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This thesis describes the thermal design and thermal testing of the development model radiative cooler for the Composite Infra-Red Spectrometer (CIRS) due for launch on the Cassini spacecraft in 1997. The radiative cooler is used to cool the instrument's Focal Plane Assembly (FPA) to approximately 80K. The FPA holds two arrays of HgCdTe detectors for the mid infra-red spectrometer of the instrument which covers the wavelength range 7μm to 17μm. The FPA is mounted from the optics on a titanium alloy tripod and is cooled conductively by the radiator via a flexible link and a cold finger. A range of thermal models of the system have been developed ranging from a simple, analytical model to a finite difference numerical model. A calorimeter was designed to perform heat leak measurements on samples of Multi- Layer Insulation (MLI) blankets to determine the number and type of shields required for the MLI blanket covering the back of the cooler radiator. A test facility incorporating a vacuum system, a space simulator target, and a simulator for the CIRS instrument was designed and constructed for testing the assembled cooler. Various configurations of the Development Model (DM) CIRS cooler were tested as components became available and the results obtained compared to the thermal model predictions. It was found that the cooler will attain a temperature of 80K in operation, but with less excess cooling power than predicted by the thermal models.

The time optimal slewing problem for flexible spacecraft is considered. We study single-axis rotational maneuvers for a simple flexible system, consisting of a rigid hub with an elastic appendage. The equations of motions are derived by Hamilton’s Principle, and a discrete nonlinear model is obtained by the assumed modes method. The problem is first solved in a discrete linearized space by parameter optimization. Optimality is verified by Pontryagin’s Maximum Principle. The linear solutions are then used to obtain time optimal solutions for the non-linear problem by a multiple-shooting algorithm. Although this approach is applicable to arbitrary boundary conditions, this work is confined, almost exclusively, to rest-to-rest maneuvers. These maneuvers are shown to possess some interesting symmetric and asymptotic properties. The problem is further analyzed in infinite-dimensional space, and the convergence of the finite-dimensional approximations is studied. Finally, a soft version of the time optimal slewing problem is considered, where the control is bounded only by a penalty term in the cost functional. A perturbation technique is applied to further simplify this problem. / Ph. D.

Titanium matrix composites have been extensively evaluated for their potential to replace conventional superalloys in high temperature structural applications, with significant weight-savings while maintaining comparable mechanical properties. The purpose of this investigation is the evaluation of a gamma titanium aluminide alloy with nominal composition Ti-46.5Al-4(Cr,Nb,Ta,B)at.% as a matrix material for use in intermediate temperature applications (400-800㩠in future aerospace transportation systems, as very light-weight structures are needed for cost and weight reduction goals. Mechanical characterization testing was performed over the potential usable temperature range (21-800㩮 Thermal expansion behavior was evaluated, as thermal mismatch of the constituents is an expected problem in composites employing this matrix material. Monotonic testing was conducted on rolled sheet material samples to obtain material properties. The alloy exhibited good strength and stiffness retention at elevated temperatures, as well as improved toughness. Monotonic testing was also conducted on specimens exposed to elevated temperatures to determine the degradation effects of high temperature exposure and oxidation. The exposure did not significantly degrade the alloy properties at elevated temperatures; however, room temperature ductility decreased. Analytical modeling using AGLPLY software was conducted to predict the residual stress state after composite consolidation as well as the potential mechanical behavior of [0]4 laminates with a 㭍ET matrix. Silicon carbide (Ultra-SCS) and alumina (Nextel 610) fibers were selected as potential reinforcing materials for the analysis. High residual stresses were predicted due to the thermal mismatch in the materials. Laminates with Nextel 610 fibers were found to offer the better potential for a composite in this comparison as they provide a better thermal match. Coupons of SCS-6/㭍ET were manufactured with different volume fractions (10% and 20%). Both manufacturing attempts resulted in transverse cracking in the matrix from the residual thermal stress.

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The analytical theory of the “passive thermal design of spacecraft" can be divided into two parts. The first part is concerned with the description of the radiant heat transfer to spacecraft external surfaces. The second part is concerned with calculating temperature over a spacecraft when the radiant heat incident, on the spacecraft's wall, is known. The first part, the calculation of the heat incident on a spacecraft's external surfaces, has been investigated in the literature. References one, two, and three are examples of such papers. Unfortunately, the results of auch papers are either numerical or else too specialized to be of general interest for the analytical study of the thermal design of spacecraft. The second part, the calculation of temperatures over a spacecraft when the incident radiant heat is known, is also dealt with in the literature. References four and five are examples of such papers. The heat flow, occurring in the walls of spacecraft, is nonlinear because of thermal radiation and few exact solutions are known. This problem is usually attacked by "linearizing'' the nonlinear term or by directly employing power aeries. The solution of the nonlinear heat equation by the linearization process is valid only for small temperature variations. When temperature differences are large, the linearized solutions do not properly account for the nonlinear radiation terms and series error can result. When power series are employed directly to solve the nonlinear heat flow equation, the labor required to solve the time dependent problem is generally excessive because the elementary functions cannot be used efficiently. In this thesis, the radiant heat transferred to spacecraft is found by the use of Fourier series. The resulting solutions are simple and are valid for spacecraft of very general geometry. Heat transfer calculation which previously required extensive integration on electronic computers can be calculated by the results of this thesis with only trivial labor. Also, the results have the advantage of being well suited for use in the solution of the nonlinear heat transfer equation. The problem of heat flow including nonlinear radiation is also attached in this thesis. The method of solution used is closely related to the well known method of successive approximations and allows solution of nonlinear equations which do not have the classical “Small perturbation parameter.” Also, the method of solution used makes good use of the elementary functions so that time dependent problems can be solved without excessive labor. The problems solved in this thesis includes: the temperature time history of a body at uniform temperature but exposed to periodic radiative heating, the temperature time history of a body having nonuniform temperatures and exposed to periodic radiative heating, and finally the problem of linear heat flow with nonlinear boundary conditions. In each case it is shown how linearized solutions neglect the important results of nonlinear radiation heat transfer. / Master of Science