AIAA-95-2896
FOUNDATIONS OF OPTICS TECHNOLOGY FOR MARS
COLONIZATION, PAID ASTRONAUTICS, ENERGY
MILLENNIUM, INTERPLANETARY/INTERSTELLAR
SHUTTLES
J. H. Bloomer*, DISCRAFT Associates, Portland, OR
For
31st AIAA/ASME/SAE/ASEE JOINT PROPULSION CONFERENCE AND EXHIBIT, SAN DIEGO, CA.
Date
July 10-12, 1995
Site
Sheraton Harbor Island
Copyright 1995 by John H. Bloomer. Published by the American Institute of Aeronautics and Astronautics,
Inc. with permission.
*Member AIAA
Page i
CONTENTS:
FOUNDATIONS OF OPTICS TECHNOLOGY FOR MARS COLONIZATION, PAID ASTRONAUTICS, ENERGY
MILLENNIUM, INTERPLANETARY/INTERSTELLAR SHUTTLES
Pg.
ABSTRACT .......................................................................................................................................................... 1
1.
INTRODUCTION: PROPULSION AND CIVILIZATION ............................................................................................ 1
2.
HIGHWAYS IN THE SKY....................................................................................................................................... 3
3.
V/STOL AMPHIBIOUS AIRCRAFT RUNABOUT “FAMILY CARS” ........................................................................... 6
4.
PUBLIC AND PRIVATE TRANSPORTATION TO THE PLANETS .............................................................................. 7
5.
THE BEGINNING: BEAMING POWER FROM ORBIT VIA ADVANCED (LIQUID) LASER OPTICS ............................ 7
6.
MACROLASER SELF-SUBSIDY ............................................................................................................................. 8
6A.
STARFLIGHT...................................................................................................................................................... 10
7.
TRANSCENDING “CHINESE” ROCKETRY ........................................................................................................... 11
8.
NEW FREEDOM FOR SPECIFIC IMPULSE: STARPUMPED MACROLASERS DRIVING LASERPOWERED
REMOTE (PARTICLE-ACCELERATOR) ELECTRICROCKET MOTORS (LREM’S)..................................................... 12
9.
SPACE-BASED SOLAR/STELLAR MACROLASERS ............................................................................................... 13
10.
CHARACTERISTICS OF LIQUID SPACE OPTICS................................................................................................... 14
11.
MATHEMATICO-PHYSICAL FOUNDATIONS OF LIQUID SPACE OPTICS............................................................. 15
12.
LI’S DIFFERENTIAL EQUATION.......................................................................................................................... 21
13.
THE FUNDAMENTAL (MATHEMATICAL) EXPRESSION OF LIQUID SPACE OPTICS ............................................ 24
14.
CHARACTERISTICS OF SEHS-POWERED “NONSELF-CONTAINED” VS. “CHINESE” ROCKETRY.......................... 24
15.
THREE DAYS TO MARS BY MACROLASER-SUPPLIED 1 g “FERRYBOAT” ........................................................... 26
16.
COLONIZING MARS: CELESTIAL ENGINEERING TO CREATE BREATHABLE O2 .................................................. 29
17.
CELESTIAL ENGINEERING: CREATING OCEANS ON MARS BY MELTING PERMAFROST ................................... 30
18.
FIRST STELLAR FLYBY PROBES .......................................................................................................................... 31
19.
FIRST TWO-WAY INTERSTELLAR MANNED EXPEDITION .................................................................................. 33
20.
“SEED” SEHS WEIGHT & SIZE FOR INTERPLANETARY/INTERSTELLAR USE ...................................................... 33
21.
“SEED” EXPONENTIAL GROWTH PERIOD FOR INTERSTELLAR APPLICATION................................................... 35
22.
ADVANCE STUDY OF THE TARGET STAR’S VICINITY......................................................................................... 36
23.
TOTAL POWER DELIVERED BY THE SUN-ORBITING “PERMANENT” SEHS ....................................................... 36
24.
RATIO OF GWT. OF ACCELERATOR PKG. TO GWT. OF ASSOCIATED PAYLOAD ............................................... 36
25.
INTERSTELLAR SEH SPACESHIP ASSEMBLY EN ROUTE BY ROBOTS (AUTOMATION) ....................................... 36
26.
ALPHA-CENTAURI-CIRCLING “COCOONED” LAUNCH PACKAGE FOR ASTRONAUT RETURN ........................... 37
27.
MINIMUM SIZE OF “LAUNCH-FROM ALPHA CENTAURI” (L.A.C.) ROCKET PACKAGE ENCLOSING
EARTH-ASTRONAUT-RETURN “COCOON” (STOPPABLE WITH SUN-ORBITING SEHS)...................................... 38
Page ii
28.
TOTAL QUANTITY OF “STOPPABLE” PAYLOADS TO BE SENT TO A.C. IF ALL A.C.-CIRCLING
“SEHS ACCELERATOR” MATERIAL IS TO BE SENT FROM THE SOLAR SYSTEM ................................................. 38
29.
POSSIBLE MINIMUM QUANTITY TOTAL “STOPPABLE” A.C.-GOING TOOL-&-SUPPLIES PAYLOADS IF
EARTH-RETURN SEHS IS MANUFACTURED (NOT, ASSEMBLED ONLY) IN ALPHA CENTAURI ORBIT ................ 39
30.
GALLIUM COST AND EXPECTED COST REDUCTION ......................................................................................... 39
31.
RULES, ARBITRARY AND NOT ........................................................................................................................... 40
Page iii
ILLUSTRATIONS:
FOUNDATIONS OF OPTICS TECHNOLOGY FOR MARS COLONIZATION, PAID ASTRONAUTICS, ENERGY
MILLENNIUM, INTERPLANETARY/INTERSTELLAR SHUTTLES
FIG.
Pg.
1.
BEANSTALK UMBILICAL WITH AERODYNAMIC SHROUD UNITS TO 80000 FEET ALTITUDE .............................. 5
2.
AEROSPACE SUPERSHUTTLE .............................................................................................................................. 6
3.
PULSE-RAMJET BLASTWAVE JET ENGINE........................................................................................................... 8
4.
MULTI-ORBIT SOLAR MACROLASER................................................................................................................... 9
5.
ONE-MILE DIAMETER VARIABLE-FOCUS DIFFRACTION-LIMITED ORBITAL PRIMARY MIRROR ....................... 16
6.
LIQUID SURFACE IN LOWERED GRAVITY ......................................................................................................... 18
7.
LIQUID 1-G AND ZERO-G CONTACT ANGLES ................................................................................................... 18
8.
LIQUID OPTIC IN ZERO - GRAVITY .................................................................................................................... 19
9.
ERROR DUE TO FINITE LOAD ............................................................................................................................ 19
10.
ERROR δ VS. F/NO. FOR SEVERAL APERTURES AT 10-12 G LOAD ..................................................................... 25
11.
ERROR δ VS. AXIAL LOAD N FOR SEVERAL LIQUID GALLIUM MIRRORS .......................................................... 26
12.
ERROR VS. AXIAL LOAD FOR 300 INCH F/I OPTICS OF VARIOUS LIQUIDS ....................................................... 27
13.
INTERSTELLAR SHIP IN ACCELERATION PHASE ................................................................................................ 32
14.
ELECTROMAGNETIC NON-SELF-CONTAINED ROCKET MOTOR........................................................................ 33
15.
FABRICATION OF MACRO-OPTICS IN SPACE .................................................................................................... 34
16.
STOPPABLE SHIP IN BRAKING APPROACH TO TARGET STAR ........................................................................... 37
Page 1
ABSTRACT
Evidently post-‘Sixties earthly aerospace technology owns the optical tools necessary and sufficient to
access practically unlimited solar/stellar energy in space, condition it and beam it coherently down (1) to
planets like earth for manufacturing, agriculture, housing, education, recreation, medicine, science, and
(2) into space for opening up the interplanetary/interstellar era of propulsion, exploration and
colonization. Heretofore macro energy specialists have concentrated on nuclear technology to (1) first
create continuous local starlike processes of energy production then (2) seek to control them. Beginning
1966 this author/inventor recommended instead accessing, controlling and directing significant fractions
of the output of actual stars themselves (starting with the sun), using advanced visible, IR, UV laser
system optics. Potentially thereby the total problem would be halved. However cost of conventional such
(visible/IR/UV) optics is totally prohibitive because of the continuous extreme geometrical accuracy
required over extremely large (optical primary) surface areas. Author’s suggested new science of
capillary-controlled “epihydrostatic” (or surface-static) geometrical space optics –“liquid space optics”theoretically overcomes this objection, offering extremes of size (diameter up to one mile or more), of
accuracy (beyond diffraction limit) and of economy (self-forming). Key mathematical expression is the
author’s “fundamental relation of liquid space optics,” which is a direct result of the variational research
of Dr. Ta Li into configurations and perturbations of liquids in tanks in zero-g. Applied to geometrical
optics, Li’s results permit analysis and synthesis of all but arbitrarily large optical (reflective and refractive)
surfaces in space. Application of such surfaces to “space-based” (“zero-g,” i.e., flown in an orbit or
unpowered trajectory) lasers and laser systems, permits design of up-to-interstellar-order, Stellar-EnergyHandling Satellites or Spacecraft (SEHS). Such systems are self-formed and because of presumed power
marketability, potentially self-paid (plus evidently in principle, exponentially profitable). Proceeds and
energy of these “macrolasers” should be convertible into virtually unlimited wealth of essentially
arbitrary kinds, leading if desired to a (material) “millennium”. Similarly interplanetary shuttles and
repeatable relativistic interstellar expeditions, should become routine, by indicated grossly expanded
limits in resultant propulsion technology, leading logically to early exploration and colonization of both
solar and extrasolar planets.
1.
INTRODUCTION: PROPULSION AND CIVILIZATION
Propulsion is primarily concerned with the collection and controlled release of energy, as is evidently
civilization itself. Quoting Erb (Ref. 1), “The utilization of energy is a surrogate for economic productivity
and quality of life.” Relatedly O’Neill says (paraphrase from Ref. 2): “Ecological pressure of populations is
inversely proportional to wealth.”
Money is energy. Money is also the lifeblood of civilization. “Federal Reserve Bank” (of course not
“federal,” not a “reserve” and not a “bank”) debt-money (actual purpose apparently that of
circumventing the U.S. Constitution to rob and enslave most Americans) is diseased money. Therefore
Page 2
western civilization currently has a blood disease. New energy is new (wealth) money. Therefore the
current proposed new energy system is also a proposal whereby society can rid itself of its (planned
fatality, mathematically inevitable) blood disease and return itself to health (Ref’s 3,4,5).
Astronautics today is in a parlous state of decline, as is the American civilization itself, evidently
respectively due to the same reasons: (1) Loaning all new money into circulation as an interest-bearing
debt (to circumvent the U.S. Constitution), a system which creates the principal only and never the
interest; i.e., the total debt is always greater than the total money-supply, forcing constant borrowing to
keep the system alive. Eventually the process must become mathematically unworkable as society,
mortgaged to the maximum, can no longer afford to borrow. Meantime the debt creates extreme stress
on natural persons struggling to meet impossible money obligations, resulting in constant rise in the cost
of living, layoffs, family breakdowns, increased drug and alcohol use, increase in crime, general moral
breakdown, cancelling of spaceflight and space exploration programs, cancelling of advanced research &
technology programs of all kinds; (2) Setting up a few large banks to benefit in power or control over all
American money and property as the unpayable debt compounds and as general ability to meet the
constantly rising cost of living diminishes, as also the standard of living drops, as finally the people
become debt-slaves.
Consequently astronautics it is claimed holds the key to its own salvation and to that of society itself: That
key is honest money, which astronautics can create, does create, will create as it accesses/unlocks the
universe’s vast storehouse of energy. The latter task is the subject of this current treatise.
Key current technical discovery is Ta Li’s 1960 variational treatment of liquid-hydrogen physical and
surface chemistry in Apollo Spacecraft Second-Stage fuel tanks (Ref. 6). Li’s results – comprising the
foundations of “zero-gravity capillary epihydrostatics” (at least as applies to optics), after the term coined
in 1960 by his colleague, fellow North American Aviation scientist Elliott Benedikt (Ref. 7) – led to author’s
1965 precise synthetic/analytic formulation of “liquid space optics” (Ref. 8). The latter consists of
technology for fabricating ultra-low-cost, self-forming, vast, precision (diffraction-limited), unfurlable,
liquid-plastic (nominally Dow-Corning #200 silicone)-surfaced, liquid-metal (nominally pure gallium)plated, reflective (liquid “membrane”) optics and refractive (liquid) optics, in “zero-g” (orbit or
unpowered trajectory), respectively.
Today’s space technology is severely energy-limited. The present is a proposal to use the new optics to
harness the energy of stars for propulsion as well as for consumer “free” electricity, and simultaneously
to harness the mass of the cooler heavenly bodies for raw materials. Consequently it is a proposal to use
the new optics to initiate both the interstellar era in space and a quantum-leap in personal fortunes on
earth.
Page 3
2.
HIGHWAYS IN THE SKY
Cheap, affordable airflight (in mostly private vehicles) and spaceflight (initially in mostly public vehicles)
would evidently be the analogue of the historical American highway-building system, whereby in
payment for highway construction goods and services, “money” was created/floated in the “united States
of America”* in the last century, such that the u.S.A. “spent themselves into prosperity.” In a similar
fashion, palpable advantages to the public of “highways in the sky” would be a higher standard of living,
lower
prices,
abolishing
the
debt-money-linked/unConstitutionally-applied,
rigorously
unnecessary/endlessly entrapping, personal income tax (a direct tax, unapportioned, therefore needs be
voluntary or unConstitutional), high employment, dollar buys more, no inflation, affordable health care,
airflight and spaceflight for all, ever decreasing need for social programs as most American natural
persons would be able to provide for themselves (Ref’s 3,4,5 op. cit.).
The united States had a sound, productive economy – able to afford advanced invention and advanced
(aerospace and other) technology, for example – when they had a wealth-based (rather than a debtbased) money system. In that system, incidentally, implemented as well as based on gold & silver
(convenient but not necessary), money was spent into circulation, debt-free, to pay for building and
maintaining roads, a benefit to every American person.
Technology introduced here has as its main purpose the automatic collection and control/redirection of
solar/stellar energy, in narrow, coherent (laser) beams, both for civilization’s (unbounded, “free,” mainsource) utility power (therefore wealth), and also for astronautics’ (unprecedented, main-source) shuttle
interplanetary and “easy” interstellar propulsion power (therefore a wholly new era on earth is implied,
politically, economically and socially as well as technically).
The dream following may well be the one whereby we will rescue civilization, pave our streets in gold,
colonize the planets and visit the stars.
It is herein proposed that the transport (aerospace) industry should ask American natural persons to once
again instruct the united States’ government (in the same fashion as in the last century), to offer to every
American natural person (first, then to the world) free use of highways (this time, in the sky), the latter to
be paid for and maintained once again in a currency based (Constitutionally) on wealth and not on debt.
The combination of a wealth-money system with free –aerospace- highway construction, should once
again (as in the last century) point the way to riches for society –this time such as to project earth’s entire
civilization into the (literal, energy or wealth-money) “Millennium;” i.e., American and other natural
persons would spend wealth-money (not debt-money as at present) that was put into circulation to pay
for “roads” (or “highways in the sky”), nationally, globally and solar-system-wide – a benefit to every
earthborn natural person.
*so-called in the Declaration of Independence
Page 4
FIG. 1A
TO “SOLAR-CELL SEA” (DISCPLANFORM BALLOON @
80,000 FT. ALTITUDE
A
SWIVEL BEARING
NOTE I: SECTION WILL ALIGN
ITSELF WITH PREVAILING WIND
AT
THAT
ALTITUDE;
AERODYNAMIC FAIRING OFFERS
1/40TH THE DRAG THAT ROUND
UMBILICAL
OF
IDENTICAL
FRONTAL
(WIND-DIRECTION)
PROJECTED
AREA,
WOULD
OFFER
NOTE II: EACH SECTION IS
PRESSURIZED WITH HELIUM &
FITTED AT EACH END W/ A
HELIUM-TIGHT
BEARING
AROUND THE UMBILICAL,
THUS EACH SECTION ACTS
LIKE A MINIATURE HELIUM
BALLOON
NOTE
III:
BEARING-RING
WASHER
(EACH)
HARDMOUNTED TO UMBILICAL (CF.,
WELDED); THUS EACH SECTION
SUPPORTS WEIGHT OF LENGTH
OF
UMBILICAL
PASSING
THROUGH THAT SECTION
A
WEATHERVANE/AEROFAIRING
SLEEVE (HELIUM BALLOON
500’-LONG SECTION)
SECTION AA
Wind direction
SLEEVE BEARING
UMBILICAL
PACKING (HOLDS
2 CU WIRES)
CHORD
CHORD/4 = AERODYNAMIC CENTER
UMBILICAL-SUPPORT
BEARING (MOUNTED IN
BALLOON)
TO GROUND
Page 5
SOLAR COLLECTOR MIRROR IN
GEOSYNCHRONOUS EARTH ORBIT
(GEO, 22,240-MI. HIGH) OR OTHER
COLLECTION ORBITAL APPARATUS
22 mi.
LIQUID-SPACE-OPTIC
BEAMFORMER OR OTHER BEAMFORMER (GENERATES LASER OR
MICROWAVE BEAM)
LASER
OR
MICROWAVE
ELECTROMAGNETIC BEAM FROM
ORBIT (@22,240 MI.)
22,240 mi.
WAFER-SHAPED,
FLAT,
SHALLOW,
CIRCULARPLANFORM
BALLOON
(HELIUM) COVERED WITH
(“SEA OF”) SOLAR CELLS @
80,000 FT. ALTITUDE
80,000 ft.
1 mi.
80,000-FT.-LONG
UMBILICAL
TO
TAKE COLLECTED
POWER
TO
GROUND
POWER
CONDITIONING,
DISTRIBUTION
PLANT
ON
SURFACE
FIGURE. 1B. Beanstalk umbilical with aerodynamic shroud units to 80000 feet altitude
Page 6
3.
V/STOL AMPHIBIOUS AIRCRAFT RUNABOUT “FAMILY CARS”
Present technology purports to outfit all interested American natural persons (first) with equivalent
runabout “family cars of the air,” designed for operating on “highways in the sky.” These highways –built
and maintained by the traditional/historical system of creating “wealth-money” that made our nation
great (vs. current system of “debt-money”) would consist of public (1) collision-avoidance beacons, and
(2) orbital solar-energy, “free fillup,” battery-charging (for propulsion), beam-fed “solar-cell-sea” (Fig. 1)
“service stations in the sky” at 80,000 ft altitude for all “highway” motorists (Ref. 9).
The present united States debt-money system and not a lack of technology (at all) evidently is the
responsible barrier that prevents most American natural persons from owning and operating such nationwandering, globetrotting private (backyard-based) “Vertical-or-Shortfield, Take-Off and Landing” (V/STOL)
electrical aircraft, garageable, amenable to automatic control and collision-avoidance electronics,
amphibious, supersafe and highly efficient (Fig. 2).
FIG. 2.
AEROSPACE SUPERSHUTTLE: BLASTWAVE-JET-POWERED AIRCRAFT TO 80,000 FT.;
STEAMROCKET-POWERED SPACECRAFT ABOVE 80,000’
Page 7
4.
PUBLIC AND PRIVATE TRANSPORTATION TO THE PLANETS
Almost all planets and their moons, respectively, are amenable to permanent mounting of orbitalsunpumped-macrolaser-fed, solar-cell-sea, electrical-power “filling stations,” on their surfaces or in their
atmospheres. If a given celestial body has a dense atmosphere – as do earth and Venus, for example –
these “filling stations” would be tethered high in the atmosphere on special flat, wafer-shaped balloons
(Fig. 1). If a given body has a tenuous or nonexistent atmosphere – like Mars or the moon – these solarcell-sea “service stations” would be mounted on the body’s surface directly.
Electricrocket motors (ion or plasma) in non-self-contained (macrolaser-fed) propulsion systems, would
drive “ferryboat” interplanetary spacecraft, where acceleration at 1 g would be employed for half the
trip, followed by deceleration at 1 g for the other half, getting us out to most nearby planets and their
moons in two or three days.
Rocketmotor exhaust velocity for maximum “fuel” efficiency, would be adjusted to be just greater than
“ferryboat” spaceship terminal velocity at midcourse.
Purchase, ownership and maintenance of “family cars of the air,” would be the owners’ individual
responsibilities. However, in lieu of actual roadbuilding of conventional highways, free “filling stations of
the air” would be erected and maintained “full” of stored, available electrical energy (by macrolasers),
built by national contractors for wealth-money payments, traded in this historical fashion for injecting
new wealth-money into the economy.
Thus (free) “Service Stations In the Sky,” would be established on or above most planets and their moons,
and operated by the (debt-free, wealth-money-based) nation itself, as a public service (comparable to
superhighways).
5. THE BEGINNING: BEAMING POWER FROM ORBIT VIA ADVANCED (LIQUID) LASER
OPTICS
Lasers, because providing immensely tighter, narrower, longer range beams, must be preferred in the
long run, for most purposes in space, to masers, as most authors agree (Ref. 10). But laser diffractionlimited transmitter optics (just reversed, maximum-quality astronomical-telescope primary-mirror
systems in large size) have heretofore been very difficult, expensive or impossible to build –particularly
for space. In ’66, though, there appeared this author’s solution (declassified Nov. ’65 – Ref. 11), based
essentially on introducing very shallow, static, capillary-boundary-constrained, reflective, liquid-metalplated, liquid-plastic pools onto interior surfaces of “rigidized” balloons erected (inflated) in orbit (Ref’s
12, 13, 14) somewhat in the fashion of “Echo” (Ref. 15). These pools in capillary (“zero-g”) fashion, pull
themselves – their own surfaces – into precision optical mirrors (as retouched by deliberately-introduced
static masses providing self-gravitation (Ref’s 16, 17). Evidently this stable, simple, universal and
permanent approach to retouching should absorb, expand, economize and systematize the essentially
limited and expensive new “adaptive optics” method much in the news of late (Ref. 18).
Page 8
6.
MACROLASER SELF-SUBSIDY
To initiate circulation of wealth-money by creation of same (via uS’ govt’s spending it into existence to
build “superhighways in the sky” – thus using it to replace and retire today’s debt-money), this author
recommends first development of a (profit-making) “excursion-bus/excursion-boat into space” system.
Central to this scheme would be this inventor’s (proprietary) VTOL, 5,000-passenger, circular-planform
aerospace SUPERSHUTTLE vehicle (Fig. 2), evidently capable of reflexively financing its (dedicated)
MACROLASER (thereby “bootstrapping” the world’s first – general purpose – orbital MACROLASER (this
inventor’s loss leader)).
Capable of hauling sightseeing tourists to synchronous orbit and return at a suggested price of $200 (1994
“FRN”) per seat, SUPERSHUTTLE – both VTOL aircraft and beam-climbing spacecraft – would operate in
the atmosphere on “BLASTWAVE” (Ref. 19, Fig. 3), LPG-burning, jet engines, producing only water-vapor
as exhaust.
FIG. 3. PULSE-RAMJET BLASTWAVE JET ENGINE
Likewise, as spacecraft, SUPERSHUTTLE would accept energy from a synchronous-orbiting, liquid-mirror
macrolaser system, “tightened” beam (Ref. 20, Fig. 4), driving its (SUPERSHUTTLE’s) electric rockets
(“steamrockets”) which would exhaust water vapor as sole expellant.
Predicated on six excursions per day at 3 hours each (reaching Geosynchronous Earth Orbit (GEO) at
approximately 1 g continuous acceleration/braking while “leading” the target sufficiently, might require
about 1 hour, as would of course the return), SUPERSHUTTLE – if fully loaded at 5,000 passengers per
excursion – would earn about $1 million gross profit per jaunt to GEO (inflatable dollars, 1994 FRN).
Presuming net profit of one-half after expenses, salaries, spares, repairs, maintenance, life-cycle costs,
etc. are accounted for – and that perhaps half the natural persons on the planet might end up as
customers – each SUPERSHUTTLE should earn about $1.2 billion per year, indefinitely (Ref. 21), in
inflatable dollars, 1994 FRN.
Page 9
FIG. 4
1
2
4
5
7
3
8
10
6
9
11
12
13
FIGURE 4
GIGAWATT RADIATIVE OPTICAL MACROELECTRONIC (GROM) MULTI-ORBIT SOLAR MACROLASER FOR
PROJECTING FOCUSABLE BEAM ENERGY-SUPPLY TO ACCELERATING INTERSTELLAR SPACECRAFT.
At the extreme left of the figure is the great solar-collecting mirror (60 miles in dia.), (1). CO2-type
macrolaser is shown symbolically, although in actual fact, a free-electron-type (variable-wavelength in
near-infrared and perhaps 4-5 times more efficient) laser will be used. The great solar-collecting mirror
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 “lases”, i.e., transmits a coherent beam normal to the rod’s endsurface mirrors. 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 (1-mile-diameter) liquid-surface primary mirror (3). Highprecision “primary” (3) is bordered by a rigid plastic-foam boundary-ring (2). Laser energy (5), focused by
reflection from 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).
Page 10
6A. STARFLIGHT.
Starflight propulsion is surely the anticipated acme of all (proposed human) experience, ably introduced
by Mallove and Matloff in their “Starflight handbook” (Ref. 22). These authors say, “The big problem with
starflight is, of course, distance, and that is why the bulk of this handbook is devoted to methods of
interstellar propulsion. The ancillary problems of guidance and navigation, payload content, reliability,
and so on, though difficult, are relatively minor issues compared with the primary hurdle of attaining
speed (or its equivalent, sufficient energy source and sufficient specific impulse – JHB) sufficient to reduce
to tolerable lengths transit times to the stars… Starflight is not just very hard, it is very, very, very hard!”
Right on, Gene and Greg!
Considering the history of interstellar propulsion investigations, the pioneers and their critics have had
this to say:
Von Hoerner, Sebastian, concludes (1963) “…Space travel, even in the most distant future, will be
confined completely to our own planetary system, and a similar conclusion will hold for any other
civilization, no matter how advanced it may be…” (Ref. 23); Purcell, Edward, says (1963): “Let us consider
taking a trip to a place 12 light-years away, and back… It is preposterous. And remember, our conclusions
are forced on us by the elementary laws of mechanics… All this stuff about travelling around the universe
in space suits – except for local exploration… belongs back where it came from, on the cereal box.” (Ref.
24).
Marx, G., concludes (July 2, 1966) “…The laws of conservation of energy and momentum (do not
necessarily) forbid the visiting of other planetary systems in the human lifespan” (Ref. 25); Bloomer, J.H.
propounds (Oct. 13, 1966) “…We can send a probe to Alpha Centauri before 2000 A.D. Shortly following
the return of data from an Alpha Centauri probe, we may feasibly plan to send astronauts on round trips
to this…and…other…stars…before 2100 A.D.” (Ref. 33); Dyson, Freeman J. says (October 1968) “…I predict
that about 200 years from now, barring a catastrophe, the first interstellar voyages will begin” (Ref. 26);
Norem, Philip C., says (June 1969) “An overall system employing a vehicle with a solar sail, a large laser
array, and employing interactions with the interstellar magnetic field…is…capable of roundtrip
interstellar…journey of 75 to 150 years…” (Ref. 27); Forward, Robert L. says (1976): “…a more practical
method would be to carry along some mass to use as propellant and just use the laser beam to energize
the propellant…” (Ref. 28); Jackson&Whitmire say (1978) “The laser rocket…can also accelerate directly
into the beam since the ship uses the laser beam only as a source of energy and not momentum” (Ref.
29); Weiss, Pirri and Kemp say (1979): “In…(one way of)…beaming energy to a propulsion system with a
large power laser…a thermal rocket system uses energy beamed from a remote laser to heat a chemically
inert propellant” (Ref. 30).
Oliver, B.M. says (Oct. ’87): “…interstellar travel…if such travel is to be accomplished in a human lifetime,
the energy requirements are enormous…Our intent is to dispel the notion…that progress in rocket
technology can reduce by orders of magnitude the energy needed for interstellar flight…” (Ref. 31).
Page 11
We’ll certainly see that Barney Oliver was right about that! However, to reduce what I regard as his and
others’ unwarranted trepidation over “energy” (or the lack thereof), I submit that adequate (nuclear)
energy apparently is available to us after all, but just not from the (domestic, terrestrial) source we’ve set
our hopes on ever since the dramatic demonstrations of the ‘Forties!
Prime rocket figure-of-merit is of course “specific impulse”. Specific impulse in seconds is conveniently
calculated as a function of exhaust velocity in meters/sec by dividing latter by acceleration of gravity (at
earth sea-level), about 10 m/sec2 (Ref. 22, op. cit.).
Hermann Koelle’s reliable text on astronautical engineering (Ref. 32), makes very clear that for a
reasonably sized (i.e., single-stage) interstellar rocket, the exhaust velocity must slightly exceed the
desired terminal velocity of the rocket.
But if the latter is – nominal minimum for interstellar flight – say, one-quarter the speed of light, this
implies a specific impulse of about 7,500,000 seconds.
7.
TRANSCENDING “CHINESE” ROCKETRY
Principle of “Alpha Centauri Probe” (Ref. 33) was to transcend “Chinese rocketry” (or conventional
rocketry, in which all mass and all energy are carried on any given ship), by instead supplying all energy to
an interstellar rocket via starpumped (macro-) laserbeam. The latter is dispatched steadily (continuouswave) from afar (from orbit of earth, sun or star). The energy used to pump the laser would have been
furnished by a very large, associated nearby, stellar collector mirror.
Such a concept required a new optics and a new kind of rocket motor.
Conventional optical technology applied to the problem of fabricating the one-mile-wide and larger,
precision mirrors needed in orbit, would rapidly exhaust all the resources of all the countries on earth
even just trying to build such optics on the ground, let alone trying to build them in orbit (which would of
course be orders of magnitude more expensive).
No less discouraging would have been the attempt to apply conventional rocketry to interstellar flight.
Conventional rocket motors will not even approach an Isp of 7,500,000 seconds. For examples, (Ref. 22,
op. cit.), Hydrogen-Oxygen Rockets deliver Isp = 460 sec; Oxygen-Beryllium-Hydrogen Rockets would
deliver 705 seconds; ion rockets have been built to deliver specific impulses from 2,500 sec to 10,000 sec,
and at a theoretically very best might deliver ultimately 400,000 seconds. Fission nuclear rockets,
depending on core, deliver 500-700 sec. Fusion nuclear rockets would deliver 2,500 to 200,000 sec.
The absolute destruction of matter is in theory the only reaction more energetic than fusion; yet putting
aside the engineering problems (many of which have been solved by Robert Forward and others – Ref. 22,
op. cit.), we could still hope to derive from this process only about 140 times the energy available in
hydrogen fusion. This would not be nearly enough (Ref. 24, op. cit.), to both send the astronauts out and
get them home – alive.
Page 12
By why do the exhaust particles have to be accelerated by a (chemical or nuclear) reaction (aside from ion
rocketry)? What’s to keep us from using an actual laboratory particle accelerator as rocketmotor (and
supplying the energy from afar by liquid-optic macrolaserbeam)?
8.
NEW FREEDOM FOR SPECIFIC IMPULSE: STARPUMPED MACROLASERS DRIVING
LASERPOWERED REMOTE (PARTICLE-ACCELERATOR) ELECTRICROCKET MOTORS
(LREM’S)
If the particle-accelerator is (perhaps best –i.e., lightest) a linear accelerator (Ref’s 34, 35, 36), then we
might be deterred by: “One of the chief drawbacks of the linear accelerator for acceleration to very high
energies is its enormous power consumption…For example, a 100-gap, 100 MeV accelerator in which the
operating frequency is 200 Mc, the stable phase angle is 60 o, and in which gaps of 5 f capacity are
resonated by circuits having a Q of 50,000 will consume approximately 7 megawatts of power!”
However, LIQUID-OPTIC ORBITAL MACROLASER chief advantage – along with advantages of narrow beam
and advancing focus, is perhaps its truly enormous and practically unlimited CW power output (derived,
after all, straightforwardly from the sun or other star). Hence linear accelerators (either electron or
proton type) may be ideal choice for relativistic (interstellar) single-stage rocketmotors. As particles are
accelerated, their velocity vp, increases according to the relation
 
 
vp  
1
 1  
T
c
 1 
 
M0 c 2








2 1 / 2






,
where T is the kinetic energy of the particle being accelerated and M0c2 is its rest energy, c being the
velocity of light (Ref. 35).
“For T much greater than the rest energy, ,
vp
c
～1 .
(Ref. 35, op. cit.)
“Preaccelerators for linacs are often high-voltage dc or pulsed-dc accelerators, operating at from a few
hundred kilovolts to a few million volts. Protons and heavier ions have rest energies (masses) of one to
many giga electron volts so that they emerge from the preaccelerator with velocities only a small fraction
of c. Electrons have a mass of 511 keV so that even a preaccelerator of only 80 keV boosts their velocity
to 0.5c. The arrangements needed for efficient generation of accelerating waves depend markedly on the
desired wave velocity and the designs of linacs for protons or heavier ions and for electrons differ
markedly.” (Ref. 35, op. cit.).
Page 13
9.
SPACE-BASED SOLAR/STELLAR MACROLASERS
Suppose as noted our optics are fabricated of liquid surfaces in orbit (or unpowered trajectory), using
capillary (surface tension, adhesion, cohesion, contact-angle) effects according to the principles of liquid
space optics.
As has been shown in numerous studies (Ref. 21, op. cit., 37), such systems can “bootstrap” themselves
into orbit, under their own power, in a modular fashion, providing their own financing by selling collected
and redirected (laserbeamed) solar power.
Principle here is that solar/stellar energy is to be accessed in practically unlimited, desired amount in
orbit, by conventional aluminized-mylar (paraboloidal) solar-collector mirrors (say, 60 or more miles in
diameter – Ref. 33). The latter are focused on (or “pump”) a “liquid-optic” multi-orbital (probably CO2)
laser system, which converts – at about 1% efficiency – to a highly “tightened” (narrowed inversely
proportional to primary diameter) laserbeam.
Primary diameter may be one mile or more. “Diffraction-limited” quality of such a beam was shown by
present author feasibly to obtain, by application of a treatise of Dr. Ta Li (Ref’s 6, 8 op. cit.), also by
practical limitation of the cumulative residual accelerations that the (liquid-surfaced) optics are subject
to.
Essentially, (power-beaming) solar (or more generally, stellar) macrolasers derive from a suggestion of
C.H. Townes & R.N. Schwartz (Ref. 20, op. cit.), laser co-inventors, on interstellar communication. Townes
and Schwartz recommended placing your laser to “fire” backward at the focal plane (or surface) of an
astronomical telescope, whereby the (reversed) instrument projects a coherent beam with spread
inversely proportional to primary diameter (if latter is “diffraction-limited” – i.e., theoretically
unimprovable – or correct to the Raleigh limit of /4 at every point on its surface for the projected laser
light.)
Cost of fabrication of such given-quality – say diffraction-limited – (solid) optical surfaces (mirrors or
lenses) on the ground, is about proportional to the fourth power of the primary (principal) optic diameter.
Thus an earthbound observatory featuring a diffraction limited 30-inch-diameter (about the maximum
practical diffraction-limited instrument on earth) glass greenlight reflector, and costing, say, $10 million
(’94 FRN), would cost about $2 × 108 trillion if enlarged to one mile in diameter while maintaining the
same surface quality (or about $200 billion trillion on earth by this rule.)
Consequently other approaches such as computerized, segmented, sensor-feedback, “adaptive optics”
have been taken for fabricating large-diameter, astronomical-quality, solid-optics (usually reflectors) on
the ground (Ref. 18, op. cit.). However, “adaptive optics” complexity and cost limitations are still at best
evidently within perhaps only one or two orders of magnitude of the older, conventional art, which would
indicate still a minimum cost of about $2 billion trillion for a one-mile-dia. primary, diffraction-limited,
solid mirror on the ground.
In space, as demonstrated recently by the “Hubble Telescope” example, conventional (solid) optics’ costs
are going to be multiplied again, say by a factor of 20 to 200 times, indicating there is no older technology
Page 14
known today which could practically deliver, say, a 1-mile-diameter greenlight diffraction-limited
astronomical telescope mirror in orbit.
But it is precisely this order of primary mirror size, quality and location (space) that’s proposed by the
present inventor, for a Townes-type “power-beaming” (not communication), inverted astronomical
optical telescope in orbit. Latter would rely on author’s ’65 invention of “liquid space optics” [rejected by
USPATOFF in ’70 as “obvious”, oblivious to the earlier (’65) threat by USAF against inventor of fine and
imprisonment for disclosure of this “obvious” invention; USAF lifted its “secrecy order” when “inventor”
showed “invention” had already been disclosed to every major library on earth anyway (Ref. 11, op. cit.)].
Other element surfaces (i.e., secondary mirror, laser cavity-end reflector mirrors) of this (multi-orbit)
optical system, will also be liquid, as will surface of the primary mirror itself.
10.
CHARACTERISTICS OF LIQUID SPACE OPTICS
Outlines of “liquid space optics” and “solar energy handling satellites or spacecraft” (SEHS) technology
have been sketched by the present author in six earlier publications (Ref’s 8, 9, 17, 21, 33, 37), to which
the reader is referred for general introduction.
General principles are those of permitting capillary forces ruling liquids in “zero-gravity” (orbit or
unpowered trajectory) to do the work of fabrication of large astronomical optical surfaces – mirrors or
lenses. Above studies have shown surface errors locally and about the boundary can nominally be held to
less than those at the threshold of diffraction limited resolution for visible/IR/UV optics even on the order
of a mile or more in diameter. Quality of liquid optical surfaces already is ideal, a perfection which all
lapped and polished solid optical surfaces can only approach. Optical substrates in orbit or unpowered
trajectory can be expanded and rigidized from very compact and convenient packages according to wellknown military and other tested techniques.
Energy of the sun and stars can be collected, manipulated and redirected in any practical quantity, and
with ultimate precision, by sufficiently large, astronomical-quality, zero-gravity, capillary or
“epihydrostatic” optics. Both civilization itself and advanced astronautics depend critically on energy. That
energy is superabundant in space. Collecting, handling and redirecting solar/stellar energy is perhaps the
only bar both to advanced civilization and to advanced astronautics. It is believed joint invention of the
laser and of liquid space optics, may offer the desired solution for both, one that has been obscured
heretofore by excitement over discovery of nuclear physics.
For diameter size of astronomical optics there is no substitute. However cost of given-quality (e.g.,
diffraction-limited) such optics on earth goes up, as noted above, by the fourth power of the diameter. In
diffraction-limited case, this is due to difficulty of grinding a given solid surface to a /4 maximum
tolerance overall while polishing it sufficiently to result in a uniform “Beilby” layer or surface consisting of
irregular collections of small peaks or pits none of them more than a molecule or two in height or depth.
In the formation of such polished layers on solids we have the natural phenomenon corresponding to the
smoothing out of liquid irregularities under “surface tension”. For clean circular boundaries at fixed
Page 15
azimuthal inclination and zero constant acceleration therefore a sufficient amount of clean, liquid when
introduced will by energy principles exhibit a concave or convex spherical surface with fixed contact angle
at the boundary.
Such surfaces are very nearly what is needed for (reversible) astronomical telescope primary and
secondary mirrors. Contemplated is introducing a liquid plastic, which may or may not remain liquid
depending on application and “plating” it with a liquid metal, which may or may not remain liquid
depending on application. Probably best liquid plastic is Dow-Corning #200 silicone and best liquid metal
is ultrapure gallium, although many experiments are needed.
Astronomical reflective primary and secondary optics are generally spherical (Maksutov) or nearly
spherical (Cassegrainian or Newtonian, i.e. needing only a slight retouching to generate required
paraboloidal or hyperboloidal surface). Using an expression derived from Ta Li’s General Dynamics (’60,
San Diego) formulation of deviation of (liquid-solid circular-boundary) liquid surfaces from a spherical
liquid-vapor interface under given net axial acceleration, it is possible to show – for nearly-unaccelerated
space systems in thermodynamic equilibrium – that we are facing a standard isoperimetric problem (with
mobile upper limit) in the calculus of variations. Couching the rapidly-converging series solution (given by
Li) to the resulting differential equation in both physico-chemical and optical constants, one can show
arbitrary (integrated) “figuring” gravitational acceleration may be practically applied to such (reflective)
liquid surface by deliberately “sculpting” and emplacing required axially symmetric masses behind the
optics (in “zero-gravity” – i.e., orbit or unpowered trajectory in space). See Fig. 5.
A “liquid space optical” (power-beaming) orbital macrolaser apparatus, relies essentially (as above) on
introducing very shallow, static, capillary-boundary-constrained, reflective, liquid-metal-plated, liquidplastic “pools” onto interior surfaces of “rigidized” balloons erected (inflated) in orbit (Ref’s 8, 12, 13, 14
Op. Cit.), somewhat in the fashion of the “Echo” balloon satellite (Ref. 15, Op. Cit.). See Fig’s 6, 7.
These “pools” in capillary (“zero-g”) fashion, pull themselves – their own surfaces – into precision optical
mirrors or lens surfaces [as retouched be deliberately introduced static masses adjacent to and behind
the optics, providing “self-gravitation” for figuring (Ref. 8, Fig. 5)].
11.
MATHEMATICO-PHYSICAL FOUNDATIONS OF LIQUID SPACE OPTICS
Final development of the ultimate general mathematical tool of “zero-gravity epihydrostatical optics,” is
due to a treatment by Dr. Ta Li (Ref. 6). The following development by Li, illustrated in Figure 8, is
characterized by liquids in containers whose interiors are in the form of surfaces of revolution.
lv, ls, and sv are respective surface tensions at the liquid-vapor, liquid-solid, and solid-vapor interfaces,
in dyne/cm; in the plane of Fig. 8, the “horizontal boundary” is the point intersection of the surface
(liquid-vapor interface) with the toroidal annulus shown. u = u(x) is the vapor-annulus interface, v = v(x) is
the liquid-annulus interface, and z = z(x) is the liquid-vapor interface, where u=u(x) and v=v(x) are known
and z=z(x) is sought.
1
6
2
7
8
3
10
4
9
5
12 13
11
27
34
14
25
16
15
17
20
29
19
18
26
28
30
32
33
31
22
FIG. 5. ONE-MILE DIAMETER VARIABLE-FOCUS DIFFRACTION-LIMITED ORBITAL PRIMARY MIRROR (BEGIN CENTRAL SECTION)
23
21
24
Page 16
Page 17
FIG. 5. One-Mile-Diameter Variable-Focus Liquid-Surface Diffraction-limited Orbital Primary Mirror
(Conclude)
1. Solar ray impinging on 60-mile-diameter collector mirror;
2. Subliming-propellant microrocket for 1-mi.-dia. mirror attitude-control;
3. Extreme possible position of spline; spline is attached along its length to the skin of the 1-milediameter bellows;
4. Accordion-pleats around circumference of bellows;
5. Spline (attached to bellows) in opposite extreme position;
6. Solar ray after reflection from solar-collecting mirror;
7. Envelope of final focused high-energy coherent beam;
8. Short-focus surface of 1-mile-diameter liquid metal “pool”;
9. Instantaneous surface of 1-mile-dia. liquid “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 of “figuring” liquid metal mass, constrained by bellows;
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 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-collecting mirror;
26. Subliming-propellant microrocket for solar-collector attitude control;
27. Interior of bellows figuring-liquid mass for extreme short-focus configuration of liquid mirror;
28. Bottom surface of enclosure for 1-mile-diameter liquid metal mirror “pool”;
29. Membrane structure of bellows;
30. Envelope of laser beam which illuminates primary mirror;
31. Solar ray before reflection from 60-mile-diameter concentrating mirror;
32. Envelope of final focused high-energy coherent beam;
33. Solar ray after reflection from solar-concentrating mirror;
34. Long-focus position of liquid mirror surface (for interstellar-rocket laser-beamed supply of solar
energy continuously for acceleration to 1/4th speed of light).
Page 18
FIG. 6. Liquid surface in Lowered Gravity
N = VERTICAL LOAD IN g’s
θ = CONTACT ANGLE
N
FIG. 7
LIQUID
CHARACTERIZED
BY:
0o CONTACT
ANGLE
45o CONTACT
ANGLE
(AFTER PETRASH & OTTO)
90o CONTACT
ANGLE
(VAPOR)
(LIQUID)
1-G
135o CONTACT
ANGLE
(VAPOR)
180o CONTACT
ANGLE
(LIQUID)
CYLINDRICAL
TANK
ZERO-G
Page 19
FIG. 8. Liquid Optic in Zero - Gravity
FIG. 9. Error Due To Finite Load (Begin)
= R3/2
1 +2
Page 20
FIG. 9. Error Due To Finite Load (Conclude)
R3
1
2
( g0 = SEA LEVEL GRAVITY )
The total potential energy K of the system consists of the sum of surface energies and the effective (net)
central-force-field or contact-force potential energy, where
(1)
x1
x1
x1

1
/
2
1
/
2


K  2  lv   1  z 2 
xdx    1  u2 
xdx     ng 0  z 2  v 2  xdx


0
0
0

while thermodynamic equilibrium implies:
(2)
V

2
x1
 u  z xdx
,
0
where:
 = ratio of vapor volume to container volume
V = Volume of container
g0 = sea-level acceleration of gravity = -980.665 cm/sec2
=
 sv   ls
 lv
 = (liquid - vapor),
C=

 lv

 liquid   vapor
 lv
liquid = density of fluid
Page 21
vapor = density of vapor above liquid
Above equations (1) and (2) define an isoperimetric problem in the calculus of variations with a mobile
upper limit. The total energy K is to be minimized and the resulting differential equation (the “solution”)
will itself be solved to provide a fundamental design formula, or expression. This expression will be critical
both in the analysis of systematic errors engendered by net residual acceleration (all sources combined),
and in the synthesis of diffraction-limited optical surface figures by designing masses (shapes, sizes and
positions) which will deviate the optical surfaces slightly by gravitational attraction.
Minimization of K can be facilitated by introduction of the “Bond Number” (dimensionless physicochemical constant characteristic of the system):
(3)
ng 0  R 2
B0 
,
 lv
where:
n = load factor = sea level g’s of acceleration operating on system
R = radius of curvature of liquid-vapor interface.
12.
LI’S DIFFERENTIAL EQUATION
Applying standard techniques of the variational calculus, Li found the differential equation
(1)

d  d

d  d

2

 d  
1   
 
d



 
1 / 2 

  B0


,
where:
(2)
 
1
2R 
 z  z 0  

R 
B0 
,
and
(3)
  
x 

 2R 
2
.
Li’s solution for ordinate (height) z(x) of the meniscus surface at a lateral distance x (abscissa) from the
vertex, was the infinite series,
2
(4)


4
6
1
1
20
1
 x 
 x 
 x 
z  z(0)  2    4  B0     12 
B0  B0 2     ...  f (x)
2
3
3
6
 2R 
 2R 
  2R 
.
This equation is the required solution to the characteristic (Euler) differential equation of the given
variational problem.
From Figure 8 and 9, we define
Page 22


  z 0 x 1   z 0 0   z B x1   z B 0 
(5)
0
0
,
, where:
z0(x1) = ordinate of the (spherical) liquid surface at x = x1 when Bond No. B0 = 0  n = 0 or “zero
gravity”; see Fig’s 8 and 9.
z0(0) = 0 = ordinate of the vertex (at x = 0) of the (spherical) liquid surface when Bond No. B0 = 0 
n = 0 or “zero gravity”.
zB x1   ordinate of the (non-spherical) liquid surface at x = x1 when B0 ≠ 0  n ≠ 0 or “nonzero0
gravity”.
z B 0   ordinate of the (non-spherical) liquid surface at x = 0 when B0 ≠ 0  n ≠ 0 or “non-zero0
gravity”.
We also define:
R0 = radius of curvature of unperturbed liquid surface,
x1 = radius of the meniscus “aperture”,
 = f/number = focal ratio = F/D,
F = focal length,
D = aperture,
g = ng0 ,
RB = radius of curvature at the vertex of the (non-spherical) surface of revolution
0
obtained when n (and therefore B0) is finite.
Note that R0 is defined as the radius of curvature of the circular cross section when B0 = 0, and RB is
0
defined as the radius of curvature at this vertex when B0 (which is proportional to the acceleration of the
system) is a very small quantity indeed. Hence, for the systems under consideration, which are
accelerated only by such tiny forces as gravity gradient, self-gravitation and solar wind, the difference
between R0 and RB is an infinitesimal of yet higher order.
0
Neglecting this infinitesimal we may write RB = R0 .
0
Substituting the latter (close) approximation in (5) we find
(6)
where:
4


 x 
1

  ...
  R0 4  a1 

2

 2R0 


,
Page 23
(7)
a1 = 4 + B0
,
Now the ai are functions of the dimensionless parameter B0 and the ratio x/(2R0) is also dimensionless
since x and R0 are measured in the same units. Hence the units of  are the units of R0.
Combining (6) and (7) and taking the deviation  positive we have from the first non-zero term of this
rapidly convergent series:
(8)
4


 x 
1



  R0 B0 
 ... 
 2  2R0 



.
Substituting for the Bond Number, B0, from (3) of the last section, reducing and neglecting higher-order
terms, this becomes:
4
 1   n g 0 x 1
 lv R 0
2
5
(9)
  
.
If the liquid meniscus surface is considered an optical surface (reflector or refractor) we may redefine the
parameters of  as follows:
R3 = R0
= radius of curvature of an optical mirror or lens,
F = R3/2
= focal length for paraxial rays striking an optic of curvature radius R3,
D = 2x1
= aperture of optical system,
 = f/no
= focal ratio = F/D =
1  R3 
.
4  x1 
Substituting expression for x1 derived above, (9) becomes:
4
 1   n g 0 D / 2
 lv R0
2
5
(10)
  
.
Further focal ratio is:
(11)
=
1  R3  1
=
4  x1  4
 R0  R0


 D / 2  = 2D ;


Hence, substituting (11) in (10) and substituting for
(10) becomes:
10
(12)
3
D
1
    ng 0 C
.

2
 liquid   vapor
 lv
from (2) of the last section above,
Page 24
13.
THE FUNDAMENTAL (MATHEMATICAL) EXPRESSION OF LIQUID SPACE OPTICS
We may define (12) of the section above as “the fundamental expression of liquid space optics,” equally
useful for analysis and synthesis.
From (12) in last section above, note that, as you would expect, the error imposed by the net axial
acceleration, is inversely proportional to both liquid-vapor surface tension and f/no , where telescopes
with large f/no’s already have relatively flat primaries. Hence they are going to be, if liquid, less
susceptible to net axial acceleration, which is felt merely as a flattening influence anyway.
Note also from (12) above, the error will be directly proportional to net axial acceleration, ng0, and not
surprisingly will also be proportional to the third power of the liquid space optical system’s primary
diameter.
Distributing signs (removing parentheses) in Equation (5) in Section 12 and going back to Figure 7, note
that flattening at the circumference of liquid optic under a net downward acceleration, is reflected by
positive quantity  1  z 0 x 1   z B x 1  , whereas concomitant flattening at the vertex of the liquid optic
0
under the same net downward acceleration, is reflected by positive quantity,  2  z B 0   z 0 0  , such
0
that, finally, combining (summing):
10
(1)
3
D
1
 = 1 + 2    ng 0 C
,

2
Equation (1) above is plotted with f/no as parameter in Fig. 10. Note the graphed decreasing error  with
larger f/no’s. Similarly plot of Figure 11 shows  exponentially increasing with increased aperture D,
whereas plot of Figure 12 shows error vs. axial load ng0 with liquid-optic material as parameter.
14.
CHARACTERISTICS OF SEHS-POWERED “NONSELF-CONTAINED” VS. “CHINESE”
ROCKETRY
Using solar or stellar energy, astronomical and laser-optical principles, it becomes practical by
powerbeam-from-orbit “bootstrapping” (and exponential self-construction) of ultimately vast optical
systems in space, to pinpoint-transport almost arbitrarily large solar/stellar-derived energy supplies at
lightspeed to a fixed or moving receiver virtually anywhere in a solar/stellar system or outside in nearby
extrasolar/stellar space. Conventional cost yardsticks meantime, derived from solid optics on the
terrestrial surface, become totally irrelevant. Such “macro-optics” in prospect rapidly become selfsupporting, then supply a vast and exponentially increasing surplus quantity of valuable, marketable,
(“free”) energy, as desired, to specified locations on a planet or in nearby interplanetary space for the
indefinite future.
Available energy and raw material resources of a star and associated planets are effectively limitless.
Proposed technology evidently therefore could result for example here in a literal earthly “millennium”.
Similarly shuttle service to all local planets easily should be effected by such technology at one-gravity
Page 25
acceleration/deceleration or for example average of fortnightly (two-weekly) roundtrip travel times to
Moon, Mars, Venus, Jupiter, etc.
δ (IN UNITS OF 10-2 ) AT 10-12 g
FIG. 10
DATA FOR LIQUID GALLIUM AT 86o F
(C = 0.0083 sec2/cm3)
FIGURE 10: Error δ vs. f/no. for several apertures at 10-12 g load.
Finally roundtrip interstellar flight employing first the energy of the local star then that of the approached
star, becomes first feasible then exponentially faster and easier as permanent, exponentially self-grown
“stellar energy handling satellites” are emplaced respectively in orbit both of the “home” and of the
target star. Relativistic interstellar-spaceship effects become marked as SEHS size (therefore energybeam
“tightness” –range- and output) exponentially is increased. Thus roundtrip interstellar trips will become
increasingly faster (closer to lightspeed) for the travellers, but will concomitantly result in differential
(greater) aging of relatives/friends on the home planet (as predicted by Special Relativity).
Page 26
Figure 11
10-11
10-10
1
10
10-9
10-8
10-7
N = AXIAL LOAD (g’s)
(F/1)
(F/2)
100
(F/1)
(F/2)
10-1
(F/1)
R3 = 300 IN.
D = 150 IN.
R3 = 300 IN.
D = 75 IN.
R3 = 200 IN.
D = 100 IN.
R3 = 200 IN.
D = 50 IN.
R3 = 100 IN.
D = 50 IN.
(F/2)
R3 = 100 IN.
D = 25 IN.
10-2
10-3
10-11
10-10
10-9
10-8
10-7
FIGURE 11: Error δ vs. axial load N for several liquid gallium mirrors
Design of particle-accelerator rocketmotors may draw on airborne applications as in particle-beam (“Star
Wars”) weaponry research. Grossly reducing the great weight of the laboratory particle-accelerator
apparatus should be straightforward, as there need be no scientific application/conditioning whatever of
the output beam.
15. THREE DAYS TO MARS BY MACROLASER-SUPPLIED 1 g “FERRYBOAT”
Mars distance from earth (Ref. 42) ranges from 34.5 × 106 miles to 235 × 106 mi. Therefore average Mars
distance is 84 × 106 miles.
Half of Mars average distance is 42 × 106 mi.
Page 27
Figure 12
10
10-13
1
10-12
10-11
10-10
10-9
N = AXIAL LOAD (g’s)
100
SILICONE OIL,
TOLUENE,
MERCURY
PHENOL
10
-1
GALLIUM
POTASSIUM
10-2
10-3 -13
10
10-12
10-11
10-10
10-9
FIGURE 12: Error vs. axial load for 300 inch F/I optics of various liquids
Therefore at an acceleration (constant) of one gravity, the time required to reach half the average
Martian distance is 1.174 × 105 seconds = 1.4 days.
Total time, therefore, to reach Mars at its average distance from earth, will be 2.8 days (at a one-gravity
acceleration for the trip’s first half, followed by a one-gravity deceleration for the other half).
Velocity at midcourse of this Martian voyage, will be 3.78 × 106 ft/sec = 2.58 × 106 mph = u.
Exhaust velocity, therefore, to earn best or lowest (single-stage to midcourse) ratio of initial mass
(includes expellant) to payload mass is also about
v = u = 3.78 × 106 ft/sec.
Specific impulse, therefore, of expellant needed is:
Page 28
(Sp. Imp.)
=
Mars Ferryboat Expellant
3.78 × 10 6 ft/sec
32.2 ft/sec 2
= 117,400 sec.
This is perhaps a reasonable specific impulse for an ion rocket, as Mallove & Matloff note (Ref. 22, op.
cit.), “ion rockets… at a theoretically very best might deliver ultimately 400,000 sec.”
Next, let (final leg, deceleration) payload mass for the Mars Ferryboat be 1,400,000 lb (arbitrary), while
combined mass of ion rocket engine and associated local (SEHS beam) collector is also 1,400,000 lb (also
arbitrary).
Initial (earth take-off) “cocoon” (“2nd Stage” of sorts) mass is 1,400,000 lb for the payload, plus another
1,400,000 lb for the combination of rocketmotor and associated local collector mirror, plus 24,220,000 lb
for the expellant = 27,020,000 lb. total. This entire combined mass is of course the original, launch-fromearth “payload”, or “cocoon”.
Let initial ion motor plus associated local SEHS beam “secondary” collector mass, to accelerate the
“cocoon”, also equal 27,020,000 lb. Then (Ref’s 32, 33) total rocket initial mass (at earth take-off) will be
(19.3) × 2.7 × 107 lb = 5.21 × 108 lb.
Also note (Ref. 32, op. cit., Section 21-2) that approximately
v   , under these circumstances, where:

P
= specific power,
Mp
where:
P = electric propulsion power output
Mp = mass of the thrust-producing system.
Acceleration (constant at 1 g) of the 5.21 × 108 lb rocketship from earth out to 42 × 106 miles, at an
exhaust velocity about equal to midcourse velocity, requires that:
u = v = 3.78 × 106 ft/sec at midcourse.
Also time duration of acceleration (or deceleration) per leg is:
 = 1.174 × 105 sec = 1.4 days.
Then, to calculate SEHS total beampower needed to follow the Mars-going rocket continuously (and
supply all power both for acceleration and deceleration):
Exhaust Velocity, v =
P

Mp
,
where:
v = 3.8 × 106 ft/sec
Mp = 2.702 × 107 lb
 = 1.174 × 105 seconds.
Page 29
Thence:
P = 4.506 × 1012 kW.
Next note from Section 23 below, that the total “permanent” (solar-orbiting) MACROLASER output power
(nominally for interstellar work) will be: P = 1.9 × 1014 kW, such that the proportion of this power needed
for Mars “ferryboat” propulsion will be only about 4%.
Finally passenger density on the 1,400,000 lb Mars ferryboat needs to be determined, to estimate
number of passengers.
Presume that ten times the empty-wgt. density of the Boeing 747, for example, will be needed,
considering extra safety and survival gear that must be included for a Mars (3-day) voyage.
Boeing 747 empty-wgt. density is (max.) 352,711 lb empty weight per 500 passengers = 705 lb/passenger
(Ref. 39).
Mars ferryboat density if ten times greater, would yield 7,050 lb empty weight per passenger, or a limiting
number of passengers of 200 passengers per trans-Martian 3-day “ferryboat”.
16.
COLONIZING MARS: CELESTIAL ENGINEERING TO CREATE BREATHABLE O2
First, note that, “If all the permafrost under the Martian surface were melted and the topography
smoothed out, an ocean several hundred meters deep (about 1000 ft deep?-JHB) would be created,
covering the whole planet” (Ref. 40). Many independent sources in theory agree (Ref. 41).
Let’s leave half the frost (for conversion to ocean water later on), and concentrate on converting the
oxygen from the other half to free (breathable) oxygen for air (we’ll have to get the nitrogen elsewhere –
from another planet or its satellite?)
Mars’ radius is about 3390 km (Ref. 42) or 2106.5 miles (or about half of earth’s radius, which is 4000
miles).
Then half of the Mars (equivalent) H2O volume would be 7.8 × 1017 ft3.
In the electrolysis of water (Ref. 43), 96,500 coulomb will deposit one oxygen chemical equivalent (Ref.
43, op. cit.) =
16 gm
= 8 gm.
2
The 1640-mile-diameter, “permanent” MACROLASER, orbiting the sun at ¼th A.U. (from Section 23
below, as designed for propulsion), will deliver a laser output beam of 1.19 × 1014 kW from a 27-mi.-dia.
diffraction-limited primary mirror. Presume this beam can be applied (its diffused focus) on Mars to an
“electrolysis cell” delivering a current of 11.9 × 1016 coulombs/sec at one volt.
Time, then, to release all the oxygen on Mars from a planet-wide ocean 500 ft deep (by electrolysis) will
be 6.8 years.
Total weight of air on earth (Ref. 43, op. cit.) is 6 × 1015 tons, yielding at 23% oxygen by weight, about
1.317 × 1015 tons of O2 above earth.
Page 30
Total weight of released oxygen on Mars, then, will be ～5.27 × 1018 lb, whereas on earth it is fixed at
about 2.76 × 1018 lb.
Thus is 6.8 years of electrolysis, since Mars surface area is only one-quarter that of earth, while about
twice as much free oxygen would have been released above Mars as exists above earth, we would have
created eight times the oxygen per unit topographical area above Mars as earth.
Consequently to create O2 per unit area above Mars equal to that above earth, only 6.8 yrs./8  10
months of electrolysis would be needed by the 1640-mi.-dia. “permanent” MACROLASER circling the sun
in a ¼th A.U. orbit.
Conveniently then, about 15/16ths of the original Mars permafrost mass – presumed originally equivalent
to water 1000 ft deep – would be left for the purpose of manufacturing oceans on Mars.
17. CELESTIAL ENGINEERING: CREATING OCEANS ON MARS BY MELTING PERMAFROST
Water heat of fusion is 144 Btu/lb, whereas 1 Btu = 0.2931 Watt·hr.
If one-sixteenth of Mars’ permafrost bed (total presumed equivalent to a planet-wide ocean 1000 ft
deep) is converted to breathable oxygen (and hydrogen – see last section), then there would remain
1.4625 × 1018 ft3 of (equivalent) H2O on Mars (frozen).
Fusion of all this ice on Mars (presumably initially at 0o C) will require 3.851 × 1018 kW·hr of energy.
Now the 1640-mi. dia. “permanent” solar MACROLASER orbiting the sun at ¼th A.U. (from Section 23
below, as designed for interstellar propulsion), will deliver (@1% efficiency overall, at its focus),
1.19 × 1014 kW.
Total time, then, needed to hold this MACROLASER beam on Mars to melt 15/16ths of a bed of
permafrost equal to a depth of 1000 ft of water (ice) planet-wide, would be (not accounting for
interposition of mantle, etc.)
(MELT-TIME)Mars Remaining Permafrost
 4 years
Macrolaser melting of Mars permafrost might be best accomplished by boring a network of “sinkholes”
(possibly for this purpose the superhot focus of the undiffused macrolaser beam). Then one might
systematically shine the (diffused) beam on large areas near “sinkholes”, to melt and drain these areas
(most draining below ground) into the “holes”. Returning the (diffused this time) laserbeam to train on
respective “sinkholes” (in a repetitive pattern), would result in water’s being directly evaporated from
each hole by the laser heat, such that water vapor should rise and mix itself into the incipient Martian
atmosphere. Precipitation from such – clearly enormous – quantity of water vapor, should lead to the
development of lakes and oceans inside existing Martian declivities, over generally the entire surface of
the planet (i.e., those areas of sufficiently low altitude)
Page 31
18.
FIRST STELLAR FLYBY PROBES (FIG’S 13, 14)
See (IAF – Gordon & Breach, ’67, NYC – Ref. 33, op. cit.) “The Alpha Centauri Probe,” being this author’s
instrumented or robotic starship to flyby our nearest stellar neighbor in about 17 years at one-quarter the
speed of light (Fig’s 13 and 14) – proposal based on a 60-mile-diameter-overall Solar Energy Handling
Satellite (considered very impressively large in those days).
Alpha Centauri probe would have been a spaceship (picture-taking probe) driven by (ion or particleaccelerator) rocketmotor supplied by the moving focus of a sunpumped, sun-orbiting, remote laser
system featuring a 60-mile-diameter aluminized mylar solar-collecting mirror (Fig. 4), assembled
(“bootstrapped”) in LEO, self-propelled to sun orbit.
The Probe’s 1-2 meter-diameter-objective (liquid) mirror camera, would itself have been an expandable,
liquid-optic unit, as the enormous acceleration level maximum of 780 g’s was proposed to get the Probe
up to speed ( ¼c), before the supply-beam spread appreciably (and of course no conventional optics
would survive these forces).
It was proposed that the supply macrolaser be placed in a solar orbit mainly to reduce the slewing or
pointing requirements.
The Probe would have weighed 14,000 lb; its acceleration period would have been 24 hours, during which
time it would have been accelerated 200 million miles into deep space.
The Probe would have
continuously monitored its position in the Probe plane normal to the supply base, and used onboard
microrockets to maintain itself square in the center.
Alpha Centauri Probe’s functions would have been entirely under control of its computers during
approach to the Alpha Centauri 3-star system: terminal speed of 45,000 mi/sec “would require precision
photography.” “Poses” would have been transmitted to earth by the probe’s expandable microwave
transceiver antenna system.
The particle-accelerator probe rocket’s performance is governed by the well-known rocket equation (Ref.
32, op. cit.)
M0
ML

e
u
v
 u


v  v
1
 e  1
2 



,
2
where,
M0 = total original rest-mass of rocketship and expellant
ML = payload mass, in lb.
u = desired spacecraft terminal velocity, in km/sec.
 = specific power ratio 
P
, where
Mp
P = power available for thrust, in kW,
Mp = mass of thrust-producing system onboard probe, in lb.
Page 32
 = desired acceleration time (or “boost”) period, in sec.
v = expellant exhaust velocity, km/sec
Product of specific power, , and acceleration time, , is approximately proportional –with nearly
equivalent proportionality factors- to both u2 and v2 for rockets with minimum
M0
ML
.
Then for Alpha Centauri probe,
 = 2.52× 107 kW/lb,
 = 24 hr.
u = v = ¼c =  =7 × 104 km/sec.
Total mass of the (unconsumed) thrust producing system, is taken (arbitrarily) at 200 lb. Expellant (for
acceleration by particle-accelerator rocketmotor) mass is nominally in the shape of a load-bearing, frozen,
concave-meniscus “cake”, tapering in cross-section out to the limb (Fig’s 13, 14).
FIG. 13. INTERSTELLAR SHIP IN ACCELERATION PHASE
Then terminal mass Me = 14,200 lb, payload mass ML = 14,000 lb, total initial mass M0 = (19.3)ML =
270,000 lb, and expellant mass = Mf = 255,800 lb.
Principle of Alpha Centauri Probe was to transcend “Chinese Rocketry” by supplying all energy to our
interstellar rocket via macrolaserbeam dispatched steadily (CW) from afar (from orbit of earth or sun).
Page 33
The energy would be provided by the sun itself, solar energy would have been used to pump the laser,
furnished by a very large collector mirror.
FIG. 14.
Electromagnetic Non-Self-Contained
5
Rocket Motor
3
(1) Payload (stoppable or unstoppable);
4
1
(2) High-temperature ionizable (ablatable)
7
disc;
6
7
2
(3) Insulator “Spike-Bed” to fix separation
between disc and electrode mesh;
(4) Electrode mesh;
(5) Envelope of exhaust plume;
(6) Direction of impinging laser beam from
distant orbital laser beampower source;
(7) Direction of rocket exhaust.
19. FIRST TWO-WAY INTERSTELLAR MANNED EXPEDITION
Using the principles introduced above, we may identify the following possible features of the first
manned round-trip interstellar expedition:
The target will be one of the two sun-like stars of the Alpha Centauri Group (nearest stars to the Solar
System), which are about 4.3 light-years distant.
20. “SEED” SEHS WEIGHT & SIZE FOR INTERPLANETARY/INTERSTELLAR USE
A “seed” solar energy handling spaceship (SEHS) package, will transfer itself into synchronous (22,240mi.-high) earth orbit after it is first lofted as a 200,000-lb payload into LEO by a common chemical
booster. Total “seed” weight of 200,000 lb comprises all (proportional) elements such that collector
mirror will be 5,290 ft.-diameter and precision (liquid) optical primary mirror will be 88 ft. diameter (Fig.
4).
5
6
7
2
10
3
9
4
8
1
16
11
13
14
12
15
FIG. 15. FABRICATION OF MACRO-OPTICS IN SPACE (BEGIN)
17
Page 34
Page 35
FIG. 15. FABRICATION OF MACROOPTICS IN SPACE (CONCLUDE)
(1) Subliming-propellant microrocket for attitude control; this rocket is embedded in a plastic-foam
section of mirror –or lens- boundary;
(2) Astronaut engaged in extending diameter of macro-optic;
(3) Plastic unit boundary section;
(4) Channel guide for spool carrying plastic sheet for circumferential optic augmentation;
(5) Escape rocket for Apollo Command module;
(6) Apollo Command module used for initial steps in fabrication of “Gigawatt Radiative Optical
Macroelectronic” (or interstellar power-beaming) system;
(7) S-IV-B stage of jettisoned Apollo ferry vehicle;
(8) Beam of macrolaser used to “bootstrap” augmentation of latter, via SCT (Spacelink Civilian
Transport) ferry vehicle;
(9) SCT ascending via near-microwave (from free electron laser) energy, ferrying load of building
material to astronauts;
(10) Plasma exhaust of SCT;
(11) Releasable joint between two boundary-unit sections;
(12) Body of giant solar-collecting mirror;
(13) Plastic-foam unit boundary section;
(14) Channel-follower, integral with spool;
(15) Old circumference of macro-optic;
(16) Pod carrying spool wound with sheet material for construction;
(17) Unrolled sheet material for macro-optic construction.
21.
“SEED” EXPONENTIAL GROWTH PERIOD FOR INTERSTELLAR APPLICATION
Use the system of the last paragraph to exponentially “bootstrap” up to low earth orbital altitude
construction materials supplied from earth, moon or other planet or satellite (whichever is most
convenient). “Seed” satellite will be augmented by astronauts (Fig. 15) at the exponential rate of 1/80 th
per day, requiring thereby 3.2 years to grow (assemble itself) into a 1640-mile-diameter, 53.5 × 1010-lbgwt., completed “permanent” SEHS unit adequate for launching (“stoppable”) 140,000-lb payloads to
Alpha Centauri. This “permanent” SEHS, to quadruple power, must first transfer itself to a ¼ th A.U. orbit of
the sun, to keep total mass down to 1/16th and overall diameter down to ¼th the equivalent-output unit in
GEO. Its primary (liquid) central mirror will be 1/60th of 1640 miles = 27-mi.dia. Transfer will be effected
by attached (dismountable) LREM’s (Laserpowered Remote Electricrocket Motors) fed from the central
macrolaser pumped by the 1640-mi.dia. associated solar-collector mirror.
Page 36
22.
ADVANCE STUDY OF THE TARGET STAR’S VICINITY
Astronomical study of the Alpha Centauri System, will be accomplished next using the 27-mi.di.
(diffraction-ltd.) optical primary of the sun-orbiting solar energy handling satellite as an astronomical
telescope. This mirror can resolve some 320 miles at Alpha Centauri distance. Main purpose will be to
identify raw materials (by photograph and spectrograph) for construction of an A.C.-orbiting SEHS as
accelerator for the return spaceship (operating also at c/4, which is nominal speed both for stellar
approach and earth-return).
23.
TOTAL POWER DELIVERED BY THE SUN-ORBITING “PERMANENT” SEHS
Power delivered by the 1640-mi.-dia.-collector “permanent” sun-orbital accelerator in ¼th A.U. solar orbit,
will be 1.19 × 1014 kW (Compare to world energy demand predicted for 1990 by Zarem-Erway -Ref. 44- of
7.35 × 1011 kW, which is about half a percent of the former –interstellar propulsion- requirement). Note
that given power delivered by the 1640-mi.-dia. sun-orbiting system, purportedly accounts for all losses
and is estimated such as to give 1% overall efficiency (in the entire propulsion process) at the associated
rocketmotor output. Dr. Bertrand Oliver of course is amply exonerated: MACROLASER power is 200 times
total earth consumption (about), just to lob and stop astronauts and assembleable packages for a returnto-earth spaceship (out to Alpha Centauri) in about 700,000 launches! This doesn’t account for all the
energy to start & stop astronauts on the return. But since stars provide all the power, who cares?
24.
RATIO OF GWT. OF ACCELERATOR PKG. TO GWT. OF ASSOCIATED PAYLOAD
Mass of each individual “stoppable” unmanned (freight) or manned (passenger) payload dispatched to
Alpha Centauri, is set arbitrarily at 140,000 lb. The 1640-mile-diameter “permanent” accelerator orbiting
the sun, will launch a succession of (Fig. 16) “stoppable” 140,000-lb payloads to Alpha Centauri, each at a
cruise velocity of c/4 (or rocket specific impulse of 7,500,000 seconds respectively, where terminal
velocity approximately equals exhaust velocity). Ratio of gross weight of the sun-orbiting “permanent”
accelerator package, to gross weight of each dispatched flyby decelerator + payload package mass, is
about 2,000-to-1. Ratio of each flyby decelerator package mass itself to its own associated “stoppable”
payload mass, is also 2,000-to-1.
25.
INTERSTELLAR SEH SPACESHIP ASSEMBLY EN ROUTE BY ROBOTS (AUTOMATION)
Each individual “stoppable” payload’s associated (flyby) decelerator (collector-and-precision liquid optics:
laser/”pump”/primary/secondary/focus-control), is assembled by robots (by automation) in zero-g
(unpowered trajectory, i.e., while “coasting”) en route to Alpha Centauri at c/4.
Page 37
FIG. 16. Stoppable Ship in Braking Approach to Target Star:
(1) Target star (e.g., Alpha Centauri A);
(7) focused
(2) ,(3) Planets of approached star;
stellar
energy;
(4) Primary mirror of flyby braking system;
(8) Collector mirror;
(5) Diverged laser beam illuminating primary mirror;
(9) Perforation
(6) laser/diverging lens assembly;
in
collector mirror;
(10) Interplanetary space
of target star;
(11) Laser beam focused
2
by primary mirror;
1
1
(12) Braking
rocket
exhaust;
(13) Braking rocket;
10
(14) 140,000-lb
16
17
4
5
stoppable payload;
7
(15) ,(16)
6
3
8
already
(17) Direction of motion
12
11
Payloads,
orbiting target star;
9
5
Stopped
of
13
flyby
braking
system.
15
5
26.
14
ALPHA-CENTAURI-CIRCLING “COCOONED” LAUNCH PACKAGE FOR ASTRONAUT
RETURN
No “flyby decelerator” system need be included in the astronaut-return “cocoon” package dispatched
from Alpha Centauri, since the original, sun-orbiting (1640-mile-diameter) “accelerator,” may now be
used as an (astronaut-return) decelerator (i.e., by supplying solar-macrolaser energy to drive the particleaccelerator retrorocket, to slow the astronauts’ return-spaceship from c/4 to earth-orbital velocity). Thus
A.C.-orbiting, “permanent accelerator” diameter and therefore mass, needn’t be nearly as great as that of
the 1640-mi.-dia. sun-orbiting one.
Additionally, included in the “Launch-from-Alpha Centauri” (L.A.C.) rocket package, will be (only):
Page 38
(I)
a “cocoon”, where the cocoon consists just of
(1) the 140,000-lb (manned) payload, plus
(2) the payload-stopping rocketmotor’s local (secondary) collector-mirror (integral with
the payload), plus
(3) the payload-stopping (particle-accelerator) rocketmotor itself (also integral with the
payload), plus
(4) a total weight of payload-stopping expellant equal to 19.3 times the sum of items
(1), (2) and (3), where one notes that the cocoon is able to exclude the vast
deceleration primary flyby solar-collecting mirror normally required;
PLUS (II)
the cocoon-accelerating rocketmotor’s local (or secondary) collector-mirror;
PLUS (III)
the cocoon-accelerating (particle-accelerator) rocketmotor itself;
PLUS (IV)
a total weight of expellant equal to 19.3 times the sum of items (I), (II), and (III) above (to
launch the “cocoon” at c/4 toward the Solar System).
27.
MINIMUM SIZE OF “LAUNCH-FROM ALPHA CENTAURI” (L.A.C.) ROCKET PACKAGE
ENCLOSING EARTH-ASTRONAUT-RETURN “COCOON” (STOPPABLE WITH SUNORBITING SEHS)
Weight of the L.A.C. rocket package (from the last paragraph above), may be some 52 × 106 lb. (only).
Then by (rocket-formula) proportions above, “primary” collector diameter of L.A.C. earth-return
accelerator system to accelerate cocoon to c/4, will be only some 680 miles overall (at 1% efficiency, a
2000-to-1 mass ratio, and presupposing a sun-equivalent Alpha Centauri star).
Mass, then, of the 680-mile-diameter-collector, A.C.-orbiting “permanent” accelerator system (or SEHS)
to launch to earth a 52 × 106-lb (takeoff mass) spaceship enclosing a “cocooned” (manned-payloadstopping) package (comprised of payload + earth-approach secondary collector mirror + earth-approach
deceleration rocketmotor + expellant), will be (only) some 93 billion pounds.
28.
TOTAL QUANTITY OF “STOPPABLE” PAYLOADS TO BE SENT TO A.C. IF ALL A.C.CIRCLING “SEHS ACCELERATOR” MATERIAL IS TO BE SENT FROM THE SOLAR
SYSTEM
Number of “stoppable” payloads at 140,000 lb each, required to be sent to Alpha Centauri from the Solar
System for the astronauts to assemble (on the spot) into a 93 billion pound (Centauri-circling) earthreturn “permanent accelerator,” would be 660,000 loads.
Page 39
29.
POSSIBLE MINIMUM QUANTITY TOTAL “STOPPABLE” A.C.-GOING TOOL-&SUPPLIES PAYLOADS IF EARTH-RETURN SEHS IS MANUFACTURED (NOT,
ASSEMBLED ONLY) IN ALPHA CENTAURI ORBIT
Using local (A.C.) raw materials, and sending astronauts with tools, maps, essential supplies and
assemblies, to manufacture their own return accelerator (rather than merely assemble it by first
collecting 660,000 previously “stopped” payloads orbiting an A.C. sun), might require that only some 300
loads (pure guess) at 140,000 lb. each, be sent to A.C.
Progressively more refined information would be gathered and processed by the astronauts in (nearly 17
years of) real-time as they approached Alpha Centauri, by the expedient of reversing their diffractionlimited (self-assembled) deceleration primary-mirror during their near-17-year voyage, and using it for an
astronomical telescope to study the A.C. vicinity. Thus, a two-way flow of information with earth libraries
& computing facilities could be maintained for most of the voyage, to constantly update and aid
astronauts in manufacturing planning.
Astronauts would spend approximately 17.2 years in transit to Alpha Centauri, 3% of that time in
acceleration-deceleration (artificial) 1 g, the balance in centrifugal-rotation (artificial) 1 g.
30.
GALLIUM COST AND EXPECTED COST REDUCTION
The “permanent accelerator” astronomical primary mirror orbiting the sun (and all other mirrors) would
be plated with a liquid-gallium thickness of 0.003-inch (plated on liquid-plastic). If one mile in diameter
and using gallium valued at, say, $400 per ounce, 336 tons of gallium costing $4.3 billion (1994 FRN)
would be required. Creation of new wealth-money (vs. present debt-money) and grossly improved metalextraction technology (from earth, moon, other planets and their moons, asteroids, etc.) respectively by
orbital solar macrolasers, should succeed in driving gallium “price” sufficiently “low” (if we still need to be
concerned about such things!) Thus it appears predictable sufficient gallium –even for the 27-milediameter primary- should be forthcoming on acceptable terms.
c/4 cruise speed of interstellar rockets gives relativistic mass increase of 3%. With SEHS “accelerators”
orbiting both our sun and an Alpha Centauri sun, improvements in cruise speed should come rapidly,
comparable to a game of progressive interstellar hardball. It should ultimately be possible to approach
lightspeed for this one star-to-star span, while other stars evidently could be added, exponentially as-wego, to available, near-lightspeed, star-to-star spans (at least in our corner of the galaxy).
MACROLASERS for scientific research (not just astronomy, propulsion and space-travel) imply of course a
virtual explosion in scientific knowledge, no doubt such as to rapidly render the interstellar
propulsion/exploration system expanded on here (mostly from this author’s earlier publications) totally
obsolete.
Page 40
31.
RULES, ARBITRARY AND NOT
Rocket exhaust (“expellant”) velocity, must be set at approximately desired spacecraft terminal velocity,
to earn best (lowest) ratio of initial spacecraft mass M0 to payload mass ML.
No A.C.-going-or-return spaceship acceleration or deceleration would – or would need to – significantly
exceed 1 g at any time.
Each solar/stellar collector would be fabricated as a 0.0007-inch-thick, aluminized-mylar, closed, inflated,
spherical segment of one base, aluminized on the inside concave surface with overall average density 2.0.
Stellar collector efficiency = .................................................................................................................... 0.9.
(Starpumped, liquid-end-mirror) laser efficiency = ................................................................................ 0.1.
Primary and Secondary Mirror Efficiency = ........................................................................................... 0.9.
Laserpowered Remote Electricrocket Motor (LREM) “secondary” (local) collector-disc efficiency = .. 0.9.
Rocketmotor thrust-producing efficiency = ........................................................................................... 0.2.
Overall (MACROLASER-energy-supplied) nonself-contained propulsion process efficiency is ... about 1%.
Blastoff-to-Cutoff, and Retrorocket-to-“Halt” stages respectively require 88.5 days = ................. 0.24 yr.
Distances Blastoff-to-Cutoff and Retrorocket-to-“Halt” respectively are ...................... 0.18 trillion miles.
In a later study, power levels for acceleration to and deceleration from c/4, must be increased to account
for “redshift” of supply source radiation toward redder (less energetic, longer wavelength) photons.
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