Abstract
The present invention relates to the control of munitions, missiles and projectiles, in flight. The present invention further relates to systems and methods for control of munitions, missiles and projectiles in flight with the use of activatable or deployable flow effectors that remain stowed or inactive during launch or firing, and can be actuated after launch or firing on demand. More specifically, the present invention relates to systems and methods for control of munitions, missiles, and projectiles by activating and/or deactivating a control actuation system (CAS) based on measurements of an inertial measurement unit (IMU) and sensors integrated into such IMU, the IMU and sensors being at least part of a configurable guidance sensor suite (CGSS).
Claims
1. A missile, munition, or projectile containing a flight control system comprising: a missile, munition, or projectile body; a control actuation system (CAS) adapted to be placed within the body of the missile, munition, or projectile, the CAS comprising at least one deployable flow effector or control surface, and a deployment mechanism adapted to maintain a state of tension on the at least one flow effector or control surface until the tension is released and the at least one flow effector or control surface deploys, the CAS further comprising at least one component adapted to maintain the position of the at least one flow effector or control surface in position during flight after being deployed; at least one image or video sensor adapted to provide real-time image or video data; a transceiver adapted for two-way communication between the missile, munition, or projectile and a remote user interface; and a situational awareness subsystem comprising the at least one image or video sensor and the transceiver, the situation awareness subsystem adapted to transmit the real-time image or video data via the transceiver to a user to provide terminal guidance to the missile, munition, or projectile via the remote user interface.
2. The flight control system of claim 1, wherein the CAS further comprises a motor, a planetary gear, and an encoder adapted to interface with the at least one component adapted to maintain the position of the at least one flow effector or control surface in position during flight after being deployed, wherein such component is a lead screw and lead nut adapted to prevent backdrive of deployed flow effectors or control surfaces caused by aerodynamic forces during flight such that the flow effectors or control surfaces remain in position without requiring power.
3. The flight control system of claim 2, wherein the CAS and transceiver are integrated into a single enclosure adapted to be placed within the missile, munition, or projectile body.
4. The flight control system of claim 3, wherein the situational awareness subsystem is further adapted to perform target prioritization in flight, including target detection, identification, and tracking, and where the terminal guidance is based on the in-flight target prioritization.
5. The flight control system of claim 4, wherein the image or video sensor is adapted to provide real-time, in-flight video signals and the remote user interface is adapted to receive signals and data from the transceiver and to allow the user to control flight of the missile, munition, or projectile, based at least in part on the real-time, in-flight video signals from the image or video sensor, and at least in part on measured flight data from the at least one integrated IMU.
6. The flight control system of claim 3, wherein the flight control system is adapted to withstand forces greater than 20,000 g.
Description
BRIEF DESCRIPTION OF THE DRAWINGS
(1) FIGS. 1A-C. Several views of a CAS design embodiment of the present invention including A) CAS with activated or deployed flow effectors, B), a cutaway view of CAS showing the drivetrain components, and C) CAS with stowed flow effectors.
(2) FIG. 2. Cutaway view of one embodiment of a CAS depicting important components and their placement within the CAS.
(3) FIG. 3. Close up view of single axis of CAS showing interaction between motor to lead screw to flow effector barrel. Left is forward (nose), right is aft (tail).
(4) FIG. 4. Diagram of one embodiment of a flow effector or control surface deployment mechanism.
(5) FIGS. 5A-C. Picture depicting one embodiment of the control electronics for the CAS of the present invention.
(6) FIG. 6. Block diagram of control electronics and firmware architecture of one embodiment of the present invention.
(7) FIGS. 7A-D. Several views of one embodiment of the IMU of the present invention depicting various sensors, such views including A) perspective view, B) circuit diagram, C) side view, and D) top view.
(8) FIGS. 8A-D. Pinout designs for various components of one embodiment of the IMU of the present invention, including A) VN-100x, B) ADXL2780-50g, C) FXAS21002-Gyro, and STM32F405OGY.
(9) FIG. 9. Clock tree configuration for one embodiment of the IMU of the present invention.
(10) FIG. 10. Block diagram of one embodiment of the IMU of the present invention showing communication architecture between subsystems and components.
(11) FIG. 11. Graph depicting test data resulting from air gun tests of rounds fired using one embodiment of each of the CAS and IMU of the present invention.
(12) FIG. 12. Graph depicting test data resulting from a 155 mm gun launch of rounds fired using one embodiment of each of the CAS and IMU of the present invention.
(13) FIGS. 13A-D. Graphs depicting flight data from a 40 mm round launch showing IMU sensor output for LV 40 mm, over about 7 seconds, including A) accelerometer data, B) gyroscope data, C) magnetometer data, and D) Euler angle data.
(14) FIGS. 14A-C. Graphs depicting flight data from a 120 mm mortar test showing IMU sensor output, including A) acceleration data, B) angular rate data, and C) magnetic field data.
(15) FIGS. 15A-F. Various embodiments of the CAS of the present invention scaled to fit various platforms including: A) 40 mm grenade, B) 120 mm mortar, C) 155 mm ERPT, D) 40 mm grenade, E) 120 mm mortar, F) 155 mm ERPT, and G) 40 mm rocket-assist projectile.
(16) FIG. 16. Cross-sectional view of one embodiment of a missile, munition, or projectile depicting the CAS and CGSS as oriented within an enclosure of the missile, munition, or projectile.
DETAILED DESCRIPTION OF THE INVENTION AND DRAWINGS
(17) Various embodiments of the CAS of the present invention include several important components in various combinations. The components may include, but are not limited to, a motor (brushless or brushed DC motor, for example), encoder, a gear or gear system, lead screw, microcontroller, and motor driver or controller.
(18) Now referring to the figures, FIGS. 1A-C depict several views of one design embodiment of the CAS 100 of the present invention. FIG. 1A depicts the CAS 100 with activated or deployed flow effectors 105. The depicted CAS 100 design is capable of operating and successfully activating or deploying the flow effectors even under high-g conditions after firing. High-g conditions include g-force loads of greater than 20,000 g's. FIG. 1B depicts a cutaway view of the CAS 100 embodiment including placement of key components described in greater detail in FIG. 2. FIG. 1C depicts the CAS 100 embodiment with stowed flow effectors 100 within the housing of the CAS, that is the flow effectors are unactuated or undeployed. When the flow effectors 110 are stowed, they are preferably flush or sub-flush to the surface of the munition, missile or projectile, which may include being within the body of the CAS 100.
(19) FIG. 2 depicts a cutaway view of a CAS 200 embodiment showing several important components of the CAS including the control electronics 205, the motor 210, which in the depicted embodiment is a BLDC motor, encoder 215, lead screw and nut 220, flow effector or control surface barrel gear 225, flow effector or control surface 230, deployment mechanism 235, and CAS housing 240. The CAS 200 housing 240 preferably is a customized or customizable aero-shell that can provide various mounting options for the CAS 200 to be mounted into numerous types of munitions, missiles and projectiles.
(20) An important consideration for the present invention to provide precision flight control is attaining positional feedback from flow effectors or control surfaces 230, which is achieved through an encoder 215 on the motor 210. The depicted encoder 215 has 256 counts per revolution with 3 channels, and uses differential EIA RS 422 driver logic. This encoder 215 has been tested for shock successfully up to levels greater than 20,000 g's. The encoder 215 package is housed in an aluminum structure and welded to the motor 210/gear stack 225 to increase robustness.
(21) The CAS 200 includes stow/deploy capability to survive the high-spin and high-g launch environment—the flow effectors or control surfaces 230 may be deployed and/or retracted as needed.
(22) The CAS 200 preferably minimizes weight and survivability through optimal material selection. The CAS 200 housing 240 may be made of various materials known in the art to be strong yet lightweight, and in at least one embodiment the housing 240 is constructed of Grade 5 Titanium as one example, while the forward and aft bulkheads may use A17075T*. The flow effectors or control surfaces 230 are preferably made of either titanium or aluminum, but may be constructed of other materials that have the potential to increase performance in the specified aerodynamic environment.
(23) FIG. 3 depicts a close-up interior view of a single axis of one embodiment of the CAS. The motor 300, a BLDC in the depicted embodiment, interacts with the lead screw 305, which, in turn, drives the lead screw nut 310. The interaction between these components is then translated to the flow effector barrel 315 to drive the flow effector to move. The flow effector may be deployed or retracted utilizing these components.
(24) FIG. 4 is a diagram of one embodiment of the deployment mechanism for the flow effectors or control surfaces of the present invention. The flow effectors or control surfaces 400 are attached to the flow effector or control surface barrel 420 via a shoulder bolt 410. The flow effector or control surface 400 is mounted in a state of tension courtesy of a tension component, depicted in the present figure as a torsion spring with a lock 405 such that when the flow effector or control surface 400 is stowed, the torsion spring provides tension that, when released, allows the flow effector or control surface 400 to activate or deploy. A spring-loaded locking pin 415 is disposed such that, when the flow effector or control surface 400 is deployed, the locking pin moves into place or otherwise is situated to prevent the flow effector or control surface 400 from retracting unintentionally, such as due to recoil from deployment or due to high g-forces against the flow effector or control surface 400.
(25) When stowed, the mechanical design is such that the non-operational spin rate will not cause a pre-trigger deployment. This deployment mechanism minimizes components and utilizes the motors to drive the canards into a deployed state when commanded. This rotation allows the internal torsion spring to release the flow effector or control surface 400 out and they continue to travel until they reach the full-stop. At this stop, a ball-detent is used to pin and hold the flow effector or control surface in the deployed state. This mechanism reduces recoil chances and allows for smooth deployment of the flow effector or control surface into the airstream in the forward-to-aft direction, leveraging the airflow to assist in deployment.
(26) FIGS. 5A-C depict several views of one embodiment of the control electronics. FIG. 5A shows the control electronics interfaced with the forward bulkhead of the CAS. FIG. 5B shows a bottom-view of the control electronics, including a microcontroller and 422-driver. FIG. 5C depicts a top view of the control electronics including controls/drive integrated circuits for DC motors.
(27) FIG. 6 is a block diagram of control electronics and firmware architecture of one embodiment of the present invention. The control electronics, as for the depicted embodiment, support a 4-channel CAS for various munitions, missiles or projectiles. The control electronics, and thus the CAS, preferably do not initiate any messages, but rather primarily receive and respond to commands from a host system, such as the munition, missile or projectile or other components thereof. Such commands input to the control electronics and CAS may include, but are not limited to, motor commands and commands to retrieve the current motor position. Responses or outputs of the control electronics and CAs may include, but are not limited to, positioning or repositioning the motor, returning information regarding the position of the motor or invalid command responses.
(28) FIGS. 7A-D include several views of one embodiment of the IMU of the present invention depicting various sensors, such views including A) perspective view, B) circuit diagram, C) side view, and D) top view. In the various views of the IMU, sensors including multi-axis accelerometers, gyroscopes, and a separate integrated sensor suite can be seen, as well as a microcontroller. The separate integrated sensor suite includes one or more of an accelerometer, gyroscope, magnetometer and DSP.
(29) FIGS. 8A-D depict pinout designs for various components of one embodiment of the IMU of the present invention, including A) separate integrated sensor suite or package, B) accelerometer(s), C) gyroscope(s), and STM32F405OGY.
(30) FIG. 9 depicts a block diagram of clock tree configuration for one embodiment of the IMU of the present invention. This diagram shows one embodiment the IMU clock tree configuration that can be used for analysis and/or simulation, and can provide a basis for further development of the IMU.
(31) FIG. 10 is a block diagram of one embodiment of the IMU of the present invention showing communication architecture between subsystems and components
(32) FIG. 11 is a graph depicting test data resulting from air gun tests of rounds fired using the CAS and IMU of the present invention. The graph demonstrates that the CAS and IMU of the present invention exhibits high survivability and calibration ability, even under high-g effects.
(33) FIG. 12 is a graph depicting test data resulting from a 155 mm gun launch of rounds fired using one embodiment the CAS and IMU of the present invention. The test data demonstrates survivability and operation of the CAS and IMU during and after a gun-launch event in high-g conditions up to about 25,000 g's.
(34) FIGS. 13A-D Graphs depicting flight data from a 40 mm round launch showing IMU sensor output for LV 40 mm, over about 7 seconds. FIG. 13A shows accelerometer data from a live fire test utilizing a 40 mm round with an embodiment of the IMU mounted therein. FIG. 13B shows gyroscope data from a live fire test utilizing a 40 mm round with an embodiment of the IMU mounted therein. FIG. 13C shows magnetometer data from a live fire test utilizing a 40 mm round with an embodiment of the IMU mounted therein. FIG. 13D shows Euler angle data from a live fire test utilizing a 40 mm round with an embodiment of the IMU mounted therein.
(35) FIGS. 14A-C Graphs depicting flight data from a 120 mm mortar test showing IMU sensor output. FIG. 14A shows acceleration data from a 120 mm mortar test with an embodiment of the IMU mounted in the round. FIG. 14B shows angular rate data from a 120 mm mortar test with an embodiment of the IMU mounted in the round. FIG. 14C shows magnetic field data from a 120 mm mortar test with an embodiment of the IMU mounted in the round.
(36) FIGS. 15A-G are pictures of various embodiments of the CAS of the present invention scaled to fit various platforms including: 15A) 40 mm grenade, 15B) 120 MM mortar, 15C) 155 mm ERPT, 15D) 40 mm grenade, 15E) 120 mm mortar, 15F) 155 mm ERPT, and 15G) 40 mm rocket-assist projectile. FIG. 15A shows a 4-channel CAS for a 40 mm grenade. FIG. 15B shows a 4-channel CAS for a 120 mm mortar. FIG. 15C shows a 2-channel CAS for a 155 mm ERPT. FIG. 15D shows a 2-channel CAS for a 40 mm grenade. FIG. 15E shows a 2-channel CAS for a 120 mm mortar. FIG. 15F shows a 2-channel CAS for a 155 mm ERPT. FIG. 15G shows a 4-channel 40 mm rocket-assist projectile with one embodiment of a CAS of the present invention.
(37) FIG. 16 depicts a cross-sectional view of a missile, munition, or projectile. The missile, munition, or projectile comprises a body or enclosure 1600, in which are housed a control actuation system (CAS) 1605 and a configurable guidance sensor suite (CGSS) 1610. The CAS 1605 is similar to that depicted in FIGS. 1A-C, and 2-4. The CGSS 1610 depicted in the present figure is similar to the IMU unit depicted in FIGS. 7A-7D.
(38) It will be apparent to those skilled in the art that various modifications and variations can be made to the present invention without departing from the spirit and scope of the invention. Thus, it is intended that the present invention cover the modifications and variations of this invention provided they come within the scope of the appended claims and their equivalents.
