Metal working – Piezoelectric device making
Reexamination Certificate
1998-12-18
2004-05-18
Jordan, Charles T. (Department: 3644)
Metal working
Piezoelectric device making
C029S897200, C029S407050, C029S709000, C073S802000
Reexamination Certificate
active
06735838
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates generally to spacecraft structural elements and, more particularly, to a process of augmenting a spacecraft structure's intrinsic damping to relax stiffness design constraints that arise due to precision pointing requirements and lead to excessive spacecraft structural weight and volume.
2. Discussion of the Related Art
The design of spacecraft structural elements is based on certain criteria. Particularly, certain spacecraft structural elements must be designed to have the strength necessary to survive launch and deployment loads, and to meet the stiffness requirements that provide accurate and stable pointing performance of spacecraft components to meet mission requirements in the presence of on-board and external disturbances. Modern spacecraft structures are generally composite structures that are light weight and are not able to easily dissipate mechanical energy from vibrations. The stiffness requirements for spacecraft structures are determined by many factors, such as jitter suppression of payload that is forced by on-board drive motors such as stepper motors. These motors provide many functions, such as antenna pointing, IR and visible light sensor pointing, solar array drives, as well as many other applications. When the drive motor disturbance frequencies align and couple with spacecraft structural modes, large response amplitudes can result that effect the pointing performance.
The conventional practice to design and develop spacecraft structural elements to meet required mission performance generally focuses on detailed analysis and tests to verify that the structural modes and disturbance sources do not adversely couple. This is especially true in the absence of significant spacecraft structural damping, which is typically the case for modern composite and aluminum honeycomb sandwich spacecraft panel construction, and for graphite or aluminum booms and tubes.
Much of a spacecraft's weight and physical volume is in its structure. Excessive weight and volume limits the ability to store payload and drives the spacecraft to larger, more expensive launch vehicles with larger fairings and throw-weight capacities. Reductions in weight and volume can be-provided by relaxing stiffness design requirements of certain structural elements, while insuring that the necessary strength requirements are met. This typically cannot be accomplished without compromising precision pointing performance because the reduction in stiffness increases the DC (low frequency) disturbance-to-response transfer function magnitudes so that for a given disturbance magnitude, greater pointing and stability errors are produced. These peak responses are the pointing performance design drivers.
The disturbance-to-response peak transfer function magnitude may be reduced by augmenting a spacecraft structure's intrinsic damping. Certain spacecraft structural designs begin by attempting to meet pointing performance by a stiffness driven design initially, and then by adding damping to meet more stringent requirements. By reducing the disturbance peaks, structural load and strain is reduced, helping to meet strength design requirements. Providing piezoelectric sensor and actuator elements embedded within the composite spacecraft structure is an example of an effective way to dampen movements of the structure. Damping can be applied to compensate for vibrational or other loading forces on the structures. U.S. Pat. No. 5,424,596 issued to Mendenhall et al., titled “Activated Structure”, and assigned to the assignee of this application, discloses the use of piezoelectric actuator/sensor elements disposed on a spacecraft structural element that provides this type of damping. Actuator performance is typically reduced in conjunction with stiff structural host members that reduce actuation strain capability and achievable damping performance. Incorporation of actuator elements in structural elements with high stiffness thus prevents the actuators from having a significant effect on reducing peak response levels of the structure. The strength design is then compromised by the increased response level.
It is an object of the present invention to provide a process for an integrated design of a precision pointing spacecraft structure that relaxes the intrinsic stiffness of the structure necessary to meet strength and mission pointing requirements.
SUMMARY OF THE INVENTION
In accordance with the teachings of the present invention, a process for designing spacecraft structural elements is disclosed that includes increasing the spacecraft structure intrinsic damping to relax stiffness design constraints that are necessary for precision pointing requirements. The process includes specifically designing the spacecraft structural elements to have an intrinsic stiffness optimal for damping augmentation but does not meet mission pointing performance requirements. To overcome this deficiency, the structural elements are equipped with strain energy control elements that sense strain in the structural elements from on-board and external disturbances, and provide actuation of the structural elements to counteract the sensed strain. The strain energy control elements can be any suitable control element that senses strain and actuates the structural element, such as piezoelectric or electrostrictive control elements. By reducing the stiffness requirements of the structural elements, the control elements can more readily provide the desired actuation for damping purposes, and thus the weight and volume of the structural elements can be reduced over those known in the art. Controlling the strain energy in the structural elements allows the structural element to be made of materials having higher strength properties, instead of higher stiffness properties, thus allowing the structural element to meet the strength requirements to survive launch and deployment loads.
Additional objects, advantages, and features of the present invention will become apparent from the following description and appended claims, taken in conjunction with the accompanying drawings.
REFERENCES:
patent: 4817768 (1989-04-01), Schumacher
patent: 5305507 (1994-04-01), Dvorsky et al.
patent: 5424596 (1995-06-01), Mendenhall et al.
patent: 5525853 (1996-06-01), Nye et al.
patent: 4008568 (1990-09-01), None
Elliott Lee E.
Triller Michael J.
Jordan Charles T.
Miller John A.
Nguyen T.
Northrop Grumman Corporation
Warn, Burgess & Hoffmann, P.C.
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