EP2307847A1 - Boîtiers électroniques composites, à amortissement actif et à fibres piézoélectriques - Google Patents

Boîtiers électroniques composites, à amortissement actif et à fibres piézoélectriques

Info

Publication number
EP2307847A1
EP2307847A1 EP08874892A EP08874892A EP2307847A1 EP 2307847 A1 EP2307847 A1 EP 2307847A1 EP 08874892 A EP08874892 A EP 08874892A EP 08874892 A EP08874892 A EP 08874892A EP 2307847 A1 EP2307847 A1 EP 2307847A1
Authority
EP
European Patent Office
Prior art keywords
housing
signal
fibers
response
control circuit
Prior art date
Legal status (The legal status is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the status listed.)
Ceased
Application number
EP08874892A
Other languages
German (de)
English (en)
Other versions
EP2307847A4 (fr
Inventor
Andrew B. Facciano
Robert T. Moore
Gregg J. Hlavacek
Craig D. Seasly
Current Assignee (The listed assignees may be inaccurate. Google has not performed a legal analysis and makes no representation or warranty as to the accuracy of the list.)
Raytheon Co
Original Assignee
Raytheon Co
Priority date (The priority date is an assumption and is not a legal conclusion. Google has not performed a legal analysis and makes no representation as to the accuracy of the date listed.)
Filing date
Publication date
Application filed by Raytheon Co filed Critical Raytheon Co
Publication of EP2307847A1 publication Critical patent/EP2307847A1/fr
Publication of EP2307847A4 publication Critical patent/EP2307847A4/fr
Ceased legal-status Critical Current

Links

Classifications

    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/02—Stabilising arrangements
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F16—ENGINEERING ELEMENTS AND UNITS; GENERAL MEASURES FOR PRODUCING AND MAINTAINING EFFECTIVE FUNCTIONING OF MACHINES OR INSTALLATIONS; THERMAL INSULATION IN GENERAL
    • F16F—SPRINGS; SHOCK-ABSORBERS; MEANS FOR DAMPING VIBRATION
    • F16F15/00—Suppression of vibrations in systems; Means or arrangements for avoiding or reducing out-of-balance forces, e.g. due to motion
    • F16F15/005—Suppression of vibrations in systems; Means or arrangements for avoiding or reducing out-of-balance forces, e.g. due to motion using electro- or magnetostrictive actuation means
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B10/00—Means for influencing, e.g. improving, the aerodynamic properties of projectiles or missiles; Arrangements on projectiles or missiles for stabilising, steering, range-reducing, range-increasing or fall-retarding
    • F42B10/02—Stabilising arrangements
    • F42B10/025—Stabilising arrangements using giratory or oscillating masses for stabilising projectile trajectory
    • F—MECHANICAL ENGINEERING; LIGHTING; HEATING; WEAPONS; BLASTING
    • F42—AMMUNITION; BLASTING
    • F42B—EXPLOSIVE CHARGES, e.g. FOR BLASTING, FIREWORKS, AMMUNITION
    • F42B15/00—Self-propelled projectiles or missiles, e.g. rockets; Guided missiles
    • F42B15/10—Missiles having a trajectory only in the air

Definitions

  • the present invention relates to systems and methods for controlling vibration. More specifically, the present invention relates to systems and methods for suppressing vibrations in missiles.
  • missiles are typically subject to severe vibration and shock during launch egress, flight ascent, and stage separation. If these vibration and shock loads are not mitigated, various system components may be damaged, causing the missile to fail.
  • Mission success requires that the missile be able to keep the target in its field-of-view while it maneuvers itself into a position to intercept the target.
  • a primary disturbance to the missile is the divert thrust delivered by the propulsion system. This thrust force tends to deform the missile into a beam bending mode at its first natural frequency. If the missile frequency modes (including the seeker frequency mode) have natural periods less than or on the same order as the divert thruster rise time, then significant dynamic amplification and airframe ringing will occur.
  • the dynamic amplification and the airframe ringing or vibration response make target tracking particularly difficult as the optical elements within the seeker will move relative or out of phase to each other producing significant seeker line- of-sight (LOS) motion.
  • Seeker pixel resolution can be maximized by providing a very rigid missile airframe to minimize the jitter transmitted to the seeker platform.
  • Missiles typically include a guidance system that relies on an inertial measurement unit (IMU) to determine the position of the missile by measuring its acceleration and rotation.
  • IMU inertial measurement unit
  • the IMU is extremely sensitive and should be very rigidly and precisely mounted to the missile airframe, which should also be very stiff. Otherwise, the EVIU will move around and make inaccurate measurements, causing the missile to tumble out of control.
  • the entire forebody assembly should therefore be made as stiff as possible to provide a stable platform for the IMU.
  • missile systems must typically be designed to attenuate flexible body dynamics or the system could have self-exciting vibrations. In the case where these vibrations are not bounded, catastrophic structural damage and mission failure may occur. In the case where the vibrations remain finite, the additional frequency content in the actuator commands can lead to actuator failure due to overheating and mission failure.
  • digital notch filters are used to attenuate the effects of the lower frequency modes (1st, and 2nd lateral modes, 1st torsional, and fin modes) and low-pass filters to attenuate the effects of the higher frequency modes.
  • a problem with this approach is that the use of digital filters results in phase loss at low frequencies, which limits the robust performance of the flight control system.
  • the notches associated with the 1 st lateral body mode are usually the lowest frequency modes and have the greatest impact on robust performance of the flight control system.
  • the traditional approach to these problems is to physically tune the structural responses of the missile components and assemblies (including the electronics housings and mounting structures, as well as the airframe and airframe joints) to mitigate these vibration loads.
  • This process typically involves iterative, long term dynamic analyses of the individual components and assemblies.
  • This highly detailed FEM analysis results in dynamic transfer functions incorporated into system guidance simulation evaluations, where further optimization is usually necessary, resulting in tuning requirements for the airframe again per analysis, iterating the transfer function and simulation studies.
  • Several different designs may be constructed and tested at great expense before a satisfactory design is found. This procedure has proven to be extremely time consuming, wrought with errors, and has led to significant program development schedule slippages and cost overruns.
  • the need in the art is addressed by the vibration controlled housing of the present invention.
  • the novel housing includes a housing structure and a mechanism for receiving a control signal and in accordance therewith electronically tuning a structural response of the housing structure.
  • the housing structure includes a composite material containing a plurality of piezoelectric fibers adapted to generate an electrical signal in response to a deformation in the structure and to deform the structure in response to an electrical signal applied thereto.
  • a control circuit receives the sensed signal from the fibers and generates an excitation signal that is applied to the fibers to increase the stiffness or compliance of the fibers at predetermined frequencies.
  • piezoelectric fiber composites are integrated into the missile airframe, seeker housing, guidance system housing, and missile mounting structures of a missile to control various vibration loads.
  • control signal is adapted to increase compliance of the fibers at high frequencies to dampen high frequency vibrations to protect system electronics, while at the same time increase stiffness of the fibers at low frequencies to provide a stable platform for the seeker and guidance system.
  • FIG. Ia is a cross-sectional view of a missile with a vibration control system designed in accordance with an illustrative embodiment of the present invention.
  • Fig. Ib is a simplified diagram of a missile with a layer of piezoelectric fiber composite attached to the missile airframe in accordance with an illustrative embodiment of the present invention.
  • Fig. 2a is a simplified diagram of a section of an illustrative piezoelectric fiber composite sensor/actuator that can be used in a vibration controlled component of the present teachings.
  • Fig. 2b is a simplified diagram of a section of an alternative piezoelectric fiber composite sensor/actuator that can be used in a vibration controlled component of the present teachings.
  • Fig. 3 is a simplified block diagram of a vibration control circuit designed in accordance with an illustrative embodiment of the present invention.
  • Fig. 4 is an exploded view of an illustrative missile with vibration controlled components designed in accordance with an alternative embodiment of the present invention.
  • Fig. 5a is a cross-sectional view of a Kinetic Energy Interceptor (KEI) missile with vibration controlled components designed in accordance with an alternative embodiment of the present invention.
  • KAI Kinetic Energy Interceptor
  • Fig. 5b is a simplified schematic of the kill vehicle and rocket motor of the illustrative KEI missile of Fig. 5 a.
  • Fig. 5c is a three-dimensional view of the internal components of the kill vehicle with vibration controlled components designed in accordance with an illustrative embodiment of the present invention.
  • Fig. 5d is a three-dimensional view of a seeker housing designed in accordance with an illustrative embodiment of the present teachings.
  • Fig. 5e is a three-dimensional view of an illustrative interstage adapter designed in accordance with an illustrative embodiment of the present teachings.
  • Fig. 6a is an illustration showing the missile bending such that its LOS is at an angle relative to the rigid body line of the missile.
  • Fig. 6b is a graph of the missile bending angle versus time.
  • the present teachings provide a novel vibration control method that integrates piezoelectric composite technology into missile components.
  • Piezoelectric composites generate electricity when they are flexed, and flex when a current or electric field is applied.
  • signals from a flexing composite part can be used by an integrated circuit (IC) to send back an excitation signal that the composite will respond to, attenuating and dampening the vibration.
  • IC integrated circuit
  • the ability to use an integrated circuit engineered to feedback a current which induces a response in the composite gives the ability to fine tune and tailor the feedback so that certain vibration frequencies or frequency ranges can be focused on for attenuation.
  • Fig. Ia is a cross-sectional view of a missile 10 with vibration controlled components designed in accordance with an illustrative embodiment of the present invention.
  • the missile 10 includes a forebody assembly 12 that is forward of the missile warhead and/or rocket motor 14.
  • the forebody assembly 12 includes a seeker assembly 16 and guidance system 18.
  • the seeker electronics of the seeker assembly 16 are housed in a novel electronics housing 20, which contains piezoelectric fiber composite sensor/actuators 30 for electronically tuning the structural response of the housing 20 in accordance with the teachings of the present invention.
  • the electronics modules of the guidance system 18 are housed in an electronics housing 22 that contains piezoelectric fiber composite sensor/actuators 30.
  • the missile forebody 12 also includes a mounting structure 24 for mounting the electronics to the missile airframe 26.
  • the mounting structure 24 also contains piezoelectric fiber composite sensor/actuators 30 to tailor the resonance characteristics of the mounting structure 24 to avoid resonance coupling with the electronic components (of the guidance system 18 and seeker 16).
  • the mounting structure 24 is a plate or bulkhead separating the forebody 12 from the warhead and/or rocket motor 14.
  • the guidance system housing 22 is mounted to the mounting structure 24, and the seeker housing 20 is mounted to the guidance system housing 22.
  • the missile airframe 26 itself also contains piezoelectric fiber composite sensor/actuators 30 for electronically tuning airframe stiffness and compliance dynamics.
  • Fig. Ib is a simplified diagram of a missile 10 with a layer of piezoelectric fiber composite 30 attached to the missile airframe 26 in accordance with an illustrative embodiment of the present invention.
  • the piezoelectric fiber composite sensor/actuators 30 perform "self- adjusting" or vibration damping functions.
  • the piezoelectric fiber composite sensor/actuators 30 are adapted to sense changes in motion (i.e., vibrations), which produces an electrical signal that is sent to a control circuit 32.
  • the control circuit 32 measures the magnitude of the change and relays a signal back to the fiber sensor/actuators 30 that either stiffens or relaxes the fiber sensor/actuators 30, producing a self-adjusting or "smart" structure, hi an illustrative embodiment, the sensor/actuators 30 and control circuit 32 are designed to stabilize the EVIU and seeker from low frequency airframe vehicle loads while attenuating high frequency vibrations from aero-buffeting, stage separation, and rocket vector shock loads.
  • Each vibration controlled component may have its own control circuit 30, or a single control circuit 30 may be configured to control vibrations in all of the components.
  • the vibration controlled components of the present invention may include a layer of piezoelectric fiber composite 30 glued or otherwise attached to the structure (as shown in Fig. Ib), or, in the preferred embodiment, the component is fabricated using the piezoelectric fiber composite 30, such that the piezoelectric fibers are embedded within the structure itself (as shown in Fig. Ia).
  • Fig. 2a is a simplified diagram of a section of an illustrative piezoelectric fiber composite sensor/actuator 30 which can be used in a vibration controlled component of the present teachings.
  • Fig. 2b is a simplified diagram of a section of an alternative piezoelectric fiber composite sensor/actuator 30 which can be used in a vibration controlled component of the present teachings.
  • the piezoelectric fiber composite 30 includes a plurality of piezoelectric fibers 42 arranged in parallel and surrounded by a matrix material 44 such as a resin or epoxy.
  • the composite 30 includes two opposing active surfaces 46 and 48.
  • a first electrode 50 is disposed on the first active surface 46 and a second electrode 52 is disposed on the second active surface 48.
  • the electrodes 50 and 52 are coupled to the control circuit 32.
  • the electrodes 50 and 52 are interdigital electrodes (as shown in Fig. 2b).
  • the piezoelectric fibers 42 may be aligned normal to the active surfaces 46 and 48, as shown in Fig. 2a, or they may be aligned parallel to the active surfaces 46 and 48, as shown in Fig. 2b, or they may be aligned at an angle to the active surfaces 46 and 48.
  • the piezoelectric fibers 42 are PZT (lead zirconium titanate) ceramic fibers made with relaxor materials. Methods for fabricating piezoelectric fiber composites are known in the art.
  • the piezoelectric fibers 42 will produce a current when deformed or flexed (i.e., by missile vibrations), and conversely will flex when exposed to an electric current or field.
  • the electrodes 50 and 52 are adapted to sense an electrical signal generated in the fibers 42 and also to apply an electrical signal from the control circuit 32 to the fibers 42.
  • the control circuit 32 generates an electrical actuator signal that is applied to the fibers 42 by the electrodes 50 and 52.
  • the fibers 42 flex in response to the signal, introducing a strain in the structure.
  • the control circuit 32 may be configured to provide active vibration damping by receiving a sensed signal from the fibers 42 and modulating the signal to form an actuator signal that is returned to the fibers 42 to dampen vibrations.
  • Fig. 3 is a simplified block diagram of a vibration control circuit 32 designed in accordance with an illustrative embodiment of the present invention.
  • control circuit 32 is configured to include a plurality of preprogrammed modes of operation, each mode generating a different actuator signal depending on a mode selection signal provided by the guidance system of the missile.
  • the mode selection signal indicates what operational phase the missile is in (for example, pre-launch, booster phase, guided flight, etc.).
  • the structural response of the vibration controlled components can therefore be changed to adapt to different environmental conditions.
  • the guidance system does not take over navigation of the missile until after the booster phase.
  • Providing a rigid platform for the IMU and seeker sensors is therefore not as important as protecting electronics during the booster phase (and also during handling before launch) when the guidance system is not controlling navigation.
  • the control circuit 32 can be configured to generate an actuator signal that reduces stiffness of the fibers 42 and attenuates vibrations, particularly at frequencies harmful to the electronics (e.g., high frequencies).
  • the control circuit 32 can then switch to a "guidance mode", generating an actuator signal adapted to increase the stiffness of the fibers 42 to provide a stable platform.
  • the frequency responses of the components can be controlled, for example, to avoid modal coupling between structures or to attenuate vibrations at frequencies that could be detrimental to the guidance system.
  • control circuit 32 can then be switched to a mode adapted to mitigate these shock loads. After the shock event is over, the control circuit 32 can then switch back to the guidance mode.
  • the control circuit 32 includes logic 60 for receiving the mode selection signal from the guidance system and loading the parameters associated with the selected mode from memory 62.
  • actuator signal e.g., the dc voltage component, how the sensor signal should be modulated for active vibration damping, etc.
  • the parameters for each mode are determined during missile testing and then stored on a RAM module 62.
  • the control circuit 32 also includes logic 64 for receiving a sensor signal measuring the amplitude and frequency of vibrations in the component, and modulating the sensor signal to form an actuator signal adapted to attenuate the sensed vibrations.
  • the actuator signal may simply be an out-of-phase version of the sensed signal, or it may be adapted to focus on attenuating vibrations in particular frequency ranges.
  • the sensor signal may be provided by the piezoelectric fibers 42, which generate an electrical signal when a vibration is applied to them. Alternatively, a separate sensor - which may also be a piezoelectric sensor - may be attached to the structure to measure vibrations.
  • the control circuit 32 also includes logic 66 for adding a dc voltage component to the actuator signal.
  • the dc voltage increases or decreases the stiffness of the fibers 42 and controls the frequency response of the structure as appropriate for the selected mode.
  • the final actuator signal is then applied to the fibers 42.
  • the control circuit 32 may be configured to return a finely tuned excitation signal designed to focus on certain frequencies or frequency ranges for vibration attenuation.
  • the control circuit 32 may be configured to return an excitation signal adapted to increase compliance of the fibers 42 at high frequencies to provide high frequency vibration isolation to protect electronics, while at the same time increase stiffness of the fibers 42 at low frequencies to provide low frequency stiffness and strength performance to achieve guidance system IMU and seeker alignment constraints.
  • the excitation signal may also be designed to attenuate certain resonance modes, counter modal coupling phenomena, and to attenuate seeker LOS jitter and smearing.
  • Captive carry loads due to aircraft flight environments may also be attenuated by tuning the missile components to dampen the fundamental bending mode for vibration suppression.
  • the control circuit 32 is implemented in a small, interlaminated IC chip.
  • the control circuit 32 may be implemented using, for example, discrete logic circuits, FPGAs, ASICs, etc.
  • the control circuit 32 may be implemented in software executed by a microprocessor. Other implementations can also be used without departing from the scope of the present teachings.
  • the control circuit 32 does not require an external power supply. If, however, a higher power excitation signal is desired, a battery may be added to supply additional power to the control circuit 32.
  • the present teachings provide vibration control using missile components with piezoelectric fiber composites controlled by an integrated circuit adapted to dynamically tune the frequency responses of the structures.
  • Extensive and iterative structural dynamic analyses, as in prior art applications, will no longer be required, since optimized tuning of the forebody dynamics can be simply programmed into the control chip for any frequency modulation change and readily implemented.
  • the desired frequency performance of a structural component may be changed due to simulation optimization studies, guidance software and payload hardware performance characterization changes, environmental load design evolutions, and test input revisions. In the past, this usually required system design changes, including complete redesigns of several assemblies.
  • the teachings of the present invention allow for changes to be made to the structural dynamics of the system by modifying the software within the vibration control circuit to shift frequency coupling performance parameters, instead of physically altering the structure (as in the prior art).
  • This attenuation method can be integrated into the electronics housings of the seeker and guidance system to protect electronics from high frequency vibrations while providing a stable platform for sensitive seeker and IMU equipment. It can also be integrated into bulkheads and mounting structures for further attenuation of electronics vibrations for avionic and seeker housing weight reductions, instead of adding heavy structural reinforcements, passive damping mounts (i.e. rubber mounts or dash-pods), or active tuning mechanisms (such as seeker steering mirrors) to achieve the same dynamic performance. In addition, integrating piezoelectric fiber composite technology into the missile airframe improves the airframe structural performance, and provides the ability to electronically tailor missile airframe frequency responses.
  • Figs. 1, 4, and 5 show different illustrative missile designs using vibration controlled components designed in accordance with the present teachings.
  • Figs. Ia and Ib showed a design that might be used in an air-to-air or surface-to-air missile.
  • Fig. 4 shows an alternate design that might be used in an air-to-air or surface-to-air missile, such as an ESSM (Evolved Sea Sparrow Missile), and Figs. 5a-5e show a design that might be used in a Kinetic Energy Interceptor (KEI) missile.
  • ESSM Evolved Sea Sparrow Missile
  • Figs. 5a-5e show a design that might be used in a Kinetic Energy Interceptor (KEI) missile.
  • KEI Kinetic Energy Interceptor
  • Fig. 4 is an exploded view of an illustrative missile 10' with vibration controlled components designed in accordance with an alternative embodiment of the present invention.
  • the missile 10' includes a mounting structure 24' which is an axial beam attached to the missile airframe (not shown).
  • the mounting beam 24' and missile airframe both contain piezoelectric fiber composite sensor/actuators in accordance with the teachings of the present invention.
  • a plurality of electronic components, each housed in a vibration controlled electronics housing 22 containing piezoelectric fiber composite sensor/actuators, are mounted to the mounting beam 24'.
  • a seeker housing 20 containing piezoelectric fiber composite sensor/actuators is also mounted to the mounting beam 24'.
  • Fig. 5a is a cross-sectional view of a Kinetic Energy Interceptor (KEI) missile 10" with vibration controlled components designed in accordance with an alternative embodiment of the present invention.
  • a KEI missile is configured to intercept enemy missiles during their boost phase, prior to mid-course ballistic ascent where the payload is uncovered and any RVs and possible decoys are deployed.
  • Booster phase interception also implies that any toxic materials dispersed during interception whether nuclear, biological, or nerve gas agents would fall back onto the country of origin with minimal liability to the defending forces positioned in the region.
  • Time-to-target is critical to the KEI mission; therefore high performance, lightweight airframe and electronics package technologies are needed to maximize Interceptor agility.
  • the KEI missile 10" includes a two-stage booster 70, a third-stage rocket motor 14", and a kill vehicle 12".
  • Fig. 5b is a simplified schematic of the kill vehicle 12" and rocket motor
  • the kill vehicle 12 includes a seeker assembly 16, guidance system electronics 18, and a lateral propulsion system 72.
  • the kill vehicle 12" components are attached to the rocket motor 14" by an interstage adapter structure 74.
  • Fig. 5c is a three-dimensional view of the internal components of the kill vehicle 12".
  • the kill vehicle 12" includes a lateral propulsion system 72, which includes a plurality of nozzles 80 and bottles of fluid 82 attached to a mounting structure 24".
  • a forward electronics assembly 18, which includes the IMU and guidance system electronics, is attached to the forward end of the mounting structure 24".
  • the seeker assembly 16 is attached to the forward electronics assembly 18.
  • An aft electronics assembly 84 is attached to the rear of the mounting structure 24".
  • the mounting structure 24" is attached to the interstage adaptor 74.
  • the mounting structure 24" and missile airframe each contain piezoelectric fiber composite sensor/actuators 30 and a control circuit 32 adapted to tune the structural responses of the components to provide a stable platform for the seeker and IMU while attenuating high frequency vibration.
  • the forward electronics assembly 18 and aft electronics assembly 84 are each housed in an electronics housing 22 containing piezoelectric fiber composite sensor/actuators 30 and a control circuit 32 adapted to dampen vibrations in the electronics assemblies.
  • the seeker assembly 16 includes a seeker housing 20, which also contains piezoelectric fiber composite sensor/actuators 30 and a control circuit 32 for providing a stable platform for the seeker components while attenuating vibrations. Fig.
  • FIG. 5d is a three-dimensional view of a seeker housing 20 designed in accordance with an illustrative embodiment of the present teachings.
  • Divert thrust forces generated by the propulsion system 72 can cause jitter and smear dynamics that affect seeker resolution and missile guidance and navigation.
  • Fig. 6a is an illustration showing the missile bending such that its LOS is at an angle ⁇ s relative to the rigid body line of the missile.
  • Fig. 6b is a graph of the missile bending angle versus time.
  • the vibration controlled components may also be adapted to mitigate LOS jitter and smearing that occur during propulsion ignition.
  • Fig. 5e is a three-dimensional view of an illustrative interstage adapter 18.
  • the KEI interstage adaptor 74 serves many functions as a transition structure between the kill vehicle 12" and the booster stack-up. Although it is not a large structure, it should be lightweight since burnout velocity is very sensitive to weight at the front end of the interceptor. It should also be sufficiently strong and stiff to preclude excessive deflection within the kill vehicle sway space, assuring it does not impact the enveloping nosecone.
  • the interstage adaptor 74 also contains piezoelectric fiber composite sensor/actuators 30 and a control circuit 32 adapted to attenuate vibrations traveling to the kill vehicle 12" and reduce the shock and vibration environment severity for the kill vehicle 12".
  • the adaptor structure 74 can be electronically tuned to provide sufficient airframe stiffness between the kill vehicle and interceptor booster to allow IMU functionality, while compliant enough to attenuate high frequency loads from damaging sensitive kill vehicle electronics and seeker assemblies.

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  • Engineering & Computer Science (AREA)
  • General Engineering & Computer Science (AREA)
  • Physics & Mathematics (AREA)
  • Aviation & Aerospace Engineering (AREA)
  • Fluid Mechanics (AREA)
  • Acoustics & Sound (AREA)
  • Mechanical Engineering (AREA)
  • Chemical & Material Sciences (AREA)
  • Combustion & Propulsion (AREA)
  • Aiming, Guidance, Guns With A Light Source, Armor, Camouflage, And Targets (AREA)
  • Radar Systems Or Details Thereof (AREA)
  • Vibration Prevention Devices (AREA)

Abstract

L'invention porte sur un boîtier à vibration contrôlée. Le nouveau boîtier comprend une structure de boîtier et un mécanisme pour recevoir un signal de commande et, en fonction de celui-ci, un accordage électronique d'une réponse structurelle de la structure. Dans un mode de réalisation donné à titre d'illustration, la structure de boîtier comprend un matériau composite contenant une pluralité de fibres piézoélectriques aptes à générer un signal électrique en réponse à une déformation de la structure et à déformer la structure en réponse à un signal électrique appliqué à celles-ci. Un circuit de commande reçoit le signal détecté provenant des fibres et génère un signal d'excitation qui est appliqué aux fibres pour accroître la rigidité ou l'élasticité des fibres à des fréquences prédéterminées. Dans un mode de réalisation donné à titre d'illustration, le signal de commande est apte à assurer les performances de rigidité et de résistance à basse fréquence tout en atténuant des vibrations haute fréquence pour protéger des composants électroniques reçus à l'intérieur de la structure.
EP08874892.6A 2008-07-02 2008-07-02 Boîtiers électroniques composites, à amortissement actif et à fibres piézoélectriques Ceased EP2307847A4 (fr)

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PCT/US2008/008291 WO2010002373A1 (fr) 2008-07-02 2008-07-02 Boîtiers électroniques composites, à amortissement actif et à fibres piézoélectriques

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EP2307847A1 true EP2307847A1 (fr) 2011-04-13
EP2307847A4 EP2307847A4 (fr) 2013-09-11

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US8573056B1 (en) * 2010-06-04 2013-11-05 The United States Of America As Represented By The Secretary Of The Army Guided projectile with motion restricting piezoelectric actuator
KR101490886B1 (ko) 2012-12-18 2015-02-09 국방과학연구소 압전 섬유 복합재료 구조체 및 이를 이용한 다축 힘 측정 장치
CN106569205A (zh) * 2016-11-01 2017-04-19 河北汉光重工有限责任公司 一种共孔径红外/雷达复合导引头

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EP2307847A4 (fr) 2013-09-11
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