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  • Other Sources  (918)
  • SPACECRAFT DESIGN, TESTING AND PERFORMANCE  (918)
  • 1975-1979  (918)
  • 1
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    Space shuttle (1975)
    In:  CASI
    Details
    Publication Date: 2006-04-10
    Description: The transportation cost associated with space materials processing were studied to determine the feasibility of space manufacturing. The assumptions use to determining the cost estimates for a 12 year planning period are listed, and the orbitor is described in terms of payloads. Shuttle operational costs, ground operational costs, and nonrecurring investment and development costs were analyzed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Econ. Analysis of Materials Processing in Space; p 8-39
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  • 2
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    Design development of the Apollo command and service module thrust vector attitude control systems (1978)
    In:  CASI
    Details
    Publication Date: 2006-03-02
    Description: Development of the Apollo thrust vector control digital autopilot (TVC DAP) was summarized. This is the control system that provided pitch and yaw attitude control during velocity change maneuvers using the main rocket engine on the Apollo service module. A list of ten primary functional requirements for this control system are presented, each being subordinate to a more general requirement appearing earlier on the list. Development process functions were then identified and the essential information flow paths were explored. This provided some visibility into the particular NASA/contractor interface, as well as relationships between the many individual activities.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA, Washington Fourth Inter-Center Control Systems Conf.; p 63-82
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  • 3
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    Cryogenic helium 2 systems for space applications (1978)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Two cryogenic systems are described which will provide cooling for experiments to be flown on Spacelab 2 in the early 1980's. The first system cools a scanning infrared telescope by the transfer of cold helium gas from a separate superfluid helium storage dewar. The flexible design permits the helium storage dewar and transfer assembly to be designed independent of the infrared experiment. Where possible, modified commerical apparatus is used. The second cryogenic system utilizes a specially designed superfluid dewar in which a superfluid helium experiment chamber is immersed. Each dewar system employs a porous plug as a phase separator to hold the liquid helium within the dewar and provide cold gas to a vent line. To maintain the low vapor pressure of the superfluid, each system requires nearly continuous prelaunch vacuum pump service, and each will vent to space during the Spacelab 2 flight.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Thermal and Environ. Control Systems; p 507-515
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  • 4
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    Experimental control requirements for life sciences (1978)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The Life Sciences dedicated Spacelab will enable scientists to test hypotheses in various disciplines. Building upon experience gained in mission simulations, orbital flight test experiments, and the first three Spacelab missions, NASA will be able to progressively develop the engineering and management capabilities necessary for the first Life Sciences Spacelab. Development of experiments for these missions will require implementation of life-support systems not previously flown in space. Plant growth chambers, animal holding facilities, aquatic specimen life-support systems, and centrifuge-mounted specimen holding units are examples of systems currently being designed and fabricated for flight.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Thermal and Environ. Control Systems; p 467-472
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  • 5
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    Thermal integration of Spacelab experiments (1978)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The method of thermally integrating the experiments for Spacelab is discussed. The scientific payload consists of a combination of European and United States sponsored experiments located in the module as well as on a single Spacelab pallet. The thermal integration must result in accomodating the individual experiment requirements as well as ensuring that the total payload is within the Spacelab Environmental Control System (ECS) resource capability. An integrated thermal/ECS analysis of the module and pallet is performed in concert with the mission timeline to ensure that the agreed upon experiment requirements are accommodated and to ensure the total payload is within the Spacelab ECS resources.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Thermal and Environ. Control Systems; p 545-559
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  • 6
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    An assessment of thermal design procedures for Shuttle payloads (1978)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A review of the thermal design options for unmanned Shuttle payloads is presented. Because many future Shuttle payloads will have budgets and less time available for the definition of their thermal subsystems, simplified design procedures will be the most cost-effective. The thermal interface between an individual payload and the Orbiter is discussed. A simplified, modular thermal control system that will reduce the cost of providing thermal protection by minimizing both the initial procurement cost of any specialized hardware and the need for a detailed thermal interface analysis is described.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Thermal and Environ. Control Systems; p 481-486
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  • 7
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    Further development of the multipactor discharge electron source (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A modification to the multipactor electron source used in the study of the behavior of spacecraft insulating materials is described. The electron accelerating system was included within the source, freeing the region between the source and the target of fields produced by the source electrodes.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 881-886
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  • 8
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    The Qualification of a large electron irradiation facility for telecommunication satellite differential charging simulation (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The definition and procurement program is presented for the set up of a test simulating an electron environment on a model of a communication satellite to study the degradation of the thermal insulation materials during electron bombardment.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 819-833
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  • 9
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    Secondary electron effects on spacecraft charging (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Calculations were conducted to determine the effects of electron-produced secondary electrons on the net charging current and the equilibrium voltage of spacecraft surfaces immersed in hot (keV) plasmas. The effects of secondary electrons produced by the primary plasma electrons were examined for aluminum, Teflon, Mylar, and Kapton. The results indicate that it is the primary electron energy region over which the secondary electrons are emitted (rather than the ratio of secondary to primary electrons) which determines the effectiveness of the secondary electrons in limiting the net charging current. While the electron-produced secondary electrons decreased the charging current, they had little effect upon the equilibrium voltages attained.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 756-768
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  • 10
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    Parameter dependence of equilibrium charging potentials (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Equilibrium surface potentials for slab configurations (representative e.g. of a large solar power satellite) are determined under extensive parametric variations of materials, solar exposure and substorm characteristics. The dependence of the charging process for dielectric surfaces on the various material response characteristics is described. The results are used as a guide to the material selection and design of large space systems to minimize dielectric breakdowns and reduce parasitic leakage currents on SPS.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA Lewis Res. Center Spacecraft Charging Technol., 1978; p 711-733
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  • 11
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    Thermal blanket metallic film groundstrap and second surface mirror vulnerability to arc discharges (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Available data on the geosynchronous orbit energetic plasma environment were examined, and a crude model was generated to permit an estimation to be made of the number of arc discharges per year to which a thermal blanket groundstrap would be subjected. Laboratory experiments and a survey of the literature on arc discharge characteristics were performed to define typical and worst case arc discharge current waveforms. In-air tests of different groundstrap configurations to a standardized test pulse were performed and a wide variability of durability values were found. A groundstrap technique, not used thus far, was found to be far superior than the others.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 657-681
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  • 12
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    Secondary emission effects on spacecraft charging: Energy distribution considerations (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The conditions under which multiple valued solutions occur by computing the floating potential of an isolated eclipses surface on a geosynchronous orbit spacecraft were examined. Different approximations for the electron spectra during a geomagnetic substorm were used. The result indicates that if the incident electron flux has a Maxwellian energy distribution, the ratio of the secondary emitted current to the incident electron current is independent of the spacecraft potential. In this case a single value solution to the current equation occurs.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 747-755
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  • 13
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    Model for breakdown process in dielectric discharges (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A heuristic model to explain the blowoff of charge during an electron-induced dielectric discharge is presented. It is proposed that blowoff of charge is initiated by a punchthrough or a flashover. The discharge time is assumed to be governed by an LC time constant where L is the inductance of the electrons flowing in the branches of the Lichtenberg figures at an electron range below the irradiated surface and C is the capacitance between the trapped electrons and the substrate for the discharged area. Experiments to verify that blowoff is a consequence of punchthrough or flashover and to measure the punchthrough current and the variation of discharge time with the sample area and thickness were conducted.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 704-710
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  • 14
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    Area scaling investigations of charging phenomena (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The charging and discharging behavior of square, planar samples of silvered, fluorinated ethylene-propylene (FEP) Teflon thermal control tape was measured. The equilibrium voltage profiles scaled with the width of the sample. A wide range of discharge pulse characteristics was observed, and the area dependences of the peak current, charge, and pulse widths are described. The observed scaling of the peak currents with area was weaker than that previously reported. The discharge parameters were observed to depend strongly on the grounding impedance and the beam voltage. Preliminary results suggest that measuring only the return-current-pulse characteristics is not adequate to describe the spacecraft discharging behavior of this material. The seams between strips of tape appear to play a fundamental role in determining the discharging behavior. An approximate propagation velocity for the charge cleanoff was extracted from the data. The samples - 232, 1265, and 5058 square centimeters in area - were exposed at ambient temperature to a 1- to 2-nA/sq cm electron beam at energies of 10, 15, and 20 kilovolts in a 19-meter-long by 4.6-meter-diameter simulation facility at the Lewis Research Center.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 485-506
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  • 15
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    Plasma particle trajectories around spacecraft propelled by ion thrusters (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The thruster plasma is assumed to be described by a collimated energetic beam and a cloud of ionized thermal propellant produced by charge-exchange. A simple adiabatic model is used to describe the expansion of these neutral plasmas away from the source. As the pressure falls, shielding currents dissipate, and the geomagnetic field takes control of the particles. In low earth orbit, it is concluded that the vehicle easily outruns its thruster plasma. At geosynchronous altitude, the local electric fields around high voltage surfaces collect return current from the thermal plasma that appears to be limited only by the available space charge. Results appropriate to proposed electric propulsion missions and the solar power satellite are presented and operational considerations are discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 419-436
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  • 16
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    Effect of parasitic plasma currents on solar-array power output (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Solar-array voltage-current curves are calculated by assuming the existence of parasitic loads that consist of local currents of charged particles collected by the array. Three cases of interest are calculated to demonstrate how the distribution and magnitude of parasitic currents affect output. Solar array performance degradation became significant when the total parasitic current plus the load current exceeded the short-circuit current. Approximate graphical methods were useful for many applications. Power loss, which was calculated by summing the product of parasitic current and the local potential, underestimated the loss in maximum power.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 358-375
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  • 17
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    Magnetic shielding of large high-power-satellite solar arrays using internal currents (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Present concepts for solar power satellites involve dimensions up to tens of kilometers and operating internal currents up to hundreds of kiloamperes. A question addressed is whether the local magnetic fields generated by these strong currents during normal operation can shield the array against impacts by plasma ions and electrons (and from thruster plasmas) which can cause possible losses such as power leakage and surface erosion. One of several prototype concepts was modeled by a long narrow rectangular panel 2 km wide and 20 km long. The currents flow in a parallel across the narrow dimension (sheet current) and along the edge (wire currents). The wire currents accumulate from zero to 100 kiloamp and are the dominant sources. The magnetic field is approximated analytically. The equations of motion for charged particles in this magnetic field are analyzed. The ion and electron fluxes at points on the surface are represented analytically for monoenergetic distributions and are evaluated.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 376-387
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  • 18
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    Spacecraft potential control on ISEE-1 (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Active control of the potential of the ISEE-1 satellite by the use of electron guns is reviewed. The electron guns contain a special cathode capable of emitting an electron current selectable between 10 to the -8th power and 10 to the -3rd power at energies from approximately .6 to 41 eV. Results obtained during flight show that the satellite potential can be stabilized at a value more positive than the normally positive floating potential. The electron guns also reduce the spin modulation of the spacecraft potential which is due to the aspect dependent photoemission of the long booms. Plasma parameters like electron temperature and density can be deduced from the variation of the spacecraft potential as a function of the gun current. The effects of electron beam emission on other experiments are briefly mentioned.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 256-267
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  • 19
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    Electron transport model of dielectric charging (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A computer code (SCCPOEM) was assembled to describe the charging of dielectrics due to irradiation by electrons. The primary purpose for developing the code was to make available a convenient tool for studying the internal fields and charge densities in electron-irradiated dielectrics. The code, which is based on the primary electron transport code POEM, is applicable to arbitrary dielectrics, source spectra, and current time histories. The code calculations are illustrated by a series of semianalytical solutions. Calculations to date suggest that the front face electric field is insufficient to cause breakdown, but that bulk breakdown fields can easily be exceeded.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 209-238
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  • 20
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    Design, development, and flight of a spacecraft charging sounding rocket payload (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Results of the flight of the spacecraft charging sounding rocket payload are given. The payload was designed to create charging by the emission of both positive ions and electrons. The relationship between environmental parameters and changes in vehicle potential during periods of emission was also studied.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 80-90
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  • 21
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    Evolution of spacecraft attitude control concepts before 1952 (1977)
    In:  CASI
    Details
    Publication Date: 2006-01-16
    Description: The control of rotational motion of a spacecraft during its free flight regime is traced from pioneer space flight to the year 1952. Essentials of attitude control systems are reviewed and spin stabilization is examined. Other topics include passive stabilization and active closed loop control.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA, Washington Essays on the History of Rocketry and Astronautics, Vol. 1; p 156-169
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  • 22
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    F. Gomez Arias' rocket vehicle project (1977)
    In:  CASI
    Details
    Publication Date: 2006-01-16
    Description: Research done by Spanish pioneer rocket scientists in the 19th century was investigated with major emphasis placed on F. Gomez Arias' rocket vehicle project. Arias, considered the world's first designer of rocket propelled, manned aircraft, was interested in solving the problem of space navigation. Major concerns included ascent and direction of heavier-than-airmachines, as well as ascent and direction of balloons.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA, Washington Essays on the History of Rocketry and Astronautics, Vol. 1; p 176-185
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  • 23
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    Thermal/environmental control and life support system requirements for future space application (1978)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Thermal/environmental control systems concepts being considered for the power module and manned orbital facilities are described. To assure an economically viable program, the approach being pursued is utilization of existing hardware, where possible, and construction of orbital facilities in modular fashion to meet anticipated growth of space processing, astronomy, life science, solar/terrestrial observation, construction, etc. Configurations include Orbiter tended missions (Orbiter remains attached to orbiting facility) and freeflying manned or unmanned earth-orbiting facilities. Utilization of shuttle hardware, e. g., external tank, or the currently orbiting Skylab vehicle is being considered for orbital facilities. Studies include use of Spacelab hardware with a power module to permit extending on-orbit mission time beyond the capability of the current space transportation system. Power levels of 25 kW are being considered for the power module. Mission duration is not limited due to the modular approach and the concept of utilizing a logistics module to provide consumables' resupply capability to the orbital facility.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Thermal and Environ. Control Systems; p 575-582
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  • 24
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    Development of the Transient Pulse Monitor (TPM) for Scatha/P78-2 (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: An instrumentation system developed for the detection of electrical transients on space vehicles is presented. The pulse monitor measures the electron pulse environment on a spacecraft at synchronous altitude, and characterizes signals produced by arcing between differentially charged elements on the spacecraft.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 876-880
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  • 25
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    A combined spacecraft charging and pulsed X-ray simulation facility (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A spacecraft charging simulation facility constructed to investigate the response of satellite materials in a typical geomagnetic substorm environment is described. The conditions simulated include vacuum, solar radiation, and substorm electrons. A nuclear threat environment simulation using a flash X ray generator is combined with the spacecraft charging facility. Results obtained on a solar cell array segment used for a preliminary facility demonstration are presented with a description of the facility.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 854-867
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  • 26
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    Investigation of electrostatic discharge phenomena on conductive and non-conductive solar reflectors (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A sample composed of non conductive optical solar reflectors (OSR) was irradiated with low energy electrons at the DERTS facility to study the effects of charge build up on thermal control coating materials. The degration effects on this panel due to electrostatic discharges were investigated to find possible alternatives to limit the amount of damage. The following systems are evaluated: (1) non conductive OSR (non conductive adhesive) (2) non conductive OSR (conductive adhesive); and conductive OSR (conductive adhesive).
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol.; p 682-703
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  • 27
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    Effects of electron irradiation on large insulating surfaces used for European Communications Satellites (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Samples of aluminized Kapton used for passive thermal control on the VHF shield and the antenna dish of ESA's OTS satellite and its derivatives were subjected to an incident electron beam of 25 keV and irradiated for 8 hours at room temperature and at -173 C under a vacuum of the 10 to the minus 6 th power torr. Visual observations during electron irradiation, measurements of leakage current and discharge characteristics, and material degradation following completion of irradiation are discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 570-586
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  • 28
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    Electric fields in irradiated dielectrics (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: An existing model for quantitatively predicting electric field build-up in dielectrics is used to demonstrate the importance of material parameters. Results indicate that electron irradiation will produce 10 to the 6th power V/cm in important materials. Parameters which can alter this build-up are discussed. Comparison to known irradiation induced dielectric charging experiments is discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 554-569
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  • 29
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    Interactions between spacecraft and the charged-particle environment (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Spacecraft-environment interactions are defined as the responses of a spacecraft surface to a charged-particle environment. This response can influence spacecraft system performance. Interactions can be divided into two broad categories: spacecraft passive, in which the environment acts on the spacecraft; and spacecraft active, in which the spacecraft causes the interaction. Passive interactions include the spacecraft-charging phenomenon. Active interactions include the relatively new interactions arising from the use of very large spacecraft and space power systems in future missions. To illustrate active interactions, a large power system operating at elevated voltages is considered. Possible interactions are described, available experimental data are reviewed, and the effect on power system performance is estimated.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 268-294
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  • 30
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    The calculation of spacecraft potential: Comparison between theory and observation (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A simple charge balance model based on the work of DeForest was adapted for the calculation of spacecraft potentials. The model was calibrated with ATS 5 plasma data. Once calibrated, the model was used to calculate the time-varying potential that was observed as a spacecraft passes in and out of eclipse. Errors on the order of + or - 800 volts were observed over a range of 0 to -10,000 volts. Possible applications of the model to large space structures are discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 239-255
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  • 31
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    Spacecraft charging results for the DSCS-3 satellite (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Spacecraft charging results are presented for the DSCS-3 satellite for a severe geomagnetic substorm. Spacecraft charging results were obtained by the use of the electrostatic charging analysis program (ESCAP). The ESCAP computer code which can determine both the transient or steady state differential charging potentials was an engineering design tool that utilized a circuit theory approach to spacecraft charging. Using the ESCAP code, the steady state (static) differential potentials of the outer spacecraft surfaces and metallic structure were obtained for the DSCS-3 satellite when under the influence of a severe geomagnetic substorm during the local midnight-to-dawn quadrant of its geosynchronous orbital path. The results indicated that, in the steady state, most of the DSCS-3 outer surface materials will not achieve differential potentials large enough to produce an electrostatic discharge.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 158-178
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  • 32
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    Comparison of NASCAP predictions with experimental data (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The NASA charging analyzer program (NASCAP) is a three dimensional, finite element computer code capable of simulating the electrostatic charging of an arbitrary body either in a ground test tank or in the space environment. The code incorporated surface property parameters needed to simulate insulating and conducting materials. These parameters are being updated as required to bring the NASCAP predictions into correspondence with data from ground tests. NASCAP predictions are also being compared with data from the ATS 5 spacecraft. The significance of these results is discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 144-157
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  • 33
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    Description of the Space Test Program P78-2 spacecraft and payloads (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The P78-2 spacecraft and its payloads are designed to measure the environment at near synchronous altitude and the interactions of the environment on the spacecraft. A brief description of each payload is provided.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 4-10
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  • 34
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    Average plasma environment at geosynchronous orbit (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The average plasma environment at geosynchronous orbit (GSO) is derived from a whole year's worth of plasma data obtained by the UCSD electrostatic electrometer on board ATS 5. The result is primarily intended for use as a general reference for engineers designing a large spacecraft to be flown at GSO. A simple mathematical formula using a 3rd order polynomial is found to be adequate for representing the yearly averaged particle energy spectrum from 70 to 41,000 eV under different geomagnetic conditions. Furthermore, correlation analyses with the geomagnetic planetary index Kp and with the auroral electrojet index AE were carried out in the hope that the ground observations of the geomagnetic field variations can be used to predict the plasma variations in space. Unfortunately, the results indicate that such forecasting is not feasible by use of these two popular geomagnetic parameters alone.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 23-37
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  • 35
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    Prediction of spacecraft potentials at geosynchronous orbit (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-16
    Description: Two relatively straightforward techniques are outlined for determining spacecraft potentials in the limit of a 'thick sheath' surrounding the spacecraft. A statistical model of the various features of the geosynchronous environment based on ATS-5 and ATS-6 data and an analytic model capable of detailed simulation of the low energy geosynchronous environment are also discussed. The results from these two environmental models are then combined with the charging models in order to provide estimates of the relationships between the geomagnetic index and spacecraft potential. The results are compared with actual potential measurements from ATS-5 and ATS-6.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Marshall Space Flight Center Solar-Terrest. Predictions Proc., Vol. 2; p 104-118
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  • 36
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    TDRSS solar array arc discharge tests (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Solar array design guidelines for immunity to the geomagnetic substorm environment at geosynchronous altitudes are summarized. The preliminary design guidelines and recommended practices based on these test results are given. It is concluded that specific design and immunity verification problems on each spacecraft program have to be solved on an individual basis until the technology matures to an adequate level.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 834-852
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  • 37
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    Charging rates of metal-dielectric structures (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Metal plates partially covered by 0.01-centimeter-thick fluorinated ethylene-propylene (FEP) Teflon were charged in the Lewis Research Center's geomagnetic substorm simulation facility using 5-, 8-, 10-, and 12-kilovolt electron beams. Surface voltage as a function of time was measured for various initial conditions (Teflon discharged or precharged) with the metal plate grounded or floating. Results indicate that both the charging rates and the levels to which the samples become charged are influenced by the geometry and initial charge state of the insulating surfaces. The experiments are described and the results are presented and discussed. NASA charging analyzer program (NASCAP) models of the experiments have been generated, and the predictions obtained are described. Implications of the study results for spacecraft are discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 507-523
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  • 38
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    Test results for electron beam charging of flexible insulators and composites (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Flexible solar-array substrates, graphite-fiber/epoxy - aluminum honeycomb panels, and thin dielectric films were exposed to monoenergetic electron beams ranging in energy from 2 to 20 keV in the Lewis Research Center's geomagnetic-substorm-environment simulation facility to determine surface potentials, dc currents, and surface discharges. The four solar-array substrate samples consisted of Kapton sheet reinforced with fabrics of woven glass or carbon fibers. They represented different construction techniques that might be used to reduce the charge accumulation on the array back surface. Five honeycomb-panel samples were tested, two of which were representative of Voyager antenna materials and had either conductive or nonconductive painted surfaces. A third sample was of Navstar solar-array substrate material. The other two samples were of materials proposed for use on Intelsat V. All the honeycomb-panel samples had graphite-fiber/epoxy composite face sheets. The thin dielectric films were 2.54-micrometer-thick Mylar and 7.62-micrometer-thick Kapton.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 457-484
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  • 39
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    Environmental interaction implications for large space systems (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Large Space Systems (LSS) comprise a new class of spacecraft, the design and performance of which may be seriously affected by a variety of environmental interactions. The special concerns associated with spacecraft charging and plasma interactions from the LSS designer's viewpoint are addressed. Survivability of these systems under combined solar U.V., particle radiation and repeated electrical discharges is of primary importance. Additional questions regard the character of electrical discharges over very large areas, the effects of high current/voltage systems and magnitude of induced structural disturbances. A concept is described for a large scale experiment platform.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 388-407
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  • 40
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    Plasma sheath effects and voltage distributions of large high-power satellite solar arrays (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Knowledge of the floating voltage configuration of a large array in orbit is needed in order to estimate various plasma-interaction effects. The equilibrium configuration of array voltages relative to space depends on the sheath structure. The latter dependence for an exposed array is examined in the light of two finite-sheath effects. One effect is that electron currents may be seriously underestimated. The other is that a potential barrier for electrons can occur, restricting electron currents. A conducting surface is assumed on the basis of a conductivity argument. Finite-sheath effects are investigated. The results of assuming thin-sheath and thick-sheath limits on the floating configuration of a linearly connected array are studied. Sheath thickness and parasitic power leakage are estimated. Numerically computed fields using a 3-D code are displayed in the thick-sheath limit.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 341-357
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  • 41
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    Status of materials characterization studies (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: In the context of the spacecraft charging technology investigation, studies were made to characterize the response of typical spacecraft surface materials to the charging environment. The objective is to obtain an understanding of the charging and discharging behavior of such materials for the reliable prediction of spacecraft response to charging environments and as a guide for the design of future spacecraft. Materials were characterized in terms of such basic properties as resistivity and secondary emission and in terms of charging and discharging behavior in simulated charging environments.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 437-456
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  • 42
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    Space environmental effects and the solar power satellite (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Some preliminary findings regarding the interactions between the space plasma at GEO and the Marshall Space Flight Center January 1978 baseline Satellite Power Systems (SPS) design are summarized. These include the following: (1) the parasitic load will be dominated by photoelectrons and will amount to about 34 MW; (2) material of higher conductivity than kapton should be used for the solar reflector substrate and the solar cell blanket support material; (3) the satellite structure and solar reflector should be tied electrically to midpoint voltage of each solar cell array; and (4) tests should be run on the proposed solar cell cover glass material (synthetic sapphire) to determine if breakdown is expected.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 408-418
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  • 43
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    Sheath effects observed on a 10 meter high voltage panel in simulated low earth orbit plasma (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: A large (1m x 10m) flat surface of conductive material was biased to high voltage (+ or - 3000 V) to simulate the behavior of a large solar array in low earth orbit. The model array was operated in a plasma environment of 1,000 to 1,000,000/cu cm, with sufficient free space around it for the resulting plasma sheaths to develop unimpeded for 5-10 meters into the surrounding plasma. Measurements of the resulting sheath thickness were obtained. The observed thickness varied approximately as V to the 3/4 power and N to the 1/2 power. This effect appears to limit total current leakage from the test array until sheath dimensions exceed about 1 meter. Total leakage current was also measured with the array.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 315-340
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  • 44
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    Plasma Interaction Experiment (PIX) flight results (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: An auxiliary payload package called PIX (plasma interaction experiment) was launched on March 5, 1978, on the LANDSAT 3 launch vehicle to study interactions between the space charged-particle environment and surfaces at high applied positive and negative voltages. Three experimental surfaces were used in this package: a plain disk to act as a control, a disk on a Kapton sheet to determine the effect of surrounding insulation on current collection, and a small solar-array segment to evaluate the effect of distributing biased surfaces among an array of insulators. Only half of the results from the 4 hours of PIX operations were recovered. The results did verify effects found in ground simulation testing. The results of this experiment are discussed in detail.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 295-314
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  • 45
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    ATS-5 and ATS-6 potentials during eclipse (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: The ATS 5 and ATS 6 data for spacecraft charging during eclipse conditions is analyzed. The ATS 5 and ATS 6 charged to voltages greater than 100 volts for about 55 percent of the eclipse periods examined. The mean spacecraft potential during eclipse was 2 keV for ATS 5, and the highest potential measured was 10 kilovolts. For ATS 6, the mean potential during eclipse was 4 keV, the highest potential measured 20 keV. The average measured spacecraft potentials for both ATS 5 and 6 depend approximately linearly upon Kp. This relationship is due mainly to the dependence of electron current density on Kp near midnight. Spacecraft potentials at geosynchronous orbit may, to a rough approximation, thus be inferred from ground-based measurements of Kp, the planetary 3-hour index.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 38-43
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  • 46
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    Modeling of the geosynchronous plasma environment (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: An analytic simulation of the geosynchronous environment in terms of local time and the daily A sub P index is presented. The simulation is compared with actual statistical data from approximately 50 days of ATS 5 plasma data and 50 days of ATS 6 plasma data. At low levels of activity the model adequately simulates the local time variations of the plasma parameters. At high values of geomagnetic activity, the predicted magnitudes of the plasma parameters agree with the statistical results but the effects of multiple injections are evident in both the data and the simulation, biasing the local time variations.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Spacecraft Charging Technol., 1978; p 11-22
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  • 47
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    Summary of the two year NASA program for active control of ATS-5/6 environmental charging (1979)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Experiments conducted on the ATS 5 and ATS 6 which have demonstrated the feasibility of modifying or clamping the environmentally induced potential of these spacecraft are described. The results of these experiments indicate that a thermionic electron source is capable of replacing photo-emitted electrons during eclipse. However, the utility of this type of device is limited if its emission is suppressed by local electric fields. On the other hand, it is shown that a plasma source will not only serve as a substitute for photo-emitted electrons but will also suppress differential charging of isolated elements of the spacecraft which would tend to suppress electron emission. This later device is therefore capable of clamping the potential of a spacecraft without special considerations of its coupling to the ambient plasma.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Spacecraft Charging Technol., 1978; p 44-58
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  • 48
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    An application of modern control theory to an elastic spacecraft (1976)
    In:  CASI
    Details
    Publication Date: 2006-01-12
    Description: Results are presented to illustrate the application of established procedures of linear, quadratic, Gaussian optimal estimation and control to a spacecraft with dynamically significant elastic appendages. Interpretations are provided in both time domain and frequency domain, and conclusions are drawn for a wide class of problems of flexible spacecraft attitude control.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Dyn. and Control of Non-rigid Space Vehicles; 5 p
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  • 49
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    Orbiting Astronomical Observatory (OAO-A2) stabilization and control subsystem (1978)
    In:  CASI
    Details
    Publication Date: 2006-03-02
    Description: The OAO was successfully operating for approximately 11 months and has made over 2500 stellar observations. The OAO control system was a complex system which takes the 4500-lb vehicle through initial stabilization to star tracker control where a pointing accuracy of less than 1 arc minute was achieved. To obtain this pointing accuracy, the system used six orthogonally mounted gimbaled star trackers. The trackers were two gimbaled devices, having a 1-deg field of view with a + or - 43 deg gimbal excursion. The tracker could recognize and track 2.0 magnitude stars or brighter. The flight results indicated spacecraft pointing accuracies of less than 1 arc minute after tracker calibration with spacecraft jitter less than 3 arc seconds.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA, Washington Fourth Inter-Center Control Systems Conf.; p 83-125
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  • 50
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    Spacecraft-produced neutron fluxes on Skylab (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Estimates of neutron fluxes in different energy ranges are reported for the Skylab spacecraft. Detectors composed of uranium, thorium, and bismuth foils with mica as a fission track recorder, as well as boron foils with cellulose acetate as an alpha-particle recorder, were deployed at different positions in the Orbital Workshop. It was found that the Skylab neutron flux was dominated by high energy (greater than 1 MeV) contributions and that there was no significant time variation in the fluxes. Firm upper limits of 7-15 neutrons/sq cm-sec, depending on the detector location in the spacecraft, were established for fluxes above 1 MeV. Below 1 MeV, the neutron fluxes were about an order of magnitude lower. The neutrons are interpreted as originating from the interactions of leakage protons from the radiation belt with the spacecraft.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Space Science Instrumentation; 3; Nov. 197
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  • 51
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    NASCAP, a three-dimensional Charging Analyzer Program for complex spacecraft (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A computer code, NASCAP (NASA Charging Analyzer Program), has been developed by Systems, Science and Software under contract to NASA-LeRC to simulate the charging of a complex spacecraft in geosynchronous orbit. The capabilities of the NASCAP code include a fully three-dimensional solution of Poisson's equation about an object having considerable geometrical and material complexity, particle tracking, shadowing in sunlight, calculation of secondary emission, backscatter and photoemission, and graphical output. A model calculation shows how the NASCAP code may be used to improve our understanding of the spacecraft-plasma interaction.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 52
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    Prediction and measurement results of radiation damage to CMOS devices on board spacecraft (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Final results from the CMOS Radiation Effects Measurement (CREM) experiment flown on Explorer 55 are presented and discussed, based on about 15 months of observations and measurements. Conclusions are given relating to long-range annealing, effects of operating temperature on semiconductor performance in space, biased and unbiased P-MOS device degradation, unbiased n-channel device performance, changes in device transconductance, and the difference in ionization efficiency between Co-60 gamma rays and 1-Mev Van de Graaff electrons. The performance of devices in a heavily shielded electronic subsystem box within the spacecraft is evaluated and compared. Environment models and computational methods and their impact on device-degradation estimates are being reviewed to determine whether they permit cost-effective design of spacecraft.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 53
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    An analytical and experimental study for surface heat flux determination (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A numerical method by which data from a single embedded thermocouple can be used to predict the transient thermal environment for both high- and low-conductivity materials is described. The results of an investigation performed to verify the method clearly demonstrate that accurate transient surface heating conditions can be obtained from a thermocouple 1.016 cm from the surface in a low-conductivity material. Space Shuttle Orbiter thermal protection system materials having temperature- and pressure-dependent properties and typical Orbiter entry heating conditions were used to verify the accuracy of the analytical procedure. Analytically generated, as well as experimental, data were used to compare predicted and measured surface temperatures.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 14; Oct. 197
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  • 54
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    Magnetic attitude control system for dual-spin satellites (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: A closed-loop control law is developed for a dual-spin satellite control system which utilizes the interaction of the geomagnetic field with the satellite dipole parallel to the spin axis. The control law consists of the linear combination of the pitch axis component of the rate of change of the geomagnetic field and the product of the roll angle and roll axis component of the geomagnetic field. Application of the method of multiple time scales yields approximate solutions for the feedback gains in terms of the system parameters. Approximate solutions are also obtained for the response of the system to disturbance torques. A comparison of the approximate solutions and numerical solutions obtained by numerical integration of the exact equations of motion is then given.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: AIAA Journal; 13; June 197
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  • 55
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    Pioneer 10 and Pioneer 11 (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: On December 3, 1974, Pioneer 11 passed the planet Jupiter at a distance of 42,000 km. The spacecraft is now headed for an encounter with the planet Saturn in September 1979. The prime scientific objectives of the Pioneer 10 and Pioneer 11 missions are discussed along with the spacecraft payload, aspects of spacecraft design, the spacecraft power supply, and details regarding the spacecraft trajectory during the approach to Jupiter.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Science; 188; May 2
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  • 56
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    Energetic solar proton vs terrestrially trapped proton fluxes (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: The relative importance of solar and trapped proton fluxes in the consideration of shielding requirements for geocentric space missions is analyzed. Using models of these particles, their fluences encountered by spacecraft in circular orbits are computed as functions of orbital altitude and inclination, mission duration, threshold energy (10 to 100 MeV), and risk factor (for solar protons only), and ratios of solar-to-trapped fluences are derived. It is shown that solar protons predominate for low-altitude polar and very high-altitude missions, while trapped protons predominate for missions at low and medium altitudes and low inclinations. It is recommended that if the ratio of solar-to-trapped protons falls between 0.1 and 10, both fluences should be considered in planning shielding systems.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 12; Feb. 197
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  • 57
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    Large Space Telescope oscillations induced by CMG friction (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: Conditions are predicted under which limit cycles may be caused by the friction torque of the control moment gyroscope (CMG) output gimbal used in the Large Space Telescope (LST). The parameter plane analysis technique is described, and three separate models of the CMG output friction nonlinearity are used in three separate analyses. The results suggest that the Dahl model (included in model 3) is the best available friction model for use in dynamic analysis. The analysis and analog simulation indicate the absence of limit cycle behavior due to the CMG output gimbal friction nonlinearity for presently estimated LST numerical parameters and the occurrence of one stable and one unstable limit cycle if substantially larger numerical values are allowed for the CMG gimbal pivot friction parameter and/or the running friction torque.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 12; Feb. 197
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  • 58
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    Spacecraft technology (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: Developments in spacecraft technology are discussed with emphasis on the Explorer satellite program. The subjects considered include the following: (1) nutational behavior of the Explorer-45 satellite, (2) panoramic sensor development, (3) onboard camera signal processor for Explorer satellites, and (4) microcircuit development. Information on the zero gravity testing of heat pipes is included. Procedures for cleaning heat treated aluminum heat pipes are explained. The development of a five-year magnetic tape, an accurate incremental angular encoder, and a blood freezing apparatus for leukemia research are also discussed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Significant Accomplishments in Sci. and Technol.; p 139-178
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  • 59
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    Gasdynamics of the cryogenic cooler of the Nimbus F spacecraft (1976)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: This paper presents a theoretical study of the gasdynamic behavior of the system used to vent methane gas from an open-cycle cryogenic cooler carried on the Nimbus F satellite. The flow of this gas from the cryogen tank to the vent-line exit takes place at very low Reynolds number, and is strongly affected by variations in the tube cross-sectional area, substantial heat addition, and important frictional effects. All three of these factors play a role in determining the choking mass flow. Finite-difference calculations are used to establish the mass flow rates for a range of cryogen temperatures. Finite-difference calculations are also used to determine the flowfield in the plume that results when the vented gas exits from the spacecraft. The momentum-flux distributions within the plume are then used to estimate the torques applied to the spacecraft when the plume impinges on several surfaces near the exit point. These torque estimates compare favorably with flight data.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 13; Dec. 197
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  • 60
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    The SAS-3 attitude control system (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: SAS-3 uses a reaction wheel to provide torque to control the spin rate. If the wheel speed becomes too great or too small, it must be restored to its nominal rate by momentum dumping which is done by magnetic torquing against the earth's magnetic field by the satellite's magnetic coils. A small rate-integrating gyro is used to sense the spin rate so that closed loop control of the spin rate can be achieved. These various systems are described in detail including the reaction wheel system, the gyro system, along with control modes (spin rate control and the star lock mode).
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: APL Technical Digest; 14; July-Sep
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  • 61
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    The SAS-3 power and thermal systems (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: Solar array configurations of the SAS-3 are described: a configuration with two sets of coplanar panels in the horizontal and two others in the vertical position, and two other configurations with either four horizontal or four vertical sets of panels. The nickel-cadmium battery of the power subsystem is described in detail, with emphasis on voltage limits and charge-discharge characteristics. The characteristic of 'solar-only' operation in the case of damage to the battery is discussed. The thermal subsystem of SAS-3 is considered, with discussions of thermal design criteria and the thermal environment. Temperature is controlled by using internal thermal louvers that regulate the rate at which the heat load from electronic equipment is transmitted to the outer surface for dumping to space.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: APL Technical Digest; 14; July-Sep
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    Future earth orbit transportation systems/technology implications (1976)
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    Publication Date: 2011-08-16
    Description: Assuming Space Shuttle technology to be state-of-the-art, projected technological advances to improve the capabilities of single-stage-to-orbit (SSTO) derivatives are examined. An increase of about 30% in payload performance can be expected from upgrading the present Shuttle system through weight and drag reductions and improvements in the propellants and engines. The ODINEX (Optimal Design Integration Executive Computer Program) program has been used to explore design options. An advanced technology SSTO baseline system derived from ODINEX analysis has a conventional wing-body configuration using LOX/LH engines, three with two-position nozzles with expansion ratios of 40 and 200 and four with fixed nozzles with an expansion ratio of 40. Two assisted-takeoff approaches are under consideration in addition to a concept in which the orbital vehicle takes off empty using airbreathing propulsion and carries out a rendezvous with two large cryogenic tankers carrying propellant at an altitude of 6100 m. Further approaches under examination for propulsion, aerothermodynamic design, and design integration are described.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 14; Sept
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    Prediction and measurement of radiation damage to CMOS devices on board spacecraft (1976)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The CMOS Radiation Effects Measurement (CREM) experiment is presently being flown on the Explorer-55, to evaluate device performance in the actual space radiation environment and to correlate the respective measurements to on-the-ground laboratory irradiation results. The experiment contains an assembly of CMOS and P-MOS devices shielded in front over 2 pi steradian by flat slabs of aluminum of 40, 80, 150, and 300 mils (1.02, 2.04, 3.81, and 7.62 mm) thicknesses, and by a practically infinite shield in the back. Initial results from the CREM experiment are reported. Predictions of radiation damage to CMOS devices are based on standard environment models and computational techniques. Measured space damage is smaller than predicted by about a factor of 2-3 for thin shields but agrees well with predictions for thicker shields. It is not clear at this time how the trapped particle environment models or the computational methods should be modified in order to achieve better agreement between experimental results and predicted damage curves.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 64
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    In-orbit servicing (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: A concept called the low-cost modular spacecraft, in which the subsystems would be contained in replaceable modules and which would be capable of various applications, is under engineering development. Studies have shown that servicing of such a spacecraft on orbit by the Space Shuttle over an extended lifetime is more economical than three other alternatives: launching replacement satellites on conventional boosters as existing satellites fail, launching replacements by Shuttle, or launching replacement satellites and retrieving malfunctioning satellites by Shuttle. For on-orbit servicing, the Shuttle would use its Remote Manipulator system (RMS) to bring the malfunctioning satellite into its cargo bay. The Flight Support System (FSS) on the Orbiter used for servicing the satellite would include an appendage storage frame, a retention cradle, a positioning platform, a Module Exchange Mechanism (MEM), and a storage magazine. Ground simulations are under way with the Orbiter mockup.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 13; Feb. 197
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  • 65
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    The Vikings are coming (1976)
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    Publication Date: 2011-08-16
    Description: The exploration of Mars with the aid of the two Viking landers is discussed. The probable landing date for Viking Lander 1 will be July 4, 1976, and for Lander 2 September 4, 1976. The criteria used in selecting the landing site are considered along with the procedure to be employed in the final approach of the lander to the planet and the touchdown. A description is given of the studies to be conducted by the lander, taking into account the search for life and the collection of meteorological and seismological data. Attention is also given to technical data concerning the spacecraft, details regarding the software, and the ground facilities on earth which are used for the Viking project.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: IEEE Spectrum; 13; June 197
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  • 66
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    Stability bounds for the control of large space structures (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Balas (1977) has discussed the stability problem of reduced-order regulators and estimators in terms of control and observation 'spillover'. The term 'control spillover' was used to define that part of the feedback control which excites the uncontrolled (or residual) modes, and 'observation spillover' was used to define that part of the measurement which is contaminated by residual modes. In this paper, two sufficient conditions are derived via Lyapunov methods for asymptotic stability of large space structures using a class of reduced-order controllers. These conditions give allowable bounds on the spectral norms of control and observation spillover terms. The sufficient condition given by a specified inequality equation appears to be less conservative, and should be useful as a design tool for the control of large space structures.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Guidance and Control; 2; July-Aug
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  • 67
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    Contamination enhanced electrostatic discharge mechanisms (1979)
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    Publication Date: 2011-08-17
    Description: The two problems of enhanced electrostatic discharge (ESD) and contamination are discussed. It is shown that there is a synergistic relationship between them such that one enhances the probability of occurance of the other. The action of both provides substantially more deleterious affects than the effects of both separately. Mechanisms for such a relationship are discussed as well as application to large advanced technology systems.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Spacecraft Mater. in Space Environ.; p 145-158
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  • 68
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    Pointing and control of planetary spacecraft - The next 20 years (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Four basic categories of spacecraft pointing and control systems are discussed: (1) spin stabilized, (2) dual spin, (3) three-axis, and (4) momentum bias. Goals, mission requirements, and payloads should be considered for selection of control systems. A set of representative requirements, including life and environmental, vehicle/payload orientation and accuracy requirements, and control-system operational and functional requirements are discussed. Attention is given to design constraints imposed by the system selection. It is noted that sensitivity of any system design to increasing accuracy and flexibility in science-mission reprogramming will be major cost factors.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 17; Oct. 197
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  • 69
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    Engineering evaluation of existing space hardware for utilization in small self-contained payloads (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The NASA small self-contained payload program allows research and development experiments of less than 5 cu ft to be flown in the Orbiter payload bay on a space available basis. Materials processing equipment available from Apollo, Skylab, SPAR, and MEA that include such items as furnaces for metallurgical processing and crystal growth, apparatus for measuring crystallization processes, and electrophoretic separators for cell separation are being offered by NASA through this program. Eighteen items for the self-contained payload program were evaluated and many were found to be suitable.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: ESA Mater. Sci. in Space; p 205-210
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  • 70
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    Trends in the automation of planetary spacecraft (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The automation of planetary spacecraft at the Jet Propulsion Laboratory (JPL) is discussed. Factors affecting the development of spacecraft automation, such as predictable and repetitive functions and narrow time-window, are analyzed. The volume of command data transmitted to the spacecraft is considered, together with an examination of 'autonomy' (executing functions without outside control) in relation to the ground command activity needed during the mission. The role of the spacecraft's growing computational power in increasing vehicle autonomy is noted.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 17; May 1979
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  • 71
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    Transparent electrically conducting thin films for spacecraft temperature control applications (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Thin transparent films of In2O3 or In2O3 + SnO2 prepared by evaporation or sputtering have been tested for use as surface layers for spacecraft temperature control coatings. The films are intended to prevent nonuniform electric charge buildup on the spacecraft exterior. Film thicknesses of 300 to 500 A were found to be optimal in terms of durability and minimum impact on the solar absorptance and the thermal emissivity of the underlayers. As a verification of their suitability for long-duration space missions, the films were subjected to simulated solar UV plus proton irradiation in a vacuum.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Applied Optics; 18; May 15
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  • 72
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    New directions in space electronics (1977)
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    Publication Date: 2011-08-17
    Description: Five new directions in space electronics are defined to yield improvements over present means: automated operations aimed at a tenfold reduction in mission support costs, precision pointing and control, efficient data acquisition to permit the tenfold increase in information collection needed for global coverage, real-time data management, and low-cost data distribution to allow a thousand-fold increase in information availability and space-systems effectiveness. Technological developments and resources for each of the new directions are derived from space-technology and space-theme workshops. One to two order of magnitude improvements in instrument pointing stability will be necessary for Shuttle and planetary payloads of the 1980s. Low-cost user terminals will involve technological addressing both near-term and long-term applications. NASA has reoriented its research and technology-base programs to support the goals represented by the new directions in space electronics. Numerous charts and schematic diagrams supplement the text.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 15; Feb. 197
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  • 73
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    Space Shuttle's orbital environment (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Variations with the solar cycle of chemical composition, density and temperature of the orbital environment of the Space Shuttle are reviewed; micrometeoroid fluxes, ionizing radiation, photon fluxes and cosmic ray bombardment to which the Shuttle will be subjected are also mentioned. In addition, gases arising from offgassing and outgassing of contaminants (at about 300 K) from the Shuttle are considered, and particulates ejected during rocket firings or from the leakage of pressurized vessels (i.e., the cabin) are taken into account.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 74
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    Methods of shield analysis for protection against electrons in space (1977)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: It is anticipated that many future manned space operations will be radiation limited and that laminated wall structures and the use of new materials will be required to reduce radiation exposure. Methods for electron shield analysis are reviewed in light of anticipated needs in the space program. The most general method is still the Monte Carlo method, which is of limited usefulness for shield analysis due to excessive computer requirements. Methods based on energy deposition coefficients or energy transmission and reflection factors are quite accurate, but are currently limited to aluminum shield material. Analytical methods based on Mar's approximation for the electron transmission factor are relatively general and computer efficient but seriously underestimate shield requirements. A correction to methods using Mar's approximate transmission factor is derived herein and results in a slightly conservative estimate of shield requirements. Techniques for laminated shield design are still lacking.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Nuclear Technology; 35; Aug. 197
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  • 75
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    Weight optimization methods in space radiation shield design (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: An empirical relation between proton range and material density is used to examine relations between shield weight, geometry, and material composition for shielding against a space proton environment. The optimum material resulting in minimum shield weight usually lies at the extremes of either the lightest or heaviest materials. Aluminum, which has special prominence in the space program, appears universally suboptimal as a radiation shielding material. Assuming square-box geometry (rectangular prisms with two square faces), the optimum shape for the shielded objects is found to be a cube, although moderate deviations from a cube result in only a small weight penalty.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Journal of Spacecraft and Rockets; 12; Dec. 197
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  • 76
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    ATS-6 - Flight performance of the Advanced Thermal Control Flight Experiment (1975)
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    Publication Date: 2011-08-16
    Description: The Advanced Thermal Control Flight Experiment on ATS-6 was designed to demonstrate the thermal control capability of a thermal diode (one-way) heat pipe, a phase-change material for thermal storage, and a feedback-controlled heat pipe. Flight data for the different operational modes are compared to ground test data, and the performance of the components is evaluated on an individual basis and as an integrated temperature-control system.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: IEEE Transactions on Aerospace and Electronic Systems; AES-11; Nov. 197
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  • 77
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    ATS-6 - Spacecraft Attitude Precision Pointing and Slewing Adaptive Control Experiment (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: The primary objective of the Spacecraft Attitude Precision Pointing and Slewing Adaptive Control (SAPPSAC) experiment is to establish feasibility and evaluate capabilities of a ground-based spacecraft attitude control system, wherein RF command and telemetry links, together with a ground station on-line minicomputer, perform closed loop attitude control of the Applications Technology Satellite-6 (ATS-6). The ground processor is described, including operational characteristics and the controller software. Attitude maneuvers include precision pointing to fixed targets, slewing between targets, and generation of prescribed ground tracks. Test results show high performance and reliability for over 30 hours of on-line control with no serious anomalies. Attitude stabilization relative to a prescribed target has been achieved to better than 0.007 deg in pitch and roll and 0.02 deg in yaw for a period of 43 min. Ground tracks were generated which had maximum latitude/longitude deviations less than 0.15 deg from reference.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: IEEE Transactions on Aerospace and Electronic Systems; AES-11; Nov. 197
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  • 78
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    ATS-6 - Description and performance (1975)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: The Applications Technology Satellite-6 embodies a geosynchronous technology that is sensitive enough to receive signals from simple inexpensive earth-based terminals and is powerful enough to relay color television and other high-quality signals to these terminals. This was accomplished by deploying a 9.1-m parabolic reflector antenna in space, by performing precision attitude maneuvers on commands from the earth to point the reflector to selected spots on the earth's surface, and by providing communications links and necessary power to receive and rebroadcast signals transmitted from the ground. All major objectives of the ATS program have been satisfied, with specifications being met or exceeded.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: IEEE Transactions on Aerospace and Electronic Systems; AES-11; Nov. 197
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  • 79
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    Testing the Shuttle heat-protection armor (1976)
    In:  Other Sources
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    Publication Date: 2011-08-16
    Description: The article deals with the thermal protection system (TPS) designed to keep Space Shuttle structures at 350 F ratings over a wide range of temperatures encountered in orbit, but also during prelaunch, launch, deorbit and re-entry, landing and turnaround. The structure, function, fabrication, and bonding of various types of reusable surface insulation and composite materials are described. Test programs are developed for insulation, seals, and adhesion bonds; leak tests and acoustic fatigue tests are mentioned. Test facilities include arc jets, radiant heaters, furnaces, and heated tunnels. The certification tests to demonstrate TPS reusability, structural integrity, thermal performance, and endurance will include full-scale assembly tests and initial orbital flight tests.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 14; Jan. 197
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  • 80
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    Pointing and control of planetary spacecraft - The next 20 years (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: The preliminary pointing and control systems for five planetary mission groups are presented, as well as the rationale and key characteristics for each system type. The five groups entail: (1) a preliminary survey, (2) a detailed remote observation, (3) close scrutiny in deep space, (4) close scrutiny near the sun, and (5) a sample return. Attention is given to each group with respect to two- and three-axis control and various instruments for spin control. The future development of component trends and needs, electronic trends, electromechanical development, gyros and other general system trends are discussed in detail.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 17; Nov. 197
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  • 81
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    NASCAP modelling of environmental-charging-induced discharges in satellites (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A study of the charging and discharging characteristics of a typical geosynchronous satellite experiencing time-varying geomagnetic substorms, in sunlight, is conducted. The NASA Charging Analyzer Program (NASCAP) is used. An electric field criteria of 1.5 x 10 to the 5th volts/cm to initiate discharges and transfer of 67% of the stored charge is used in this study, based on ground test results. The substorm characteristics are arbitrarily chosen to evaluate effects of electron temperature and particle density (which is equivalent to current density). It has been found that while there is a minimum electron temperature for discharges to occur, the rate of discharges is dependent on particle density and duration times of the encounter. Hence, it is important to define the temporal variations in the substorm environments.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 82
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    Electron, electron-bremsstrahlung and proton depth-dose data for space-shielding applications (1979)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: A data set has been developed, consisting of depth-dose distributions for omni-directional electron and proton fluxes incident on aluminum shields. The principal new feature of this work is the accurate treatment, based on detailed Monte Carlo calculations, of the electron-produced bremsstrahlung component. Results covering the energy region of interest in space-shielding calculations have been obtained for the absorbed dose (a) as a function of depth in a semi-infinite medium, (b) at the edge of slab shields, and (c) at the center of a solid sphere. The dose to a thin tissue-equivalent detector was obtained as well as that in aluminum. Various results and comparisons with other work are given.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 83
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    The International Ultraviolet Explorer: Case study in spacecraft design (1979)
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    Publication Date: 2011-08-17
    Description: The International Ultraviolet Explorer (IUE) is a geosynchronous scientific satellite that was conceived as an international space observatory capable of measuring UV spectra of faint celestial bodies. Simple operational procedures allow the astronomers to joystick the spaceborne telescope about the sky, using familiar ground-based observatory techniques. The present paper deals with the IUE project objectives, the technical problems, constraints, trade-offs, and the problem solving techniques used in the IUE program.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 84
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    Some results from 1/8-scale Shuttle model vibration studies (1978)
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    Publication Date: 2011-08-17
    Description: Highlights of experimental and analytical vibration studies of a 1/8-scale structural dynamic model of the Space Shuttle are presented. The Space Shuttle is a launch vehicle with elements assembled in an asymmetric manner. Responses of the assembled vehicle are characterized by directional coupling and high modal density at low frequencies. Effects of distortion of structure near element interfaces are shown to be significant and predictable with highly detailed mathematical models. Acquisition of modal data by single-point random excitation is shown to be viable for these complex structures. Element studies reveal large liquid-structure interactions and a wide range of structural damping.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
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  • 85
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    Cryogenic cooling for spacecraft sensors, instruments, and experiments (1978)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Several disciplines requiring in-space cryogenic cooling are identified including high-energy, gamma-ray, and IR astronomy, relativity missions, and superconducting devices. Radiant coolers are limited in terms of temperature ranges and cooling loads. Other spacecraft cryogenic systems include stored solid cryogenic coolers using materials such as hydrogen, neon, argon, and methane. Two such cooler designs are described including one for the Nimbus F limb radiance inversion radiometer and one for the Nimbus G limb infrared monitoring of the atmosphere. Suggestions for increasing the performance of solid cryogenic coolers are made, such as a multimission cooler, a mechanical refrigerator, Stirling-cycle refrigerators, and Vuilleumier mechanized coolers. Techniques for obtaining cryogenic cooling in the milli-K range are identified as dilution refrigeration and adiabatic demagnetization.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 16; Nov. 197
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  • 86
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    Insulator edge voltage gradient effects in spacecraft charging phenomena (1978)
    In:  Other Sources
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    Publication Date: 2011-08-17
    Description: Insulating surfaces on geosynchronous satellites have been charged by geomagnetic substorms to a point where discharges occur. The electromagnetic pulses from these discharges couple into satellite electronic systems disrupting operations. Laboratory tests conducted on insulator charging have indicated that discharges appear to be initiated at insulator edges where voltage gradients can exist. An experimental investigation has been conducted to measure edge voltage gradients on silvered Teflon samples as they are charged by monoenergetic electron beams. It has been found that the surface voltage at insulator edges can be approximated by an exponential expression based on an electron current density balance. Using this expression at known breakdown conditions results in a discharge voltage gradient down the insulator edge to ground of about 150,000 V/cm.
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  • 87
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    Practical design of low-cost large space structures (1978)
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    Publication Date: 2011-08-17
    Description: The practical design of low-cost space structures involves the use of approaches that reduce the cost of the design and development effort itself. Such approaches include: (1) design with criteria arrived at rationally; (2) design for simplicity, repeatability, and modularity; (3) assembly without adjustments; (4) design for testability on the ground; (5) the attainment of structural efficiency by configuration and material choice rather than by squeezing down on the design margins; and (6) prefabrication and preassembly before launch. Attention is given to truss platform designs, and the Seasat Extendible Support Structure is discussed as a particular example.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Astronautics and Aeronautics; 16; Oct. 197
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  • 88
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    Proceedings of the Spacecraft Charging Technology Conference: Executive Summary (1977)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: Aerospace environments are reviewed in reference to spacecraft charging. Modelling, a theoretical scheme which can be used to describe the structure of the sheath around the spacecraft and to calculate the charging currents within, is discussed. Materials characterization is considered for experimental determination of the behavior of typical spacecraft materials when exposed to simulated geomagnetic substorm conditions. Materials development is also examined for controlling and minimizing spacecraft charging or at least for distributing the charge in an equipotential manner, using electrical conductive surfaces for materials exposed to space environment.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: NASA. Lewis Res. Center Proc. of the Spacecraft Charging Technol. Conf.; p 9-20
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  • 89
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    A way to compare costs of building rotating structures in space (1976)
    In:  CASI
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    Publication Date: 2016-06-07
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Lunar Science Inst., Abstracts of Papers Presented at a Special Session of the Seventh Annual Lunar Science Conference on Utilization of Lunar Materials and Expertise for Large Scale Operations in Space; p 116-119
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  • 90
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    Testing of typical spacecraft materials in a simulated substorm environment (1977)
    In:  CASI
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    Publication Date: 2016-06-07
    Description: The test specimens were spacecraft paints, silvered Teflon, thermal blankets, and solar array segments. The samples, ranging in size from 300 to 1000 sq cm were exposed to monoenergetic electron energies from 2 to 20 keV at a current density of 1 NA/sq cm. The samples generally behaved as capacitors with strong voltage gradient at their edges. The charging characteristics of the silvered Teflon, Kapton, and solar cell covers were controlled by the secondary emission characteristics. Insulators that did not discharge were the spacecraft paints and the quartz fiber cloth thermal blanket sample. All other samples did experience discharges when the surface voltage reached -8 to -16kV. The discharges were photographed. The breakdown voltage for each sample was determined and the average energy lost in the discharge was computed.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Proc. of the Spacecraft Charging Technol. Conf.; p 431-457
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  • 91
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    Space shuttle digital flight control system (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The space shuttle digital, fly by wire, flight control system presents an interesting challenge in avionics system design. In residence in each of four redundant general purpose computers at lift off are the guidance, navigation, and control algorithms for the entire flight. The mission is divided into several flight segments: first stage ascent, second stage ascent; abort to launch site, abort once around; on orbit operations, entry, terminal area energy management; and approach and landing. The FCS is complicated in that it must perform the functions to fly the shuttle as a boost vehicle, as a spacecraft, as a reentry vehicle, and as a conventional aircraft. The crew is provided with both manual and automatic modes of operations in all flight phases including touchdown and rollout.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: Advanced Control Technol. and its Potential for Future Transport Aircraft; p 271-294
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  • 92
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    Viking GC/MS mechanisms design and performance (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The Viking Lander gas chromatograph/mass spectrometer will analyze pyrolyzed samples of the Martian surface for organic content. The surface-sample loader and pyrolyzer assembly (SSPLA) is described, along with the major problems encountered during design and testing. Three mechanisms were developed to implement the required SSLPA functions: (1) a soil loader that forces soil from a filled rotating funnel into each of three ovens located on a carriage, (2) a Geneva drive for rotating and precisely indexing the ovens to receive sample, and (3) a toggle-clamp mechanism for sealing the ovens by forcing circular double knife edges into gold sealing surfaces.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 208-217
    Format: application/pdf
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  • 93
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    High stability deployable boom (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Meaningful magnetic field vector measurements in space require accurate placement of a magnetometer beyond the magnetic fields of a spacecraft. The design and development of a deployable boom is described which ensures accurate deployment of an instrument package and maintains high stability after extension.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 143-154
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  • 94
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    The Helios experiment 5 antenna mechanism (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The Experiment 5 Antenna deployment problem onboard Helios A, the failure analysis, and changes in design, test, and operation which led to a successful deployment of both antennas during the early Helios B mission phase are described.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 133-142
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  • 95
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    Caging Mechanism for a drag-free satellite position sensor (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: A disturbance compensation system for satellites based on the drag-free concept was mechanized and flown, using a spherical proof mass and a cam-guided caging mechanism. The caging mechanism controls the location of the proof mass for testing and constrains it during launch. Design requirements, design details, and hardware are described.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 125-132
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  • 96
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    Pin puller impact shock attenuation (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: Design of a pin arresting mechanism for a pyrotechnically actuated pin puller is reviewed. The investigative approach is discussed and the impact shock test results for various candidate designs are presented. The selected pin arresting design reduced the peak value of the shock response spectrum by five to one.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 71-79
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  • 97
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    The Helios mechanical despin drive assembly for the high-gain antenna reflector (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: 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.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 80-89
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  • 98
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    Apollo-Soyuz test project docking system (1976)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: The United States and Soviet Union in July 1975 successfully completed a joint space mission utilizing each country's spacecraft and the compatible docking system designed and fabricated by each country. The compatible docking system is described, along with the extensive research, development, and testing leading up to the successful mission. It also describes the formulation and implementation of methods for breaking the language barrier, bridging the extensive distances for communication and travel, and adjusting to each country's different culture during the three-year development program.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: JPL 10th Aerospace Mech. Symp.; p 26-37
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  • 99
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    The Skylab Parasol (1975)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An alternate thermal protection system for the Skylab space station cluster was deemed necessary shortly after launch when severe damage occured to the orbital workshop as it passed through the period of maximum dynamic pressure. Data indicated that the combination micrometeoroid and thermal shield on the sun facing side of the vehicle had been destroyed. The simultaneous design, fabrication, and test of the Skylab Parasol System which was completed in 6 days of around-the-clock effort is described. The Skylab Parasol design was selected as the primary sunshade device based on simplicity of design, ability to fabricate quickly, and early demonstration of a working prototype which featured semiautomatic deployment from within the workshop under shirt sleeve environment conditions.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: 9th Aerospace Mech. Symp.; p 99-113
    Format: application/pdf
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  • 100
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    An active nutation damper for spacecraft (1977)
    In:  CASI
    Details
    Publication Date: 2016-06-07
    Description: An active nutation damping device, consisting of an angular accelerometer, a dc-motor-driven flywheel, and associated electronics, developed for spacecraft use is described. This damping system was used on the LAGEOS spacecraft to control nutation buildup during the long coast period (approximately 75 minutes) after the third stage separation. The damper package was evaluated and proven on a three-axis gas-bearing simulator that duplicated the LAGEOS spacecraft critical flight dynamics. In addition, a failure analysis of the damper assembly was performed. Performance of the damper during the LAGEOS flight has confirmed the preflight evaluation and analysis.
    Keywords: SPACECRAFT DESIGN, TESTING AND PERFORMANCE
    Type: The 11th Aerospace Mech. Symp.; p 133-145
    Format: application/pdf
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