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Experimental investigation of the effects of exposure of insulation materials to ground and space environments. A total of 20 materials were investigated. Mechanical and thermophysical data were obtained after exposures and were compared with control data. The most significant result of this program is the demonstration that the majority of the materials were not substantially changed by the environments.
The thermophysical properties (absorptance and emittance) of five commonly used spacecraft thermal control coatings have been measured in situ after exposure to 1.5 and 3 KeV mercury ions. Sample temperature was varied from 20 to 150 C. The results showed negligible change in emittance but significant degradation of the solar absorptance at fluences of 10 to the 17th power ions/sq cm. Plots of absorptance as a function of ion fluence are presented, along with pre- and post-exposure ex situ spectral absorptance measurements. The widely used Z93 white paint exhibited quantitative degradation from mercury ions similar to that reported by other experimenters for hydrogen and helium ions.
Consideration of those aspects of the radio emission from the moon which are directly important in interpreting the emission observations. Thermophysical models of the moon can be constructed which explain the emission to the accuracy of the measurements. The variations during lunations of the radio brightness temperatures measured in the wavelength range from 1 mm to about 5 cm can be reproduced by homogeneous models with a thermal parameter in the range from 800 to 1200 cgs units, a dielectric constant about equal to 2, and a loss tangent of about 0.008.
Review of the thermophysical properties of granular materials or silicates believed to simulate the lunar surface layer. Emphasis is placed on thermal conductivity data and the effects of material and environmental variables on the thermal conductivity. There are three basic mechanisms of heat transfer in particulate materials: conduction by the gas contained in the void spaces between the particles; conduction within the solid particles and across the interparticle contacts; and thermal radiation within the particles, across the void spaces between particle surfaces, and between void spaces themselves. Gas and solid conduction, thermal radiation, and the interaction between conduction and radiation are considered.
A program phase to characterize advanced composite materials for a large reflector support truss on the ATS F & G spacecraft is described. The selection of a Hercules Incorporated, 2002M graphite fiber reinforced epoxy material was based on criteria of spacecraft system requirements and the potential of this material to meet these requirements. The objective of this phase was to develop materials data required for development, design, fabrication, test, and flight of a graphite-fiber, reinforced-plastic spacecraft structure. Testing within a temperature range from -300 F to +200 F covered the generation of data for physical, mechanical, thermophysical, and space environmental properties for the selected material. Additional testing covered adhesive bonded joint materials within the temperature ranges of the spacecraft environment. Descriptions of the spacecraft, reflector support truss, design, requirements, materials, tests, and developed data are presented.
The thermophysical properties of the fines from the Apollo 12 landing site have been determined as a function of their relevant parameters. These properties include the thermal conductivity, thermal diffusivity, directional reflectance and emittance. The density used was the same as that observed from the returned core-tube samples and so should be close to the true density of the surface layer at the Apollo 12 site. The measured properties are used to calculate the diurnal temperature variation of the moon's surface as well as for several depths below the surface. The maximum surface of 389 K is obtained at lunar noon while the minimum temperature of 86.1 K is obtained at sunrise. It is shown that the most significant effects on temperature, as compared with previous calculations, are caused by using the directional reflectance which controls the amount of solar energy absorption during the day in place of a constant hemispherical reflectance. The results are compared with previous analyses and remote measurements.
Progress of the research accomplished during fiscal year 1972 is reported. The summaries presented include: (1) background analysis and coordination, (2) surface properties of rock in simulated lunar environment, (3) rock failure processes, strength and elastic properties in simulated lunar environment, (4) thermal fragmentation, and thermophysical and optical properties in simulated lunar environment, and (5) use of explosives on the moon.
The thermophysical and mechanical properties of closed pore insulation (CPI) were measured after exposure to 25 simulated reentry thermal cycles. In addition, mechanical properties were obtained at elevated temperatures before and after cycling. The properties of CPI were not compromised by the cycling. High temperature creep studies were done on three CPI compositions (4, 8, and 12 Wt% CoO additive). CPI-4 had the best creep resistance at temperatures up to 1363 K.
A numerical digital computer program was developed to calculate the heat-transfer coefficients for both sides of a one-dimensional finite slab subject to the boundary conditions ascribed to the phase-change coating heat-transfer technique. In a typical tunnel test situation where a thin wing was exposed to heating on both sides, the data reduction procedures for a semi-infinite slab gave heat-transfer coefficients as much as 375 percent too high on the side with the lowest heating. The results from the one-dimensional finite-slab procedure are presented in the form of correction factors to the solution for a semi-infinite slab in terms of parameters normally used with the phase-change heat-transfer technique. These correlations are not restricted to slab thickness or thermophysical properties and are easily used to obtain accurate data on thin model sections.
Two types of melt segregation effects were studied: (1) evaporative segregation, or segregation due to surface evaporation; and (2) freezing segregation, or segregation due to liquid-solid phase transformation. These segregation effects are closely related. In fact, evaporative segregation always precedes freezing segregation to some degree and must often be studied prior to performing meaningful solidification experiments. This is particularly true since evaporation may cause the melt composition, at least at the critical surface regions or layers to be affected manyfold within seconds so that the surface region or layer melting point and other thermophysical properties, nucleation characteristics, base for undercooling, and critical velocity to avoid constitutional supercooling, may be completely unexpected. An important objective was, therefore, to develop the necessary normal evaporation equations for predicting the compositional changes within specified times at temperature and to correlate these equations with actual experimental data collected from the literature.
Novel attachment and strain-measuring techniques were developed to obtain valid mechanical and thermophysical property data to above 1530 K in air on nonmetallic insulators and their coatings. The lever-arm extensometer developed used for measuring thermal and/or mechanical strain. Thermal expansion measurements were made on a tensile load to define their softening points. Creep characteristics of the insulators were monitored at elevated temperatures and found to become significant above 1090 K (1500 F).
The objectives of the experiment are discussed. These objectives were: (1) measure the thermophysical properties of the Martian surface on a global scale; (2) isolate large-scale irregularities in the global picture; and (3) isolate small-scale hot or cold spots on the surface.
Revision of the International Practical Temperature Scale requires that there be changes for all accurately tabulated thermophysical values. Revised reference data for thermocouples have been generated in a program carried out by the National Bureau of Standards. The new reference data reflect not only revisions in the temperature scale, but also slight changes in the materials themselves and improvements in data fitting methods. A new NBS monograph that contains tables, analytic expressions, various approximations, and explanatory text has been prepared. A general discussion of the project and some specific examples are given.
Tests were conducted to determine the effects of electrostatic propulsion beam divergence effects on spacecraft surfaces. The subjects discussed are: (1) sensitive surfaces on the ATS 6 spacecraft, (2) the cesium ion source and testing facility, (3) cesium ion effects on thermophysical properties, and (4) simulated charge-exchange ion exposure. The compatibility of the ATS 6 ion engine experiment with the engineering subsystems and other experiments aboard the ATS 6 spacecraft was analyzed.
The square root of the product of thermophysical properties q, c and k, where p is density, c is specific heat and k is thermal conductivity, is determined directly on a test specimen such as a wind tunnel model. The test specimen and a reference specimen of known specific heat are positioned at a given distance from a heat source. The specimens are provided with a coating, such as a phase change coating, to visually indicate that a given temperature was reached. A shutter interposed between the heat source and the specimens is opened and a motion picture camera is actuated to provide a time record of the heating step. The temperature of the reference specimen is recorded as a function of time. The heat rate to which both the test and reference specimens were subjected is determined from the temperature time response of the reference specimen by the conventional thin-skin calorimeter equation.
Materials and design technology of the all-silica LI-900 rigid surface insulation (RSI) thermal protection system (TPS) concept for the shuttle spacecraft is presented. All results of contract development efforts are documented. Engineering design and analysis of RSI strain arrestor plate material selections, sizing, and weight studies are reported. A shuttle prototype test panel was designed, analyzed, fabricated, and delivered. Thermophysical and mechanical properties of LI-900 were experimentally established and reported. Environmental tests, including simulations of shuttle loads represented by thermal response, turbulent duct, convective cycling, and chemical tolerance tests are described and results reported. Descriptions of material test samples and panels fabricated for testing are included. Descriptions of analytical sizing and design procedures are presented in a manner formulated to allow competent engineering organizations to perform rational design studies. Results of parametric studies involving material and system variables are reported. Material performance and design data are also delineated.
The thermochemical response of Coating 313 when exposed to a fuel fire environment was studied to provide a tool for predicting the reaction time. The existing Aerotherm Charring Material Thermal Response and Ablation (CMA) computer program was modified to treat swelling materials. The modified code is now designated Aerotherm Transient Response of Intumescing Materials (TRIM) code. In addition, thermophysical property data for Coating 313 were analyzed and reduced for use in the TRIM code. An input data sensitivity study was performed, and performance tests of Coating 313/steel substrate models were carried out. The end product is a reliable computational model, the TRIM code, which was thoroughly validated for Coating 313. The tasks reported include: generation of input data, development of swell model and implementation in TRIM code, sensitivity study, acquisition of experimental data, comparisons of predictions with data, and predictions with intermediate insulation.
In extended space missions, foods will be heated to enhance the psychological as well as the physiological well-being of the crew. In the low-gravity space environment natural convection is essentially absent so that the heat transfer within the food is by conduction alone. To prevent boiling in reduced pressure environments the maximum temperature of the heating system is severely limited. The Skylab food-heating system utilizes a tray with receptables for the food containers. The walls of the receptacles are lined with thermally controlled, electrical-resistance, blanket-type heating elements. A finite difference model is employed to perform parametric studies on the food-heating system. The effects on heating time of the (1) thermophysical properties of the food, (2) heater power level, (3) initial food temperatures, (4) container geometry, and (5) heater control temperature are presented graphically. The optimal heater power level and container geometry are determined.