JP2011526989A - Piezoelectric fiber active damped composite electronic housing - Google Patents

Abstract

Vibration control housing. The new housing includes a housing structure and a mechanism that receives a control signal and electronically tunes the structural response of the housing structure in accordance with the control signal. In an exemplary embodiment, the housing structure includes a composite material composed of a plurality of piezoelectric fibers, the plurality of piezoelectric fibers generating an electrical signal in response to deformation in the housing structure, and the plurality of piezoelectric fibers The housing structure is configured to deform in response to an electrical signal applied to the housing. The control circuit receives a sensing signal from the piezoelectric fiber and generates an excitation signal. An excitation signal is applied to the piezoelectric fiber to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency. In an exemplary embodiment, the control signal is configured to provide low frequency stiffness and strength performance while attenuating high frequency vibrations to protect electronics housed within the housing structure. ing.
[Selection] Figure 1a

Description

  The present invention relates to a system and method for controlling vibration. More specifically, the present invention relates to a system and method for suppressing vibration in missiles.

  In general, in highly dynamic environments, missiles are subject to intense vibrations and shocks during launch egress, flight ascent and stage detachment. If these vibration and impact loads are not mitigated, various system components can be damaged and cause missile failure.

  Successful missions require the missile to steer itself to a position where it will intercept the target while the missile can maintain the target in its field of view. The main hindrance to missiles is the divert thrust provided by the propulsion system. This thrust force has the property of deforming the missile into the beam bending mode at the primary natural frequency. When the missile frequency mode (including the seeker frequency mode) has a natural period less than or equal to the rise time of the divert thruster, significant dynamic amplification and airframe ringing occurs.

  Dynamic amplification and airframe ringing or vibration response make target tracking difficult. The reason is that the optical elements within the seeker move relative to each other or out of phase and the seeker line-of-sight (LOS) moves significantly. By providing a very stiff missile airframe and minimizing jitter transferred to the Seeker platform, the Seeker pixel resolution can be maximized.

  In addition, the missile must be able to accurately determine its own position in order to calculate the flight path to intercept the target. Missiles typically include a guidance system. The guidance system relies on an inertial measurement unit (IMU) to determine the missile position by measuring the acceleration and rotation of the missile. The IMU is extremely delicate and should be mounted very firmly and accurately on the missile airframe. The missile airframe should also be very rigid. Otherwise, the IMU will move around, making measurements inaccurate and making the missile uncontrollable. Therefore, the entire front fuselage assembly should be as rigid as possible to provide a stable platform for the IMU.

  Unfortunately, stiffening the fuselage for better performance of the IMU and seeker can result in undesirable high frequency vibrations due to rocket motor ignition, stage detachment, aerodynamic buffeting, and acoustic loading. And impact loads can be transmitted. When these vibration loads are coupled to an electronic component, the electronic component can be severely damaged, resulting in missile failure. Further, generally, as the structure becomes rigid, the mass and weight increase, affecting the maneuverability and range of the missile.

  For example, efforts to make a structure more compliant by isolating electronic components using rubber mounts can dampen high frequency vibrations, but if the structure is made too compliant, the IMU will have a missile trajectory. Accurate displacement and rotation cannot be read out. Thus, an important issue faced when packaging electronics is to provide sufficient isolation from separation and divert impact loads and to allow the IMU platform to function while meeting strength and weight requirements. This is a trade-off with providing sufficient stiffness.

  In addition, missile systems typically must be designed to damp the dynamics of the flexible body. Otherwise, the system may vibrate. If these vibrations are not suppressed, the structure can be catastrophic and the mission can fail. If the vibration remains finite and a frequency component is added to the actuator command, the actuator may fail due to overheating and the mission may fail. Currently, digital notch filters are used to reduce the effects of lower frequency modes (primary and secondary transverse modes, primary torsion modes, fin modes). Furthermore, a low-pass filter is used to weaken the influence of higher frequency modes. The problem with this approach is that using a digital filter results in phase loss at low frequencies. This limits the robust performance of the control system. The notches associated with the primary transverse body mode are usually the lowest frequency modes and have the greatest impact on the robust performance of the control system.

  Traditional approaches to these problems reduce these vibrational loads by physically tuning the structural response of missile components and assemblies (including electronics housing, mounting structure, and fuselage interface). To do. In general, this process involves repeated dynamic analysis of individual components and assemblies over a long period of time. This very detailed FEM analysis will incorporate a dynamic transfer function into the evaluation of the system-guided simulation. Usually, if further optimization is required, the requirements for the aircraft will be adjusted again for each analysis, and the transfer function and simulation studies will be repeated. Several different designs may be built and tested at great expense until a satisfactory design is found. This procedure is very time consuming and has been found to work with errors, leading to significant delays in program development schedules and cost overruns.

  Accordingly, there is a need in the art for an improved system or method that reduces missile vibration loads that is simpler, less expensive and less time consuming than conventional approaches.

  What is needed for this technique is addressed by the vibration control housing of the present invention. The new housing includes a housing structure and a mechanism that receives a control signal and electronically tunes the structural response of the housing structure in accordance with the control signal.

  In an exemplary embodiment, the housing structure includes a composite material having a plurality of piezoelectric fibers, the plurality of piezoelectric fibers generating an electrical signal in response to deformation in the housing structure and applied to the plurality of piezoelectric fibers. The housing structure is configured to be deformed in response to the electrical signal. The control circuit receives a sensing signal from the piezoelectric fiber and generates an excitation signal. An excitation signal is applied to the piezoelectric fiber to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency.

  In accordance with the teachings of the present invention, piezoelectric fiber composites are integrated into the missile fuselage, seeker housing, guidance system housing, and missile mounting structure to control various vibration loads. In an exemplary embodiment, the control signal increases the compliance of the piezoelectric fiber at high frequencies, weakens the high frequency vibrations and protects the system electronics, while at the same time increasing the stiffness of the piezoelectric fiber at low frequencies. And is configured to provide a stable platform for the seeker and guidance system.

1 is a cross-sectional view of a missile with a vibration control system designed according to an exemplary embodiment of the present invention. 1 is a simplified illustration of a missile with a layer of piezoelectric fiber composite material attached to the missile airframe, in accordance with an exemplary embodiment of the present invention. FIG. 1 is a simplified diagram of a portion of an exemplary piezoelectric fiber composite sensor / actuator that can be used in the vibration control component taught in the present invention. FIG. FIG. 4 is a simplified diagram of a portion of another piezoelectric fiber composite sensor / actuator that can be used in the vibration control component taught in the present invention. FIG. 2 is a simplified block diagram of a vibration control circuit designed in accordance with an exemplary embodiment of the present invention. FIG. 4 is an exploded view of an exemplary missile with vibration control components designed in accordance with another embodiment of the present invention. FIG. 3 is a cross-sectional view of a Kinetic Energy Interceptor (KEI) missile with a vibration control component designed in accordance with another embodiment of the present invention. FIG. 5b is a simplified schematic diagram of the exemplary KEI missile grenade and rocket motor of FIG. 5a. FIG. 3 is a three-dimensional view of the internal components of a grenade with a vibration control component designed in accordance with an exemplary embodiment of the present invention. FIG. 3 is a three-dimensional view of a seeker housing designed in accordance with an exemplary embodiment taught in the present invention. 3 is a three dimensional view of an exemplary interstage adapter designed in accordance with exemplary embodiments taught in the present invention. FIG. FIG. 5 is a diagram showing missile bending where the LOS of the missile is at an angle with respect to the hard body line of the missile. 2 is a graph of missile bending angle versus time.

  In order to disclose the advantageous teachings of the present invention, exemplary embodiments and exemplary applications will now be described with reference to the accompanying drawings.

  Although the present invention has been described herein with reference to exemplary embodiments for particular applications, it should be understood that the invention is not limited thereto. Those skilled in the art having access to the teachings provided herein will understand further modifications, applications, and embodiments within the scope of the invention and further areas where the invention will be very useful. Will.

  The teachings of the present invention provide a novel vibration control method. This novel vibration control method integrates piezoelectric composite technology into missile components. Piezoelectric composites generate electricity when bent and bend when a current or electric field is applied. Using this technique, an integrated circuit (IC) uses a signal from a portion of a bent composite material to send back an excitation signal, and the composite material dampens vibration in response to the excitation signal. Can be weakened. This has a net reinforcing effect. Using piezocomposites to build missile components can aid in efforts to optimize weight by allowing lighter designs to achieve the same strength as undamped designs in addition to vibration control . In addition, integrated circuits designed to feed back current and cause the current to cause a response in the composite material can be used, so the feedback can be fine-tuned and adapted to a specific vibration frequency or frequency range to dampen. Can focus.

  FIG. 1a is a cross-sectional view of a missile 10 with a vibration control component designed in accordance with an exemplary embodiment of the present invention. Missile 10 includes a front fuselage assembly 12. The front fuselage assembly 12 is in front of the missile warhead and / or rocket motor 14. The front fuselage assembly 12 includes a seeker assembly 16 and a guidance system 18. The seeker electronics of the seeker assembly 16 are housed in a new electronics housing 20. In accordance with the teachings of the present invention, the novel electronics housing 20 includes a piezoelectric fiber composite sensor / actuator 30 that electronically tunes the structural response of the housing 20. Similarly, the electronic device module of the guidance system 18 is accommodated in the electronic device housing 22. The electronics housing 22 contains a piezoelectric fiber composite sensor / actuator 30.

  Further, the missile front fuselage 12 includes a mounting structure 24 for mounting electronic equipment on the missile fuselage 26. In accordance with the teachings of the present invention, the mounting structure 24 also includes a piezoelectric fiber composite sensor / actuator 30 that adjusts the resonant characteristics of the mounting structure 24 so that the electronic component (of the induction system 18 and seeker 16) Resonant coupling is avoided. In the exemplary embodiment of FIG. 1 a, the mounting structure 24 is a plate or bulkhead that separates the front fuselage 12 from the warhead and / or rocket motor 14. The guidance system housing 22 is mounted on a mounting structure 24. Seeker housing 20 is mounted on housing 22 of the guidance system.

  In the preferred embodiment, the missile fuselage 26 itself also includes a piezoelectric fiber composite sensor / actuator 30 to electronically tune the stiffness and compliance dynamics of the fuselage. According to an exemplary embodiment of the present invention, FIG. 1 b is a simplified illustration of a missile 10 with a layer of piezoelectric fiber composite 30 attached to the missile fuselage 26.

  The piezoelectric fiber composite sensor / actuator 30 performs an “automatic adjustment” or vibration damping function. The piezoelectric fiber composite sensor / actuator 30 is configured to sense changes in motion (ie, vibration), generate an electrical signal, and send the electrical signal to the control circuit 32. The control circuit 32 measures the magnitude of the change and relays the signal back to the fiber sensor / actuator 30 to cure or mitigate the fiber sensor / actuator 30 to create an automatic adjustment or “smart” structure. Generate. In the exemplary embodiment, the sensor / actuator 30 and control circuit 32 stabilize the IMU and seeker from low frequency airframe vehicle loads, while aerobatting, stage decoupling, and rocket vector Designed to damp high frequency vibrations from impact loads. Each vibration control component (seeker housing 20, induction housing 22, mounting structure 24, and fuselage 26) has its own control circuit 30, or a single control circuit 30 provides vibration in all of the components. It can be configured to control.

  The vibration control component of the present invention may include a layer of piezoelectric fiber composite 30 that is adhered or otherwise attached to the structure (see FIG. 1b). Alternatively, in a preferred embodiment, the component is made using a piezoelectric fiber composite 30 so that the piezoelectric fiber is embedded within the structure itself (see FIG. 1a).

  FIG. 2a is a simplified diagram of a portion of an exemplary piezoelectric fiber composite sensor / actuator 30 that can be used in the vibration control component taught in the present invention. FIG. 2b is a simplified diagram of a portion of an alternative piezoelectric fiber composite sensor / actuator 30 that can be used in the vibration control component taught in the present invention. The piezoelectric fiber composite material 30 includes a plurality of piezoelectric fibers 42. The plurality of piezoelectric fibers 42 are arranged in parallel and surrounded by a matrix material 44 such as resin or epoxy. The composite material 30 includes two active surfaces 46, 48 facing each other. The first electrode 50 is disposed on the first active surface 46. The second electrode 52 is disposed on the second active surface 48. The electrodes 50 and 52 are connected to the control circuit 32. In the exemplary embodiment, the electrodes 50, 52 are interdigital electrodes (see FIG. 2b). The piezoelectric fiber 42 is aligned perpendicular to the active surfaces 46, 48, as shown in FIG. 2a, but parallel to the active surfaces 46, 48, as shown in FIG. 2b. Or may be aligned at an angle with respect to the active surfaces 46, 48. In the exemplary embodiment, piezoelectric fiber 42 is a lead zirconium titanate (PZT) ceramic fiber made of a relaxor material.

  Methods for making piezoelectric fiber composite materials are known in the art. For example, reference should be made to US Pat. No. 6,620,287, entitled “LARGE-AREA FIBER COMPOSITE WITH HIGH FIBER CONSISTENCY”. This teaching is incorporated herein by reference. Using known manufacturing methods for composite structures, piezoelectric fibers can be integrated into missile components at low cost.

  Piezoelectric fiber 42 generates a current when it is deformed or bent (ie, by missile vibration), and conversely, it bends when exposed to a current or electric field. Electrodes 50 and 52 are configured to sense the electrical signal generated in fiber 42 and further apply the electrical signal from control circuit 32 to fiber 42.

  The control circuit 32 generates an electric actuator signal. An electrical actuator signal is applied to the fiber 42 by the electrodes 50 and 52. In response to the signal, the fiber 42 bends and distorts the structure. Thus, by controlling the voltage of the actuator signal applied to the fiber 42, it is possible to control the stiffness of the structure and further adjust the frequency response of the structure. In addition, the control circuit 32 receives the sensing signal from the fiber 42, modulates the sensing signal to form an actuator signal, and returns the actuator signal to the fiber 42 to attenuate vibrations, thereby activating active vibration damping. Can be configured to provide.

  FIG. 3 is a simplified block diagram of a vibration control circuit 32 designed in accordance with an exemplary embodiment of the present invention. In the exemplary embodiment, control circuit 32 is configured to include a plurality of pre-programmed operating modes. Each mode generates a different actuator signal in response to a mode selection signal provided by the missile guidance system. The mode selection signal indicates which operational phase the missile is in (eg, pre-launch, booster phase, guided flight, etc.).

  Thus, the structural response of the vibration control component can be altered to adapt to various environmental conditions. For example, in certain applications, the guidance system does not take over missile maneuvers until the booster phase is complete. Providing a rigid platform for the IMU and seeker sensor when the guidance system is not controlling maneuvering is as important as protecting the electronics during the booster phase (and also during pre-launch operations) is not. During this period, the control circuit 32 can be configured to generate actuator signals that reduce the stiffness of the fiber 42 and, in particular, attenuate vibrations at frequencies that are harmful to electronic equipment (eg, high frequencies). When the guidance system attempts to take over steering control, the control circuit 32 switches to a “guidance mode” to increase the stiffness of the fiber 42 and provide an actuator signal configured to provide a stable platform. appear. Furthermore, by applying an actuator signal to the component at an appropriate dc voltage level, for example, to avoid mode coupling between structures or to dampen vibrations at frequencies that can adversely affect the guidance system. The frequency response can be controlled.

  In addition, certain events such as stage detachment and divert propulsion thrust can generate large impact loads. Large impact loads can cause uncertain IMU and / or seeker sensor readings. These events are typically very short, on the order of a few milliseconds. During these events, it may be convenient to turn off the guidance system and ignore uncertain readouts. The control circuit 32 can then be switched to a mode configured to reduce these impact loads. After the end of the impact event, the control circuit 32 can be switched back to the guidance mode.

  In the exemplary embodiment shown in FIG. 3, the control circuit 32 includes logic 60 that receives a mode selection signal from the guidance system and sets parameters associated with the selected mode. Load from memory 62. These parameters define which actuator signal should be generated (eg, dc voltage component, how to modulate the sensor signal for active vibration damping, etc.). In the exemplary embodiment, the parameters for each mode are determined during missile testing and then stored in RAM module 62.

  In addition, the control circuit 32 includes logic 64 that receives sensor signals that measure the amplitude and frequency of vibrations in the component and modulates the sensor signals to attenuate the sensed vibrations. To form an actuator signal. The actuator signal may simply be a signal that is out of phase with the sense signal, or it may be a signal configured to focus on damping vibrations in a particular frequency range. The sensor signal can be provided by the piezoelectric fiber 42. When vibration is applied to the piezoelectric fiber 42, the piezoelectric fiber 42 generates an electrical signal. Alternatively, another sensor may be attached to the structure for measuring vibration. Another sensor may also be a piezoelectric sensor.

  In addition, the control circuit 32 includes logic 66, which adds a dc voltage component to the actuator signal. The dc voltage increases or decreases the stiffness of the fiber 42 to control the frequency response of the structure appropriately for the selected mode. The final actuator signal is then applied to the fiber 42.

  The control circuit 32 can be configured to return a fine-tuned excitation signal that is generated to focus on a particular frequency or frequency range to dampen vibrations. In an exemplary embodiment, the compliance of the fiber 42 at high frequencies is increased to isolate high frequency vibrations and protect electronics, while at the same time increasing the stiffness of the fiber 42 at low frequencies to reduce low frequency The control circuit 32 can be configured to provide stiffness and strength performance and return an excitation signal configured to meet the IMU and seeker placement constraints of the guidance system. Further, the excitation signal can be designed to attenuate certain resonant modes, cancel mode coupling phenomena, and attenuate seeker LOS jitter and smearing. In addition, the captive carry load due to the flight environment of the aircraft can be attenuated by tuning the missile components to weaken the fundamental bending mode to suppress vibrations.

  In the preferred embodiment, the control circuit 32 is implemented in a small interlaminated IC chip. The control circuit 32 may be implemented using, for example, a discrete logic circuit, FPGA, ASIC, or the like. Instead, the control circuit 32 can be implemented in software executed by a microprocessor. In addition, other implementations may be used without departing from the teachings of the present invention.

  Since the piezoelectric fiber composite material 30 generates electric pulses by itself during vibration, the control circuit 32 does not require an external power source. However, if a higher power excitation signal is desired, a battery may be added to supply additional power to the control circuit 32.

  Accordingly, the teachings of the present invention provide vibration control that uses a missile component with a piezoelectric fiber composite and controls the piezoelectric fiber composite with an integrated circuit configured to dynamically tune the frequency response of the structure. provide. To change any frequency modulation, the control chip can simply be programmed and easily implemented with optimized tuning of the front fuselage dynamics, so that a wide range of iterations can be achieved, as in the prior art application. Dynamic analysis of the structure is no longer necessary. In a general missile development effort, structural components will be studied through simulation optimization studies, changes in performance characteristics of guidance software and payload hardware, environmental load design development, and test input modifications. The desired frequency performance may vary. This normally required system design has changed in the past. For example, some assemblies have been completely redesigned. The teachings of the present invention do not physically change the structure (as in the prior art), but rather modify the software in the vibration control circuit to shift the frequency coupling performance parameters, thereby changing the structural dynamics of the system. Can be changed.

  This attenuation method can be integrated into the electronic housing of the seeker and guidance system to provide a stable platform for the sensing seeker and IMU equipment while protecting the electronic equipment from high frequency vibrations. In addition, instead of adding heavy structural reinforcements, passive damping mounts (ie rubber mounts or dashpots), or active adjustment mechanisms (eg seeker steering mirrors), this damping method can be applied to the bulkhead and mounting structure. It can be integrated to further damp electronic equipment vibrations and reduce the weight of avionic and seeker housings to achieve the same dynamic performance. Furthermore, integrating piezoelectric fiber composite technology into the missile airframe can improve the structural performance of the airframe and electronically adapt the frequency response of the missile airframe.

  The teachings of the present invention can be applied to any type of missile. 1, 4 and 5 illustrate various exemplary missile designs using vibration control components designed in accordance with the teachings of the present invention. Figures 1a and 1b show designs that may be used for air-to-air or surface-to-air missiles. FIG. 4 shows another design that may be used for air-to-air or surface-to-air missiles, such as Evolved Sea Sparrow Missile (ESSM). Figures 5a-5e illustrate designs that may be used for Kinetic Energy Interceptor (KEI) missiles.

  FIG. 4 is an exploded view of an exemplary missile 10 ′ with vibration control components designed in accordance with another embodiment of the present invention. In this embodiment, the missile 10 'includes a mounting structure 24'. The mounting structure 24 'is an axial beam attached to a missile airframe (not shown). In accordance with the teachings of the present invention, both the mounting beam 24 'and the missile airframe include piezoelectric fiber composite sensors / actuators. A plurality of electronic components are mounted on the mounting beam 24 ', and each electronic component is housed in a vibration control housing 22 containing a piezoelectric fiber composite sensor / actuator. In addition, a seeker housing 20 containing a piezoelectric fiber composite sensor / actuator is mounted on the mounting beam 24 '.

  FIG. 5a is a cross-sectional view of a kinetic energy interceptor (KEI) missile 10 ″ with a vibration control component designed in accordance with another embodiment of the present invention. KEI missiles are configured to intercept enemy missiles during the boost phase prior to ballistic ascent in intermediate orbits where the payload is uncovered and any RV and viable decoy is deployed. ing. In addition, the booster stage of interception is the lowest duty of defense forces deployed in the territory, whether any toxic material dispersed during the interception, whether nuclear weapons, biological weapons, or nerve gas weapons, Suggest returning to the country of origin. For KEI missions, time to target is critical. Therefore, to maximize the agility of interceptor bullets, high performance and lightweight airframe and electronics packaging technologies are required. As shown in FIG. 5a, the KEI missile 10 "includes a two-stage booster 70, a third-stage rocket motor 14", and a shootout 12 ".

  FIG. 5b is a simplified schematic diagram of the grenade 12 '' and rocket motor 14 '' of the KEI missile 10 '' of FIG. 5a. The grenade 12 ″ includes a seeker assembly 16, guidance system electronics 18, and a lateral propulsion system 72. The components of the grenade 12 ″ 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 shootout 12 ″. The grenade 12 ″ includes a lateral propulsion system 72. The lateral propulsion system 72 includes a fluid bottle 82 attached to the mounting structure 24 ″ and a plurality of nozzles 80. The front electronics assembly 18 includes the IMU and guidance system electronics. The front electronics assembly 18 is attached to the front end of the mounting structure 24 ''. The seeker assembly 16 is attached to the front electronics assembly 18. The aft electronics assembly 84 is attached to the rear of the mounting structure 24 ''. The mounting structure 24 ″ is attached to the interstage adapter 74.

  In accordance with the teachings of the present invention, the mounting structure 24 '' and the missile airframe (not shown) each include a piezoelectric fiber composite sensor / actuator 30 and a control circuit 32, wherein the piezoelectric fiber composite Sensor / actuator 30 and control circuit 32 are configured to tune the structural response of the components to provide a stable platform for the seeker and IMU, while damping high frequency vibrations. The front electronics assembly 18 and the tail electronics assembly 84 are each housed in an electronics housing 22. The electronics housing 22 includes a piezoelectric fiber composite sensor / actuator 30 and a control circuit 32 such that the piezoelectric fiber composite sensor / actuator 30 and control circuit 32 attenuate vibrations in the electronics assembly. It is configured. Seeker assembly 16 includes a seeker housing 20. Seeker housing 20 also includes a piezoelectric fiber composite sensor / actuator 30 and control circuit 32 to provide a stable platform for the seeker component while damping vibration. FIG. 5d is a three-dimensional view of a seeker housing 20 designed in accordance with an exemplary embodiment of the present teachings.

The divert thrust force generated by the propulsion system 72 can cause jitter and smear dynamics. Jitter and smear dynamics affect seeker resolution, missile guidance, and maneuvering. FIG. 6a shows a missile bend such that the LOS of the missile is at an angle ΔΘ _S with respect to the missile's hard body line. FIG. 6b is a graph of missile bending angle versus time. Further in accordance with the teachings of the present invention, the vibration control component can be configured to mitigate LOS jitter and smear that occurs during propulsion ignition.

  FIG. 5 e is a three-dimensional view of an exemplary interstage adapter 18. The KEI interstage adapter 74 serves many functions as a relay structure between the grenade 12 ″ and the stack-up of the booster. The interstage adapter 74 is not a large structure, but should be lightweight because the burning rate is very sensitive to the weight of the front of the interceptor. Furthermore, the interstage adapter 74 should be sufficiently strong and rigid so as to eliminate excessive deflection in the grenade sway space. This ensures that the covering nose cone is not impacted.

  In accordance with the teachings of the present invention, the interstage adapter 74 also includes a piezoelectric fiber composite sensor / actuator 30 and a control circuit 32, the piezoelectric fiber composite sensor / actuator 30 and the control circuit 32 being It is configured to attenuate the vibration transmitted to 12 ″ and reduce the severity of the impact and vibration environment for the grenade 12 ″. Furthermore, the resonance characteristics can be adjusted in order to avoid resonant coupling of the grenade / adapter. The adapter structure 74 is electronically tuned to provide sufficient airframe stiffness between the shooter and interceptor booster to allow the IMU function, while the delicate grenade electronics and It is most important to be able to be sufficiently compliant to attenuate high frequency loads so as not to damage the seeker assembly.

  Thus, the present invention has been described herein with reference to a particular embodiment for a particular application. Those skilled in the art having access to the teachings of the present invention will recognize additional modifications, applications, and embodiments within the scope of the present invention.

  Accordingly, the claims are intended to cover any and all such applications, modifications, and embodiments within the scope of the present invention.

  10 ... missile, 12 ... front fuselage assembly, 12 "... grenade, 14, 14" ... missile warhead and / or rocket motor, 16 ... seeker assembly, 18 ... Induction system, 20,22 ... Electronic equipment housing, 24,24 '' ... Mounting structure, 26 ... Missile airframe, 30 ... Sensor / actuator of piezoelectric fiber composite material, 42. ..Piezoelectric fiber, 44 ... Matrix material, 46,48 ... Active surface, 50 ... First electrode, 52 ... Second electrode, 70 ... Booster, 72 ... Horizontal Directional propulsion system, 74 ... Interstage adapter, 80 ... Nozzle, 82 ... Bottle of fluid, 84 ... Electronics assembly.

Claims (40)

A housing structure having a plurality of piezoelectric fibers disposed on or in the upper part;
A control circuit that generates a control signal to electronically tune the structural response of the housing structure in response to a sensor signal provided by the piezoelectric fiber;
A housing comprising:
The piezoelectric fiber is configured to generate the sensor signal in response to deformation in the housing structure;
Furthermore, the piezoelectric fiber is configured to deform the housing structure in response to the control signal that tunes the structural response of the housing structure. The piezoelectric fiber can be configured to increase or decrease the stiffness of the housing structure in response to the control signal;
The at least some of the piezoelectric fibers that generate the sensor signal in response to the deformation include the same piezoelectric fiber as the piezoelectric fiber that deforms the housing structure in response to the control signal. housing.   The housing of claim 1, wherein the control circuit is configured to generate the control signal based on a frequency component of the sensor signal to tune the frequency response of the housing structure to the piezoelectric fiber. The control circuit includes a plurality of operation modes,
Each operating mode is configured to generate a different control signal that provides a different structural response,
The operation mode includes a booster mode and a guidance mode,
During the booster mode, the control circuit is configured to generate a control signal that reduces stiffness by increasing compliance of the piezoelectric fiber to dampen vibration at higher frequencies;
2. The housing of claim 1, wherein during the induction mode, the control circuit is configured to generate a control signal that increases the stiffness of the piezoelectric fiber at a lower frequency.   The housing of claim 1, wherein the piezoelectric fiber is embedded in a composite material attached to the housing structure.   The housing of claim 1, wherein the housing structure is made from a composite material comprising the piezoelectric fiber.   During the booster mode, the control circuit is configured to generate the control signal that attenuates vibration at a higher frequency by modulating the control signal based on vibration sensed by the piezoelectric fiber. The housing of claim 4.   8. The housing of claim 7, wherein the control circuit is configured to switch to the guidance mode immediately before a guidance system takes over steering control. The operation mode further includes a stage separation mode,
9. The housing of claim 8, wherein during the stage disconnect mode, the control circuit is configured to generate a control signal that mitigates shock associated with stage disconnect. In order to provide the sensor signal to the control circuit and provide the control signal to the piezoelectric fiber to the control circuit, the plurality of piezoelectric fibers include one or more electrodes,
The housing of claim 1, wherein the control circuit includes logic to generate the control signal configured to dampen vibrations of the housing structure at a predetermined frequency.   12. The housing of claim 11, wherein the control circuit includes logic for generating the control signal configured to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency.   The housing of claim 12, wherein the control signal is configured to increase compliance of the piezoelectric fiber at high frequencies to attenuate high frequency vibrations to which the devices contained within the housing structure react sensitively.   The control signal is further configured to increase the stiffness of the piezoelectric fiber at low frequencies so that the housing structure provides a stable platform for devices contained within the housing structure. The housing of claim 13.   The housing of claim 11, wherein the piezoelectric fiber is further configured to sense movement in the housing structure and generate a sensor signal in response to the movement.   The housing of claim 15, wherein the control circuit is configured to modulate the sensor signal to generate a control signal configured to attenuate vibrations sensed by the sensor signal. The control circuit includes a plurality of operation modes,
Each operating mode is configured to generate a different control signal that provides a different structural response,
The operation mode includes a booster mode and a guidance mode,
During the booster mode, the control circuit is configured to generate the control signal that reduces stiffness by increasing compliance of the piezoelectric fiber to dampen vibrations at higher frequencies;
12. The housing of claim 11, wherein during the inductive mode, the control circuit is configured to generate the control signal that increases the stiffness of the piezoelectric fiber at a lower frequency. The control circuit includes:
19. The housing of claim 18, comprising means for receiving a signal for selecting one of the operating modes and generating the control signal corresponding to the selected mode in accordance with the signal.   The housing of claim 1, wherein the housing is an electronic device housing.   The housing of claim 1, wherein the housing is a missile airframe. An electronic device housing comprising a housing structure and a control circuit,
The housing structure is made of a composite material including a plurality of piezoelectric fibers;
The plurality of piezoelectric fibers are:
In response to deformation in the housing structure, generating an electrical signal;
In response to an excitation signal applied to the piezoelectric fiber, the housing structure is configured to be deformed,
The control circuit includes:
Receiving the electrical signal from the piezoelectric fiber;
Modulating the electrical signal;
Forming an excitation signal configured to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency to tune the frequency response of the housing structure;
The excitation signal is configured to be applied to the piezoelectric fiber,
The electrical signal includes a frequency component associated with the deformation of the housing structure;
The electronic device housing, wherein the control circuit generates the excitation signal to tune the frequency response of the housing structure based on the frequency component. A missile fuselage comprising a fuselage structure and a control circuit,
The fuselage structure is made of a composite material that includes a plurality of piezoelectric fibers;
The plurality of piezoelectric fibers are:
In response to deformation in the airframe structure, an electrical signal is generated,
In response to an excitation signal applied to the piezoelectric fiber, the airframe structure is configured to be deformed,
The control circuit includes:
Receiving the electrical signal from the piezoelectric fiber;
Modulating the electrical signal;
Forming an excitation signal configured to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency to tune the frequency response of the airframe structure;
The excitation signal is configured to be applied to the piezoelectric fiber,
The electrical signal includes a frequency component related to the deformation of the airframe structure;
The missile airframe, wherein the control circuit generates the excitation signal that tunes the frequency response of the airframe structure based on the frequency component. A mounting structure comprising a mounting structure and a control circuit,
The mounting structure is made of a composite material including a plurality of piezoelectric fibers;
The plurality of piezoelectric fibers are:
In response to deformation in the mounting structure, an electrical signal is generated,
In response to an excitation signal applied to the piezoelectric fiber, the mounting structure is configured to be deformed,
The control circuit includes:
Receiving the electrical signal from the piezoelectric fiber;
Modulating the electrical signal;
Forming an excitation signal configured to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency to tune the frequency response of the mounting structure;
The excitation signal is configured to be applied to the piezoelectric fiber,
The electrical signal includes a frequency component related to the deformation of the mounting structure;
The mounting structure, wherein the control circuit generates the excitation signal that tunes the frequency response of the mounting structure based on the frequency component. A control circuit for controlling vibration in a structure including a piezoelectric fiber,
The piezoelectric fiber is
In response to the deformation in the structure, a sensor signal is generated,
Configured to deform the structure in response to an excitation signal applied to the piezoelectric fiber;
The control circuit includes a first circuit and a second circuit,
The first circuit receives the sensor signal including a frequency component associated with the deformation in the structure;
The second circuit modulates the sensor signal to form an excitation signal configured to electronically tune the structural response of the structure based on the frequency component of the sensor signal;
At least a portion of the piezoelectric fiber that generates the sensor signal in response to the deformation is the same piezoelectric fiber as the piezoelectric fiber that deforms the structure in response to the excitation signal applied to the piezoelectric fiber. circuit. The control circuit includes a plurality of operation modes,
Each mode of operation is configured to generate a different control signal that provides a different structural response,
The operation mode includes a booster mode and a guidance mode,
During the booster mode, the second circuit is configured to generate the excitation signal that reduces stiffness by increasing compliance of the piezoelectric fiber to dampen vibrations at higher frequencies. ,
26. The control circuit of claim 25, wherein during the induction mode, the second circuit is configured to generate the excitation signal that increases the stiffness of the piezoelectric fiber at a lower frequency.   27. The control circuit of claim 26, wherein the control circuit further includes a circuit that receives a signal that selects one of the operating modes. Missile aircraft,
A guidance system to control the missile flight path;
A first housing containing the guidance system;
A control circuit;
Mounting structure,
A missile comprising:
The first housing includes a plurality of piezoelectric fibers,
The plurality of piezoelectric fibers are:
In response to deformation in the first housing, generating a sensor signal;
In response to an excitation signal applied to the plurality of piezoelectric fibers, the first housing is configured to be deformed,
The control circuit includes:
Responsive to a frequency component associated with the deformation of the mounting structure to generate the excitation signal configured to tune the structural response of the first housing;
It is configured to apply the excitation signal to the piezoelectric fiber,
The mounting structure is a missile in which the first housing is mounted on a body of the missile. The control circuit includes a plurality of operation modes,
Each mode of operation is configured to generate a different control signal that provides a different structural response,
The operation mode includes a booster mode and a guidance mode,
During the booster mode, the control circuit is configured to generate an excitation signal that reduces stiffness by increasing compliance of the piezoelectric fiber to dampen vibration at higher frequencies;
30. The missile of claim 28, wherein during the induction mode, the control circuit is configured to generate an excitation signal that increases the stiffness of the piezoelectric fiber at a lower frequency.   30. The missile of claim 29, wherein the control circuit is configured to receive a signal from the guidance system to select one of the operating modes.   30. The missile of claim 28, wherein the control circuit is configured to receive the sensor signal from the piezoelectric fiber and modulate the sensor signal based on the frequency component to form the excitation signal. The missile fuselage includes a second plurality of piezoelectric fibers;
The second plurality of piezoelectric fibers includes:
In response to deformation of the missile body, generating a second sensor signal;
30. The missile of claim 28, configured to deform the missile fuselage in response to a second excitation signal applied to the second plurality of piezoelectric fibers. The mounting structure includes a third plurality of piezoelectric fibers;
The third plurality of piezoelectric fibers includes:
Generating a third sensor signal in response to the deformation in the mounting structure;
33. The missile of claim 32, configured to deform the mounting structure in response to a third excitation signal applied to the third plurality of piezoelectric fibers. The control circuit includes:
Providing an excitation signal to the piezoelectric fiber in the first housing, the missile body, and the mounting structure;
34. The missile of claim 33, configured to synchronize structural responses in the first housing, the missile fuselage, and the mounting structure. The control circuit is configured to provide an excitation signal;
The excitation signal is
Increase the compliance of the piezoelectric fiber at high frequency, insulate high frequency vibration, protect the induction system electronics,
35. The missile of claim 34, configured to increase the stiffness of the piezoelectric fiber at low frequencies to provide a stable platform for the guidance system.   30. The missile of claim 28, wherein the missile further includes a seeker assembly that senses a signal from a missile target. The missile further includes a second housing that houses the seeker assembly;
The second housing includes a plurality of piezoelectric fibers;
The plurality of piezoelectric fibers are:
In response to deformation in the second housing, generating a sensor signal;
40. The missile of claim 36, configured to deform the second housing in response to an excitation signal applied to the plurality of piezoelectric fibers.   38. The missile of claim 37, wherein the control circuit is configured to provide an excitation signal to the second housing.   40. The missile of claim 38, wherein the excitation signal is configured to attenuate line-of-sight jitter and smearing at the seeker assembly. A method for controlling vibration in a missile,
The missile has a piezoelectric fiber integrated into the structural components of the missile;
The method
Receiving a sensor signal from the piezoelectric fiber that measures a change in motion in the structural component;
Forming an excitation signal configured to increase the stiffness or compliance of the piezoelectric fiber at a predetermined frequency to modulate the sensor signal to tune the structural response of the structural component;
Applying the excitation signal to the piezoelectric fiber,
Generating the excitation signal to tune the structural response of the structural component based on a frequency component of the sensor signal. Receiving from the guidance system an operational mode signal indicative of one of a plurality of operational modes comprising at least a booster mode and a guidance mode;
Providing a different excitation signal for each of the operating modes to provide different structural responses;
41. The method of claim 40, further comprising: During the booster mode, the method includes generating the excitation signal to reduce stiffness by increasing compliance of the piezoelectric fiber to dampen vibrations at higher frequencies;
42. The method of claim 41, wherein during the induction mode, the method includes generating the excitation signal that increases the stiffness of the piezoelectric fiber at a lower frequency.

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