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Helicopter ice protection design criteria was developed and technological shortcoming in meeting helicopter mission requirements is that of helicopter rotor blade ice protection. Airframe components are protected using existing technology while the rotor blade protected using the cyclic electrothermal deicing concept.
The possibility of lightning striking the Space Shuttle during liftoff is considered and the lightning protection system designed by the Martin Marietta Corporation for the external tank (ET) portion of the Shuttle is discussed. The protection system is based on diverting and/or directing a lightning strike to an area of the spacecraft which can sustain the strike. The ET lightning protection theory and some test analyses of the system's design are reviewed including studies of conductivity and thermal/stress properties in materials, belly band feasibility, and burn-through plug grounding and puncture voltage. The ET lightning protection system design is shown to be comprised of the following: (1) a lightning rod on the forward most point of the ET, (2) a continually grounded, one inch wide conductive strip applied circumferentially at station 371 (belly band), (3) a three inch wide conductive belly band applied over the TPS (i.e. the insulating surface of the ET) and grounded to a structure with eight conductive plugs at station 536, and (4) a two inch thick TPS between the belly bands which are located over the weld lands.
An investigation to determine the effectiveness of icing protection afforded by air-heating hollow steel unpartitioned propeller blades has been conducted In the NACA Cleveland icing research tunnel. The propeller used was a production model modified with blade shank and tip openings to permit internal passage of heated air. Blade-surface and heated-air temperatures were obtained and photographic observations of Ice formations were made with variations In icing intensity and heating rate to the blades. For the conditions of Icing to which the propeller was subjected, it was found that adequate ice protection was afforded with a heating rate of 40 1 000 Btu per hour per blade. With less than 40,000 Btu per hour per blade, ice protection failed because of significant ice accretions on the leading edge. The chordwise distribution of heat was unsatisfactory with most of the available heat dissipated well back of the leading edge on both the thrust and camber face's instead of at the leading edge where it was most needed. A low utilization of available heat for icing protection is indicated by a beat-exchanger effectiveness of approximately 47 percent.
A metallic thermal protection system (TPS) concept the multiwall designed for temperature and pressure at Shuttle body point 3140 where the maximum surface temperature is approximately 811 K was tested to evaluate thermal performance and structural integrity. A two tile model of titanium multiwall and a model consisting of a low temperature reusable surface insulation (LRSI) tiles were exposed to 25 simulated thermal and pressure Shuttle entry missions. The two systems performed the same, and neither system deteriorated during the tests. It is indicated that redesign of the multiwall tiles reduces tile thickness and/or weight. A nine tile model of titanium multiwal was tested for radiant heating and aerothermodynamics. Minor design changes that improve structural integrity without having a significant impact on the thermal protection ability of the titanium multiwall TPS are identified. The capability of a titanium multiwall thermal protection system to protect an aluminum surface during a Shuttle type entry trajectory at locations on the vehicle where the maximum surface temperature is below 811 K is demonstrated.
A low cost coating for protecting metallic base system substrates from high temperatures, high gas velocity oxidation, thermal fatigue and hot corrosion is described. The coating is particularly useful for protecting vanes and blades in aircraft and land based gas turbine engines. A lacquer slurry comprising cellulose nitrate containing high purity silicon powder is sprayed onto the superalloy substrates. The silicon layer is then aluminized to complete the coating. The Si-Al coating is less costly to produce than advanced aluminides and protects the substrate from oxidation and thermal fatigue for a much longer period of time than the conventional aluminide coatings. While more expensive Pt-Al coatings and physical vapor deposited MCrAlY coatings may last longer or provide equal protection on certain substrates, the Si-Al coating exceeded the performance of both types of coatings on certain superalloys in high gas velocity oxidation and thermal fatigue. Also, the Si-Al coating increased the resistance of certain superalloys to hot corrosion.
Lightning protection system consisting of two layers of aluminum foil separated by layer of dielectric material protects graphite/epoxy composite structures on aircraft. Protective layer is secondarily applied lightning protection system, prime advantage of which is nullification of thermal and right angle effect of lightning arc attachment to graphite/epoxy laminate.
The DC wiring system for a photovoltaic power plant presents a number of unique challenges to be overcome by the plant designers. There are a number of different configurations that the grounding of the DC wiring system can take, and the choice will affect the number and type of protective devices required to ensure safety of personnel and protection of equipment. The major grounding and fault protection considerations that must be taken into account when selecting the basic overall circuit configuration are summarized. The inherent advantages and disadvantages of each type of circuit grounding (resistance or solid) along with the personnel safety and equipment protection issues for each of these grounding methods are presented.
The aerothermodynamics and thermal protection of aerobraking orbital transfer vehicles are studied with a view toward design methodology and technology requirements. A particular class of geometries is investigated: the ellipsoidally blunted raked-off elliptic cone. Discussed are the influence of various geometrical parameters on the trajectory and the heat flux as well as heating prediction methods ranging from an engineering correlation formula to computational solutions of the Navier-Stokes equations. Finite-rate catalytic thermal protection materials and gas radiation are considered. Estimates of the specific mass of possible thermal protection systems indicate that support structure is a large contributor to the aerobrake mass as compared with the thermal protection insulation itself. Sizing curves are included which enable the designer to estimate the size aerobrake required from heating considerations for different mass vehicles having aerobrakes of several specific masses.
The meteoroid and man made space debris environments of an Earth orbital manned space operations center are discussed. Protective shielding thickness and design configurations for providing given levels of no penetration probability were also calculated. Meteoroid/debris protection consists of a radiator/shield thickness, which is actually an outer skin, separated from the pressure wall, thickness by a distance. An ideal shield thickness, will, upon impact with a particle, cause both the particle and shield to vaporize, allowing a minimum amount of debris to impact the pressure wall itself. A shield which is too thick will crater on the outside, and release small particles of shield from the inside causing damage to the pressure wall. Inversely, if the shield is too thin, it will afford no protection, and the backup must provide all necessary protection. It was concluded that a double wall concept is most effective.
The results and outcome of thermal tests conducted to evaluate the performance of the protective coat of paint on the solid rocket booster (SRB) thermal protection system are discussed. A problem was uncovered during a series of tests on the SRB instrumentation islands in AEDC Tunnel C on 13 January 1979. The white protective paint or the Turco coating on the thermal protection system panels began to flow soon after the panels were exposed to the flow. This presented a serious problem especially since the critical pressure sensing, parachute opening baroswitches located on the frustum of the SRB were most likely to be contaminated by the paint flowing down the sides of the SRB nose cone. Because the first two flight articles were already completed, it was necessary to find a solution to the existing paint problem. It was found that all the coatings tested, except the Hypalon, had similar undesirable flow characteristics. Also even the Hypalon, which did not flow, would bubble up and disintegrate when it was applied on top of the new Turco. Recently, the Turco coating was removed from an MSA-:11 panel by dissolving the paint with a certain agent. This was done in two ways, by dissolving and removing almost all of the paint on one side of the panel and dissolving and removing about 50% of the paint on the other. The panel was then coated with Hypalon and tested as before in the Hot Gas Facility. No evidence of any paint flow nor any adverse performance of MSA was observed.
Sixteen materials consisting of metallizations, silicones, and FEP Teflon were applied as protective coatings to selected spacecraft material surfaces and exposed on STS-41G to the LEO atomic oxygen environment. Evaluations of their protective effectiveness were made through assessing their mass loss/gain characteristics, maintenance of base material optical properties, and imperviousness to atomic oxygen attack. Generally, all coatings provided some degree of protection for the underlying material. In some cases the overcoat appeared to be too thin thereby providing inadequate protection.
This paper discusses the fault protection tesing of the Attitude and Articulation Control Subsystem (AACS) of the Galileo spacecraft. The need for an autonomous fault protection system on an interplanetary spacecraft is discussed. Galileo requirements for the detection and response of specific hardware failures is discussed along with the fault protection software design and implementation. The test beds and test methods used for fault protection testing are described. The paper concludes with a presentation of the results of this testing with an emphasis on requirement and design changes that were made as a result of these tests.
The thermal protection systems for the shuttle/Centaur would have had to provide fail-safe thermal protection during prelaunch, launch ascent, and on-orbit operations as well as during potential abort. The thermal protection systems selected used a helium-purged polyimide foam beneath three rediation shields for the liquid-hydrogen tank and radiation shields only for the liquid-oxygen tank (three shields on the tank sidewall and four on the aft bulkhead). A double-walled vacuum bulkhead separated the two tanks. The liquid-hydrogen tank had one 0.75-in-thick layer of foam on the forward bulkhead and two layers on the larger area sidewall. Full scale tests of the flight vehicle in a simulated shuttle cargo bay that was purged with gaseous nitrogen gave total prelaunch heating rates of 88,500 Btu/hr and 44,000 Btu/hr for the liquid-hydrogen and -oxygen tanks, respectively. Calorimeter tests on a representative sample of the liquid-hydrogen tank sidewall thermal protection system indicated that the measured unit heating rate would rapidly decrease from the prelaunch rate of approx 100 Btu/hr/sq ft to a desired rate of less than 1.3 Btu/hr/sq ft once on orbit.
Accidental overwriting of files or of memory regions belonging to other programs, browsing of personal files by superusers, Trojan horses, and viruses are examples of breakdowns in workstations and personal computers that would be significantly reduced by memory protection. Memory protection is the capability of an operating system and supporting hardware to delimit segments of memory, to control whether segments can be read from or written into, and to confine accesses of a program to its segments alone. The absence of memory protection in many operating systems today is the result of a bias toward a narrow definition of performance as maximum instruction-execution rate. A broader definition, including the time to get the job done, makes clear that cost of recovery from memory interference errors reduces expected performance. The mechanisms of memory protection are well understood, powerful, efficient, and elegant. They add to performance in the broad sense without reducing instruction execution rate.
Described here are the LEC, Inc. Spline Ball Ionizer (trademark) (SBI) and the Spline Ball Terminal (trademark) (SBT) families and how their use in a conventional lightning protection system can achieve strike prevention in a conventional setting. A properly completed standards based system which is composed of the SBI and SBT in a hybrid system will provide two modes of lightning protection: (1) a stroke prevention mode that reduces the risk of a strike to the protected facility in proportion to the size of that facility and the size and number of SBI/SBT's used; and (2) a Stroke Collector-Diverter System that is far superior to any system now in use because it collects strokes entering the protected area from any direction and angle.
Through existing treaty obligations of the United States, NASA is committed to exploring space while avoiding biological contamination of the planets, and to the protection of the earth against harm from materials returned from space. Because of the similarities between Mars and earth, plans for the exploration of Mars evoke discussions of these Planetary Protection issues. U.S. Planetary Protection Policy will be focused on the preservation of these goals in an arena that will change with the growth of scientific knowledge about the Martian environment. Early opportunities to gain the appropriate data will be used to guide later policy implementation. Because human presence on Mars will result in the end of earth's separation from the Martian environment, it is expected that precursor robotic missions will address critical planetary protection concerns before humans arrive.
Characterization of the behavior of atomic oxygen interaction with materials on the Long Duration Exposure Facility (LDEF) assists in understanding of the mechanisms involved. Thus the reliability of predicting in-space durability of materials based on ground laboratory testing should be improved. A computational model which simulates atomic oxygen interaction with protected polymers was developed using Monte Carlo techniques. Through the use of an assumed mechanistic behavior of atomic oxygen interaction based on in-space atomic oxygen erosion of unprotected polymers and ground laboratory atomic oxygen interaction with protected polymers, prediction of atomic oxygen interaction with protected polymers on LDEF was accomplished. However, the results of these predictions are not consistent with the observed LDEF results at defect sites in protected polymers. Improved agreement between observed LDEF results and predicted Monte Carlo modeling can be achieved by modifying of the atomic oxygen interactive assumptions used in the model. LDEF atomic oxygen undercutting results, modeling assumptions, and implications are presented.
The heat requirements for the icing protection of two radome configurations have been studied over a range of design icing conditions. Both the protection limits of a typical thermal protection system and the relative effects of the various icing variables have been determined. For full evaporation of all impinging water, an effective heat density of 14 watts per square inch was required. When a combination of the evaporation and running wet surface systems was employed, a heat requirement of 5 watts per square inch provided protection at severe icing and operating conditions.