Space Power Resources Manufacuting and Dvelopmnt Volume 13 Numbers 3/4
SPACE POWER Published under the auspices of the Council for Economic and Social Studies on behalf of the SUNSAT Energy Council. Editor: Dr. Gay E. Canough, ETM Solar Works, Inc. Associate Editors: Fred Koomanoff, Dept, of Energy, USA Andrew Hall Cutler, Minerva Labs, USA Richard Boudreault, Consultant, Montreal, Canada Lars Broman, SERC, Sweden William C. Brown, Massachusetts, USA Lucien Deschamps, Paris, France Ben Finney, U of Hawaii, USA Peter Glaser, Aurther D. Little, Inc. USA Dieter Kassing, ESTEC, The Netherlands Mikhail Ya. Marov, U of North Carolina, USA Gregg Maryniak, International Space Power Program, USA Makoto Nagatomo, ISAS, Japan Viorel Badescu, Polytechnic University of Bucharest John R. Page, U of New South Wales, Australia Tanya Sienko, NASDA, Tsukuba, Japan Space Power is a quarterly, international journal for the presentation, discussion and analysis of advanced concepts, initial treatments and ground-breaking basic research on the technical, economic and societal aspects of: large-scale space-based solar power, space resources utilization, space manufacturing, space colonization, and other areas related to the development and use of space for the benefit of humanity. Recent subject coverage: • history and status of national space power programs • technologies for large-scale space power e.g. solar power satellites • systems aspects of large-scale power, e.g. SPS and central space power utilities • potential extraterrestrial resources for use in space-based manufacturing • lunar and planetary science for understanding space resource location and availability • plasma and other space environment interactions with large space structures • medical, psychological, sociological and cultural aspects of human presence in space • forms of advanced space propulsion and power technologies and systems. Space Power is published four times per year. These four issues constitute one volume. An annual index and title-page is bound in the December issue. 1994 is volume 13 ISSN = 0883-6272 Editorial Correspondence: Dr. Gay E. Canough, Space Power c/o ETM Solar Works, Inc., PO Box 67, Endicott, NY 13761, phone/fax = (607) 785-6499 e-mail (Internet): CANOUGH@BINGVAXA.CC.BINGHAMTON.EDU radio call sign: KB20XA. Business Correspondence including orders, subscriptions, advertisements, back issues and off prints should be addressed to the publisher: Council for Economic and Social Studies, 1133 13th NW, suite 2-C, Washington DC 20005,202 371 2700, fax 1523 Subscriptions: libraries: $288/year, individuals: $155/yr., additional $25 for airmail Cover: An artist's conception of a solar power satellite constructed of lunar materials, on station over the Earth's equator. Reproduced by courtesy of Charles L. Owen, Institute of Design, Illinois Institute of Technology, 10 W. 35th St. Chicago IL 60616.
Developing a Space Power Brayton System V.F. PRISNYAKOV, LN. STATSENKO, A.I. KONDRATIEV, V.L. MARKOV, B.E. PETROV, V.A. GABRINETS * Summary: This paper presents the result of preliminary development on a dynamic solar power system suitable for space applications. From the results of this development two schemes that meet most demands have been chosen. Introduction This study has been carried out at Dniepropetrovsk State University to generate and evaluate advanced solar dynamic Brayton cycle engines functioning within the range of 3 to 5 kWt. Solar dynamic systems are of great importance for the future. These systems offer the advantages of safe, reliable power production as compared to other space energy systems. Design concepts of both the systems and the components are presented along with the results of first-order evaluation of the effects of major system design parameters electric power output. The performance of such a system is presented in Table 1. * State University, Dniepropetrovsk, 72 Gargarin Av 320625 Dniepropetrovsk 10 Ukraine This paper was presented at the SPS ‘91 conference. Re-printed with permission from SEE TABLE 1. Technical performance of dynamic Brayton cycle energy plant
Figure 1. Concept schemes of space power dynamic system. In this schemes all paraboloid reflectors can be deployed in external size. 1. paraboloid reflector, plate and cross formed radiator; 2. paraboloid reflector and cylindrical radiator; 3. paraboloid reflector and divided heat receiver and heat storage system; 4. paraboloid reflector with a Kassegrain type collector and tube ray radiator; 5. facet reflector and tube arrayed radiator; 6. annular paraboloid reflector and cylindrical radiator; 7. parabolic-cylindrical reflector and tube arrayed radiator; 8. parabolic-cylindrical reflector and cylindrical radiator
General Analyses In order to meet these demands, 8 different models of power systems were analyzed during the preliminary study. All of them are given in Figure 1. Individual Analysis For various reasons, all schemes except for those having number 3 and 4 have been rejected. Scheme number 3 is shown in detail in Figure 2. where: 1. tube arrayed network of radiator; 2. collector tube of rotating unit and radiator; 3. wheel of radial flow compressor; 4. recuperator; 5. alternator rotor; 6. alternator stator; 7. gas bearings; 8. wheel of radial flow turbine 9. heat storage system; 10. aperture window; 11. heat receiver structure; 12. heat receiver; 13. framework of power station; 14. central reflector of radiator; 15. deployable petal shaped part of reflector Special mounting devices on the reflector allowed the transfer of the petal-form parts in the radial direction with respect to the framework. The usefid surface of the reflector consists of the surface of the central paraboloid reflector and the deployable petal-shaped part of the reflector. The rotating unit included the radial outflow compressor, the radial inflow turbine and the alternator rotor. The compressor and the Figure 2. Package scheme with paraboloid reflector.
turbine are mounted on the common shaft, like a console relative to the position of the alternator. The recuperator is an annular cylinder placed around the rotating unit. It is a gas-to-gas plate-and-fm counterflow heat exchanger. The recuperator design is shown in Figure 3. Figure 3. Recuperator design This design avoids using long tubes. The radiator is made as a radially-arrayed tube framework and is a rigid construction produced by welding. The heat receiver is coupled with the thermal energy storage (TES) system. In this device the collected solar energy is directly transferred to thermal energy and then extracted by the working fluid that is passing through an inner volume. The working fluid is filled with ball-shaped capsules with a heat-storing medium. The working fluid transports the extracted heat energy to the power system. TES, along with the heat receiver, is located in the central conically shaped body. The inner volume of TES is divided by distance lattices. Conclusion The deficiencies of such a system are follows: • it is necessary to decrease the tangential component of outlet velocity to about zero value, which was caused by design of the plate-and-fin and design of the heat exchanger. This led to increasing pressure drop; • use of radial outflow turbine is less efficient than the radial inflow turbine;
• complicated fabrication: the hermetic junction between glass of window and metal of receiver at high temperature. Analysis of Facet Scheme Another scheme was designed in which a facet collector was used instead of a monolithic paraboloid-shaped mirror. Each facet was mounted with special devices on common frame and was positioned separately with respect to the main focus of the whole system. In order to produce a reflecting surface, composite materials were used to manufacture the reflecting surface framework. A layer of aluminum and a protective coating were vacuum deposited on the reflecting surface. All facets are subdivided into six types according to focus length. The package scheme with Kassegrain type collector is shown in Figure 4. There are: 1. tube arrayed frame that reinforce the secondary mirror; 2. fin arrayed tube channel of the radiator; 3. central mirror of collector; 4. petal shaped deployable panel of collector; 5. secondary mirror; 6. receiver with the thermal storage system; 7. framework of the power system; 8. recuperator; 9. rotating unit; 10. the mounting device of petal form part of reflector; Figure 4. Package scheme with Kassegrain type collector.
The central part of the reflector in this scheme is similar to the previous one. The only difference lies in the manner of manufacturing the central mirror. In the latter case it's made with three layers is rigidly connected with the framework central part of energy system. The secondary mirror is installed on three hollow rods. Two of them are used as ducts for a coolant that rejects heat from the secondary mirror. The working medium passes in sequence through the recuperator and spiral-shaped duct situated on the back side of mirror. The chamber of the heat receiver is combined with the thermal storage system and all this is placed in the central part of the energy setup. The specific feature of this system is the additional surface of the channels that receive solar energy straight at the inlet of such a combined receiver. A rotating assembly is made as one unit and consists of a radial outflow compressor and a radial inflow turbine. Analysis The advantages of this system are as follows: • convenient packaging of the heat receiver with high performance; • good packaging of the collector- receiver subsystem as in the previous scheme; • Brayton power system with Kassegrain type collector satisfying better the imposed limit in maximal dimension. The system also possesses the following disadvantages: • secondary mirror reduces the reflecting efficiency of the collector; • necessity to cool the secondary mirror increases the value of pressure drop. As the result of preliminary comparative evaluation, the Kassegrain type collector is preferable. The mirrors with short focal length meet the demand of the size limit. The use of the hollow receiver combined with thermal energy storage decreases heat losses. There is no need to deploy anything for the majority of the subsystems, except for the petal-shaped part of collector. The other units are tightly connected. General mounting and adjustment can be fulfilled at plants on the Earth. Research of Working Process in High Temperature Storage System. The results of the development of latent heat storage systems in aforementioned dynamic space systems are also described here. The heat storage system is combined here with the heat receiver. The maximal temperature in this system will approximately be in range 1150 to 1175 K. Fluoride lithium (LiF) will be used as a phase change medium (PCM). The main feature of this PCM is its transparency. One of the main problems arising while developing such systems is low heat transfer rate in thermal energy storage (TES) due to low heat conductivity of the PCM. Devices, which enhance heat transfer rate, are offered to solve this problem. The purpose of this investigation was better understanding of the working process that goes on during the melting of semitransparent materials at high temperature. Experiments were performed with a cylindrical heat storage enclosure. The cross section of this enclosure is shown in Figure 5.
The testing apparatus consisted of an isolated cylinder volume of 0.45 m length and 0.18 m inner diameter. Five lamps each with a wolfram filament, were installed inside this volume. Electrical power for each lamp was 1 kWt. It was possible to use any number of lamps due to special switch. Visual observation made it possible to study the distribution of liquid and solid phase during melting-freezing cycles. The empty volume with a gas and the same volume with solid PCM during freezing process had a darker color with respect to the capsule volume. The measurements of surface temperature were made with thermocouples placed circumferentially on the outer and the inner surface of the capsule. Before the experiment, every capsule was filled with liquid PCM in a special device and was sealed also in special device by welding in argon. The properties of the PCM and the capsule material are shown in table 3. FIGURE 5. Cross section of heat storage capsule with PCM. TABLE 2 The dimension of heat storage capsule and weight of PCM Table 3 Properties of the PCM and the capsule material
The average heat losses are evaluated at every level of the capsule temperature, using experimental data on the cooling rate, correlated according to: The cooling rate is the at switched off lamp condition. Its value is measured experimentally. The results of calculation according to the previous correlation for various temperature levels are given in Figure 6. These results comprise a heat flux surface density obtained by dividing the average value of heat losses by the capsule surface area. Figure 6. Overall heat loses as function of capsule temperature. The data at temperatures over 900 K are not given in this figure because in this case we have the heat effect of the fusion process of the PCM in the surface temperature. The temperature level presented in Figure 6 was taken as a mean value between the readings of the thermocouples installed on the outer and inner surfaces. The rate of heating was calculated in the same way. The analyses of the data, given in Figure 6, shows that the uniformity of the temperature fields, circumferrentially around the capsule, occurs only in case of radiant heat put in from one side of the capsule only. In the other set of experiments, in order to eliminate the radiant heat flux from upper flat surfaces of the capsule, three thermocouples were installed (in the middle one temperature fields were measured) into the experimental set up.
These experiments were made to evaluate the capsule heat flux in the radial direction and temperature gradients in the radiant direction. Temperature variations in this case are given in Figure 7. Figure 7. Temperature variation at capsule outer (1,3) and inner (2,4) surfaces during heating; 1, 2 are the surface temperature of a single capsule; 3, 4 are the temperature of capsule with shield The analyses of the data given in Figure 7 shows that shields cause the increase of the temperature gradient in the capsule in the radial direction. It was also observed that the melting process begins 2 to 3 minutes latter. The experiments with the heating of the capsule were also made when 1,2, 3, 4 or 5 lamps were switched on in succession. The results of the heating rate at 600 K are given in table 4. Table 4 Heating rates at 600K They testified to the assumption that, in the condition of our experiment, there exists additional input of the radiant flux from different lamps. The melting process went on in the capsule when 4 or 5 lamps were switched on. Temperature variation versus time on the outer and inner surfaces of the capsule at 5 switched on lamps is shown in Figure 7 (curve 1 and 2). The experiments were repeated with a capsule made of the same material as the aforementioned capsules but not hollow inside. The temperature variation in this case are given in Figure 8.
FIGURE 8. Temperature variation at outer (1,3) and inner (2,4) surface of capsule with PCM (1,2) and the capsule fabricated from stainless steel (3,4) These data revealed insignificant differences between temperature fields from the externally heating surface and the inner side of the capsule with the PCM as well as the presence of such a difference on the opposite side. The uniformity of the temperature field in the capsule with PCM was approximately the same as the uniformity of the capsule made entirely of steel. When the PCM was freezing, the difference in capsule with PCM was more than in the one made of steel. When PCM was melting this difference was smaller. These data certify that radiant heat transfer plays an important part in the heat transfer processing in the medium of semitransparent materials. Twenty melting cycles were performed in capsules with PCM. The maximal temperature was 1175 K on the outer surface and 1160 K on the inner one. For the experiment without shields, the temperature difference was 5 to 10 K less. In both cases these differences were less than in case with capsule made of stainless steel only. Judging by a darker color in the top part of the capsule during its cooling, we have made the conclusion that liquid PCM is collected in the lower part of the capsule. The hollow volume is connected to pipes for filling. The analyses of temperature curves showed that during 90 percent of freezing time, the surface temperature change was in the 1175 to 1125 K range. The duration of this process was 3.5 minutes for every capsule. The heat flow range in this case was 25000 to 36000 Wt/m.
Foundations of Liquid Space Optics for Astronomy, Solar Power Satellites and Interplanetary Shuttles J. H. Bloomer! Invited Review Summary: Energy of the sun and stars is vast, as is the raw material of the planets. But how can we use these resources? Can they be concentrated, converted and transported? How can the abundant energy in space be used for interplanetary propulsion, mining the planets and week ending on their moons, and enriching mankind beyond its wildest dreams—in agriculture, medicine, construction, education, leisure, space exploration, ecology, and economy? Already we are used to concentrating and controlling the energy of stars, passively, and our tool is the telescope. But every optical instrument is reversible. The telescope can as well send as receive. However, the cost of such a conventional space telescope as we'd need for sending out converted solar energy (and we'd need it in space, in orbit) would be enormous. However, the tool of the orbital aluminized mylar solarcollecting mirror for concentrating solar energy in space has been demonstrated since the mid-sixties, as has the laser, for supplying converted solar energy in narrow, tight, efficient, long range beams—the bigger your sending telescope diameter, the "tighter" (or longer range) the beams. So we're back to the telescope, or its optics. That's the stumbling block. This present article proposes that the solution lies in application of some space research of the sixties that emerged in connection with the Apollo Moon-landing Project. This research, unremarkably enough, had to do with the chore of identifying the location of liquid hydrogen fuel in an Apollo second-stage rocket fuel tank, in zero gravity. It will be shown herein how potentially to apply that research—in terms of what I call "The Fundamental Expression of Liquid Space Optics"—along with some other of our scientific history's hard-won lore on liquids, space, optics, propulsion and energy, to selfconstruct at ultimately less than zero cost (i.e., turn a handsome profit in the bargain)-the required orbital, optical tools. Those tools—orbital telescope optics and laser end-mirrors, as it turns out-promise to be super large, super low cost (or negative in cost) and at the same time, super high precision, according to a succession of brainstorms or inventions of this author, as laid out collectively in the following treatise. All are intended for the public domain as "loss leaders," or non-proprietary, i.e., free-for-all, intellectual properties. Introduction, Beilby Layers An optical method of collecting/handling/transporting energy of nearby stellar sources (in arbitrary quantity for arbitrary wavelength range) continuously in space for terminals in space and on associated nearby planetary surfaces, is desired. Radially and transversely moving or stationary terminals, within practical cost/time-frame/engineering limitations, would be necessary. t DISCRAFT Corp., Portland, OR 97233 Copyright ©1994 by John H. Bloomer. Published by Space Power with permission.
For diameter-size of the precision (giant) optics required for both astronomy and electromagnetic energy-handling, there is no substitute. However, the cost of "difiractionlimited" (or geometrically optically unimprovable for given wavelength-and-diameter [1]) optics on Earth, goes up by the fourth power of the diameter [2,3], In the visible-band diffraction-limited case, for example, this is due to difficulty of grinding a given solid surface to a "Rayleigh Limit" or V4 (or 5 x 10"^ in. for green light) maximum tolerance overall while polishing it sufficiently to exhibit a uniform "Beilby" layer or surface consisting of an irregular collection of small peaks or pits none of them more than a molecule or two in height or depth [4,5], In the formation of such smooth, polished, "Beilby" layers on solids in the process of optical finishing, we have results corresponding to the smoothing out of liquid surface irregularities under "surface tension." FIGURE 1 R.W. Wood’s Mercury Mirror Telescope The feasibility of using liquid-surfaced optical construction principles for astronomical quality mirrors was first demonstrated by R. W. Wood in 1908. Using a vertical axis, spinning (therefore dynamic paraboloidal), 20 inch diameter dish of mercury, Dr. Wood resolved 5 arc second double stars near the zenith with a hand held magnifier. Capillary Properties of Liquids Meniscus properties of liquids long have been known in physical chemistry and surface chemistry. For example (Figure 2), the well known Dupre-Young equation establishes the relationship among solid-vapor, solid-liquid and liquid-vapor contact angles for a clean liquid in contact with a clean, flat solid. Hysteresis of contact angles (Figure 3)
is another well known effect which must be accounted for in liquid optics. Pressure balance exhibited by air bubbles in liquid (Figure 4), by the meniscus liquid surface in a drinking straw (Figure 5), and in a water glass (Figure 6) are, respectively, highly pertinent. FIGURE 2 Durre-Young Equation applied to wetting and non-wetting droplets Figure 3 Hysteresis of contact angle for a 3-phase boundary
Non-Spherical Zero-G Liquid Surfaces In actual application of such "capillary" surfaces in "zero-gravity" in space, moreover, a number of new observations arise, together with a new set of measuring instruments (Figures 7 and 8). Figure 4 Pressure balance in an air bubble Thus non-sphericitv of the meniscus surface must be accounted for in cases of distorted (nominally and practically chosen as toroidal) boundaries, as in Figure 9. Here the contact angle is constant, the mean curvature is constant over the surface, but clearly the liquid surface is not spherical. In another significant case of non-sphericity with a zero contact angle in a liquidhydrogen fuel tank in zero-gravity, the liquid surface is described by a mathematical infinite series involving the hyperbolic cosine (Figure 10). Behavior of liquids in zero-g, is further exhibited by Petrash & Otto's illustrations of spherical tanks part-filled with liquids of various contact angles (Figure 11) and of rectangular channels holding liquids of various contact angles (Figure 12), and by E. Benedikt's illustrations of the surface of a single liquid in an infinitely long rectangular trough under various gravitational-equivalent load-factors (Figure 13). Theory and a number of zero-gravity orbital experiments (Figures 11 and 12) and much research (Figure 10, [6, 7]), have established that in space, in zero-gravity (i.e., in orbit or unpowered trajectory), the surface of a free liquid in a container will pull itself into a minimum potential energy configuration. In general, an “epihydrostatic” surface is a liquid surface operated on by zero net acceleration along each of any chosen set of three orthoganal axes. Li [6] and others have shown that the “epihydrostatic” surface of a liquid
Figure 6 Pressure balance in a water glass FIGURE 5 Pressure resultants on capillary menisci
Figure 8 NSL optical Epihydrogoniometer FIGURE 7 Jones and Ray precision tensiometer
Figure 10 Liquid-vapor interface in a fuel tank FIGURE 9 Non-sphericity of zero-g meniscus in rectangular tank
Figure 12 Experimental configurations of liquids in 1-g and zero-g Figure 11 Configurations in liquids in zero-g
Figure 13 Menisci for water for various load factors N (after E. Benedikt) mass that is free floating or adhering to a precisely plane surface or to interior surface of a precisely cylindrical "pot" (with a precisely circular cross section)--will be precisely spherical. For clean precisely circular boundaries at fixed inclination over an arbitrary azimuth direction and at zero relative constant acceleration ("zero-gravity") therefore, a sufficient amount of clean, non-reactive, optical (reflective or refractive) liquid when introduced will by energy principles quickly exhibit a precise, concave or convex spherical surface with fixed constant contact angle at all points on the boundary. Characteristics & Scope of Liquid Optics in Space Such automatically ground and polished precision surfaces are very nearly what is needed for astronomical and energy-handling primary and secondary mirrors. "Liquid space optics" technology as conceived here uses no spinning. Static principles of adhesion, cohesion and surface tension of liquids-"epihydrostatics"--are relied on exclusively. Very large, precision optics are difficult, expensive or impossible to build on Earth at one gravity. In space they should be natural, straightforward and "easy." Liquid space optics purports to offer characteristic diffraction-limited quality in orbital mirrors up to one mile in diameter or larger, which are invulnerable to ordinary space hazards at costs which are orders of magnitude below those of conventional fabrication technology (when latter offers feasibility at any price). Optical substrates for liquid space optics are expanded and rigidized from folded plastic materials as shown in Figure 14. Then lofted, stored liquid is introduced onto these solid surfaces, where spreading is accounted for by principles of adhesion (Figure 15).
Final (spherical) curvature of (reflective or refractive) optics is effected by fixed contact angle of (a shallow, liquid optical) "pool" at the circumferential boundary (Figures 16 and 18), and by "elevation" of "liquid-level" inside boundary-ring. Of course liquids cannot "leak" out of such "pools", as shown in Figure 17. Toroidal, circumferential boundary rings are "poured" in advance of the latter step (Figure 18), and generally solidified prior to introducing the liquid mirror (for example) material. Self-Fabrication of liquid mirrors in orbit One option is the introduction of (i.e., "pumping" shown in Figures 16 and 19) a settable liquid into a circular boundary "trough" (Figure 17, 18, 20, 21 and 22). The liquid is then solidified. A liquid mirror plastic is then applied (Figure 23) which may or may not remain liquid depending on the application. It is then plated with a liquid metal which also may or may not remain liquid depending on the application. "A precision circular boundary (toroidal) trough" is needed to avoid "long-period" boundary errors (definitely impairing mirror surface sphericity, identified in Figures 9 and 24), while "short-period" boundary errors (identified in Figure 25) are precluded from affecting mirror sphericity by Young-LaPlace considerations (Figure 25). Probably best liquid plastic is Dow Coming silicone #200 [8, 9, 10], and presumably best liquid-metal is ultra-pure gallium [11, 12, 13], although many experiments are needed. Precise satellite overall temperature control will of course be needed at all times, implemented by a system like that illustrated in Figures 26 and 27. Variable-focus liquid orbital primary mirror via self-gravitation Next the important case of energy-handling variable-focus primary mirrors (Figures 28 and 29) is treated. The gravitational attraction of a ballast-mass "figuring" liquid (probably mercury) held in a (static) rotationally-symmetric, variable geometry, concentric enclosure fixed behind the mirror, is used for "fine tuning" the mirror focal length. Also note that such a system may for example be used as well for continuously varying focal length (in transmitting energy) to follow interplanetary/interstellar spaceships. That is, the primary mirror surface may gravitationally be altered (fine tuned) continuously from spherical to parabolic to plane to hyperbolic, via adjusting the size and shape (while maintaining rotational symmetry) of a gravitationally attracting bladder containing a massive liquid—probably mercury—stored behind the mirror. A variable focus primary mirror precision surface control is illustrated in Figure 28, where a "bellows" is carried on (bonded to) a complete set of radial splines encircling the mirror back structure. These splines are pivoted respectively at their centers, on a concentric raised ring encircling the mirror halfway from center to limb. A massive "figuring" liquid (like mercury - Figures 15 and 16) is carried in the (flattened) bladder between the bellows and the mirror.
FIGURE 14 Erection sequence Figure 15 F/I Orbiting eye
Figure 16 Hydraulic system and focusing mechanism FIGURE 17 Why fluids cannot leak out of boundaries
FIGURE 18 Enlarged detail of corrector rim Figure 19 Reversible screw pump for transporting liquids and adjusting optics and focus
Figure 20 Minimization of boundary errors via liquid ring FIGURE 21 Detail of corrector rim showing fabrication sequence
Figure 22 Corrector erection sequence Figure 23 Space-built precision boundary rings
FIGURE 24 Surface tension, boundary errors, cohesive force in zero-gravity FIGURE 25 Impossibility of short-period errors in liquid surfaces
Figure 26 Skin section FIGURE 27 Transit Meet 86 phase change material thermal control parameters
Thus, to gravitationally force on the liquid mirror a converging (parabolic) "free" liquid surface, one with minimum focal-length: The splines are rocked or pivoted such that outer ends are brought in contact with the mirror's limb or circumference, while concomitantly, teeter-totter-wise, the opposite ends (at mirror center) are rotated out a maximum distance away from the center. In this (converging) case, the attached, circular, flattened bladder carrying the (gravitationally) "figuring" liquid, is maximally thickened at its center while it is maximally thinned at its circumference. By local gravitational attraction of the "figuring" mass, the primary mirror's entire center is pulled infinitesimally backward (or maximum "fine-tuning" paraboloidal curvature is gravitationally achieved, as will be shown by application of Ta Li's analysis to synthesis, discussed below). Hyperbolic (coarser, convex) surface control, when required, is effected by the technique of first adding or withdrawing liquid from the shallow liquid mirror "pool". One thereby axially translates the contact ring to a given, desired "altitude" on the toroidal mirror boundary-ring (Figure 18), which will give a greater or lesser mirror radius of curvature, depending on whether the translation rotates the contact angle out of (away from) the mirror surface, or into (toward) the minor surface, respectively (Figure 18). "Altitude" of mirror surface intersection on circumferential, toroidal boundaryring, is adjusted by adding or withdrawing liquid, until desired (paraboloidal or hyperboloidal or even plane) mirror curvature is achieved. "Fine tuning" of this surface is then effected by squeezing the "gravitational figuring" liquid (probably mercury), more from the limb to the center (decreases focal length) or vice-versa (increases focal length), then locking splines at corresponding angle. Figuring telescope/laser liquid mirrors for orbit; perturbations Astronomical and energy-handling reflective primary and secondary liquid optics generally will be nearly spherical (e.g., needing only a slight retouching from the "epihydrostatic" spherical figure to generate required parabolic or hyperbolic surfaces [2, 3]), or purely spherical [14], Using an expression (the "Fundamental Expression of Liquid Space Optics," derived below), developed from Dr. Ta Li's paper on what Dr. Elliott Benedikt [7] first called "epihydrostatics", one can calculate shapes, sizes and positions of masses designed to lend the required retouching (to render spherical liquid optical surfaces, Cassegrainian or Newtonian-that is, paraboloidal). [6] Dr. Li showed in his General Dynamics ('60, San Diego) formulation for deviation of (spherical-Boundary) liquid surfaces from a spherical liquid-vapor interface under given small net axial acceleration (in zero-gravity for example, orbit or unpowered trajectory in space), that we are facing a standard isoperimetric problem (with a mobile upper limit) in the calculus of variations. Couching the rapidly-converging series solution (given by Li, op. cit.) to the resulting differential equation in both physico-chemical and optical parameters, one can show arbitrary (integrated) "figuring" gravitational acceleration, may be practically applied to such (reflective), almost arbitrarily large diameter, liquid surfaces.
FIGURE 28 One-mile-diameter variable-focus liquid-surface primary mirror (1) Solar ray; (2) Subliming-propellant micro-rocket for altitude control; (3) Extreme possible position of spline if spline is attached along its length to the skin of the 1-mile-diameter bellows; (4) Accordion-pleats around circumference of bellows; (5) Spline (attached to bellows) in opposite extreme position; (6) Solar ray after reflection from solar mirror; (7) Envelope of final focused high-energy coherent beam; (8) Short-focus surface of 1-mile-diameter liquid metal mirror “pool’’; (9) Instantaneous liquid surface of 1-mile-diameter “primary” mirror; (10) Envelope of laser beam which illuminates primary mirror; (11) Structure of posterior of primary mirror; (12) Pivot bearing for spline; (13) Interior liquid metal mass, constrained by bellows, for figuring; (14) Cylinder containing piston used for varying bellows shape; (15) Piston used for varying bellows shape; (16) Piston position for extreme long-focus of liquid mirror; (17) Piston position for extreme short-focus of liquid mirror; (18) Interior of reservoir for massive liquid; (19) Flexible conduit for massive liquid; (20) Flexible conduit for position for extreme long-focus; (21) Plastic-foam boundary ring; (22) Toroidal surface of cast-plastic mirror boundary, fabricated in space; (23) Channel for plastic; liquid plastic is “cast” by surface tension forces, results in toroidal surface; (24) Aluminized mylar material of solar collector mirror; (25) Plastic-foam structural border of solar mirror; (26) Subliming-propellant micro-rocket for solar mirror altitude control; (27) Interior of bellows for extreme short-focus configuration of liquid mirror; (28) Bottom surface of enclosure for 1-mile-diameter liquid metal mirror “pool”; (29) Membrane structure for bellows; (30) Envelope of laser beam which illuminates primary mirror; (31) Solar ray before reflection from solar mirror; (32) Solar ray after reflection from solar mirror; (34) Long-focus position of liquid mirror (surface).
FIGURE 29 Multi-orbit solar macrolaser. At the extreme left is the great solar mirror (1). It focuses solar energy onto the semi-silvered collar-like pumping mechanism (9). Energy trapped by multiple reflections in collar (9) is transferred to the transparent, hollow laser cylindrical cavity (8). Rod (8) is maintained concentric with collar (9) by strut-supports (10). The laser cavity is filled with a gas which absorbs the reflected solar energy and “lasers", i.e., transmits a coherent beam normal to the rod's end-surfaces. The end-mirror nearest the great solar mirror (1) is partially silvered, so that a portion of the coherent energy in the rod continuously escapes. The “escaping” beam is diverged by “secondary” lens (6). The latter is rigidly attached to laser-rod (8) and “pump” (9), by strut-supports (7). The diverged coherent beam (4) illuminates the large (l-mile diameter) liquid-surface “primary” mirror (3). High-precision “primary” (3) is bordered by a rigid plastic foam boundary-ring (2). Laser energy (5), is focused by reflection form primary (3), passes through the empty interior of collar (9) and emerges in the form of focused high energy coherent beam (11). The beam (11) supplies energy at or near its focus to disc-like craft (12), which might carry a protected payload as shown at (13).
The mechanism (See "Gravitational 'Figuring Example" below), is that of deliberately emplacing the required axially symmetric "sculptured" masses behind the optics (in "zero-gravity" —i.e., orbit or unpowered trajectory in space) [15, 16]. These masses' sizes, shapes and positions would be designed by a special local gravitational acceleration integration program. Of course for such "epihydrostatical" or capillary (liquid space) optics, even the slight forces lent by gravitational gradient, solar radiation, internal (machineiy or human) movements, atmospheric drag, etc., become highly significant and must be accounted for (Figure 30). Ultimate applications of liquid space optics Using solar (or in general, stellar) energy, astronomical and liquid-mirror laser- optical system principles, it becomes practical by macrolaser power-beam-from-orbit "bootstrapping" (and exponential self-construction) of ultimately vast optical systems in space, to precisely transport virtually anywhere, almost arbitrarily large (stellar-derived), energy supplies. These quantities would be conveyed at light speed to a fixed or (by variable focal length—Figures 28 and 29) moving receiver virtually anywhere in the stellar system (if any) or outside in extrastellar space (depending on latter definition) [17, 18, 19, 20, 21, 22], Figure 30 Moments affecting vehicles in Earth orbit
Macrolasers for millennium, interplanetary mining and interstellar flight Conventional cost yardsticks meantime, derived from solid optics on the ground, become totally irrelevant. For example such "macro-optics" in prospect rapidly become self-supporting, then supply a vast and exponentially increasing surplus quantity of (valuable, marketable, "free") energy, as desired, to specified locations on Earth or in space for the indefinite future. Available energy and raw material resources of the sun and planets are effectively limitless for purposes of this technology. Herein-proposed systems could result in a new, high-economic-growth "millennium". Similarly, interplanetary shuttle service could be provided by such technology at a comfortable one-gravity acceleration/deceleration, resulting in an average of two-week round trip travel times to Mars and Venus. Finally round trip interstellar flight employing first the energy of our own sun, then that of the approached star, becomes first feasible then exponentially faster and easier as permanent, exponentially self-grown "stellar energy handling satellites" (SEHS or MACROLASERS-Figure 28, 29) are emplaced respectively in orbit both of our own sun and of the target star. Relativistic interstellar-spaceship effects become marked as SEHS size (therefore energy beam "tightness" (range) and output) exponentially is increased. Thus round trip interstellar trips [19, 20, 22] will become increasingly faster (closer to light speed) for the travelers, but of course will concomitantly result in differential (greater) aging of relatives/friends at home on Earth (as predicted by Special Relativity). Limitations of conventional rocketry Today's conventional rocketry, in which the entire amount of both energy and mass needed for propulsion are carried offers little hope for comfortable, relativistic, round trip interstellar flight for humans. For practical (single-stage) manned interstellar operation, rocket exhaust velocity must slightly exceed desired rocket terminal velocity. The latter is arbitrarily selected here as one quarter of the speed of light. But the resulting 7,500,000+ sec specific impulse cannot [23] be attained by any conceivable rocket motor short of one employing the absolute destruction of matter (still inadequate for interstellar manned - humane - travel as Purcell [24] has shown, since with even matter/antimatter rockets, no provision is possible for the return voyage). New interstellar rocket motor: the particle accelerator Present proposal however puts the matter in a wholly new light (no pun intended!). The rocket motor proposed here is simply the well-known, laboratory particle accelerator, but one pared down for interstellar space travel in a fashion comparable to its earth-orbital application in particle-beam weaponry. Here, the "linear accelerator" [25, 26] may be best choice among such devices, because the distant beam power SEHS apparatus can supply ample power to run it. For rocket motors, there need be no scientific
application/conditioning whatever of the output beam, which reduses the required mass significantly. Financing the liquid-space-optic millennium: recreation and patriotism Getting started with the present technology for propulsion and power might best be enabled by (1) the recreation and power industries' respectively venturing an independent commercial enterprise after the fashion of the proposal listed immediately below, and/or (2) patriotic U.S. natural persons’ chosing to invest in such open and accountable space ventures rather than continuing to yield to massive legal deception [27.28], The latter is characterized by litigational entrapment embodied first in the 14th Constitutional Amendment then continuing with the Internal Revenue codes. The combination presently is netting SI trillion/yr in un-owed “taxes”. Such legal wording is shown [28 op.cit.] “always in the final analysis Constitutional and stopping short of actual fraud, but limitlessly deceptive and endlessly entrapping, enjoined by certain general residual faith and trust in the goodwill of government” (George Washington’s warning not withstanding). Read correctly, it appears U.S. codes and laws all strictly reflect the Constitutional mandate that “no U.S. natural person owes any tax whatever on wages, salary or other personal receipt of value” [27,28 op.cit.). Moreover, “all references and representations to the contrary not withstanding, such claims are purely and only for the (immensely successful) purposes of deception. Similarly, it appears that the $1 trillion + “federal” receipts gained by deception in the fashion given above, most are going “into the Federal Reserve Corp, and not into the U.S. Treasury at all” [27 op.cit.]. These receipts are “unnecessary to any legitimate (Constitutional) purpose of government.” [27,28 op. cit.] These “surplus” funds could be used for space development. Commercial jaunt to orbit for the general public Proposed here is the construction of the interplanetary/interstellar hi-speed propulsion system with private capital, by first implementing application of another invention of this present author, namely the VTOL, 5,000-seat, circular-platform aerospace SUPERSHUTTLE. This shuttle would haul sightseeing tourists to synchronous (22,240- mile-high) orbit and return at a suggested price of $200 per seat, by accepting energy from a dedicated, synchronous-orbiting MACROLASER system as above. The fully-loaded SUPERSHUriLE [21, 22] would earn about $1 million gross per jaunt to GEO. If the net profit is one-half, the SUPERSHUTTLE should earn about $1.2 billion net per vehicle per year, indefinitely. Or one-quarter of the latter sum would be gained just by selling attendant MACROLASER power at $0.05 per kilowatt-hour yearlong. Required attendant MACROLASER power (for SUPERSHU riLE) would be delivered by a (calculated) 1%- efficient MACROLASER system of seven miles overall (collector) diameter.
FIGURE 31 Supershuttle for everyone-in-orbit enterprise Of course delivery of solar power from GEO to the ground-or even to SUPERSHUTTLE (at the 80,000-ft.-high beginning and end of each jaunt into space)-- requires new technology, as present day laser or microwave systems deposit power inside the atmosphere, with possibly negative environmental impact to birds, planes, people, other fauna, and flora. New technology for power transmission from space The optical MACROLASER, stellar energy handling satellites, relying on the new "liquid space optics" technology, could solve the problem of economically supplying all but limitless "free" solar energy to Earth—at least to the outermost reaches of the atmosphere. There yet remains the problem of transmitting that energy safely, continuously and efficiently, to the ground. Obviously the expedient of "shining (laser or maser or even concentrated sunlight) beams down through the atmosphere," is fraught with potential risks to both physical and life forms, let alone serious attenuation of the beam itself. A desirable alternative would be a system for converting the beam to a high voltage- low current electrical output at 80,000 ft. altitude (above all clouds, birds, planes, geography, etc.), and then sending the electricity to the ground via a 16-mile-long pair of "umbilical" solid rods of, say, copper. One mile diameter "solar cell sea" at 80000 feet Such a system could be implemented with an 80,000-ft-high receiver consisting of an expanse or "sea" of solar cells. A 22-mile-collector-diameter MACROLASER system in geosynchronous Earth orbit feeding the 16-mile-high solar-cell "sea" at 1% efficiency is
assumed [20], The corresponding nominal, design-rule [20], optical primary diameter (for MACROLASER-downlink, beam output) would be l/60th x 22 miles = 1936 ft. If the given MACROLASER system should illuminate a "solar cell sea such as to produce an irradiance of 5 times normal sunlight, then if GaAs cells are used in the "Quadrispectral Converter" mode, the best efficiency = 43% [30]. Under these circumstances, a 1-mile-diameter receiver expanse is required. The 1936-ft.-diameter-primary output beam must of course slightly diverge to illuminate a 1-mile-diameter ("sea of solar cells") receiver 22,240 miles below. Such divergence could be handled easily by the primary mirror's "variable-focus" apparatus as described above (Figure 28). Of course for interplanetary/interstellar propulsion applications of the same MACROLASER beam, focusing as well as diverging geometry of the primary mirror (probably) will be required. 80000 feet high balloon carrying "solar cell sea" Meanwhile ideal balloon support for the 1-mile-diameter "solar cell sea" is a single, flat, circular-platform, shallow, wafer-shaped helium balloon which carries the solar cells on its upper surface; one that is one mile in diameter (Figure 32). Assuming the balloon is l/10th of a mile thick, the volume is 1.156 x 10^ ft3. Conservatively [31, 32] this balloon will support net 3.354 x 10$ lb. Let's say this capacity (of 3-1/3rd million lb. net) is used for supporting the weight of all solar cells and power conditioning equipment used to achieve a platform-output voltage of 500,000 Vdc. "Beanstalk" umbilical to 80000 feet The weight of the pair of continuous 16-mile-long, approximately 1-1/2-inch- diameter (arbitrary), copper umbilical rods could be carried inside a continuous series of 160 individually pivotable, 500-ft.-long, enveloping vertical wing-shaped aerodynamic- shroud/helium-balloon units (Figure 32). Rod weight (total) will be: 160 x 6407 lb. = 1025120 lb., where through each 500-ft.-long section passes a pair of 1-1/2-inch-diameter copper rod weighing (inside that section) 6407 lb. Let’s consider only the enveloping (low-drag, "weathervaning"), helium-balloon section #1, supporting the first 500 ft. of umbilical (at ground level), and section #160, supporting the last 500 ft. of umbilical (at 80000 ft.). Shroud Unit #1 will of course have length 500 ft. and chord 73 ft. and thickness 9 ft., for a thickness ratio about 12%. The horizontal cross-section area is 500 ft.3. The net sea-level buoyancy under these circumstances will be some 17400 lb. [31, 32], Beanstalk umbilical self-supported by aerodynamic shroud helium balloons By the ratios inferred from [31], 17400 lb. of gross vertical balloon-lift should be adequate to lift 170% more than the gross weight of copper rod in the section. It is believed
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that the balloon fabric and miscellaneous will not weigh more than 150% of the gross weight of 500 ft. of 1-1/2-inch-diameter copper rods. To maximize the strength of the composite umbilical, 10 strands per cable (positive and negative) would be used. The wire would be A.W.G. 0000 gage, hard-drawn copper wire (0.460 inch diameter), rated for 325 amps per strand. Thus, the over-all ratings would be 3250 amps [36]. Since as shown below, 27100 amps are to be carried, which is not necessarily a problem, because they are bare wires in open air. The tensile strength of 0000 hard-drawn copper wire is 49000 psi. Therefore, the dual cable system could support 163000 pounds. This means that each 500 ft. section (which weighs 6407 lbs) will support it’s own weight, plus 24 additional sections. Drag on balloon airfoils contiguous with umbilical “Beanstalk” umbilicals clearly would need to be mounted in relatively wind-free areas. To maximize independence from unavoidable wind forces, balloon vertical-lift must be maximized, while balloon (lateral) drag must be minimized. Pressure maximum windspeed of 68.18 mph (= 100 ft/s), and a 12%-thick (vertical-balloon) airfoil section, the drag of each 500 ft umbilical-balloon section will be given by [37]: where: Let the horizontal cross-sectional area of the balloon (upended) “wing” sections be 500 ft2, (yielding net vertical lift of bottom-most and each additional balloon, respectively, of 17400 lb). Then approximately for 12%-thick bottom-most balloon wing, the chord is 73 ft and the (maximum) thickness is 9 ft. Balloon “wing” sections will need to grow proportionally in volume, as one moves toward the top of the Beanstalk (so that each supports 17400 lb). Also, above the altitude of most winds, linear sections should grow more circular in cross-section. Estimating the drag force at or near sea level for 12%-thick, smooth wing section, the drag coefficient may be assumed about CD = 0.004 [38]. Drag, in pounds, under all the above conditions, on the sea-level, vertical balloon wing (low-drag) section (where S = 73ftx9ft = 657 ft2) will be » 31 lbs. Higher wind speeds dramatically increase drag. For example, at 200 ft/s (136 mph), drag ~ 125 lbs. Similarly, extreme increase in drag coefficient (as by damage to or distortion of balloon wings) could inordinately increase drag. Generally, increased drag evidently will be experienced as increased tension force in the (copper) umbilical cables. However, at 163000 lb maximum permissible tensile load for each balloon supported segment, it is difficult to conceive of natural circumstances imposing excessive tensile stress.
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