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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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The engineering test unit of the electron-proton spectrometer was thermal/vacuum tested. The tests confirmed the thermal analysis, and verified that the thermal design would control the temperatures of the instrument.
A second surface mirror type coating, vapor deposited silver on FEP Teflon, was used as a thermal control surface for one of the experiments aboard the Imp-I spacecraft. This coating was selected to obtain the low operating temperature required for this experiment. Initial flight temperature of this thermal control surface was -70.5 C, very close to the predicted value of -73 C and at a very satisfactory level. Since temperatures within the spacecraft interior are not at this desired low level, the detectors had to be mounted exterior to the spacecraft with a good view of space, preferably in an area shaded from sunlight. When this latter preference proved unobtainable, the detectors were mounted on an aluminum plate located on the exterior of the spacecraft, parallel to the spin axis but rotating about the solar vector. The mounting plate was approximately 6.5 inches by 7.5 inches by 0.125 inches thick. To achieve the desired temperature level with the mounting plate in such a location, the thermal design had to minimize not only the effects of the relatively warm spacecraft environment but also the effects of the incident solar energy.
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.
Combustion instability, which results from a coupling of the combustion process and the fluid dynamics of the engine system, was investigated. The design of devices which reduce coupling (combustion chamber baffles) and devices which increase damping (acoustic absorbers) are described. Included in the discussion are design criteria and recommended practices, structural and mechanical design, thermal control, baffle geometry, baffle/engine interactions, acoustic damping analysis, and absorber configurations.
A unique mechanism developed for conducting friction and wear experiments in orbit is described. The device is capable of testing twelve material samples simultaneously. Parameters considered critical include: power, weight, volume, mounting, cleanliness, and thermal designs. The device performed flawlessly in orbit over an eighteen month period and demonstrated the usefulness of this design for future unmanned spacecraft or shuttle applications.
The structural and thermal design of a hydrogen fueled regeneratively cooled three dimensional fixed geometry scramjet was examined. An exploratory study was conducted at Mach 5.3 in the 7-inch Mach 7 pilot tunnel to investigate the unstart phenomena and to provide the experimental data base required to predict the design pressure loads. The test results indicate that the peak pressures occurred during the transient unstart and not during steady state started or unstarted flow conditions. The local peak pressures can be conservatively predicted by normal shock wave theory as the peak approaches the pressure that would exist behind a stationary normal shock with an upstream Mach number equal to the area weighted local Mach number for the normal started condition.
Helios is the German-American solar probe which comes within 0.31 Astronomical Units of the sun. A special thermal design and a dry lubrication system have resulted in successful operation up to now, with Helios having finished its first orbit around the sun.
The Large Space Telescope will employ a 70 mm format SEC type television camera tube as the data sensor for one or more scientific instruments. The characteristics of this tube are described. The current performance is characterized by an MTF of 50% at 20 line pairs per mm and a signal-to-noise ratio that is about one-half the theoretical based on the square root of the number of photoelectrons. The paper presents some astronomical results and the status of the environmental testing of the tube. Other topics such as radiation sensitivity and thermal design are also discussed.
Thermal Vacuum and Thermal Balance tests were successfully conducted on the high power Communications Technology Satellite (CTS). The mission required testing for two distinct thermal environments; spin stabilized at 60 rpm with a minimum of 55 watts of dissipated power, and three-axis stabilized with as much as a tenfold increase of dissipation. The test program verified the adequacy of the spacecraft thermal design and demonstrated functional operation over a wide temperature range for all of the components. Special test hardware incorporating water-cooling systems was required to accomodate the heatpipes and high RF output power. Objectives, test facilities, special hardware, spacecraft configurations, and results of the tests are discussed. Preliminary flight results and conclusions drawn from the test experience are reported.
The propulsion module comprises six to eight 30-cm thruster and power processing units, a mercury propellant storage and distribution system, a solar array ranging in power from 18 to 25 kW, and the thermal and structure systems required to support the thrust and power subsystems. Launch and on-orbit configurations are presented for both modular approaches. The propulsion module satisfies the thermal design requirements of a multimission set including: Mercury, Saturn, and Jupiter orbiters, a 1-AU solar observatory, and comet and asteroid rendezvous. A detailed mass breakdown and a mass equation relating the total mass to the number of thrusters and solar array power requirement is given for both approaches.
The cryogenic refrigerator thermal design calculations establish design approach and basic sizing of the machine's elements. After the basic design is defined, effort concentrates on matching the thermodynamic design with that of the heat transfer devices (heat exchangers and regenerators). Typically, the heat transfer device configurations and volumes are adjusted to improve their heat transfer and pressure drop characteristics. These adjustments imply that changes be made to the active displaced volumes, compensating for the influence of the heat transfer devices on the thermodynamic processes of the working fluid. Then, once the active volumes are changed, the heat transfer devices require adjustment to account for the variations in flows, pressure levels, and heat loads. This iterative process is continued until the thermodynamic cycle parameters match the design of the heat transfer devices. By examing several matched designs, a near-optimum refrigerator is selected.
Research in computerized structural sizing technology was reviewed. Areas covered include: overall design; structural subassembly design; thermal structures; and stiffened panels. In each case, sample results are presented.
Thirteen experiments have been selected for development for the multidisciplinary Spacelab 2 mission. They include three solar physics instruments and an atmospheric physics investigation mounted on the Instrument Pointing System (IPS), a two-axis gimballed X-ray experiment, an infrared telescope scanned about a single axis, and several fixed instruments. Integration of the IPS with its instruments and integration of the entire experiment complement into a flight mission has been a most demanding, complex task which has evolved as the Shuttle carrier vehicle, Spacelab, the IPS, and the experiments have matured. Mechanical interactions between all of these elements, thermal design restrictions, Shuttle and IPS operational constraints, mission duration, pointed instrument collision avoidance, and many other factors have impacted the scientific and engineering objectives of all of the experiments. Not all of the impacts have been adverse, however, and four types of joint science between previously unrelated experiments have so far been generated. This paper summarizes the current state of mission planning, particularly as it relates to the achievement of the scientific and engineering objectives of the pointed experiments. It discusses some of the compromises that have been made to satisfy the diverse requirements and capabilities of the hardware, the planning to date for joint solar observations from the IPS, and some system problems which remain.
A study was performed of cooling methods for a space-borne, earth observing infrared optical instrument, AMTS. Major requirements on the thermal design are an optics temperature below 200 K, a detector array temperature below 75 K, orbital lifetime of 3 to 5 years, a near polar, sun synchronous orbit with altitude near 800 km. Power dissipation of the detectors is 38 mW, in the optics compartment 1.4 W. Large radiative coolers positioned so as to be shielded from sun, spacecraft and earth result in predicted optics temperature of 156 K and detector temperature of 63 K.
The SCIM (Silicon-Coating-By-Inverted Meniscus) process to produce SOC (silicon-on-ceramic) substrates has been investigated for various growth conditions and substrate velocities (4-30 cm/min). Slotted mullite-based substrates (10-cm-wide by 100-cm-long) have been coated with smooth, continuous silicon layers, with thicknesses in the range from 100 to 300 microns. Thermal stress (which can be a problem at low velocities) is prevented by proper thermal design. The highest SCIM-coated SOC cell efficiencies to date are 7.5% (AM1, AR) as compared to 10.5% (AM1, AR) for dip-coated SOC cells. Substantial improvements in cell efficiency are expected when high purity parts are installed in the SCIM-coater.
The areas of scientific interest addressed by the Ultraviolet Imaging Telescope (UIT) as part of the OSS-3 (Office of Space Science) experiment complement and the use of the Space Shuttle to make possible extended observations impose constraints in the optical, mechanical, and thermal design areas which are significantly more difficult than those encountered with sounding rocket payloads. Yet, in order that the UIT be cost-effective, it has been designed using the technology proven in the prior sounding rocket flights as a basis. The UIT had its beginnings in a series of eleven Goddard Space Flight Center sounding rocket payloads dating from 1973. Optically and mechanically it is derived from the 13-inch telescope of Rocket 26.008. Its detector design is related to that of Rocket 25.017, launched in 1977. Rocket 27.062, to be launched in 1982, will be the second flight of the 15-inch telescope serving as the direct prototype of the UIT. This evolution of the UIT from the rocket payload configuration, the nature of the new design requirements, and the manner in which the requirements were met are discussed.
Pressure measurements on the National Transonic Facility models will be made using electronically scanned pressure instrumentation. The system consists of pressure modules, a pressure calibration standard, and a system controller. The pressure modules, which must be operated above -18 C, will be housed in model-integrated designed thermally controlled containers. The Pathfinder 1 192-channel pressure package is described. Recent and planned developments to reduce the pressure module's volume per channel ratio are discussed, including the 48-channel module and a proposed 32-channel module that would be 2.54 by 2.54 by 8.13 cm. Pressure transducers capable of operating at cryogenic temperatures for dynamic and static pressure measurements are discussed.
A simulated section of the main ring of the Space Telescope, (ST) the part that holds the primary mirror, was manufactured and tested for thermal stability characteristics in a complex structural assembly. The other telescope subassemblies are also mounted on the main ring, which is constructed of a titanium alloy. The entire ring is wrapped in heater elements around the outside diameter to reduce the external temperature gradient effects by two orders of magnitude. A thermal vacuum test was performed using a stainless steel alloy tailored to represent the titanium ring's thermal characteristics. The tests examined the thermal performance and controllability of the test ring, the effects of a 24-hr slew manuever that produced worst-case environment changes, hot and cold simulation of predicted ST orbital thermal environments, and the effects of heater failure. Pointing and thermal control features were verified, but a heater failure on the 'on' side was found to quickly lead to exceeding the thermal design limits.