Methods and apparatuses for aerial interception of aerial threats

Abstract

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.

Description

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.

BRIEF DESCRIPTION OF THE DRAWINGS

FIGS. 1A and 1B illustrate a helicopter as an aerial platform that may be under attack from an aerial threat and coverage areas that may be employed to sense when such a threat is present;

FIGS. 2A and 2B illustrate a conventional dispenser in which an eject vehicle (EV) according to one or more embodiments of the present disclosure may be placed;

FIG. 3 illustrates systems that may be present on a helicopter and that may intercommunicate according to one or more embodiments of the present disclosure;

FIG. 4 illustrates an exploded view of an eject vehicle showing various elements of the EV according to one or more embodiments of the present disclosure;

FIGS. 5A-5C illustrate the eject vehicle of FIG. 4 as it may be configured during various stages of an intercept mission according to one or more embodiments of the present disclosure;

FIGS. 6A-6C illustrate various propulsion and thruster elements that may be included with one or more embodiments of the present disclosure;

FIG. 7 illustrates various electrical and communication connections that may be present on an EV while it is disposed on the mobile platform prior to launch;

FIG. 8 is a block diagram illustrating elements that may be present on the eject vehicle according to one or more embodiments of the present disclosure;

FIG. 9A is a block diagram illustrating elements that may be present on the aerial platform according to one or more embodiments of the present disclosure;

FIG. 9B is a perspective view of a radar module that may be present on the aerial platform according to one or more embodiments of the present disclosure;

FIGS. 10A and 10B are diagrams illustrating radar scanning beams during an acquisition mode and a tracking mode, respectively;

FIG. 11 is a spectrum diagram illustrating possible Doppler spectrum regions where various aerial vehicles may be detected;

FIG. 12 is a simplified flow diagram illustrating some of the processes involved in one or more embodiments of the present disclosure;

FIG. 13 illustrates an example flight path for the eject vehicle and an aerial threat during an intercept process;

FIG. 14 illustrates two aerial vehicles flying in a formation and various radar sectors that may be covered by the aerial vehicles;

FIG. 15 is a simplified side view of a kill vehicle illustrating the principle forces of interest on the kill vehicle;

FIGS. 16A and 16B illustrate a nose thruster module;

FIG. 17 is a simplified block diagram of an attitude control loop;

FIG. 18 is a simplified flowchart of a joint adaptive polarization and roll angle estimator (JAPRAE) algorithm;

FIG. 19 illustrates the JAPRAE algorithm in the form of a conventional servo loop;

FIG. 20 illustrates a non-limiting example of a set of simulation results of the JAPRAE algorithm of FIGS. 18 and 19;

FIG. 21 illustrates a set of position results of an extended Kalman filter (EKF) fusion algorithm simulation;

FIG. 22 illustrates a set of velocity results of the EKF fusion algorithm simulation of FIG. 21;

FIG. 23 illustrates a set of attitude results of the EKF fusion algorithm simulation of FIG. 21;

FIG. 24 illustrates an attitude control loop;

FIG. 25 illustrates a controller input-output topology;

FIG. 26 illustrates an example of a search procedure that may be used to find a thruster to fire; and

FIG. 27 illustrates a mass and center of gravity (CG) location update algorithm.

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 in FIG. 3 are designated with a format of 3xx, where 3 indicates FIG. 3 and xx designates the unique element.

It should be understood that any reference to an element herein using a designation such as “first,” “second,” and so forth does not limit the quantity or order of those elements, unless such limitation is explicitly stated. Rather, these designations may be used herein as a convenient method of distinguishing between two or more elements or instances of an element. Thus, a reference to first and second elements does not mean that only two elements may be employed or that the first element must precede the second element in some manner. In addition, unless stated otherwise, a set of elements may comprise one or more elements.

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 1B illustrate a helicopter as an aerial platform 100 that may be under attack from an aerial threat 120 and coverage areas 140 that 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 in FIG. 1A, the aerial threat 120 may be shot by an attacker 110 toward the aerial platform 100.

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 platforms 100 due 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. 1B illustrates coverage areas 140 in which one or more embodiments of the present disclosure may detect an incoming aerial threat 120 and perform active countermeasures using one or more embodiments of the present invention to remove the aerial threat 120 before it can damage the aerial platform 100. 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 2B illustrate a dispenser 200 configured as a conventional CMDS (e.g., an AN/ALE-47) in which an eject vehicle 400 (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 vehicles 400 may also be placed in the AN/ALE-47 and ejected therefrom under control of the AN/ALE-47 and other electronics on the aerial platform 100 (FIGS. 1A and 1B). The eject vehicle 400 may be configured as a substantially cylindrical vehicle to be placed in a tubular dispenser 210 and ejection may be controlled from control wiring 220 connected to the dispenser 200. Moreover, the dispenser 200 may be configured to hold both the passive countermeasures for which it was originally designed, as well as one or more eject vehicles 400 according to embodiments of the present disclosure.

While some embodiments of the eject vehicle 400 may be configured to be disposed in an AN/ALE-47, other types of dispensers 200 or other types of carriers for the eject vehicle 400 may also be used. Moreover, the tubular dispenser 210 is illustrated with a circular cross section. However, other cross sections may be used, such as, for example, square, hexagonal, or octagonal.

FIG. 3 illustrates systems that may be present on a helicopter frame 300 and that may intercommunicate according to one or more embodiments of the present disclosure. The helicopter frame 300 and systems described are used as specific examples to assist in giving details about embodiments of the present disclosure. In the specific example of FIG. 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, and 320D) are disposed near four corners of the helicopter frame 300. A central processor 360 may be used to control and coordinate the four MAWSs (320A, 320B, 320C, and 320D).

Two AN/ALE-47 dispensers (200A and 200B) are positioned on outboard sides of the helicopter frame 300, each of which may contain one or more eject vehicles 400. As shown in FIG. 3, there are four eject vehicles 400 on each side labeled EV1 through EV4 on one side and labeled EV5-EV8 on the other side. The AN/ALE-47 dispensers (200A and 200B) are each controlled by an AN/ALE-47 sequencer (350A and 350B), which are, in turn, controlled by the central processor 360.

According to one or more embodiments of the present disclosure four radar modules (900A, 900B, 900C, and 900D) are included to augment and connect with the AAR-47s and communicate with the eject vehicles 400. These radar modules 900 (see FIG. 9A) are configured to detect and track relatively small incoming aerial threats (e.g., an RPG) as well as the outgoing eject vehicles 400. Moreover, the radar modules 900 can send wireless communications (340A, 340B, 340C, and 340D) to the eject vehicles 400 both before and after they are ejected from the dispensers (200A and 200B). The radar modules 900, and eject vehicles 400 may each include unique identifiers, such as, for example, a media access control (MAC) address. The radar modules 900 may 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 vehicle 400 into a sector that may damage or interfere with another helicopter.

The control processors, such as the central processor 360, the MAWSs 320, the radar modules 900, the sequencers 350, and the dispensers 200 may be configured to form an ad hoc network and include the eject vehicles 400.

The specific example of FIG. 3 is 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 radar module 900A may be positioned on one side of the helicopter frame 300 and another radar module 900C may be positioned on another side of the helicopter frame. In such a case, the radar modules 900 would be configured to provide hemispherical coverage areas. These radar modules 900 may be controlled by, communicate with, or a combination thereof, a different central processor 360 configured specifically for embodiments of the present disclosure. Moreover, the eject vehicles 400 may be disposed in different carriers or different dispensers from the AN/ALE-47 dispensers (200A and 200B) shown in FIG. 3.

When embodiments of the present disclosure are used as illustrated in FIG. 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 inches×5.5 inches 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 attacker 110 (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 threats 120 can be detected and tracked and multiple outgoing eject vehicles 400 can be tracked. In addition, to increase a probability of destroying an incoming aerial threat 120, multiple eject vehicles 400 may be launched, directed toward, and detonated proximate the same aerial threat 120.
  • 6) Finally, eject vehicles 400 can 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 vehicle 400 of FIG. 2B), including an ejection mechanism with an impulse charge that can fit in, and can be launched by, the AN/ALE-47 chaff/flare dispenser 200. Some embodiments of the present disclosure include the radar module 900 covering a 90-degree sector or more (i.e., with a 90-degree sector, each helicopter platform would use four radar modules 900).

When referring to the radar module 900 herein (e.g., as shown in FIG. 3), it should be understood that in some embodiments the radar module 900 may perform the operations described herein in combination with other electronics and processors on the aerial platform 100. As such, the radar modules 900 may be used to: 1) search, acquire, and track incoming aerial threats 120, 2) launch the active kinetic countermeasure (i.e., eject vehicle 400), 3) track the outgoing eject vehicle 400 with respect to the incoming aerial threat 120, 4) point and guide the eject vehicle 400 toward the incoming aerial threat 120, 5) command detonate the eject vehicle 400, and 6) geo-locate the attacker 110, all in less than about one second. In one configuration, at least two AN/ALE-47 dispensers 200 would be used in conjunction with the four radar modules 900 such that each dispenser 200 provides hemispherical coverage.

The radar modules 900 may be configured as pulse Doppler radar modules 900 to 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 threat 120, the associated radar module 900 may then direct the launch and guidance of an eject vehicle 400 from an AN/ALE-47 dispenser 200 that covers that sector. The eject vehicle 400 may be command guided to the target by the radar module 900 and command detonated. The radar modules 900 may be configured as an addition to the existing AN/AAR-47 system and may use its existing interface for launching of the eject vehicle 400.

Some of the embodiments of the present disclosure may be configured to deploy an eject vehicle 400 that fits in a standard dispenser 200 but 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 dispenser 200. The radar modules 900 may then guide the eject vehicle 400 to accurately intercept the aerial threat 120 within about 330 milliseconds and thus reduce the requirement of carrying a large warhead.

FIG. 4 illustrates an exploded view of an eject vehicle 400 showing various elements of the eject vehicle 400 according to one or more embodiments of the present disclosure. Reference may also be made to FIGS. 1A-3 in describing features and operations of the eject vehicle 400. The eject vehicle 400 is a lightweight guided projectile that, in some embodiments, may be designed to be launched from chaff/flare dispensers. The eject vehicle 400 may intercept and destroy incoming aerial threats 120 at ranges sufficient to prevent damage to the host aerial platform 100. The eject vehicle 400 may be packaged in a cartridge containing an impulse charge and interface electronics designed to fit the AN/ALE-47 dispenser magazine.

The eject vehicle 400 includes an ejection piston 780 configured to transmit the energy of an impulse cartridge 750 (described below in connection with FIG. 7) to the eject vehicle 400 and launch the eject vehicle 400 away from the aerial platform 100 to a distance safe enough for the eject vehicle 400 to begin performing alignment and interception maneuvers.

A rocket motor 420 may be used to propel the eject vehicle 400 toward the aerial threat 120 after the eject vehicle 400 has been rotated such that a longitudinal axis of the eject vehicle 400 is pointed in the general direction of the aerial threat 120. A first set of folding fins 482 may be attached to the rocket motor 420 and configured to deploy once the eject vehicle 400 has exited the dispenser 200. The folding fins 482 are small and configured to provide stability to the eject vehicle 400 during its flight path rather than as control surfaces for directing the fight path.

An airframe shell 430 may be configured to contain a warhead 440, a divert thruster module 610, a nose thruster module 620 (may also be referred to herein as an alignment thruster module 620), an electronics module 450, and a battery 452. An airframe nose 490 may be configured to attach to the airframe shell 430 to protect the electronics module 450 and provide a somewhat aerodynamic nose for the eject vehicle 400.

A safe and arm module 460 may be included within the airframe shell 430 and configured to safely arm the warhead 440 when the eject vehicle 400 is a safe distance away from the aerial platform 100.

FIGS. 5A-5C illustrate the eject vehicle 400 of FIG. 4 as it may be configured during various stages of an intercept mission according to one or more embodiments of the present disclosure. Stage 1, in FIG. 5A, illustrates the eject vehicle 400 in a cartridge 710 (FIG. 7) and includes the ejection piston 780, the rocket motor 420, the airframe shell 430, and the airframe nose 490.

Stage 2, in FIG. 5B, illustrates the eject vehicle 400 after it has been dispensed and shows the rocket motor 420, the airframe shell 430, and the airframe nose 490. FIG. 5B also illustrates the folding fins 482 deployed near the end of the rocket motor 420 and wireless communication antennas 890 deployed near the airframe nose 490.

Stage 3, in FIG. 5C illustrates the eject vehicle 400 after the rocket motor 420 has burned and been detached from the airframe shell 430. At this stage, the eject vehicle 400 may be referred to as a terminal vehicle and includes the airframe nose 490, the wireless communication antennas 890, and the airframe shell 430. Still within the airframe shell 430 are the warhead 440, the divert thruster module 610, the alignment thruster module 620, the electronics module 450, the battery 452, and the safe and arm module 460. After the rocket motor 420 is detached, a second set of folding fins 484 are deployed from the airframe shell 430 to stabilize the eject vehicle 400 during the remainder of the flight to intercept the aerial threat 120. This second set of folding fins 484 are used to replace the first set of folding fins 482 that were attached to the rocket motor 420, which has been detached from the airframe shell 430 during stage 3.

In addition, after the rocket motor 420 is detached, one or more corner reflectors 470 are exposed. The corner reflector 470 may be configured with sharp angles to enhance radar detection of the eject vehicle 400 by a radar module 900 on the aerial platform 100. For example, the corner reflector 470 may be configured as an interior angle of a small cube shape, which will enhance radar detection.

Returning to FIG. 4, the alignment thruster module 620 is offset from a center of mass of the eject vehicle 400 such that an initial pitch maneuver can be performed to align the longitudinal axis of the eject vehicle 400 along an intercept vector pointed toward the aerial threat 120. This alignment maneuver is performed prior to the burn of the rocket motor 420.

The divert thruster module 610 is 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 threat 120. The terminal vehicle may be referred to herein as the eject vehicle 400 and it should be understood what is being referred to based on the context of the discussion.

The warhead 440 may be command detonated when the radar module 900 on the aerial platform 100 determines that the eject vehicle 400 has 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 threat 120 at a nominal distance of about 50 meters when the aerial threat 120 is launched from a range of about 100 meters.

FIGS. 6A-6C illustrate various propulsion and thruster elements that may be included with one or more embodiments of the present disclosure. FIG. 6A illustrates a nose thruster module 620 with four nose thrusters 622 (two are hidden) arranged around a periphery of the nose thruster module 620. These nose thrusters 622 (also referred to herein as alignment thrusters 622) are positioned to generate a perpendicular force on the eject vehicle 400 relative to the longitudinal axis and are offset from the center of mass of the eject vehicle 400 so that an initial pitch maneuver can be performed to rotate and align the longitudinal axis of the eject vehicle 400 along an intercept vector pointed toward the aerial threat 120. In this embodiment, the four nose thrusters 622 are orthogonally arranged giving two opportunities to adjust the pitch of the eject vehicle 400 in each direction. Of course, other embodiments may include fewer or more alignment thrusters 622.

FIG. 6B illustrates a divert thruster module 610 with eight divert thrusters 612 (five are hidden) arranged around a periphery of the divert thruster module 610. These divert thrusters 612 are positioned to generate a perpendicular force on the eject vehicle 400 relative to the longitudinal axis and are positioned near the center of mass of the eject vehicle 400 so that the divert thrusters 612 will move the eject vehicle 400 laterally to a slightly different travel path, while substantially maintaining the same pitch. Thus, the divert thrusters 612 can modify the flight path of the eject vehicle 400 to correct for minor errors in the initial pitch maneuvers pointing directly toward the aerial threat 120. In this embodiment, eight divert thrusters 612 are used giving eight opportunities to adjust the flight path of the eject vehicle 400 during its flight toward the aerial threat 120. Of course, other embodiments may include fewer or more divert thrusters 612.

FIG. 6C illustrates a thruster 650 configured to expel a gas through a nozzle 652 to create a lateral force. The thruster 650 may be controlled from a thrust signal 654, which may be connected to the electronics module 450 of the eject vehicle 400. The thruster 650 is one example of a type of thruster that may be used for both the divert thrusters 612 and the alignment thrusters 622.

FIG. 7 illustrates various electrical and communication connections that may be present on the eject vehicle 400 while it is disposed on the aerial platform 100 (FIGS. 1A and 2B) prior to launch. A cartridge 710 includes a cartridge flange 720 such that the cartridge 710 may be securely placed in a dispenser 200 (FIG. 2A). An end cap 790 may be positioned over the cartridge 710 to hold the eject vehicle 400 within the cartridge 710. An impulse cartridge 750 is positioned near the base of the cartridge flange 720 and is configured to fire in response to a fire command signal 755 from the radar module 900 (FIG. 3) or other electronics on the aerial platform 100. An ejection piston 780 is positioned between the impulse cartridge 750 and the eject vehicle 400 and is configured to transmit the energy of the firing impulse cartridge 750 to the eject vehicle 400 and propel the eject vehicle 400 out of the dispenser 200 and safely away from the aerial platform 100.

A power signal 740 and a ground signal 730 may run along or through the cartridge to an antenna spring contact 745 and a ground spring contact 735, respectively. The ground spring contact 735 is configured to flexibly couple with a ground patch 738 on the eject vehicle 400 to provide a ground for the eject vehicle 400 electronics while the eject vehicle 400 is in the cartridge 710. The antenna spring contact 745 is configured to flexibly couple with the antenna 890 on the eject vehicle 400 and a power signal on the eject vehicle 400 to provide power and direct communication for the eject vehicle 400 electronics while the eject vehicle 400 is in the cartridge 710. The cartridge 710 may include a cartridge antenna 760 that may be coupled to the antenna 890 of the eject vehicle 400 by the antenna spring contact 745. Thus, the eject vehicle 400 may communicate wirelessly 795 with electronics onboard the aerial platform 100 through the antenna 890 on the eject vehicle 400 or through the cartridge antenna 760.

FIG. 8 is a block diagram illustrating elements that may be present on the eject vehicle 400 according to one or more embodiments of the present disclosure. A microcontroller 810 may be coupled to a memory 820, which is configured to hold instructions for execution by the microcontroller 810 and data related to command and control of the eject vehicle 400. The microcontroller 810 may 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 vehicle 400. The memory 820 may be any suitable combination of volatile and non-volatile memory configured to hold data and computing instructions related to operation of the eject vehicle 400.

One or more antennas 890 may be configured to provide a communication link with electronics (e.g., the radar module 900) onboard the aerial platform 100. As non-limiting examples, the communication link may be configured using Wi-Fi or WiMAX frequencies and protocols. A diversity combiner 880 may be used to combine signals from multiple antennas.

A communication transceiver 870 (e.g., a Wi-Fi transceiver) may be coupled to the diversity combiner 880 and be configured to transmit and receive frequencies to and from the diversity combiner 880. A communication modem 860 (e.g., a Wi-Fi modem) may be coupled to the communication transceiver 870 and be configured to package and modulate communication information for communication transmission as well as demodulate and extract information from communication reception. The microcontroller 810 receives information from the communication modem 860 and may perform operations related to the received information. In addition, based on processes performed on the microcontroller 810, information may be sent to the communication modem 860 for transmission through the one or more antennas 890.

The microcontroller 810 may be coupled to a thrust controller 830, which interfaces with the alignment thrusters 622 and the divert thrusters 612 (FIGS. 6A and 6B). A warhead fusing interface 840 may be provided to interface to the warhead 440 (FIG. 4), the safe and arm module 460 (FIG. 4) or a combination thereof, for arming and control of detonation of the warhead 440.

A roll sensor 850 and a vertical reference 855 may be used in combination to determine the attitude of the eject vehicle 400 as well as a spin rate and spin position of the eject vehicle 400 and communicate such information to the microcontroller 810. Other types of sensors, such as, for example, accelerometers and magnetometers may also be used for this purpose.

FIG. 9A is a block diagram illustrating elements that may be present on the aerial platform 100 according to one or more embodiments of the present disclosure. The electronics module and functions thereof on the aerial platform 100 may be contained within a radar module 900, as illustrated in FIG. 9B. Alternatively, some of the function may be within the radar module 900 while other functions may be located in different places on the aerial platform 100 such as, for example, the central processor 360 (FIG. 3). The various modules used to control the radar module 900 and the eject vehicle 400 and determine other information related thereto may be collectively referred to herein as an “onboard system.”

FIG. 9B is a perspective view of the radar module 900 that may be present on the aerial platform 100 according to one or more embodiments of the present disclosure. The radar module 900 includes an azimuth scan radar antenna 920, an elevation scan radar antenna 940, and a wireless communication link antenna 960.

The azimuth scan radar antenna 920 is included in an azimuth radar subsystem, which includes a diplexer 922 for combining radar sent and reflected radar received. A radio frequency (RF) up/down converter 925 converts the radar frequencies sent from a digital synthesizer 930 and converts the radar frequencies received for use by a digital receiver 935.

The elevation scan radar antenna 940 is included in an elevation radar subsystem similar to the azimuth radar subsystem, but configured for the elevation direction. The elevation radar subsystem includes a diplexer 942 for combining radar sent and reflected radar received. A radio frequency (RF) up/down converter 945 converts the radar frequencies sent from a digital synthesizer 950 and converts the radar frequencies received for use by a digital receiver 955.

The wireless communication link antenna 960 may be configured to provide a communication link with electronics onboard the eject vehicle 400. As non-limiting examples, the communication link may be configured using Wi-Fi or WiMAX frequencies and protocols. A wireless communication subsystem includes a communication transceiver 965 (e.g., a Wi-Fi transceiver) coupled to the wireless communication link antenna 960 and configured to transmit and receive frequencies to and from the antenna 960. A communication modem 970 (e.g., a Wi-Fi modem) may be coupled to the communication transceiver 965 and be configured to package and modulate communication information for communication transmission as well as demodulate and extract information from communication reception.

A sector processor 910 communicates with the elevation radar subsystem, the azimuth radar subsystem, and the wireless communication subsystem. The sector processor 910 may communicate helicopter navigation information 912 from other electronics on the aerial platform 100. Referring also to FIG. 3, the sector processor 910 may also communicate with the dispenser 200 (e.g., one or more ALE-47s) using communication signal 914 and the missile approach warning system 320 (e.g., one or more AAR-47s) using communication signal 916. The sector processor 910 performs a number of functions to detect and track aerial threats 120, control and track the eject vehicle 400, as well as other functions related to the active protection system. In some embodiments, communication between the dispenser 200 and the sector processor 910 may be accomplished through the missile approach warning system 320.

The sector processor 910 in combination with the radar subsystems can detect and track incoming aerial threats 120 (e.g., RPGs). Based on the tracking of the incoming aerial threat 120, and in combination with navigation information from the aerial platform 100, the sector processor 910 can extrapolate to a geo-location of the attacker 110, from where the aerial threat 120 was launched. The aerial platform 100 may act on this geo-location or transmit the geo-location to other aerial platforms or ground-based platforms for follow-up actions.

The sector processor 910 may be configured to send launch commands to the dispenser 200 on communication signal 914 to launch one or more eject vehicles 400 to intercept one or more detected aerial threats 120. The sector processor 910 may also calculate required pitch adjustments that should be performed by the eject vehicle 400 after it has been ejected and is safely away from the aerial platform 100.

Once the eject vehicle 400 is launched, the sector processor 910 may be configured to track the eject vehicle 400 and send guidance commands (i.e., divert commands) to the eject vehicle 400 so the eject vehicle 400 can perform divert maneuvers to adjust its flight path toward the aerial threat 120. The sector processor 910 may also be configured to determine when the eject vehicle 400 will be near enough to the aerial threat 120 to destroy the aerial threat 120 by detonation of the warhead 440 on the eject vehicle 400. Thus, a detonation command may be sent to the eject vehicle 400 instructing it to detonate, or instructing it to detonate at a detonation time after receiving the command.

Referring to FIGS. 10A, 10B, 9, and 3, the radar modules 900 may 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 threat 120 will be launched at relatively short ranges, typically on the order of 100 m. The radar modules 900 are 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 vehicle 400 to intercept the aerial threat 120. 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. 10A illustrates 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 threat 120 when it is within range.

FIG. 10B illustrates a track mode. In FIG. 10B, two sequential azimuth scans and two sequential elevation scans are shown that pinpoint a first location 1010 of the eject vehicle 400. In addition, two sequential azimuth scans and two sequential elevation scans are shown that pinpoint a second location 1020 of the aerial threat 120. With this location information, the sector processor 910 can derive relative position information that can be used to provide divert commands to the eject vehicle 400 to more closely intercept the aerial threat 120.

FIG. 11 is a spectrum diagram illustrating possible Doppler spectrum regions where various aerial vehicles may be detected. As non-limiting examples, FIG. 11 illustrates a ground clutter spectrum 1110, a spectrum 1120 for the eject vehicle 400 (i.e., PRJ in FIG. 11), a spectrum 1130 that may be indicative of an RPG, and a spectrum 1140 that may be indicative of a MANPAD. Of course, other aerial threats and their associated spectrums may also be identified.

FIG. 12 is a simplified flow diagram illustrating some of the processes 1200 involved in one or more embodiments of the present disclosure. The processes may be loosely considered as an acquisition phase 1210, a pre-launch phase 1220, an align and launch phase 1240, a guidance phase 1260, a divert phase 1270, and a detonation phase 1280.

Operation block 1212 indicates that continuous radar scans are performed looking for incoming aerial threats. Decision block 1214 indicates that the process loops until a target is detected. While not shown, during this phase the radar modules 900 may 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 with FIG. 14.

If a target is detected, the process 1200 enters the pre-launch phase 1220. Operation block 1222 indicates that the sector processor 910 uses the range and travel direction of the incoming aerial threat 120 to calculate a threat direction to the incoming aerial threat 120 and an intercept vector pointing from a deployed eject vehicle 400 to a projected intercept point where the eject vehicle 400 would intercept the incoming aerial threat 120. Operation block 1224 indicates that the intercept vector is sent to the eject vehicle 400. The intercept vector may be sent to the eject vehicle 400 in a number of forms. The actual directional coordinates may be sent and the eject vehicle 400 would be responsible for determining the proper pitch maneuvers to perform. Alternatively, the sector processor 910 may determine the proper pitch maneuvers that the eject vehicle 400 should perform after launch and send only pitch commands (e.g., start and burn times for each alignment thruster 622) to be used during the pitch maneuvers. While FIG. 12 indicates 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 phase 1210 and pre-launch phase 1220, the eject vehicle 400 remains in the dispenser 200 and connected to power. An RF communication link may be in operation through the eject vehicle 400 antenna via a transmission line inside the dispenser 200.

The process enters the align and launch phase 1240 after the intercept vector is determined. Operation block 1242 indicates the impulse cartridge 750 is fired to propel the eject vehicle 400 from the dispenser 200 and safely away from the aerial platform 100.

Operation block 1244 indicates that the pitch maneuvers are performed to align the eject vehicle 400 with 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 vehicle 400 along 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 vehicle 400 using the divert thrusters 612 during the guidance phase 1260. After ejection, the folding fins 482 will deploy and the communication link antennas 960 will deploy and wireless communication between the eject vehicle 400 and the radar module 900 may commence.

Operation block 1246 indicates that the rocket motor 420 will fire, which accelerates the eject vehicle 400 to about 160 meters/second and imposes a spin rate on the eject vehicle 400 of about 10 Hertz. Upon exhaustion, the rocket motor 420 and folding fins 482 will separate and the terminal vehicle (TV) is exposed. With separation of the TV, the second set of folding fins 484 deploy and the corner reflector 470 is exposed.

During the guidance phase 1260, the process will perform a track and divert loop in order to adjust the flight path of the eject vehicle 400 to more closely intercept the aerial threat 120. Operation block 1262 indicates that the sector processor 910 will track the eject vehicle 400 and aerial threat 120 as discussed above with reference to FIGS. 9A-10B. Decision block 1264 indicates that the sector processor 910 will determine if a divert maneuver is required to intercept the incoming aerial threat 120 and estimate the direction of divert thrust required.

A divert phase 1270 includes operations to cause the eject vehicle 400 to modify its course. Operation block 1272 indicates that the divert direction and time, if required, are sent to the eject vehicle 400.

The divert process takes into account the rotation of the eject vehicle 400 and the direction of the desired divert thrust. This rotation adds a complication to the selection and fire time determination of the proper divert thruster 612, but also ensures that all of the available divert thrusters 612 can be used to divert the eject vehicle 400 in any desired direction substantially perpendicular to the travel direction of the eject vehicle 400. Operation block 1274 indicates that the processor on the eject vehicle 400 will select the divert thruster 612 to be fired and determine the firing time based on the divert angle received from the sector processor 910 and its internal attitude sensors.

Operation block 1276 indicates that the appropriate divert thruster 612 is fired at the appropriate fire time to move the eject vehicle 400 laterally along a diversion vector to adjust the flight path of the eject vehicle 400. As a non-limiting example, each divert thruster 612 may be capable of correcting for about two degrees of error from the initial pointing of the eject vehicle 400 during the pitch maneuver. Thus, when the divert thrusters 612 are fired when the eject vehicle 400 is in the correct rotational position, the process can slide the travel direction vector of the eject vehicle 400 toward the path of the aerial threat 120. Moreover, the process can fire in any circular direction and can fire multiple divert thrusters 612 in the same direction to repeatedly move the eject vehicle 400 in the same direction.

While FIG. 12 indicates the guidance phase 1260 and the detonation phase 1280 as operating sequentially, they also may operate in parallel. During the detonation phase 1280, operation block 1282 indicates that the sector processor 910 determines an optimum intercept time when the eject vehicle 400 will be at its closest point to the aerial threat 120. Operation block 1284 indicates that a detonation command may be sent to the eject vehicle 400. This detonation command may be in the form of a detonation time for the eject vehicle 400 to count out or it may be in the form of an immediate command for the eject vehicle 400 to perform as soon as the command is received.

Operation block 1286 indicates that the warhead 440 on the eject vehicle 400 is detonated at the intercept time responsive to the detonation command received from the sector processor 910.

FIG. 13 illustrates an example flight path for the eject vehicle 400 and an aerial threat 120 during 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 vehicle 400 receives its coordinate commands from the radar module 900 and is then ejected from the port chaff dispenser 200 at an angle of 90 degrees to the helicopter axis.

During period 1310, the eject vehicle 400 separates to a distance of about two meters from the helicopter. During period 1320, the nose thrusters 622 pitch the eject vehicle 400 to the approximate approach angle of the incoming RPG (e.g., within about ±10° accuracy). The rocket motor 420 then fires to accelerate the eject vehicle 400 to approximately 160 meters/second and is then separated from the remaining terminal vehicle upon exhaustion.

During period 1330, the radar module 900 transmits a series of divert commands to the eject vehicle 400, which fires the divert thrusters 612 to correct the trajectory of the eject vehicle 400 and intercept the RPG. A radar command is finally sent to the eject vehicle 400 to detonate the warhead 440 when 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 warhead 440.

FIG. 14 illustrates 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. In FIG. 14, a first helicopter 1410 is monitoring a first radar sector 1410A, a second radar sector 1410B, a third radar sector 1410C, and a fourth radar sector 1410D.

A second helicopter 1420 near the first helicopter 1410 is monitoring a fifth radar sector 1420A, a sixth radar sector 1420B, a seventh radar sector 1420C, and an eighth radar sector 1420D. If an aerial threat approaches from a direction indicated by arrow 1430 it may be detected by the third radar sector 1410C of the first helicopter 1410 and the seventh radar sector 1420C of the second helicopter 1420. If the first helicopter 1410 attempts to launch an eject vehicle, it may cause damage to the second helicopter 1420. However, using communication between the various wingman platforms, sectors of responsibility can be identified. Thus, for the direction indicated by arrow 1430, the first helicopter 1410 can determine that the third radar sector 1410C will be covered by the seventh radar sector 1420C of the second helicopter 1420. As a result, while this formation continues, the first helicopter 1410 does not respond to threats in its third radar sector 1410C.

Returning to FIG. 9B, the radar module 900 may also be referred to herein more generically as an engagement management module (EMM) 900. As discussed above, in some embodiments, an aerial platform 100 (FIGS. 1A and 1B) may be configured with four engagement management modules 900, an example of which is shown in FIG. 3.

The engagement management modules 900 may 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 to FIGS. 4 through 5C, the eject vehicle 400 may also be referred to herein as an intercept vehicle 400 and a kill vehicle (KV) 400. FIG. 5C illustrates the eject vehicle 400 after the rocket motor 420 has burned and been detached from the airframe shell 430. At this stage, the eject vehicle 400 may be referred to as a terminal vehicle, a terminal section, or a terminal section of the kill vehicle 400. Still within the terminal section are the warhead 440, the divert thruster module 610, the alignment thruster module 620, the electronics module 450 the battery 452, and the safe and arm module 460.

In operation, radar on the EMM 900 detects and tracks an aerial threat (e.g., an RPG) launched at the helicopter and launches one or more KVs 400 from the AN/ALE-47 countermeasure dispenser to intercept the incoming RPG. Following launch, the KV 400 executes a series of pitch maneuvers using nose thrusters (i.e., in the alignment thruster module 620) 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 thrusters 610. Finally, the KV 400 warhead 440 is command detonated when the KV 400 reaches 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 in FIG. 15.

FIG. 15 is a simplified side view of a kill vehicle (KV) 400 illustrating the principle forces of interest on the kill vehicle (KV) 400. Ideally, the divert thruster's 610 center of thrust, X_CT, would be aligned with the KV 400 center of gravity (CG) to minimize any moments introduced by its thrust force (F_DT). However, even with balancing during manufacturing, there will be some migration of the CG as divert thrusters 610 and alignment thrusters 622 (FIGS. 6A and 6B and also referred to herein as “pitch thrusters 622”) and “attitude control thrusters 1604” (FIG. 16A) are fired and propellant is expended. Also, when the divert thrusters 610 are fired, airflow along a body 1510 of the KV 400 is disrupted, known as the jet interaction (JI) effect, which introduces a high pressure region 1506 in front of the divert thrusters 610 and a low pressure region 1508 aft of the divert thrusters 610. This causes a moment that can be represented as a force F_JI at location X_JI. The other principle aerodynamic component is the normal force F_N introduced as a function of the angle of attack α, which is the angle between a body longitudinal axis 1520 and the velocity vector V. The normal force F_N acts at the center of pressure X_CP.

One or more tail fins 1524 can be used to add aerodynamic stability, but the time constants associated with these tail fins generally are not fast enough to stabilize the KV 400 during its short flight time. Instead, or in addition to the tail fins 1524, a plurality of small micro-thrusters 1604 (FIG. 16A) are added to the nose thruster module 620 (FIGS. 4 and 16A) to implement an active attitude control. This introduces F_μ at location X_μ.

FIGS. 16A and 16B illustrate a nose thruster module 620. The nose thruster module 620 may include a plurality of pitch thrusters 622 and a plurality of attitude control thrusters 1604. FIG. 16A illustrates a plurality of cylinders 1606 and 1608 configured for storing propellant grains. The larger cylinders 1606 correspond to the pitch thrusters 622 while the smaller cylinders 1608 correspond to the attitude control thrusters 1604. FIG. 16B illustrates a plurality of nozzles 1610 and 1612, each configured to divert a flow from one of the plurality of cylinders 1606, 1608. Thus, the combination of cylinders 1606 and nozzles 1610 create the pitch thrusters 622 and the combination of cylinders 1608 and nozzles 1612 create the attitude control thrusters 1604.

Referring again to FIG. 15, the various moments described earlier may cause the KV 400 to 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 thrusters 612 normal to the velocity vector V for proper guidance maneuvering and in directing the warhead 440 blast toward the RPG at the point of detonation.

FIG. 17 is a simplified block diagram of an attitude control loop 1700. The EMM 900 radar generates position {circumflex over (d)}_KV and velocity {circumflex over (ν)}_KV track data for the KV 400 and sends this information to the KV 400 via a command link radio (CLR) 1706. The EMM 900 radar also transmits a roll angle {circumflex over (ϕ)}_TX via a transmit antenna 1804 (FIG. 18) to the CLR 1706 to the KV 400 over the CLR 1706. A joint adaptive polarization and roll angle estimator (JAPRAE) algorithm 1710, resident in the KV 400, processes the polarized CLR 1706 signal received by two orthogonal linear receive antennas 1802 (FIG. 18) on the KV 400 to estimate a roll angle {circumflex over (ϕ)}_KV of the KV 400. An onboard inertial management unit (IMU) 1714 measures an acceleration a_KV and an angular velocity ω_KV of the KV 400. This information is processed by an extended Kalman filter (EKF) 1720 to generate fused estimates of the position {tilde over (d)}_KV velocity {tilde over (ν)}_KV, attitude [{tilde over (ϕ)}_KV, {tilde over (θ)}_KV, {tilde over (ψ)}_KV], and angular velocity {tilde over (ω)}_KV of the KV 400.

The attitude control algorithm 1730 is designed to fire the nose mounted attitude control thrusters 1604 to 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 algorithm 1730 selects the next available attitude control thruster 1604 that comes into position due to the body spin at operational block 1732, and times the firing to coincide with the desired direction of thrust at operational block 1734. The attitude control algorithm 1730 maintains knowledge of which attitude control thrusters 1604 have been fired and which are available at operational block 1736.

The KV 400 has an onboard IMU 1714 that measures the angular rate and acceleration. The IMU 1714 may 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 IMU 1714 also includes an FPGA that contains the EKF 1720 to fuse the output of the sensors into 3-dimensional position, velocity, angular velocity and attitude.

FIG. 18 is a simplified flowchart of the joint adaptive polarization and roll angle estimator (JAPRAE) algorithm 1710. The JAPRAE algorithm 1710 may process the CLR 1706 signals received, for example, by orthogonal linearly polarized receive antennas 1802, to adapt the two receive signals in a beam former 1806 (also referred to herein as “the dual polarized receiver 1806”) to match the incoming polarization to eliminate polarization mismatch loss induced by body rotation. The JAPRAE algorithm 1710 also processes the CLR 1706 signals received to provide an estimate of the KV 400 roll angle {circumflex over (ϕ)}_rx.

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

$\begin{array}{cc}{E}_{tx}\left(t\right)=\left[u_{tx\varphi 1},u_{tx\varphi 2}\right]\left[\begin{array}{c}{p}_1\left(t\right)\\ p_2\left(t\right)\end{array}\right]E_{tx}{e}^{i{\omega }_{0}t}=U_{tx}p{E}_{tx}{e}^{i{\omega }_{0}t}& \left(1\right)\end{array}$ where u_txϕ1 and u_txϕ2 are orthogonal unit vectors that define the direction of the electric field components, p_{1hu (t)} and p₂^(t) are the complex components of a unit vector that define the projection of the electric field vector on the basis vectors, and U and p are the matrix of basis unit vectors and the polarization vector, respectively. Similarly, the polarization of an antenna can be defined by a 2×1 complex vector q. The signal received by an antenna with polarization q from a signal with polarization p is simply the inner product of p and q:x_rx(t)=q(t)^Hp(t)E(t)e^iβ(t)  (2)

If polarizations p and q are defined in different coordinate systems, U_tx and U_rx, then (2) becomes:x_rx(t)=q(t)^HU_rx^TU_txp(t)E(t)e^iβ(t)  (3)

Assume the polarization of a source antenna 1804 and the receive antennas 1802 can be represented as p_o and q_o in terms of a natural antenna coordinate system. Let A_tx(t) and A_rx (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:E_tx(t)=A_tx(t)U_txp_oE(t)e^iβ(t) q(t)=A_rx(t)U_rxq_o.  (4)Then, (3) becomes:x_rx(t)=q_o^HU_rx^TA_rx^T(t)A_tx(t)U_txp_pE(t)e^iβ(t)  (5)

In the JAPRAE system 1710 there are two orthogonal receive antennas 1802, with polarizations q_o1 and q_o2. Thus, the signal generated at the input of the dual polarized receiver 1806 is given by:

$\begin{array}{cc}\begin{array}{c}{x}_{rx}\left(t\right)=\left[\begin{array}{c}{x}_{rx1}\left(t\right)\\ x_{rx2}\left(t\right)\end{array}\right]=\left[\begin{array}{c}{q}_{o1}^H{U}_{rx}^T{A}_{rx}^T\left({\varphi }_{rx}\right)A_{tx}\left({\varphi }_{tx}\right)U_{tx}{p}_o\\ q_{o2}^H{U}_{rx}^T{A}_{rx}^T\left({\varphi }_{rx}\right)A_{tx}\left({\varphi }_{tx}\right)U_{tx}{p}_o\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}\\ =\left[\begin{array}{c}{a}_1\\ a_2\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}=aE\left(t\right)e^{i\beta \left(t\right)}\end{array}& \left(6\right)\end{array}$

In equation (6) the transformations are explicitly represented as rotations of the transmitter ϕ_tx and receiver ϕ_rx around the line of sight. In the case where the transmit antenna 1804 and the receive antennas 1802 are linearly polarized, it can be shown that this reduces to the following form:

$\begin{array}{cc}{x}_{rx}\left(t\right)=\left[\begin{array}{c}\cos \left({\varphi }_{tx}-{\varphi }_{rx}\right)\\ -\sin \left({\varphi }_{tx}-{\varphi }_{rx}\right)\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}=\left[\begin{array}{c}\cos \left(\delta \right)\\ -\sin \left(\delta \right)\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}& \left(7\right)\end{array}$

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

$\begin{array}{cc}\begin{array}{c}{y}_{rx}\left(t\right)=\left[\begin{array}{c}{y}_{rx1}\left(t\right)\\ y_{rx2}\left(t\right)\end{array}\right]=\left[\begin{array}{cc}\cos \left(\hat{\delta }\right)& -\sin \left(\hat{\delta }\right)\\ \sin \left(\hat{\delta }\right)& \cos \left(\hat{\delta }\right)\end{array}\right]\left[\begin{array}{c}\cos \left(\delta \right)\\ -\sin \left(\delta \right)\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}\\ =\left[\begin{array}{c}\cos \left(\hat{\delta }\right)\cos \left(\delta \right)+\sin \left(\hat{\delta }\right)\sin \left(\delta \right)\\ \sin \left(\hat{\delta }\right)\cos \left(\delta \right)-\cos \left(\hat{\delta }\right)\sin \left(\delta \right)\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}\\ =\left[\begin{array}{c}\cos \left(\hat{\delta }-\delta \right)\\ \sin \left(\hat{\delta }-\delta \right)\end{array}\right]E\left(t\right)e^{i\beta \left(t\right)}\end{array}& \left(8\right)\end{array}$

Note that if {circumflex over (δ)}=δ, y_rx1=Ee^iβ(t) and y_rx0=0. This amounts to a case where γ_rx1 is 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)

$\begin{array}{cc}\begin{array}{c}r\left(t\right)=\frac{y_{\mathrm{rx}2}\left(t\right)}{y_{\mathrm{rx}1}\left(t\right)}=\frac{\sin \left(\hat{\delta }-\delta \right)E\left(t\right)e^{i\beta \left(t\right)}}{\cos \left(\hat{\delta }-\delta \right)E\left(t\right)e^{i\beta \left(t\right)}}=\tan \left(\hat{\delta }-\delta \right)\\ =\tan \left(\hat{\delta }-\delta \right)\end{array}& \left(9\right)\end{array}$

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

$\begin{array}{cc}{e}_{\delta }\left(t\right)=\hat{\delta }\left(t\right)-\delta \left(t\right)={\tan }^{-1}\left(\frac{y_2\left(t\right)}{y_1\left(t\right)}\right)& \left(10\right)\end{array}$

FIG. 19 illustrates the JAPRAE algorithm 1710 (FIG. 17) in the form of a conventional servo loop 1900. The closed loop may force y_ex2 to zero, hence {circumflex over (δ)}→δ. Under closed loop conditions, the KV 400 roll angle is estimated by:{circumflex over (θ)}_rx(t)={circumflex over (θ)}_tx(t)−{circumflex over (δ)}(t)  (11)

The transmitter roll angle, {circumflex over (ϕ)}_tx can be transmitted to the KV 400 via the data link 1806.

FIG. 20 illustrates a non-limiting example of a set of simulation results 2000 of the JAPRAE algorithm 1710 (FIG. 17) tracking the roll angle of a KV 400 body spinning on its longitudinal axis 1520 (FIG. 15) at 2200 degrees per second. A top plot 2002 of FIG. 20 illustrates an overlay of true and estimated roll angles. A bottom plot 2004 illustrates a difference or error between the overlay of true and estimated roll angles. The set of simulation results 2000 are shown without additive noise and illustrate the performance of the servo loop 1900 (FIG. 19).

Note that the solution for equation 10 is ambiguous by n radians. Therefore, the loop 1900 of FIG. 19 should 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)=[d^T(k),v^T(k),a^T(k)]^T  (12)

Transition and measurement equations include a transition matrix, F_k, a plant noise coupling matrix, G_k, a measurement matrix, H_k, state transition noise w_k and measurement noise V_k.x_k+1=Fx_k+G_kw_k Z_k=H_kx_k+V_k  (13)

The state transition noise w_k is i.i.d. zero mean Gaussian with covariance R_k, and the measurement noise V_k is also i.i.d. zero mean Gaussian with covariance Q_k. The subscript k indicates the variables might be time varying.

The expanded form of these equations is given by:

$$\begin{gathered} \begin{array}{cc}\left[\begin{array}{c}{d}_{k+1}\\ v_{k+1}\\ a_{k+1}\end{array}\right]=\left[\begin{array}{ccc}{I}_3& T_s{I}_3& 0.5T_s^2{I}_3\\ 0_3& I_3& T_s{I}_3\\ 0_3& 0_3& I_3\end{array}\right]\left[\begin{array}{c}{d}_k\\ v_k\\ a_k\end{array}\right]+\left[\begin{array}{c}0.5T_s^2{I}_3\\ T_s{I}_3\\ I_3\end{array}\right]\left[\begin{array}{c}{w}_{d,k}\\ w_{v,k}\\ w_{wa,k}\end{array}\right]\text{ \\ [zd,kzv,kza,k]=[I3030303I3030303I3][dkvkak]+[vd,kvv,kva,k]}& \left(14\right)\end{array} \end{gathered}$$ where T_s is the sample interval and I₃ and 0₃ are 3×3 unity and zero matrices, respectively.

The Kalman filter 1720 (FIG. 17) involves two acts: the projection act and the update act.

The Projection Act is Given by:{circumflex over (x)}_k|k+1=F_k{circumflex over (x)}_k−1|k−1 {circumflex over (P)}_k|1+1=F_k{circumflex over (P)}_k−1|k−1F_k^T+G_kQ_kG_k^T  (15)The update act is given by:L_k=P_k|k−1H_k^T[H_k{circumflex over (P)}_k|k−1H_k^T+R_k]⁻¹ {circumflex over (x)}_k|k=F{circumflex over (x)}_k|k−1+L_k[Z_k−H_k{circumflex over (x)}_k|k−1]{circumflex over (P)}_k|k={circumflex over (P)}_k|k−1−{circumflex over (P)}_k|k−1H_k^T[H_k{circumflex over (P)}_k|k−1H_k^T+R_k]⁻¹H_k^T{circumflex over (P)}_k|k−1^T  (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)=[θ^T(k),ω^T(k)]^T  (17)

The transition and measurement equations have a similar form as the translational equations except the transition equation is nonlinear.x_k+1=F_k(y_k)+G_kw_k Z_k+1=H_ky_k+ν_k  (18)

The transition noise, w_k is i.i.d. zero mean Gaussian with covariance R_k and the measurement noise v_k is i.i.d. Gaussian with covariance Q_k. The nonlinear property is shown in the expanded form.

$$\begin{gathered} \begin{array}{cc}\left[\begin{array}{c}{\overline{x}}_{\theta ,k+1}\\ {\overline{x}}_{\omega ,k+1}\end{array}\right]=\left[\begin{array}{cc}{I}_3& T_{s}B\left({\theta }_k\right)\\ 0_3& I_3\end{array}\right]\left[\begin{array}{c}{\stackrel{\_}{x}}_{\theta ,k}\\ {\stackrel{\_}{x}}_{\omega ,k}\end{array}\right]+\left[\begin{array}{c}{0}_3\\ I_3\end{array}\right]\left[\begin{array}{c}{\stackrel{\_}{w}}_{\theta ,k}\\ {\stackrel{\_}{w}}_{\omega ,k}\end{array}\right]\text{ \\ [z\_ϕ,kz\_ω,k]=[10000003I3][x¯θ,kx¯ω,k]+[v\_ϕ,kv\_ω,k]}& \left(19\right)\end{array} \end{gathered}$$

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

$\begin{array}{cc}B\left({\theta }_k\right)=\left[\begin{array}{ccc}1& \sin \left(\varphi \right)\tan \left(\theta \right)& \cos \left(\varphi \right)\tan \left(\theta \right)\\ 0& \cos \left(\varphi \right)& -\sin \left(\varphi \right)\\ 0& \sin \left(\varphi \right)\sec \left(\theta \right)& \cos \left(\varphi \right)\sec \left(\theta \right)\end{array}\right]& \left(20\right)\end{array}$

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

$\begin{array}{cc}{\stackrel{\_}{F}}_k=\frac{\partial {\stackrel{\_}{F}}_k\left(y_k\right)}{\partial {\stackrel{\_}{x}}_k}=\left[\frac{\partial {\stackrel{\_}{F}}_k\left(y_k\right)}{\partial \theta }\frac{\partial {\stackrel{\_}{F}}_k\left(y_k\right)}{\partial \omega }\right]& \left(21\right)\end{array}$

The two derivatives are:

$\frac{\partial {\stackrel{\_}{F}}_k\left(y_k\right)}{\partial \theta }=\left[\begin{array}{ccc}1+T_s\left(\cos \left(\varphi \right)\tan \left(\theta \right){\omega }_{y,k}-\sin \left(\varphi \right)\tan \left(\theta \right){\omega }_{z,k}\right)& T_s\left(\sin \left(\varphi \right){\sec }^2\left(\theta \right){\omega }_{y,k}+\cos \left(\varphi \right){\sec }^2\left(\theta \right){\omega }_{z,k}\right)& 0\\ -T_s\sin \left(\varphi \right){\omega }_{y,k}-T_s\cos \left(\varphi \right){\omega }_{z,k}& 1& 0\\ T_s\left(\cos \left(\varphi \right)\sec \left(\theta \right){\omega }_{y,k}-\sin \left(\varphi \right)\sec \left(\theta \right){\omega }_{z,k}\right)& T_s\left(\sin \left(\varphi \right)\sec \left(\theta \right)\tan \left(\theta \right){\omega }_{y,k}+\cos \left(\varphi \right)\sec \left(\theta \right)\tan \left(\theta \right){\omega }_{z,k}\right)& 1\end{array}\right]\text{⁠}\frac{\partial {\stackrel{\_}{F}}_k\left(y_k\right)}{\partial \omega }=T_{s}B\left({\theta }_k\right)$

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=F_k{circumflex over (x)}_k−1|k−1 {circumflex over (P)}_k|k+1=F_k{circumflex over (P)}_k−1|k−1+G_kQ_kG_k^T  (22)

The update step for the EKF is given by:L_k={circumflex over (P)}_k|k−1H_k^T[H_k{circumflex over (P)}_k|k−1H_k^T+R_k]⁻¹ {circumflex over (x)}_k|k=F{circumflex over (x)}_k|k−1+L_k[Z_k−H_k{circumflex over (x)}_k|k−1]{circumflex over (P)}_k|k={circumflex over (P)}_k|k−1−{circumflex over (P)}_k|k−1H_k^T[H_k{circumflex over (P)}_k|k−1H_k^T+R_k]⁻¹H_k^T{circumflex over (P)}_k|k−1^T  (23)

FIGS. 21, 22, and 23 illustrate sets of simulation results of the EKF fusion [0163] algorithm using the EMM 900 radar, JAPRAE 1710 and IMU 1714 (FIG. 17) inputs for a typical KV 400 flight profile. FIG. 21 illustrates 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. 22 illustrates 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. 23 illustrates 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. 24 illustrates an attitude control loop 2400. The attitude control loop 2400 utilizes an uplink and onboard EKF outputs to compute firing commands to the plurality of attitude control thrusters (ACT) 1604 to maintain the KV 400 body axis 1520 aligned with the velocity vector V (FIG. 15).

In FIG. 24, {tilde over (d)}_KV and {tilde over (ν)}_KV are 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)}_KV is body rate vector. The projectile mass, CG location and inertia tensor are updated whenever there is a divert thruster 610 or an attitude control thruster 1604 firing. The attitude control loop 2400 may include an attitude control algorithm 2402. The attitude control algorithm 2402 may compute how much attitude control thrust is needed at the attitude thruster station 620 and 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 2404 then searches for an ACT to fire at operational block 2404. If the selected ACT is at the desired ignition position, the attitude thruster fire control 2406 electronics squib the selected ACT to fire at operational block 2406. Otherwise, the process exits the control loop. Finally, the attitude thruster units remaining function 2410 keeps track of which ACTs are spent and which are available for use. This information is fed back to the mass property estimator function 2412 and the attitude thruster select function 2404.

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 (ϕ)}_KV) attitude angles. These angles are used to compute the NED to body directional cosine matrix C_n^b as

$\begin{array}{cc}{C}_n^b=\left[\begin{array}{ccc}1& 0& 0\\ 0& \cos {\tilde{\varphi }}_{\mathrm{KV}}& \sin {\tilde{\varphi }}_{\mathrm{KV}}\\ 0& -\sin {\tilde{\varphi }}_{\mathrm{KV}}& \cos {\tilde{\varphi }}_{\mathrm{KV}}\end{array}\right]·\left[\begin{array}{ccc}\cos {\tilde{\theta }}_{\mathrm{KV}}& 0& -\sin {\tilde{\theta }}_{\mathrm{KV}}\\ 0& 1& 0\\ \sin {\tilde{\theta }}_{\mathrm{KV}}& 0& \cos {\tilde{\theta }}_{\mathrm{KV}}\end{array}\right]·\left[\text{⁠}\begin{array}{ccc}\cos {\tilde{\psi }}_{\mathrm{KV}}& \sin {\tilde{\psi }}_{\mathrm{KV}}& 0\\ -\sin {\tilde{\psi }}_{\mathrm{KV}}& \cos {\tilde{\psi }}_{\mathrm{KV}}& 0\\ 0& 0& 1\end{array}\right]& \left(24\right)\end{array}$

Act 2: The projectile velocity in the NED frame, {tilde over (V)}_KV, is tracked by EMM 900 radar and uplinked to the projectile via RF communication links. The velocity is then transformed into the body frame.{tilde over (V)}_KV^b=C_n^b·{tilde over (V)}_KV  (25)

Act 3: The total angle of attack α_total and aerodynamic roll angles ϕ_aero is 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.

$\begin{array}{cc}{\alpha }_{\mathrm{total}}={\tan }^{-1}\left(\frac{\sqrt{{{\tilde{V}}_{\mathrm{KV}}^b\left(2\right)}^2+{{\tilde{V}}_{\mathrm{KV}}^b\left(3\right)}^2}}{{\tilde{V}}_{\mathrm{KV}}^b\left(1\right)}\right){\phi }_{\mathrm{aero}}={\tan }^{-1}\left(\frac{{\tilde{V}}_{\mathrm{KV}}^b\left(2\right)}{{\tilde{V}}_{\mathrm{KV}}^b\left(3\right)}\right)& \left(26\right)\end{array}$

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 (ω)}_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)}_KV sin {tilde over (ϕ)}_KV  (27)

Act 5: Compute the KV 400 heading 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.

$\begin{array}{cc}{\psi }_{\mathrm{cmd}}={\tan }^{-1}\left(\frac{{\tilde{V}}_{\mathrm{KV}}\left(2\right)}{{\tilde{V}}_{\mathrm{KV}}\left(1\right)}\right){\theta }_{\mathrm{cmd}}={\tan }^{-1}\left(\frac{-{\tilde{V}}_{\mathrm{KV}}\left(3\right)}{\sqrt{{{\tilde{V}}_{\mathrm{KV}}\left(1\right)}^2+{{\tilde{V}}_{\mathrm{KV}}\left(2\right)}^2}}\right)& \left(28\right)\end{array}$

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 K_p,outer to form attitude angle rate command.{dot over (ψ)}_cmd=K_p,outer(ψ_cmd−{tilde over (ψ)}_KV){dot over (θ)}_cmd=K_p,outer(θ_cmd−{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 K_p,inner to form attitude acceleration commands.{umlaut over (ψ)}_cmd=K_p,inner({dot over (ψ)}_cmd−{dot over (ψ)}_KV){umlaut over (θ)}_cmd=K_p,inner({dot over (θ)}_cmd−{dot over (θ)}_KV)  (30)

FIG. 25 illustrates 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 algorithm 1710. The controller gain Kp_outer and Kp_inner are to be and optimized for various engagement scenarios. The symbol Euler_d represents the Euler rate vector └{dot over (ψ)}_KV, {dot over (θ)}_KV, {dot over (θ)}_KV┘. Euler_dd represents the Euler acceleration vector {umlaut over (└)}_Kv, {umlaut over (θ)}_KV, {umlaut over (ϕ)}_KV┘. F_cmd is the output of the controller containing the magnitude of the attitude thrust force command and its orientation relative to the body roll axis 1520. 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 KV 400 mass, CG location and inertia tensor will be described below with respect to FIG. 27. The total aerodynamic moment in the body frame, ignoring the rolling moment, may be computed as:

$\begin{array}{cc}{M}^b=\mathrm{QSd}\left\{\begin{array}{c}0\\ C_m\\ C_n\end{array}\right\}{C}_m=\left({\stackrel{\_}{X}}_{\mathrm{CP}}-{\stackrel{\_}{X}}_{\mathrm{CG}}\right)*C_N+C_{m_q}\stackrel{\_}{q}+C_A{\stackrel{\_}{Z}}_{\mathrm{CG}}{C}_n=\left({\stackrel{\_}{X}}_{\mathrm{CP}}-{\stackrel{\_}{X}}_{\mathrm{CG}}\right)*C_Y+C_{n_r}\stackrel{\_}{r}-C_A{\stackrel{\_}{Y}}_{\mathrm{CG}}& \left(31\right)\end{array}$

Where:

• • Q is dynamic pressure
  • S is reference area
  • d is reference diameter
  • C_m is total pitching moment coefficient
  • C_n is total yawing moment coefficient)
  • X_CP is normalized center of pressure location along the body x-axis)
  • X_CG is normalized center of gravity location along the body x-axis
  • Y_CG is CG offset from the body x-axis
  • Z_CG is CG offset from the body x-axis
  • C_A is axial drag-force coefficient
  • C_N is normal force coefficient
  • C_Y is lateral force coefficient
  • C_mg is pitch damping coefficient
  • C_n, is yaw damping coefficient
  • q is normalized pitching rate
  • r is normalized yawing 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:

$\begin{array}{cc}{M}_{cmd}=-\left[\begin{array}{ccc}0& 0& 0\\ 0& -\sin {\tilde{\varphi }}_{\mathrm{KV}}& -\cos {\tilde{\varphi }}_{\mathrm{KV}}\\ 0& -\cos {\tilde{\varphi }}_{\mathrm{KV}}& \sin {\tilde{\varphi }}_{\mathrm{KV}}\end{array}\right]\left\{\begin{array}{c}0\\ c_2\\ c_3\end{array}\right\}{I}_t.& \left(32\right)\end{array}$

I_t is the estimated moment of inertia about the lateral axis and, C₂ and C₃ are computed as:c₂=COS {tilde over (θ)}_KV{umlaut over (ψ)}_cmd−sin {tilde over (θ)}_KV{dot over (θ)}_KV{dot over (ψ)}_KV−({tilde over (ω)}_KV(2)cos {tilde over (ϕ)}_KV−{tilde over (ω)}_KV(3)sin {tilde over (ϕ)}_KV){dot over (ψ)}_KV−(I_t−I_a)({tilde over (ω)}_KV(1){tilde over (ω)}_KV(3)sin {tilde over (ϕ)}_KV+{tilde over (ω)}_KV(1){tilde over (ω)}_KV(2)cos {tilde over (ϕ)}_KV)C₃={umlaut over (θ)}_cmd+({tilde over (ω)}_KV(2)sin {tilde over (ϕ)}_KV+{tilde over (ω)}_KV(3)cos {tilde over (ϕ)}_KV){dot over (ϕ)}_KV−(I_t−I_a)({tilde over (ω)}_KV(1){tilde over (ω)}_KV(3)cos {tilde over (ϕ)}_KV−{tilde over (ω)}_KV(1){tilde over (ω)}KV(2)Sin {tilde over (ϕ)}_KV  (33)

I_a denotes the estimated moment of inertia about the roll axis.

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

$\begin{array}{cc}{F}_{cmd}=\left\{\begin{array}{c}0\\ M_{cmd}\left(3\right)-M^b\left(3\right)\\ -M_{cmd}\left(2\right)+M^b\left(2\right)\end{array}\right\}\frac{1}{X_{\mathrm{AT}}-X_{\mathrm{CG}}}& \left(34\right)\end{array}$

The Magnitude of this Command is:F_total=∥F_cmd∥  (35)

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

$\begin{array}{cc}{\varphi }_{cmd}={\tan }^{-1}\left(\frac{F_{cmd}\left(3\right)}{F_{cmd}\left(2\right)}\right)& \left(36\right)\end{array}$

Act 11: If the thrust force command, F_total, 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. 26 illustrates an example of a search procedure that may be used to find a thruster to fire. A plurality of attitude thrusters 1, 2, 3, 4, 5, 6, 7, and 8 are shown with circles. A white circle, such as attitude thrusters 1, 4, 6, and 8, indicates the attitude thruster has been spent. A shaded circle, such as attitude thrusters 2, 3, 5, and 7, 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 V^b and 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 ϕ_aero such 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, in FIG. 26, attitude thruster number 8 is about to reach the ignition command position; but it has already been spent, so it cannot be selected. Thruster number 7 is available but it's too soon to be fired so it will not be selected, either. However, as the KV 400 processes and nutates, the ignition command angle ϕ_ignit,cmd and thruster number 7 may 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 thruster 610 and the attitude control thruster 1604 propellants to the KV 400 total mass, the CG location and the inertia tensor will migrate as any divert thruster 610 or an attitude control thruster 1604 is 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. 27 illustrates a mass and CG location update algorithm 2700. The mass and CG location update algorithm 2700 may be configured to update the total mass, CG location and inertia tensor as a divert thruster 610 or attitude control thruster 1604 is firing. The definitions of the symbols are:{dot over (m)}_i≡i^th thruster mass burn rate<0ir_i≡i^th thruster CG locationm≡total massr_CGCg locationI_xy Inertia tensor element xy  (37)

The following equations are computed recursively:

$\begin{array}{cc}\Delta m_i={\stackrel{.}{m}}_i·\mathrm{dt}<0m\left(t_k\right)=m\left(t_{k-1}\right)+\Delta m_i{r}_{\mathrm{CG}}\left(t\right)=\frac{m\left(t_{k-1}\right)r_{\mathrm{CG}}\left(t_{k-1}\right)+\Delta m_i{r}_i}{m\left(t_k\right)}{I}_{\mathrm{xx}}\left(t_k\right)=I_{\mathrm{xx}}\left(t_{k-1}\right)-\sum \Delta m_i\left(y_i^2+z_i^2\right)I_{\mathrm{yy}}\left(t_k\right)=I_{\mathrm{yy}}\left(t_{k-1}\right)-\sum \Delta m_i\left(x_i^2+z_i^2\right)I_{\mathrm{zz}}\left(t_k\right)=I_{\mathrm{zz}}\left(t_{k-1}\right)-\sum \Delta m_i\left(x_i^2+y_i^2\right)I_{\mathrm{xy}}\left(t_k\right)=I_{\mathrm{xy}}\left(t_{k-1}\right)+\sum \Delta m_i{x}_i{y}_i{I}_{\mathrm{xz}}\left(t_k\right)=I_{\mathrm{xz}}\left(t_{k-1}\right)+\sum \Delta m_i{x}_i{z}_i{I}_{\mathrm{yz}}\left(t_k\right)=I_{\mathrm{yz}}\left(t_{k-1}\right)+\sum \Delta m_i{y}_i{z}_i& \left(38\right)\end{array}$

To a large extent, this detailed description has focused on a particular type of intercept vehicle (e.g., the eject vehicle 400). However, engagement management systems described herein may be used with many types of intercept vehicles 400 in which the engagement management system can track the intercept vehicle 400, alter the course of the intercept vehicle 400, determine when to detonate the intercept vehicle 400, or combinations thereof using commands communicated between the engagement management system and the intercept vehicle 400.

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.

Claims

1. An active protection system, comprising: an aircraft;an aerial vehicle carried by the aircraft, the aerial vehicle comprising: a rocket motor configured to accelerate the aerial vehicle in a direction having a component parallel to a longitudinal axis of the aerial vehicle; anda terminal section including alignment thrusters configured to cause the aerial vehicle to perform a pitch maneuver by rotating the aerial vehicle about an axis orthogonal to the longitudinal axis of the aerial vehicle prior to the rocket motor being fired, the rocket motor configured to detach from the terminal section after accelerating the aerial vehicle.

2. The active protection system of claim 1, wherein the terminal section further comprises divert thrusters configured to divert the terminal section in a direction transverse to the longitudinal axis of the aerial vehicle subsequent to the rocket motor being fired and detached from the terminal section.

3. The active protection system of claim 1, wherein the terminal section further comprises a nose thruster module proximate an airframe nose, the airframe nose at a front end of the aerial vehicle and the nose thruster module including the alignment thrusters.

4. The active protection system of claim 1, wherein the terminal section further comprises: a controller; andone or more sensors operably coupled to the controller, wherein the controller is configured to, subsequent to the rocket motor detaching from the terminal section, initiate flight corrections to the aerial vehicle during flight based on feedback from the one or more sensors.

5. The active protection system of claim 4, further comprising a guidance system carried by the aircraft and configured to transmit guidance signals to the aerial vehicle during flight of the aerial vehicle.

6. The active protection system of claim 1, further comprising a detection system carried on the aircraft and comprising at least one sensor configured to detect aerial threats within threat range of the aircraft.

7. The active protection system of claim 6, wherein the detection system is further configured to track aerial threats detected by the at least one sensor.

8. The active protection system of claim 6, further comprising an aerial vehicle carrier disposed on the aircraft.

9. A method of neutralizing an aerial threat, comprising: performing a first action of a launch sequence comprising an aerial vehicle leaving a platform;subsequent to the first action of the launch sequence, performing a second action of the launch sequence, the second action comprising performing a pitch maneuver by rotating the aerial vehicle about an axis orthogonal to a longitudinal axis of the aerial vehicle;subsequent to the second action of the launch sequence, firing a rocket motor of the aerial vehicle to cause the aerial vehicle to accelerate in a direction having a component parallel to the longitudinal axis of the aerial vehicle;subsequent to firing the rocket motor, detaching the rocket motor from a terminal section of the aerial vehicle.

10. The method of claim 9, wherein performing the first action of the launch sequence comprises releasing the aerial vehicle from the platform.

11. The method of claim 9, wherein performing the first action of the launch sequence comprises propelling the aerial vehicle from the platform.

12. The method of claim 9, wherein performing the first action of the launch sequence comprises launching the aerial vehicle via a dispenser of the platform.

13. The method of claim 9, further comprising performing a divert maneuver with one or more divert thrusters of the aerial vehicle to divert the aerial vehicle one or more times subsequent to: firing the rocket motor; and detaching the rocket motor from the terminal section.

14. The method of claim 9, further comprising performing the second action of the launch sequence after the aerial vehicle is a particular distance away from the platform.

15. The method of claim 9, wherein performing the first action of the launch sequence comprises releasing the aerial vehicle from a carrier disposed on the platform into air deflected by the platform.

16. The method of claim 9, further comprising initiating the first action of the launch sequence responsive to identifying the aerial threat.

17. A method of countering an aerial threat, comprising: causing an aerial vehicle to exit a vicinity of an aerial platform;causing the aerial vehicle to perform a pitch maneuver by rotating the aerial vehicle about an axis orthogonal to a longitudinal axis of the aerial vehicle;subsequent to causing the aerial vehicle to perform the pitch maneuver, causing a rocket motor of the aerial vehicle to fire and accelerate the aerial vehicle in a direction having a component parallel to the longitudinal axis of the aerial vehicle; andsubsequent to firing the rocket motor, causing the rocket motor to detach from a terminal section of the aerial vehicle.

18. The method of claim 17, further comprising identifying the aerial threat as a potential threat to at least one object.

19. The method of claim 18, wherein the at least one object is one of the aerial platform, a ground vehicle, another aerial platform, or a ground structure.

20. The method of claim 18, further comprising causing a payload of the aerial vehicle to detonate upon interception of the aerial threat with the aerial vehicle.