Abstract:
A process for designing spacecraft structural elements ( 20, 30 ) that increases 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 ( 20, 30 ) to have a stiffness that is intrinsically not suitable to meet mission pointing performance requirements in order to reduce weight and volume. To overcome this deficiency, the structural elements ( 20, 30 ) are equipped with strain energy control elements ( 44 ) that sense strain in the structural elements ( 20, 30 ) from on-board and external disturbances, and provide actuation of the structural elements ( 20, 30 ) to counteract the sensed strain. The strain energy control elements ( 44 ) can be any suitable control element that senses strain and actuates the structural element ( 20, 30 ), such as piezoelectric electric or electrostrictive control elements. By reducing the stiffness requirements of the structural elements ( 20, 30 ), the control elements ( 44 ) can more readily provide a desired actuation for damping purposes in order to meet pointing performance requirements, and thus the weight and volume of the structural elements ( 20, 30 ) can be reduced over those known in the art. Relaxing the stiffness requirements of the structural elements ( 20, 30 ) allows the structural element ( 20, 30 ) to be made of materials having higher strength properties, instead of higher stiffness properties, thus allowing the structural element ( 20, 30 ) to meet the strength requirements to survive launch and deployment loads.

Description:
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&#39;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&#39;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&#39;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. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a perspective plan view of a spacecraft having a specially designed payload structure, according to an embodiment of the present invention; 
     FIG. 2 is a perspective view of a known payload structure that can be attached to the spacecraft shown in FIG. 1; 
     FIG. 3 is a graph showing pointing error, where pointing error response amplitude is on the vertical axis and frequency is on the horizontal axis, of a known spacecraft structural element that is stiffness designed with no controlled damping; 
     FIG. 4 is the graph of FIG. 3 showing the pointing error of a known spacecraft structural element that is stiffness designed with controlled damping augmentation; 
     FIG. 5 is a specially designed payload structure, according to an embodiment of the present invention, that is attached to the spacecraft shown in FIG. 1; 
     FIG. 6 shows a spacecraft structural element incorporating an optimized strain energy design of the invention; 
     FIG. 7 is the graph of FIG. 3 showing the pointing error of a spacecraft structural element of the invention without controlled damping; and 
     FIG. 8 is the graph of FIG. 3 showing the pointing error of a spacecraft structural element of the invention including controlled damping augmentation. 
    
    
     DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     The following description of the preferred embodiments directed to an optimized strain energy structure for a spacecraft is merely exemplary in nature and is in no way intended to limit the invention or its applications or uses. 
     According to the invention, a process is provided to design spacecraft structures to survive launch, deployment and mission pointing requirements by using strain based sensors and actuators, such as piezoelectric elements or electrostrictive elements, and a damping control system to increase spacecraft structure damping, thus allowing a reduction in structural weight and volume of the spacecraft, while at the same time adding robustness to the spacecraft structural design. This is accomplished by using an integrated design process where augmented damping levels are used in the initial structural element sizing. The structural elements are sized to meet the launch load, deployment load, and on-orbit strength requirements, but not to have the pointing stiffness requirements as those in the prior art. The structural elements can be designed to meet the launch, deployment and on-board strength requirements with the damping control system on or off. Maximum damping is introduced into the structure by an optimization of the spacecraft structure strain energy distribution. 
     By designing the structure to be less stiff at target actuator locations, according to the invention, more efficient actuation is achieved. By optimally introducing strain energy into the actuator and sensor target locations, maximum actuation force, and therefore damping may be obtained. Actuator locations are determined as a concurrent outcome of the optimization process. Strain based actuators and sensors are then sized to provide optimal coverage in the increased strain energy target area. Strain energy distributions are optimized using a finite element representation of the structure. Modal strain energy distributions are mapped spatially and cross-sectional properties, wall thicknesses, materials, and other stiffness parameters are optimally adjusted to reduce weight and volume, and to localize strain energy in modes that are coupled to disturbance inputs. 
     This integrated approach is performed by iteration on a cost function that weighs objectives, such as actuation figure of merit and structure weight subject to strength and mission pointing performance constraints. Precision pointing performance is obtained by greatly reducing the structural response peaks from disturbance source locations in conjunction with standard disturbance mitigation techniques as appropriate, such as isolation, wheel or cryo-cooler piston balancing, etc. Also, the strain based sensor/actuator system has a coefficient of thermal expansion optimally matched to the structural element to prevent thermal distortion of the structural baseline that would otherwise degrade pointing performance and stability without thermal control. The synergistic optimization of actuator location and structural strain energy is obtained by reducing tube or panel cross-sections which allows smaller, lighter fittings. The reduction in weight of payload support structure will facilitate smaller drive motors saving additional weight and volume. Tailoring of the strain energy distribution may also be accomplished by utilizing known spacecraft materials with higher strength properties instead of higher stiffness properties, which are usually associated with lower strength. This additionally helps to meet the strength design goals while optimizing damping performance. 
     The spacecraft structural design process of the invention recognizes that reduced weight and volume of the spacecraft structure may be obtained by the relaxation of stiffness design constraints on spacecraft precision pointing platforms while maintaining, or even improving, pointing performance. Advantages include the capacity to carry more payload, carry less fuel for attitude orientation (pointing) or use fuel more conservatively, downsize the launch vehicle required for boost, and to minimize the analysis/design/test cycle time typically required to carefully partition structural modes from harmonic disturbance sources and then validate on-orbit performance. The system may further be used to reduce strength requirements by utilizing the active damping system during launch and deployment. The reduction in peak jitter response in the transfer function will produce initial designs that are more robust to changes that occur later in the design cycle. Careful partitioning of disturbance frequencies and structural modes is no longer critical. This reduces schedule and cost burdens brought on by last minute wholesale redesign. 
     The spacecraft structural design process of the invention can better be understood by viewing FIGS. 1-8. FIG. 1 shows a perspective plan view of a spacecraft  10 , including a spacecraft body  12  and solar array panels  14  attached to the body  12  by booms  16 . An instrument support structure  18  includes three boom tube structural elements  20  connected to the spacecraft body  12  by fittings  22  at one end, and a structural element  24  at an opposite end. A more detailed view of the support structure  18  separated from the spacecraft  10  is shown in FIG. 2. A harmonic drive motor  26  is connected to the support structure fitting element  24 , and an elbow fitting  28  is connected to the drive motor  26 . A boom tube structural element  30  is connected to the fitting  28  at one end, and a payload element  32  is connected to an opposite end of the element  30 . The configuration of the support structure  18  is common in the art, and is intended to represent any type of support structure mounted to a spacecraft for pointing and directing the payload element  32  in a desirable manner. The boom tube elements  20  and  30  can be any boom tube suitable for the particular application of the structure  18 , and can be made of any suitable material or composite consistent with spacecraft structures. The elements  20  and  30  can be cylindrical, rectangular, or any suitable geometric configuration, and can be hollow, thin walled, solid, or formed of a honeycomb structure, as would be well understood to those skilled in the art. 
     The payload element  32  is intended to represent any type of suitable element such as a sensor, antennae, telescope, etc. Actuation of the motor  26  causes the structural element  30  to change position, thus altering the pointing direction of the payload element  32 . As discussed above, the operation of the various drive motors, actuators, etc. necessary for the operation of the spacecraft  10  may cause vibrational modes in the spacecraft  10  that adversely effects the pointing direction of the element  32 . 
     FIG. 3 is a graph with disturbance frequency on the horizontal axis and pointing error response on the vertical axis to illustrate the pointing performance of the element  32  for various structure designs discussed below. A pointing performance goal is shown as a horizontal dotted line  38 , and represents a maximum pointing error response at all frequencies for a specific mission requirement, in that any point below the performance goal line  38  is acceptable for that particular mission. A graph line  40  illustrates the pointing error of the element  32  for the structure  18  of the prior art design over the range of disturbance frequencies that exist on the spacecraft  10 . For higher frequencies, no vibrational disturbance is generated to cause a pointing error. 
     For the known support structure  18 , the design stiffness of the structural elements  20  and  30  provide an acceptable pointing of the element  32 , except in two frequency bands where the disturbance response peaks above the goal line  38 . At these disturbance frequencies, the vibration of the spacecraft couples with the structural elements in the support structure  18  to cause directional pointing errors of element  32  to be unacceptable for that particular mission. These peaks represent the disturbance response peaks referred to above. Therefore, the design of the support structure  18  does not satisfy the entire required pointing performance. 
     As mentioned above, it is known to equip the existing stiff support structures for spacecraft designs with actuator/sensor elements to add damping to the structural elements of the spacecraft, as discussed in the &#39;596 patent. FIG. 4 shows the same graph as in FIG. 3, but overlayed with a graph where actuator/sensor elements are provided on the structural elements of the support structure  18  to provide controlled damping. The pointing error for a “closed-loop” type system of this type that includes the actuator/sensor elements on the structural elements of the support structure  18 , basically follows the same pointing error as the “open-loop” pointing error of the structural elements of the support structure  18  without the actuator/sensor elements, except for a slight reduction at the peak locations. Although there is a decrease in the pointing or peaks at the disturbance frequencies that caused the element  32  to point in an unacceptable direction, this decrease in the peaks is minimal, and does not reduce the amplitude below the performance goal line  38 . Because of the high stiffness design of these structural elements, the actuation of the structural elements is minimal. Therefore, the existing support structure  18  would not significantly benefit from these types of actuator/sensor elements. 
     As discussed above, the present invention proposes first reducing the stiffness of the structural elements of the support structure  18 . Reducing the stiffness of the structural elements can be accomplished in different ways, such as by reducing the structural cross-section and thicknesses of the elements, reducing the fitting sizes, using increased strength materials that typically have reduced stiffness, reducing drive motor size, etc. FIG. 5 shows a support structure  18 ′ incorporating these changes to replace the support structure  18 . In this figure, like elements are labeled with the same reference numeral and a prime. Each of the elements of the support structure  18 ′ are smaller in size than the corresponding element in the support structure  18 . Thus, the elements of the support structure  18 ′ have a reduced stiffness than those elements in the support structure  18 . The stiffness of the structural elements of the support structure  18 ′ would not satisfy the pointing requirements. Also, in certain designs, the strength of the structural elements of the support structure  18 ′ may not satisfy launch and deployment loads. 
     To compensate for the lack of stiffness in the structural elements of the support structure  18 ′, the structural elements  20 ′ and  30 ′ are equipped with a plurality of strain energy control elements  44  optimized for this example application. The strain energy control elements  44  can be any suitable control element that senses the strain in the elements  20 ′ and  30 ′, and provides a compensating, or damping, actuation to the sensed strain that causes the structural elements  20 ′ and  30 ′ to maintain their intended position. For example, the control elements  44  can be piezoelectric ceramics, such as lead zirconate titanate (PZT), or electrostrictive ceramics, such as lead molybdenum niobate (PMN). U.S. Pat. No. 5,305,507 issued to Dvorsky et al., title “Method For Encapsulating A Ceramic Device For Embedding In Composite Structures”, discloses actuator/sensor control elements suitable for this application. The control elements  44  would be electrically connected to a damping control circuit (not shown) that converts the sensed strain in the structural elements to a representative voltage, and provides voltage control signals for actuating the elements in response to the sensed strain. The particular control elements  44  used would be matched to the coefficient of thermal expansion of the structural elements to minimize thermal distortion of the elements  44  causing additional pointing errors. The specific optimization design would determine the size of the control elements  44 , the number of control elements  44 , the location of the control elements  44  on the structural elements  20 ′ and  30 ′, the length of the control elements  44 , whether the control elements are located outside or within the structural elements  20 ′ and  30 ′, etc. The specific type of control element and control scheme for compensation for strain in the structural elements  20 ′ and  30 ′ forms no part of the invention, as such a design would be optimized depending on the particular mission requirements and structural configuration. 
     FIG. 6 shows a spacecraft structural element  46  that includes four strain-energy control elements  48  symmetrically positioned around and outside of the structural element  46  at one end thereof. The structural element  46  is shown relative to a graph that shows the modal strain energy along the length of the element  46 . The control elements  48  are positioned at an end of the structural element  46  where the maximum strain energy would be located, such as where the element  46  is connected to the spacecraft  10 . The control elements  48  introduce bending and axial damping into the modes where disturbances cause structural pointing errors. A graph line  50  shows the energy profile along the length of the structural element  46  for a stiff structural element of the type in the prior art, and a graph line  52  shows the energy profile for an optimized structural element of the invention. As is apparent, strain on the structural element  46  has less energy for the stiff structural element than for the optimized structural element of the invention. Therefore, sensing and actuation of the control elements  48  would have a greater effect on the damping in the structural element  46  for the optimal structural element. 
     FIG. 7 shows the graph of FIG. 3 including a pointing performance graph line  54  representing the pointing performance of the reduced structural elements without using the compensating effect of the control elements  44  and  48 . In other words, because the intrinsic stiffness of the structural elements of the support structure  18 ′ is significantly reduced in the design of the present invention, the pointing performance of the structural elements decreases without the use of the control elements  44  and  48 . However, because the strength of the structural elements can be increased using different, but smaller sized materials, the strength of the structural elements can actually be increased. With these types of structural elements, flexibility has been introduced in the system and DC pointing error from disturbances actually increases, as indicated by the peaks above the goal line  38 . Strength requirements are met for this design, but mission pointing performance goals are not. By providing the damping control circuit that controls the control elements  44  and/or  48  with the reduced size structural elements of structure  18 ′, the damping of the structural elements can be increased and the performance requirements can be met. This is shown in FIG. 8 as pointing performance graph line  56 . 
     The foregoing discussion discloses and describes merely exemplary embodiments of the present invention. One skilled in the art will readily recognize from such discussion, and from the accompanying drawings and claims, that various, changes, modifications and variations can be made therein without departing from the spirit and scope of the invention as defined in the following claims.