules900may then guide the eject vehicle400to accurately intercept the aerial threat120within about 330 milliseconds and thus reduce the requirement of carrying a large warhead.

FIG. 4illustrates an exploded view of an eject vehicle400showing various elements of the eject vehicle400according to one or more embodiments of the present disclosure. Reference may also be made toFIGS. 1A-3in describing features and operations of the eject vehicle400. The eject vehicle400is a lightweight guided projectile that, in some embodiments, may be designed to be launched from chaff/flare dispensers. The eject vehicle400may intercept and destroy incoming aerial threats120at ranges sufficient to prevent damage to the host aerial platform100. The eject vehicle400may be packaged in a cartridge containing an impulse charge and interface electronics designed to fit the AN/ALE-47 dispenser magazine.

The eject vehicle400includes an ejection piston780configured to transmit the energy of an impulse cartridge750(described below in connection withFIG. 7) to the eject vehicle400and launch the eject vehicle400away from the aerial platform100to a distance safe enough for the eject vehicle400to begin performing alignment and interception maneuvers.

A rocket motor420may be used to propel the eject vehicle400toward the aerial threat120after the eject vehicle400has been rotated such that a longitudinal axis of the eject vehicle400is pointed in the general direction of the aerial threat120. A first set of folding fins482may be attached to the rocket motor420and configured to deploy once the eject vehicle400has exited the dispenser200. The folding fins482are small and configured to provide stability to the eject vehicle400during its flight path rather than as control surfaces for directing the fight path.

An airframe shell430may be configured to contain a warhead440, a divert thruster module610, a nose thruster module620(may also be referred to herein as an alignment thruster module620), an electronics module450, and a battery452. An airframe nose490may be configured to attach to the airframe shell430to protect the electronics module450and provide a somewhat aerodynamic nose for the eject vehicle400.

A safe and arm module460may be included within the airframe shell430and configured to safely arm the warhead440when the eject vehicle400is a safe distance away from the aerial platform100.

FIGS. 5A-5Cillustrate the eject vehicle400ofFIG. 4as it may be configured during various stages of an intercept mission according to one or more embodiments of the present disclosure. Stage1, inFIG. 5A, illustrates the eject vehicle400in a cartridge710(FIG. 7) and includes the ejection piston780, the rocket motor420, the airframe shell430, and the airframe nose490.

Stage2, inFIG. 5B, illustrates the eject vehicle400after it has been dispensed and shows the rocket motor420, the airframe shell430, and the airframe nose490.FIG. 5Balso illustrates the folding fins482deployed near the end of the rocket motor420and wireless communication antennas890deployed near the airframe nose490.

Stage3, inFIG. 5Cillustrates the eject vehicle400after the rocket motor420has burned and been detached from the airframe shell430. At this stage, the eject vehicle400may be referred to as a terminal vehicle and includes the airframe nose490, the wireless communication antennas890, and the airframe shell430. Still within the airframe shell430are the warhead440, the divert thruster module610, the alignment thruster module620, the electronics module450, the battery452, and the safe and arm module460. After the rocket motor420is detached, a second set of folding fins484are deployed from the airframe shell430to stabilize the eject vehicle400during the remainder of the flight to intercept the aerial threat120. This second set of folding fins484are used to replace the first set of folding fins482that were attached to the rocket motor420, which has been detached from the airframe shell430during stage3.

In addition, after the rocket motor420is detached, one or more corner reflectors470are exposed. The corner reflector470may be configured with sharp angles to enhance radar detection of the eject vehicle400by a radar module900on the aerial platform100. For example, the corner reflector470may be configured as an interior angle of a small cube shape, which will enhance radar detection.

Returning toFIG. 4, the alignment thruster module620is offset from a center of mass of the eject vehicle400such that an initial pitch maneuver can be performed to align the longitudinal axis of the eject vehicle400along an intercept vector pointed toward the aerial threat120. This alignment maneuver is performed prior to the burn of the rocket motor420.

The divert thruster module610is positioned substantially near a center of mass of the terminal vehicle and is used to laterally divert the terminal vehicle from its current flight path to make minor corrections to the flight path in order to more accurately intercept the aerial threat120. The terminal vehicle may be referred to herein as the eject vehicle400and it should be understood what is being referred to based on the context of the discussion.

The warhead440may be command detonated when the radar module900on the aerial platform100determines that the eject vehicle400has reached the closest point of approach (nominally about 15 cm). The use of thrusters, provide the fast reaction times that may be needed to intercept the aerial threat120at a nominal distance of about 50 meters when the aerial threat120is launched from a range of about 100 meters.

FIGS. 6A-6Cillustrate various propulsion and thruster elements that may be included with one or more embodiments of the present disclosure.FIG. 6Aillustrates a nose thruster module620with four nose thrusters622(two are hidden) arranged around a periphery of the nose thruster module620. These nose thrusters622(also referred to herein as alignment thrusters622) are positioned to generate a perpendicular force on the eject vehicle400relative to the longitudinal axis and are offset from the center of mass of the eject vehicle400so that an initial pitch maneuver can be performed to rotate and align the longitudinal axis of the eject vehicle400along an intercept vector pointed toward the aerial threat120. In this embodiment, the four nose thrusters622are orthogonally arranged giving two opportunities to adjust the pitch of the eject vehicle400in each direction. Of course, other embodiments may include fewer or more alignment thrusters622.

FIG. 6Billustrates a divert thruster module610with eight divert thrusters612(five are hidden) arranged around a periphery of the divert thruster module610. These divert thrusters612are positioned to generate a perpendicular force on the eject vehicle400relative to the longitudinal axis and are positioned near the center of mass of the eject vehicle400so that the divert thrusters612will move the eject vehicle400laterally to a slightly different travel path, while substantially maintaining the same pitch. Thus, the divert thrusters612can modify the flight path of the eject vehicle400to correct for minor errors in the initial pitch maneuvers pointing directly toward the aerial threat120. In this embodiment, eight divert thrusters612are used giving eight opportunities to adjust the flight path of the eject vehicle400during its flight toward the aerial threat120. Of course, other embodiments may include fewer or more divert thrusters612.

FIG. 6Cillustrates a thruster650configured to expel a gas through a nozzle652to create a lateral force. The thruster650may be controlled from a thrust signal654, which may be connected to the electronics module450of the eject vehicle400. The thruster650is one example of a type of thruster that may be used for both the divert thrusters612and the alignment thrusters622.

FIG. 7illustrates various electrical and communication connections that may be present on the eject vehicle400while it is disposed on the aerial platform100(FIGS. 1A and 2B) prior to launch. A cartridge710includes a cartridge flange720such that the cartridge710may be securely placed in a dispenser200(FIG. 2A). An end cap790may be positioned over the cartridge710to hold the eject vehicle400within the cartridge710. An impulse cartridge750is positioned near the base of the cartridge flange720and is configured to fire in response to a fire command signal755from the radar module900(FIG. 3) or other electronics on the aerial platform100. An ejection piston780is positioned between the impulse cartridge750and the eject vehicle400and is configured to transmit the energy of the firing impulse cartridge750to the eject vehicle400and propel the eject vehicle400out of the dispenser200and safely away from the aerial platform100.

A power signal740and a ground signal730may run along or through the cartridge to an antenna spring contact745and a ground spring contact735, respectively. The ground spring contact735is configured to flexibly couple with a ground patch738on the eject vehicle400to provide a ground for the eject vehicle400electronics while the eject vehicle400is in the cartridge710. The antenna spring contact745is configured to flexibly couple with the antenna890on the eject vehicle400and a power signal on the eject vehicle400to provide power and direct communication for the eject vehicle400electronics while the eject vehicle400is in the cartridge710. The cartridge710may include a cartridge antenna760that may be coupled to the antenna890of the eject vehicle400by the antenna spring contact745. Thus, the eject vehicle400may communicate wirelessly795with electronics onboard the aerial platform100through the antenna890on the eject vehicle400or through the cartridge antenna760.

FIG. 8is a block diagram illustrating elements that may be present on the eject vehicle400according to one or more embodiments of the present disclosure. A microcontroller810may be coupled to a memory820, which is configured to hold instructions for execution by the microcontroller810and data related to command and control of the eject vehicle400. The microcontroller810may be any suitable microcontroller, microprocessor, or custom logic configured to directly execute, or execute responsive to software instructions, processes related to operation of the eject vehicle400. The memory820may be any suitable combination of volatile and non-volatile memory configured to hold data and computing instructions related to operation of the eject vehicle400.

One or more antennas890may be configured to provide a communication link with electronics (e.g., the radar module900) onboard the aerial platform100. As non-limiting examples, the communication link may be configured using WiFi or WiMax frequencies and protocols. A diversity combiner880may be used to combine signals from multiple antennas.

A communication transceiver870(e.g., a WiFi transceiver) may be coupled to the diversity combiner880and be configured to transmit and receive frequencies to and from the diversity combiner880. A communication modem860(e.g., a WiFi modem) may be coupled to the communication transceiver870and be configured to package and modulate communication information for communication transmission as well as demodulate and extract information from communication reception. The microcontroller810receives information from the communication modem860and may perform operations related to the received information. In addition, based on processes performed on the microcontroller810, information may be sent to the communication modem860for transmission through the one or more antennas890.

The microcontroller810may be coupled to a thrust controller830, which interfaces with the alignment thrusters622and the divert thrusters612(FIGS. 6A and 6B). A warhead fuzing interface840may be provided to interface to the warhead440(FIG. 4), the safe and arm module460(FIG. 4) or a combination thereof, for arming and control of detonation of the warhead440.

A roll sensor850and a vertical reference855may be used in combination to determine the attitude of the eject vehicle400as well as a spin rate and spin position of the eject vehicle400and communicate such information to the microcontroller810. Other types of sensors, such as, for example, accelerometers and magnetometers may also be used for this purpose.

FIG. 9Ais a block diagram illustrating elements that may be present on the aerial platform100according to one or more embodiments of the present disclosure. The electronics module and functions thereof on the aerial platform100may be contained within a radar module900, as illustrated inFIG. 9B. Alternatively, some of the function may be within the radar module900while other functions may be located in different places on the aerial platform100such as, for example, the central processor360(FIG. 3). The various modules used to control the radar module900and the eject vehicle400and determine other information related thereto may be collectively referred to herein as an “onboard system”.

FIG. 9Bis a perspective view of the radar module900that may be present on the aerial platform100according to one or more embodiments of the present disclosure. The radar module900includes an azimuth scan radar antenna920, an elevation scan radar antenna940, and a wireless communication link antenna960.

The azimuth scan radar antenna920is included in an azimuth radar subsystem, which includes a diplexer922for combining radar sent and reflected radar received. A Radio Frequency (RF) up/down converter925converts the radar frequencies sent from a digital synthesizer930and converts the radar frequencies received for use by a digital receiver935.

The elevation scan radar antenna940is included in an elevation radar subsystem similar to the azimuth radar subsystem, but configured for the elevation direction. The elevation radar subsystem includes a diplexer942for combining radar sent and reflected radar received. A Radio Frequency (RF) up/down converter945converts the radar frequencies sent from a digital synthesizer950and converts the radar frequencies received for use by a digital receiver955.

The wireless communication link antenna960may be configured to provide a communication link with electronics onboard the eject vehicle400. As non-limiting examples, the communication link may be configured using WiFi or WiMax frequencies and protocols. A wireless communication subsystem includes a communication transceiver965(e.g., a WiFi transceiver) coupled to the wireless communication link antenna960and configured to transmit and receive frequencies to and from the antenna960. A communication modem970(e.g., a WiFi modem) may be coupled to the communication transceiver965and be configured to package and modulate communication information for communication transmission as well as demodulate and extract information from communication reception.

A sector processor910communicates with the elevation radar subsystem, the azimuth radar subsystem, and the wireless communication subsystem. The sector processor910may communicate helicopter navigation information912from other electronics on the aerial platform100. Referring also toFIG. 3, the sector processor910may also communicate with the dispenser200(e.g., one or more ALE-47s) using communication signal914and the Missile Approach Warning System320(e.g., one or more AAR-47s) using communication signal916. The sector processor910performs a number of functions to detect and track aerial threats120, control and track the eject vehicle400, as well as other functions related to the active protection system. In some embodiments, communication between the dispenser200and the sector processor910may be accomplished through the Missile Approach Warning System320.

The sector processor910in combination with the radar subsystems can detect and track incoming aerial threats120(e.g., RPGs). Based on the tracking of the incoming aerial threat120, and in combination with navigation information from the aerial platform100, the sector processor910can extrapolate to a geo-location of the attacker110, from where the aerial threat120was launched. The aerial platform100may act on this geo-location or transmit the geo-location to other aerial platforms or ground-based platforms for follow-up actions.

The sector processor910may be configured to send launch commands to the dispenser200on communication signal914to launch one or more eject vehicles400to intercept one or more detected aerial threats120. The sector processor910may also calculate required pitch adjustments that should be performed by the eject vehicle400after it has been ejected and is safely away from the aerial platform100.

Once the eject vehicle400is launched, the sector processor910may be configured to track the eject vehicle400and send guidance commands (i.e., divert commands) to the eject vehicle400so the eject vehicle400can perform divert maneuvers to adjust its flight path toward the aerial threat120. The sector processor910may also be configured to determine when the eject vehicle400will be near enough to the aerial threat120to destroy the aerial threat120by detonation of the warhead440on the eject vehicle400. Thus, a detonation command may be sent to the eject vehicle400instructing it to detonate, or instructing it to detonate at a detonation time after receiving the command.

FIGS. 10A and 10Bare diagrams illustrating radar scanning beams during an acquisition mode and a tracking mode, respectively. Referring toFIGS. 10A, 10B, 9, and 3, the radar modules900may be mounted in close proximity to the existing AN/ALR-47 missile warning receiver (MWR) installations to provide 360 degrees spatial coverage while minimizing wiring modifications to the helicopter. It is anticipated that an aerial threat120will be launched at relatively short ranges, typically on the order of 100 m. The radar modules900are designed to detect and track the low radar cross section (typically −15 dBsm) of an RPG fired from any aspect angle, within 30 milliseconds of launch, and out to a range of at least 300 meters. The radars operate in the Ka-Band to minimize the antenna size yet provide the precision angular measurements needed to guide the eject vehicle400to intercept the aerial threat120. A high pulse-repetition-frequency pulse Doppler waveform provides radial velocity measurements as well as the clutter rejection needed to operate in close proximity to the ground while detecting low radar cross section targets. Pulse compression may be used to achieve precision range measurements as well as increasing the transmit duty cycle to best utilize the capabilities of existing Ka-Band solid-state power amplifiers. The antennas generate a pair of orthogonal fan beams, providing a continuous track-while-scan capability to minimize detection latency and provide multiple target track capability. Beam scanning can be accomplished using a frequency scan method to eliminate the need for expensive phase shifters.

FIG. 10Aillustrates an acquisition mode wherein an elevation radar generates an elevation fan beam extending in the vertical direction that sweeps in the horizontal direction and an azimuth radar generates an azimuth fan beam extending in the horizontal direction that sweeps in the vertical direction. Thus, an entire 90-degree scan sector can be covered by the radar systems to quickly detect and acquire an incoming aerial threat120when it is within range.

FIG. 10Billustrates a track mode. InFIG. 10B, two sequential azimuth scans and two sequential elevation scans are shown that pinpoint a first location1010of the eject vehicle400. In addition, two sequential azimuth scans and two sequential elevation scans are shown that pinpoint a second location1020of the aerial threat120. With this location information, the sector processor910can derive relative position information that can be used to provide divert commands to the eject vehicle400to more closely intercept the aerial threat120.

FIG. 11is a spectrum diagram illustrating possible Doppler spectrum regions where various aerial vehicles may be detected. As non-limiting examples,FIG. 11illustrates a ground clutter spectrum1110, a spectrum1120for the eject vehicle400(i.e., PRJ inFIG. 11), a spectrum1130that may be indicative of an RPG, and a spectrum1140that may be indicative of a MANPAD. Of course, other aerial threats and their associated spectrums may also be identified.

FIG. 12is a simplified flow diagram illustrating some of the processes1200involved in one or more embodiments of the present disclosure. The processes may be loosely considered as an acquisition phase1210, a pre-launch phase1220, an align and launch phase1240, a guidance phase1260, a divert phase1270, and a detonation phase1280.

Operation block1212indicates that continuous radar scans are performed looking for incoming aerial threats. Decision block1214indicates that the process loops until a target is detected. While not shown, during this phase the radar modules900may also be detecting distance and angle to wingman platforms (i.e., other aerial platforms) in the vicinity. Using communication between the various wingman platforms, sectors of responsibility can be identified as discussed more fully below in connection withFIG. 14.

If a target is detected, the process1200enters the pre-launch phase1220. Operation block1222indicates that the sector processor910uses the range and travel direction of the incoming aerial threat120to calculate a threat direction to the incoming aerial threat120and an intercept vector pointing from a deployed eject vehicle400to a projected intercept point where the eject vehicle400would intercept the incoming aerial threat120. Operation block1224indicates that the intercept vector is sent to the eject vehicle400. The intercept vector may be sent to the eject vehicle400in a number of forms. The actual directional coordinates may be sent and the eject vehicle400would be responsible for determining the proper pitch maneuvers to perform. Alternatively, the sector processor910may determine the proper pitch maneuvers that the eject vehicle400should perform after launch and send only pitch commands (e.g., start and burn times for each alignment thruster622) to be used during the pitch maneuvers. WhileFIG. 12indicates that the intercept vector or pitch commands are sent before launch, some embodiments may be configured such that this information can be sent after launch.

During the acquisition phase1210and pre-launch phase1220, the eject vehicle400remains in the dispenser200and connected to power. An RF communication link may be in operation through the eject vehicle400antenna via a transmission line inside the dispenser200.

The process enters the align and launch phase1240after the intercept vector is determined. Operation block1242indicates the impulse cartridge750is fired to propel the eject vehicle400from the dispenser200and safely away from the aerial platform100.

Operation block1244indicates that the pitch maneuvers are performed to align the eject vehicle400with the already determined intercept vector. The pitch maneuver is a two-stage process that sequentially executes an azimuth rotation and an elevation rotation to align the longitudinal axis of the eject vehicle400along the intercept vector. The pitch maneuver does not have to be exact. As a non-limiting example, offsets of up to about 10 to 15 degrees may be corrected during flight of the eject vehicle400using the divert thrusters612during the guidance phase1260. After ejection, the folding fins482will deploy and the communication link antennas890will deploy and wireless communication between the eject vehicle400and the radar module900may commence.

Operation block1246indicates that the rocket motor420will fire, which accelerates the eject vehicle400to about 160 meters/second and imposes a spin rate on the eject vehicle400of about 10 Hertz. Upon exhaustion, the rocket motor420and folding fins482will separate and the Terminal Vehicle (TV) is exposed. With separation of the TV, the second folding fins484deploy and the corner reflector470is exposed.

During the guidance phase1260, the process will perform a track and divert loop in order to adjust the flight path of the eject vehicle400to more closely intercept the aerial threat120. Operation block1262indicates that the sector processor910will track the eject vehicle400and aerial threat120as discussed above with reference toFIGS. 9A-10B. Decision block1264, indicates that the sector processor910will determine if a divert maneuver is required to intercept the incoming aerial threat120and estimate the direction of divert thrust required.

A divert phase1270includes operations to cause the eject vehicle400to modify its course. Operation block1272indicates that the divert direction and time, if required, are sent to the eject vehicle400.

The divert process takes into account the rotation of the eject vehicle400and the direction of the desired divert thrust. This rotation adds a complication to the selection and fire time determination of the proper divert thruster612, but also ensures that all of the available divert thrusters612can be used to divert the eject vehicle400in any desired direction substantially perpendicular to the travel direction of the eject vehicle400. Operation block1274indicates that the processor on the eject vehicle400will select the divert thruster612to be fired and determine the firing time based on the divert angle received from the sector processor910and its internal attitude sensors.

Operation block1276indicates that the appropriate divert thruster612is fired at the appropriate fire time to move the eject vehicle400laterally along a diversion vector to adjust the flight path of the eject vehicle400. As a non-limiting example, each divert thruster612may be capable of correcting for about two degrees of error from the initial pointing of the eject vehicle400during the pitch maneuver. Thus, when the divert thrusters612are fired when the eject vehicle400is in the correct rotational position, the process can slide the travel direction vector of the eject vehicle400toward the path of the aerial threat120. Moreover, the process can fire in any circular direction and can fire multiple divert thrusters612in the same direction to repeatedly move the eject vehicle400in the same direction.

WhileFIG. 12indicates the guidance phase1260and the detonation phase1280as operating sequentially, they also may operate in parallel. During the detonation phase1280, operation block1282indicates that the sector processor910determines an optimum intercept time when the eject vehicle400will be at its closest point to the aerial threat120. Operation block1284indicates that a detonation command may be sent to the eject vehicle400. This detonation command may be in the form of a detonation time for the eject vehicle400to count out or it may be in the form of an immediate command for the eject vehicle400to perform as soon as the command is received.

Operation block1286indicates that the warhead440on the eject vehicle400is detonated at the intercept time responsive to the detonation command received from the sector processor910.

FIG. 13illustrates an example flight path for the eject vehicle400and an aerial threat120during an intercept process. In this example, a typical RPG and EV trajectory example are shown. The RPG is launched at a range of about 100 meters and 30 degrees left of the nose of the helicopter. The eject vehicle400receives its coordinate commands from the radar module900and is then ejected from the port chaff dispenser200at an angle of 90 degrees to the helicopter axis.

During period1310, the eject vehicle400separates to a distance of about two meters from the helicopter. During period1320, the nose thrusters622pitch the eject vehicle400to the approximate approach angle of the incoming RPG (e.g., within about ±10° accuracy). The rocket motor420then fires to accelerate the eject vehicle400to approximately 160 meters/second and is then separated from the remaining terminal vehicle upon exhaustion.

During period1330, the radar module900transmits a series of divert commands to the eject vehicle400, which fires the divert thrusters612to correct the trajectory of the eject vehicle400and intercept the RPG. A radar command is finally sent to the eject vehicle400to detonate the warhead440when the terminal vehicle reaches the closest point of approach (CPA). The guidance algorithm may be configured to produce a maximum CPA of about 30 centimeters, which is well within the lethal 0.6-meter kill radius of the warhead440.

FIG. 14illustrates two aerial vehicles flying in a formation and various radar sectors that may be covered by the aerial vehicles. A significant concern is the presence of wingman helicopters and the potential damage caused by accidental targeting. The system presented has the capability of tracking and recognizing the adjacent helicopters and networking with their associated active protection systems to avoid collateral damage by handing off sectors covered by other platforms. InFIG. 14, a first helicopter1410is monitoring a first radar sector1410A, a second radar sector1410B, a third radar sector1410C, and a fourth radar sector1410D.

A second helicopter1420near the first helicopter1410is monitoring a fifth radar sector1420A, a sixth radar sector1420B, a seventh radar sector1420C, and an eighth radar sector1420D. If an aerial threat approaches from a direction indicated by arrow1430it may be detected by the third radar sector1410C of the first helicopter1410and the seventh radar sector1420C of the second helicopter1420. If the first helicopter1410attempts to launch an eject vehicle, it may cause damage to the second helicopter1420. However, using communication between the various wingman platforms, sectors of responsibility can be identified. Thus, for the direction indicated by arrow1430, the first helicopter1410can determine that the third radar sector1410C will be covered by the seventh radar sector1420C of the second helicopter1420. As a result, while this formation continues, the first helicopter1410does not respond to threats in its third radar sector1410C.

Returning toFIG. 9B, the radar module900may also be referred to herein more generically as an Engagement Management Module (EMM)900. As discussed above, in some embodiments, an aerial platform100(FIGS. 1A and 1B) may be configured with four engagement management modules900, an example of which is shown inFIG. 3.

The engagement management modules900may be used as part of a Helicopter Active Protection System (HAPS), but may also be used in other types of aerial vehicles, ground vehicles, water vehicles, and stationary deployments.

Returning toFIGS. 4 through 5C, the eject vehicle400may also be referred to herein as an intercept vehicle400and a kill vehicle (KV)400.FIG. 5Cillustrates the eject vehicle400after the rocket motor420has burned and been detached from the airframe shell430. At this stage, the eject vehicle400may be referred to as a terminal vehicle, a terminal section, or a terminal section of the kill vehicle400. Still within the terminal section are the warhead440, the divert thruster module610, the alignment thruster module620, the electronics module450the battery452, and the safe and arm module460.

In operation, radar on the EMM900detects and tracks an aerial threat (e.g., an RPG) launched at the helicopter and launches one or more KVs400from the AN/ALE-47 Countermeasure Dispenser to intercept the incoming RPG. Following launch, the KV400executes a series of pitch maneuvers using nose thrusters (i.e., in the alignment thruster module620) to align the body axis with the estimated intercept point, uses a boost motor to accelerate to high speed to intercept the RPG at maximum range, and executes a series of commands for lateral guidance maneuvers using a set of divert thrusters610. Finally, the KV400warhead440is command detonated when the KV400reaches the closest point of approach (CPA) computed by an EMM guidance processor.

The flight time of the KV is typically on the order of 300 to 500 msec. During the short flight time, the KV is exposed to strong moments due to divert thruster offsets from the center of gravity and to strong aerodynamic moments due principally to the Jet Interaction (JI) effect when the divert thrusters are fired. The principle forces of interest are shown inFIG. 15.

FIG. 15is a simplified side view of a kill vehicle (KV)400illustrating the principle forces of interest on the kill vehicle (KV)400. Ideally, the divert thruster's610center of thrust, XCT, would be aligned with the KV400center of gravity (CG) to minimize any moments introduced by its thrust force (FDT). However, even with balancing during manufacturing, there will be some migration of the CG as divert thrusters610and alignment thrusters622(FIGS. 6A and 6Band also referred to herein as “pitch thrusters622”) and “attitude control thrusters1604” (FIG. 16A) are fired and propellant is expended. Also, when the divert thrusters610are fired, airflow along a body1510of the KV400is disrupted, known as the Jet Interaction (JI) effect, which introduces a high pressure region1506in front of the divert thrusters610and a low pressure region1508aft of the divert thrusters610. This causes a moment that can be represented as a force FJIat location XJI. The other principle aerodynamic component is the normal force FNintroduced as a function of the angle of attack α, which is the angle between a body longitudinal axis1520and the velocity vector V. The normal force FNacts at the center of pressure XCP.

One or more tail fins1524can be used to add aerodynamic stability, but the time constants associated with these tail fins generally are not fast enough to stabilize the KV400during its short flight time. Instead, or in addition to the tail fins1524, a plurality of small micro-thrusters1604(FIG. 16A) are added to the nose thruster module620(FIGS. 4 and 16A) to implement an active attitude control. This introduces Fμat location Xμ.

FIGS. 16A and 16Billustrate a nose thruster module620. The nose thruster module620may include a plurality of pitch thrusters622and a plurality of attitude control thrusters1604.FIG. 16Aillustrates a plurality of cylinders1606and1608configured for storing propellant grains. The larger cylinders1606correspond to the pitch thrusters622while the smaller cylinders1608correspond to the attitude control thrusters1604.FIG. 16Billustrates a plurality of nozzles1610and1612, each configured to divert a flow from one of the plurality of cylinders1606,1608. Thus, the combination of cylinders1606and nozzles1610create the pitch thrusters622and the combination of cylinders1608and nozzles1612create the attitude control thrusters1604.

Referring again toFIG. 15, The various moments described earlier may cause the KV400to become unstable and likely to tumble without stabilization. Maintaining the angle of attack α within a nominal 10 to 20 degrees is useful to maintain the divert thrusters612normal to the velocity vector V for proper guidance maneuvering and in directing the warhead440blast toward the RPG at the point of detonation.

FIG. 17is a simplified block diagram of an attitude control loop1700. The EMM900radar generates position {circumflex over (d)}KVand velocity {circumflex over (v)}KVtrack data for the KV400and sends this information to the KV400via a command link radio (CLR)1706. The EMM900radar also transmits a roll angle {circumflex over (φ)}TXvia a transmit antenna1804(FIG. 18) to the CLR1706to the KV400over the CLR1706. A Joint Adaptive Polarization and Roll Angle Estimator (JAPRAE) algorithm1710, resident in the KV400, processes the polarized CLR1706signal received by two orthogonal linear receive antennas1802(FIG. 18) on the KV400to estimate a roll angle {circumflex over (φ)}KVof the KV400. An onboard Inertial Management Unit (IMU)1714measures an acceleration a, and an angular velocity ωKVof the KV400. This information is processed by an Extended Kalman Filter (EKF)1720to generate fused estimates of the position {tilde over (d)}KV, velocity {tilde over (v)}KV, attitude [{tilde over (φ)}KV, {tilde over (θ)}KV, {tilde over (ψ)}KV], and angular velocity {tilde over (ω)}KVof the KV400.

The Attitude Control Algorithm1730is designed to fire the nose mounted attitude control thrusters1604to maintain the angle of attack α (FIG. 15) within preset limits. Once it is determined that an attitude correction thrust is required, the Attitude Control Algorithm1730selects the next available attitude control thruster1604that comes into position due to the body spin at operational block1732, and times the firing to coincide with the desired direction of thrust at operational block1734. The Attitude Control Algorithm1730maintains knowledge of which attitude control thrusters1604have been fired and which are available at operational block1736.

The KV400has an onboard IMU1714that measures the angular rate and acceleration. The IMU1714may include a set of 3-axis gyros, 3-axis accelerometers, 3-axis magnetometers, and a Field Programmable Gate Array (FPGA) to partially process the sensor information. The sensors may be solid-state MEMS that are machined on a small circuit board. The IMU1714also includes an FPGA that contains the EKF1720to fuse the output of the sensors into 3-dimensional position, velocity, angular velocity and attitude.

FIG. 18is a simplified flowchart of the Joint Adaptive Polarization and Roll Angle Estimator (JAPRAE) algorithm1710. The JAPRAE algorithm1710may process the CLR1706signals received, for example, by orthogonal linearly polarized receive antennas1802, to adapt the two receive signals in a beam former1806(also referred to herein as “the dual polarized receiver1806”) to match the incoming polarization to eliminate polarization mismatch loss induced by body rotation. The JAPRAE algorithm1710also processes the CLR1706signals received to provide an estimate of the KV400roll angle {circumflex over (φ)}rx.

The electric vector field of a propagating signal can be represented by a pair of orthogonal components:

If polarizations p and q are defined in different coordinate systems, Utxand Urx, then (2) becomes:
xrx(t)=q(t)HUrxTUtxp(t)E(t)eiβ(t)(3)

Assume the polarization of a source antenna1804and the receive antennas1802can be represented as poand qoin terms of a natural antenna coordinate system. Let Atx(t) and Arx(t) be the time varying coordinate transformations from the natural coordinate system to some inertial system. In this system the electric field vector and receiver polarization vectors are given by:
Etx(t)=Atx(t)UtxpoE(t)eiβ(t)
q(t)=Arx(t)Urxqo.  (4)

In the JAPRAE system1710there are two orthogonal receive antennas1802, with polarizations qo1and qo2. Thus, the signal generated at the input of the dual polarized receiver1806is given by:

In equation (4) the transformations are explicitly represented as rotations of the transmitter φtxand receiver φrxaround the line of sight. In the case where the transmit antenna1804and the receive antennas1802are linearly polarized, it can be shown that this reduces to the following form:

Consider a signal y(t) generated by performing a rotation on the vector x(t).

Note that if {circumflex over (δ)}=δ, yrx1=Eeiβ(t)and yrx0=0. This amounts to a case where yrX1is the output of a beamformer that matches the polarization of the incident signal to maximize the receive signal power. It can be shown that this also maximizes the signal-to-noise ratio.

Now form the ratio of y2(k) to y1(k)

This forms the basis for generating an error signal for an adaptive polarization loop.

FIG. 19illustrates the JAPRAE algorithm1710(FIG. 17) in the form of a conventional servo loop1900. The closed loop may force yrx2to zero, hence {circumflex over (δ)}→δ. Under closed loop conditions, the KV400roll angle is estimated by:
{circumflex over (θ)}rx(t)={circumflex over (θ)}tx(t)−{circumflex over (δ)}(t)  (6)

The transmitter roll angle, {circumflex over (φ)}tx, can be transmitted to the KV400via the data link1806.

FIG. 20illustrates a non-limiting example of a set of simulation results2000of the JAPRAE algorithm1710(FIG. 17) tracking the roll angle of a KV400body spinning on its longitudinal axis1520(FIG. 15) at 2200 degrees per second. A top plot2002ofFIG. 20illustrates an overlay of true and estimated roll angles. A bottom plot2004illustrates a difference or error between the overlay of true and estimated roll angles. The set of simulation results2000are shown without additive noise and illustrate the performance of the servo loop1900(FIG. 19).

Note that the solution for equation 10 is ambiguous by π radians. Therefore, the loop1900ofFIG. 19should be initialized near the correct angle to avoid the ambiguity.

A Kalman Filter methodology is used to fuse the sensor data. The translational motion involves a set of linear equations and the conventional Kalman Filter (KF) can be used, whereas the rotational motion involves a set of nonlinear equations and the conventional Extended Kalman Filter (EKF) can be used.

A 9×1 discrete time translational motion state vector, x(k), includes 3-D spatial components comprising a position vector, d(k), a velocity vector, v(k), and an acceleration vector, a(k), as follows:
x(k)=[dT(k),vT(k),aT(k)]T(12)

Transition and measurement equations include a transition matrix, Fk, a plant noise coupling matrix, Gk, a measurement matrix, Hk, state transition noise wkand measurement noise vk.
xk+1=FXk+Gkwk
zk=Hkxk+vk(13)

The state transition noise wkis i.i.d. zero mean Gaussian with covariance Rk, and the measurement noise vkis also i.i.d. zero mean Gaussian with covariance Qk. The subscript k indicates the variables might be time varying.

The expanded form of these equations is given by:

[dk+1vk+1ak+1]=[I3Ts⁢I30.5⁢Ts2⁢I303I3Ts⁢I30303I3]⁡[dkvkak]+[0.5⁢Ts2⁢I3Ts⁢I3I3]⁡[wd,kwv,kwwa,k]⁢[zd,kzv,kza,k]=[I3030303I3030303I3]⁡[dkvkak]+[vd,kvv,kva,k](7)
where Tsis the sample interval and I3and 03are 3×3 unity and zero matrices, respectively.

The Kalman filter1720(FIG. 17) involves two acts: the projection act and the update act.

The projection act is given by:
{circumflex over (x)}k|k+1=Fk{circumflex over (x)}k−1|k−1
{circumflex over (P)}k|k+1=Fk{circumflex over (P)}k−1|k−1FkT+GkQkGkT(15)

The update act is given by:
Lk={circumflex over (P)}k|k−1HkT[Hk{circumflex over (P)}k|k−1HkT+Rk]−1
{circumflex over (x)}k|k=F{circumflex over (x)}k|k−1+Lk[zk−Hk{circumflex over (x)}k|k−1]
{circumflex over (P)}k|k={circumflex over (P)}k|k−1−{circumflex over (P)}k|k−1HkT[Hk{circumflex over (P)}k|k−1HkT+Rk]−1HkT{circumflex over (P)}k|k−1T(16)

The 9×1 discrete time translational motion state vector,x(k) comprises 3-D spatial components including attitude, θ(k), and angle rate, ω(k). An overbar is used to distinguish the rotational variables from the translational variables:
x(k)=[θT(k),ωT(k)]T(17)

The transition and measurement equations have a similar form as the translational equations except the transition equation is nonlinear.
xk+1=Fk(yk)+Gkwk
zk+1=Hkyk+vk(18)

The transition noise,wkis i.i.d. zero mean Gaussian with covarianceRkand the measurement noisevkis i.i.d. Gaussian with covarianceQk. The nonlinear property is shown in the expanded form.

The nonlinear matrix B(θk) converts angle rate to Euler angle rate and is given by:

The EKF method is similar to the KF method except the transition equation is linearized using the Jacobian.

The two derivatives are:

From this point on, the EKF method is similar to the KV method using the linearized components.

The projection step for the EKF is given by:
{circumflex over (x)}k|k+1=Fk{circumflex over (x)}k−1|k−1
{circumflex over (P)}k|k+1=Fk{circumflex over (P)}k−1|k−1FkT+GkQkGkT(22)

The update step for the EKF is given by:
Lk={circumflex over (P)}k|k−1HkT[Hk{circumflex over (P)}k|k−1HkT+Rk]−1
{circumflex over (x)}k|k=F{circumflex over (x)}k|k−1+Lk[zk−Hk{circumflex over (x)}k|k−1]
{circumflex over (P)}k|k={circumflex over (P)}k|k−1−{circumflex over (P)}k|k−1HkT[Hk{circumflex over (P)}k|k−1HkT+Rk]−1HkT{circumflex over (P)}k|k−1T(23)

FIGS. 21, 22, and 23illustrate sets of simulation results of the EKF fusion algorithm using the EMM900radar, JAPRAE1710and IMU1714(FIG. 17) inputs for a typical KV400flight profile.FIG. 21illustrates position results of the simulation. A top plot illustrates true position values, a middle plot illustrates estimated position values, and a bottom plot illustrates position error. An x position is indicated by solid lines or blue lines, a y position is indicated by dotted lines or green lines, and a z position is indicated by dashed lines or red lines.

FIG. 22illustrates velocity results of the simulation. A top plot illustrates true velocity values, a middle plot illustrates estimated velocity values, and a bottom plot illustrates velocity error. An x velocity is indicated by solid lines or blue lines, a y velocity is indicated by dotted lines or green lines, and a z velocity is indicated by dashed lines or red lines.

FIG. 23illustrates attitude results of the simulation. A top plot illustrates true attitude values, a middle plot illustrates estimated attitude values, and a bottom plot illustrates attitude error. An x attitude is indicated by solid lines or blue lines, a y attitude is indicated by dotted lines or green lines, and a z attitude is indicated by dashed lines or red lines.

FIG. 24illustrates an attitude control loop2400. The attitude control loop2400utilizes an uplink and onboard EKF outputs to compute firing commands to the plurality of Attitude Control Thrusters (ACT)1604to maintain the KV400body axis1520aligned with the velocity vector V (FIG. 15).

InFIG. 24, {tilde over (d)}KVand {tilde over (v)}KVare projectile position and velocity vectors in the North-East-Down (NED) frame. [{tilde over (φ)}KV, {tilde over (θ)}KV, {tilde over (ψ)}KV] are body Euler angles and {tilde over (ω)}KVis body rate vector. The projectile mass, CG location and inertia tensor are updated whenever there is a divert thruster610or an attitude thruster1604firing. The attitude control loop2400may include an Attitude Control Algorithm2402. The Attitude Control Algorithm2402may compute how much attitude control thrust is needed at the attitude thruster station620and which direction the thrust centroid should point to. If the Attitude Control Thrust (ACT) exceeds a certain fraction of the ACT force, then the algorithm may initiate a command to fire an ACT. The Attitude Thruster Select function then searches for an ACT to fire at operational block2404. If the selected ACT is at the desired ignition position, the Attitude Thruster Fire Control electronics squib the selected ACT to fire at operational block2406. Otherwise, the process exits the control loop. Finally, the Attitude Thruster Units Remaining function2410keeps track of which ACTs are spent and which are available for use. This information is fed back to the Mass Property Estimator function2412and the Attitude Thruster Select function2404.

Operation may be as explained in acts 1 through 12 below:

Act 1: The onboard EKF algorithm tracks the projectile yaw ({tilde over (ψ)}KV), pitch ({tilde over (θ)}KV) and roll ({tilde over (ω)}KV) attitude angles. These angles are used to compute the NED to body directional cosine matrix Cnbas

Act 2: The projectile velocity in the NED frame, {tilde over (V)}KV, is tracked by EMM900radar and uplinked to the projectile via RF communication links. The velocity is then transformed into the body frame.
{tilde over (V)}KVb=Cnb·{tilde over (V)}KV(25)

Act 3: The total angle of attack αtotaland aerodynamic roll angles φaerois then computed. The total angle of attack should be small at all times. The aerodynamic angle tells where the total angle of attack is relative to the body roll axis.

Act 4: The attitude changing rate cannot be measured for the purpose of rate feedback, but the synthetic rates may be derived by using the relationship between the attitude rate and the estimated body rates and attitude angles as
{dot over (ψ)}KV=({tilde over (ω)}KV(2)sin {tilde over (φ)}KV+{tilde over (ω)}KV(3)cos {tilde over (φ)}KV)sec{tilde over (θ)}KV
{dot over (θ)}KV={tilde over (ω)}KV(2)cos {tilde over (φ)}KV−{tilde over (ω)}KV(3)sin {tilde over (φ)}KV
{dot over (φ)}KV={tilde over (ω)}KV(1)+{dot over (ψ)}KVsin {tilde over (θ)}KV(27)

Act 5: Compute the KV400heading and flight path angles in the NED frame. These are the desired body yaw and pitch attitude angles that we want to control. Note that the body roll angle is free.

Act 6: Compute attitude error by subtracting the estimated attitude angles from the commanded attitude angles above. This error is then multiplied by a proportional gain Kp,outerto form attitude angle rate command.
{dot over (ψ)}cmd=Kp,outer(ψcmd−{tilde over (ψ)}KV)
{dot over (θ)}cmd=Kp,outer(θcmd−{tilde over (θ)}KV)  (29)

Act 7: Compute attitude rate errors by subtracting the attitude rate derived in Act 4 above from the rate commands above. These rate errors are multiplied by a proportional gain Kp,innerto form attitude acceleration commands.
{umlaut over (ψ)}cmd=Kp,inner({dot over (ψ)}cmd−{dot over (ψ)}KV)
{umlaut over (θ)}cmd=Kp,inner({dot over (θ)}cmd−{dot over (θ)}KV)  (9)

FIG. 25illustrates a controller input-output topology. The yaw and pitch commands are derived from Act 5. The estimated body attitude angles, Euler, and body rates, ω, are provided by the JAPRAE algorithm1710. The controller gain Kpouterand Kpinnerare to be optimized for various engagement scenarios. The symbol Eulerdrepresents the Euler rate vector [{dot over (ψ)}KV, {dot over (θ)}KV, {dot over (φ)}KV]. Eulerddrepresents the Euler acceleration vector [{umlaut over (ψ)}KV, {umlaut over (θ)}KV, {umlaut over (φ)}KV]. And, Fcmdis the output of the controller containing the magnitude of the attitude thrust force command and its orientation relative to the body roll axis1520. The function that computes the force command is described in Acts 8 through 10.

Act 8: Because the air vehicle is aerodynamically unstable at all flight regimes, angle of attack will grow if unchecked. For this reason, the aerodynamic moment to be countered must be estimated, in addition to the moment needed to produce the attitude acceleration of Act 7. The estimation is done by storing a complete set of aerodynamic coefficient tables on board the projectile and computing the total yawing and pitching moments about the estimated CG location on the flight. The algorithm to estimate the KV400mass, CG location and inertia tensor will be described below with respect toFIG. 27. The total aerodynamic moment in the body frame, ignoring the rolling moment, may be computed as:

Q is dynamic pressure

S is reference area

d is reference diameter

Cmis total pitching moment coefficient

XCPis normalized center of pressure location along the body x-axis

XCGis normalized center of gravity location along the body x-axis

YCGis CG offset from the body x-axis

ZCGis CG offset from the body x-axis

CNis normal force coefficient

Cyis lateral force coefficient

qis normalized pitching rate

Act 9: The instantaneous moment about the CG location required to produce the attitude acceleration which will force the body to align with the velocity vector V (FIG. 15) can now be calculated as:

Iadenotes the estimated moment of inertia about the roll axis.

Act 10: Given the estimated CG location, XCG, and the attitude nozzle location, XAT, we can convert the above moment command to thrust force command as:

The magnitude of this command is:
Ftotal=∥Fcmd∥  (35)

And, the roll orientation relative to the body y-axis is:

Act 11: If the thrust force command, Ftotal, is greater than a threshold value, usually a fraction of the attitude thruster force, a thruster search algorithm may be initiated to find a thruster to fire. Otherwise, no attitude thruster may be fired.

FIG. 26illustrates an example of a search procedure that may be used to find a thruster to fire. A plurality of attitude thrusters1,2,3,4,5,6,7, and8are shown with circles. A white circle, such as attitude thrusters1,4,6, and8, indicates the attitude thruster has been spent. A shaded circle, such as attitude thrusters2,3,5, and7, indicates the attitude thruster has not been spent.

From Act 3, a total angle of attack is found between the body x-axis and the velocity vector Vband the aerodynamic angle relative to the body z-axis. To reduce the total angle of attack, an attitude thruster may be fired 180 degrees opposite to φaerosuch that the resultant force F pushes the body nose toward the velocity vector. Knowing the duration of the attitude thruster burn and time delays from the thruster command generation to actual thruster ignition, may enable calculation of the proper roll angle to issue the attitude thruster squib command. If there happens to be an attitude thruster at the right place at the moment, that thruster will be selected to fire. For example, inFIG. 26, attitude thruster number8is about to reach the ignition command position; but it has already been spent, so it can't be selected. Thruster number7is available but it's too soon to be fired so it will not be selected, either. However, as the KV400precesses and nutates, the ignition command angle φignit,cmdand thruster number7may line up and be selected to fire.

Act 12: The attitude control loop may keep track of which attitude thrusters are available to use in an availability list. When a certain attitude thruster is commanded to fire, it is removed from the availability list.

Because of the high mass ratio of the divert thruster610and the attitude thruster1604propellants to the KV400total mass, the CG location and the inertia tensor will migrate as any divert thruster610or an attitude thruster1604is firing. This is the so-called CG migration problem and it becomes more important toward the end of flight when most of the thrusters are spent.

FIG. 27illustrates a mass and CG location update algorithm2700. The mass and CG location update algorithm2700may be configured to update the total mass, CG location and inertia tensor as a divert thruster610or attitude thruster1604is firing. The definitions of the symbols are:
{dot over (m)}i≡iththruster mass burn rate<0
ri≡iththrusterCGlocation
m≡total mass
rCG≡Cglocation
Ixy≡Inertia tensor elementxy(37)

The following equations are computed recursively:

To a large extent, this detailed description has focused on a particular type of intercept vehicle (e.g., the eject vehicle400). However, engagement management systems described herein may be used with many types of intercept vehicles400in which the engagement management system can track the intercept vehicle400, alter the course of the intercept vehicle400, determine when to detonate the intercept vehicle400, or combinations thereof using commands communicated between the engagement management system and the intercept vehicle400.

Moreover, while embodiments of the present disclosure may be particularly suitable for use on aerial platforms, they may also be used in other types of mobile platforms like ground-based mobile platforms such as, for example, tanks, armored personnel carriers, personnel carriers (e.g., Humvee and Stryker vehicles) and other mobile platforms capable of bearing embodiments of the present disclosure. Moreover, embodiments of the present disclosure may be used for relatively stationary ground-based personnel protection wherein a mobile platform may not be involved. Accordingly, embodiments of the disclosure are not limited to aerial applications.

While the present disclosure has been described herein with respect to certain illustrated embodiments, those of ordinary skill in the art will recognize and appreciate that the present invention is not so limited. Rather, many additions, deletions, and modifications to the illustrated and described embodiments may be made without departing from the scope of the invention as hereinafter claimed along with their legal equivalents. In addition, features from one embodiment may be combined with features of another embodiment while still being encompassed within the scope of the invention as contemplated by the inventor.