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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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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1973

Thermal engineering research

A theoretical investigation of gas flow inside a multilayer insulation system has been made for the case of the broadside pumping process. A set of simultaneous first-order differential equations for the temperature and pressure of the gas mixture was obtained by considering the diffusion mechanism of the gas molecules through the perforations on the insulation layers. A modified Runge-Kutta method was used for numerical experiment. The numerical stability problem was investigated. It has been shown that when the relaxation time is small compared with the time period over which the gas properties change appreciably, the set of differential equations can be replaced by a set of algebraic equations for solution. Numerical examples were given, and comparisons with experimental data were made.

Shih, C. C.↗
NASA NTRS1974

Development of a reusable flightweight cryogenic storage system

A goldized-Kapton Superfloc multilayer insulation (MLI) coupled with an MLI purge and repressurization subsystem has been developed for use on reusable space vehicles. The subsystems were installed on an 87-inch diameter oblate spheroid liquid hydrogen tank and subjected to repeated space shuttle flight simulations to determine the effect upon system performance.

Walburn, A. B.↗
NASA NTRS1974

An evaluation of Orbital Workshop passive thermal control surfaces

The optical properties of selected Orbital Workshop thermal control surfaces are discussed from the time of their installation through the end of the Skylab missions. The surfaces considered are the goldized Kapton tape on the habitation area sidewall, the S-13G white paint on the Workshop aft skirt, and the multilayer insulation system on the forward dome of the habitation area. A quantitative assessment of the effects of exposure to the ascent and orbital environments is made including the effects of rocket exhaust plume contamination. Although optical property degradation of the external surfaces was noted, satisfactory thermal performance was maintained throughout the Skylab missions.

Daniels, D. J.↗
NASA NTRS1974

Cryogenic thermal control technology summaries

A summarization and categorization is presented of the pertinent literature associated with cryogenic thermal control technology having potential application to in-orbit fluid transfer systems and/or associated space storage. Initially, a literature search was conducted to obtain pertinent documents for review. Reports determined to be of primary significance were summarized in detail. Each summary, where applicable, consists of; (1) report identification, (2) objective(s) of the work, (3) description of pertinent work performed, (4)major results, and (5) comments of the reviewer (GD/C). Specific areas covered are; (1) multilayer insulation of storage tanks with and without vacuum jacketing, (2) other insulation such as foams, shadow shields, microspheres, honeycomb, vent cooling and composites, (3) vacuum jacketed and composite fluid lines, and (4) low conductive tank supports and insulation penetrations. Reports which were reviewed and not summarized, along with reasons for not summarizing, are also listed.

Stark, J. A.↗
NASA NTRS1975

Design of a heat pipe governed thermal control system for the Solar Electric Propulsion Stage /SEPS/

A 2200-w capacity spacecraft heat rejection system designed for the SEPS and utilizing heat pipe radiator panels has been investigated. The total thermal control system consists of two radiator panels connected to the heat source by variable conductance heat pipes (VCHP's). The system was designed to operate in the 223 to 333 temperature range. The radiators have an emittance of 0.88 at their operational temperature and a fin efficiency of approximately 80 percent. The radiators are thermally isolated from the SEPS and environment by multilayer insulation and thermal shields. Butane was selected as the working fluid for the VCHP because of its low freezing point (135), which is necessary to prevent diffusion freezeout of the liquid during the cold outbond missions. Helium was selected for the control gas. This paper describes the VCHP system, discusses the system design parameters and presents the results of the analyses.

Ruttner, L. E.↗
NASA NTRS1976

Environmental charging of spacecraft-tests of thermal control materials for use on the global positioning system flight space vehicle. Part 2: Specimen 6 to 9

The NASA/USAF program on the Environmental Charging of Spacecraft Surfaces consists, in part, of experimental efforts directed toward evaluating the response of materials to the environmental charged particle flux. Samples of thermal blankets of the type to be used on the Global Positioning System Flight Space Vehicles were tested to determine their response to electron flux. The primary result observed was that no discharges were obtained with the quartz-fiber-fabric-covered multilayer insulation specimen. The taped aluminized polyester grounding system used on all specimens did not appear to grossly deteriorate with time; however, the specimens require specific external pressure to maintain constant grounding system resistance.

Stevens, N. J.↗
NASA NTRS1976

Thermal control of the International Ultraviolet Explorer

The International Ultraviolet Explorer (IUE) is a large astronomical observatory scheduled to be placed in a three-axis stabilized synchronous orbit in the fourth quarter of 1977. The thermal control system consists of multilayer insulation, ammonia-filled grooved heat pipes, bimetallic actuated louvers, and assorted commandable heaters. The spacecraft must operate over a 135-deg solar aspect range and must be able to survive an 85-min.-zero power eclipse. It has a design lifetime of three years. To verify the thermal design, an Engineering Test Unit (ETU) was subjected to a thermal balance test in the Solar Environment Simulator (SES) at the Goddard Space Flight Center (GSFC). Heater skins were utilized to simulate four solar aspect angles, and various internal power settings were used to simulate different operational cases. Temperatures obtained corresponded within 5C of predicted values, thus verifying the thermal analytical model. In addition, two 85-min.-eclipse periods were successfully completed verifying that the thermal design of the IUE was adequate for this mission requirement.

Skladany, J. T.↗
NASA NTRS1977

Thermal design of the IUE hydrazine auxiliary propulsion system

The International Ultraviolet Explorer is a large astronomical observatory scheduled to be placed in a three-axis stabilized synchronous orbit in the fourth quarter of 1977. The Hydrazine Auxiliary Propulsion System (HAPS) must perform a number of spacecraft maneuvers to achieve a successful mission. This paper describes the thermal design which accomplishes temperature control between 5 and 65 C for all orbital conditions by utilizing multilayer insulation and commandable component heaters. A primary design criteria was the minimization of spacecraft power by the selective use of the solar environment. The thermal design was carefully assessed and verified in both spacecraft thermal balance and subsystem solar simulation testing.

Joseph T. Skladany↗
NASA NTRS1979

OTV orbital tanking systems

Orbital transfer of cryogenic propellants could benefit spacecraft and Orbital Transfer Vehicle (OTV) missions in the 1980s by supplying main propulsion, attitude control, or other fluid systems. The Space Shuttle can operate as a tanker when equipped with cryogenic propellant storage and orbital transfer systems. The key technologies are multilayer insulation, capillary propellant acquisition, zero-g gaging, orbital chilldown, and possibly large flight weight dewars. The technologies and operations could be realistically demonstrated using a Centaur that has been integrated with the Shuttle. Orbital refueling capability can enhance the usefulness of the whole Shuttle program

Heald, D. A.↗
NASA NTRS1981

Propellant management for low thrust chemical propulsion systems

Low-thrust chemical propulsion systems (LTPS) will be required for orbital transfer of large space systems (LSS). The work reported in this paper was conducted to determine the propellant requirements, preferred propellant management technique, and propulsion system sizes for the LTPS. Propellants were liquid oxygen (LO2) combined with liquid hydrogen (LH2), liquid methane or kerosene. Thrust levels of 100, 500, and 1000 lbf were combined with 1, 4, and 8 perigee burns for transfer from low earth orbit to geosynchronous earth orbit. This matrix of systems was evaluated with a multilayer insulation (MLI) or a spray-on-foam insulation. Vehicle sizing results indicate that a toroidal tank configuration is needed for the LO2/LH2 system. Multiple perigee burns and MLI allow far superior LSS payload capability. Propellant settling, combined with a single screen device, was found to be the lightest and least complex propellant management technique.

Hamlyn, K. M.↗
NASA NTRS1981

Feasibility study for long lifetime helium dewar

A feasible concept for a launchable three year lifetime helium dewar was investigted. Current helium dewar designs were examined to see where the largest potential reductions in parasitic heat loads can be made. The study was also devoted to examining support concepts. The support concept chosen, a passive orbital disconnect strut (PODS), has an orbital support conductance that is lower by more than an order of magnitude over current tension band supports. This lower support conductance cuts the total dewar weight in half for the same three year life time requirements. Effort was also concentrated on efficient wire feed through designs and vapor cooling of the multilayer insulation, supports, wire feed throughs and plumbing penetrations. A single stage helium dewar vs. dual stage dewars with a guard cryogen of nitrogen or neon was examined. The single stage dewar concept was selected. Different support concepts were analyzed from which the PODS support concepts was chosen. A preliminary design of the dewar was thermally and structurally analyzed and laid out including system weights, thermal performance and performance sensitivities.

Parmley, R. T.↗
NASA NTRS1981

Design study of imaging techniques for the Starprobe mission

A 10 cm aperture, off-axis aplanatic gregorian telescope is investigated as a candidate instrument for imaging from Starprobe in the 115 nm to 900 nm wavelength range. Focal lengths of 300 cm (f/30) and 150 cm (f/15) are compared in terms of optical performance, compatibility with the Starprobe spacecraft and response to the thermal environment in a 20-hour time interval, centered on perihelion. Ultraviolet/visible imaging from Starprobe concluded should be technically feasible. A 10 cm aperture telescope with focal length close to 300 cm is recommended for further study. This instrument should be capable of a detector-limited angular resoluton of about 2 arc sec. The estimated overall dimensions should not exceed 105 cm by 40 cm by 25 cm and its mass be less than 28 kg. The recommended concept assumes solid blank ULE mirrors, a graphite-epoxy optical bench, and an aluminum thermal enclosure, covered with multilayer insulation. These materials must be carefully selected to minimize the risk of UV photopolymerization of contaminants at the mirror surfaces during observtions.

Bottema, M.↗
NASA NTRS1982

Transit Thermal Control Design for Galileo Entry Probe for Planet Jupiter

A totally passive design was completed for the thermal control of the Galileo entry probe during its transit to the planet Jupiter. The design utilizes radio isotope heater units, multilayer insulation blankets and a thermal radiator in conjunction with a design conductance support structure to achieve both the required storage and critical initial planet atmosphere entry temperatures. The probe transit thermal design was completed and verified based on thermal vacuum testing of a prototype probe thermal test model.

Haverly, George C.↗
NASA NTRS1982

Contaminant characterization of five satellite materials

An extensive laboratory test program was performed to characterize outgassing of five satellite materials. The materials were Chemglaze Z-306 over 9922 primer, M-773 adhesive, multilayer insulation, polyurethane foam, and silverized Teflon. The souces were prepared to the specifications of a typical satellite program. Dynamic thermogravimetric mass loss characteristics of these five materials were obtained in vacuum with a beam microbalance. The temperature of the material was linearly raised from 25 C to over 650 C while monitoring the mass loss, rate of mass loss, temperature, and the composition of the outgassed material by residual gas analysis. Isothermal source emission/capture coefficients/reemission parameters were obtained with an array of quartz crystal microbalances (QCM). A detailed test matrix in which the temperatures of the QCMs (-160 C, - 100 C, -40 C, and +10 C) and the source material (125 C, and 90 C, and 50 C) were varied was performed. The array of QCMs was also used to measure the spatial distribution of the source emission. The effects of vacuum ultraviolet radiation on the deposition and reemission parameters was determined.

Muscari, J. A.↗
NASA NTRS1984

Advanced passive radiator for spaceborne cryogenic cooling

A novel design to improve the cooling capability of spaceborne cryogenic passive radiators is described. The design is based on the use of lightweight angled radiation shields, low-conductance structural supports, and a separate detachable launch-support system to reduce the parasitic heat leaks from the warm spacecraft to the cold radiator. The effectiveness of this design is demonstrated by thermal-vacuum-chamber experiments which indicate that the angled-radiation-shield assembly has an effective emittance that is an order of magnitude lower than that of the best multilayer insulation used in flight. Performance predictions based on the experiments and analytical model indicate that an advanced passive-radiator design based on this technology would be between 10 and 57% of the size and 10 and 36% of the mass of the best state-of-the-art passive radiators and would make lower temperatures (< 60 K) and larger heat loads practical. The cooling requirements of many new spaceborne instruments could be accommodated by application of this new passive-radiator design.

Steven Bard↗
NASA NTRS1986

Hubble Space Telescope meteoroid-debris protection analysis

A system level failure could occur if the Hubble Space Telescope's (ST) capability to operate as a facility on-orbit is critically reduced or when a significant reduction in the quality of science data is registered. Failure could occur if a meteoroid/debris impact damages a component of a major support subsystem or if a meteoroid/debris penetration causes straylight contamination in the light shield, forward shell, aft shroud, or through the aperture door. The ST was analyzed to find the probability of no critical penetration. This probability value was found to be 92.25% for a two-year service life. A straylight leakage repair technique was recommended for the aft shroud, the region found most likely to be critically penetrated. Fozar tape and multilayer insulation blankets are suggested as posible repair materials.

Horn, J.↗
NASA NTRS1985

Thermal analysis of the IUE - Seven years later

The International Ultraviolet Explorer (IUE) is a large astronomical observatory which was launched on January 26, 1978. The observatory was placed into a three-axis stabilized, eccentric synchronous orbit with an inclination of about 29 degrees. It was designed for a three year lifetime, but is currently still operational. This paper will briefly review the thermal design which consists of multilayer insulation, ammonia-filled grooved heat pipes, bimetallic actuated louvers, and assorted commandable heaters. It will then discuss the seven years of thermal data and compare the results with prelaunch predictions. Finally an attempt will be made to explain the unexpected and continuing rise in temperature of the hydrazine auxiliary propulsion system which has increased over twenty degrees centigrade since launch.

Skladany, J. T.↗