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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.
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A soil impact test was conducted on a 880-pound 2-foot diameter sphere model. The impact area consisted of back filled desert earth and rock. The impact generated a crater 5 feet in diameter by 5 feet deep. It buried itself a total of 15 feet - as measured to the bottom of the model. After impact the containment vessel was pressure checked. No leaks were detected nor cracks observed.
A program was initiated to develop improved types of terminal seals for aerospace Ni-Cd batteries. The approach used has not involved attempts, such as employed elsewhere, to improve the ceramic-to-metal seal that is now extensively employed for this application. Rather the approach has been directed toward the development and evaluation of new types of seals. Of prime interest in this initial investigation has been the Ziegler type of compression seal and in particular the injection molded version developed by the Bell Telephone Laboratories (BTL). A number of these units were designed, fabricated, and evaluated on an accelerated life test under a simulated battery environment. Results have shown that there are no major problems involved in scaling up the BTL small-size (5-amp) seal to a larger size (up to 50-amp) seal suitable for most JPL flight batteries. Five out of five such seals successfully completed over 10 months of continuous thermal cycling without developing any leaks greater than 1.8 x to 10 to the minus 9th power atm-cc-He/s.
An operating test series was applied to two explosive actuated, normally open, titanium valves. There were no failures. Tests included: proof pressure and external leakage test, gross leak test, post actuation leakage test, and burst pressure test.
Four Arde conospheroid metallic diaphragms were tested at NASA to evaluate their capability for use in the orbit adjust propulsion subsystem (OAPS) of the Atmosphere Explorer spacecraft (AE's C, D, and E). The diaphragms will be used for positive propellant expulsion and spacecraft center of mass (c.m.) control. A leak-free cycle life capability of nine reversals was demonstrated. The diaphragms rolled smoothly from ring to ring in a predictable manner on the first reversal. Varying amounts of diaphragm cocking and ring skipping were observed on subsequent reversals. The diaphragm pressure differential did not exceed 10 psid during any reversal. Cycle life capability, reversal mode, and pressure differential were not affected by sudden reversals, environmental tests, or 18,000 partial reversals. An expulsion efficiency of approximately 97 percent was demonstrated. The results of these tests show that metallic diaphragms can be used as an effective means of positive fuel expulsion; however, to achieve spacecraft c.m. control, the diaphragm must not be reversed prior to flight.
The flight performance of the Mariner 71 attitude control subsystem is discussed. Each phase of the mission is delineated and the attitude control subsystem is evaluated within the observed operational environment. Performance anomalies are introduced and discussed within the context of general performance. Problems such as the sun sensor interface incompatibility, gas valve leaks, and scan platform dynamic coupling effects are given analytical considerations.
Three silicide coated columbium alloys and two cobalt alloys were subjected to identical simulated reentry profiling exposures in both static (controlled vacuum leak) and dynamic (hypersonic plasma shear) environments. Primary emphasis in the columbium alloy evaluation was on the Cb752 and C129Y alloys with a lesser amount on FS85. Commercial silicide coatings of the R512E and VH109 formulations were used. The coated specimens were intentionally defected to provide the types of coating flaws that are expected in service. Temperatures were profiled up to peak temperatures of either 2350 F or 2500 F for 15 minutes in each cycle.
The operating characteristics of a bypass control valve seal and bearing life cycle tests are reported. Data from the initial assembly, leak, torque, and deflection tests are included along with the cycle life test results and conclusions. The equipment involved was to be used in the nuclear engine for the rocket vehicles program.
Results are presented of the failure investigation conducted on resistors produced by the Vamistor Divison, Wagner Electric Corporation. This failure investigation included; failure analyses, chemical and metallurgical analyses, failure mechanism studies, seal leak analyses, and nondestructive stress tests. The data, information, conclusions, and recommendation can be helpful in assessing current usage of these resistors.
An analysis of transient pressures in externally pressurized cryogenic hydrogen and oxygen tanks was conducted and the effects of design variables on pressure response determined. The analysis was conducted with a computer program which solves the compressible viscous flow equations in two-dimensional regions representing the tank and external loop. The external loop volume, thermal mass, and heat leak were the dominant design variables affecting the system pressure response. No significant temperature stratification occurred in the fluid contained in the tank.
The NASA/MSFC multilayer diffusion models are discribed which are used in applying meteorological information to the estimation of toxic fuel hazards resulting from the launch of rocket vehicle and from accidental cold spills and leaks of toxic fuels. Background information, definitions of terms, description of the multilayer concept are presented along with formulas for determining the buoyant rise of hot exhaust clouds or plumes from conflagrations, and descriptions of the multilayer diffusion models. A brief description of the computer program is given, and sample problems and their solutions are included. Derivations of the cloud rise formulas, users instructions, and computer program output lists are also included.
A general graded life detection approach ranging from environmental and in situ observations to specific metabolic experiments on Mars is reported. A zero dead volume leak and computer mediated data acquisition system is also discussed. Soil analyses were also conducted.
Thin metallic liners that provide leak-free service in cryogenic propulsion systems are overwrapped with a glass-fiber composite that provides strength and protection from handling damage. The resultant tube is lightweight, strong and has a very low thermal flux. Several styles of tubing ranging from 5 to 38 cm in diameter and up to 305 cm long were fabricated and tested at operating temperatures from 294 to 21 K and operating pressures up to 259 N/sq cm. The primary objective for the smaller sizes was thermal performance optimization of the propulsion system while the primary objective of the larger sizes was weight optimization and to prove fabricability. All major program objectives were met resulting in a design concept that is adaptable to a wide range of aerospace vehicle requirements. Major items of development included: bonding large diameter aluminum end fittings to the thin Inconel liner; fabrication of a 38 cm diameter tube from 0.008 cm thick Inconel; and evaluation of tubing which provides essentially zero quality propellant in a very short period of time resulting in a lower mass of propellant expended in chilldown.
A study program has been conducted to explore new methods of achieving zero leakage, separable and semipermanent, ducting joints for space flight vehicles. The study consisted of a search of literature of existing zero leakage methods, the generation of concepts of new methods of achieving the desired zero leakage criteria and the development of detailed analysis and design of a selected concept. Other techniques of leak detection were explored with a view toward improving this area.
A space simulation procedure was developed for studying rocket plume contamination effects using a 5-lb bipropellant reaction control system thrustor. Vacuum chamber pressures of 0.00003 torr (70 miles altitude) were achieved with the thrustor firing in pulse trains consisting of eight pulses - 50 msec on, 100 msec off, and seven minutes between pulse trains. The final vacuum was achieved by cooling all vacuum chamber surfaces to liquid-helium temperature and by introducing a continuous argon leak of 48 std. cc/sec into the test chamber. Fast time response liquid flow rate measurements showed that large variations occurred in the ratio of oxidizer to fuel flow for pulse-on times up to 120 msec. These variations could lead to poor combustion efficiency and the production of contamination.
Test failures of heat pipes occur when the functional performance is unable to match the expected design limits or when the power applied to the heat pipe (in the form of heat) is distributed unevenly through the system, yielding a large thermal gradient. When a thermal gradient larger than expected is measured, it normally occurs in the evaporator or condenser sections of the pipe. Common causes include evaporator overheating, condenser dropout, noncondensable gas formation, surge and partial recovery of evaporator temperatures, masking of thermal profiles, and simple malfunctions due to leaks and mechanical failures or flaws. Examples of each of these phenomena are described along with corresponding failure analyses and corrective measures.
The pressures indicated by two tubulated ionization gages, one pointing to a spinning spacecraft undergoing thermal vacuum test and the other the walls of the chamber, have been used in a computer program to calculate important parameters of flow kinetics in the vacuum chamber. These parameters calculated as a function of time are: the self-contamination of the spacecraft (defined as the return of outgassed molecules on its critical surfaces either in orbit or while undergoing vacuum test); the spacecraft outgassing including leaks from sealed compartments; and the gas pumping performance of the vacuum chamber. The test indicated the feasibility of this type of evaluation and the improvements in instrumentations and arrangements needed for future tests.
A satellite experiment to test the predicted relativistic precession of a gyroscope moving through a gravitational field has been under development at Stanford University for a number of years. The instruments will be located in a liquid helium bath to insure dimensional stability and to permit using superconducting types of circuitry for readout of the gyro orientation. Heat leaks in the system cause the helium to boil and the resultant helium gas is used for attitude control. The principal subject of the paper is the design and experimental evaluation of an electromagnetically actuated differential thrustor which was built and tested at Stanford University. The results are unique because most propulsion systems operate on-off in order to utilize propellant efficiently. In this case, the gas must flow continually to provide cooling, and the requirements are primarily for low power, small volume, and high reliability.
Four Arde conospheroid metallic diaphragms were tested to evaluate their capability for use in the orbit adjust propulsion subsystem (OAPS) of the Explorer spacecraft. The diaphragms will be used for positive propellant expulsion and spacecraft center of mass control. A leak-free cycle life capability of nine reversals was demonstrated. The diaphragms rolled smoothly from ring to ring in a predictable manner on the first reversal. Varying amounts of diaphragm cocking and ring skipping were observed on subsequent reversals. The diaphragm pressure differential did not exceed 7 N/sq cm during any reversal. Cycle life capability, reversal mode, and pressure differential were not affected by sudden reversals, environmental tests, or 18,000 partial reversals. An expulsion efficiency of approximately 97 percent was demonstrated. The results of these tests show that metallic diaphragms can be used as an effective means of positive fuel expulsion; however, to achieve spacecraft center of mass control, the diaphragm must not be reversed prior to flight.