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Early design studies showed that the UV spectrometer thermal design margins were very small so that an experimental confirmation of the analytical model would be desirable. At that time the prototype unit was scheduled too far downstream to be of value, so a separate thermal model was built for use in verifying the analytical model.
This manual details the input instructions to the data bank, and explanation of the program and its output. The data bank was developed in satisfaction of two of the study tasks, the equipment thermal requirement catalog and the equipment characteristics and constraints catalog. The data bank contains 109 components within space tug avionics system. Other systems were not included in the data bank due to the available information, however, with some program modification, other systems could be incorporated into the data bank program. The data bank was developed and checked out and is compatible with the Univac 1108, and the CDC 6500 operating systems. The data contained in the data bank is general in content with emphasis on the component thermal design. The data is applicable to any spacecraft program where the components contained in the data bank can be applied in satisfaction of the system and subsystem requirements.
The Space Tug Thermal Control Study contained two tasks associated with the cataloging of equipment thermal requirements, physical characteristics and constraints. In satisfaction of these tasks a Data Bank program was developed to provide a means of standardizing the method of cataloging while using the computer to handle the data and format the data into the desired catalogues. During the course of the study 109 components were catalogued and included in the Data Bank. A standardized method was selected for describing each component. Each subsystem of the Avionics System is described on a subsystem header page which describes the types of components included within the subsystem, the quantity requirements, target weights, target power and relative comments. The individual components listed within each subsystem are not necessarily a complete list of candidate items but do represent several of the presently available components for consideration in a Tug application.
Results are presented of an analysis which compares the performance predictions of a thermal model of a multi-panel modular radiator system with thermal vacuum test data. Comparisons between measured and predicted individual panel outlet temperatures and pressure drops and system outlet temperatures have been made over the full range of heat loads, environments and plumbing arrangements expected for the shuttle radiators. Both two sided and one sided radiation have been included. The model predictions show excellent agreement with the test data for the maximum design conditions of high load and hot environment. Predictions under minimum design conditions of low load-cold environments indicate good agreement with the measured data, but evaluation of low load predictions should consider the possibility of parallel flow instabilities due to main system freezing. Performance predictions under intermediate conditions in which the majority of the flow is not in either the main or prime system are adequate although model improvements in this area may be desired. The primary modeling objective of providing an analytical technique for performance predictions of a multi-panel radiator system under the design conditions has been met.
The results of a thermal-structural test program to verify the performance of a metallic/radiative Thermal Protection System (TPS) under reentry conditions are presented. This TPS panel is suitable for multiple reentry, high L/D space vehicles, such as the NASA space shuttle, having surface temperatures up to 1200 C (2200 F). The TPS panel tested consists of a corrugation-stiffened, beaded-skin TD Ni-20Cr metallic heat shield backed by a flexible fibrous quartz and radiative shield insulative system. Test conditions simulated the critical heating and aerodynamic pressure environments expected during 100 repeated missions of a reentry vehicle. Temperatures were measured during each reentry cycle; heat-shield flatness surveys to measure permanent set of the metallic components were made every 10 cycles. The TPS panel, in spite of localized surface failures, performed its designated function.
The 100 mission lifetime requirement for the space shuttle orbiter vehicle dictates a unique set of requirements for the Thermal Protection System (TPS) thermal and heat flow instrumentation. This paper describes the design and development of such instrumentation with emphasis on assessment of the accuracy of the measurements when the instrumentation is an integral part of the TPS. The temperature and heat flow sensors considered for this application are described and the optimum choices discussed. Installation techniques are explored and the resulting impact on the system error defined.
It was demonstrated that cryogenic propellants can be stored unvented in space long enough to accomplish a Saturn orbiter mission after 1,200-day coast. The thermal design of a hydrogen-fluorine rocket stage was carried out, and the hydrogen tank, its support structure, and thermal protection system were tested in a vacuum chamber. Heat transfer rates of approximately 23 W were measured in tests to simulate the near-Earth portion of the mission. Tests to simulate the majority of the time the vehicle would be in deep space and sun-oriented resulted in a heat transfer rate of 0.11 W.
Nine inch diameter samples of the material which will provide thermal and acoustic protection between the nozzle and outer skirt on the space shuttle solid rocket boosters were subjected to heating tests to determine the porosity of the material and the thermal response to a step change in heating. For the porosity measurements a steady state flow of air at 70 F, 500 F, and 1000 F was passed through a sample of the curtain material and measurements of the flow rates were made at different pressure drops across the sample. For the transient measurements, a sample of the curtain material was subjected to a step change in temperature as air was passed through the sample. Measurements of the heat flow through the sample were made as a function of time after the input of the heat pulse. The sample consisted of three layers of curtain panels. Each panel was made of combinations of quartz and fiberglass cloth between which a fiberfrax filler material had been stitched. The hardware design and test procedures were described. Data are provided in engineering units for the flow conditions and and temperatures at which measurements were conducted.
The siting issues associated with small, dispersed solar thermal power plants for utility/small community applications of less than 10 MWe are reported. Some specific requirements are refered to the first engineering experiment for the Small Power Systems Applications (SPSA) Project. The background for the subsequent issue discussions is provided. The SPSA Project and the requirements for the first engineering experiment are described, and the objectives and scope for the report as a whole. A overview of solar thermal technologies and some technology options are discussed.
The thermal conductivity (TC) of tungsten-fiber-reinforced superalloys was determined for two composite systems by using a thermal conductivity standard from the National Bureau of Standards and a comparator and technique developed for that purpose. The results were compared with TC data for the nickel-base alloy MAR-M200. The technique lends itself to applications involving thin specimens, such as thin-walled turbine blades. The TC's of the composite systems were considerably higher in both the longitudinal and transverse directions than that of the monolithic superalloys used as the matrices.
User instructions are given for performing linear and nonlinear steady state and transient thermal analyses with SPAR thermal analysis processors TGEO, SSTA, and TRTA. It is assumed that the user is familiar with basic SPAR operations and basic heat transfer theory.
The preliminary design for a prototype small (20 kWe) solar thermal electric generating unit was completed, consisting of several subsystems. The concentrator and the receiver collect solar energy and a thermal buffer storage with a transport system is used to provide a partially smoothed heat input to the Stirling engine. A fossil-fuel combustor is included in the receiver designs to permit operation with partial or no solar insolation (hybrid). The engine converts the heat input into mechanical action that powers a generator. To obtain electric power on a large scale, multiple solar modules will be required to operate in parallel. The small solar electric power plant used as a baseline design will provide electricity at remote sites and small communities.
A study was made of the comparative life of plasma sprayed ZrO2-Y2O3 thermal barrier coatings on NiCrAlY bond coats on Rene 41 in short (4 min) and long (57 min) thermal cycles to 1040 C in a 0.3 Mach flame. Short cycles greatly reduced the life of the ceramic coating in terms of time at temperature as compared to longer cycles. Appearance of the failed coating indicated compressive failure. Failure occurred at the bond coat-ceramic coat junction. At heating rates greater than 550 kw/sq m, the calculated coating detachment stress was in the range of literature values of coating adhesive/cohesive strength. Methods are discussed for decreasing the effect of high heating rate by avoiding compressive stress.
The activities of the Point-Focusing Thermal and Electric Applications (PETEA) project for the fiscal year 1979 are summarized. The main thrust of the PFTEA Project, the small community solar thermal power experiment, was completed. Concept definition studies included a small central receiver approach, a point-focusing distributed receiver system with central power generation, and a point-focusing distributed receiver concept with distributed power generation. The first experiment in the Isolated Application Series was initiated. Planning for the third engineering experiment series, which addresses the industrial market sector, was also initiated. In addition to the experiment-related activities, several contracts to industry were let and studies were conducted to explore the market potential for point-focusing distributed receiver (PFDR) systems. System analysis studies were completed that looked at PFDR technology relative to other small power system technology candidates for the utility market sector.
An analysis of thermal cycle life data for 22 thermal barrier coated (TBC) specimens was conducted. The Zr02-8Y203/NiCrAlY plasma spray coated Rene 41 rods were tested in a Mach 0.3 Jet A/air burner flame. All specimens were subjected to the same coating and subsequent test procedures in an effort to control three parametric groups; material properties, geometry and heat flux. Statistically, the data sample space had a mean of 1330 cycles with a standard deviation of 520 cycles. The data were described by normal or log-normal distributions, but other models could also apply; the sample size must be increased to clearly delineate a statistical failure model. The statistical methods were also applied to adhesive/cohesive strength data for 20 TBC discs of the same composition, with similar results. The sample space had a mean of 9 MPa with a standard deviation of 4.2 MPa.
The reported study had the purpose to examine variations in cyclic life and in adhesive/cohesive coating strength. Possible effects of heating rate, stress reversal, temperature level, and ceramic deposition methods on coating life were also investigated. Life cycle data for 22 thermal barrier coated rods were examined and found to be statistically modeled by normal or log-normal distributions. The sample mean was 1330 cycles with a standard deviation of 520 cycles. Adhesive/cohesive pulloff strength data for 20 thermal barrier coated flat head piston specimens were taken. The average pulloff stress was 9 MPa with a standard deviation of 4.2 MPa. It was found that variations in heating rate can produce significant variations in the life cycle data.
Because the OSS-1 pallet contained a variety of instruments with irregular surface geometry and properties which limited predictability, the total absorbed flux on thermal canister radiators was measured to determine heat rejection capability. The instrumentation and sensor cup design and locations for the thermal canister experiment are illustrated. Graphs show flux sensor history in hot and coal orbits. Kapton erosion is also considered. Results show that the flux levels measured in all STS attitudes are higher than predictions. In cold and moderate attitudes, flight results are a factor of two to three higher than predicts. In hot attitude, much better agreement occurred. It is concluded that in cold or moderate attitudes other sources may be contributing to added inputs (albedo, Earth shine, shuttle background). In hot attitude, smaller differences could be attributed to coatings assumptions or calculation uncertainty.
Solar absorptance and thermal emittance of spacecraft materials are critical parameters in determining spacecraft temperature control. Because thickness, surface preparation, coatings formulation, manufacturing techniques, etc. affect these parameters, it is usually necessary to measure the absorptance and emittance of materials before they are used. Absorptance and emittance data for many common types of thermal control coatings, are together with some sample spectral data curves of absorptance. In some cases for which ultraviolet and particle radiation data are available, the degraded absorptance and emittance values are also listed.