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The conceptual design and mass properties are presented for the CLL in viewgraph format. The spacecraft structural mass is given for orbital assembly, thermal insulation, integrated propulsion, power generation, avionics, environment control, and pyrotechnics and landing system. The mass is given of the lander as well as the transfer stage.
Discussed here are the activities of the Research Triangle Institute (RTI) Technology Applications Team for the period 1 October 1990 through 30 September 1991. Topics researched include automated data acquisition and analysis of highway pavement cracking, thermal insulation for refrigerators, the containment of paint removed from steel structures, improved technologies for Kuwait oil well control, sprayed zinc coatings for corrosion control of reinforcing steel in bridges, and the monitoring and life support of medically fragile children in the educational setting.
Relationships defining the ballistic limit of Space Station Freedom's (SSF) dual wall protection systems have been determined. These functions were regressed from empirical data found in Marshall Space Flight Center's (MSFC) Hypervelocity Impact Testing Summary (HITS) for the velocity range between three and seven kilometers per second. A stepwise linear least squares regression was used to determine the coefficients of several expressions that define a ballistic limit surface. Using statistical significance indicators and graphical comparisons to other limit curves, a final set of expressions is recommended for potential use in Probability of No Critical Flaw (PNCF) calculations for Space Station. The three equations listed below represent the mean curves for normal, 45 degree, and 65 degree obliquity ballistic limits, respectively, for a dual wall protection system consisting of a thin 6061-T6 aluminum bumper spaced 4.0 inches from a .125 inches thick 2219-T87 rear wall with multiple layer thermal insulation installed between the two walls. Normal obliquity is d(sub c) = 1.0514 v(exp 0.2983 t(sub 1)(exp 0.5228). Forty-five degree obliquity is d(sub c) = 0.8591 v(exp 0.0428) t(sub 1)(exp 0.2063). Sixty-five degree obliquity is d(sub c) = 0.2824 v(exp 0.1986) t(sub 1)(exp -0.3874). Plots of these curves are provided. A sensitivity study on the effects of using these new equations in the probability of no critical flaw analysis indicated a negligible increase in the performance of the dual wall protection system for SSF over the current baseline. The magnitude of the increase was 0.17 percent over 25 years on the MB-7 configuration run with the Bumper II program code.
The stability of the planar interface of a structureless solid growing from a depositing component dilute in a carrier fluid is studied when the main solute transport mechanism is thermal (Soret) diffusion. A linear stability analysis, carried out in the limit of low growth Peclet number, leads to a dispersion relation which shows that the planar front is unstable either when the thermal diffusion factor of the condensing component is positive and the latent heat release is small or when the thermal diffusion factor is negative and the solid grows over a thermally-insulating substrate. Furthermore, the influence of interfacial energy effects and constitutional supersaturation in the vicinity of the moving interface is analyzed in the limit of very small Schmidt numbers (small solute Fickian diffusion). The analysis is relevant to physical vapor deposition of very massive species on cold surfaces, as in recent experiments of organic solid film growth under microgravity conditions.
An account is given of the advantages of ceramics-based materials systems in thermostructural applications in aircraft propulsion, spacecraft propulsion and power systems, aerospace (transatmospheric) vehicles, and space structures. Prototype component applications are evaluated with a view to further technical requirements entailed by the implementation of such ceramic components. Typical near-term applications are the Space Shuttle's thermal protection system and high temperature aircraft thermal insulators; a major long-term development goal is that of integrally bladed turbine rotors in aircraft gas turbines. The fiber technology required for highly loaded/high temperature operations is not yet mature; use temperatures are restricted to 1100-1300 C.
Analytical and experimental studies were completed to assess the feasibility of using high-temperature heat pipes to cool hypersonic engine components. This new approach involves using heat pipes to transport heat away from the combustor, nozzle, or inlet regions, and to reject it to the environment by thermal radiation from an external heat pipe nacelle. For propulsion systems using heat pipe radiation cooling (HPRC), it is possible to continue to use hydrocarbon fuels into the Mach 4 to Mach 6 speed range, thereby enhancing the economic attractiveness of commercial or military hypersonic flight. In the second-phase feasibility program recently completed, it is found that heat loads produced by considering both convection and radiation heat transfer from the combustion gas can be handled with HPRC design modifications. The application of thermal insulation to ramburner and nozzle walls was also found to reduce the heat load by about one-half and to reduce peak HPRC system temperatures to below 2700 F. In addition, the operation of HPRC at cruise conditions of around Mach 4.5 and at an altitude of 90,000 ft lowers the peak hot-section temperatures to around 2800 F. An HPRC heat pipe was successfully fabricated and tested at Mach 5 conditions of heat flux, heat load, and temperature.
Early lunar technologies will probably use a common lunar material as ore. They will be robust to minor fluctuations in feedstock composition and will not require appreciable feedstock beneficiation such as rock grinding or mineral concentration. Technologies using unprocessed soil and indifferent to its composition will have the advantage. Nevertheless, the size and grade of the ore body must be confirmed for even the most indiscriminate process. Simple uses such as heaping unprocessed lunar soil for thermal insulation or radiation shielding onto a habitat require that we know the depth of the regolith, the size distributions of its soils, the locations of large boulders, and the ease of excavation. Costs of detailed site surveys trade against restrictions on site selection and conservative engineering design to accommodate unknown conditions of a poorly explored site. Given the above considerations, we consider briefly some abundant lunar materials, their proposed uses, and technologies for their preparation, with particular attention to the Taurus-Littrow site.
The Naval Surface Warfare Center has developed processes for the preparation of mullite (3(Al2O3)(dot)2(SiO2)) whiskers and mullite-whisker felt. Three patents on the technology were issued in 1990. The processes are based on chemical reactions between AlF3, Al2O3, and SiO2. The felt is formed in-situ during the processing of shaped powdered precursors. It consists of randomly oriented whiskers which are mutually intergrown forming a rigid structure. The microstructure and properties of the felt and size of the whiskers can be modified by varying the amount of Al2O3 in the starting mixture. Loose mullite whiskers can be used as a reinforcement for polymer-, metal-, and ceramic-matrix composites. The felt can be used as preforms for fabricating composite materials as well as for thermal insulation and high temperature, chemically stable filters for liquids (melts) and gases.
The objective of this experiment is to develop an understanding of the physical processes leading to spacecraft glow phenomena. The emphasis is to be on surface temperature and altitude effects. A complete understanding of the phenomena could be used to accomplish the following: (1) characterize optical instrument backgrounds; (2) provide guidelines for thermal insulations; (3) characterize material selection for flight optics and associated spacecraft; and (4) affect flight-operation altitude selection for relevant missions.
A high power, 28 V, 330 A h, active lithium thionyl chloride battery has been developed for use as main and payload power sources on an expendable launch vehicle. Nine prismatic cells, along with the required electrical components and a built-in heater system, are efficiently packaged resulting in significant weight savings over presently used silver-zinc batteries. The high rate capability is achieved by designing the cells with a large electrochemical surface area and impregnating an electrocatalyst, polymeric phthalocyanine, into the carbon cathodes. Passivation effects are reduced with the addition of sulfur dioxide into the thionyl chloride electrolyte solution. The results of conducting a detailed thermal analysis are utilized to establish the heater design parameters and the thermal insulation requirements of the battery. An analysis of cell internal pressure and vent characteristics clearly illustrates the margins of safety under different operating conditions. Performance of fresh cells is discussed using polarization scan and discharge data at different rates and temperatures. Self-discharge rate is estimated based upon test results on cells after storage. Results of testing a complete prototype battery are described.
Some of the considerations involved in the design of aircraft fuel tanks for liquid hydrogen are discussed herein. Several of the physical properties of metals and thermal insulators in the temperature range from ambient to liquid-hydrogen temperatures are assembled. Calculations based on these properties indicate that it is possible to build a large-size liquid-hydrogen fuel tank which (1) will weigh less then 15 percent of the fuel weight, (2) will have a hydrogen vaporization rate less than 30 percent of the cruise fuel-flow rate, and (3) can be held in a stand-by condition and readied for flight in a short time.
Optoelectronic tool measures gaps between tiles and steps between faces of tiles. Eliminates need for difficult, tedious, time-consuming, and error-prone manual measurements. Automates measurements of steps and gaps to within 0.006 in. and records data automatically. Developed for inspecting thermal-insulation tiles on Space Shuttle. Adaptable to inspection of any other settings where steps and gaps critical.
Tool converts partly formed towpreg into preconsolidated towpreg ribbon. In technique, towpreg coated with thermoplastic powder converted into preconsolidated ribbon without use of plasticizers or solvents. Relying on melt flow and proper application of tension, friction, and cooling, technique ensures both consistent distribution of polymer within, and consistent cross section of ribbon. Ceramic preconsolidation tool includes heated and cooled halves thermally insulated from each other.
A thermal barrier coating system consists of two layers of a zirconia-yttria ceramic. The first layer is applied by low pressure plasma spraying. The second layer is applied by conventional atmospheric pressure plasma spraying. This facilitates the attachment of a durable thermally insulating ceramic coating directly to the surface of a highly oxidation resistant NiAl-based intermetallic alloy after the alloy has been preoxidized to promote the formation of a desirable Al2O3 scale.
ATTAP activities during the past year included test-bed engine design and development, ceramic component design, materials and component characterization, ceramic component process development and fabrication, ceramic component rig testing, and test-bed engine fabrication and testing. Significant technical challenges remain, but all areas exhibited progress. Test-bed engine design and development included engine mechanical design, combustion system design, alternate aerodynamic designs of gasifier scrolls, and engine system integration aimed at upgrading the AGT-5 from a 1038 C (1900 F) metal engine to a durable 1372 C (2500 F) structural ceramic component test-bed engine. ATTAP-defined ceramic and associated ceramic/metal component design activities completed include the ceramic gasifier turbine static structure, the ceramic gasifier turbine rotor, ceramic combustors, the ceramic regenerator disk, the ceramic power turbine rotors, and the ceramic/metal power turbine static structure. The material and component characterization efforts included the testing and evaluation of seven candidate materials and three development components. Ceramic component process development and fabrication proceeded for the gasifier turbine rotor, gasifier turbine scroll, gasifier turbine vanes and vane platform, extruded regenerator disks, and thermal insulation. Component rig activities included the development of both rigs and the necessary test procedures, and conduct of rig testing of the ceramic components and assemblies. Test-bed engine fabrication, testing, and development supported improvements in ceramic component technology that permit the achievement of both program performance and durability goals. Total test time in 1991 amounted to 847 hours, of which 128 hours were engine testing, and 719 were hot rig testing.
A unique plug-type heat flux gage probe was tested in the NASA Ames Research Center 2x9 turbulent flow duct facility. The probe was fabricated by welding a miniature dual active surface heat flux gage body to the end of a hollow metal cylindrical bolt containing a metal inner tube. Cooling air flows through the inner tube, impinges onto the back of the gage body and then flows out through the annulus formed between the inner tube and the hollow bolt wall. Heat flux was generated in the duct facility with a Huels arc heater. The duct had a rectangular cross section and one wall was fabricated from 2.54 centimeter thick thermal insulation rigid surface material mounted onto an aluminum plate. To measure heat flux, the probe was inserted through the plate and insulating materials with the from of the gage located flush with the hot gas-side insulation surface. Absorbed heat fluxes measured with the probe were compared with absorbed heat fluxes measured with six water-cooled reference calorimeters. These calorimeters were located in a water-cooled metal duct wall which was located across from the probe position. Correspondence of transient and steady heat fluxes measured with the reference calorimeters and heat flux gage probe was generally within a satisfactory plus or minus 10 percent. This good correspondence was achieved even though the much cooler probe caused a large surface temperature disruption of 1000K between the metal gage and the insulation. However, this temperature disruption did not seriously effect the accuracy of the heat flux measurement. A current application for dual active surface heat flux gages is for transient and steady absorbed heat flux, surface temperature and heat transfer coefficient measurements on the surface of an oxidizer turbine inlet deflector operating in a space shuttle test bed engine.
Closed-loop servo control subsystem incorporated, as modification, into system controlling flows of two ingredients mixed and sprayed to form thermally insulating foams on large tanks. Provides steady flows at specified rates. Foams produced smoother and of higher quality. Continued use of system results in substantial reduction in cost stemming from close control of application of foam and consequent reduced use of material.
Technique of multivariate quantitative chemical analysis devised for use in determining relative proportions of two components mixed and sprayed together onto object to form thermally insulating foam. Potentially adaptable to other materials, especially in process-monitoring applications in which necessary to know and control critical properties of products via quantitative chemical analyses of products. In addition to chemical composition, also used to determine such physical properties as densities and strengths.