ALBERT

Description: This document presents in viewgraph form techniques for the optimal design of aerospace vehicles which take into account vehicle shape, aerodynamic performance, and weight minimization. Methods include multilevel optimization with linear decomposition, multi-constraint optimization, global sensitivity matrix, and calculation of sensitivity derivatives using both first and second order derivatives. The effects of including flexibility in the optimization is discussed.

Description: Future utilization of space will require large space structures in low-Earth and geostationary orbits. Example missions include: Earth observation systems, personal communication systems, space science missions, space processing facilities, etc., requiring large antennas, platforms, and solar arrays. The dimensions of such structures will range from a few meters to possibly hundreds of meters. For reducing the cost of construction, launching, and operating (e.g., energy required for reboosting and control), it will be necessary to make the structure as light as possible. However, reducing structural mass tends to increase the flexibility which would make it more difficult to control with the specified precision in attitude and shape. Therefore, there is a need to develop a methodology for designing space structures which are optimal with respect to both structural design and control design. In the current spacecraft design practice, it is customary to first perform the structural design and then the controller design. However, the structural design and the control design problems are substantially coupled and must be considered concurrently in order to obtain a truly optimal spacecraft design. For example, let C denote the set of the 'control' design variables (e.g., controller gains), and L the set of the 'structural' design variables (e.g., member sizes). If a structural member thickness is changed, the dynamics would change which would then change the control law and the actuator mass. That would, in turn, change the structural model. Thus, the sets C and L depend on each other. Future space structures can be roughly divided into four mission classes. Class 1 missions include flexible spacecraft with no articulated appendages which require fine attitude pointing and vibration suppression (e.g., large space antennas). Class 2 missions consist of flexible spacecraft with articulated multiple payloads, where the requirement is to fine-point the spacecraft and each individual payload while suppressing the elastic motion. Class 3 missions include rapid slewing of spacecraft without appendages, while Class 4 missions include general nonlinear motion of a flexible spacecraft with articulated appendages and robot arms. Class 1 and 2 missions represent linear mathematical modeling and control system design problems (except for actuator and sensor nonlinearities), while Class 3 and 4 missions represent nonlinear problems. The development of an integrated controls/structures design approach for Class 1 missions is addressed. The performance for these missions is usually specified in terms of (1) root mean square (RMS) pointing errors at different locations on the structure, and (2) the rate of decay of the transient response. Both of these performance measures include the contributions of rigid as well as elastic motion.

Description: A fundamental problem facing controls-structures analysts is a means of determining the trade-offs between structural design parameters and control design parameters in meeting some particular performance criteria. Developing a general optimization-based design methodology integrating the disciplines of structural dynamics and controls is a logical approach. The objective of this study is to develop such a method. Classical design methodology involves three phases. The first is structural optimization, wherein structural member sizes are varied to minimize structural mass, subject to open-loop frequency constraints. The next phase integrates control and structure design with control gains as additional design variables. The final phase is analysis of the 'optimal' integrated design phase considering 'real' actuators and 'standard' member sizes. The control gains could be further optimized for fixed structure, and actuator saturation constraints could be imposed. However, such an approach does not take full advantage of opportunities to tailor the structure and control system design as one system.

Description: The JPL Control/Structure Interaction Program is developing new analytical methods for designing micro-precision spacecraft with controlled structures. One of these, the Conceptual Design Tool, will illustrate innovative new approaches to the integration of multi-disciplinary analysis and design methods. The tool will be used to demonstrate homogeneity of presentation, uniform data representation across analytical methods, and integrated systems modeling. The tool differs from current 'integrated systems' that support design teams most notably in its support for the new CSI multi-disciplinary engineer. The design tool will utilize a three dimensional solid model of the spacecraft under design as the central data organization metaphor. Various analytical methods, such as finite element structural analysis, control system analysis, and mechanical configuration layout, will store and retrieve data from a hierarchical, object oriented data structure that supports assemblies of components with associated data and algorithms. In addition to managing numerical model data, the tool will assist the designer in organizing, stating, and tracking system requirements.

Description: Ever since the United States space program started some forty years ago, there have been many ideas on how the U.S. should proceed to explore space. Throughout the years, many innovative designs have surfaced for transfer vehicles, space stations, and surface bases. Usually the difference in designs are due to differences in mission objectives and requirements. The problem for Mars is how to choose an architecture for human travel to Mars and what kind of base construction to design for Mars that will be reliable and cost effective. Eventually, if the Space Exploration Initiative is to become a reality, NASA will have to select and fund a single mission architecture involving manned and unmanned Mars fly-by precursors, a Mars landing vehicle, and, ultimately, the plan for constructing a Mars base. The decision to commit to a single architecture is a vital one and, therefore, the design issues, the decision making process, and the analysis tools must be available to explore all of the options that are available. A large part of any space mission architecture is the Earth-to-Mars transfer vehicle. The decision on the type of transfer vehicle to design is a crucial one. The many options must take into account the constraints encountered when assembling the vehicle in earth orbit such as effective joining methods, test and evaluation methods, preventative maintenance measures, etc. Therefore, the process of trading off various designs must include every facet of that design. The on-orbit assembly/construction constraints will drive designs and architectures. This viewgraph presentation highlights the above critical issues so that designs may be evaluated from these viewpoints. Evaluating designs from the issues contained in this paper will help decision makers detect inadequate designs. Stressing these issues in the evaluation procedure will have a great impact on the decisions of future space mission transfer vehicles and consequent architectures.

Description: A cooperative United States/Japan study was made for one year from 1987 to 1988 regarding the feasibility of the Tropical Rainfall Measuring Mission (TRMM). As part of this study a phase-A-level design of spacecraft for TRMM was developed by NASA/GSFC, and the result was documented in a feasibility study. The phase-A-level design is developed for the TRMM satellite utilizing a multimission spacecraft.

Description: A spacecraft-shielding technique is investigated in which the geometrical configuration and material used are emphasized. Ultrathin spaced shield elements are employed to repeatedly shock the impacting projectile to a high energy state that causes melting and vaporization. The ratio of the thickness of the elements to projectile diameter corresponds to a relatively small percentage of debris-plume mass that can be withstood by the backsheet. The strength of the backsheet is thereby reduced and employed in a specific configuration that prevents the debris plume from destroying successive sheets before the particulates reach the sheet. The primary benefit is weight reduction of 30 percent when compared to a 'Whipple shield' fabricated with the same material. The concept is shown to be effective against all impact types tested and produces minimal secondary debris.

Description: Averaging methods are applied to analyzing the transient response associated with adaptive control of large space structures. Using a dominant mode approximation to the plant, an analytical bound is found on the envelope of the adaptive response, characterizing many of the features of the response useful for control design (e.g., peak values, quadratic costs, settling times, etc.). An optimal adaptive design problem is formulated based on minimizing the product of the settling time and peak torque requirement. The resulting nonlinear constrained optimization problem is solved in closed form, and several properties of the optimal adaptive design are discussed.

Description: An approach to the integrated design of linear controlled structures, which is being investigated as part of NASA's controlled structures interaction (CSI) methodology development program, is presented. The approach is integrated in the sense that the structure and its controller are simultaneously designed. The design methodology uses constrained nonlinear optimization procedures based on analytically obtained gradients of the structural responses. Design of the controller is based on the so-called Q-parameterization theory, which parameterizes all closed-loop input/output maps achievable with stabiizing linear controllers. Very general objective and constraint functions are possible, and structural shape can be included in the design variables. This method has been partially implemented and demonstrated; early findings are reported.

Description: The problem of controller design for flexible spacecraft is addressed. Model-based compensators, which rely on the knowledge of the system parameters to tune the state estimator, are considered. The instability mechanisms resulting from high sensitivity to parameter uncertainties are investigated. Dissipative controllers, which use collocated actuators and sensors, are also considered, and the robustness properties of constant-gain dissipative controllers in the presence of unmodeled elastic-mode dynamics, sensor/actuator nonlinearities, and actuator dynamics are summarized. In order to improve the performance without sacrificing robustness, a class of dissipative dynamic compensators is proposed and is shown to retain robust stability in the presence of second-order actuator dynamics if acceleration feedback is employed. A class of dissipative dynamic controllers is proposed which consists of a low-authority, constant-gain controller and a high-authority dynamic compensator. A procedure for designing an optimal dissipative dynamic compensator is given which minimizes a quadratic performance criterion. Such compensators offer the promise of better performance while still retaining robust stability.